A hybrid wing layout aircraft

Abstract

The present invention discloses a hybrid wing layout aircraft, belonging to the field of aviation equipment technology, which includes: a fuselage, wings, control surfaces, and a power unit for providing flight power and control, characterized in that: the power unit includes a left tilting power mechanism arranged on the left side of the fuselage, a right tilting power mechanism symmetrically arranged on the right side, and an adjustment mechanism that enables the two tilting power mechanisms to tilt synchronously, the left tilting power mechanism and the right tilting power mechanism are used to provide power and flight control during vertical flight and horizontal flight, the adjustment mechanism is used for the conversion of vertical flight and horizontal flight modes, and the control surface is used for roll and pitch control during horizontal flight. The aircraft layout mode of the present invention, as well as the setting of synchronous tilting and left and right tilting power mechanisms, successfully solve the problem that the rotor wing of the hybrid layout aircraft is difficult to control under aerodynamic interference.

Description

Technical Field

[0001] The present invention relates to a tilt-rotor twin-rotor aircraft, in particular to a hybrid-wing layout aircraft, and belongs to the technical field of aviation equipment. Background Art

[0002] At present, most small and light unmanned aerial vehicles are based on multi-rotor layout configurations. They are flexible and maneuverable during flight and have been widely used in production and life. However, the problem with the multi-rotor configuration is that, due to the fact that its lift source is only provided by the rotor, there is a large loss of power when flying at high speeds, and the resistance is large, and the maximum flight speed is relatively low. Due to the power loss and the large resistance when flying at high speeds, the endurance time is also relatively short. The main principle is:

[0003] a. When a multi-rotor aircraft flies forward horizontally, the front flow has an inclined angle with the propeller disc. The flow leads to low propeller propulsion efficiency, severe aerodynamic interference, large resistance, and difficulty in increasing the maximum flight speed.

[0004] b. Compared with fixed-wing aircraft, multi-rotor aircraft have different principles when flying forward horizontally. The horizontal component of the propeller thrust of the multi-rotor aircraft is used to overcome the greater resistance, and the vertical component is used to overcome its own gravity. Fixed-wing aircraft use the aerodynamic lift of the fixed wing to overcome gravity, and the propeller thrust only overcomes the lower resistance. Therefore, the energy consumption of multi-rotor aircraft is much higher than that of fixed-wing aircraft at the same flight speed, resulting in high energy consumption and short flight time.

[0005] However, tilt-rotor aircraft have the ability to take off and land vertically and fly horizontally, combining the advantages of rotorcraft taking off and landing vertically and fixed-wing aircraft flying at high speed and for a long flight time. Therefore, they are widely favored and attract the attention of the industry. However, the existing tilt-rotor aircraft have the following problems: heavy weight, sparse and not compact structure, and complex tilt mechanism. The main reasons are:

[0006] a. The layout of existing tilt-rotor aircraft is mostly normal layout, and the tilt-rotor or propeller is mostly arranged at the wing tip or in front of the wing. In order to reduce the aerodynamic coupling interference between the propeller and the wing, the layout is sparse and not compact enough;

[0007] b. Existing tilt-rotor aircraft generally have a variable pitch flapping mechanism to provide sufficient pitch control torque, or use a tilt-rotor tri-rotor configuration by adding a propeller at the front or tail to provide sufficient pitch control torque.

[0008] For example, CN103192990A discloses a flying wing layout aircraft, which includes: a fuselage and wings, ailerons located at the outer section of the wings, an elevator located at the rear of the fuselage, a double vertical tail located at the rear of the fuselage, a sleeve that crosses the fuselage, two front propeller pods located on both sides of the fuselage and in front of the wings, which are fixedly connected to the sleeve, and a rear propeller located below the double vertical tail, wherein the power tilt of the two front propellers is realized by the rotation of the sleeve, and when flying horizontally, the front propeller disc is vertical, providing a The front propeller has a pull force in front. During vertical takeoff and landing, the propeller disk of the front propeller is horizontal with the plane of the aircraft, providing an upward thrust. This thrust and the lift generated by the rear propeller jointly bear the weight of the aircraft. During short takeoff and landing, the deflection state of the front propeller is between the deflection state during vertical takeoff and landing and horizontal forward flight. The thrust vector points forward and upward, and the thrust of the rear propeller remains upward. When the taxiing speed reaches a certain level, the lift generated by the front propeller, the rear propeller and the integrated fuselage and wings is greater than the weight of the aircraft, and the aircraft can take off. The layout of this aircraft is to set two front propellers in front of the wing. In order to avoid aerodynamic interference of the propeller on the wing, the wing is designed with a large sweep angle, and two rear propellers are set at the tail. Therefore, its structure is difficult to achieve miniaturization.

[0009] CN116573179A discloses a box-type layout tilt-rotor micro unmanned aerial vehicle, characterized in that it includes a fuselage, which includes a fuselage shell, which is a complete integrated shell, and the fuselage shell is connected to a box-type wing and a tilt-rotor, the box-type wing includes a front wing and a rear wing, the front wing and the rear wing are connected by a wing tip plate to form a box-shaped layout; the tilt-rotor is fixedly connected to both sides of the middle part of the fuselage shell by connecting rods.

[0010] A common problem in CN116573179A and CN103192990A is that when the front propeller switches between the vertical take-off and landing state and the horizontal flight state, since the propeller disc has a certain turning radius, it is impossible to ensure that its pulling force line and the center of gravity of the aircraft remain coincident in both flight states. Figure 1 , 2 , the figure shows the positional relationship between the propeller tension line and the center of gravity of the existing aircraft tilt-rotor horizontal flight mode and vertical flight mode, where point a is the center of gravity of the aircraft, d is the rotation axis of the tilt motor, that is, the tilt center, and c is the tension line. In the vertical flight mode, the tension line c is vertically upward, and the center of gravity a is located on the tension line, but when in the horizontal flight mode, its tension line c is located in the horizontal direction of the rotor center. At this time, the tension line cannot remain coincident. Therefore, in different flight states, a complex flapping and pitch-changing mechanism is required to achieve stable pitch control, or control compensation is performed through the rotor or rudder at the tail. However, no matter which method is used, on the one hand, the structure will be complicated, and on the other hand, the efficiency will be reduced.

[0011] Another problem with CN116573179A is that it adopts a negatively staggered wing but uses an upward-pull propeller arrangement. In the vertical flight mode, the front wing is easily subject to the strong aerodynamic impact interference of the slipstream behind the propeller, resulting in large propeller-wing coupling interference.

