A method for realizing flight control of a delta-wing aircraft

Abstract

The present invention discloses a method for realizing the flight control of a delta-wing aircraft, belonging to the field of aviation technology. The delta-wing aircraft has a large-area shell-shaped wide-chord wing. There are multiple rows of driving devices arranged above and below the shell-shaped wing to generate laminar driving airflows on the wing surface. The numerous driving devices divide the wing and the driving devices into several zones with the aerodynamic center of the large-area shell-shaped wide-chord wing as the origin. By adjusting the power of the driving devices above the shell-shaped wing in a certain zone and / or the power of the driving devices below, that is, adjusting the pressure of the laminar driving airflow on the upper surface of the shell-shaped wing in this zone and / or the pressure of the laminar driving airflow on the lower surface, thereby changing the force on the upper wing surface of the shell-shaped wing in this zone and / or the force on the lower wing surface. By changing the forces on the upper and lower surfaces of the wing surfaces in each zone of the wing, the flight attitude of the large-area shell-shaped wide-chord wing is adjusted to realize the flight control of the delta-wing aircraft.

Description

Technical Field

[0001] The invention relates to a method for realizing flight control of an armored wing aircraft, in particular to a method for realizing vertical take-off and landing and flight of a fixed-wing aircraft, and belongs to the field of aviation technology. Background Art

[0002] Manned aircraft can be divided into two categories: fixed-wing and rotary-wing. Fixed-wing aircraft are common in horizontal take-off and landing aircraft (such as jet airliners), while rotary-wing aircraft are common in vertical take-off and landing aircraft (such as helicopters). In terms of the current technical level, fixed-wing aircraft can fly at high speeds and are easy to operate, but they need to rely on runways for take-off and landing. Rotary-wing aircraft can take off and land vertically, do not need to rely on runways for take-off and landing, and have strong adaptability, but their mechanisms are out of balance, their controls are complex, their flight speeds are slow, and their fuel efficiency is low.

[0003] At present, there are still some inherent defects in vertical take-off and landing aircraft that cannot be overcome. First, the efficiency is low. The thrust-to-weight ratio of the vertical take-off and landing aircraft currently in use is less than 1, and the flight efficiency is low. Compared with the thrust-to-weight ratio of fixed-wing aircraft, which is generally greater than 5, the gap is very obvious; second, due to the limitation that the absolute speed of the wing tip must be less than the speed of sound, the theoretical speed of multi-rotor aircraft cannot exceed 120 kilometers per hour, and the flight speed is limited; third, the flapping of the rotor blades produces mechanical vibration, which increases the wear of the hinge and has low reliability; fourth, the rotor diameter and speed are limited by the wing tip speed cannot exceed the speed of sound. The maximum rotor diameter is generally more than ten meters, and the maximum take-off weight of the aircraft is limited and cannot be increased; fifth, many new composite vertical take-off and landing aircraft try to use fixed wings, but they face the contradiction that small-area fixed wings have limited effects and large-area fixed wings will block the downwash airflow during vertical take-off and landing.

[0004] In short, the above defects of vertical take-off and landing aircraft come from the contradiction that the wings have to take into account both vertical take-off and landing and horizontal flight modes. The method of providing drive devices on the upper and lower parts of the wings and dividing the large-area wings and drive devices into several partitions for separate control can realize the flight control of the large-area wings, making the A-wing aircraft well compatible with the two flight modes of vertical take-off and landing and horizontal flight, with more efficient operation, safer flight, more reasonable structure and more stable performance, which is a new development direction of vertical take-off and landing aircraft. Summary of the invention

[0005] The technical problem to be solved by the present invention is to address the issues of low flight efficiency, small payload, complex control, poor safety, and mutual conflict between the horizontal flight mode and the vertical takeoff and landing mode in existing vertical takeoff and landing aircraft. A method for realizing the flight control of a delta-wing aircraft is proposed. Driving devices are provided on the upper and lower parts of the wing. At the same time, the large-area wing and the driving devices are divided into several zones for separate control. By adjusting the power of the driving devices, the pressure of the driving air flow is changed, thereby changing the pressure of the driving air flow on the wing surface of this zone. By changing the pressure on the upper and lower surfaces of each zone of the wing, the flight attitude of the large-area shell-shaped wide-chord wing is adjusted to realize the flight control of the delta-wing aircraft.

[0006] To solve the above technical problem, the present invention provides a method for realizing the flight control of a delta-wing aircraft. The delta-wing aircraft has a large-area shell-shaped wide-chord wing. Driving devices are fixedly installed on the upper and lower parts of the shell-shaped wing respectively. The driving air flow generated by the upper driving device flows along the upper surface of the shell-shaped wing, and the driving air flow generated by the lower driving device flows along the lower surface of the shell-shaped wing. The driving devices are arranged in multiple rows along the wing chord direction and are densely distributed along the wing span direction. The driving devices generate a layered driving air flow on the wing surface. The thickness of the layered driving air flow is less than one-fourth of the chord length (width) of the large-area shell-shaped wide-chord wing. With the aerodynamic center of the large-area shell-shaped wide-chord wing as the origin, the wing and the driving devices are divided into several zones. By adjusting the power of the upper driving device and / or the lower driving device of the shell-shaped wing in a certain zone, that is, adjusting the pressure of the layered driving air flow on the upper surface and / or the pressure of the layered driving air flow on the lower surface of the shell-shaped wing in this zone, thereby changing the force on the upper wing surface and / or the force on the lower wing surface of the shell-shaped wing in this zone. By changing the forces on the upper and lower surfaces of each zone of the wing, the flight attitude of the large-area shell-shaped wide-chord wing is adjusted to realize the flight control of the delta-wing aircraft.

[0007] The large-area shell-shaped wide-chord wing is designed as a combined type, which is horizontally divided into several tilting wings and several fixed wings. Each partition includes several tilting wings and several fixed wings. All the tilting wings in each partition are connected to the fuselage through the same tilting mechanism and are controlled by the same tilting signal. The tilting wings of each partition are respectively controlled by their own tilting signals and driven by their own tilting mechanisms; the tilting wings and the fixed wings are arranged alternately like zebra crossings. Through the tilting of the tilting wings, they can be combined with the fixed wings as a whole to achieve horizontal flight, or the tilting wings can be opened like a shutter to form a strip-shaped airflow channel to achieve vertical takeoff and landing; a row of driving devices are densely arranged along the span direction on the upper and lower parts of the tilting wings. The driving devices tilt synchronously with the tilting wings. A large number of small-diameter driving devices are densely arranged adjacent to the wing surface to generate layered driving airflow along the wing surface; during the vertical takeoff and landing stage, the tilting wings tilt downward to form several vertical strip-shaped airflow channels on the large-area shell-shaped wide-chord wing. The driving devices on each section of the tilting wings are connected in parallel front and back. Through the strip-shaped airflow channels, air is inhaled from above the fixed wings and exhausted below the fixed wings to control the flight of each partition in the quadrotor mode; during the transition flight stage and the horizontal flight stage, by adjusting the tilting angle of each partition and the intensity of the layered driving airflow on the upper and lower surfaces of each tilting wing, the lift of each partition of the wing can be changed to achieve the flight control of the winged aircraft.

