Fixed-wing aircraft and attitude control method thereof

Abstract

The present invention discloses a fixed-wing aircraft, including an aircraft body, on which a flight control computer and an angle of attack sensor, a sideslip angle sensor and a slope sensor connected to the flight control computer are provided; main wings are symmetrically installed on both sides of the aircraft body, and their leading edges are swept back; the trailing edge of the main wing is swept forward; aerodynamic control surfaces whose deflection is controlled by the flight control computer are symmetrically installed at the trailing edge positions of the main wings on both sides; tail fins whose rotation is controlled by the flight control computer are symmetrically installed at both sides of the tail of the aircraft body, and the rotation axis of the tail fin is swept back; the sweep angle of the rotation axis of the tail fin is greater than or equal to the sweep angle of the leading edge of the main wing, and the chord plane where the tail fin is located is parallel to the chord plane where the leading edge of the main wing is located, or the angle between the two is less than 15°. The present invention can achieve effective yaw stabilization and yaw control of fixed-wing aircraft without vertical tail through aerodynamic surface control in all the aforementioned modes.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft, and in particular to a fixed-wing aircraft without a vertical tail and an attitude control method thereof. Background Art

[0002] The successfully manufactured fixed-wing aircraft without vertical tail and the related technical solutions that have been made public are mainly for military use. The main advantages of aircraft without vertical tail are: 1. After the vertical tail, which is the radar reflection plane, is removed, the appearance of the aircraft is simplified, and the RCS value of the aircraft is effectively reduced; 2. After the vertical tail is removed, the drag source is reduced, and the lift-to-drag ratio of the whole aircraft is improved.

[0003] Products that have been successfully manufactured and technical solutions have been proposed, such as the US stealth bomber B-2, B-21, X-36 (technical verification aircraft), X47B (unmanned aerial vehicle), etc. After the vertical tail that provides yaw stability and yaw control torque for the aircraft during flight is removed, the yaw stability and yaw control torque of the aircraft during flight will be equivalently provided by other devices. The common technical measures of the above-mentioned aircraft are that a split drag rudder is set at the trailing edge of the outer end of the aircraft's main wing and parallel to the leading edge of the main wing. The lower part of the drag rudder is deflected downward, and the upward torque generated is offset by the downward torque generated by the upward deflected part of the drag rudder. The drag torques generated by the upper and lower parts of the drag rudder jointly generate a yaw torque around the center of gravity of the aircraft to achieve yaw stability and yaw control. The limitation of this technical solution is that when the aircraft is in a steady cruise flight and a small-bank yaw maneuvering mode, its angle of attack is at a relatively small state, and the upward deflected part of the drag rudder is in the unseparated, high-velocity incoming flow at the rear of the upper surface of the wing, and the aerodynamic efficiency is relatively high, and sufficient downward torque and drag torque can be generated, so the drag rudder can generate the control torque required to achieve yaw control. When the design requirements of a large-angle-of-attack and large-bank maneuvering aircraft are put forward for the aircraft, the split drag rudder will not be able to effectively provide yaw stability and control torque, and the reason is that when the aircraft is in a large-angle-of-attack mode, especially when it is close to the stall angle of attack, the incoming flow at the rear of the upper surface of the wing partially becomes turbulent, and the speed pressure of the upward deflected part of the drag rudder is reduced, and it is unable to generate sufficient downward torque and effectively balance the upward torque generated by the downward deflected part of the drag rudder, so the split drag rudder cannot achieve effective yaw control of the aircraft in the large-angle-of-attack mode. Furthermore, when the aircraft is in the post-stall mode, the split drag rudder is completely ineffective and the attitude control of the aircraft can only be achieved through vector thrust. Summary of the invention

[0004] Purpose of the invention: In order to overcome the functional limitations of the above-mentioned prior art solutions in terms of yaw stabilization and maneuverability, the present invention provides a fixed-wing aircraft and an attitude control method thereof which can achieve effective yaw stabilization and maneuverability through aerodynamic surface control in all the above-mentioned modes.

