CN110015415B - Double-shaft tilting four-rotor aircraft - Google Patents

Double-shaft tilting four-rotor aircraft Download PDF

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Publication number
CN110015415B
CN110015415B CN201910382221.2A CN201910382221A CN110015415B CN 110015415 B CN110015415 B CN 110015415B CN 201910382221 A CN201910382221 A CN 201910382221A CN 110015415 B CN110015415 B CN 110015415B
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aircraft
unit
rotor
controller unit
module
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CN110015415A (en
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雷瑶
叶艺强
王金利
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Fuzhou University
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Fuzhou University
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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C27/00—Rotorcraft; Rotors peculiar thereto
    • B64C27/04—Helicopters
    • B64C27/08—Helicopters with two or more rotors
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C27/00—Rotorcraft; Rotors peculiar thereto
    • B64C27/52—Tilting of rotor bodily relative to fuselage
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
    • B64U10/00—Type of UAV
    • B64U10/10—Rotorcrafts
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
    • 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
    • G05D1/0825—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability using mathematical models
    • G—PHYSICS
    • G08—SIGNALLING
    • G08C—TRANSMISSION SYSTEMS FOR MEASURED VALUES, CONTROL OR SIMILAR SIGNALS
    • G08C17/00—Arrangements for transmitting signals characterised by the use of a wireless electrical link
    • G08C17/02—Arrangements for transmitting signals characterised by the use of a wireless electrical link using a radio link
    • H—ELECTRICITY
    • H04—ELECTRIC COMMUNICATION TECHNIQUE
    • H04W—WIRELESS COMMUNICATION NETWORKS
    • H04W76/00—Connection management
    • H04W76/10—Connection setup

Abstract

The invention provides a double-shaft tilting four-rotor aircraft, wherein a fuselage comprises a plurality of power arms uniformly arranged at the periphery of the fuselage, and a sensor unit, a controller unit, a control distributor unit, a vehicle dynamics system module and a wireless communication module which are arranged at the fuselage; the wireless communication module can establish wireless communication with the ground control station; the power arm comprises a rotor wing and a tilting mechanism; the rotor is driven by a driving motor at a motor fixing seat supported by the tilting mechanism; the tilting mechanism is connected with the controller unit; the controller unit adjusts the rotor position and rotor orientation via a tilt mechanism to adjust the flight of the aircraft; according to the invention, the rotating speeds and the tilting angles of the four rotors can be automatically regulated according to the input instruction, the self flight state is fed back by utilizing the stability of the closed-loop control system, the self flight regulation and control work is automatically completed, and the purposes of flight pitching, rolling and yawing are achieved; the high efficiency and the high reliability of the flight regulation are ensured.

Description

Double-shaft tilting four-rotor aircraft
Technical Field
The invention relates to the technical field of small unmanned aircrafts, in particular to a double-shaft inclined four-rotor aircraft.
Background
The double-shaft tilting rotor wing configuration is that the rotation plane of the rotor wing and the plane of the machine body form a certain included angle through the rotation of the supporting arm and the motor fixing seat, and the rotor wing of the aircraft can achieve the effect of tilting around the driving motor due to the characteristic of double-shaft tilting, so that pitching, rolling and yawing movement are easy to perform, and the purpose of flexible flying of the aircraft is achieved. Most of the supporting arms and motor fixing seats of the traditional four-rotor aircraft are fixedly arranged on an aircraft body, four rotors of the aircraft are in rotary motion in the same plane, the rotors can realize pitching, rolling and yawing motions only by changing the rotating speed of a driving motor, four variable-speed driving motors and an adjusting control system are needed, the strain function is poor, the process is not stable enough, and if the aircraft needs to complete the above 3 motions in time, the first method is to additionally install a driver at a nose or additionally install a driver and a tail pipe at a tail part to meet the flight requirement. However, this approach, with the increase in equipment, also entails an increase in body weight, requiring more rotors or lengthening propellers to maintain the hovering motion of the aircraft, by which the reliability of the aircraft is enhanced, but its flexibility is difficult to guarantee. The second method is a method using a variable pitch propeller which maintains a high flexibility of the aircraft in terms of trying to maintain the weight of the aircraft, but which requires the use of more complex propeller and servo motor systems, which is complex to implement, has a high failure rate, and may be accompanied by personnel damage in the event of failure. The method for realizing flexible flying movement of the multi-rotor aircraft is low in lifting force and weight, or complex in structure and low in efficiency, and the wide application of the method in various fields is greatly restricted.
