CN115465443A - Fixed-wing aircraft and attitude control method thereof - Google Patents
Fixed-wing aircraft and attitude control method thereof Download PDFInfo
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- CN115465443A CN115465443A CN202110652656.1A CN202110652656A CN115465443A CN 115465443 A CN115465443 A CN 115465443A CN 202110652656 A CN202110652656 A CN 202110652656A CN 115465443 A CN115465443 A CN 115465443A
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- 238000000034 method Methods 0.000 title claims description 11
- RZVHIXYEVGDQDX-UHFFFAOYSA-N 9,10-anthraquinone Chemical compound C1=CC=C2C(=O)C3=CC=CC=C3C(=O)C2=C1 RZVHIXYEVGDQDX-UHFFFAOYSA-N 0.000 claims abstract description 65
- 230000006641 stabilisation Effects 0.000 claims abstract description 7
- 238000011105 stabilization Methods 0.000 claims abstract description 7
- 230000001133 acceleration Effects 0.000 claims description 35
- 238000005096 rolling process Methods 0.000 claims description 10
- 238000005336 cracking Methods 0.000 description 2
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C9/00—Adjustable control surfaces or members, e.g. rudders
- B64C9/32—Air braking surfaces
- B64C9/323—Air braking surfaces associated with wings
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
- B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
- B64F5/60—Testing or inspecting aircraft components or systems
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
- G05D1/0816—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability
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Abstract
The invention discloses a fixed-wing aircraft, which comprises an aircraft body, wherein a flight control computer, and an attack angle sensor, a sideslip angle sensor and a gradient sensor which are connected with the flight control computer are arranged on the aircraft body; the main wings are symmetrically arranged on two sides of the aircraft body, and the front edges of the main wings are swept backwards; the trailing edge of the main wing is swept forward; pneumatic control surfaces which are controlled to deflect by a flight control computer are symmetrically arranged at the positions of the rear edges of the main wings on the two sides; the tail fins are symmetrically arranged at the two sides of the tail of the aircraft body and controlled to rotate by a flight control computer, and the rotating axes of the tail fins are swept backwards; the sweep angle of the rotation axis of the tail fin is larger 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 included angle of the chord plane and the chord plane is smaller than 15 degrees. The invention can realize effective yaw stabilization and yaw control on the fixed wing aircraft under the condition of no vertical tail wing through pneumatic surface control in all the modes.
Description
Technical Field
The invention relates to the technical field of aircrafts, in particular to a fixed-wing aircraft without a vertical tail wing and an attitude control method thereof.
Background
The prior art successfully manufactured fixed wing aircraft without vertical tail and the related technical scheme disclosed mainly comprise military. The main advantages of the aircraft without the vertical tail are that: 1. after the vertical tail serving as a radar reflecting plane is removed, the appearance of the aircraft is simplified, and the RCS value of the aircraft is effectively reduced; 2. after the vertical tails are removed, the resistance sources are reduced, and the lift-drag ratio of the whole machine is improved.
Products and proposed solutions are available which have been successfully manufactured, such as U.S. invisible bombers B-2,B-21, X-36 (technical validation machine), X47B (unmanned aerial vehicle), etc. After the vertical tail fin that provides the yaw stability and yaw maneuvering moment for the aircraft during flight is removed, the yaw stability and yaw maneuvering moment for the aircraft during flight will be equivalently provided by other means. The common technical measure of the aircraft is that a split type resistance rudder is arranged at the rear edge of the outer end of the main wing of the aircraft, which is parallel to the front edge of the main wing, the lower part of the split type resistance rudder deflects downwards, the generated upward moment is offset by the downward moment generated by the upward deflection part of the resistance rudder, and the resistance moments generated by the upper part and the lower part of the resistance rudder jointly generate a yaw moment around the gravity center of the aircraft to realize yaw stability and yaw control. The technical scheme has the limitations that: when the aircraft is in a constant cruising flight mode and a small-gradient yawing maneuvering mode, the attack angle of the aircraft is in a small state, the upward deflecting part of the resistance rudder is in an incoming flow which is not separated at the rear part of the upper surface of the wing and has high flow velocity, the aerodynamic efficiency is high, and sufficient downward moment and resistance moment can be generated, so that the resistance rudder can generate the control moment required by realizing yawing control. When the design requirement of a large attack angle and large gradient maneuvering airplane is provided for an aircraft, the cracking resistance rudder cannot effectively provide yaw stability and control moment, and the reason is as follows: when the aircraft is in a large-attack-angle mode, particularly close to a stall attack angle, incoming flow at the rear part of the upper surface of the wing becomes turbulent partially, the pressure of the upward deflection part of the resistance rudder is reduced, so that sufficient downward moment cannot be generated, and the upward moment generated by the downward deflection part of the resistance rudder cannot be effectively balanced, so that the cracking resistance rudder cannot realize effective yaw control of the aircraft in the large-attack-angle mode. Further, when the aircraft is in the over-stall mode, the cracked resistance rudder is completely disabled, and attitude control of the aircraft will only be achieved through vector thrust.