[0008] The most basic partition method of the partition is four partitions. Taking the aerodynamic center of the large-area shell-shaped wide-chord wing as the origin, the wing and the driving devices are divided into 4 partitions, namely the front left partition, the rear left partition, the front right partition and the rear right partition; on the basis of the 4-partition division, two partitions in the front or the rear are merged into one partition to form a 3-partition division, or each partition is subdivided into two partitions to form an 8-partition division, or each partition is subdivided into 4 partitions to form a 16-partition division, or the two partitions in the front are subdivided into 4 partitions and the two partitions in the rear are merged into one partition to form a 5-partition division, and so on.

[0009] The driving device adopts distributed drive and is composed of a large number of small-diameter motor propellers or ducted fans or jet engines or air nozzles or small engines; all the driving devices in each partition rotate in the same direction; a large number of tilting wings in each partition tilt synchronously under the control of the same tilting signal; when divided into four partitions, there are 4 independent tilting mechanisms respectively controlled by 4 tilting signals; the upper driving devices of each partition are controlled by one power signal, and the lower driving devices of each partition are controlled by another power signal; in short, each partition has 3 flight control signals, namely 1 tilting signal, 1 power signal of the upper driving device, and 1 power signal of the lower driving device; the winged aircraft controls the flight by adjusting the tilting signal and the power signal of each partition to change the force on each part of the upper and lower surfaces of the wing.

[0010] During the vertical takeoff and landing phase of the said A-wing aircraft, the tilting angle of each tilting wing section is tilted downward by 90°, and the driving device is vertically upward. Each partition is controlled according to a quadcopter. By changing the power of the driving device in each partition, the lift of each partition is adjusted while balancing the torque to control the flight. During the transition flight phase of the A-wing aircraft, the flight is controlled by adjusting the tilting angle of each partition and the power of the upper and lower driving devices. During the takeoff transition flight phase, the tilting wings gradually rotate upward, and the driving device gradually rotates horizontally forward. In the high-power state of the quadcopter, the driving device generates a forward pulling force to drive the A-wing aircraft forward. The tilting wing with a positive angle of attack generates an angle-of-attack lift when moving forward. The angle-of-attack lift is greater than the vertical lift loss caused by the forward tilt of the driving device, and the A-wing aircraft obtains a tendency to move upward. At the same time, the different air flow states in front of each tilting wing section cause the angle-of-attack lift obtained by the front-row tilting wings to be greater than that of the rear-row tilting wings, and the A-wing aircraft generates a nose-up moment. Therefore, during the takeoff transition flight phase, the control needs to gradually reduce the power of the driving device in each partition and finally transition to horizontal flight control. During the landing transition flight phase, the tilting wings gradually tilt downward, and the driving device gradually moves vertically upward. The inertia of horizontal flight keeps the A-wing aircraft moving forward. The tilting wing with a positive angle of attack generates an angle-of-attack lift when moving forward. The different air flow states in front of each tilting wing section cause the angle-of-attack lift obtained by the front-row tilting wings to be greater than that of the rear-row tilting wings, and the A-wing aircraft generates a nose-up moment. In the low-power state of horizontal flight, the insufficient lift obtained by the A-wing aircraft causes a tendency to move downward. Therefore, during the landing transition flight phase, the control needs to gradually increase the power of the driving device in each partition and finally transition to quadcopter control.

[0011] During the transition flight phase of the said A-wing aircraft, the pitch control of the A-wing aircraft is achieved by changing the power of the driving devices in the front and rear row partitions to adjust the lift difference of the wings in the front and rear row partitions. The nose-up control is achieved by increasing the power of the driving device in the front row partition and reducing the power of the driving device in the rear row partition in equal proportion. The nose-down control is achieved by reducing the power of the driving device in the front row partition and increasing the power of the driving device in the rear row partition in equal proportion.

[0012] During the transition flight phase of the said A-wing aircraft, the roll control of the A-wing aircraft is achieved by changing the power of the driving devices in the left and right side partitions to adjust the lift difference of the wings on the left and right sides. The right roll control of the A-wing aircraft is achieved by increasing the power of the lower driving device in the left side partition and reducing the power of the upper driving device in the left side partition in equal proportion, while reducing the power of the lower driving device in the right side partition and increasing the power of the upper driving device in the right side partition in equal proportion. The left roll control of the A-wing aircraft is achieved by reducing the power of the lower driving device in the left side partition and increasing the power of the upper driving device in the left side partition in equal proportion, while increasing the power of the lower driving device in the right side partition and reducing the power of the upper driving device in the right side partition in equal proportion.

[0013] During the transition flight phase of the said A-wing aircraft, the yaw control of the A-wing aircraft is achieved by changing the power of the drive devices in the left and right partitions to adjust the lift difference between the left and right drive devices. The right yaw control of the A-wing aircraft is achieved by increasing the power of the drive device in the left partition and reducing the power of the drive device in the right partition in equal proportion, and by increasing the power of the drive device on the upper part of the left wing and reducing the power of the drive device on the lower part of the left wing in equal proportion to offset the lift increment generated by the tilting wing angle of attack after increasing the power of the left drive device. At the same time, increase the power of the drive device on the lower part of the right wing and reduce the power of the drive device on the upper part of the right wing in equal proportion to make up for the lift loss generated by the tilting wing angle of attack after reducing the power of the right drive device; The left yaw control is achieved by increasing the power of the drive device in the right partition and reducing the power of the drive device in the left partition in equal proportion, and by increasing the power of the drive device on the upper part of the right wing and reducing the power of the drive device on the lower part of the left wing in equal proportion to offset the lift increment generated by the tilting wing angle of attack after increasing the power of the right drive device. At the same time, increase the power of the drive device on the lower part of the left wing and reduce the power of the drive device on the upper part of the left wing in equal proportion to make up for the lift loss generated by the tilting wing angle of attack after reducing the power of the left drive device.