[0012] Based on the above shortcomings of tilt-rotor aircraft, a dual-tilt-rotor and fixed-wing hybrid layout aircraft that can overcome the above shortcomings and ensure that the thrust line and the center of gravity of the aircraft can coincide in both vertical flight mode and horizontal flight mode and have higher efficiency and higher speed has become a goal pursued by technical personnel in this field. Summary of the invention

[0013] The primary purpose of the present invention is to solve the problem that it is difficult for a tiltrotor aircraft to ensure the relative position requirements of the tension line or thrust line and the center of gravity in different flight modes, that is, it is impossible to ensure that the tension line or thrust line and the center of gravity coincide in horizontal flight and vertical flight states.

[0014] The second purpose of the present invention is to solve the problem that the existing aircraft adopts a negatively staggered wing combined with an upward-pull propeller arrangement, which causes the front wing to be easily subject to the strong aerodynamic impact interference of the slipstream behind the propeller, resulting in large propeller-wing coupling interference.

[0015] To achieve the above-mentioned purpose of the invention, the technical solution of the present invention is: a hybrid wing layout aircraft, which includes: a fuselage, wings, control surfaces and a power unit, characterized in that: the power unit is used to provide power and flight control during vertical flight and horizontal flight; the power unit is located between the front wing and the rear wing of the wing, and the power unit includes two independently tiltable left-turn power mechanisms and right-turn power mechanisms symmetrically arranged on both sides of the fuselage, and an adjustment mechanism for installing the left-turn power mechanism and the right-turn power mechanism and being able to adjust the left-turn power mechanism and the right-turn power mechanism to ensure that the tension or thrust line coincides with the center of gravity in both vertical flight and horizontal flight.

[0016] As a preferred embodiment, the adjustment mechanism includes a tilt servo mounting seat installed on the load-bearing structure of the fuselage, a tilt servo fixedly installed on the tilt servo mounting seat, and a bracket hinged on the tilt servo mounting seat, the hinge axis of the bracket is directly connected to the rotating shaft of the tilt servo, and can rotate from a vertical position to a forward horizontal position around the center of the rotating shaft of the tilt servo under the drive of the tilt servo; the left tilt power mechanism and the right tilt power mechanism are arranged at both ends of the bracket, and when the bracket is in the vertical position, the center of gravity a of the aircraft and the tilt center d of the left tilt power mechanism and the right tilt power mechanism are located on the same vertical plane, and in the horizontal flight mode, when the bracket is rotated to the horizontal position, the center of gravity a of the aircraft and the tilt center d of the left tilt power mechanism and the right tilt power mechanism are located in the same horizontal plane, and in the horizontal flight mode, the vertical flight mode and the rotation process of the adjustment mechanism, the tension / thrust line c can always pass through the center of gravity a of the aircraft.

[0017] As a preferred solution, the bracket includes a load-bearing crossbeam, a rotating bracket arranged in the middle of the load-bearing crossbeam, the lower part of the rotating bracket is hinged to the two sides of the tilting servo mounting seat, and the left tilting power mechanism and the right tilting power mechanism are arranged at the two ends of the load-bearing crossbeam;

[0018] The left rotating support and the right rotating support are connected to the rotating shafts of the left and right servos through the shaft holes at the upper ends, and a left tilting shaft base and a right tilting shaft base for supporting the ends of the rotating shafts of the left and right servos are respectively provided at the two ends of the load-bearing crossbeam.

[0019] As a preferred solution, the center of the hinge axis of the rotating bracket of the bracket and the tilt-servo mounting seat, that is, the center b of the tilt-servo rotating axis coincides with the center of gravity a of the aircraft.

[0020] As a preferred solution, the left tilting power mechanism includes a left propeller, a left motor, a left rotating support and a left servo, and the right tilting power mechanism includes a right propeller, a right motor, a right rotating support and a right servo; the left servo and the right servo are symmetrically arranged on the load-bearing crossbeam, the left motor is installed on the left rotating support, the left propeller is installed on the left motor, the right motor is installed on the right rotating support, and the right propeller is installed on the right motor; the left rotating support and the right rotating support are respectively connected to the left servo and the right servo, and can rotate around the center of the rotating shaft of the left servo and the right servo parallel to the load-bearing crossbeam under the drive of the left servo and the right servo.

[0021] As a preferred solution, the left servo and the right servo are directly mounted on the two ends of the load-bearing beam by means of left servo buckles and right servo buckles or bonding. This connection method has a simple structure, reduces the number of installation parts, can reduce the weight of the aircraft, and increase the flight time.

[0022] As a preferred solution, the left and right servos can also be installed in another way: the left and right servos are respectively installed under the load-bearing crossbeam through a left servo mounting seat and a right servo mounting seat symmetrically arranged on the left and right sides of the load-bearing crossbeam; a left tilting shaft base and a right tilting shaft base are symmetrically arranged on the same side of both ends of the load-bearing crossbeam, and the left rotating support and the right rotating support can be rotatably installed on the left tilting shaft base and the right tilting shaft base, respectively, and the left and right servos drive the left rotating support and the right rotating support to rotate around the axis of the left tilting shaft base and the right tilting shaft base. This installation method of the left and right servos has a more stable structure and better overall strength.

[0023] As a preferred solution, in order to reduce the assembly of parts, the rotating bracket is fixedly connected to the load-bearing beam at the upper end or is integrally formed on the load-bearing beam.

[0024] As a preferred solution, the bracket is also provided with an electronic speed regulator and a power supply and communication line for adjusting the rotation speed of the left motor and the right motor. This optimized design makes the structure more compact and the space utilization rate better.

[0025] In order to achieve the second purpose of the present invention and reduce the coupling interference between the propeller and the wing, as a preferred solution, the front wing and the rear wing are arranged in a negative staggered manner with the front wing being low and the rear wing being high, and the left propeller and the right propeller are arranged in a push-down manner in the vertical flight mode, forming a compact layout with minimal aerodynamic interference in the vertical flight mode.

[0026] In order to improve the overall strength of the wing and make it more suitable for reducing the weight of the wing, and at the same time protect the power unit from hurting the operator, as a preferred solution: the wing includes a front wing, a rear wing, a left end wing, and a right end wing, the fuselage is directly connected to the front wing, the rear wing is connected to the front wing through the left end wing and the right end wing, and the rear wing is connected to the load-bearing structure of the fuselage through the vertical tail; the two ends of the front wing and the rear wing are respectively connected to the left end wing and the right end wing.

