CN112960100A - Aircraft and control method thereof - Google Patents
Aircraft and control method thereof Download PDFInfo
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- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
- B64C1/0009—Aerodynamic aspects
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- B—PERFORMING OPERATIONS; TRANSPORTING
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- B64C29/00—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft
- B64C29/02—Aircraft capable of landing or taking-off vertically, e.g. vertical take-off and landing [VTOL] aircraft having its flight directional axis vertical when grounded
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- B—PERFORMING OPERATIONS; TRANSPORTING
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- B64C—AEROPLANES; HELICOPTERS
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
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- B64C39/00—Aircraft not otherwise provided for
- B64C39/12—Canard-type aircraft
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
- B64C2001/0045—Fuselages characterised by special shapes
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C39/00—Aircraft not otherwise provided for
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Abstract
An aircraft and a control method thereof comprise a fuselage, a propeller, a motor, a canard wing, wings, a control system and an energy management system, wherein the canard wing is trapezoidal and is positioned in the central axis position of the front part of the fuselage and is higher than the wings; the wing is a modified triangular wing, the arc of the front edge is excessive, the middle and rear sections of the wing body are lower than the canard wing and are in a lower single wing layout, and the canard wing and the wing form a canard layout, so that the coupled vortex lift-increasing is realized. The control method comprises the following steps: in the takeoff stage, the aircraft takes off vertically in a two-axis mode, and after the airspeed exceeds the critical airspeed, the aircraft flies in a fixed wing mode from the two-axis mode; in the landing stage, the aircraft enters a gliding route in a fixed wing mode, and the aircraft flies at a reduced speed by lowering the altitude; when the speed is lower than the critical airspeed, the aircraft is converted from a fixed wing mode to a two-axis mode; the aircraft keeps the altitude flying to a landing point in a two-axis mode; after reaching the landing point, the aircraft droops and lands.
Description
Technical Field
The invention relates to an aircraft, in particular to an aircraft with double-launch vector canard wing layout vertical take-off and landing and a control method thereof.
Background
At present, the small aircraft mainly adopts three layout modes of non-foldable wings, foldable wings and high-speed cross missile wings.
The fixed wing is adopted in the non-foldable wing aircraft, the aircraft is huge in size, generally only can be launched in a rocket boosting box type, the aircraft is huge in size and complex to use, and single soldier use and multi-platform deployment are inconvenient. Such aircraft are fast in flight, difficult to find objects, and must be used by long-trained operators. Generally, the parachute is used for one time or is opened and landed, the parachute landing recovery process is complex and uncontrollable, and the parachute cannot be used repeatedly and rapidly after the state is confirmed again for a long time.
The folding wing aircraft is small in size, mortar type high-overload catapult launching is adopted, then wings are unfolded to carry out cruise missions, overload is large during launching, impact on missile-borne equipment is large, the cost of the whole aircraft is high, and large-scale equipment of the aircraft is not facilitated. The aircraft has high flying speed and high target searching difficulty, and is not beneficial to rapid investigation of operators. In addition, the aircraft cannot be safely recycled, and the target can only be self-destroyed if not found, so that the aircraft cannot be recycled.
The high-speed cross missile wing aircraft adopts an anti-tank missile technology, has higher cost and high flying speed, is difficult to execute a reconnaissance search task, and can only complete initial target reconnaissance by other means and then launch the aircraft to strike. Moreover, the size of the aircraft is huge, the missile wing is huge to expand, the pressure of storage and transportation on individual soldiers is large, and the individual soldier independent combat cannot be finished.
Disclosure of Invention
In order to solve the technical problems, the invention provides an aircraft with double-engine vector canard wing layout vertical take-off and landing, which has the advantages of high-speed and low-speed flight performance, easy vertical autonomous take-off and landing, simple structure, and convenient production and recovery.