[0014] During the horizontal flight phase, the said A-wing aircraft enters the fixed-wing flight mode, and the tilting angles of each section of the tilting wings are all maintained at zero, and the flight is controlled by adjusting the power of the drive device; The pitch control of the A-wing aircraft is achieved by adjusting the power of the drive device to increase or decrease the lift of the wings in the front row partition and at the same time decrease or increase the lift of the wings in the rear row partition; The roll control of the A-wing aircraft is achieved by adjusting the power of the drive device to change the lift difference between the wings in the left and right partitions. The right roll is achieved by increasing the lift of the wings in the left partition and / or reducing the lift of the wings in the right partition, and the left roll is achieved by increasing the lift of the wings in the right partition and / or reducing the lift of the wings in the left partition; The yaw control of the A-wing aircraft is achieved by adjusting the power of the left and right drive devices to change the pull difference between the left and right drive devices.

[0015] The driving device is symmetrically arranged above and below the tilting wing with the tilting axis as the center, dividing the tilting wing into two parts: the tilting wing in front of the tilting axis and the tilting wing behind the tilting axis. The tilting axis is installed in the middle and front of the tilting wing and is located between one-half and one-fourth of the chord length of the tilting wing, so that the tilting wing behind the tilting axis is larger than the tilting wing in front of the tilting axis, ensuring that the tilting wing can automatically rotate upward and level under the action of the rising air current during free fall, realizing gliding forced landing; the driving device generates a layered driving air current on the surface of the tilting wing, and the thickness of the driving air current is less than half of the chord length of the tilting wing. The driving air current generates a negative pressure on the wing surface of the tilting wing in front of the tilting axis and a positive pressure on the wing surface of the tilting wing behind the tilting axis. Since the wing surface of the tilting wing behind is larger than that of the tilting wing in front, and at the same time, due to the small negative pressure of the divergent air current in front of the driving device and the large positive pressure of the convergent air current behind the driving device, the driving air current generates a positive pressure on the wing surface.

[0016] The number of the tilting wings is 2 to n segments, the chord length of the tilting wing is more than 2 times the height of the driving device, and the more the number of segments of the tilting wing, the smaller the chord length, and the more prominent the advantages of the A-wing aircraft; the number of fixed wings is 0 to n + 1 segments, the chord length of the fixed wing is less than 3 times the height of the driving device, and the fewer the number of segments of the fixed wing, the smaller the chord length, and the more prominent the advantages of the A-wing aircraft.

[0017] When the A-wing aircraft controls the attitude by adding or subtracting the lift of the wing, it maintains the stability of the overall power of the tilting wing and the balance of torque by the way of inversely increasing or decreasing the power of the upper driving device and the power of the lower driving device; when controlling the yaw by adding or subtracting the pulling force on the left or right side of the wing, it maintains the stability of the overall power on both sides of the wing and the balance of torque by the way of inversely increasing or decreasing the power of the driving devices in the left front partition and the right rear partition and / or inversely increasing or decreasing the power of the driving devices in the right front partition and the left rear partition.

[0018] The center of gravity of the A-wing aircraft of the present invention is lower than the aerodynamic center of the wing. It adopts a large-area wide-chord wing, reducing the wing loading to less than 10 kg per square meter. Due to the multi-segment tilting reducing the moment of inertia and increasing the tilting speed, the safety height required for the flight mode conversion of the A-wing aircraft is zero, and it can tilt at takeoff, greatly shortening the vertical takeoff and landing time and improving the energy efficiency; when approaching the ground during vertical takeoff and landing, due to the constraint of the large-area wing, the downward driving air current gathers under the large-area wide-chord wing, generating an air cushion effect, which can effectively improve the vertical takeoff and landing ability of the A-wing aircraft at the moment of takeoff and landing, thereby improving the flight efficiency of the A-wing aircraft; the wing configuration of the shutter-type multi-segment tilting has a small moment of inertia, fast downward tilting, and a small side area, unloading the side wind pressure in time and improving the crosswind stability of the large-area wide-chord wing during vertical takeoff and landing.

[0019] The shell-shaped wide-chord wing with a large area and low wing loading of the Jia wing aircraft of the present invention is divided into multiple zones. Driving devices are symmetrically arranged on the upper and lower parts of the shell-shaped tilting wing in each zone, generating multiple pairs of clamping forces acting up and down on the wing surfaces of each zone. Therefore, by adjusting the power of the driving devices, the force state of the wing surfaces can be changed, thereby controlling the flight attitude of the Jia wing aircraft. Since the driving devices are fixedly connected to the tilting wings and tilt synchronously, this control method works continuously along with the tilting of the tilting wings, showing coherence. These forces and moments of force act continuously during the three stages of vertical takeoff and landing, transitional flight, and horizontal flight. The magnitude of their values, the direction of the vectors, and the effect of flight control do not undergo sudden changes, improving the maneuverability and safety of the Jia wing aircraft.

[0020] During the three stages of vertical takeoff and landing, transitional flight, and horizontal flight of the Jia wing aircraft, for pitch, the power of the driving devices is inversely proportional to add or subtract in the front and rear row zones to change the lift of the wing surfaces in the front and rear row zones; for roll, the power of the driving devices is inversely proportional to add or subtract in the left and right side zones to change the lift of the wing surfaces in the left and right side zones; for yaw, the power of the driving devices is inversely proportional to add or subtract in the left and right side zones to change the pulling force in the left and right side zones. The flight control parameters (power of the driving devices) and the generated effects (for pitch, the lift difference of the wing surfaces in the front and rear row zones; for roll, the lift difference of the wing surfaces in the left and right side zones; for yaw, the pulling force difference in the left and right side zones) do not undergo sudden changes due to the change of flight mode and flight stage, showing consistency, continuity, and stability. In contrast, for tilt-rotor aircraft such as the Osprey V-22, if the power of the left and right rotors is inversely proportional to add or subtract, it will roll during vertical takeoff and landing and yaw during horizontal flight, and the control effect undergoes a sudden change. Another example is quadrotor aircraft such as the Ehang 184. If the power of the left and right rotors is inversely proportional to add or subtract, it will yaw during vertical takeoff and landing and roll during horizontal flight, and the control effect also undergoes a sudden change. This sudden change in control effect poses a major hidden danger to flight safety. The Jia wing aircraft fundamentally overcomes this hidden danger.

