Non-stall floating wing mechanism and adjusting method thereof

Abstract

The invention discloses a non-stall floating wing mechanism and an adjusting method thereof. The non-stall floating wing mechanism comprises a fuselage and wings, the two ends of the wing main beam respectively extend out of the fuselage, penetrate through the floating connecting plates and are rotationally connected with the wing main beam shaft holes; the force transmission main beam is located in the fuselage, the two ends of the force transmission main beam extend out of the fuselage respectively, and the force transmission main beam is rotationally connected with force transmission main beam shaft holes in the floating connecting plates; one end of the linear steering engine is hinged to the floating connecting plates, and the other end of the linear steering engine is hinged to the wings; wherein the included angle between the rotating center connecting line of the wing main beam shaft hole and the force transmission main beam shaft hole and the vertical line of the chord line of the wing is theta, and the theta angle is changed by adjusting the length of the linear steering engine. The wing attack angle can be separated from the pitching attitude angle of the fuselage, and the wing trim attack angle is adjustable during flight, so that wing stall-free flight and high-aerodynamic-efficiency flight of the aircraft are realized.

Description

Technical Field

[0001] The invention relates to the technical field of aircraft, and in particular to a non-stall floating wing mechanism and an adjustment method thereof. Background Art

[0002] At present, the wings of fixed-wing aircraft are fixedly connected to the fuselage. There is a risk of stalling due to excessive angle of attack during flight, which may cause the aircraft to crash. At the same time, there is a stall problem, which affects flight safety. At the same time, on tail-seat vertical take-off and landing aircraft, when the wing is converted to vertical, there is an intermediate transition state of wing stall, which affects flight control and cannot smoothly complete the vertical mode conversion. In the prior art, there are also some aircraft with free wings. The free wings are mainly used to reduce gust loads. Due to the balancing torque problem, its maximum lift-to-drag ratio is relatively low and cannot meet flight requirements.

[0003] Therefore, it is urgent to solve the above problems. Summary of the invention

[0004] Purpose of the invention: The first purpose of the present invention is to provide a non-stall floating wing mechanism, which can realize the separation of the wing angle of attack and the fuselage pitch attitude angle, and the wing trim angle of attack is adjustable during flight, thereby realizing non-stall flight of the aircraft wing and high aerodynamic efficiency flight.

[0005] A second object of the present invention is to provide a method for adjusting a non-stall floating wing mechanism.

[0006] Technical solution: To achieve the above purpose, the present invention discloses a non-stall floating wing mechanism, including a fuselage and a wing, and also including a wing main beam located in the fuselage and having two ends extending out of the fuselage and rotatably connected to the wing main beam shaft hole through a floating connecting plate, a force transmission main beam located in the fuselage and having two ends extending out of the fuselage and rotatably connected to the force transmission main beam shaft hole on the floating connecting plate, and a linear servo with one end hinged to the floating connecting plate and the other end hinged to the wing, wherein the angle between the rotation center line of the wing main beam shaft hole and the force transmission main beam shaft hole and the perpendicular line of the wing chord line is angle θ, and the size of angle θ can be changed by adjusting the length of the linear servo.

[0007] Optionally, the wing main beam axis hole is located at the wing focus or pressure center, or the distance between the wing main beam axis hole and the wing focus or pressure center does not exceed 5% of the root chord length.

[0008] Optionally, the shaft hole of the force transmission main beam is located at the center of gravity of the fuselage, or the distance between the shaft hole of the force transmission main beam and the center of gravity of the fuselage does not exceed 5% of the root chord length.

[0009] Optionally, the force transmission main beam axis hole is located in the front and lower part of the wing main beam axis hole.

[0010] Optionally, one end of the linear servo is hinged to the floating connecting plate through a hinge point, and the other end is hinged to the wing rear fulcrum on the wing through a wing rear fulcrum beam.

[0011] Optionally, the floating connecting plate rotates counterclockwise around the main force transmission beam axis hole by an angle α, which is the wing attack angle α during horizontal flight, tan(θ+α)=D / L, D is the corresponding wing resistance, and L is the corresponding wing lift.

[0012] Optionally, a rear fulcrum member is provided on the fuselage for limiting the maximum installation angle of the wing relative to the fuselage.

[0013] Optionally, the force transmission main beam is provided with a bearing connected to the fuselage.

[0014] Optionally, a rotation damper connected to the fuselage is provided on the force transmission main beam.