The technical scheme of the invention is as follows: an aircraft comprises a fuselage, propellers, a motor, canard wings, a control system and an energy management system, wherein the fuselage is cylindrical and barrel-shaped, and the head and the tail of the fuselage are in a rectifying cone shape, so that the induced resistance is reduced; the duck wing is trapezoidal and is positioned in the central axis position of the front part of the fuselage and is higher than the wings; the wings are in shape-modified delta wings, the front edges of the wings are in arc transition, the wings are arranged at the middle and rear sections of the fuselage, are lower than the canard wings and are in a lower single wing layout, and the canard wings and the wings form a canard layout to realize the coupling vortex lift increase;
the sweep angle of the duck wing is 40 degrees, 50 degrees, 60 degrees or 70 degrees, and the sweep angle of the wing is 40 degrees, 50 degrees or 60 degrees;
preferably, the sweep angle of the duck wing is 70 degrees, and the sweep angle of the wing is 40 degrees;
the control system comprises a sensing acquisition unit, a main control processor, a calculation node, a communication module and a motion coprocessor, wherein the main control processor, the calculation node, the communication module and the motion coprocessor are installed on an airborne flight control board card, the sensing acquisition unit comprises a GPS and/or Beidou navigation system, an inertial navigation system, an electronic compass, an airspeed tube, a thermometer, a vibration sensor and a structural deformation sensor, the GPS and/or Beidou system acquires position data of an aircraft, the inertial navigation device acquires attitude data of the aircraft, the electronic compass is used for acquiring course data, the airspeed tube acquires airspeed data of the aircraft, and the vibration sensor and the structural deformation sensor acquire body structural data;
the main control processor is used for operating the timing operation system, resolving positioning, orientation and attitude, executing a flight control algorithm, receiving a communication signal and debugging, the computing node is used for aircraft mode identification and system reconstruction, the motion coprocessing is used for motor driving and canard wing control, the main control processor, the computing node, the motion coprocessor and the sensing acquisition unit form a multi-core distributed networking hardware architecture, and data are interacted between the main control processor and the motion coprocessor through a high-speed bus communication mechanism; a star information internet of things is formed between the main control processor and the sensing acquisition unit; the master control processor, the computing nodes and the motion coprocessor form a master-slave multi-core decision center through a high-speed bus.
A control method of an aircraft comprises a takeoff stage, wherein the aircraft takes off vertically in a two-axis mode, climbs to 2m, and the wind direction is judged according to a hovering inclination angle after the aircraft rotates for a circle along the course; the aircraft flies in a two-axis mode and is accelerated against the wind; when the airspeed exceeds the critical airspeed, the aircraft flies in a fixed wing mode converted from a two-axis mode, and the takeoff is finished; in the landing stage, the aircraft enters a gliding route in a fixed wing mode, and the aircraft flies at a reduced speed by lowering the altitude; when the speed is lower than the critical airspeed, the aircraft is converted from a fixed wing mode to a two-axis mode; the aircraft keeps the altitude flying to a landing point in a two-axis mode; after reaching the landing point, the aircraft is suspended and landed; when the flight mode is switched in the task execution process, the acceleration and deceleration actions are directly carried out, and different flight modes are switched according to the airspeed.
The invention has the beneficial effects that:
1. simple lifting
The aircraft can take off only by manual or vertical assistance of a frame during taking off and can also be thrown and flown after being skilled. When the aircraft lands, the aircraft automatically completes 1.5m near-ground hovering and is manually recovered. Or the tail part can be automatically grounded, and then the controllable front end is slowly grounded.
2. High speed performance
The cross section area and the wingspan of the aircraft are small, the resistance is small, and the aircraft can fly at high speed due to abundant power. Under the power of 6s, the maximum thrust of the full throttle exceeds 3.4kg, the thrust-weight ratio is 2.26, and the flat flight speed of the aircraft exceeds 40m/s (possibly 50 m/s). Due to the adoption of double-engine vector active control and a larger wing area, the high-speed state active control attitude aerodynamic stability is excellent.