[0027] As a preferred solution, the trailing edge of the rear wing is provided with a left elevon and a right elevon for controlling the roll angle and pitch angle of the aircraft during level flight, and a servo 1 and a servo 2 are installed in the vertical tail for driving the left elevon and the right elevon to deflect.

[0028] As a preferred embodiment, it further includes an autonomous control system, which includes a flight control subsystem, a mission control subsystem, a camera and a lithium battery. The flight control subsystem and the mission control subsystem are installed at the rear of the fuselage's load-bearing structure, and the camera and the lithium battery are installed at the front of the fuselage's load-bearing structure.

[0029] As a preferred solution, in which: in the vertical flight mode, the adjustment mechanism is rotated to a vertical position; by differentiating the pulling force of the left propeller of the left tilt power mechanism and the right propeller of the right tilt power mechanism, a rolling moment can be generated to achieve rolling motion; by simultaneously increasing or decreasing the tilt angles of the left propeller and the right propeller, a pitch moment can be generated to achieve pitch motion; by differentiating the tilt angles of the left propeller and the right propeller, a yaw moment can be generated to achieve yaw motion.

[0030] As a preferred solution, in which: in the fixed-wing horizontal flight mode, the forward rotating adjustment mechanism is turned to a horizontal position, and the deflection angles of the left elevon and the right elevon are differentiated, so as to generate a rolling moment; at the same time, the deflection angles of the left elevon and the right elevon are increased or decreased, so as to generate a pitching moment; and the pulling force of the left propeller of the left-tilt power mechanism and the right propeller of the right-tilt power mechanism are differentiated, so as to generate a yaw moment.

[0031] The present invention arranges left and right tilting power mechanisms between the front wing and the rear wing of the aircraft, and adds an adjustment mechanism for installing the left and right tilting power mechanisms, which can adjust the left and right tilting power mechanisms to ensure that the tension or thrust line coincides with the center of gravity in both vertical flight and horizontal flight. Specifically, a tilt servo mounting seat and a tilt servo mounted on the fuselage, and a bracket for supporting the left and right tilting power mechanisms are added, and the bracket can be rotatably mounted on the tilt servo mounting seat, so that the bracket can rotate between a vertical position and a horizontal position with the left and right tilting power mechanisms. When the bracket is in the vertical position, that is, at this time, the tension line or thrust line of the left and right tilting power mechanisms coincides with the center of gravity of the aircraft. When the aircraft is in the horizontal flight mode, the bracket is rotated forward to the horizontal plane position where the center of gravity is located, so that the tension line or thrust line of the left and right tilting power mechanisms coincides with the center of gravity of the aircraft.

[0032] Furthermore, the present invention designs the bracket as a load-bearing beam and a rotating bracket arranged in the middle of the load-bearing beam, the left and right tilting power mechanisms are arranged at both ends of the load-bearing beam, and the lower part of the rotating bracket is hinged to the two sides of the tilting servo mounting seat, so that the bracket can drive the left and right tilting power mechanisms to rotate. In particular, it is preferred to design the hinge center position of the rotating bracket to be at the center of gravity of the aircraft, so that when the bracket is in the vertical position and the horizontal position, the tension line or the thrust line can be accurately coincident with the center.

[0033] Furthermore, the present invention designs the wings to be in a negative staggered layout and designs the propellers of the power unit to be in a push-down type, so that the propeller-wing coupling interference is minimized in the vertical flight mode and there is no propeller-wing coupling interference during the tilting process.

[0034] Furthermore, the present invention designs the wing into a layout in which the front and rear wings are connected by the end wings, and combines the power of dual propellers, so that the micro-aircraft can not only take off and land vertically and hover like a multi-rotor aircraft, but also fly forward horizontally like a fixed-wing aircraft, with a fast flight speed, high efficiency and low energy consumption; this layout has a natural shielding and protection effect on the two propellers, which can prevent the propellers from harming the operator during use, and also protect the propellers from loss when the aircraft collides, allowing the aircraft to collide to a certain extent; this layout has the advantages of compact layout and small geometric dimensions; when the aircraft flips over and falls upside down on the ground, with the support of the connected wings, the propellers can be rotated in the opposite direction to achieve forward rolling, and the aircraft can be adjusted to a forward posture to take off again.

[0035] Furthermore, the aircraft of the present invention can generate the rolling moment, pitching moment or yaw moment required for rolling, pitching and yaw motion in vertical flight and horizontal flight by changing the tilt angle of the adjustment mechanism, and the tilt angle and rotation speed of the left tilting power mechanism and the right tilting power mechanism. The rolling moment and pitching moment required for rolling and pitching in horizontal flight can be generated by changing the deflection angle of the left elevon and the right elevon. BRIEF DESCRIPTION OF THE DRAWINGS

[0036] Figure 1 , 2 This is a schematic diagram of the relationship between the tension line and the center of gravity of the existing tilt-rotor aircraft in horizontal flight mode and vertical flight mode; this existing technology should be a tension line, rotor-type pull-up arrangement

[0037] Figure 3 It is a schematic diagram of the three-dimensional structure of a hybrid wing layout aircraft in a vertical flight mode according to the present invention;

[0038] Figure 4 It is a schematic diagram of the three-dimensional structure of a hybrid wing layout aircraft in a horizontal flight mode according to the present invention;

[0039] Figure 5 , 6 It is a three-dimensional structural schematic diagram and a three-dimensional exploded schematic diagram of the power unit of the first embodiment of the present invention;

[0040] Figure 7 , 8 It is a schematic diagram of the relationship between the thrust line and the center of gravity of the power unit described in the present invention in the horizontal flight mode and the vertical flight mode.

[0041] Fig. 9 , 10 11 is a schematic diagram of the aircraft of the present invention adopting a negative staggered box wing layout combined with a push-down propeller layout, a schematic diagram of the aerodynamic coupling state during vertical flight, and a schematic diagram of the aerodynamic interference state during the tilting process;

[0042] Fig.12 , 13 14 is a schematic diagram of the aircraft of the present invention adopting an orthogonal staggered wing layout combined with a push-down propeller layout, a schematic diagram of the aerodynamic coupling state during vertical flight, and a schematic diagram of the aerodynamic interference state during the tilting process;

[0043] Fig.15 , 16 17 is a schematic diagram of the aircraft of the present invention adopting a negative staggered wing layout combined with a pull-up propeller layout, a schematic diagram of the aerodynamic coupling state during vertical flight, and a schematic diagram of the aerodynamic interference state during the tilting process;

[0044] Fig.18 , 19 20 is a schematic diagram of the aircraft of the present invention adopting an orthogonal staggered wing layout combined with an upward-pull propeller layout, a schematic diagram of the aerodynamic coupling state during vertical flight, and a schematic diagram of the aerodynamic interference state during the tilting process;

[0045] Fig.21 , 22 It is a schematic diagram of the forces acting on the propeller and the control surface of the hybrid wing layout aircraft during flight control;

[0046] Fig.23 , 24 It is a three-dimensional structural schematic diagram and a three-dimensional exploded schematic diagram of the power unit of the second embodiment of the present invention.