[0015] Based on the same inventive concept, the present invention discloses a method for adjusting a non-stall floating wing mechanism, comprising the following steps:

[0016] When the aircraft takes off and is in a flight condition that requires high lift, the linear servo is shortened, the angle θ is reduced, and the angle α is increased;

[0017] When switching from climb to cruise, in the condition of medium lift and high flight efficiency, the linear servo is extended to increase the angle θ, thereby reducing the angle α to the angle of attack corresponding to the maximum lift-to-drag ratio of the wing;

[0018] When lowering the altitude, in low lift flight conditions, the linear servo is extended to increase the angle θ, thereby reducing the angle α.

[0019] Beneficial effects: Compared with the prior art, the present invention has the following significant advantages: the wing angle of attack of the present invention is separated from the pitch attitude angle of the fuselage, and the wing trim angle of attack is adjustable during flight, so that the aircraft maintains a high lift-to-drag ratio flight state, thereby achieving non-stall flight of the aircraft wing and high aerodynamic efficiency flight; on vertical take-off and landing aircraft, the wing of the present invention can adapt to changes in flight direction, continuously generate lift, and avoid the intermediate transition state of stall; the rear fulcrum member of the present invention limits the maximum installation angle of the wing relative to the fuselage, and unidirectionally supports the trailing edge of the wing. When the aircraft lowers its head and dives at a large angle, the floating connecting plate contacts the rear fulcrum member, the wing no longer floats, and the angle of attack changes synchronously with the pitch angle of the fuselage, thereby realizing a high-speed dive function. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Figure 1 It is a schematic diagram of the principle of the present invention;

[0021] Figure 2 A schematic diagram of the angle θ in the initial state of the present invention;

[0022] Figure 3 Schematic diagram of the angle θ when the linear servo is shortened in the present invention;

[0023] Figure 4 Schematic diagram of the angle θ when the linear servo is extended in the present invention;

[0024] Figure 5 Schematic diagram of balancing the wing attack angle α in the present invention;

[0025] Figure 6 It is a schematic diagram of the present invention used on a fixed-wing aircraft;

[0026] Figure 7 It is a structural schematic diagram of the present invention. DETAILED DESCRIPTION

[0027] In order to make the purpose, technical solution and advantages of the present invention more clearly understood, the present invention is further described in detail below in conjunction with the accompanying drawings and specific embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not used to limit the present invention.

[0028] It should be understood that the present invention can be implemented in different forms and should not be construed as being limited to the embodiments set forth herein. On the contrary, providing these embodiments will make the disclosure thorough and complete and fully convey the scope of the present invention to those skilled in the art. In the accompanying drawings, the sizes and relative sizes of the components may be exaggerated for clarity. The same reference numerals throughout represent the same components.

[0029] like Figure 1 and Figure 2 As shown, the present invention discloses a non-stall floating wing mechanism, the floating wing mechanism of the present invention connects the wing and the fuselage, transmits and balances the wing aerodynamic load and the fuselage gravity and aerodynamic load. The floating wing mechanism comprises a fuselage, a wing 10, a floating connecting plate 20, a wing main beam 211, a force transmission main beam 221, a linear servo 41, a wing rear fulcrum beam 111 and a rear fulcrum member 31, wherein the wing main beam 211 is located in the fuselage, and the two ends of the wing main beam 211 extend out of the fuselage respectively, and the two ends of the wing main beam 211 respectively pass through a floating connecting plate 20 and are rotatably connected to the wing main beam shaft holes 21 on the wings on both sides, and the wing can rotate around the wing main beam shaft hole 21. The force transmission main beam 221 is located in the fuselage, and both ends of the force transmission main beam 221 extend out of the fuselage respectively. The two ends of the force transmission main beam 221 are rotatably connected to the force transmission main beam shaft holes 22 on the floating connecting plates 20 on both sides respectively, and the floating connecting plates 20 can rotate around the force transmission main beam shaft holes 22.

[0030] One end of the linear servo 41 is hinged to the floating connection plate 20 through the hinge point 23, and the other end of the linear servo 41 is hinged to the wing rear fulcrum 11 on the wing through the wing rear fulcrum beam 111, and the wing rear fulcrum beam 111 is parallel to the wing main beam 211. The angle θ between the rotation center connection line of the wing main beam shaft hole 21 and the force transmission main beam shaft hole 22 of the linear servo 41 and the vertical line of the wing chord line is hinged, and the wing rear fulcrum 11 and the hinge point 23 are both hinged. When the linear servo is not adjusted, the relative position of the wing 10 and the floating connection plate 20 remains fixed; by adjusting the length of the linear servo 41, the wing angle of attack α can be adjusted, so that the wing angle of attack is at a larger lift-to-drag ratio during flight, thereby achieving high aerodynamic efficiency flight.

