CN110316368A - A kind of distributed-power tilting rotor wing unmanned aerial vehicle and its control method - Google Patents
A kind of distributed-power tilting rotor wing unmanned aerial vehicle and its control method Download PDFInfo
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- CN110316368A CN110316368A CN201910269423.6A CN201910269423A CN110316368A CN 110316368 A CN110316368 A CN 110316368A CN 201910269423 A CN201910269423 A CN 201910269423A CN 110316368 A CN110316368 A CN 110316368A
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C19/00—Aircraft control not otherwise provided for
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C27/00—Rotorcraft; Rotors peculiar thereto
- B64C27/22—Compound rotorcraft, i.e. aircraft using in flight the features of both aeroplane and rotorcraft
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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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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
- G05B13/00—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion
- G05B13/02—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric
- G05B13/04—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators
- G05B13/042—Adaptive control systems, i.e. systems automatically adjusting themselves to have a performance which is optimum according to some preassigned criterion electric involving the use of models or simulators in which a parameter or coefficient is automatically adjusted to optimise the performance
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Abstract
The invention discloses a kind of distributed-power tilting rotor wing unmanned aerial vehicle and its control methods.The distributed-power tilting rotor wing unmanned aerial vehicle includes the first rotor system being arranged on wing and the second rotor system being arranged at rudder-vator；First rotor system includes 4 tilting rotors, this 4 tilting rotors are arranged in two-by-two on left side wing and right side wing；Second rotor system includes 2 rotors, this 2 rotors are separately positioned on the two sides of V-arrangement wing.The control method of the unmanned plane includes the control strategy of rotor Model control rule, fixed-wing Model control rule and transition mode.The present invention had not only been able to achieve the optimal power matching of dynamical system, but also can provide robust for each state of gyroplane and reliably control.
Description
Technical field
The invention belongs to air vehicle technique fields, in particular to a kind of tilting rotor wing unmanned aerial vehicle.
Background technique
Tilt rotor aircraft can as helicopter VTOL, hovering and as fixed-wing it is quickly solid
Determine wing flight, has the characteristics that high-efficient, speed is fast, voyage is remote.Therefore tiltrotor tool has a broad prospect of the use.
The hot spot that tiltrotor is always studied both at home and abroad faces aerodynamic characteristic complexity, it is difficult to the technical problems such as control:
It takes off vertically and has a long way to go with engine power demand under fixed-wing offline mode, it is difficult to realize that optimal power matches；Meanwhile together
Aerodynamic instability caused by step is verted also is difficult to control gyroplane.
Therefore, this field structure simple tiltrotor aircraft new there is still a need for research, and can be each shape of gyroplane
The control of state provides reliable solution.
Summary of the invention
In order to solve the technical issues of above-mentioned background technique proposes, the present invention is directed to propose a kind of distributed-power verts
Rotor wing unmanned aerial vehicle and its control method.
In order to achieve the above technical purposes, the technical solution of the present invention is as follows:
A kind of distributed-power tilting rotor wing unmanned aerial vehicle, including fuselage, wing and rudder-vator further include being arranged in wing
On the first rotor system and the second rotor system for being arranged at rudder-vator；First rotor system includes 4 rotations of verting
The wing, this 4 tilting rotors are arranged in two-by-two on left side wing and right side wing；Second rotor system includes 2 rotors, this
2 rotors are separately positioned on the two sides of V-arrangement wing.
Further, in the first rotor system the rotary shaft of each tilting rotor and unmanned plane center of gravity overlapping of axles, verting
In the process, the center of gravity approximation of unmanned plane is constant.
Further, each tilting rotor has limiting device, so that the tilt angle of tilting rotor is limited in 0~
Within the scope of 90 °.
Further, the propeller of the rotor in the second rotor system uses variable-pitch propeller.