[0047] Description of the reference numerals: fuselage 1, load-bearing structure 11, nose cover 12, wing 2, front wing 21, rear wing 22, left end wing 23, right end wing 24, vertical tail 25, control surface 3, left elevator 31, right elevator 32, steering gear 1 33, steering gear 2 34; power unit 4, left tilting power mechanism 41, left propeller 411; left motor 412; left rotating support 413, left steering gear 414; right tilting power mechanism 42, right propeller 421, right motor 422, right rotating support 423, right steering gear 424 ; Adjustment mechanism 43, tilt servo 431, tilt servo mounting seat 432, left servo buckle 4321, right servo buckle 4322, bracket 433, load-bearing beam 4331, rotating bracket 4332, left servo mounting seat 4333, right servo mounting seat 4334, electronic speed regulator 4335, power supply and communication line 4336, left tilt shaft base 4337, right tilt shaft base 4338; autonomous control system 5, flight control subsystem 51, mission control subsystem 52, camera 53, lithium battery 54. DETAILED DESCRIPTION

[0048] The preferred embodiments of the present invention are described in detail below with reference to the accompanying drawings, but the preferred embodiments cannot be used to limit the protection scope of the present invention.

[0049] Example 1, see Figure 3-6 , the figure shows a hybrid wing layout aircraft according to the present invention, which includes: a fuselage 1, wings 2, control surfaces 3 and a power unit 4, the fuselage 1 includes a load-bearing structure 11 for installing various components and a nose cover 12 arranged at the head of the fuselage, the wings 2 include a front wing 21 and a rear wing 22, wherein: the power unit 4 is located between the front wing 21 and the rear wing 22, the power unit 4 is arranged on the load-bearing structure 11 of the fuselage 1, and the power unit 4 is used to provide power and flight control during hovering vertical flight and horizontal flight; the power unit 4 includes a left tilting power mechanism 41 and a right tilting power mechanism 42, the left The tilting power mechanism 41 and the right tilting power mechanism 42 are symmetrically arranged on both sides of the fuselage 1, and can be tilted independently. The power unit 4 also includes an adjustment mechanism 43, which is installed on the load-bearing structure 11 of the fuselage 1. The left tilting power mechanism 41 and the right tilting power mechanism 42 are installed on the adjustment mechanism 43, and can be tilted synchronously with the adjustment mechanism 43. Therefore, the left tilting power mechanism 41 and the right tilting power mechanism 42 can be adjusted to ensure that the tension line / thrust line c coincides with the center of gravity a of the aircraft in both the hovering vertical flight state and the horizontal flight state, as shown in FIG. Figure 7 , 8 shown.

[0050] See also Figure 5-6, the figure shows a specific structure of the power unit 4, which includes: a left tilt power mechanism 41, a right tilt power mechanism 42 and an adjustment mechanism 43, wherein the adjustment mechanism 43 includes a tilt servo mounting seat 432 fixed on the load-bearing structure 11 of the fuselage 1, and a tilt servo 431 is installed on the servo mounting seat 432. In this embodiment, the tilt servo mounting seat 432 is designed to be bench-shaped, with four pillars and a top plate, and the tilt servo 431 is installed under the top plate. Obviously, this is a better way, and it can be replaced by other equivalent bracket structures. A bracket 433 is hinged on the tilt servo mounting base 432, and the rotating shaft of the tilt servo 431 is directly connected to the hinge axis of the bracket 433, specifically, it is connected to the hinge axis of the rotating bracket 4332 of the bracket 433 (described in detail later), and is used to drive the bracket 433 to rotate around the center of the tilt servo rotating shaft from a vertical position (set to 0 degree position) to a forward horizontal position (90 degree position), thereby driving the left tilt power mechanism 41 and the right tilt power mechanism 42 to tilt synchronously. It can rotate up to 120 degrees at most. As the best embodiment, when the bracket 433 is in a vertical position, the center of gravity a of the aircraft and the tilt center d of the left tilt power mechanism 41 and the right tilt power mechanism 42 are located in the same plumb plane. In the vertical flight mode, it can be ensured that the tension line / thrust line c coincides with the center of gravity a of the aircraft (that is, in the same plumb plane). As the best implementation method, the center of gravity a of the aircraft and the center of rotation b of the tilt servo (the center of the hinge axis of the bracket 433) coincide. In the horizontal flight mode, the bracket 433 is rotated forward to a horizontal position. At this time, the tilt center d of the left tilt power mechanism 41 and the right tilt power mechanism 42 can also pass through the center of gravity a of the aircraft. Obviously, if the center of gravity a of the aircraft is offset a little above or below the center of the hinge axis of the bracket 433, the rotation position of the bracket 433 is adjusted to the horizontal plane where the center of gravity is located, and the tension line / thrust line c also coincides with the center of gravity a of the aircraft during horizontal flight.

[0051] In this embodiment, the bracket 433 includes a load-bearing crossbeam 4331 and a rotating bracket 4332. The rotating bracket 4332 has two hinged arms, the lower ends of which are hinged to the two sides of the tilting servo mounting seat 432. The middle position of the load-bearing crossbeam 4331 is horizontally installed on the upper part of the rotating bracket 4332. The horizontal position refers to the horizontal direction perpendicular to the fuselage axis. The left tilting power mechanism 41 and the right tilting power mechanism 42 are symmetrically arranged at both ends of the load-bearing crossbeam 4331. It should be noted that in this embodiment, the two hinged arms of the rotating bracket 4332 are directly formed at the middle position of the load-bearing crossbeam 4331 at one time, so that the structure is more concise. It can obviously also be an independent component, with the upper end connected to the middle of the load-bearing crossbeam 4331 and the lower end hinged to the tilting servo mounting seat 432.