[0031] The floating connecting plate 20 rotates counterclockwise around the force transmission main beam shaft hole 22 by an angle α, which is the wing attack angle α during horizontal flight, tan(θ+α)=D / L, D is the corresponding wing resistance, and L is the corresponding wing lift. The present invention can adjust the trim attack angle α by adjusting the length of the linear servo 41, so that the wing attack angle is at a larger lift-to-drag ratio during flight, thereby achieving high aerodynamic efficiency flight. In this state, the flight rise and fall are adjusted by adjusting the flight speed, increasing the speed lift, increasing the height rise, reducing the speed lift, and reducing the height descent.

[0032] The angle between the rotation center connecting line of the wing main beam shaft hole 21 and the force transmission main beam shaft hole 22 and the vertical line of the wing chord line is angle θ, and the length of the linear servo 41 is adjusted to change the angle θ. The wing main beam shaft hole 21 is located at the wing focus or pressure center position, or the distance between the wing main beam shaft hole 21 and the wing focus or pressure center position does not exceed 5% of the root chord length, that is, the point of action of the main aerodynamic force of the wing, so that the linear servo 41 adjusts the load required for the wing to be minimized. The wing 10 has a rotational degree of freedom relative to the wing main beam shaft hole 21.

[0033] The force transmission beam shaft hole 22 is located at the center of gravity of the fuselage, or the distance between the force transmission beam shaft hole 22 and the center of gravity of the fuselage does not exceed 5% of the root chord length, thereby reducing the longitudinal trim moment of the aircraft. The force transmission beam shaft hole 22 is located in the front and lower part of the wing beam shaft hole 21. The floating connecting plate 20 has a rotational freedom relative to the force transmission beam shaft hole 22, that is, it has a rotational freedom relative to the fuselage.

[0034] The fuselage is provided with a rear fulcrum member 31 for limiting the maximum installation angle of the wing relative to the fuselage. The rear fulcrum member 31 is installed on the fuselage to limit the maximum installation angle of the wing relative to the fuselage. The rear fulcrum member 31 supports the trailing edge of the wing in one direction. When the aircraft lowers its head and dives at a large angle, the floating connecting plate 20 contacts the rear fulcrum member 31, the wing no longer floats, and the change of the angle of attack is synchronized with the pitch angle of the fuselage, realizing the high-speed dive function.

[0035] A bearing 32 and a rotation damper 33 are provided on the force transmission main beam 221 , and both the bearing 32 and the rotation damper 33 are fixed on the fuselage.

[0036] like Figure 6 The figure shows an embodiment of a floating wing mechanism on a fixed-wing aircraft. In the figure, one wing and the middle fuselage are made transparent. The floating wing mechanism is located between the wing and the fuselage, and the above floating wing function is realized on the aircraft. In order to more clearly show the arrangement of the floating wing mechanism, one wing and the middle fuselage are hidden, and the figure is enlarged, as shown in FIG. Figure 7 As shown, after the wing main beam 211 passes through the wing main beam shaft hole 21 from the wing 10, it passes through the fuselage to transfer the wing lift to the floating wing mechanism, and the wing bending moment is self-balanced by the wing main beam 211; the floating connecting plate 20 transfers the wing aerodynamic load to the force transmission main beam 221 through the force transmission main beam shaft hole 22, and the force transmission main beam 221 transfers the wing aerodynamic load to the fuselage through the bearing 32, and the rotating damper 33 is used to increase the rotation damping of the force transmission main beam 221; the rear fulcrum member 31 is installed on the fuselage inwardly, and contacts with the floating connecting plate 20 when the aircraft dives at a large angle with its head down outward, so that the wing no longer floats, and the angle of attack changes synchronously with the pitch angle of the fuselage, realizing the high-speed dive function; the wing rear fulcrum 11 is connected to the rear end of the wing through the wing rear fulcrum beam 111 to transfer the load of the linear servo 41. The floating wing mechanism can also be used for tail-seat vertical take-off and landing drones. The present invention avoids the problem of wing stall and improves flight safety. In a vertical take-off and landing aircraft, the wing can adapt to changes in flight direction and continuously generate lift, thereby avoiding a transition state during stall. The present invention adjusts the trim angle of attack to keep the aircraft in a high lift-to-drag ratio flight state, thereby improving flight efficiency.