3. High low-speed detection capability
The aircraft adopts double vector active control and a larger wing area, the canard wing with a large elevation angle has obvious vortex pulling, the close-range coupling is obvious in lift increasing, and the aircraft can fly at a larger attack angle and a lower speed. After the head-on angle of attack is exceeded, the aircraft can be converted into a two-axis mode for hovering flight, and the whole flight envelope of the aircraft cannot stall due to the rapid response of the high-voltage motor. When the aircraft performs target detection, the aircraft can search targets at a large attack angle at a low speed, a two-axis mode can be switched to hover after suspicious targets are found, the targets are accurately identified, and operators can independently search the targets without training.
4. Simple and reliable structure
The whole machine action components are two motors and two canard wings, a traditional control surface is not provided, and the power control structure is centralized and convenient for machine body design and assembly. The wing has no control surface, can be integrally formed, and the wing profile is well kept and is convenient to process and assemble. The fuselage is the barrel type, and control structure is concentrated, and the equipment cabin is great, makes things convenient for modularization task load.
5. Low cost and quick production
The photoelectric task load and control power equipment of the aircraft are all installed at the front end, so that the integrated design is facilitated, and the control circuit and the action mechanism are integrated. The whole machine has fewer parts, lower processing cost, no instant impact overload in the whole flying process, reliable quality of the main control assembly and low cost, and can be produced quickly. The production assembly does not have complicated wiring and precise installation steps, can be produced quickly, and even can be assembled on site, so that the application range is greatly expanded.
Drawings
FIG. 1 is a perspective view of the aircraft of the present invention;
FIG. 2 is a front view of the aircraft of the present invention;
FIG. 3 is a side view of the aircraft of the present invention;
FIG. 4 is a schematic representation of fluid simulation data for a canard and airfoil of the present invention;
FIG. 5 is a topological relationship diagram of the control system, the motor, the canard wing and the energy management system of the present invention;
FIG. 6 is a block diagram of the control system components of the present invention;
FIG. 7 is a flow chart of a control strategy of the present invention;
FIG. 8 is a flow chart of the vertical take-off and landing operation of the present invention;
FIG. 9 is a state change diagram of the two-axis mode of the present invention converted into the fixed-wing mode;
FIG. 10 is a state diagram illustrating the fixed wing mode of the present invention being converted into a two-axis mode;
FIG. 11 is a diagram of hovering state in a two-axis mode according to the present invention;
FIG. 12 is a rolling state diagram in the two-axis mode of the present invention;
FIG. 13 is a view of the two-axis mode of the present invention in a pitch state;
FIG. 14 is a side view of the heading control in the two-axis mode of the present invention;
FIG. 15 is a top view of the heading control in the two-axis mode of the present invention;
FIG. 16 is a horizontal forward state view of the fixed wing mode of the present invention;
FIG. 17 is a fixed wing mode heading control state diagram of the present invention;
FIG. 18 is a fixed wing mode pitch state diagram of the present invention;
FIG. 19 is a fixed wing mode roll state diagram of the present invention;
in the figure, 1, a fuselage; 2. a propeller; 3. a motor; 4. duck wings; 5. an airfoil.
Detailed Description
The invention is further described with reference to the following figures and examples.
Embodiments of the invention referring to fig. 1 and 5, an aircraft comprises a fuselage, propellers, motors, canard wings, a control system and an energy management system,
the machine body is cylindrical, and the head and the tail of the machine body are in a rectifying conical shape, so that the induced resistance is reduced; the duck wing is trapezoidal and is positioned in the central axis position of the front part of the fuselage and is higher than the wings; the wing is a modified triangular wing, the front edge of the wing is in arc transition, the wing is installed at the middle and rear sections of the fuselage and is lower than the canard wing, the canard wing and the wing form a canard layout, the coupled vortex lift-up is realized, the detached vortex generated by the front wing of the canard wing and the attached vortex of the wing form favorable interference during flying, the flow field of the wing is improved, the wing lift force is increased, the vortex breakage is delayed, and the comprehensive aerodynamic performance of the aircraft is improved.