[0021] The wing of the delta-wing aircraft of the present invention is in the shape of a thin shell and can be simplified to a flat wing without an airfoil. It relies on the lift generated by the angle of attack for flight. Therefore, the lift-to-drag ratio K of the delta-wing aircraft is the ratio of the delta-wing area to the horizontal windward area (drag area) Sd of the delta-wing aircraft, that is, the lift-to-drag ratio K of the delta-wing aircraft is: K = Cl / Cd = S / Sd; the wing loading of the delta-wing aircraft is relatively low, and the delta-wing area is relatively large. Generally, the horizontal windward area (drag area) Sd of the delta-wing aircraft can be made less than one-tenth of the delta-wing area, and the lift-to-drag ratio of the delta-wing aircraft can reach more than 10, that is: Sd ≤ 0.1S, K ≥ 10; the aspect ratio λ of the delta wing is the ratio of the wingspan l to the chord b, that is: λ = l / b. Generally, the aspect ratio λ of the delta-wing aircraft ≤ 1. At this time, the induced drag is relatively large. The induced drag is reduced by the large V-shaped winglets to form a V-shaped stabilizer; the delta-wing aircraft also has a characteristic parameter, the straightness ratio δ (delta). The straightness ratio δ of the delta-wing aircraft is the ratio of the wingspan l of the delta-wing aircraft to the diameter d of the drive device of the delta-wing aircraft, that is: δ = l / d, which represents the ratio relationship between the thickness of the laminar flow drive and the wing of the delta-wing aircraft. The larger the straightness ratio, the thinner the driving air flow, and the stronger the characteristics of the delta-wing aircraft. Generally, the straightness ratio of the delta-wing aircraft is 6 ≤ δ ≤ 20.

[0022] Therefore, compared with the existing aircraft, the present invention significantly improves the flight control ability of the delta-wing aircraft, can fully take into account both vertical takeoff and landing and horizontal flight conditions, realizes a smooth transition and continuous adjustment of flight control, and has the advantages of natural mode conversion, high flight efficiency, low wing loading, safety and stability, simple structure, easy operation, low cost, and easy popularization. Brief Description of the Drawings

[0023] Figure 1 It is a front view schematic diagram of the horizontal flight state of the delta-wing aircraft of the present invention.

[0024] Figure 2 It is a side view schematic diagram of the horizontal flight state of the delta-wing aircraft of the present invention.

[0025] Figure 3 It is a top view schematic diagram of the horizontal flight state of the delta-wing aircraft of the present invention.

[0026] Figure 4 It is a front view schematic diagram of the vertical takeoff and landing state of the delta-wing aircraft of the present invention.

[0027] Figure 5 It is a side view schematic diagram of the vertical takeoff and landing state of the delta-wing aircraft of the present invention.

[0028] Figure 6 It is a top view schematic diagram of the vertical takeoff and landing state of the delta-wing aircraft of the present invention.

[0029] Figure 7 It is a four-zone schematic diagram of the delta-wing aircraft of the present invention.

[0030] Figure 8 It is a schematic diagram of the control principle for the upward movement during the transitional flight of the Type A wing aircraft of the present invention.

[0031] Figure 9 It is a schematic diagram of the control principle for the downward movement during the transitional flight of the Type A wing aircraft of the present invention.

[0032] Figure 10 It is a schematic diagram of the control principle for the right yaw movement during the transitional flight of the Type A wing aircraft of the present invention.

[0033] Figure 11 It is a schematic diagram of the control principle for the left yaw movement during the transitional flight of the Type A wing aircraft of the present invention.

[0034] Figure 12 It is a schematic diagram of the control principle for the right roll movement during the transitional flight of the Type A wing aircraft of the present invention.

[0035] Figure 13 It is a schematic diagram of the control principle for the left roll movement during the transitional flight of the Type A wing aircraft of the present invention.

[0036] Figure 14 It is a schematic diagram of the control principle for the upward movement during the horizontal flight of the Type A wing aircraft of the present invention.

[0037] Figure 15 It is a schematic diagram of the control principle for the downward movement during the horizontal flight of the Type A wing aircraft of the present invention.

[0038] Figure 16 It is a schematic diagram of the control principle for the right yaw movement during the horizontal flight of the Type A wing aircraft of the present invention.

[0039] Figure 17 It is a schematic diagram of the control principle for the left yaw movement during the horizontal flight of the Type A wing aircraft of the present invention.

[0040] Figure 18 It is a schematic diagram of the control principle for the right roll movement during the horizontal flight of the Type A wing aircraft of the present invention.

[0041] Figure 19 It is a schematic diagram of the control principle for the left roll movement during the horizontal flight of the Type A wing aircraft of the present invention.

[0042] Figure 20 It is a schematic diagram of the shell-shaped tilting wing and drive device of the Type A wing aircraft of the present invention.

[0043] Figure 21 It is a side view schematic diagram of the single-seat sports Type A wing aircraft of Embodiment 2 of the present invention.

[0044] Figure 22 It is a top view schematic diagram of the single-seat sports Type A wing aircraft of Embodiment 2 of the present invention.

[0045] Figure 23 It is a side view schematic diagram of the three-seat family Type A wing aircraft of Embodiment 3 of the present invention.

[0046] Figure 24It is a top-down schematic view of the three-seat family-type delta-wing aircraft in Embodiment 3 of the present invention.

[0047] Figure 25 It is a side view schematic of the five-seat business-type delta-wing aircraft in Embodiment 4 of the present invention.

[0048] Figure 26 It is a top-down schematic view of the five-seat business-type delta-wing aircraft in Embodiment 4 of the present invention.

[0049] Figure 27 It is a three-zone schematic diagram of Embodiment 5 of the present invention.

[0050] Figure 28 It is a sixteen-zone schematic diagram of Embodiment 6 of the present invention.

[0051] In the figure: 1 - fuselage, 2 - wing, 3 - drive device, 4 - tilting wing, 5 - fixed wing, 6 - tilting mechanism, 7 - winglet, 8 - landing gear, 9 - nose shell-shaped fixed wing, 10 - tail shell-shaped fixed wing, 11 - aerodynamic center of the wing, 12 - left front zone, 13 - left rear zone, 14 - right front zone, 15 - right rear zone, 16 - left front tilting mechanism, 17 - left rear tilting mechanism, 18 - right front tilting mechanism, 19 - right rear tilting mechanism, 20 - tilting axis, 21 - tilting wing at the front of the tilting axis, 22 - tilting wing at the rear of the tilting axis, 23 - upper drive device, 24 - lower drive device, 25 - V-shaped wing, 26 - vertical tail. Detailed implementation manners

[0052] The following further describes in detail the specific implementation manners of the present invention in conjunction with the accompanying drawings. For technologies or products not specified in the embodiments, they are all existing technologies or conventional products that can be obtained by purchase. Structural features not yet described can all be implemented using conventional structures.