[0052] See also Figure 5 , 6The left tilting power mechanism 41 includes a left propeller 411, a left motor 412, a left rotating support 413 and a left servo 414, and the right tilting power mechanism 42 includes a right propeller 421, a right motor 422, a right rotating support 423 and a right servo 424; the left servo 414 and the right servo 424 are symmetrically arranged at the two ends of the load-bearing crossbeam 4331, the left motor 412 is installed on the left rotating support 413, the left propeller 411 is installed on the left motor 412, the right motor 422 is installed on the right rotating support 423, and the right propeller 421 is installed on the right motor 422; the left rotating support 413 is connected to the left servo 414 perpendicularly in the direction of the load-bearing crossbeam 4331, and can rotate around the center of the rotation axis of the left servo 414 parallel to the load-bearing crossbeam 4331 under the drive of the left servo 414, so as to change the tilt angle of the left propeller 411. Similarly, the right rotating support 423 is connected to the right servo 424 perpendicularly in the direction of the load-bearing crossbeam 4331, and can rotate around the center of the rotation axis of the right servo 424 parallel to the load-bearing crossbeam 4331 under the drive of the right servo 424, so as to change the tilt angle of the right propeller 421.

[0053] In this embodiment, the left servo 414 and the right servo 424 are directly mounted on the load-bearing crossbeam 4331 through the left servo buckle 4321 and the right servo buckle 4322, and can also be fixed on the load-bearing crossbeam 4331 by bonding, welding, etc. This structure makes the installation of the left and right servos simpler and reduces the weight of the aircraft.

[0054] In order to improve the connection strength between the left and right servos and the left and right rotating supports, a left-tilted rotating shaft base 4337 and a right-tilted rotating shaft base 4338 for supporting the rotating shafts of the left servo 423 and the right servo 424 are respectively provided at both ends of the load-bearing beam 4331. The rotating shafts of the left servo 423 and the right servo 424 respectively pass through the axial holes at the upper ends of the left rotating support 413 and the right rotating support 423 and extend into the left-tilted rotating shaft base 4337 and the right-tilted rotating shaft base 4338. The left-tilted rotating shaft base 4337 and the right-tilted rotating shaft base 4338 can provide free end support for the rotating shafts of the left servo 423 and the right servo 424, thereby improving the connection rigidity between them and the left and right rotating supports.

[0055] As the best way, the center of the hinge axis of the rotating bracket 4332 and the tilt servo mounting seat 432, that is, the center b of the tilt servo axis coincides with the center of gravity a of the aircraft, that is, the center b passes through the center of gravity a. At this time, when the rotating bracket 4332 is rotated to a vertical position, when the left tilt power mechanism 41 and the right tilt power mechanism 42 are in a vertical flight mode, the tension line / thrust line c of the rotor just passes through the center of gravity a. When flying horizontally, the rotating bracket 4332 is rotated forward to a horizontal position, at which time the tension line / thrust line c of the rotor also just passes through the center of gravity a. At the same time, during the tilting process, it can be ensured that the tension line / thrust line c also coincides with the center of gravity a of the aircraft.

[0056] See also Figure 7 , the figure shows that the aircraft of this embodiment is in the horizontal flight mode, the tilt steering gear shaft center b (the hinge axis center of the rotating bracket 4332) coincides with the aircraft gravity center a, at this time, the rotating bracket 4332 rotates forward to a horizontal position, and the left and right propellers (axis lines) are arranged horizontally, which can ensure that the thrust line c coincides with the aircraft gravity center a; see Figure 8 In the vertical flight mode, the rotating bracket 4332 rotates backward to a vertical position, and the left and right propellers (axis lines) are arranged vertically. At this time, it can also be ensured that the tension line / thrust line c coincides with the center of gravity a. The coincidence means that the center of gravity is located in the vertical plane (vertical flight mode) or horizontal plane (horizontal flight mode) where the tension line or thrust line is located.

[0057] Further, as a preferred solution, the bracket 433 is also provided with an electronic speed regulator 4335 and a power supply communication line 4336 for adjusting the rotation speed of the left motor 412 and the right motor 422. Specifically, the load-bearing crossbeam 4331 is in a plate shape and is arranged on the surface of the load-bearing crossbeam 4331.

[0058] In order to solve the problem of large propeller-wing coupling interference caused by the propeller adopting the pull-up type combined with the negative staggered wing arrangement in the prior art, see Fig. 9 —— Fig. 20 The figure shows the aerodynamic interference schematic diagram of four combination schemes of positive staggered and negative staggered box-wing layouts and propeller push-down and pull-up arrangements.

[0059] Fig. 9 , 10 11 is a schematic diagram of the start coupling of the negative staggered box wing layout combined with the push-down propeller layout. It can be seen from the figure that this scheme, Fig.10 In the vertical flight mode, the propeller-wing aerodynamic coupling condition occurs. The left tilting power mechanism 41 and the right tilting power mechanism 42 tilt forward, and the rear wing is in the weak induced interference of the inflow in front of the propeller. The propeller-wing coupling interference is small, and the aerodynamic interference is minimal. Fig.11This is the aerodynamic coupling between the propeller and the wing during the tilting process. It can be seen that there is no interference between the propeller and the wing during the tilting process.

[0060] Fig.12 , 13 14 is a schematic diagram of an orthogonal staggered box wing layout combined with a push-down propeller layout. Fig.13 This is the aerodynamic coupling situation in the vertical flight mode. At this time, the front wing is in the weak induced interference of the inflow in front of the propeller, and the propeller-wing coupling interference is small; Fig.14 This is the aerodynamic interference of the propeller when it tilts backward. At this time, it is easy to collide with the rear wing, and it is difficult to switch to fixed-wing mode. At the same time, it can be seen that there is strong propeller-wing interference with the rear wing during the tilting process.

[0061] Fig.15 , 16 17 is a schematic diagram of a negative staggered box wing layout combined with an upward-pull propeller layout. The prior art CN116573179A adopts this layout. Fig.16 This is the aerodynamic coupling situation in the vertical flight mode. The front wing is easily subject to the strong aerodynamic impact interference of the slipstream behind the propeller, and the propeller-wing coupling interference is large. Fig.17 It is the aerodynamic interference of the propeller when it tilts backward, and the interference of the propeller wing is smaller.

[0062] Fig.18 , 19 20 is a schematic diagram of an orthogonal staggered box wing layout combined with an upward-pull propeller layout. Fig.19 In the vertical flight mode, the rear wing is easily affected by the strong aerodynamic impact of the slipstream behind the propeller, and the propeller-wing coupling interference is large. Fig. 20 This is the aerodynamic coupling situation during the tilting process. At this time, there is strong propeller interference with the rear wing.

