CN107176298B

CN107176298B - Aircraft flight control method and aircraft

Info

Publication number: CN107176298B
Application number: CN201710393612.5A
Authority: CN
Inventors: 何春旺
Original assignee: Zhuhai Pan Lei Intelligent Technology Co Ltd
Current assignee: Zhuhai Pan Lei Intelligent Technology Co Ltd
Priority date: 2017-05-28
Filing date: 2017-05-28
Publication date: 2020-01-03
Anticipated expiration: 2037-05-28
Also published as: CN107176298A

Abstract

The invention provides a flight control method of an aircraft, wherein the aircraft comprises an aircraft body, an aircraft arm, a power unit and a flight control device, wherein the power unit is arranged on the aircraft body through the aircraft arm, and the flight control device is arranged in the aircraft body; the power unit comprises a first rotor wing and a second rotor wing which are symmetrically distributed on two sides of the aircraft body, and further comprises a third rotor wing and a fourth rotor wing which are symmetrically distributed on two sides of the aircraft body and can tilt, and the respective tilting shafts of the third rotor wing and the fourth rotor wing are approximately distributed along the length direction of the aircraft; and the tilting driving device is used for driving the third rotor and the fourth rotor to tilt. The flight control method comprises the following steps: the tilting driving device receives a tilting control instruction of the flight control device; the tilting driving device drives the third rotor wing and the fourth rotor wing to tilt between a first position and a second position; when third rotor and fourth rotor tilt, first rotor and second rotor are located preset position. The aircraft is beneficial to air flow concentration or diffusion, and overcomes the defects of large weight, large energy consumption and reduced maneuverability of the conventional aircraft.

Description

Aircraft flight control method and aircraft

Technical Field

The invention relates to an aircraft flight control method and an aircraft using the control method.

Background

The aircraft comprises an unmanned aerial vehicle piloting aircraft and a manned aircraft, wherein the unmanned aerial vehicle piloting aircraft is called an unmanned aerial vehicle for short, and is an unmanned aerial vehicle controlled by radio remote control equipment or an embedded program. The present application is more extensive is rotor type unmanned aerial vehicle, and its power comes from the lift that the high-speed rotation of rotor provided, realizes unmanned aerial vehicle's different actions such as going up and down to hover through the different rotational speed of controlling every rotor.

In order to improve the stability of the flight of a multiaxial aircraft, multiaxial aircraft with simultaneously changing wheelbases have appeared. However, the flight mode of the aircraft with synchronous change increases the control device, so that the weight of the aircraft is increased, the energy consumption is high, and the maneuvering performance is reduced.

Disclosure of Invention

The invention provides an aircraft flight control method and an aircraft using the control method, and solves the problems of low maneuverability and high energy consumption.

In order to achieve the purpose, the invention provides an aircraft flight control method, wherein an aircraft comprises a fuselage, a horn, a power unit and a flight control device, the power unit is installed on the fuselage through the horn, and the flight control device is installed in the fuselage; the power unit comprises a first rotor and a second rotor which are symmetrically distributed on two sides of the machine body; the power unit also comprises a third rotor wing and a fourth rotor wing which can tilt and are symmetrically distributed on two sides of the aircraft body, and the respective tilting shafts of the third rotor wing and the fourth rotor wing are distributed along the length direction of the aircraft; and the tilting driving device is used for driving the third rotor and the fourth rotor to tilt.

The flight control method comprises the following steps:

s1, receiving a tilting control instruction of a flight control device by a tilting driving device;

s2, the tilting driving device drives the third rotor wing and the fourth rotor wing to tilt between a first position and a second position;

s3, when third rotor and fourth rotor vert, first rotor and second rotor are located preset position.

The rotor discharges the interior below of air current directional aircraft when the rotor is located the first position, moves the rotor and discharges the outer below of air current directional aircraft when being located the second position.

According to the scheme, the aircraft flight control method is that the flight control system can actively control the tilting positions of the third rotor and the fourth rotor, and when the airflow direction is flexibly changed and the exhausted airflow is directed to the inner lower part, the aircraft flight control method is favorable for airflow concentration, strong power and rapid take-off speed; when the exhaust airflow points to the outer lower part, the supporting span of the aircraft is increased, which is beneficial to the stable flight of the aircraft; obviously improves the maneuverability of the flying of the aircraft, reduces the control devices, reduces the weight of the aircraft and reduces the energy consumption.