As shown in fig. 1-3, the duck-type layout is an advanced aerodynamic layout, and the design concept of the vortex separation flow configuration is successfully utilized in the aerodynamic aspect, the duck-type layout adopts close-range coupling, the duck wings fully and effectively utilize the vortex aerodynamic force, so that positive trim lift force is provided, and the direct force control of the flight attitude can be realized under the coordination of other control surfaces, so that the layout has extremely strong control capability under a large attack angle. The aerodynamic force of the duck-shaped layout is mainly influenced by the geometrical shapes of the duck wings and the main wings, the high lift initial attack angle of the layout is improved along with the increase of the sweepback angle of the main wings, and the high lift effect is more obvious on the small sweepback main wings.
As shown in fig. 4, the aircraft employs a close-coupled canard configuration in which the sweep angle of the canard and wing is determined by fluid simulation. In order to simplify the calculation, the propeller slipstream is neglected to carry out pneumatic simulation, and the canard swept-back of 40 degrees, 50 degrees, 60 degrees and 70 degrees and the wing swept-back of 40 degrees, 50 degrees and 60 degrees are respectively selected according to experience to carry out simulation. The simulation shows that the lift coefficient of the canard wing combined with the wing with the sweepback angle is greatly improved compared with that of the pure sweepback wing. Wherein, the combination of 70 degrees of duck wing sweepback and 40 degrees of wing sweepback has two slight fluctuations of lift coefficient at an attack angle of 0-54 degrees except at 30-36 degrees and 50-54 degrees, and the change of the overall lift coefficient is most gentle. The aircraft is mainly applied to a cruise task, relates to complex pneumatic and flight mode conversion, and is suitable for selecting the most stable and reliable pneumatic layout, so that the pneumatic layout of 70 degrees of canard wing sweep and 40 degrees of wing sweep is selected according to a simulation result.
As shown in fig. 5-6, the control system includes a sensing and collecting unit, a main control processor, a computing node, a communication module and a motion coprocessor, wherein the main control processor, the computing node, the communication module and the motion coprocessor are mounted on the airborne flight control board card, the sensing and collecting unit includes a GPS and/or beidou navigation system, an inertial navigation system, an electronic compass, an airspeed tube, an air thermometer, a vibration sensor and a structural deformation sensor, the GPS and/or beidou navigation system obtains position data of the aircraft, the inertial navigation device obtains attitude data of the aircraft, the electronic compass is used for obtaining heading data, the airspeed tube obtains airspeed data of the aircraft, the vibration sensor and the structural deformation sensor obtain body structural data,
the main control processor and the motion coprocessor are used for acquiring position data, attitude data, course data, airspeed data and body structure data through the sensing Internet of things and calculating, and are connected with the motion coprocessor through a high-speed bus to drive the canard wing and the motor to adjust the state of the aircraft; the two motors are arranged in the middle section of the front edge of the canard wing, the canard wing is completely soaked in the airflow of the propeller during flying, and the control system carries out vector control on the motors;
vector control means that motor and screw are directly installed on duck wing, and duck wing action directly changes the pulling force direction, directly produces corresponding control moment. Different from the traditional fixed wing pneumatic control surface, the propellers of the aircraft can generate direct control torque after rotating, the aircraft is not influenced by the flight airspeed of the aircraft, and the control torque in the whole flight envelope line has no control force failure.
As shown in fig. 5, the main control processor is used for operating a timing operation system, resolving positioning, orientation and attitude, executing a flight control algorithm, receiving a communication signal and debugging, the computing node is used for aircraft mode identification and system reconstruction, the motion coordination processing is used for motor driving and duck wing control, the main control processor, the computing node, the motion coprocessor and the sensing acquisition unit form a multi-core distributed networking hardware architecture, and data are interacted between the main control processor and the motion coprocessor through a high-speed bus communication mechanism; a star information internet of things is formed between the main control processor and the sensing acquisition unit; the master control processor forms a master-slave multi-core decision center with the computing nodes and the motion coprocessor through a high-speed bus;
the sensing acquisition unit and the main control processor form a star information internet of things, the main control processor actively schedules required sensing information in a normal state, the sensing acquisition unit can seize a network communication token according to priority in an emergency, the communication port triggers hardware interruption of the main control processor, and related emergency processing programs are responded in real time.
