CN110015415A - A kind of bi-axial tilt quadrotor - Google Patents
A kind of bi-axial tilt quadrotor Download PDFInfo
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- CN110015415A CN110015415A CN201910382221.2A CN201910382221A CN110015415A CN 110015415 A CN110015415 A CN 110015415A CN 201910382221 A CN201910382221 A CN 201910382221A CN 110015415 A CN110015415 A CN 110015415A
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- 230000005484 gravity Effects 0.000 claims description 12
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- 238000006243 chemical reaction Methods 0.000 claims description 5
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- 230000033228 biological regulation Effects 0.000 abstract description 6
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- 230000037396 body weight Effects 0.000 description 1
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- 210000003746 feather Anatomy 0.000 description 1
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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
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- 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
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64U—UNMANNED AERIAL VEHICLES [UAV]; EQUIPMENT THEREFOR
- B64U10/00—Type of UAV
- B64U10/10—Rotorcrafts
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
- G05D1/0816—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability
- G05D1/0825—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft to ensure stability using mathematical models
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- 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
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- H—ELECTRICITY
- H04—ELECTRIC COMMUNICATION TECHNIQUE
- H04W—WIRELESS COMMUNICATION NETWORKS
- H04W76/00—Connection management
- H04W76/10—Connection setup
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Abstract
The present invention proposes a kind of bi-axial tilt quadrotor, fuselage includes uniformly multiple power arms set on fuselage week at, and sensor unit, controller unit, control dispenser unit, vehicle dynamics system module, wireless communication module at fuselage；Wireless communication module can establish wireless telecommunications with ground control station；The power arm includes rotor and leaning device；Rotor is with the driving motor driving of the motor fixing seat supported by leaning device；The leaning device is connected with controller unit；The controller unit adjusts rotor position and rotor towards the flight to adjust aircraft through leaning device；The present invention can instruct the revolving speed and rotor tilt angle of four rotors of automatic adjustment according to input, utilize the stability of closed-loop control system, the state of flight of itself is fed back, the work of self flight regulation is automatically performed, achievees the purpose that flight pitching, rolling and yaw；Both the high efficiency and high reliability of flight regulation be ensure that.
Description
Technical field
The present invention relates to small-sized unmanned aircraft technical field, especially a kind of bi-axial tilt quadrotor.
Background technique
The configuration of bi-axial tilt formula rotor refers to rotation by support arm and motor fixing seat, make the rotational plane of rotor with
The rotor configuration that forms an angle with body planar shaped, since the aircraft has the characteristics that bi-axial tilt, then the aircraft
Rotor can be completed to be easy to carry out pitching, rolling and yawing rotation, realize aircraft around the inclined effect of driving motor surrounding
The purpose flexibly flown.The support arm and motor fixing seat of most of traditional quadrotor are all fixedly mounted at body
On, four rotors of aircraft all make rotating motion in the same plane, and rotor can only pass through the revolving speed of change driving motor
Realize pitching, rolling and yawing rotation, need four variable speed drive motors and regulation and control system, and strain function it is poor and
Not stable enough in the process, if aircraft needs to complete above 3 kinds of movements in time, first method is to need to add at head
Driver and tail pipe is installed at dress driver or tail additional to require to reach flight.However, this method increasing with equipment,
Body weight needs more rotors or lengthens propeller to maintain the hovering campaign of aircraft, pass through this also along with increase
Although the method that kind increases rotor enhances the reliability of aircraft, but its flexibility is difficult to be guaranteed.Second method
It is the method using feather propeller, this method is according to the weight for trying holding aircraft, to maintain the height of aircraft
Flexibility, but this method needs to implement complex, failure using more complicated propeller and servo electrical machinery system
Rate is high, and may be with personal damage in the event of failure.The method of the above-mentioned flexible sporting flying of realization multi-rotor aerocraft or
Lift and weight is relatively low or structure is complicated, low efficiency, this greatly constrains it in the extensive use of every field.
Summary of the invention
The present invention proposes a kind of bi-axial tilt quadrotor, and four rotors of automatic adjustment can be instructed according to input
Revolving speed and rotor tilt angle, remote control can be carried out, and closed loop can be utilized according to different flight environment of vehicle conditions
The stability of control system is fed back the state of flight of itself, is automatically performed the work of self flight regulation, can be led to
Tilting rotor is crossed to achieve the purpose that flight pitching, rolling and yaw；Both the high efficiency and high reliability of flight regulation be ensure that,
The size that aircraft can be reduced again, to improve flexibility.