Disclosure of Invention
The invention provides a double-shaft tilting four-rotor aircraft, which can automatically adjust the rotating speeds and the tilting angles of four rotors according to input instructions, can perform remote control, can feed back the self-flying state according to different flying environment conditions by utilizing the stability of a closed-loop control system, and automatically complete the self-flying regulation and control work, and can achieve the purposes of flying pitching, rolling and yawing by tilting the rotors; the high efficiency and the high reliability of the flight regulation and control are guaranteed, and the size of the aircraft can be reduced, so that the flexibility is improved.
The invention adopts the following technical scheme.
A biaxial tilting four-rotor aircraft, the fuselage (1) of the aircraft comprises a plurality of power arms uniformly arranged at the periphery of the fuselage, and a sensor unit (2), a controller unit (3), a control distributor unit (4), a vehicle dynamics system module (5) and a wireless communication module (6) which are arranged at the fuselage; the wireless communication module can establish wireless communication with a ground control station (7);
the power arm comprises a rotor (108) and a tilting mechanism (115); the rotor is driven by a driving motor (107) at a motor fixing seat (105) supported by the tilting mechanism; the tilting mechanism is connected with the controller unit; the controller unit adjusts rotor position and rotor orientation via a tilt mechanism to adjust the flight of the aircraft.
The number of the power arms is four.
The tilting mechanism is connected with a servo motor (106) at the machine body through the initial end of a supporting arm (103), and is hinged with the middle part of a motor fixing seat through the tail end of the supporting arm; the servo motor drives the supporting arm to rotate so as to enable the rotor wing to tilt; a fixed block (111) for fixing the steering engine (109) is arranged at the support arm; the steering engine is connected with the lower end of the motor fixing seat through a connecting rod mechanism (104), and the steering engine drives the motor fixing seat to swing through the connecting rod mechanism so that the rotor wing generates yaw force.
The connecting rod mechanism is a parallelogram mechanism; the tilt angle α of the rotor can be adjusted between 0 ° and 90 °, and the tilt angle α can be 0 ° but cannot be 90 °.
The fuselage further comprises landing gear (101) and electrical equipment bay (102); a gravity sensor (112), a gyroscope (113), remote sensing (114) and a plurality of batteries (116) are arranged in the electric equipment cabin; the output shaft of the servo motor and the contact position of the rotor support arm and the electrical equipment cabin (102) are connected through a bearing (110).
The sensor unit (2) is responsible for the connection work between the sensor unit and the controller unit (3), receives the measurement information of the sensor unit and converts the received measurement information into an electric signal according to a certain rule to be output to the controller unit (3); the controller unit (3) performs corresponding processing and sends out control command signals after finishing receiving the information transmitted by the sensor unit (2) and the wireless communication module (5);
the control distributor unit (4) completes distribution work of control command signals of the controller unit (3), sequentially distributes command information received by the controller unit (3) to four servo motors (106), four driving motors (107) and four steering engines (109) of the aircraft, and realizes adjustment of flight states;
the vehicle dynamics system module (5) comprises an actuator unit (501) and an inductor unit (502), the inductor unit (502) comprises a gravity sensing module (503), a perturbation sensing module (504), a gravity sensor (111), a gyroscope (113) and remote sensing (114), the instruction for controlling the distributor unit (4) is executed, and the vehicle dynamics system module performs real-time gravity and perturbation external force measurement through the inductor unit (502) and feeds back to the controller unit (3) to form a closed loop system;
the wireless communication module (6) comprises a transmitting module (601) and a receiving module (602), the working modes of the transmitting module (601) and the receiving module (602) can be converted and do not interfere with each other, the transmitting module (602) is connected with the controller unit (3), and the receiving module (602) is in communication with the ground control station (7) in a wireless mode so that the ground control station (7) can execute azimuth instruction operation on the aircraft.