Based on the control method of above-mentioned distributed-power tilting rotor wing unmanned aerial vehicle, including following control law:
(1) rotor Model control is restrained:
Using two-stage PID control, first according to desired roll anglePitching angle thetad, yaw angle ψdWith actual rolling
AnglePitching angle theta, the margin of error of yaw angle ψ construct first order PID control, i.e. attitude angle closed-loop control is restrained:
In above formula, pd、qd、rdFor expectation angular velocity in roll, expectation rate of pitch and expectation yaw rate；eθ=θd- θ, eψ=ψd-ψ；kθp、kψpIt is roll angle, pitch angle and the scale parameter for yawing angle error；
According to desired angular speed (pd,qd,rd) with the margin of error of actual angular speed (p, q, r), construct second level PID control,
I.e. attitude angular velocity closed-loop control is restrained:
In above formula, Δ δp、Δδq、ΔδrIt is the PID output of three attitude angular velocities；ep=pd- p, eq=qd- q, er=rd-
r；kpp、kqp、krpIt is the scale parameter of three attitude angular velocities；, kpd、kqd、krdIt is the differential parameter of three attitude angular velocities；
(2) fixed-wing Model control is restrained:
Including vertical passage control law and interconnection control law, wherein vertical passage control law:
In above formula, δeControl amount is exported for elevator；Δ H=Hc- H, HcFor altitude instruction, H is actual height；KHpFor height
Spend ring proportional component coefficient, KHiFor height ring integral element coefficient；For the proportional component coefficient of angular speed ring, KωyAngular speed
The differentiation element coefficient of ring；For pitch angle, ωyFor pitch rate；T is the sampling period, and k is sampling period number；
α is the switching coefficient of integral term:
Wherein, Δ H (t) is the Δ H value of current time t, and ε is the difference in height judgment threshold of setting；
Interconnection control law:
When lateral deviation is away from D > M:
δa=KD1·D+Kψ·Δψ+Kωx·ωx
When lateral deviation is away from D≤M:
Wherein, δaControl amount, δ are exported for aileronrControl amount is exported for rudder；ωx、ωzRespectively angular velocity in roll,
Yaw rate；Δ ψ=ψd-ψ；β is the switching coefficient of integral term；KD1、KD2Respectively segment processing when lateral deviation away from ratio system
Number；Kψ、KγFor the proportional component coefficient of angular speed ring；Kωz、KωxThe differentiation element coefficient of angular speed ring；M be setting lateral deviation away from
Threshold value.
Further, it needs to switch in rotor mode and fixed-wing mode at any time during verting, including from rotation
Wing mode is to the transition mode of fixed-wing mode conversion and from fixed-wing mode to the transition mode of rotary wings mode conversion, transition
The control strategy of mode is as follows:
Rotor control system works simultaneously with fixed-wing steerable system, and control weight changes with a square alternating for air speed,
Until being transformed into a certain mode completely.
Further, it is divided into 2 steps from rotor mode to the process of the transition mode of fixed-wing mode conversion:
Step 1: outermost two tilting rotors tilt forward on the wing of two sides, during verting, rotor and fixation
The wing uses mixing mode, and control weight is with square variation of air speed, and air speed is bigger, and fixed-wing control weight is higher, this stage
Rotor control is controlled using six rotors；When two tilting rotors of outermost vert completion and air speed reaches V1, rotor control adopt
With quadrotor scheme control, and two tilting rotors of outermost is allowed to accelerate to provide forward speed and Heading control；
Step 2: when unmanned plane reaches the safe speed and height of setting, intermediate two tilting rotors are tilted forward, are controlled
Weight continues square variation with preceding winged air speed, when the process time that verts is greater than T1And speed reaches V2, then it is assumed that it has verted
At.
Further, in rotor mode into fixed-wing mode conversion process, if mode switches back to rotor mode, institute
There is tilting rotor to return rapidly to incline, the control of six rotors is called in rotor control, and rotor mode weight is set as 100%.