The present invention uses following technical scheme.
A kind of bi-axial tilt quadrotor, the fuselage (1) of the aircraft include uniformly being set to fuselage week at
Multiple power arms, and the sensor unit (2) at fuselage, controller unit (3), control dispenser unit (4), vehicle
Dynamic system module (5), wireless communication module (6)；The wireless communication module can establish channel radio with ground control station (7)
News；
The power arm includes rotor (108) and leaning device (115)；Motor of the rotor to be supported by leaning device
Driving motor (107) driving at fixing seat (105)；The leaning device is connected with controller unit；The controller unit
Rotor position and rotor are adjusted towards the flight to adjust aircraft through leaning device.
The power arm quantity is four.
The leaning device is connected with support arm (103) beginning with servo motor at fuselage (106), with support arm end with
The middle part of motor fixing seat is hinged；The servo motor driving support arm rotation is so that rotor verts；It is equipped at the support arm
For fixing the fixed block (111) of steering engine (109)；The steering engine is connected through link mechanism (104) with the lower end of motor fixing seat,
The steering engine makes rotor generate yaw forces through the swing of link mechanism driving motor fixing seat.
The link mechanism is parallelogram mechanism；The inclination angle alpha of the rotor can carry out between 0 ° and 90 °
Adjustment, inclination angle alpha can be 0 ° but cannot be 90 °.
The fuselage further includes undercarriage (101) and electrical compartment (102)；Gravity-feed tank is equipped in the electrical compartment
Answer device (112), gyroscope (113), remote sensing (114) and several batteries (116)；The output shaft and rotor branch of the servo motor
Brace and the contact position of electrical compartment (102) are attached by bearing (110).
The sensor unit (2) is responsible for the connection work between sensor unit and controller unit (3), receives sensing
The metrical information received is simultaneously for conversion into electric signal according to certain rules and exported to controller unit by the metrical information of device unit
(3)；The controller unit (3) completes to carry out after the transmission information of receiving sensor unit (2) and wireless communication module (5) pair
The processing answered simultaneously issues control command signal；
Control dispenser unit (4) completes sharing out the work for the control command signal of controller unit (3), will receive
Command information to controller unit (3) carries out orderly four servo motors (106) for distributing to aircraft, four driving motors
(107) and four steering engines (109), the adjustment of state of flight is realized；
The vehicle dynamics system module (5) includes actuator unit (501) and sensor unit (502), inductor
Unit (502) include gravity sensing module (503), perturbation induction module (504), gravity sensor (111), gyroscope (113),
Remote sensing (114), execute control dispenser unit (4) instruction, vehicle dynamics system module by sensor unit (502) into
The real-time gravity of row and perturbation external force measure and feed back to controller unit (3) to form closed-loop system；
The wireless communication module (6) includes transmitting module (601) and receiving module (602), the transmitting module (601)
It can conversion and non-interference, transmitting module (602) and controller unit mutually simultaneously with the operating mode of receiving module (602)
(3) be connected, between receiving module (602) and ground control station (7) wirelessly communication contact so that the ground control station
(7) orientation is executed to aircraft and instructs operation.
The sensor unit (2) is an IMU by SBG system and IG-500, this unit has an embedded place
Device is managed, for exporting the posture and position data of filtering；The controller unit (3) is controlled based on a PD for having 3 SISD
Device processed, the chipset maximal criterion for processing the Digilent of plate are used to read signal；Controller unit (2) is from sensor unit
(2) serial information is read with MAX32 in output signal, instructs servo motor (106), driving motor by issuing pwm signal
(107) and steering engine (109):
It is defended between receiving module (602) and ground control station (7) in the wireless communication module (6) by Spektrum
Star receiver and DX6i RC transmitter are communicated.