[0063] Therefore, the present invention has concluded through theoretical calculations and experiments that the front wing 21 and the rear wing 22 are arranged in a negative staggered manner with the front low and the rear high, and the left propeller 411 and the right propeller 421 are arranged in a push-down manner in the vertical flight mode, forming a compact layout with minimal aerodynamic interference in the vertical flight mode and during the tilting process.

[0064] Furthermore, a left elevon 31 and a right elevon 32 are arranged at the trailing edge of the rear wing 22, and a servo 1 33 and a servo 2 34 are installed in the vertical tail 25 to drive the left elevon 31 and the right elevon 32 to deflect and control the roll angle and pitch angle of the aircraft when it is in level flight.

[0065] Furthermore, the aircraft also includes an autonomous control system 5, which includes a flight control subsystem 51 and a mission control subsystem 52. The flight control subsystem 51 and the mission control subsystem 52 are installed at the rear of the load-bearing structure 11 of the fuselage 1, and the camera 53 and the lithium battery 54 are installed at the front of the fuselage load-bearing structure 11.

[0066] In this embodiment, in order to further optimize the structure of the aircraft and make its layout more compact, as a preferred solution, the wing of the present invention preferably includes: a front wing 21, a rear wing 22, a left end wing 23, and a right end wing 24. The fuselage 1 is directly connected to the front wing 21, the rear wing 22 is connected to the front wing 21 through the left end wing 23 and the right end wing 24, and the rear wing 22 is connected to the fuselage load-bearing structure 11 through the vertical tail 25. The front wing 21, the rear wing 22, the left end wing 23, and the right end wing 24 together form a connecting wing layout. By designing the wing as a connecting wing layout, combined with the power unit 4 located between the front wing 21 and the rear wing 22, on the one hand, the structure of the aircraft of the present invention can be made compact, and at the same time, the power unit 4 can adjust the left tilting power mechanism 41 and the right tilting power mechanism 42 through the adjustment mechanism 43 to ensure that the thrust line of its power passes through the center of gravity of the aircraft in both the horizontal flight state and the vertical flight state, so as to ensure that the aircraft is in the best power state.

[0067] The dual-tilt-rotor and fixed-wing hybrid layout aircraft described in this embodiment can fly horizontally and vertically. When in the vertical flight mode, by changing the tilt angle of the adjustment mechanism 43, as well as the tilt angle and rotation speed of the left tilt power mechanism 41 and the right tilt power mechanism 42, the rolling moment, pitch moment or yaw moment required for the roll, pitch and yaw movements in vertical and horizontal flight can be generated.

[0068] Specifically, see Fig.21 , the thrust of the left propeller 411 is T L , the thrust of the right propeller 421 is T R , the left propeller rotates at an angle of δ L , the rotation angle of the right propeller is δ R , the tilt angle of the tilt servo 431 is δ M , the distance between the rotating shaft of the tilting servo 431 and the rotating shaft of the left and right motors 412 and 422 is d, and the center of gravity a is located at the center b of the rotating shaft of the tilting servo (i.e., the center of the hinge axis of the bracket). In vertical flight mode (δ M =90°), ignoring the components of the forward and lateral forces on the aircraft, the generation mechanism of the vertical forces and moments in three directions of the aircraft is:

[0069] F z =-T L cosδ L -TR cosδ R

[0070] M x =T L L b cosδ L -T R L b cosδ R

[0071] M y =T L dsinδ L +T R dsinδ R

[0072] M z =-T L L b sinδ L +T R L b sinδ R

[0073] Assume the total flight weight of the aircraft is m. When the aircraft is in vertical hovering flight, T L =T R =mg / 2,δ L =δ R =0°,δ M =90°.

[0074] When the aircraft needs to roll, a rolling moment M is required. x , at this time, increase the left propeller thrust T L , reduce the right propeller thrust T R , that is, a right rolling moment is generated; increase the right propeller thrust T R , reduce the left propeller thrust T L , that is, a left rolling moment is generated.

[0075] When the aircraft needs to make a pitch motion, a pitch moment My needs to be generated. At this time, the tilt angles of the left and right motors are increased at the same time. L and δ R , according to M y =T L dsinδ L +T R dsinδ R , will generate a positive pitch moment; at the same time, reduce the left and right motor tilt angles δ L and δ R ,in accordance with

[0076] M y =T Ldsinδ L +T R dsinδ R , a negative nose-down pitching moment will be generated.

[0077] When the aircraft needs to yaw, a yaw moment Mz needs to be generated. At this time, the left motor tilt angle δ is increased. L , reduce the right tilt angle δ R , according to M z =-T L L b sinδ L +T R L b sinδ R , a negative left yaw moment will be generated; reduce the left motor rotation angle δ L , increase the right motor rotation angle δ R , according to M z =-T L L b sinδ L +T R L b sinδ R , a positive right yaw moment will be generated.

[0078] See also Fig. 22 , in fixed-wing horizontal flight mode (δ M =δ L =δ R =0°), the left rudder surface (left elevator 31) arranged behind the rear wing 22 has a deflection angle of δ aL , the right rudder surface (right elevator 32) deflection angle is δ aR , the two propellers generate forward thrust to overcome flight resistance. The left propeller thrust is T L , the right propeller thrust is T R , the torque control mechanism of the aircraft in this flight mode is:

[0079] M x =K1ρV 2 (δ aL -δ aR )

[0080] M y =K2ρV 2 (-δ aL -δ aR )

[0081] M z =T L L b -T R L b

[0082] Among them, ρ is the air density, V is the flight speed, K1 is the roll control coefficient, K2 is the pitch control coefficient, and K1 and K2 are positive constants.

[0083] When the aircraft needs to roll, a rolling moment M is required. x At this time, the left rudder surface (left elevator 31) deflects by an angle δ aL , reduce the right side (right elevator 32) deflection angle δ aR , that is, a positive right rolling moment is generated; increasing the right rudder surface (right elevator 32) deflection angle δ aR , reduce the deflection angle δ of the left rudder surface (left elevator 31) aL , that is, a negative left rolling moment is generated.

[0084] When the aircraft needs to make a pitch motion, a pitch moment My is generated. At this time, the deflection angles δ of the left and right control surfaces (left elevon 31 and right elevon 32) are increased at the same time. aL and δ aR , will produce a negative head-down pitching moment; at the same time, reduce the deflection angle δ of the left and right rudder surfaces (left elevon 31 and right elevon 32) aL and δ aR , a positive nose-up pitching moment will be generated.