Further, the aircraft body is 800mm long, and the wingspan of the wing is 500 mm.
Further, the propeller adopts an F60kv1750 motor matched with a 6040 two-blade propeller, and the full throttle thrust-weight ratio of a 6s lithium battery exceeds 2.
Furthermore, the duck wing is combined with the machine body through the rotating shaft, the control system drives the rotating shaft to rotate, and the duck wing rotates along with the rotating shaft;
furthermore, the wings are integrally formed and made of high-density foam materials.
Furthermore, the action parts of the aircraft are two motors and two canard wings, no control surface is arranged, and a control system and a mission load are arranged in the aircraft body.
The aircraft is high in pneumatic efficiency, simple, free of a traditional pneumatic control surface and a vertical tail wing, small in resistance, strong in double-engine power, high in thrust-weight ratio and suitable for high-speed flight. The aircraft has no vertical tail, low height, convenient storage and transportation, no traditional control surface, simple structure and convenient large-scale production. Vortex pulled out by the canard wing can be coupled with vortex on the upper surface of the wing to improve the lift force of the wing, the speed is further reduced to zero, a two-axis mode is adopted for flying, the lift force is generated by motor thrust, the wing is almost vertical, and the lift force is not generated.
As shown in fig. 7-8, a method of controlling an aircraft,
in the takeoff stage, the aircraft takes off vertically in a two-axis mode, climbs to 2m, and the wind direction is judged according to the hovering inclination angle after the aircraft rotates for one circle along the course; the aircraft flies in a two-axis mode and is accelerated against the wind; when the airspeed exceeds the critical airspeed, the aircraft flies in a fixed wing mode converted from a two-axis mode, and the takeoff is finished;
in the landing stage, the aircraft enters a gliding route in a fixed wing mode, and the aircraft flies at a reduced speed by lowering the altitude; when the speed is lower than the critical airspeed, the aircraft is converted from a fixed wing mode to a two-axis mode; the aircraft keeps the altitude flying to a landing point in a two-axis mode; after reaching the landing point, the aircraft is suspended and landed;
when the flight mode is switched in the task execution process, the acceleration and deceleration actions are directly carried out, and different flight modes are switched according to the airspeed.
As shown in fig. 9, the process of switching from the two-axis mode to the fixed-wing mode is as follows: the aircraft keeps a hovering state and starts to be switched, the canard wings deflect forwards in a linkage mode, the aircraft accelerates gradually in a large attack angle posture, and when the airspeed of the aircraft exceeds a safe airspeed, the aircraft flies in a fixed wing mode gradually.
As shown in fig. 10, the process of switching from the fixed-wing mode to the two-axis mode is as follows: the aircraft starts to convert with fixed wing horizontal flight, the canard linkage deflects upwards, the aircraft attack angle gradually increases, the airspeed gradually decreases, and the aircraft gradually hovers for the vertical two-axis of two-axis mode by level transition.
The judgment standard for controlling mode switching is as follows: and the two-axis mode control is carried out when the air speed is less than or equal to the critical air speed, and the fixed wing mode control is carried out when the critical air speed is exceeded.
As shown in fig. 11, in the two-axis mode, the two motors are used for controlling the height in a linkage manner, differentially controlling the rolling of the aircraft, controlling the pitching of the aircraft in a linkage manner by using the canard wings, and differentially controlling the course by using the canard wings; the aircraft is approximately vertical to the horizontal plane, the lift force is completely from the propeller, the flight state of the aircraft is similar to that of a helicopter, the aircraft can hover, ascend and descend, and fly at a low speed front, back, left and right, and the power consumption is large;
as shown in FIG. 12, the aircraft roll is controlled by the double-motor differential, and the double-motor differential rotates to generate a lift difference and control the aircraft roll to move transversely.
As shown in fig. 13, the canard wing is in linkage control over the pitching of the aircraft, the canard wing is in linkage control in the same direction, and the lift force generated by the double motors is deflected to control the pitching motion of the aircraft.