[0005] Technical solution:

[0006] A fixed-wing aircraft, an aircraft body, on which a flight control computer and an angle of attack sensor, a sideslip angle sensor and a slope sensor connected to the flight control computer are arranged; main wings of the aircraft body are symmetrically mounted on both sides of the aircraft body, and the leading edges of the main wings are swept back;

[0007] The trailing edge of the main wing is swept forward; aerodynamic control surfaces whose deflection is controlled by a flight control computer are symmetrically installed at the trailing edge positions of the main wing on both sides; tail fins whose rotation is controlled by a flight control computer are symmetrically installed at both sides of the tail of the aircraft body, and the rotation axis of the tail fin is swept backward;

[0008] The sweep angle of the tail fin rotation axis is greater than or equal to the sweep angle of the main wing leading edge, and the chord plane where the tail fin is located is parallel to the chord plane where the main wing leading edge is located, or the angle between the two is less than 15°;

[0009] The rotation axis of the tail fin is arranged between the leading edge and the trailing edge of the tail fin.

[0010] The aerodynamic control surface is a flaperon, and the rotation axis of the flaperon is arranged on the hinge line of the main wing.

[0011] The aerodynamic control surfaces adopt flaps and ailerons. Ailerons are rotatably installed at positions of the trailing edges of the left and right sides of the main wing away from the aircraft body, and flaps are rotatably installed at positions of the trailing edges of the left and right sides of the main wing close to the aircraft body. At the same time, actuators for controlling the rotation of the flaps and the ailerons are installed in the aircraft body, and deflection sensors for collecting the rotation angles of the flaps and the ailerons are installed in the aircraft body.

[0012] The leading edge of the tail fin is parallel to its trailing edge, and its outer edge is parallel to its inner edge; the outer edge of the tail fin is parallel to the edge strip of the aircraft body, and the leading edge of the main wing is parallel to the leading edge of the tail fin.

[0013] A method for attitude control of a fixed-wing aircraft, comprising:

[0014] (1) Yaw stability:

[0015] (11) When the aircraft is in the straight-ahead mode with its angle of attack less than the stall angle of attack and the flight control computer is not sending control command signals to the control surface actuators:

[0016] (111) When the sideslip angle acceleration of the aircraft acquired by the sideslip angle sensor is zero, the aerodynamic control surface on the main wing does not deflect, and at this time, the tail fin is controlled by the flight control computer to be in a real-time zero angle of attack state with respect to the local flow field;

[0017] (112) When the aircraft is disturbed and generates sideslip angle acceleration, the sideslip angle sensor transmits the real-time deflection angle acceleration signal collected by the sensor to the flight control computer;

[0018] The flight control computer calculates the rotation angle of the corresponding side tail fin and the corresponding side aerodynamic control surface based on the calculation, and controls the corresponding side tail fin based on the calculation, so that its upper surface is in a negative angle of attack posture in the local flow field, and its deflection angle sensor collects its rotation angle in real time; at this time, the other side tail fin is still kept in a real-time zero angle of attack state under the control of the flight control computer; at the same time, the flight control computer controls the trailing edge of the corresponding side aerodynamic control surface to deflect downward, and its deflection angle sensor collects its rotation angle in real time, while the other side aerodynamic control surface remains in place;

[0019] When the angular acceleration or sideslip angle of the aircraft nose deviation disappears, the corresponding side tail fin returns to the real-time zero angle of attack state, and the corresponding side aerodynamic control surface returns to the initial position;

[0020] (12) When the aircraft is in the post-stall mode with an angle of attack greater than the stall angle and the flight control computer does not send a control command signal to the control surface actuators;