Disclosure of Invention
The purpose of the invention is as follows: in order to overcome the functional limitations of yaw stabilization and control in the prior art, the invention provides a fixed-wing aircraft and an attitude control method thereof, wherein effective yaw stabilization and control can be realized in all modes through pneumatic surface control.
The technical scheme is as follows:
a fixed wing aircraft comprises an aircraft body, a flight control computer, an attack angle sensor, a sideslip angle sensor and a gradient sensor, wherein the attack angle sensor, the sideslip angle sensor and the gradient sensor are connected with the flight control computer; the main wings of the aircraft body are symmetrically arranged on two sides of the aircraft body, and the front edges of the main wings are swept backwards;
the trailing edge of the main wing is swept forward; pneumatic control surfaces which are controlled to deflect by a flight control computer are symmetrically arranged at the positions of the rear edges of the main wings on the two sides; the tail fins are symmetrically arranged at the two sides of the tail of the aircraft body and controlled to rotate by a flight control computer, and the rotating axes of the tail fins are swept backwards;
the sweep angle of the rotation axis of the tail fin is larger 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 included angle of the chord plane and the chord plane is smaller than 15 degrees.
The pneumatic control surface adopts a flaperon, and a rotating shaft of the flaperon is arranged on a hinge line of the main wing.
The pneumatic control surface adopts a flap and an aileron, the ailerons are respectively and rotatably installed at the positions, far away from the aircraft body, of the rear edges of the left side and the right side of the main wing, and the flaps are respectively and rotatably installed at the positions, close to the aircraft body, of the rear edges of the left side and the right side of the main wing; simultaneously, install respectively in the aircraft body and be used for control the flap with the aileron pivoted actuator, just be used for gathering respectively in the aircraft body the flap with the declination sensor of aileron turned angle is installed.
The front edge of the tail fin is parallel to the rear edge of the tail fin, and the outer edge of the tail fin is parallel to the inner edge of the tail fin; and the outer edge of the tail fin is parallel to the edge strips of the aircraft body, and meanwhile, the front edge of the main wing is parallel to the front edge of the tail fin.
The rotating shaft of the tail fin is arranged between the front edge and the rear edge of the tail fin.