[0053] Embodiment 1: As Figure 1-20 shown, for the method of realizing the flight control of the delta-wing aircraft, the delta-wing aircraft has a large-area shell-shaped wide-chord wing. The upper and lower parts of the shell-shaped wing are respectively fixedly installed with drive devices. The drive devices are arranged in 2 rows along the chord direction of the wing and are densely distributed along the span direction of the wing. The thickness of the laminar drive airflow generated by the drive devices on the wing surface is less than 0.2 times the chord length of the large-area shell-shaped wide-chord wing. The large-area shell-shaped wide-chord wing is designed as a combined type, transversely divided into 4 tilting wings and 5 fixed wings. With the aerodynamic center of the large-area shell-shaped wide-chord wing as the origin, the wing and the drive devices are divided into four zones. Both the left front zone and the right front zone include 2 tilting wings and 2 fixed wings. Both the left rear zone and the right rear zone include 2 tilting wings and 3 fixed wings. By adjusting the power of the upper and / or lower drive devices of the shell-shaped wing in each zone, the force on the wing in this zone is changed to adjust the flight attitude of the large-area shell-shaped wide-chord wing, thereby realizing the flight control of the delta-wing aircraft.

[0054] The drive device adopts distributed drive, which consists of small-diameter motor propellers. It forms a layered driving air flow along the wing surface. The layered driving air flow under the wing surface generates pressure difference lift and forced downwash lift, improving the flight efficiency. The small-diameter propellers are paired with even smaller motors to improve the thrust-to-weight ratio. The technical route of "ants moving house" is adopted to set up hundreds of shafts of propellers to solve the contradiction between small propellers and large payload, while increasing the propeller solidity, reducing the disc loading, significantly reducing the vibration and noise of the armor-wing aircraft, and improving the load capacity and flight efficiency of the armor-wing aircraft.

[0055] The drive device is symmetrically arranged above and below the tilt wing with the tilt axis as the center. The tilt axis is installed in the front part of the middle of the tilt wing and at 0.45 of the chord length of the tilt wing, making the tilt wing behind the tilt axis larger than the tilt wing in front of the tilt axis, ensuring that the tilt wing can automatically rotate upward and level under the action of the rising air flow during free fall to achieve gliding forced landing. The drive device generates a layered driving air flow on the surface of the tilt wing, producing a positive pressure on the wing surface.

[0056] All the drive devices in each partition rotate in the same direction. Many tilt wings in each partition are synchronously driven by a set of tilt mechanisms. The upper drive devices in each partition are controlled by one power signal, and the lower drive devices in each partition are controlled by another power signal. Each partition has a total of 3 flight control signals, namely 1 tilt signal, 1 upper drive device power signal, and 1 lower drive device power signal. The armor-wing aircraft controls flight by adjusting the tilt signals and power signals of each partition to change the forces on each part of the upper and lower surfaces of the wing.

[0057] During the vertical takeoff and landing phase of the A-wing aircraft, the tilt angle of each tilting wing section is tilted downward by 90°, and the driving device is vertically upward. Each partition is controlled according to a quadcopter. By changing the power of the driving device in each partition, the lift of each partition is adjusted while balancing the torque to control the flight. During the transition flight phase of the A-wing aircraft, the flight is controlled by adjusting the tilt angle of each partition and the power of the upper and lower driving devices. During the takeoff transition flight phase, the tilting wings gradually rotate upward, and the driving device gradually rotates horizontally forward. In the high-power state of the quadcopter, the driving device generates a forward pulling force to drive the A-wing aircraft forward. The tilting wing with a positive angle of attack moving forward generates an angle of attack lift, and the angle of attack lift is greater than the loss of vertical lift caused by the forward tilt of the driving device. The A-wing aircraft obtains a tendency to move upward. At the same time, the different airflow states in front of each tilting wing section cause the angle of attack lift obtained by the front-row tilting wings to be greater than that of the rear-row tilting wings, and the A-wing aircraft generates a pitching moment. Therefore, during the takeoff transition flight phase, the control needs to gradually reduce the power of the driving device in each partition and finally transition to horizontal flight control. During the landing transition flight phase, the tilting wings gradually tilt downward, and the driving device gradually moves vertically upward. The inertia of horizontal flight keeps the A-wing aircraft moving forward. The tilting wing with a positive angle of attack moving forward generates an angle of attack lift. The different airflow states in front of each tilting wing section cause the angle of attack lift obtained by the front-row tilting wings to be greater than that of the rear-row tilting wings, and the A-wing aircraft generates a pitching moment. In the low-power state of horizontal flight, the insufficient lift obtained by the A-wing aircraft leads to a tendency to move downward. Therefore, during the landing transition flight phase, the control needs to gradually increase the power of the driving device in each partition and finally transition to quadcopter control.

[0058] During the transition flight phase of the A-wing aircraft, the pitch control of the A-wing aircraft is achieved by changing the lift difference between the front-row and rear-row partition wings. The upward pitch control is achieved by increasing the power of the driving device in the front-row partition and reducing the power of the driving device in the rear-row partition in equal proportion. The downward pitch control is achieved by reducing the power of the driving device in the front-row partition and increasing the power of the driving device in the rear-row partition in equal proportion.

[0059] During the transition flight phase of the A-wing aircraft, the roll control of the A-wing aircraft is achieved by adjusting the lift difference between the left and right wings. The right roll control of the A-wing aircraft is achieved by increasing the power of the lower driving device in the left partition and reducing the power of the upper driving device in equal proportion, while reducing the power of the lower driving device in the right partition and increasing the power of the upper driving device in equal proportion. The left roll control of the A-wing aircraft is achieved by reducing the power of the lower driving device in the left partition and increasing the power of the upper driving device in equal proportion, while increasing the power of the lower driving device in the right partition and reducing the power of the upper driving device in equal proportion.

[0060] During the transition flight phase of the A-wing aircraft, the yaw control of the A-wing aircraft is achieved by adjusting the pulling force difference between the left and right drive devices. The right yaw control of the A-wing aircraft is achieved by increasing the power of the drive device in the left partition while reducing the power of the drive device in the right partition in equal proportion, and by increasing the power of the upper drive device on the left and reducing the power of the lower drive device to offset the lift increment, while increasing the power of the lower drive device on the right and reducing the power of the upper drive device to make up for the lift loss; the left yaw control is achieved by increasing the power of the drive device in the right partition while reducing the power of the drive device in the left partition in equal proportion, and by increasing the power of the upper drive device on the right and reducing the power of the lower drive device to offset the lift increment, while increasing the power of the lower drive device on the left and reducing the power of the upper drive device to make up for the lift loss.