[0021] (121) When the aircraft is not disturbed by the outside world and the yaw acceleration of the aircraft collected by the sideslip angle sensor is zero, the aerodynamic control surface does not deflect. At this time, the flight control computer controls the leading edges of the left and right tail fins to deflect downward, so that the tail fins on both sides are in an adaptive real-time zero angle of attack state in the local flow field;

[0022] (122) When the aircraft is subjected to external disturbances causing the aircraft body to generate sideslip angle acceleration, the sideslip angle sensor transmits the real-time sideslip angle acceleration signal it collects to the flight control computer, and the flight control computer calculates the angle at which the leading edges of the two side tail fins are simultaneously deflected in the dorsal direction of the sideslip, and controls the tail fin actuator to actuate to a specific deflection angle. At the same time, the tail fin deflection angle sensor collects the tail fin deflection angle and transmits the signal to the flight control computer; at the same time, after the angular acceleration or sideslip angle of the aircraft nose disappears, the flight control computer controls the left and right tail fins to return to the adaptive local flow field zero angle of attack state;

[0023] (2) Yaw control:

[0024] (21) When the aircraft is in a mode where the angle of attack is less than the stall angle of attack, and a yaw control command is input, causing the aircraft to perform yaw control;

[0025] The aircraft rolls inward in the yaw direction, and the aircraft rolling slope is collected by the slope sensor. When the specified slope is reached, the roll command is released, the aerodynamic control surface returns to the initial position, and the aircraft enters the corresponding yaw circling maneuver at the corresponding angle of attack and thrust level. At this time, the flight control computer controls the corresponding aerodynamic control surface and the corresponding tail fin to flip downward, so that the upper surface of the corresponding tail fin is tilted in the local flow field to present a negative angle of attack posture. At this time, the tail fin on the other side is still controlled by the flight control computer to maintain a real-time zero angle of attack state; at this time, the downward moment component generated by the corresponding tail fin and the upward moment component generated by the corresponding aerodynamic control surface are equal in value and opposite in direction, and they offset each other to maintain the stability of the lateral side of the aircraft body and maintain the original slope;

[0026] (22) When the aircraft is in the post-stall mode, before the yaw control command is input, the left and right tail fins are still controlled by the flight controller to maintain a real-time zero angle of attack state in the local flow field:

[0027] After the yaw control command is transferred to the flight control computer, the left and right tail fin actuators are controlled by the flight control computer, so that the leading edges of the left and right tail fins are deflected to the corresponding directions at the same angle at the same time. The left and right tail fins jointly generate lateral force moments in the corresponding directions, thereby realizing the deflection of the aircraft body.

[0028] Beneficial effects: The fixed-wing aircraft of the present invention can achieve effective yaw stabilization and yaw control through aerodynamic surface control in all the aforementioned modes, realizes yaw stabilization and yaw control of the fixed-wing aircraft without a vertical tail, and completes an equivalent replacement of the vertical tail function of the fixed-wing aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] Figure 1 The figure is a schematic structural diagram of a fixed-wing aircraft according to an embodiment of the present invention.

[0030] Figure 2 The figure is a schematic structural diagram of a fixed-wing aircraft according to another embodiment of the present invention.

[0031] Figure 3 It is a flow chart of the posture control method of the present invention.

[0032] Wherein, 1 is the aircraft body, 11 is the main wing, 12 is the tail fin, 13 is the flaperon, 14 is the flap, and 15 is the aileron;

[0033] AB represents the side strip of the aircraft body, BC represents the leading edge of the main wing, DE represents the leading edge of the tail fin, FH represents the trailing edge of the tail fin, EF represents the outer edge of the tail fin, and GH represents the inner edge of the tail fin. DETAILED DESCRIPTION

[0034] The present invention is further explained below in conjunction with the accompanying drawings and specific embodiments.