An attitude control method of a fixed-wing aircraft, comprising:
(1) And (3) yaw stabilization:
(11) When the aircraft is in a straight-ahead mode in which the aircraft angle of attack is less than the stall angle of attack, and the flight control computer does not send the maneuver instruction signals to the control surface actuators:
(111) When the sideslip angular acceleration of the aircraft acquired by the sideslip angular sensor is zero, the pneumatic control surface on the main wing does not deflect, and the tail fin and the local flow field are controlled to be in a real-time zero-attack-angle state by the flight control computer;
(112) When the aircraft is disturbed to generate sideslip angular acceleration, the sideslip angular sensor transmits a real-time deflection angle acceleration signal acquired by the sideslip angular sensor into the flight control computer;
the flight control computer calculates the rotation angles of the corresponding side tail fins and the corresponding side pneumatic control surface according to the rotation angles, and controls the corresponding side tail fins respectively according to the rotation angles so that the upper surface of the corresponding side tail fins is in a negative attack angle posture in a local flow field, and meanwhile, the deflection angle sensor acquires the rotation angles in real time; at the moment, the tail fin on the other side still keeps a real-time zero attack angle state under the control of a flight control computer; meanwhile, the rear edge of the pneumatic control surface on the corresponding side is controlled to deflect downwards through a flight control computer, meanwhile, the deflection angle sensor acquires the rotation angle of the pneumatic control surface in real time, and the pneumatic control surface on the other side keeps in the original position;
when the offset angular acceleration or sideslip angle of the aircraft nose disappears, the corresponding side tail fin returns to a real-time zero attack angle state, and the corresponding side pneumatic control surface returns to an initial position;
(12) When the aircraft is in an over-stall mode in which the incidence is larger than the stall incidence angle and the flight control computer does not send a direct flight mode of an operation command signal to the control surface actuator;
(121) When the aircraft is not disturbed by the outside and the yaw acceleration of the aircraft is zero acquired by the sideslip angle sensor, the pneumatic control surface does not deflect, and the flight control computer controls the downward deflection of the front edges of the left and right tail fins at the moment so that the tail fins on the two sides are in a self-adaptive real-time zero-attack-angle state in a local flow field;
(122) When the aircraft is disturbed by the outside to enable the aircraft body to generate sideslip angular acceleration, the sideslip angular acceleration sensor transmits a real-time sideslip angular acceleration signal acquired by the sideslip angular acceleration sensor into the flight control computer, the flight control computer calculates the angle of the front edges of the two lateral tail fins deflecting towards the sideslip back side direction at the same time according to the sideslip angular acceleration signal, controls the tail fin actuator to actuate to a specific deflection angle, and simultaneously the tail fin deflection angle sensor acquires a tail fin deflection angle and transmits the signal to the flight control computer; meanwhile, after the angular acceleration or sideslip angle of the aircraft nose offset disappears, the flight control computer controls the left tail fin and the right tail fin to return to a self-adaptive local flow field zero attack angle state;
(2) Yaw manipulation:
(21) When the incidence angle of the aircraft is smaller than the stall incidence angle, when a yaw control command is input, the aircraft is subjected to yaw control;
the aircraft rolls towards the inner side of the yaw direction, the rolling gradient of the aircraft is collected through a gradient sensor, when the specified gradient is reached, the rolling instruction is removed, the pneumatic control surface returns to the initial position, the aircraft enters corresponding yaw hovering maneuver under corresponding attack angle and thrust level, at the moment, the flight control computer controls the corresponding pneumatic control surface and the corresponding tail fin to turn downwards, so that the upper surface of the corresponding tail fin is in a negative attack angle attitude in a local flow field in an inclined manner, and at the moment, the tail fin on the other side is still kept in a real-time zero attack angle state under the control of the flight control computer; at the moment, the downward moment component generated by the corresponding tail fin and the upward moment component generated by the corresponding pneumatic control surface have the same numerical value and opposite directions, and are mutually counteracted to keep the stability of the transverse side of the aircraft body and the original gradient;
(22) When the aircraft is in an over-stall mode, before a yaw control command is input, the left and right tail fins still keep a real-time zero-attack-angle state in a local flow field under the control of the flight controller:
after the yaw control instruction is transferred into the flight control computer, the flight control computer controls the left and right tail fin actuators to enable the front edges of the left and right tail fins to deflect towards corresponding directions at equal angles at the same time, and the left and right tail fins generate lateral force moments in corresponding directions together to achieve deflection of the aircraft body.
Has the advantages that: the fixed wing aircraft can realize effective yaw stability and yaw control in all the modes through pneumatic surface control, realize the yaw stability and yaw control of the fixed wing aircraft under the condition of no vertical tail, and finish the equivalent substitution of the vertical tail function of the fixed wing aircraft.
Drawings
Fig. 1 is a schematic structural view of a fixed-wing aircraft according to an embodiment of the present invention.
Fig. 2 is a schematic structural view of a fixed-wing aircraft according to another embodiment of the present invention.
FIG. 3 is a flow chart of a method for controlling attitude according to the present invention.