[0061] During the horizontal flight phase, the A-wing aircraft enters the fixed-wing flight mode, and the tilt angles of each tilt wing are all kept at zero, and the flight is controlled by adjusting the power of the drive device; the pitch control of the A-wing aircraft is achieved by increasing or reducing the wing lift in the front row partition while reducing or increasing the wing lift in the rear row partition; the roll control of the A-wing aircraft is achieved by changing the power of the drive device to adjust the wing lift difference between the left and right partitions. The right roll is achieved by increasing the wing lift in the left partition and / or reducing the wing lift in the right partition, and the left roll is achieved by increasing the wing lift in the right partition and / or reducing the wing lift in the left partition; the yaw control of the A-wing aircraft is achieved by changing the power of the left and right drive devices to adjust the pulling force difference between the left and right drive devices.

[0062] When the A-wing aircraft controls its attitude by increasing or decreasing the wing lift, the overall power of the tilt wing is kept stable and the torque is balanced by the inverse proportional increase and decrease of the power of the upper drive device and the lower drive device; when controlling the yaw by increasing or decreasing the pulling force on the left or right side of the wing, the overall power on the left and right sides of the wing is kept stable and the torque is balanced by the inverse proportional increase and decrease of the power of the drive devices in the left front partition and the right rear partition and / or the inverse proportional increase and decrease of the power of the drive devices in the right front partition and the left rear partition.

[0063] Example 2: As Figure 1-22As shown in the figure, the method for realizing the flight control of the A-wing aircraft. The single-seat sports A-wing aircraft has a large-area shell-shaped wide-chord wing, and drive devices are fixedly installed on the upper and lower parts of the shell-shaped wing respectively. With the aerodynamic center of the large-area shell-shaped wide-chord wing as the origin, the wing and the drive devices are divided into four zones. The large-area shell-shaped wide-chord wing is designed as a combined type, horizontally divided into 2 tilt wings and 3 fixed wings. The left front zone and the right front zone both include 1 tilt wing and 2 fixed wings, and the left rear zone and the right rear zone both include 1 tilt wing and 1 fixed wing. The aspect ratio of the single-seat sports A-wing aircraft is λ = l / b = 2, and the chord length of the tilt wing is 1.5 meters. One row of propellers with a diameter of 0.7 meters is arranged on the upper and lower parts of each tilt wing, and the number of propellers in each row is 6, with a total of 48 propellers. The propellers adopt FLUXER CB2 ultra-light propellers 2812, and the weight of each propeller is 50g. Solar cells are arranged on the A-wing, using single-crystal thin-silicon solar photovoltaic cells, with silver-plated wires on the surface and PET+EV encapsulation. The rest is the same as that of Embodiment 1.

[0064] Embodiment 3: As Figure 1-20 shown in FIGS. 23-24, the method for realizing the flight control of the A-wing aircraft. The 3-seat family A-wing aircraft has a large-area shell-shaped wide-chord wing, and drive devices are fixedly installed on the upper and lower parts of the shell-shaped wing respectively. With the aerodynamic center of the large-area shell-shaped wide-chord wing as the origin, the wing and the drive devices are divided into four zones. The large-area shell-shaped wide-chord wing is designed as a combined type, horizontally divided into 6 tilt wings and 3 fixed wings. The left front zone and the right front zone both include 3 tilt wings and 2 fixed wings, and the left rear zone and the right rear zone both include 3 tilt wings and 1 fixed wing. The aspect ratio of the 3-seat family A-wing aircraft is λ = l / b = 0.7, and the chord length of the tilt wing is 1.5 meters. One row of propellers with a diameter of 0.7 meters is arranged on the upper and lower parts of each tilt wing, and the number of each row is 5, with a total of 120 propellers. The propellers adopt FLUXER CB2 ultra-light propellers 2812, and the weight of each propeller is 50g. Solar cells are arranged on the A-wing, using single-crystal thin-silicon solar photovoltaic cells, with silver-plated wires on the surface and PET+EV encapsulation. The rest is the same as that of Embodiment 1.

[0065] Embodiment 4: As Figure 1-20As shown in FIGS. 25 - 26, the method for controlling the flight of the Jia wing aircraft. The 5 - seat business - type Jia wing aircraft has a large - area shell - shaped wide - chord wing, and driving devices are fixedly installed on the upper and lower parts of the shell - shaped wing respectively. With the aerodynamic center of the large - area shell - shaped wide - chord wing as the origin, the wing and the driving devices are divided into four zones. The large - area shell - shaped wide - chord wing is designed as a combined type, horizontally divided into 6 tilt - wing segments and 3 fixed - wing segments. The left - front zone and the right - front zone both include 3 tilt - wing segments and 2 fixed - wing segments, and the left - rear zone and the right - rear zone both include 3 tilt - wing segments and 1 fixed - wing segment. The aspect ratio of the 5 - seat business - type Jia wing aircraft is λ = l / b = 1, and the chord length of the tilt - wing is 1.5 m. One row of propellers with a diameter of 0.7 m is arranged on the upper and lower parts of each tilt - wing, with 8 propellers in each row, and a total of 192 propellers are provided. The propellers adopt FLUXER CB2 ultra - light propellers 2812, and the weight of a single propeller is 50 g. Solar cells are installed on the Jia wing, using single - crystal thin - silicon solar photovoltaic cells, with silver - plated wires on the surface and PET + EV encapsulation. The rest is the same as in Embodiment 1.

[0066] Embodiment 5: As Figure 1-20 As shown in FIGS. 27, the method for controlling the flight of the Jia wing aircraft. With the aerodynamic center of the large - area shell - shaped wide - chord wing as the origin, the wing and the driving devices are first divided into four zones. On the basis of the four zones, the two rear zones are combined into one zone, and a total of 3 zones are divided. The large - area shell - shaped wide - chord wing is designed as a combined type, horizontally divided into 3 tilt - wing segments, and each zone only includes 1 tilt - wing segment. The driving devices adopt distributed driving, which are composed of ducted fans. By adjusting the power of the driving devices on the upper part and / or the lower part of the shell - shaped wing in each zone, the force on the upper and lower surfaces of the wing in each zone is changed, so as to adjust the flight attitude of the large - area shell - shaped wide - chord wing and realize the flight control of the Jia wing aircraft. The rest is the same as in Embodiment 1.