[0035] like Figure 1 , 2 As shown, the fixed-wing aircraft of the present invention comprises an aircraft body 1 and a main wing 11 mounted on the aircraft body 1, a flight control computer is arranged in the aircraft body 1, and an angle of attack sensor, a sideslip angle sensor and a slope sensor are respectively mounted on the aircraft body 1; the leading edge of the main wing 11 is swept back, and the trailing edge is swept forward. A flaperon 13 is rotatably mounted at a symmetrical position on the trailing edge of the main wing 11 through a rotating shaft, and a flaperon actuator for controlling the deflection of the flaperon 13 is mounted on the left and right main wings 11, and a flaperon deflection angle sensor for collecting the rotation angle of the flaperon 13 is mounted on the flaperon actuator; a tail fin 12 is rotatably mounted on both sides of the aircraft body 1 at symmetrical positions on both sides of the tail of the aircraft body 1 through a rotating shaft, the tail fin 12 rotating shaft is located between the leading edge DE and the trailing edge FH on the tail fin and is swept back (its sweep angle is greater than or equal to the sweep angle of the leading edge DE), the rotating shaft is connected to the aircraft body 1 through the inner edge GH of the tail fin, and the chord plane where the tail fin 12 is located is parallel to the main wing 11.

[0036] The chord planes where the leading edges are located are parallel or the angle between the two is less than 15°; a tail fin actuator for controlling the rotation of the tail fin 12 is installed in the aircraft body 1, and a tail fin deflection sensor for collecting the rotation angle of the tail fin 12 is installed on the actuator.

[0037] Among them, the angle of attack sensor, sideslip angle sensor, slope sensor, flaperon angle sensor and tail fin angle sensor are all connected to the flight control computer, and the data signals collected by them are sent to the flight control computer; the flaperon actuator and the tail fin actuator are all connected to the flight control computer, and their movements are controlled by the flight control computer, thereby controlling the free deflection of the flaperon 13 and the tail fin 12 respectively.

[0038] In the present invention, the rotation axis of the flaperon 13 is arranged on the hinge line of the main wing 11 .

[0039] In a specific embodiment of the present invention, the leading edge DE of the tail fin is parallel to the trailing edge FH of the tail fin, and the outer edge EF of the tail fin is parallel to the inner edge GH of the tail fin. The outer edge EF (or the inner edge GH) of the tail fin 12 is parallel to the side strip AB of the aircraft body 1, and the leading edge BC of the main wing 11 is parallel to the leading edge DE (or the trailing edge FH) of the tail fin 12.

[0040] In the present invention, the rotation axis of the tail fin 12 is coplanar with the tail fin 12 and is disposed between the leading edge DE of the tail fin and the trailing edge FH of the tail fin.

[0041] In another embodiment of the present invention, ailerons 15 are rotatably installed at positions of the trailing edges of the left and right main wings 11 away from the aircraft body 1 through a rotating shaft, and flaps 14 are rotatably installed at positions of the trailing edges of the left and right main wings 11 close to the aircraft body 1 through a rotating shaft. At the same time, actuators for controlling the rotation of the flaps 13 and the ailerons 15 are installed in the aircraft body 1, and angle sensors for collecting the rotation angles of the flaps 14 and the ailerons 15 are installed in the aircraft body 1.

[0042] In the present invention, when the aircraft is in a state of no yaw or no control instruction input at all angles of attack within the flight envelope, the flight control computer controls the tail fin to be in an adaptive real-time zero angle of attack state with the local incoming flow at any angle of attack.

[0043] An example of a fixed-wing aircraft attitude control method of the present invention includes:

[0044] (1) Yaw stability

[0045] (11) When the aircraft is in the straight-ahead mode with an angle of attack less than the stall angle of attack;

[0046] (111) When the sideslip angle sensor detects that the sideslip angle of the aircraft is zero, the flaperon 13 on the main wing 11 does not deflect, and the flight control computer controls the tail fin to be in an adaptive real-time zero angle of attack state with the local incoming flow;

[0047] (112) When the aircraft is disturbed and produces sideslip angular acceleration, take left slip as an example:

[0048] The aircraft generates an angular acceleration that deflects to the right, and the sideslip angle sensor transmits the real-time sideslip angle acceleration signal it collects to the flight control computer. The flight control computer calculates the rotation angle of the left tail fin and the left flaperon based on this, and transmits it to the left tail fin actuator and the left flaperon actuator respectively, so that the leading edge of the left tail fin deflects downward from the initial position (no new control command input), so that the upper surface of the left tail fin is inclined in the local flow field and presents a negative angle of attack attitude, and at the same time its deflection angle sensor collects its rotation angle in real time; at this time, the left tail fin generates three torque components to the right, downward, and backward respectively; the trailing edge of the left flaperon deflects downward under the command of the flight control computer, and at the same time its deflection angle sensor collects its rotation angle in real time; the flight control computer and the bank angle sensor control the left tail fin to generate The downward moment component generated is equal to the upward moment component generated by the same side flap aileron and is opposite in direction. The two moment components cancel each other out, making the roll angular acceleration collected by the slope sensor zero, so as to maintain the lateral stability of the aircraft body and maintain the original slope; at this time, the right tail fin is still controlled by the flight control computer to maintain the real-time zero angle of attack state; the right flap aileron has no control input and remains in place; the backward drag component generated by the left flap aileron and the backward drag moment component and the right force moment component generated by the left tail fin jointly generate a yaw recovery angular acceleration that deflects the nose to the left, so as to maintain the yaw stability of the aircraft itself; when the angular acceleration or sideslip angle of the aircraft nose deviating to the right disappears, the left tail fin returns to the adaptive real-time zero angle of attack state, and the left flap aileron returns to its initial position.

[0049] The controls when sliding right are opposite, but only target the right tail fin and right flaperon.

[0050] (12) When the aircraft is in an angle of attack exceeding the stall angle of attack (i.e., in the post-stall mode), if there is a disturbance such as a crosswind that causes the aircraft nose to deflect by an angular acceleration, still taking the aircraft nose deflecting to the right as an example:

[0051] The sideslip angle sensor transmits the sideslip angle acceleration signal it collects to the flight control computer, which calculates the deflection angles of the left and right tail fins based on the signal and sends instructions to the left and right tail fin actuators to make the leading edges of the left and right tail fins deflect to the right at the same angle at the same time, generating an angular acceleration that causes the aircraft nose to deflect to the left to eliminate the aircraft nose deviation trend. At the same time, the deflection angle sensors of the left and right tail fins respectively collect their rotation angles in real time; when the angular acceleration or sideslip angle of the aircraft nose deviation disappears, the left and right tail fins return to the adaptive zero angle of attack state under the instruction of the flight control computer.

[0052] (2) Yaw control

[0053] (21) In the mode where the aircraft's angle of attack is less than the stall angle of attack, when a yaw control command is input to cause the aircraft to perform yaw control, taking yaw to the left as an example:

[0054] When the aircraft rolls to the left and generates a specific slope, the slope sensor collects the aircraft's roll slope. When the specified slope is reached, the roll command is released and the flaperons involved in the roll control return to their initial positions. The aircraft enters a left yaw hovering maneuver under the corresponding angle of attack and thrust. To eliminate the generated left slip angle, the flight control computer controls the left flaperon actuator and the left tail fin actuator, thereby deflecting the left flaperon and the left tail fin, causing the leading edge of the left tail fin to adaptively deflect downward in the real-time zero angle of attack state, and the upper surface of the left tail fin is tilted in the local flow field to present a negative angle of attack attitude, which respectively generates right , downward, and backward torque components. At the same time, the flight control computer sends a control command to the left flaperon actuator to control the trailing edge of the left flaperon to deflect downward from the initial position, generating upward and backward resistance torque components. The flight control computer and the bank angle sensor control the downward torque component generated by the left tail fin to be equal in value and opposite in direction to the upward torque component generated by the left flaperon, so as to maintain the existing slope of the lateral side of the aircraft body on the principle of keeping the rolling angular acceleration at zero. At this time, the right tail fin is still kept in a real-time zero angle of attack state under the control of the flight control computer, and the right flaperon is not deflected.