Wherein, 1 is an aircraft body, 11 is a main wing, 12 is a tail fin, 13 is a flaperon, 14 is a flap, and 15 is an aileron;
AB denotes the aircraft body strake, BC denotes the leading edge of the main wing, DE denotes the leading edge of the tail fin, FH denotes the trailing edge of the tail fin, EF denotes the outer edge of the tail fin, and GH denotes the inner edge of the tail fin.
Detailed Description
The invention is further elucidated with reference to the drawings and the embodiments.
As shown in fig. 1 and 2, the fixed-wing aircraft of the present invention includes an aircraft body 1 and a main wing 11 installed on the aircraft body 1, a flight control computer is installed in the aircraft body 1, and an attack angle sensor, a sideslip angle sensor and a gradient sensor are respectively installed on the aircraft body 1; the main wing 11 is swept back at the leading edge and swept forward at the trailing edge. The flaperon 13 is rotatably arranged at the symmetrical position of the rear edge of the main wing 11 through a rotating shaft, flaperon actuators for controlling the flaperon 13 to perform deflection are arranged on the left main wing 11 and the right main wing 11, and flaperon deflection angle sensors for collecting the rotation angle of the flaperon 13 are arranged on the flaperon actuators; the tail fins 12 are rotatably arranged on two sides of the aircraft body 1 at symmetrical positions of two sides of the tail of the aircraft body 1 through rotating shafts, the rotating shafts of the tail fins 12 are positioned between the front edge DE and the rear edge FH and are swept backward (the sweep angle of the tail fins is more than or equal to that of the front edge DE), the rotating shafts are connected with the aircraft body 1 through the inner edges GH of the tail fins, and the chord plane where the tail fins 12 are positioned is parallel to the chord plane where the front edge of the main wing 11 is positioned or the included angle between the chord plane and the chord plane is less than 15 degrees; 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 angle sensor for acquiring the rotation angle of the tail fin 12 is installed on the actuator.
Wherein, the attack angle sensor, the sideslip angle sensor, the gradient sensor, the flaperon deflection angle sensor and the tail fin deflection angle sensor are all connected with the flight control computer, and send the acquired data signals to the flight control computer; the flaperon actuator and the tail fin actuator are both connected with the flight control computer, and the movement of the flaperon actuator and the tail fin actuator is controlled by the flight control computer, so that the flaperon 13 and the tail fin 12 are controlled to deflect freely.
In the present invention, the rotation axis of the flap 13 is arranged on the hinge line of the main wing 11.
In the exemplary 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. And the outer edge EF (or the inner edge GH) of the tail fin 12 is parallel to the edge strip AB of the aircraft body 1, while 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.
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 and the trailing edge FH of the tail fin.
In another embodiment of the present invention, ailerons 15 are rotatably mounted on the positions of the rear edges of the left and right main wings 11 away from the aircraft body 1 through rotating shafts, respectively, flaps 14 are rotatably mounted on the positions of the rear edges of the left and right main wings 11 close to the aircraft body 1 through rotating shafts, respectively, actuators for controlling the rotation of the flaps 13 and the ailerons 15 are mounted in the aircraft body 1, respectively, and declination sensors for acquiring the rotation angles of the flaps 14 and the ailerons 15 are mounted in the aircraft body 1, respectively.
In the invention, when the aircraft is in a state without yaw or operation instruction input under all angles of attack in the flight envelope, the flight control computer controls the tail fin to be in a self-adaptive real-time zero-angle-of-attack state with local incoming flow under any angle of attack.