In a further scheme, during take-off, the third rotor wing and the fourth rotor wing are close to the first position; after takeoff, the third rotor and the fourth rotor are proximate to the second position. Therefore, the flight control method of the aircraft can adjust the rotor wing according to different requirements of environmental conditions, and improves the adaptability of the aircraft.

Further, the diameter of the third rotor and the fourth rotor is smaller than the diameter of the first rotor and the second rotor. Thus, the aircraft tail weight is further reduced to enhance the maneuverability of the aircraft control.

The further scheme is that the wheel base of the third rotor and the fourth rotor is smaller than the wheel base of the first rotor and the second rotor. Therefore, the air flow is concentrated, and the driving capability of the aircraft is enhanced.

The protection frame comprises an upper wire net, a lower wire net and a side wall protection frame; the power unit also comprises a driving motor; the upper wire mesh, the lower wire mesh and the side wall protection frame form an accommodating space of the protection frame, and the rotor wing is arranged in the accommodating space; the rotor wing is arranged on a rotating shaft of the driving motor, and the driving motor is arranged on the horn in a tilting way through the mounting seat; the middle part of the upper wire net is arranged on the machine arm.

According to the scheme, the driving rotor wing is protected in the accommodating space of the protective frame of the aircraft, when the aircraft is subjected to rigid impact, for example, the aircraft falls on the ground, the impact force is transmitted to the upper wire net from the lower wire net through the side wall protective frame or transmitted to the upper wire net from the side wall protective frame, and then the impact force is transmitted to the driving unit or the horn through the upper wire net; through multistage transmission, the influence of rigid impact on the fuselage and the power unit is obviously weakened, and because the driving rotor wing is suspended in the accommodating space of the protective frame, the damage of impact force on the power unit is greatly weakened.

The further proposal is that the outer end of the machine arm is provided with an installation seat positioned at the upper side and an installation plate positioned at the lower side; the middle part of the upper wire net is fixed between the mounting seat and the mounting plate; the driving motor is fixed on the upper side of the mounting seat and is directly connected with the horn, so that the rotary inertia caused by the protective frame is reduced.

In another further scheme, the driving motor is fixed in the accommodating space of the protective frame on the lower side of the mounting plate. The influence of impact on the rotor, the driving motor and the machine body can be further reduced, and the heat dissipation, installation, maintenance and replacement of the motor are facilitated.

The further scheme is that the inclination driving device comprises a steering engine and an inclination driving connecting rod; the inclined connecting rod comprises a first driving connecting rod, a second driving connecting rod, a third driving connecting rod and a fourth driving connecting rod, and all the driving connecting rods are sequentially hinged end to end; one end of the inclined driving connecting rod is hinged with the protection frame, and the other end of the inclined driving connecting rod is hinged with the steering engine. It is thus clear that the scheme is that slope actuating mechanism can the free control protective frame position state, and drive protective frame and third rotor, fourth rotor vert between primary importance and second place around the hinge, make the aircraft adapt to multiple complicated flight environment at any time, simple structure moreover, easily processing, the connecting rod constitutes laborsaving lever, can control verting of rotor with less rudder, is favorable to reducing aircraft weight.

Drawings

FIG. 1 is a schematic view of an aircraft controlled by a first embodiment aircraft flight control method

FIG. 2 is a schematic sectional view of the boom, the bezel and the power unit in the first embodiment;

FIG. 3 is a schematic view of another form of mounting of the motor in the first embodiment;

FIG. 4 is a schematic view of the tilt driving mechanism in the first embodiment;

FIG. 5 is a schematic view of a first position of the bezel in the first embodiment;

FIG. 6 is a schematic view of a second position of the bezel in the first embodiment;

fig. 7 is a schematic illustration of the arm distribution of an aircraft in a second embodiment.

Detailed Description

The invention is described below with reference to specific embodiments and with reference to the drawings.

First embodiment

As shown in fig. 1 and 2, the aircraft 10 mainly includes a fuselage 1, a horn 20, a fender frame 3, and a power unit 4. The horn 20 includes a first horn 21, a second horn 22, a third horn 23, and a fourth horn 24. The power unit 4 comprises a first rotor 411 and a second rotor 412 which are symmetrically distributed on two sides of the machine body, the first rotor 411 and the second rotor 412 have equal diameters, and the pitch angles are opposite; the power unit further comprises a third rotor 423 and a fourth rotor 414 which are symmetrically distributed on two sides of the fuselage, the tilting axes of the third rotor 413 and the fourth rotor 414 are distributed along the length direction 0 of the aircraft, and the diameters of the third rotor 423 and the fourth rotor 414 are equal.