[0055] The controls for right yaw maneuver are opposite, but only for the right tail fin and right flaperon.

[0056] (22) When the aircraft's angle of attack exceeds the stall angle of attack, that is, when it is in the post-stall mode, before the yaw control command is input, the left and right tail fins are still controlled by the flight controller to maintain a real-time zero angle of attack state in the local flow field, so as to make the aircraft nose deflect to the left: the flight control computer controls the left and right tail fin actuators, so that the leading edges of the left and right tail fins deflect to the right at the same angle at the same time, and the left and right tail fins jointly generate a leftward lateral force moment, thereby realizing the leftward deflection of the aircraft body.

[0057] In the present invention, for the fixed-wing aircraft of another embodiment described above, during the specific yaw stabilization and yaw control process, the flight control computer applies control similar to the aforementioned flaperon 13 to the flaps 14 and ailerons 15 thereof.

[0058] The present invention realizes yaw stabilization and yaw control of a fixed-wing aircraft without a vertical tail through the control method of the wing surface system in the above-mentioned modes, and completes the equivalent replacement of the vertical tail function of the fixed-wing aircraft.

[0059] The preferred embodiments of the present invention are described in detail above, but the present invention is not limited to the specific details in the above embodiments. Within the technical concept of the present invention, various equivalent transformations (such as quantity, shape, position, etc.) can be made to the technical scheme of the present invention, and these equivalent transformations all belong to the protection scope of the present invention.

Claims

1. A fixed-wing aircraft, an aircraft body, on which is provided a flight control computer and an angle of attack sensor, a sideslip angle sensor and a slope sensor connected to the flight control computer; main wings of the aircraft body are symmetrically mounted on both sides of the aircraft body, and the leading edges of the main wings are swept back; Features: The trailing edge of the main wing is swept forward; aerodynamic control surfaces whose deflection is controlled by a flight control computer are symmetrically installed at the trailing edge positions of the main wing on both sides; tail fins whose rotation is controlled by a flight control computer are symmetrically installed at both sides of the tail of the aircraft body, and the rotation axis of the tail fin is swept backward; The sweep angle of the tail fin rotation axis is greater than or equal to the sweep angle of the main wing leading edge, and the chord plane where the tail fin is located is parallel to the chord plane where the main wing leading edge is located, or the angle between the two is less than 15°; The rotation axis of the tail fin is arranged between the leading edge and the trailing edge of the tail fin.

2. The fixed-wing aircraft according to claim 1, Features: The aerodynamic control surface is a flaperon, and the rotation axis of the flaperon is arranged on the hinge line of the main wing.

3. The fixed-wing aircraft according to claim 1, Features: The aerodynamic control surfaces adopt flaps and ailerons. Ailerons are rotatably installed at positions of the trailing edges of the left and right sides of the main wing away from the aircraft body, and flaps are rotatably installed at positions of the trailing edges of the left and right sides of the main wing close to the aircraft body. At the same time, actuators for controlling the rotation of the flaps and the ailerons are installed in the aircraft body, and deflection sensors for collecting the rotation angles of the flaps and the ailerons are installed in the aircraft body.

4. The fixed-wing aircraft according to claim 1, Features: The leading edge of the tail fin is parallel to its trailing edge, and its outer edge is parallel to its inner edge; the outer edge of the tail fin is parallel to the edge strip of the aircraft body, and the leading edge of the main wing is parallel to the leading edge of the tail fin.