An example of a fixed-wing aircraft attitude control method of the present invention includes:
(1) Yaw stabilization
(11) When the aircraft is in a straight-ahead mode with an angle of attack smaller than a stall angle of attack;
(111) When the sideslip angle of the aircraft acquired by the sideslip angle sensor is zero, the flaperon 13 on the main wing 11 does not deflect, and the tail fin and local incoming flow are controlled to be in a self-adaptive real-time zero-attack-angle state by the flight control computer;
(112) When the aircraft is disturbed to generate sideslip angular acceleration, take the left sideslip as an example:
the aircraft generates angular acceleration deflecting rightwards, the sideslip angle sensor transmits real-time sideslip angle acceleration signals acquired by the sideslip angle sensor into the flight control computer, the flight control computer calculates the rotation angles of the left tail fin and the left flaperon according to the rotation angles and respectively transmits the rotation angles to the left tail fin actuator and the left flaperon actuator, the front edge of the left tail fin is enabled to deflect downwards from an initial position (without new operation instruction input), the upper surface of the left tail fin is enabled to be in a negative attack angle posture in an inclined mode in a local flow field, and meanwhile, the deflection angle sensor acquires the rotation angle in real time; at the moment, the left tail fin generates three moment components of right, downward and backward directions respectively; the left flaperon deflects downwards at the rear edge under the instruction of a flight control computer, and meanwhile, a deflection angle sensor acquires the rotation angle of the left flaperon in real time; the flight control computer and the slope angle sensor control the downward moment component generated by the left tail fin and the upward moment component generated by the same side flap to be equal in value and opposite in direction, and the two moment components are mutually offset to enable the rolling angle acceleration acquired by the slope sensor to be zero so as to keep the stability of the transverse side of the aircraft body and the original slope; at the moment, the right tail fin still keeps a real-time zero attack angle state under the control of a flight control computer; the right flaperon has no operation input and keeps the original position; the yaw recovery angular acceleration which enables the aircraft head to deflect leftwards is generated by the backward resistance component generated by the left flaperon, the backward resistance moment component generated by the left tail fin and the right force moment component together, and the yaw stability of the aircraft is kept; when the angular acceleration or the sideslip angle of the aircraft nose deviated to the right disappears, the left skeg returns to the self-adaptive real-time zero attack angle state, and the left flaperon returns to the initial position.
The control in right sideslip is the opposite, only for the right tail fin and the right flaperon.
(12) When the aircraft is in a state that the attack angle of the aircraft exceeds a stall attack angle (namely, in an over-stall mode), if the aircraft nose generates deflection angular acceleration due to disturbance of factors such as crosswind, the aircraft nose is still deflected to the right as an example:
the sideslip angle sensor transmits the collected sideslip angle acceleration signals to the flight control computer, the flight control computer calculates the deflection angles of the left tail fin and the right tail fin according to the sideslip angle acceleration signals, and sends instructions to the left tail fin actuator and the right tail fin actuator to enable the front edges of the left tail fin and the right tail fin to deflect rightwards at the same time at the same angle, so that the acceleration of the left deflection angle of the aircraft nose is generated to eliminate the deviation trend of the aircraft nose, and meanwhile, the deflection angle sensors of the left tail fin and the right tail fin respectively collect the rotation angles of the aircraft nose in real time; and when the angular acceleration or the sideslip angle of the aircraft nose offset disappears, the left tail fin and the right tail fin return to the self-adaptive zero attack angle state under the instruction of the flight control computer.
(2) Yaw steering
(21) In a mode that the attack angle of the aircraft is smaller than the stall attack angle, when a yaw control command is input to enable the aircraft to carry out yaw control, taking left yaw as an example:
when the aircraft rolls to the left to generate a specific gradient, acquiring the rolling gradient of the aircraft through a gradient sensor, when the specified gradient is reached, removing a rolling instruction and returning the flap aileron participating in rolling operation to an initial position, enabling the aircraft to enter a left yawing hovering mode under the action of corresponding attack angle and thrust, and controlling a left flap actuator and a left tail fin actuator by a flight control computer to eliminate a generated left sliding angle, so that the left flap and the left tail fin deflect, leading edges of the left tail fin deflect downwards in a self-adaptive real-time zero-attack-angle state, the upper surface of the left tail fin is obliquely in a negative-attack-angle posture in a local flow field to generate right, downward and backward torque components respectively, meanwhile, the flight control computer sends a control instruction to the left flap actuator to control the trailing edge of the left flap to deflect downwards from the initial position, so as to generate upward and backward resistance torque components, the flight control computer and the slope angle sensor control the downward torque generated by the left flap to be equal to the upward torque generated by the left flap, and keep the directions of the upward gradient components of the left flap as opposite to the rolling acceleration of the aircraft body, and the rolling acceleration principle of the left flap; at the moment, the right tail fin still keeps a real-time zero attack angle state under the control of a flight control computer, and the right flaperon does not deflect.