The sensor unit (2) is an IMU consisting of an SBG system and an IG-500, which unit has an embedded processor for outputting filtered attitude and position data; the controller unit (3) is based on a PD controller with 3 SISDs, which processes the chip set maximum criterion of Digilent of the board for reading the signals; the controller unit (2) reads serial information from the output signal of the sensor unit (2) at MAX32 and instructs the servo motor (106), the driving motor (107) and the steering engine (109) by sending PWM signals:
the receiving module (602) in the wireless communication module (6) communicates with the ground control station (7) through a Spekdrum satellite receiver and a DX6i RC transmitter.
The absolute rotating speed of the output shaft of the driving motor is the sum of the angular speed of the aircraft in three directions of XYZ and the vector sum of the rotating speeds of the steering engine and the servo motor, the physical characteristics of four supporting arms of the aircraft are consistent, one supporting arm is taken for carrying out stress analysis, and the absolute rotating speed of the output shaft of the driving motor is determined
Angular acceleration of
Intermediate vectorA unit vector representing a reference frame j, represented by a reference frame i; p, q and r represent the angular velocities of the aircraft in three directions of the reference coordinate system XYZ, respectively; η and γ represent tilt angles of the biaxial during servo motion; η 'γ' represents the tilt angular velocity of the biaxial during servo motion; omega is expressed as the rotational speed of the drive motor;
using Euler's equation 3 M＝ 3 I 3 α+ 3 ω× 3 I 3 Omega (formula 3), calculating and obtaining torque of output shaft of driving motor 3 M, then utilize rotation matrix R 3to1 Calculating the moment of the aircraft itself in the reference system 1 M Gvro The calculation formula is as follows:
the thrust torque of the connecting rod structure is as follows:
wherein L is the length of the parallelogram mechanism; h is the height of the parallelogram mechanism;
the thrust and torque coefficients of the rotor are defined as:
T＝ρA(ΩR) 2 C T (6) And q=ρa (Ω R) 2 RC Q (equation 7);
c in the formula T And C o Is the thrust and torque coefficient of the connecting rod structure.
The total thrust moment of the aircraft is:
the absolute stress condition of the aircraft in six degrees of freedom is as follows:
wherein X, Y and Z are expressed as absolute forces of the aircraft on the XYZ axes, respectively; l, M and N are expressed as absolute moment of action of the aircraft on XYZ axes, respectively; u, v and w are expressed as the angular velocity of the aircraft as a whole in XYZ axes, respectively; when the aircraft is under comprehensive stress, the absolute angular rates of the aircraft in three directions of XYZ coordinate axes are as follows:
where Φ ', θ ' and ψ ' are expressed as the absolute angular velocities of the aircraft in three directions of the XYZ coordinate axes, respectively; phi is the included angle between the central axis of the aircraft body and the Z axis of the longitudinal axis; θ is the angle between the projection line of the central axis of the aircraft fuselage on the XY plane and the X axis.
According to the novel double-shaft tilting four-rotor aircraft provided by the invention, the rotor wing can be tilted by utilizing the rotation of the supporting arm and the motor fixing seat, then the purpose of free flight strain is achieved through the tilting rotor wing, the self flight state can be fed back by utilizing the stability of the closed-loop control system according to different flight environment conditions, the self flight regulation and control work is automatically completed, and the application of a plurality of power system devices is reduced due to the optimization of the body structure of the aircraft, so that the scale size of the aircraft is reduced to a certain extent, and the aircraft becomes more flexible to fly. The novel double-shaft tilting four-rotor aircraft has the outstanding advantages of convenience, safety and intelligence, and has wide application prospects in military and civil use.
Drawings
The invention is described in further detail below with reference to the attached drawings and detailed description:
FIG. 1 is a schematic diagram of an operation control system of the present invention;
FIG. 2 is a schematic top view of an aircraft according to the present invention;
FIG. 3 is a schematic view of the tilting mechanism of the present invention;
FIG. 4 is a schematic diagram of a vehicle dynamics system module of the present invention;
FIG. 5 is a schematic diagram of the communication of information between the controller unit and the ground control station of the present invention;
FIG. 6 is a schematic illustration of the communication between the controller unit and each of the actuator and sensor units of the present invention;
figure 7 is a schematic view of the reference frame of an aircraft according to the present invention (with one rotor in a tilted state) during flight.