[0037] like Figure 2 As shown, the horizontal state of the wing 10 is taken as the initial state. At this time, the incoming air flow is also in the horizontal direction. As the initial state of calculation, the angle θ is the angle between the rotation center line of the wing main beam shaft hole 21 and the force transmission main beam shaft hole 22 and the vertical line of the wing chord line. θ1 is the θ angle after shortening the linear actuator 41, and θ2 is the θ angle after extending the linear actuator 41. Figure 3 and 4 As shown in FIG. 1 , the wing 10 and the floating connecting plate 20 rotate counterclockwise around the main beam shaft hole 22 under the action of the airflow and reach a balance. Under the action of the resistance D of the wing 10, the wing rotates backward around the axis, and the lift L of the wing 10 drives the wing to rotate clockwise around the axis. Figure 1 , Figure 2 and Figure 5 As shown, balance is achieved after the rotation angle is α, and the angle complies with the mathematical law formula tan(θ+α)=D / L. The adjustment of the linear servo 41 follows the following law: increasing the θ angle will reduce the trim α angle, and decreasing the θ angle will increase the trim α angle.

[0038] The present invention provides a method for adjusting a non-stall floating wing mechanism, comprising the following steps:

[0039] When the aircraft takes off and is in a flight condition that requires high lift, the linear servo is shortened, the angle θ is reduced, and the angle α is increased;

[0040] When switching from climb to cruise, in the condition of medium lift and high flight efficiency, the linear servo is extended to increase the angle θ, thereby reducing the angle α to the angle of attack corresponding to the maximum lift-to-drag ratio of the wing;

[0041] When lowering the altitude, in low lift flight conditions, the linear servo is extended to increase the angle θ, thereby reducing the angle α.

Claims

1. A non-stall floating wing mechanism, comprising a fuselage and a wing (10), characterized in that: The invention also comprises a wing main beam (211) located in the fuselage and having two ends extending out of the fuselage respectively and passing through a floating connection plate (20) to be rotatably connected to a wing main beam shaft hole (21); a force transmission main beam (221) located in the fuselage and having two ends extending out of the fuselage respectively and being rotatably connected to a force transmission main beam shaft hole (22) on the floating connection plate (20); and a linear servo (41) having one end hinged to the floating connection plate (20) and the other end hinged to the wing, wherein the angle between the rotation center line of the wing main beam shaft hole (21) and the force transmission main beam shaft hole (22) and the perpendicular line of the wing chord line is an angle θ, and the size of the angle θ can be changed by adjusting the length of the linear servo (41).

2. The non-stall floating wing mechanism according to claim 1, characterized in that: The wing main beam axis hole (21) is located at the wing focus or pressure center position, or the distance between the wing main beam axis hole (21) and the wing focus or pressure center position does not exceed 5% of the root chord length.

3. The non-stall floating wing mechanism according to claim 1, characterized in that: The force transmission main beam shaft hole (22) is located at the center of gravity of the fuselage, or the distance between the force transmission main beam shaft hole (22) and the center of gravity of the fuselage does not exceed 5% of the root chord length.

4. The non-stall floating wing mechanism according to claim 1, characterized in that: The force transmission main beam shaft hole (22) is located at the front lower part of the wing main beam shaft hole (21).

5. The non-stall floating wing mechanism according to claim 1, characterized in that: One end of the linear servo (41) is hinged to the floating connection plate (20) via a hinge point (23), and the other end is hinged to the wing rear fulcrum (11) on the wing via a wing rear fulcrum beam (111).

6. The non-stall floating wing mechanism according to claim 1, characterized in that: The floating connecting plate (20) rotates counterclockwise around the force transmission main beam shaft hole (22) at an angle α, which is the wing attack angle α during horizontal flight, tan (θ+α) = D / L, D is the corresponding wing resistance, and L is the corresponding wing lift.

7. The non-stall floating wing mechanism according to claim 1, characterized in that: The fuselage is provided with a rear fulcrum member (31) for limiting the maximum installation angle of the wing relative to the fuselage.

8. The non-stall floating wing mechanism according to claim 1, characterized in that: The force transmission main beam (221) is provided with a bearing (32) connected to the fuselage.

9. The non-stall floating wing mechanism according to claim 1, characterized in that: The force transmission main beam (221) is provided with a rotation damper (33) connected to the fuselage.

10. A method for adjusting a non-stall floating wing mechanism according to any one of claims 1 to 9, characterized in that: The steps include: When the aircraft takes off and is in a flight condition that requires high lift, the linear servo is shortened, the angle θ is reduced, and the angle α is increased; When switching from climb to cruise, in the condition of medium lift and high flight efficiency, the linear servo is extended to increase the angle θ, thereby reducing the angle α to the angle of attack corresponding to the maximum lift-to-drag ratio of the wing; When lowering the altitude, in low lift flight conditions, the linear servo is extended to increase the angle θ, thereby reducing the angle α.

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