[0067] Embodiment 6: As Figure 1-20As shown in Figure 28, in the method for controlling the flight of the first wing aircraft, the driving devices are arranged in 8 rows along the chord direction of the wing and are densely distributed along the span direction of the wing. With the aerodynamic center of the large-area shell-shaped wide-chord wing as the origin, the wing and the driving devices are divided into sixteen zones, namely the left front zone 1, the left front zone 2, the left front zone 3, the left front zone 4, the right front zone 1, the right front zone 2, the right front zone 3, the right front zone 4, the left rear zone 1, the left rear zone 2, the left rear zone 3, the left rear zone 4, the right rear zone 1, the right rear zone 2, the right rear zone 3, and the right rear zone 4. The large-area shell-shaped wide-chord wing is designed as a combined type, horizontally divided into 8 tilting wings and 9 fixed wings. Both the left front zone 1 and the right front zone 1 include 1 tilting wing and 2 fixed wings, and the remaining zones all include 1 tilting wing and 1 fixed wing. The driving devices adopt distributed driving and are composed of air nozzles. By adjusting the power of the driving devices on the upper part and / or the lower part of the shell-shaped wing in each zone, the force on the upper and lower surfaces of the wing in each zone is changed, so as to adjust the flight attitude of the large-area shell-shaped wide-chord wing and realize the flight control of the first wing aircraft. The rest is the same as in Embodiment 1.

[0068] The technical content of the present invention has been described above in conjunction with the accompanying drawings, but the protection scope of the present invention is not limited to the above content. Within the scope of knowledge possessed by those of ordinary skill in the art, various changes can be made to the technical content of the present invention without departing from the purpose of the present invention. Any modifications, equivalent replacements, improvements, etc. made within the spirit and principle of the present invention shall be included within the protection scope of the present invention.

Claims

1. A method for implementing flight control of a wing-in-ground effect vehicle, characterized in that: The first wing aircraft has a large-area shell-shaped wide-chord wing. Driving devices are fixedly installed on the upper and lower parts of the shell-shaped wing respectively. The driving air flow generated by the upper driving device flows along the upper surface of the shell-shaped wing, and the driving air flow generated by the lower driving device flows along the lower surface of the shell-shaped wing. The driving devices are arranged in multiple rows along the chord direction of the wing and are densely distributed along the span direction of the wing. The driving devices generate layered driving air flow on the wing surface, and the thickness of the layered driving air flow is less than one-fourth of the chord length of the large-area shell-shaped wide-chord wing. Taking the aerodynamic center of the large-area shell-shaped wide-chord wing as the origin, the wing and the driving devices are divided into several partitions with numerous driving devices. By adjusting the power of the upper driving device and / or the lower driving device of the shell-shaped wing in a certain partition, that is, adjusting the pressure of the layered driving air flow on the upper surface and / or the pressure of the layered driving air flow on the lower surface of the shell-shaped wing in this partition, the force on the upper wing surface and / or the force on the lower wing surface of the shell-shaped wing in this partition can be changed. By changing the forces on the upper and lower surfaces of each partition of the wing, the flight attitude of the large-area shell-shaped wide-chord wing is adjusted to achieve the flight control of the first wing aircraft. The large-area shell-shaped wide-chord wing is designed as a combined type, which is transversely divided into several tilt wings and several fixed wings. Each partition includes several tilt wings and several fixed wings. All the tilt wings in each partition are connected to the fuselage through the same tilting mechanism and are controlled by the same tilting signal. The tilt wings in each partition are respectively controlled by their own tilting signals and are driven by their own tilting mechanisms. The tilt wings and the fixed wings are arranged alternately like zebra crossings. Through the tilting of the tilt wings, they can be combined with the fixed wings as a whole to achieve horizontal flight, and the tilt wings can be opened like a shutter to form a strip-shaped air flow channel to achieve vertical takeoff and landing. Driving devices are respectively arranged along the span direction on the upper and lower parts of the tilt wings, and the driving devices tilt synchronously with the tilt wings. During the vertical takeoff and landing stage, the tilt wings tilt downward to form a vertical strip-shaped air flow channel on the large-area shell-shaped wide-chord wing. The driving devices on each section of the tilt wings are connected in parallel front and back. Air is inhaled from above the fixed wings and exhausted below the fixed wings through the strip-shaped air flow channel, and each partition is controlled to fly in the quadrotor mode. During the transition flight stage and the horizontal flight stage, by adjusting the tilting angle of each partition and the intensity of the layered driving air flow on the upper and lower surfaces of each tilt wing surface, the lift of each partition of the wing can be changed to achieve the flight control of the first wing aircraft.

2. The method for realizing the flight control of the A-wing aircraft according to claim 1, characterized in that: The partition method is four partitions. Taking the aerodynamic center of the large-area shell-shaped wide-chord wing as the origin, the numerous driving devices are divided into 4 partitions, namely the left front partition, the left rear partition, the right front partition and the right rear partition.

3. The method for realizing the flight control of the A-wing aircraft according to claim 2, characterized in that: The driving device adopts distributed drive and is composed of numerous small-diameter motor propellers or ducted fans or jet engines or air nozzles or small engines. All the driving devices in each partition rotate in the same direction. All the numerous tilt wings in each partition tilt synchronously under the control of the same tilting signal. When divided into four partitions, there are 4 independent tilting mechanisms respectively controlled by 4 tilting signals. The upper driving device in each partition is controlled by one power signal, and the lower driving device in each partition is controlled by another power signal. In summary, each partition has three flight control signals, namely one tilting signal, one power signal for the upper driving device, and one power signal for the lower driving device. The first wing aircraft controls flight by adjusting the tilting signals and power signals of each partition to change the forces on each part of the upper and lower surfaces of the wings.

4. The method for realizing the flight control of the alpha-wing aircraft according to claim 2, characterized in that: During the vertical takeoff and landing phase of the first wing aircraft, the tilting angle of each tilting wing section is tilted downward by 90°, and the driving device is vertically upward. Each partition is controlled according to a quadcopter. By changing the power of the driving device in each partition, the lift of each partition is adjusted while balancing the torque to control flight. During the transition flight phase of the first wing aircraft, flight can be controlled by adjusting the tilting angle and the power of the driving device in each partition. During the takeoff transition flight phase, the tilting wings gradually rotate upward, and the driving device gradually rotates horizontally forward. In the high-power state of the quadcopter, the driving device generates a forward pulling force to drive the first wing aircraft forward. The tilting wings with a positive angle of attack generate an angle-of-attack lift when moving forward. The angle-of-attack lift is greater than the loss of vertical lift caused by the forward tilt of the driving device, so the first wing aircraft has a tendency to move upward. At the same time, the different air flow states in front of each tilting wing section cause the angle-of-attack lift obtained by the front-row tilting wings to be greater than that of the rear-row tilting wings, and the first wing aircraft generates a nose-up moment. Therefore, during the takeoff transition flight phase, the control needs to gradually reduce the power of the driving device in each partition and finally transition to horizontal flight control. During the landing transition flight phase, the tilting wings gradually tilt downward, and the driving device gradually turns vertically upward. The inertia of horizontal flight keeps the first wing aircraft moving forward. The tilting wings with a positive angle of attack generate an angle-of-attack lift when moving forward. The different air flow states in front of each tilting wing section cause the angle-of-attack lift obtained by the front-row tilting wings to be greater than that of the rear-row tilting wings, and the first wing aircraft generates a nose-up moment. In the low-power state of horizontal flight, the insufficient lift obtained by the first wing aircraft leads to a tendency to move downward. Therefore, during the landing transition flight phase, the control needs to gradually increase the power of the driving device in each partition and finally transition to quadcopter control.