[0085] When the aircraft needs to yaw, it needs to generate a yaw moment Mz. At this time, increase the left motor pull T L , reduce the right motor pulling force T R , will produce a positive right yaw moment; reduce the left motor pull T L , increase the right motor pulling force T R , a negative left yaw moment will be generated.

[0086] Example 2, see Fig.23 , 24The present embodiment is different from the previous embodiment in that: the load-bearing crossbeam 4331 is in the shape of a horizontally arranged flat plate, and a left servo mounting seat 4333 and a right servo mounting seat 4334 are symmetrically arranged on both sides of a rotating bracket 4332 integrally formed under the load-bearing crossbeam 4331, and the left servo mounting seat 4333 and the right servo mounting seat 4334 are fixed under the load-bearing crossbeam 4331. Obviously, the left servo mounting seat 4333 and the right servo mounting seat 4334 can also be integrally formed on the load-bearing crossbeam 4331; the left servo 414 and the right servo 424 are mounted on the left servo mounting seat 4333 and the right servo mounting seat 4334. Two left-leaning rotating shaft bases 4337 and two right-leaning rotating shaft bases 4338 are symmetrically provided at the two ends below the load-bearing crossbeam 4331. The left rotating support 413 is rotatably installed between the two left-leaning rotating shaft bases 4337 on the left side through the left rotating shaft portion 4131. Similarly, the right rotating support 423 can be rotatably installed between the right-leaning rotating shaft bases 4338 through the right rotating shaft portion 4231. The left rotating support 413 also has a left motor mounting seat 4132 vertically arranged on the rotating shaft portion 4131 for mounting the left motor 412. Similarly, the structure of the right rotating support 423 is the same as that of the left rotating support 413 and is symmetrically arranged. The vertical fixed installation refers to the vertical arrangement of the left motor mounting seat 4132 and the rotating shaft portion 4131. The output shafts of the left servo 414 and the right servo 424 are respectively connected to the ends of the rotating shaft 4131 of the left rotating support 413 and the right rotating support 423, and are used to drive the left rotating support 413 and the right rotating support 423 to rotate around the center line of the left tilting rotating shaft base 4337 and the right tilting rotating shaft base 4338. In this embodiment, the electronic speed regulator 4335 and the power supply communication line 4336 are arranged on the upper surface of the load-bearing crossbeam 4331, making full use of the space and making the structure compact.

[0087] The left-tilt power mechanism 41 and the right-tilt power mechanism 42 can rotate in the left-tilt rotation axis base 4337 and the right-tilt rotation axis base 4338, and can rotate synchronously with the rotating bracket 4332 and the load-bearing beam 4331 around the rotation axis center b of the tilt servo (the hinge axis center of the rotating bracket 4332); thereby, the left-tilt power mechanism 41 and the right-tilt power mechanism 42 can be adjusted to make the tension line / thrust line c coincide with the center of gravity a of the aircraft in both the vertical flight mode and the horizontal flight mode.

[0088] In this embodiment, the left servo mounting seat 4333 and the right servo mounting seat 4334 are separately provided, so that the installation of the left and right servos is more secure. At the same time, the left and right servos are located on the lower surface of the load-bearing beam 4331. When the adjustment mechanism 43 rotates forward to a horizontal position, the left and right servos are just located behind the flat load-bearing beam 4331, that is, in a non-windward position, thereby reducing the wind resistance of the small servo body. At the same time, the left and right rotating supports 413 and 423 are respectively connected through two left-leaning rotating shaft bases 4337 and right-leaning rotating shaft bases 4338, that is, both ends of the rotating shaft part of each rotating bracket have supports, so that its structure is more stable.

[0089] The above description is only illustrative and not restrictive of the present invention. The present invention aims to provide a hybrid wing layout aircraft. Those skilled in the art understand that many modifications, changes or equivalents may be made without departing from the spirit and scope defined by the claims, such as changing the structure of the bracket, rotating the bracket and the connection method of the load-bearing beam, etc., but all of these will fall within the scope of protection of the present invention.

Claims

1. A hybrid wing layout aircraft, comprising: A fuselage (1), wings (2), control surfaces (3), and a power unit (4), characterized in that: the power unit (4) is used to provide power and flight control during vertical flight and horizontal flight; the power unit (4) is located between the front wing (21) and the rear wing (22) of the wing (2); the power unit (4) only includes two independently tiltable left-tilt power mechanisms (41) and right-tilt power mechanisms (42) symmetrically arranged on both sides of the fuselage (1); an adjustment mechanism (43) for installing the left-tilt power mechanism (41) and the right-tilt power mechanism (42) and being able to adjust the left-tilt power mechanism (41) and the right-tilt power mechanism (42) to ensure that the tension or thrust line coincides with the center of gravity in both vertical flight and horizontal flight; the adjustment mechanism ( The invention relates to a tilting servo engine mounting seat (432) mounted on a load-bearing structure (11) of a fuselage (1), a tilting servo engine (431) fixedly mounted on the tilting servo engine mounting seat (432), and a bracket (433) hinged on the tilting servo engine mounting seat (432). The hinge axis of the bracket (433) is directly connected to the rotating shaft of the tilting servo engine (431) and can rotate from a vertical position to a forward horizontal position around the center of the rotating shaft of the tilting servo engine under the drive of the tilting servo engine (431); the left tilting power mechanism (41) and the right tilting power mechanism (42) are arranged at two ends of the bracket (433), and the adjustment mechanism (43) can ensure that the tension line or the thrust line can always pass through the center of gravity of the aircraft in the horizontal flight mode, the vertical flight mode and the rotation process.

2. A hybrid wing layout aircraft according to claim 1, characterized in that: When the bracket (433) is in a vertical position, the center of gravity of the aircraft and the tilt centers of the left tilt power mechanism (41) and the right tilt power mechanism (42) are located on the same vertical plane, and in a horizontal flight mode, when the bracket (433) is rotated to a horizontal position, the center of gravity of the aircraft and the tilt centers of the left tilt power mechanism (41) and the right tilt power mechanism (42) are located in the same horizontal plane.

3. A hybrid wing layout aircraft according to claim 2, characterized in that: The support (433) includes a load-bearing crossbeam (4331), a rotating support (4332) arranged in the middle of the load-bearing crossbeam (4331), the lower part of the rotating support (4332) is hinged to the two sides of the tilting servo mounting seat (432), and the left tilting power mechanism (41) and the right tilting power mechanism (42) are arranged at the two ends of the load-bearing crossbeam (4331).