The absolute revolving speed of the output shaft of the driving motor be angular speed of the aircraft on tri- directions XYZ, steering engine and
The physical features of the sum of the vector of servo motor revolving speed, four support arms of aircraft are consistent, and a support arm is taken to carry out stress point
Analysis, the absolute revolving speed of the output shaft of driving motor
Angular acceleration is
Vector in formulaThe unit vector for indicating reference system j, is indicated with reference system i；P, q and r respectively indicates flight
Angular speed of the device on tri- directions reference frame XYZ；η and γ indicates the tilt angle of twin shaft when servo motion；η ' γ ' table
The angular speed that verts of twin shaft when showing servo motion；Ω is expressed as the revolving speed of driving motor；
Utilize Eulerian equation3M=3I3α+3ω×3I3ω (formula 3) calculates the torque for acquiring driving motor output shaft3M,
Then spin matrix R is utilized3to1Calculate the torque for acquiring the aircraft itself in referential1MGvro, its calculation formula is:
The thrust torques of link mechanism are as follows:
L is the length of parallelogram mechanism in formula；H is the height of parallelogram mechanism；
The thrust and torque coefficient of rotor is defined as:
T=ρ A (Ω R)2CT(6) and Q=ρ A (Ω R)2RCQ(formula 7)；
C in formulaTAnd CoFor link mechanism thrust and torque coefficient.
Then aircraft gross thrust torque are as follows:
Aircraft absolute stress condition on six-freedom degree are as follows:
X, Y and Z are expressed as absolute active force of the aircraft on XYZ axis in formula；L, M and N is expressed as flying
Absolute opplied moment of the device on XYZ axis；U, v and w is expressed as the whole angular speed on XYZ axis of aircraft；Aircraft
Absolute angular speed in comprehensive stress condition, on three directions of XYZ coordinate axis are as follows:
Φ ', θ ' and Ψ ' are expressed as absolute angular speed of the aircraft on three directions of XYZ coordinate axis in formula；Φ
For the angle of aircraft fuselage central axis and longitudinal axis Z axis；θ is projection line and X-axis of the aircraft fuselage central axis on X/Y plane
Angle.
A kind of novel biaxial that the present invention provides tilts quadrotor, can use support arm and motor fixing seat
Rotation makes rotor vert, and then achievees the purpose that flight free strain by tilting rotor, and can be according to different flight environment of vehicle
Condition is fed back the state of flight of itself using the stability of closed-loop control system, is automatically performed self flight regulation
Work reduces fly to a certain extent since the optimization of the housing construction of aircraft reduces the utilization of some dynamical system equipment
The scale size of row device becomes more flexible flight.The flight for realizing quadrotor is intelligent, and manipulator only needs to input
Instruction can be obtained by its requirement, also avoid the operation error bring of manipulator dangerous, the novel biaxial that the present invention provides inclines
Convenient, the safe and intelligent outstanding advantages of oblique quadrotor have a broad prospect of the use in dual-use upper tool.
Detailed description of the invention
The present invention is described in more detail with reference to the accompanying drawings and detailed description:
Attached drawing 1 is operation control system schematic diagram of the invention；
Attached drawing 2 is the vertical view of aircraft of the present invention to schematic diagram；
Attached drawing 3 is leaning device schematic diagram of the invention；
Attached drawing 4 is vehicle dynamics system module diagram of the invention；
Attached drawing 5 is information interchange schematic diagram between controller unit and ground control station of the invention；
Attached drawing 6 is the communication schematic diagram between controller unit and each actuator unit and sensor unit of the invention；
(one of rotor is in vert reference frame schematic diagram when attached drawing 7 is aircraft flight of the present invention
State).
In attached drawing: 1- fuselage；2- sensor unit；3- controller unit；4- controls dispenser unit；5- dynamics of vehicle
System module；6- wireless communication module；7- ground control station；
101- undercarriage；102- electrical compartment；103- support arm；104- link mechanism；105- motor fixing seat；106-
Servo motor；107- driving motor；108- rotor；109- steering engine；110- bearing；111- fixed block；112- gravity sensor；
113- gyroscope；114- remote sensing；115- leaning device；116- battery；
501- actuator unit；502- sensor unit；503- gravity sensing module；504- perturbation induction module；
601- transmitting module；602- receiving module.
Specific embodiment
As shown in figs. 1-7, a kind of bi-axial tilt quadrotor, the fuselage 1 of the aircraft include uniformly being set to machine
Body week along place multiple power arms, and sensor unit 2 at fuselage, controller unit 3, control dispenser unit 4,
Vehicle dynamics system module 5, wireless communication module 6；The wireless communication module can establish channel radio with ground control station 7
News；
The power arm includes rotor 108 and leaning device 115；The rotor is fixed with the motor supported by leaning device
Driving motor 107 at seat 105 drives；The leaning device is connected with controller unit；The controller unit is through dumper
Structure adjusts rotor position and rotor towards the flight to adjust aircraft.