The control during yaw steering to the right is reversed, only for the right tail fin and the right flaperon.
(22) When the aircraft attack angle exceeds the stall attack angle, namely the aircraft is in the over-stall mode, before the yaw control command is input, the left and right tail fins still keep a real-time zero-attack-angle state in a local flow field under the control of the flight controller, so that the aircraft nose deflects to the left as an example: the left and right tail fin actuators are controlled by the flight control computer, so that the front edges of the left and right tail fins deflect rightwards at the same time at equal angles, and the left and right tail fins jointly generate leftward lateral force moment to realize leftward deflection of the aircraft body.
In the present invention, for the fixed wing aircraft of the other embodiment described above, the flight control computer applies control to its flaps 14 and ailerons 15 similar to the aforementioned flaperons 13 during yaw settling and yaw maneuvering.
The invention realizes the yaw stabilization and yaw control of the fixed wing aircraft under the condition of no vertical tail wing through the control method of the wing surface system in each mode, and completes the equivalent substitution of the vertical tail wing function of the fixed wing aircraft.
Although the preferred embodiments of the present invention have been described in detail, the present invention is not limited to the details of the foregoing embodiments, and various equivalent changes (such as number, shape, position, etc.) may be made within the technical spirit of the present invention, and these equivalent changes are all within the scope of the present invention.
Claims (6)
1. A fixed wing aircraft comprises an aircraft body, a flight control computer, an attack angle sensor, a sideslip angle sensor and a gradient sensor, wherein the attack angle sensor, the sideslip angle sensor and the gradient sensor are connected with the flight control computer; the main wings of the aircraft body are symmetrically arranged on two sides of the aircraft body, and the front edges of the main wings are swept backwards; the method is characterized in that:
the trailing edge of the main wing is swept forward; pneumatic control surfaces which are controlled to deflect by a flight control computer are symmetrically arranged at the positions of the rear edges of the main wings on the two sides; the tail fins are symmetrically arranged at the two sides of the tail of the aircraft body and controlled to rotate by a flight control computer, and the rotating axes of the tail fins are swept backwards;
the sweep angle of the rotation axis of the tail fin is larger 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 included angle of the chord plane and the chord plane is smaller than 15 degrees.
2. The fixed-wing aircraft of claim 1, wherein: the pneumatic control surface adopts a flaperon, and a rotating shaft of the flaperon is arranged on a hinge line of the main wing.
3. The fixed-wing aircraft of claim 1, wherein: the pneumatic control surface adopts a flap and an aileron, the ailerons are respectively and rotatably installed at the positions, far away from the aircraft body, of the rear edges of the left side and the right side of the main wing, and the flap is respectively and rotatably installed at the positions, close to the aircraft body, of the rear edges of the left side and the right side of the main wing; simultaneously, install respectively in the aircraft body and be used for control the flap with the aileron pivoted actuator, just be used for gathering respectively in the aircraft body the flap with the declination sensor of aileron turned angle is installed.
4. The fixed-wing aircraft of claim 1, wherein: the front edge of the tail fin is parallel to the rear edge of the tail fin, and the outer edge of the tail fin is parallel to the inner edge of the tail fin; and the outer edge of the tail fin is parallel to the edge strips of the aircraft body, and meanwhile, the front edge of the main wing is parallel to the front edge of the tail fin.
5. The fixed-wing aircraft according to claim 1, wherein: the rotating shaft of the tail fin is arranged between the front edge and the rear edge of the tail fin.