As shown in fig. 14-15, the dual canard wing differential control aircraft course, the dual canard wing differential control, the dual motor driven propellers generate differential tension to control the aircraft course.
In a fixed wing mode, two motors are used for controlling airspeed in a linkage mode and controlling the course in a differential mode, canard wings are used for controlling pitching in a linkage mode, and canard wings are used for controlling rolling in a differential mode; the aircraft is approximately parallel to the horizontal plane, the lift force mainly comes from the wings, the flying state of the aircraft is similar to that of a fixed wing, the aircraft can fly forwards at a high speed, the electric energy consumption is low, and the flying efficiency is high. The fixed-wing mode control corresponds to a two-axis mode horizontally rotated by 90 °.
As shown in fig. 16, in the fixed wing mode, the dual motors are controlled in a linkage manner to push the aircraft to fly horizontally forward.
As shown in fig. 17, in the fixed wing mode, the dual-motor differential control, the dual-motor tension difference controls the aircraft heading.
As shown in fig. 18, in the fixed wing mode, the dual canard wing differential linkage control causes the dual-motor tension to deflect in the pitching direction, thereby controlling the pitching of the aircraft.
As shown in fig. 19, in the fixed wing mode, the dual canard wing differential control, the dual motors are deflected in opposite pitch directions, and the aircraft is controlled to roll.
When the two-axis vertical hovering mode and the fixed wing cruise mode are converted, the power is sufficient, the motor canard forms vector control, the distance between the gravity center of the whole aircraft and the canard is large, the aerodynamic area of the wing is large, the lift force is sufficient, and the aerodynamic damping is large, so that the control torque of the aircraft is sufficient, the aircraft is controllable in various states, and the congenital aerodynamic defects of the traditional aircraft such as stall and the like do not exist.
The above-described embodiment merely represents one embodiment of the present invention, but is not to be construed as limiting the scope of the present invention. It should be noted that, for a person skilled in the art, several variations and modifications can be made without departing from the inventive concept, which falls within the scope of the present invention.
Claims (10)
1. An aircraft comprises a fuselage, propellers, a motor, canard wings, a control system and an energy management system, wherein the fuselage is cylindrical, and the head and the tail of the fuselage are in a rectifying cone shape; the duck wing is trapezoidal and is positioned in the central axis position of the front part of the fuselage and is higher than the wings; the wings are in shape-modified delta wings, the front edges of the wings are in arc transition, the wings are arranged at the middle and rear sections of the fuselage, are lower than the canard wings and are in a lower single wing layout, and the canard wings and the wings form a canard layout to realize the coupling vortex lift increase; the sweep angle of the duck wing is 40 degrees, 50 degrees, 60 degrees or 70 degrees, and the sweep angle of the wing is 40 degrees, 50 degrees or 60 degrees;
the control system comprises a sensing acquisition unit, a main control processor, a calculation node, a communication module and a motion coprocessor, wherein the main control processor, the calculation node, the communication module and the motion coprocessor are installed on an airborne flight control board card, the sensing acquisition unit comprises a GPS and/or Beidou navigation system, an inertial navigation system, an electronic compass, an airspeed tube, a thermometer, a vibration sensor and a structural deformation sensor, the GPS and/or Beidou system acquires position data of an aircraft, the inertial navigation device acquires attitude data of the aircraft, the electronic compass is used for acquiring course data, the airspeed tube acquires airspeed data of the aircraft, and the vibration sensor and the structural deformation sensor acquire body structural data;
the main control processor is used for operating the timing operation system, resolving positioning, orientation and attitude, executing a flight control algorithm, receiving a communication signal and debugging, the computing node is used for aircraft mode identification and system reconstruction, the motion coprocessing is used for motor driving and canard wing control, the main control processor, the computing node, the motion coprocessor and the sensing acquisition unit form a multi-core distributed networking hardware architecture, and data are interacted between the main control processor and the motion coprocessor through a high-speed bus communication mechanism; a star information internet of things is formed between the main control processor and the sensing acquisition unit; the master control processor, the computing nodes and the motion coprocessor form a master-slave multi-core decision center through a high-speed bus.