The power arm quantity is four.
The leaning device is connected with 103 beginning of support arm with servo motor 106 at fuselage, with support arm end and motor
The middle part of fixing seat is hinged；The servo motor driving support arm rotation is so that rotor verts；It is equipped with and is used at the support arm
The fixed block 111 of fixed steering engine 109；The steering engine is connected through link mechanism 104 with the lower end of motor fixing seat, the steering engine warp
The swing of link mechanism driving motor fixing seat makes rotor generate yaw forces.
The link mechanism is parallelogram mechanism；The inclination angle alpha of the rotor can carry out between 0 ° and 90 °
Adjustment, inclination angle alpha can be 0 ° but cannot be 90 °.
The fuselage further includes undercarriage 101 and electrical compartment 102；Gravity sensor is equipped in the electrical compartment
112, gyroscope 113, remote sensing 114 and several batteries 116；The output shaft and flight support arm of the servo motor with electrically set
The contact position in standby cabin 102 is attached by bearing 110.
The sensor unit 2 is responsible for the work that contacts between sensor unit and controller unit 3, receiving sensor list
The metrical information received is simultaneously for conversion into electric signal according to certain rules and exported to controller unit 3 by the metrical information of member；Institute
Corresponding processing is carried out concurrently after stating the transmission information of the completion receiving sensor unit 2 of controller unit 3 and wireless communication module 5
Control command signal out；
The control dispenser unit 4 completes sharing out the work for the control command signal of controller unit 3, will receive control
The command information of device unit 3 processed carries out four servo motors, 106, four driving motors 107 and four for orderly distributing to aircraft
A steering engine 109, realizes the adjustment of state of flight；
The vehicle dynamics system module 5 includes actuator unit 501 and sensor unit 502, sensor unit 502
Including gravity sensing module 503, perturbation induction module 504, gravity sensor 111, gyroscope 113, remote sensing 114, control is executed
The instruction of dispenser unit 4, vehicle dynamics system module carries out real-time gravity by sensor unit 502 and perturbation external force is surveyed
Determine and feeds back to controller unit 3 to form closed-loop system；
The wireless communication module 6 includes transmitting module 601 and receiving module 602, the transmitting module 601 and reception mould
The operating mode of block 602 can simultaneously mutually conversion and it is non-interference, transmitting module 602 is connected with controller unit 3, reception mould
Between block 602 and ground control station 7 wirelessly communication contact so that the ground control station 7 to aircraft execute orientation refer to
Enable operation.
The sensor unit 2 is an IMU by SBG system and IG-500, this unit has an embedded processing
Device, for exporting the posture and position data of filtering；The controller unit 3 is the PD control device for having 3 SISD based on one,
The chipset maximal criterion of its Digilent for processing plate is used to read signal；Controller unit 2 is exported from sensor unit 2 to be believed
Serial information is read with MAX32 in number, instructs servo motor 106, driving motor 107 and steering engine 109 by issuing pwm signal；
Pass through Spektrum satellite reception between receiving module 602 and ground control station 7 in the wireless communication module 6
Device and DX6i RC transmitter are communicated.
The absolute revolving speed of the output shaft of the driving motor be angular speed of the aircraft on tri- directions XYZ, steering engine and
The physical features of the sum of the vector of servo motor revolving speed, four support arms of aircraft are consistent, and a support arm is taken to carry out stress point
Analysis, the absolute revolving speed of the output shaft of driving motor
Angular acceleration is
Vector in formulaThe unit vector for indicating reference system j, is indicated with reference system i；P, q and r respectively indicates flight
Angular speed of the device on tri- directions reference frame XYZ；η and γ indicates the tilt angle of twin shaft when servo motion；η ' γ ' table
The angular speed that verts of twin shaft when showing servo motion；Ω is expressed as the revolving speed of driving motor；
Utilize Eulerian equation3M=3I3α+3ω×3I3ω (formula 3) calculates the torque for acquiring driving motor output shaft3M,
Then spin matrix R is utilized3to1Calculate the torque for acquiring the aircraft itself in referential1MGvro, its calculation formula is:
The thrust torques of link mechanism are as follows:
L is the length of parallelogram mechanism in formula；H is the height of parallelogram mechanism；
The thrust and torque coefficient of rotor is defined as:
T=ρ A (Ω R)2CT(6) and Q=ρ A (Ω R)2RCQ(formula 7)；
C in formulaTAnd CoFor link mechanism thrust and torque coefficient.