The protective frame 3 comprises an upper net 31, a side wall protective frame 32 and a lower net 33; the power unit 4 comprises a rotor 41 and a driving motor 42; the mount 5 includes a mount 51, a mounting plate 52, and a power mount 53. The upper wire mesh 31, the side wall protective frame 32 and the lower wire mesh 33 are combined to form an accommodating space 34 of the power protective frame 3; rotor 41 is disposed within receiving space 34 without contacting the surroundings. Rotor 41 is mounted on the rotation shaft of drive motor 42, drive motor 42 is disposed in power mount 53, and drive motor 42 and power mount 53 are simultaneously mounted on the mount. The upper wire net 31 is installed between the installation seat 51 and the installation plate 52, and the upper wire net 31 is tightly fixed by the assembly of the installation seat 51 and the installation plate 52. The middle of the protection frame 3 passing through the upper wire net 31 is mounted on the horn 2.

It can be seen from this solution that the protective frame 3 protects the rotor 41 in its accommodation space 34, and when the aircraft 10 is subjected to a rigid impact, for example, when it falls on the ground, the impact force is transmitted from the lower wire mesh 33 through the side wall protective frame 32 to the upper wire mesh 31 or from the side wall protective frame 32 to the upper wire mesh 31, and then the impact force is transmitted through the upper wire mesh 31 to the drive unit 4 and the horn 2; due to multiple transmission, the influence of rigid impact on the fuselage 1 and the power unit 4 is obviously weakened, and because the rotor 41 is suspended in the accommodating space 34 of the protective frame 3, the damage of impact force on the power unit 4 is greatly weakened, and the damage of impact force on the fuselage 1 is weakened at the same time. In addition, because the horn 2 is directly connected to the power unit 4, under the same condition, the structure can obviously reduce the rotational inertia of the power unit and improve the response speed of the power unit 4 to the aircraft 10.

As shown in fig. 3, it is preferable that the driving motor 422 is also mounted on the mounting plate 52 and is mounted with the rotor 41 in the accommodating space 34, so as to facilitate installation and maintenance of the driving motor 422 and the rotor 41, and further facilitate heat dissipation of the driving motor 422. Therefore, the installation position of the driving motor can be flexibly set according to different working requirements and structural requirements of the aircraft.

As shown in fig. 4, a tilt driving device is disposed in the horn 20, and the tilt driving device includes a steering gear 61, a first driving link 62, a second driving link 63, a third driving link 64, and a fourth driving link 65. The arm 20 is provided with a first hinge device 71 and a second hinge device 72. The protection frame 3 is hinged with the second hinge device 72 on the horn 20 through the power mounting base 53, the hinge shaft is arranged along the width direction of the horn 20, and the width direction of the horn 20 is the direction perpendicular to the paper surface, so that the protection frame 3 can freely tilt inwards and outwards. The first driving connecting rod 62, the second driving connecting rod 63, the third driving connecting rod 64 and the fourth driving connecting rod 65 are sequentially hinged, the other end of the first driving connecting rod 62 is hinged with the steering engine 61, the other end of the fourth motor driving connecting rod 65 is hinged with the protective frame 3, the middle of the third driving connecting rod 64 is hinged with the first hinge device 71, a labor-saving connecting rod is formed, and the whole inclination driving device is positioned and installed. According to different flight requirements of the aircraft 10, the steering engine 61 can be controlled by a flight control device (not shown) to drive the first driving connecting rod 62, the second driving connecting rod 63, the third driving connecting rod 64 and the fourth driving connecting rod 65, so that the position states of the protective frame 3 and the driving rotor 41 can be actively controlled, and the protective frame can tilt between a first position and a second position. Fig. 4 shows the horizontal state of the protection frame 3. When the first drive connecting rod anticlockwise rotation of flight control device control steering wheel 61 work drive, drive the motion of whole slope drive arrangement simultaneously, drive protecting frame 3 and use second hinge means 72 as the fixed point, vert around the hinge, order about protecting frame 3 and lean out, be in the second position, as shown in fig. 5, make the supporting span grow to aircraft 10, be favorable to aircraft 10 flight steadily, prevent collision or aircraft 10 turns on one's side suddenly in narrow space. When the flight control device controls the steering engine 61 to work and drive the first driving connecting rod to rotate clockwise, the whole inclination driving device is driven to move simultaneously, the protection frame 3 is driven to incline around the hinge shaft by taking the second hinge device 72 as a fixed point, the protection frame 3 is driven to incline inwards, and the protection frame is located at a first position, as shown in fig. 6, the airflow concentration is facilitated, the power is strong, and the takeoff speed is rapid.