In the accompanying drawings: 1-a fuselage; a 2-sensor unit; 3-a controller unit; 4-a control dispenser unit; 5-a vehicle dynamics system module; 6-a wireless communication module; 7-a ground control station;
101-landing gear; 102-an electrical equipment bay; 103-supporting arms; 104-a linkage mechanism; 105-a motor fixing seat; 106-a servo motor; 107-driving a motor; 108-rotor wing; 109-steering engine; 110-a bearing; 111-fixing blocks; 112-a gravity sensor; 113-gyroscopes; 114-remote sensing; 115-tilting mechanism; 116-cell;
501-an actuator unit; 502-an inductor unit; 503-a gravity sensing module; 504-a perturbation sensing module;
601-a transmitting module; 602-a receiving module.
Detailed Description
As shown in fig. 1-7, a biaxial tilting quadrotor aircraft, the fuselage 1 of which comprises a plurality of power arms uniformly arranged at the periphery of the fuselage, and a sensor unit 2, a controller unit 3, a control distributor unit 4, a vehicle dynamics system module 5 and a wireless communication module 6 arranged at the fuselage; the wireless communication module can establish wireless communication with the ground control station 7;
the power arm includes a rotor 108 and a tilting mechanism 115; the rotor is driven by a drive motor 107 at a motor mount 105 supported by the tilting mechanism; the tilting mechanism is connected with the controller unit; the controller unit adjusts rotor position and rotor orientation via a tilt mechanism to adjust the flight of the aircraft.
The number of the power arms is four.
The tilting mechanism is connected with a servo motor 106 at the machine body through the initial end of a supporting arm 103, and is hinged with the middle part of a motor fixing seat through the tail end of the supporting arm; the servo motor drives the supporting arm to rotate so as to enable the rotor wing to tilt; the support arm is provided with a fixing block 111 for fixing the steering engine 109; the steering engine is connected with the lower end of the motor fixing seat through the connecting rod mechanism 104, and the steering engine drives the motor fixing seat to swing through the connecting rod mechanism so that the rotor wing generates yaw force.
The connecting rod mechanism is a parallelogram mechanism; the tilt angle α of the rotor can be adjusted between 0 ° and 90 °, and the tilt angle α can be 0 ° but cannot be 90 °.
The fuselage also includes landing gear 101 and electrical equipment bay 102; a gravity sensor 112, a gyroscope 113, a remote sensing 114 and a plurality of batteries 116 are arranged in the electrical equipment cabin; the output shaft of the servo motor and the contact position of the rotor support arm and the electrical equipment bay 102 are connected by bearings 110.
The sensor unit 2 is responsible for the connection work between the sensor unit and the controller unit 3, receives the measurement information of the sensor unit and converts the received measurement information into an electric signal according to a certain rule and outputs the electric signal to the controller unit 3; the controller unit 3 performs corresponding processing and sends out control command signals after finishing receiving the information transmitted by the sensor unit 2 and the wireless communication module 5;
the control distributor unit 4 completes distribution work of control command signals of the controller unit 3, sequentially distributes the command information received by the controller unit 3 to four servo motors 106, four driving motors 107 and four steering engines 109 of the aircraft, and realizes adjustment of flight states;
the vehicle dynamics system module 5 comprises an actuator unit 501 and an inductor unit 502, wherein the inductor unit 502 comprises a gravity sensing module 503, a perturbation sensing module 504, a gravity sensor 111, a gyroscope 113 and a remote sensing 114, executes instructions for controlling the distributor unit 4, and performs real-time gravity and perturbation external force measurement through the inductor unit 502 and feeds back to the controller unit 3 to form a closed loop system;
the wireless communication module 6 comprises a transmitting module 601 and a receiving module 602, the working modes of the transmitting module 601 and the receiving module 602 can be converted simultaneously and do not interfere with each other, the transmitting module 602 is connected with the controller unit 3, and the receiving module 602 is in communication with the ground control station 7 in a wireless mode so that the ground control station 7 can execute azimuth instruction operation on the aircraft.