Then aircraft gross thrust torque are as follows:
Aircraft absolute stress condition on six-freedom degree are as follows:
X, Y and Z are expressed as absolute active force of the aircraft on XYZ axis in formula；L, M and N is expressed as flying
Absolute opplied moment of the device on XYZ axis；U, v and w is expressed as the whole angular speed on XYZ axis of aircraft；Aircraft
Absolute angular speed in comprehensive stress condition, on three directions of XYZ coordinate axis are as follows:
Φ ', θ ' and Ψ ' are expressed as absolute angular speed of the aircraft on three directions of XYZ coordinate axis in formula；Φ
For the angle of aircraft fuselage central axis and longitudinal axis Z axis；θ is projection line and X-axis of the aircraft fuselage central axis on X/Y plane
Angle.
Embodiment:
For aircraft in turning flight posture, controller unit controls leaning device, and servo motor drives support arm rotation
So that rotor vertically verts by the direction perpendicular to support arm, steering engine is swung through link mechanism driving motor fixing seat, makes rotor
It is outwardly or inwardly verting at plane where fuselage, the lift outbound course of change of flight device sends out the flight attitude of aircraft
It is raw to change.
Claims (9)
1. a kind of bi-axial tilt quadrotor, it is characterised in that: the fuselage (1) of the aircraft includes uniformly being set to fuselage
All multiple power arms at, and sensor unit (2), controller unit (3), control dispenser unit at fuselage
(4), vehicle dynamics system module (5), wireless communication module (6)；The wireless communication module can be built with ground control station (7)
Vertical wireless telecommunications；
The power arm includes rotor (108) and leaning device (115)；The rotor is fixed with the motor supported by leaning device
Driving motor (107) driving at seat (105)；The leaning device is connected with controller unit；The controller unit is through inclining
Oblique mechanism adjusts rotor position and rotor towards the flight to adjust aircraft.
2. a kind of bi-axial tilt quadrotor according to claim 1, it is characterised in that: the power arm quantity is
Four.
3. a kind of bi-axial tilt quadrotor according to claim 1, it is characterised in that: the leaning device is to prop up
Brace (103) beginning is connected with servo motor at fuselage (106), is hinged with support arm end and the middle part of motor fixing seat；Institute
Servo motor driving support arm rotation is stated so that rotor verts；The fixation for fixing steering engine (109) is equipped at the support arm
Block (111)；The steering engine is connected through link mechanism (104) with the lower end of motor fixing seat, and the steering engine drives through link mechanism
Motor fixing seat swing makes rotor generate yaw forces.
4. a kind of bi-axial tilt quadrotor according to claim 3, it is characterised in that: the link mechanism is flat
Row quadrangular mechanism；The inclination angle alpha of the rotor can be adjusted between 0 ° and 90 °, inclination angle alpha can for 0 ° but
It cannot be 90 °.
5. a kind of bi-axial tilt quadrotor according to claim 3, it is characterised in that: the fuselage further included
Fall frame (101) and electrical compartment (102)；It is equipped in the electrical compartment gravity sensor (112), gyroscope (113), distant
Feel (114) and several batteries (116)；The output shaft and flight support arm of the servo motor and electrical compartment (102)
Contact position is attached by bearing (110).
6. a kind of bi-axial tilt quadrotor according to claim 5, it is characterised in that: the sensor unit
(2) the connection work being responsible between sensor unit and controller unit (3), the metrical information of receiving sensor unit will simultaneously connect
The metrical information received, which is for conversion into electric signal according to certain rules and exports, gives controller unit (3)；The controller unit (3) is complete
At carried out after the transmission information of receiving sensor unit (2) and wireless communication module (5) it is corresponding processing and issue control command
Signal；
Control dispenser unit (4) completes sharing out the work for the control command signal of controller unit (3), will receive control
The command information of device unit (3) processed carries out orderly distributing to four servo motors (106) of aircraft, four driving motors
(107) and four steering engines (109), the adjustment of state of flight is realized；
The vehicle dynamics system module (5) includes actuator unit (501) and sensor unit (502), sensor unit
It (502) include gravity sensing module (503), perturbation induction module (504), gravity sensor (111), gyroscope (113), remote sensing
(114), the instruction of control dispenser unit (4) is executed, vehicle dynamics system module carries out real by sensor unit (502)
Shi Chongli and perturbation external force measure and feed back to controller unit (3) to form closed-loop system；
The wireless communication module (6) includes transmitting module (601) and receiving module (602), the transmitting module (601) and is connect
The operating mode for receiving module (602) can while be converted mutually and non-interference, transmitting module (602) and controller unit (3)
Be connected, between receiving module (602) and ground control station (7) wirelessly communication contact so that the ground control station (7)
Orientation is executed to aircraft and instructs operation.