6. An attitude control method of a fixed-wing aircraft to which the fixed-wing aircraft of any one of claims 1 to 5 is applied, characterized in that: the method comprises the following steps:
(1) And (3) yaw stabilization:
(11) When the aircraft is in a straight-ahead mode in which the aircraft angle of attack is less than the stall angle of attack, and the flight control computer does not send the maneuver instruction signals to the control surface actuators:
(111) When the sideslip angular acceleration of the aircraft acquired by the sideslip angular sensor is zero, the pneumatic control surface on the main wing does not deflect, and the tail fin and the local flow field are controlled to be in a real-time zero-attack-angle state by the flight control computer;
(112) When the aircraft is disturbed to generate sideslip angular acceleration, the sideslip angular sensor transmits a real-time deflection angle acceleration signal acquired by the sideslip angular sensor into the flight control computer;
the flight control computer calculates the rotation angles of the corresponding side tail fin and the corresponding side pneumatic control surface according to the rotation angles, and respectively controls the corresponding side tail fin according to the rotation angles so that the upper surface of the corresponding side tail fin is in a negative attack angle posture in a local flow field, and meanwhile, the deflection angle sensor acquires the rotation angles in real time; at the moment, the tail fin on the other side still keeps a real-time zero attack angle state under the control of a flight control computer; meanwhile, the rear edge of the pneumatic control surface on the corresponding side is controlled to deflect downwards through a flight control computer, meanwhile, a deflection angle sensor acquires the rotation angle of the pneumatic control surface in real time, and the pneumatic control surface on the other side keeps in the original position;
when the offset angular acceleration or sideslip angle of the aircraft nose disappears, the corresponding side tail fin returns to a real-time zero attack angle state, and the corresponding side pneumatic control surface returns to an initial position;
(12) When the aircraft is in an over-stall mode in which the incidence is larger than the stall incidence angle and the flight control computer does not send a direct flight mode of an operation command signal to the control surface actuator;
(121) When the aircraft is not disturbed by the outside and the yaw acceleration of the aircraft is zero acquired by the sideslip angle sensor, the pneumatic control surface does not deflect, and the flight control computer controls the downward deflection of the front edges of the left and right tail fins at the moment so that the tail fins on the two sides are in a self-adaptive real-time zero-attack-angle state in a local flow field;
(122) When the aircraft is disturbed by the outside to enable the aircraft body to generate sideslip angular acceleration, the sideslip angular acceleration sensor transmits a real-time sideslip angular acceleration signal acquired by the sideslip angular acceleration sensor into the flight control computer, the flight control computer calculates the angle of the front edges of the two lateral tail fins deflecting towards the sideslip back side direction at the same time according to the sideslip angular acceleration signal, controls the tail fin actuator to actuate to a specific deflection angle, and simultaneously the tail fin deflection angle sensor acquires a tail fin deflection angle and transmits the signal to the flight control computer; meanwhile, after the angular acceleration or sideslip angle of the aircraft nose offset disappears, the flight control computer controls the left tail fin and the right tail fin to return to a self-adaptive local flow field zero attack angle state;
(2) Yaw manipulation:
(21) When the incidence angle of the aircraft is smaller than the stall incidence angle, when a yaw control command is input, the aircraft is subjected to yaw control;
the aircraft rolls towards the inner side of the yaw direction, the rolling gradient of the aircraft is collected through a gradient sensor, when the specified gradient is reached, the rolling instruction is removed, the pneumatic control surface returns to the initial position, the aircraft enters corresponding yaw hovering maneuver under corresponding attack angle and thrust level, at the moment, the flight control computer controls the corresponding pneumatic control surface and the corresponding tail fin to turn downwards, so that the upper surface of the corresponding tail fin is in a negative attack angle attitude in a local flow field in an inclined manner, and at the moment, the tail fin on the other side is still kept in a real-time zero attack angle state under the control of the flight control computer; at the moment, the downward moment component generated by the corresponding tail fin and the upward moment component generated by the corresponding pneumatic control surface have the same numerical value and opposite directions, and are mutually counteracted to keep the stability of the transverse side of the aircraft body and the original gradient;
(22) When the aircraft is in an over-stall mode, before a yaw control command is input, the left and right tail fins still keep a real-time zero-attack-angle state in a local flow field under the control of the flight controller:
after the yaw control instruction is transferred into the flight control computer, the flight control computer controls the left and right tail fin actuators to enable the front edges of the left and right tail fins to deflect towards corresponding directions at equal angles at the same time, and the left and right tail fins generate lateral force moments in corresponding directions together to achieve deflection of the aircraft body.
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