The sensor unit 2 is an IMU consisting of SBG system and IG-500, which has an embedded processor for outputting filtered attitude and position data; the controller unit 3 is based on a PD controller with 3 SISDs, which processes the chip set maximum criteria of Digilent of the board for reading the signals; the controller unit 2 reads serial information from the sensor unit 2 output signal at MAX32 and instructs the servo motor 106, the driving motor 107 and the steering engine 109 by sending PWM signals;
the receiving module 602 in the wireless communication module 6 communicates with the ground control station 7 through a Spektrum satellite receiver and a DX6i RC transmitter.
The absolute rotating speed of the output shaft of the driving motor is the sum of the angular speed of the aircraft in three directions of XYZ and the vector sum of the rotating speeds of the steering engine and the servo motor, the physical characteristics of four supporting arms of the aircraft are consistent, one supporting arm is taken for carrying out stress analysis, and the absolute rotating speed of the output shaft of the driving motor is determined
Angular acceleration of
Intermediate vectorA unit vector representing a reference frame j, represented by a reference frame i; p, q and r represent the angular velocities of the aircraft in three directions of the reference coordinate system XYZ, respectively; η and γ represent tilt angles of the biaxial during servo motion; η 'γ' represents the tilt angular velocity of the biaxial during servo motion; omega is expressed as the rotational speed of the drive motor;
using Euler's equation 3 M＝ 3 I 3 α+ 3 ω× 3 I 3 Omega (formula 3), calculating and obtaining torque of output shaft of driving motor 3 M, then utilize rotation matrix R 3to1 Calculating the moment of the aircraft itself in the reference system 1 M Gvro The calculation formula is as follows:
the thrust torque of the connecting rod structure is as follows:
wherein L is the length of the parallelogram mechanism; h is the height of the parallelogram mechanism;
the thrust and torque coefficients of the rotor are defined as:
T＝ρA(ΩR) 2 C T (6) And q=ρa (Ω R) 2 RC Q (equation 7);
c in the formula T And C o Is of a connecting rod structure and thrust andtorque coefficient.
The total thrust moment of the aircraft is:
the absolute stress condition of the aircraft in six degrees of freedom is as follows:
wherein X, Y and Z are expressed as absolute forces of the aircraft on the XYZ axes, respectively; l, M and N are expressed as absolute moment of action of the aircraft on XYZ axes, respectively; u, v and w are expressed as the angular velocity of the aircraft as a whole in XYZ axes, respectively; when the aircraft is under comprehensive stress, the absolute angular rates of the aircraft in three directions of XYZ coordinate axes are as follows:
where Φ ', θ ' and ψ ' are expressed as the absolute angular velocities of the aircraft in three directions of the XYZ coordinate axes, respectively; phi is the included angle between the central axis of the aircraft body and the Z axis of the longitudinal axis; θ is the angle between the projection line of the central axis of the aircraft fuselage on the XY plane and the X axis.
Examples:
when the aircraft changes the flight attitude, the controller unit controls the tilting mechanism, the servo motor drives the supporting arm to rotate so as to enable the rotor wing to vertically tilt in the direction perpendicular to the supporting arm, the steering engine drives the motor fixing seat to swing through the connecting rod mechanism, so that the rotor wing tilts inwards or outwards at the plane of the aircraft body, the lift output direction of the aircraft is changed, and the flight attitude of the aircraft is changed.