7. a kind of bi-axial tilt quadrotor according to claim 6, it is characterised in that: the sensor unit
(2) it is an IMU by SBG system and IG-500, this unit has an embeded processor, for exporting the posture of filtering
And position data；The controller unit (3) is the PD control device for having 3 SISD based on one, processes the Digilent of plate
Chipset maximal criterion be used to read signal；Controller unit (2) is read from sensor unit (2) output signal with MAX32
Serial information is taken, by issuing pwm signal instruction servo motor (106), driving motor (107) and steering engine (109)；
It is connect between receiving module (602) and ground control station (7) in the wireless communication module (6) by Spektrum satellite
It receives device and DX6i RC transmitter is communicated.
8. a kind of bi-axial tilt quadrotor according to claim 3, it is characterised in that: the driving motor it is defeated
The absolute revolving speed of shaft is angular speed, steering engine and the vector of servo motor revolving speed the sum of of the aircraft on tri- directions XYZ, is flown
The physical features of four support arms of row device are consistent, take a support arm carry out force analysis, the output shaft of driving motor it is exhausted
To revolving speed
Angular acceleration is
Vector in formulaThe unit vector for indicating reference system j, is indicated with reference system i；P, q and r respectively indicate aircraft
Angular speed on tri- directions reference frame XYZ；η and γ indicates the tilt angle of twin shaft when servo motion；η ' γ ' is indicated
The angular speed that verts of twin shaft when servo motion；Ω is expressed as the revolving speed of driving motor；
Utilize Eulerian equation3M=3I3α+3ω×3I3ω (formula 3) calculates the torque for acquiring driving motor output shaft3M, then
Utilize spin matrix R3to1Calculate the torque for acquiring the aircraft itself in referential1MGyro, its calculation formula is:
The thrust torques of link mechanism are as follows:
L is the length of parallelogram mechanism in formula；H is the height of parallelogram mechanism；
The thrust and torque coefficient of rotor is defined as:
T=ρ A (Ω R)2CT (6) and Q=ρ A (Ω R)2RCQ(formula 7)；
C in formulaTAnd CQFor link mechanism thrust and torque coefficient.
Then aircraft gross thrust torque are as follows:
9. a kind of bi-axial tilt quadrotor according to claim 8, it is characterised in that: aircraft is in six freedom
Absolute stress condition on degree are as follows:
X, Y and Z are expressed as absolute active force of the aircraft on XYZ axis in formula；L, M and N is expressed as aircraft and exists
Absolute opplied moment on XYZ axis；U, v and w is expressed as the whole angular speed on XYZ axis of aircraft；Aircraft is comprehensive
When closing stress condition, the absolute angular speed on three directions of XYZ coordinate axis are as follows:
Φ ', θ ' and Ψ ' are expressed as absolute angular speed of the aircraft on three directions of XYZ coordinate axis in formula；Φ is winged
The angle of row device fuselage central axis and longitudinal axis Z axis；θ is the folder of projection line and X-axis of the aircraft fuselage central axis on X/Y plane
Angle.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910382221.2A CN110015415B (en) | 2019-05-09 | 2019-05-09 | Double-shaft tilting four-rotor aircraft |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910382221.2A CN110015415B (en) | 2019-05-09 | 2019-05-09 | Double-shaft tilting four-rotor aircraft |
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| Publication Number | Publication Date |
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| CN110015415A true CN110015415A (en) | 2019-07-16 |
| CN110015415B CN110015415B (en) | 2024-02-09 |
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| CN201910382221.2A Active CN110015415B (en) | 2019-05-09 | 2019-05-09 | Double-shaft tilting four-rotor aircraft |
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Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
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