Preferably, the aircraft is provided with a first rotor and a second rotor which are symmetrically distributed on two sides of the fuselage, the first rotor and the second rotor have equal diameters, opposite pitch angles and are not installed.

Second embodiment

The present embodiment differs from the first embodiment in the arrangement of the arms and rotors of the aircraft.

As shown in fig. 7, the first horn 415, the second horn 416, the third horn 417, and the fourth horn 418 are uniformly distributed along the circumferential direction of the aircraft. The tilt axes of the rotor on the third horn 417 and the rotor on the fourth horn 418 are distributed substantially along the length of the aircraft. Therefore, the aircraft can generate backward airflow component, thereby increasing the forward thrust without greatly inclining the aircraft body, and improving the flight speed and the capability of adapting to the environment.

The flight control method of the aircraft comprises the following steps:

The foregoing is a more detailed description of the present invention, taken in conjunction with the specific preferred embodiments thereof, and it is not intended that the invention be limited to the specific embodiments shown and described. For those skilled in the art to which the invention pertains, equivalent substitutions or obvious modifications may be made without departing from the spirit of the invention, and the same properties or uses are deemed to fall within the scope of the invention as defined by the claims as filed.

Claims

1. The aircraft flight control method is characterized in that the aircraft comprises a fuselage, a horn, a power unit and a flight control device, wherein the power unit is installed on the fuselage through the horn, and the flight control device is installed in the fuselage;

the power unit comprises a first rotor and a second rotor which are symmetrically distributed on two sides of the fuselage;

the power unit further comprises a third tilting rotor and a fourth tilting rotor which are symmetrically distributed on two sides of the fuselage, and the tilting shafts of the third tilting rotor and the fourth tilting rotor are distributed along the length direction of the aircraft;

the tilting driving device is used for driving the third rotor and the fourth rotor to tilt;

the protection frame comprises an upper wire mesh, a lower wire mesh and a side wall protection frame;

the middle part of the upper wire net is arranged on the machine arm;

the power unit further comprises a driving motor;

the driving motor is tiltably arranged on the machine arm through a mounting seat;

a tilting driving device serving as the tilting driving device is arranged in the machine arm;

the inclination driving device comprises a steering engine and an inclination driving connecting rod; the inclined connecting rod comprises a first driving connecting rod, a second driving connecting rod, a third driving connecting rod and a fourth driving connecting rod, and all the driving connecting rods are sequentially hinged end to end; one end of the inclined driving connecting rod is hinged with the protection frame, and the other end of the inclined driving connecting rod is hinged with the steering engine;

the flight control method comprises the following steps:

s1, the tilting driving device receives a tilting control command of the flight control device;

s2, the tilt driving device drives the third rotor and the fourth rotor to tilt between a first position and a second position;

s3, when the third rotor and the fourth rotor tilt, the first rotor and the second rotor are located at preset positions.

2. The aircraft flight control method of claim 1, wherein:

at takeoff, the third and fourth rotors are proximate to the first position;

after takeoff, the third rotor and the fourth rotor are proximate to the second position.

3. An aircraft using the aircraft flight control method of claim 1 or 2, wherein:

the diameters of the third rotor and the fourth rotor are smaller than the diameters of the first rotor and the second rotor.

4. The aircraft of claim 3, wherein:

the wheelbase of the third rotor and the fourth rotor is smaller than the wheelbase of the first rotor and the second rotor.

5. The aircraft of claim 3 or 4, wherein:

the upper wire mesh, the lower wire mesh and the side wall protection frame form an accommodating space of the protection frame, and the rotor wing is arranged in the accommodating space; the rotor is installed on the rotation axis of the driving motor.

6. The aircraft of claim 5, wherein:

the mounting seat is positioned at the outer end of the machine arm, and a mounting plate is arranged on the lower side of the mounting seat;

the middle part of the upper wire net is fixed between the mounting seat and the mounting plate;

the driving motor is fixed on the upper side of the mounting seat.

7. The aircraft of claim 6, wherein:

the driving motor is fixedly arranged on the lower side of the mounting plate.