2. The aircraft of claim 1, wherein: the sweep angle of the duck wing is 70 degrees, and the sweep angle of the wing is 40 degrees.
3. The aircraft of claim 1, wherein: the main control processor and the motion coprocessor are used for acquiring position data, attitude data, course data, airspeed data and body structure data through the sensing Internet of things and calculating, and are connected with the motion coprocessor through a high-speed bus to drive the canard wing and the motor to adjust the state of the aircraft; the two motors are arranged in the middle section of the front edge of the canard wing, the canard wing is completely soaked in the airflow of the propeller during flying, and the control system carries out vector control on the motors.
4. The aircraft of claim 1, wherein: the sensing acquisition unit and the main control processor form a star information internet of things, the main control processor actively schedules required sensing information in a normal state, the sensing acquisition unit can seize a network communication token according to priority in an emergency, the communication port triggers hardware interruption of the main control processor, and related emergency processing programs are responded in real time.
5. The aircraft of claim 1, wherein: the duck wing is combined with the machine body through the rotating shaft, the control system drives the rotating shaft to rotate, and the duck wing rotates along with the rotating shaft; the action parts of the aircraft are two motors and two canard wings without control surfaces, and a control system and a mission load are arranged in the aircraft body.
6. A method of controlling an aircraft according to any one of claims 1 to 5, characterized in that:
in the takeoff stage, the aircraft takes off vertically in a two-axis mode, climbs to 2m, and the wind direction is judged according to the hovering inclination angle after the aircraft rotates for one circle along the course; the aircraft flies in a two-axis mode in an upwind acceleration mode; when the airspeed exceeds the critical airspeed, the aircraft flies in a fixed wing mode converted from a two-axis mode, and the takeoff is finished;
in the landing stage, the aircraft enters a gliding route in a fixed wing mode, and the aircraft flies at a reduced speed by lowering the altitude; when the speed is lower than the critical airspeed, the aircraft is converted from a fixed wing mode to a two-axis mode; the aircraft keeps the altitude flying to a landing point in a two-axis mode; after reaching the landing point, the aircraft is suspended and landed;
when the flight mode is switched in the task execution process, the acceleration and deceleration actions are directly carried out, and different flight modes are switched according to the airspeed.
7. The control method according to claim 6, characterized in that:
the process of converting the two-axis mode into the fixed wing mode is as follows: the aircraft keeps a hovering state and starts to be switched, the canard wings deflect forwards in a linkage manner, the aircraft gradually accelerates in a large attack angle posture, and when the airspeed of the aircraft exceeds a safe airspeed, the aircraft gradually excessively flies in a fixed wing mode;
the process of converting the fixed wing mode into the two-axis mode is as follows: the aircraft starts to convert with fixed wing horizontal flight, the canard linkage deflects upwards, the aircraft attack angle gradually increases, the airspeed gradually decreases, and the aircraft gradually hovers for the vertical two-axis of two-axis mode by level transition.
8. The control method according to claim 7, wherein the determination criterion for controlling the mode switching is: and the two-axis mode control is carried out when the air speed is less than or equal to the critical air speed, and the fixed wing mode control is carried out when the critical air speed is exceeded.
9. The control method according to claim 7, characterized in that: in a two-axis mode, the two motors are used for controlling the height in a linkage manner, differentially controlling the rolling of the aircraft, controlling the pitching of the aircraft in a linkage manner by using the canard wings, and differentially controlling the course by using the canard wings; the aircraft is approximately vertical to the horizontal plane, the lifting force is totally from the propeller, and the aircraft can hover, ascend and descend, and fly forwards, backwards, leftwards and rightwards at a low speed.
10. The control method according to claim 7, characterized in that: under the fixed wing mode, two motors are used for controlling airspeed in a linkage mode and controlling course in a differential mode, canard wing is used for controlling pitching in a linkage mode, and canard wing is used for controlling rolling in a differential mode; the aircraft is approximately parallel to the horizontal plane, the lift force is mainly from the wings, and the aircraft can fly forwards at high speed.
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