5. A method for attitude control of a fixed-wing aircraft using any one of claims 1 to 4, Features: include: (1) Yaw stability: (11) When the aircraft is in the straight-ahead mode with its angle of attack less than the stall angle of attack and the flight control computer is not sending control command signals to the control surface actuators: (111) When the sideslip angle acceleration of the aircraft acquired by the sideslip angle sensor is zero, the aerodynamic control surface on the main wing does not deflect, and at this time, the tail fin is controlled by the flight control computer to be in a real-time zero angle of attack state with respect to the local flow field; (112) When the aircraft is disturbed and generates sideslip angle acceleration, the sideslip angle sensor transmits the real-time deflection angle acceleration signal collected by the sensor to the flight control computer; The flight control computer calculates the rotation angle of the corresponding side tail fin and the corresponding side aerodynamic control surface based on the calculation, and controls the corresponding side tail fin accordingly, so that its upper surface presents a negative angle of attack attitude in the local flow field, and at the same time, its deflection angle sensor collects its rotation angle in real time; At this time, the tail fin on the other side is still controlled by the flight control computer to maintain a real-time zero angle of attack state; at the same time, the flight control computer controls the trailing edge of the aerodynamic control surface on the corresponding side to deflect downward, and its deflection angle sensor collects its rotation angle in real time, while the aerodynamic control surface on the other side remains in place; When the angular acceleration or sideslip angle of the aircraft nose deviation disappears, the corresponding side tail fin returns to the real-time zero angle of attack state, and the corresponding side aerodynamic control surface returns to the initial position; (12) When the aircraft is in the post-stall mode with an angle of attack greater than the stall angle and the flight control computer does not send a control command signal to the control surface actuators; (121) When the aircraft is not disturbed by the outside world and the yaw acceleration of the aircraft collected by the sideslip angle sensor is zero, the aerodynamic control surface does not deflect. At this time, the flight control computer controls the leading edges of the left and right tail fins to deflect downward, so that the tail fins on both sides are in an adaptive real-time zero angle of attack state in the local flow field; (122) When the aircraft is subjected to external disturbances causing the aircraft body to generate sideslip angle acceleration, the sideslip angle sensor transmits the real-time sideslip angle acceleration signal it collects to the flight control computer, and the flight control computer calculates the angle at which the leading edges of the two side tail fins are simultaneously deflected in the dorsal direction of the sideslip, and controls the tail fin actuator to actuate to a specific deflection angle. At the same time, the tail fin deflection angle sensor collects the tail fin deflection angle and transmits the signal to the flight control computer; at the same time, after the angular acceleration or sideslip angle of the aircraft nose disappears, the flight control computer controls the left and right tail fins to return to the adaptive local flow field zero angle of attack state; (2) Yaw control: (21) When the aircraft is in a mode where the angle of attack is less than the stall angle of attack, and a yaw control command is input, causing the aircraft to perform yaw control; The aircraft rolls inward in the yaw direction, and the aircraft roll slope is collected by the slope sensor. When the specified slope is reached, the roll command is released, the aerodynamic control surface returns to the initial position, and the aircraft enters the corresponding yaw circling maneuver at the corresponding angle of attack and thrust level. At this time, the flight control computer controls the corresponding aerodynamic control surface and the corresponding tail fin to flip downward, so that the upper surface of the corresponding tail fin is tilted in the local flow field to present a negative angle of attack attitude. At this time, the tail fin on the other side is still controlled by the flight control computer to maintain a real-time zero angle of attack state; at this time, the downward torque component generated by the corresponding tail fin and the upward torque component generated by the corresponding aerodynamic control surface are equal in value and opposite in direction, and they offset each other to maintain the stability of the lateral side of the aircraft body and maintain the original slope; (22) When the aircraft is in the post-stall mode, before the yaw control command is input, the left and right tail fins are still controlled by the flight controller to maintain a real-time zero angle of attack state in the local flow field: After the yaw control command is transferred to the flight control computer, the left and right tail fin actuators are controlled by the flight control computer, so that the leading edges of the left and right tail fins are deflected to the corresponding directions at the same angle at the same time. The left and right tail fins jointly generate lateral force moments in the corresponding directions, realizing the deflection of the aircraft body.

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