5. The method for realizing the flight control of the A-wing aircraft according to claim 2, wherein: During the transition flight phase of the first wing aircraft, the pitch control of the first wing aircraft is achieved by changing the power of the driving devices in the front-row and rear-row partitions to adjust the lift difference between the wings in the front-row and rear-row partitions. The up-pitch control is achieved by increasing the power of the driving device in the front-row partition and reducing the power of the driving device in the rear-row partition in equal proportion. The down-pitch control is achieved by reducing the power of the driving device in the front-row partition and increasing the power of the driving device in the rear-row partition in equal proportion. During the transition flight phase of the first wing aircraft, the roll control of the first wing aircraft is achieved by changing the power of the driving devices in the left and right partitions to adjust the lift difference between the wings on the left and right sides. The right roll control of the first wing aircraft is achieved by increasing the power of the lower driving device in the left partition and reducing the power of the upper driving device in the left partition in equal proportion, while reducing the power of the lower driving device in the right partition and increasing the power of the upper driving device in the right partition in equal proportion. The left roll control of the first wing aircraft is achieved by reducing the power of the lower driving device in the left partition and increasing the power of the upper driving device in the left partition in equal proportion, while increasing the power of the lower driving device in the right partition and reducing the power of the upper driving device in the right partition in equal proportion. During the transition flight phase of the first winged aircraft, the yaw control of the first winged aircraft is achieved by changing the power of the driving devices in the left and right partitions to adjust the difference in pulling force between the left and right driving devices. The right yaw control of the first winged aircraft is achieved by increasing the power of the driving device in the left partition while proportionally reducing the power of the driving device in the right partition, and by increasing the power of the driving device above the left wing while proportionally reducing the power of the driving device below the left wing to offset the lift increment generated by the tilting wing angle of attack after increasing the power of the left driving device. At the same time, the power of the driving device below the right wing is increased while the power of the driving device above the right wing is proportionally reduced to compensate for the lift loss generated by the tilting wing angle of attack after reducing the power of the right driving device. The left yaw control is achieved by increasing the power of the driving device in the right partition while proportionally reducing the power of the driving device in the left partition, and by increasing the power of the driving device above the right wing while proportionally reducing the power of the driving device below the left wing to offset the lift increment generated by the tilting wing angle of attack after increasing the power of the right driving device. At the same time, the power of the driving device below the left wing is increased while the power of the driving device above the left wing is proportionally reduced to compensate for the lift loss generated by the tilting wing angle of attack after reducing the power of the left driving device.

6. The method for implementing flight control of the A-wing aircraft according to claim 2, characterized in that: During the horizontal flight phase, the first winged aircraft enters the fixed-wing flight mode, and the tilting angles of all tilting wings are kept at zero, and the flight is controlled by adjusting the power of the driving devices. The pitch control of the first winged aircraft is achieved by adjusting the power of the driving devices to increase or decrease the lift of the wings in the front row partition while decreasing or increasing the lift of the wings in the rear row partition. The roll control of the first winged aircraft is achieved by adjusting the power of the driving devices to change the lift difference between the wings in the left and right partitions. The right roll is achieved by increasing the lift of the wings in the left partition and / or decreasing the lift of the wings in the right partition. The left roll is achieved by increasing the lift of the wings in the right partition and / or decreasing the lift of the wings in the left partition. The yaw control of the first winged aircraft is achieved by adjusting the power of the left and right driving devices to change the difference in pulling force between the left and right driving devices.

7. The method for implementing the flight control of the A-wing aircraft according to claim 2, wherein: The driving devices are symmetrically arranged above and below the tilting wing with the tilting axis as the center, dividing the tilting wing into two parts: the tilting wing in front of the tilting axis and the tilting wing behind the tilting axis. The tilting axis is installed in the middle and front of the tilting wing and is located between one-half and one-fourth of the chord length of the tilting wing, so that the tilting wing behind the tilting axis is larger than the tilting wing in front of the tilting axis, ensuring that the tilting wing can automatically rotate upward and level under the action of the updraft during free fall to achieve gliding and forced landing. The driving devices generate layered driving airflows on the surface of the tilting wing. The thickness of the driving airflow is less than half of the chord length of the tilting wing. The driving airflow generates negative pressure on the wing surface of the tilting wing in front of the tilting axis and positive pressure on the wing surface of the tilting wing behind the tilting axis. Since the wing surface of the tilting wing behind is larger than that of the tilting wing in front, and at the same time, due to the divergence of the airflow in front of the driving device with a small negative pressure and the convergence of the airflow behind the driving device with a large positive pressure, the driving airflow generates positive pressure on the wing surface.

8. The method for realizing the flight control of the Jia wing aircraft according to claim 2, characterized in that: The number of the tilting wings is 2 to n segments. The chord length of the tilting wings is greater than 2 times the height of the driving device. The more the number of segments of the tilting wings, the smaller the chord length, and the more prominent the advantages of the A-wing aircraft. The number of the fixed wings is 0 to n+1 segments. The chord length of the fixed wings is less than 3 times the height of the driving device. The fewer the number of segments of the fixed wings, the smaller the chord length, and the more prominent the advantages of the A-wing aircraft.

9. The method for realizing the flight control of the Jia wing aircraft according to any one of claims 1 to 8, characterized in that: When the A-wing aircraft controls the attitude by adding or subtracting the wing lift, the overall power of the tilting wings is kept stable and the torque is balanced by the inverse proportional increase and decrease of the power of the driving device on the upper part of the wing and the power of the driving device on the lower part of the wing. When controlling the yaw by adding or subtracting the pulling force on the left or right side of the wing, the overall power on both sides of the wing is kept stable and the torque is balanced by the inverse proportional increase and decrease of the power of the driving devices in the left front partition and the right rear partition and / or the inverse proportional increase and decrease of the power of the driving devices in the right front partition and the left rear partition.

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