Claims (7)

1. A dual-axis tiltrotor aircraft, characterized by: the aircraft body (1) comprises a plurality of power arms uniformly arranged at the periphery of the aircraft body, and a sensor unit (2), a controller unit (3), a control distributor unit (4), a vehicle dynamics system module (5) and a wireless communication module (6) which are arranged at the aircraft body; the wireless communication module can establish wireless communication with a ground control station (7);
the power arm comprises a rotor (108) and a tilting mechanism (115); the rotor is driven by a driving motor (107) at a motor fixing seat (105) supported by the tilting mechanism; the tilting mechanism is connected with the controller unit; the controller unit adjusts the rotor position and rotor orientation via a tilt mechanism to adjust the flight of the aircraft;
the tilting mechanism is connected with a servo motor (106) at the machine body through the initial end of a supporting arm (103), and is hinged with the middle part of a motor fixing seat through the tail end of the supporting arm; the servo motor drives the supporting arm to rotate so as to enable the rotor wing to tilt; a fixed block (111) for fixing the steering engine (109) is arranged at the support arm; the steering engine is connected with the lower end of the motor fixing seat through a connecting rod mechanism (104), and the steering engine drives the motor fixing seat to swing through the connecting rod mechanism so as to enable the rotor wing to generate yaw force;
the absolute rotating speed of the output shaft of the driving motor is the sum of the angular speed of the aircraft in three directions of XYZ and the vector sum of the rotating speeds of the steering engine and the servo motor, the physical characteristics of four supporting arms of the aircraft are consistent, one supporting arm is taken for carrying out stress analysis, and the absolute rotating speed of the output shaft of the driving motor is determined
Angular acceleration of
Intermediate vector i i j i j j i k j A unit vector representing a reference frame j, represented by a reference frame i; p, q and r represent the angular velocities of the aircraft in three directions of the reference coordinate system XYZ, respectively; η and γ represent tilt angles of the biaxial during servo motion; η 'γ' represents the tilt of the biaxial during servo motionAngular velocity; omega is expressed as the rotational speed of the drive motor;
using Euler's equation 3 M＝ 3 I 3 α+ 3 ω× 3 I 3 Omega (formula 3), calculating and obtaining torque of output shaft of driving motor 3 M, then utilize rotation matrix R 3to1 Calculating the moment of the aircraft itself in the reference system 1 M Gyro The calculation formula is as follows:
the thrust torque of the connecting rod structure is as follows:
wherein L is the length of the parallelogram mechanism; h is the height of the parallelogram mechanism;
the thrust and torque coefficients of the rotor are defined as:
T＝ρA(ΩR) 2 C T (6) And q=ρa (Ω R) 2 RC Q (equation 7);
c in the formula T And C Q The thrust and torque coefficients are of a connecting rod structure;
the total thrust moment of the aircraft is:
2. the dual-axis tiltrotor aircraft according to claim 1, wherein: the number of the power arms is four.
3. The dual-axis tiltrotor aircraft according to claim 1, wherein: the connecting rod mechanism is a parallelogram mechanism; the tilt angle α of the rotor can be adjusted between 0 ° and 90 °, and the tilt angle α can be 0 ° but cannot be 90 °.
4. The dual-axis tiltrotor aircraft according to claim 1, wherein: the fuselage further comprises landing gear (101) and electrical equipment bay (102); a gravity sensor (112), a gyroscope (113), remote sensing (114) and a plurality of batteries (116) are arranged in the electric equipment cabin; the output shaft of the servo motor and the contact position of the rotor support arm and the electrical equipment cabin (102) are connected through a bearing (110).
5. The dual-axis tiltrotor aircraft according to claim 4, wherein: the sensor unit (2) is responsible for the connection work between the sensor unit and the controller unit (3), receives the measurement information of the sensor unit and converts the received measurement information into an electric signal according to a certain rule to be output to the controller unit (3); the controller unit (3) performs corresponding processing and sends out control command signals after finishing receiving the information transmitted by the sensor unit (2) and the wireless communication module (6);
the control distributor unit (4) completes distribution work of control command signals of the controller unit (3), sequentially distributes command information received by the controller unit (3) to four servo motors (106), four driving motors (107) and four steering engines (109) of the aircraft, and realizes adjustment of flight states;
the vehicle dynamics system module (5) comprises an actuator unit (501) and an inductor unit (502), the inductor unit (502) comprises a gravity sensing module (503), a perturbation sensing module (504), a gravity sensor (112), a gyroscope (113) and remote sensing (114), the instruction for controlling the distributor unit (4) is executed, and the vehicle dynamics system module performs real-time gravity and perturbation external force measurement through the inductor unit (502) and feeds back to the controller unit (3) to form a closed loop system;
the wireless communication module (6) comprises a transmitting module (601) and a receiving module (602), the working modes of the transmitting module (601) and the receiving module (602) can be converted and do not interfere with each other, the transmitting module (601) is connected with the controller unit (3), and the receiving module (602) is in communication with the ground control station (7) in a wireless mode so that the ground control station (7) can execute azimuth instruction operation on the aircraft.