4. A hybrid wing layout aircraft according to claim 3, characterized in that: The rotating bracket (4332) of the bracket (433) and the hinge axis center of the tilting servo mounting seat (432), the center of the tilting servo rotating axis and the center of gravity of the aircraft coincide with each other.

5. The hybrid wing layout aircraft according to claim 3, characterized in that: The left tilting power mechanism (41) comprises a left propeller (411), a left motor (412), a left rotating support (413) and a left steering gear (414); the right tilting power mechanism (42) comprises a right propeller (421), a right motor (422), a right rotating support (423) and a right steering gear (424); the left steering gear (414) and the right steering gear (424) are symmetrically arranged on the load-bearing crossbeam (4331); the left motor (412) is installed on the left rotating support (413); the left propeller (421) and the right motor (422) are symmetrically arranged on the load-bearing crossbeam (4331); the left motor (412) is installed on the left rotating support (413); the left propeller (421) and the right motor (422) are symmetrically arranged on the load-bearing crossbeam (4331); the left propeller (421) and the right motor (422) are symmetrically arranged on the load-bearing crossbeam (4331); the left propeller (421) and the right propeller ... The propeller (411) is installed on the left motor (412), the right motor (422) is installed on the right rotating support (423), and the right propeller (421) is installed on the right motor (422); the left rotating support (413) and the right rotating support (423) are respectively connected to the left steering gear (414) and the right steering gear (424), and can rotate around the rotating shaft center of the left steering gear (414) and the right steering gear (424) parallel to the load-bearing crossbeam (4331) under the drive of the left steering gear (414) and the right steering gear (424).

6. A hybrid wing layout aircraft according to claim 5, characterized in that: The left servo (414) and the right servo (424) are directly mounted on the two ends of the load-bearing crossbeam (4331) by means of a left servo buckle (4321) and a right servo buckle (4322) or by bonding; The left rotating support (413) and the right rotating support (423) are connected to the rotating shafts of the left servo (414) and the right servo (424) through the shaft holes at the upper ends, and a left tilting rotating shaft base (4337) and a right tilting rotating shaft base (4338) for supporting the ends of the rotating shafts of the left servo (414) and the right servo (424) are respectively provided at the two ends of the load-bearing crossbeam (4331).

7. The hybrid wing layout aircraft according to claim 5, characterized in that: The left servo (414) and the right servo (424) are respectively mounted under the load-bearing crossbeam (4331) through a left servo mounting seat (4333) and a right servo mounting seat (4334) which are symmetrically arranged under the load-bearing crossbeam (4331); a left-tilted rotating shaft base (4337) and a right-tilted rotating shaft base (4338) are symmetrically arranged on the same side of both ends of the load-bearing crossbeam (4331); the left rotating support (413) and the right rotating support (423) can be rotatably mounted on the left-tilted rotating shaft base (4337) and the right-tilted rotating shaft base (4338), respectively; the left servo (414) and the right servo (424) drive the left rotating support (413) and the right rotating support (423) to rotate around the axis of the left-tilted rotating shaft base (4337) and the right-tilted rotating shaft base (4338).

8. A hybrid wing aircraft according to any one of claims 3 to 7, characterized in that: The rotating bracket (4332) is fixedly connected at the upper end or integrally formed on the load-bearing crossbeam (4331).

9. A hybrid wing layout aircraft according to any one of claims 5 to 7, characterized in that: The bracket (433) is provided with an electronic speed regulator (4335) for adjusting the rotation speed of the left motor (412) and the right motor (422) and a power supply communication line (4336).

10. A hybrid wing aircraft according to any one of claims 5 to 7, characterized in that: The front wing (21) and the rear wing (22) are arranged in a negative staggered manner with the front wing being lower and the rear wing being higher, and the left propeller (411) and the right propeller (421) are arranged in a push-down manner in the vertical flight mode, forming a compact layout with minimal aerodynamic interference in the vertical flight mode.

11. A hybrid wing layout aircraft according to any one of claims 2 to 7, characterized in that: The wing (2) comprises a front wing (21), a rear wing (22), a left end wing (23), and a right end wing (24); the fuselage (1) is directly connected to the front wing (21); the rear wing (22) is connected to the front wing (21) via the left end wing (23) and the right end wing (24); and the rear wing (22) is connected to the load-bearing structure (11) of the fuselage (1) via a vertical tail (25); The two ends of the front wing (21) and the rear wing (22) are respectively connected to the left end wing (23) and the right end wing (24).

12. A hybrid wing aircraft according to any one of claims 5 to 7, characterized in that: The trailing edge of the rear wing (22) is provided with a left elevon (31) and a right elevon (32) for controlling the rolling angle and the pitching angle of the aircraft during level flight. A first servo (33) and a second servo (34) are installed in the vertical tail (25) for driving the left elevon (31) and the right elevon (32) to deflect.

13. A hybrid wing aircraft according to any one of claims 2 to 7, characterized in that: It further comprises an autonomous control system (5), wherein the autonomous control system (5) comprises a flight control subsystem (51), a mission control subsystem (52), a camera (53) and a lithium battery (54), wherein the flight control subsystem (51) and the mission control subsystem (52) are installed at the rear of the load-bearing structure (11) of the fuselage (1), and the camera (53) and the lithium battery (54) are installed at the front of the load-bearing structure (11) of the fuselage.

14. A hybrid wing aircraft according to any one of claims 5 to 7, characterized in that: In the vertical flight mode, the adjustment mechanism (43) is rotated to a vertical position; the pulling force of the left propeller (411) of the left tilting power mechanism (41) and the right propeller (421) of the right tilting power mechanism (42) are differentiated to generate a rolling moment and realize a rolling motion; the tilting angles of the left propeller (411) and the right propeller (421) are simultaneously increased or decreased to generate a pitching moment and realize a pitching motion; the tilting angles of the left propeller (411) and the right propeller (421) are differentiated to generate a yaw moment and realize a yaw motion.

15. The hybrid wing layout aircraft according to claim 12, characterized in that: In the fixed-wing horizontal flight mode, the adjustment mechanism (43) is rotated forward to a horizontal position, and the deflection angles of the left elevator (31) and the right elevator (32) are differentiated to generate a rolling moment; the deflection angles of the left elevator (31) and the right elevator (32) are increased or decreased at the same time to generate a pitching moment; and the pulling force of the left propeller (411) of the left tilting power mechanism (41) and the right propeller (421) of the right tilting power mechanism (42) are differentiated to generate a yaw moment.

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