6. The dual-axis tiltrotor aircraft according to claim 5, wherein: the sensor unit (2) is an IMU consisting of an SBG system and an IG-500, which unit has an embedded processor for outputting filtered attitude and position data; the controller unit (3) is based on a PD controller with 3 SISDs, which processes the chip set maximum criterion of Digilent of the board for reading the signals; the controller unit (3) reads serial information from the output signal of the sensor unit (2) by MAX32 and instructs the servo motor (106), the driving motor (107) and the steering engine (109) by sending PWM signals;
the receiving module (602) in the wireless communication module (6) communicates with the ground control station (7) through a Spekdrum satellite receiver and a DX6i RC transmitter.
7. The dual-axis tiltrotor aircraft according to claim 1, wherein: the absolute stress condition of the aircraft in six degrees of freedom is as follows:
wherein X, Y and Z are expressed as absolute forces of the aircraft on the XYZ axes, respectively; l, M and N are expressed as absolute moment of action of the aircraft on XYZ axes, respectively; u, v and w are expressed as the angular velocity of the aircraft as a whole in XYZ axes, respectively; when the aircraft is under comprehensive stress, the absolute angular rates of the aircraft in three directions of XYZ coordinate axes are as follows:
where Φ ', θ ' and ψ ' are expressed as the absolute angular velocities of the aircraft in three directions of the XYZ coordinate axes, respectively; phi is the included angle between the central axis of the aircraft body and the Z axis of the longitudinal axis; θ is the angle between the projection line of the central axis of the aircraft fuselage on the XY plane and the X axis.
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Families Citing this family (3)

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JP7085892B2 (en) * 2018-05-14 2022-06-17 川崎重工業株式会社 Aircraft and how to control the aircraft
CN110920909A (en) * 2019-11-22 2020-03-27 南京航空航天大学 Flight control method of double-engine-driven variable-pitch multi-rotor aircraft
CN111258324B (en) * 2020-01-19 2023-08-18 沈阳无距科技有限公司 Multi-rotor unmanned aerial vehicle control method and device, multi-rotor unmanned aerial vehicle and storage medium

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN105346695A (en) * 2015-11-20 2016-02-24 浙江大学 Quad-rotor mini-type underwater vehicle
CN105882959A (en) * 2015-02-13 2016-08-24 空中客车防卫和太空有限责任公司 Aircraft capable of vertical takeoff
EP3170746A1 (en) * 2015-11-19 2017-05-24 HW Aviation AG Method for controlling a multicopter and device for carrying out the method
CN106892094A (en) * 2017-01-22 2017-06-27 南京航空航天大学 A kind of individually controllable four rotor unmanned aircraft of space six degree of freedom and its control method
CN109562824A (en) * 2016-11-08 2019-04-02 深圳市大疆创新科技有限公司 Motor and unmanned plane with the motor
CN210027899U (en) * 2019-05-09 2020-02-07 福州大学 Four rotor crafts of biax slope

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN105882959A (en) * 2015-02-13 2016-08-24 空中客车防卫和太空有限责任公司 Aircraft capable of vertical takeoff
EP3170746A1 (en) * 2015-11-19 2017-05-24 HW Aviation AG Method for controlling a multicopter and device for carrying out the method
CN105346695A (en) * 2015-11-20 2016-02-24 浙江大学 Quad-rotor mini-type underwater vehicle
CN109562824A (en) * 2016-11-08 2019-04-02 深圳市大疆创新科技有限公司 Motor and unmanned plane with the motor
CN106892094A (en) * 2017-01-22 2017-06-27 南京航空航天大学 A kind of individually controllable four rotor unmanned aircraft of space six degree of freedom and its control method
CN210027899U (en) * 2019-05-09 2020-02-07 福州大学 Four rotor crafts of biax slope

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