Further, it is divided into 2 steps from fixed-wing mode to the process of the transition mode of rotor mode conversion:
Step 1: outermost two tilting rotors vert backward on the wing of two sides, and fixed-wing and rotor use mixing at this time
Mode, control weight with air speed square variation, air speed is smaller, rotor control weight it is higher, when vert complete when and air speed
Reach V1, rotor calling quadrotor control；
Step 2: intermediate two tilting rotors vert backward, complete until verting, and call the control of six rotors.
Further, in fixed-wing mode into rotor mode conversion process, if mode switches back to fixed-wing mode, institute
There is tilting rotor to lean forward rapidly, rapidly return back to fixed-wing mode, fixed-wing mode weight is set as 100%.
By adopting the above technical scheme bring the utility model has the advantages that
The present invention is promoted on power using distributed, is solved VTOL and is asked with power match when cruising flight
Topic；Use more rotors substep on strategy verting and vert, at the same can also be realized using part tilting rotor motor-driven enhancing and
Fitful wind mitigates side-jet control, finally plays the redundancy backup effect of multiple rotors, faults-tolerant control when achievement unit separated motor failure.
Therefore the present invention have speed is fast, voyage is remote, task is flexible, can VTOL many advantages, such as, greatly improve flight safety
Property and flight quality.
Detailed description of the invention
Fig. 1 is the overall structure diagram of unmanned plane of the present invention；
Fig. 2 is tilting rotor schematic diagram in the present invention；
Fig. 3 is vertical passage control principle block diagram in the present invention；
Fig. 4 is interconnection control principle block diagram in the present invention；
Fig. 5 is the flow chart of transition mode in the present invention.
Specific embodiment
Below with reference to attached drawing, technical solution of the present invention is described in detail.
As shown in Figure 1, a kind of distributed-power tilting rotor wing unmanned aerial vehicle, including fuselage 1, wing 2 and rudder-vator 3.Also
Including the second rotor system 5 at the first rotor system 4 being set on the wing and empennage.
As shown in Fig. 2, the first rotor system includes four tilting rotors that can be verted being arranged on wing and use
In the inclining rotary mechanism to vert, including motor 5, blade 6 and the motor 7 that verts.
The present embodiment uses optimal technical scheme, and each inclining rotary mechanism has limiting device, within the scope of 0~90 °, in Fig. 2
Left and right be respectively tilt angle be 0 and 90 ° of schematic diagram.
The present embodiment uses optimal technical scheme, by the gravity axis of the rotary shaft of rotor and unmanned plane in the first rotor system
It is overlapped, during verting, the center of gravity approximation of unmanned plane is constant.
The present embodiment uses optimal technical scheme, and the propeller in the second rotor system uses variable-pitch propeller, in rotor
The longitudinal stability of unmanned plane is controlled under mode.
In the present invention, number of motors is more in the first rotor system, has certain redundancy for aircraft safety.
Present invention is alternatively directed to above-mentioned distributed-power tilting rotor wing unmanned aerial vehicles to devise control method.
(1) rotor Model control is restrained:
Using two-stage PID control, first according to desired roll anglePitching angle thetad, yaw angle ψdWith actual rolling
AnglePitching angle theta, the margin of error of yaw angle ψ construct first order PID control, i.e. attitude angle closed-loop control is restrained:
In above formula, pd、qd、rdFor expectation angular velocity in roll, expectation rate of pitch and expectation yaw rate；eθ=θd- θ, eψ=ψd-ψ；kθp、kψpIt is roll angle, pitch angle and the scale parameter for yawing angle error；
According to desired angular speed (pd,qd,rd) with the margin of error of actual angular speed (p, q, r), construct second level PID control,
I.e. attitude angular velocity closed-loop control is restrained:
In above formula, Δ δp、Δδq、ΔδrIt is the PID output of three attitude angular velocities；ep=pd- p, eq=qd- q, er=rd-
r；kpp、kqp、krpIt is the scale parameter of three attitude angular velocities；, kpd、kqd、krdIt is the differential parameter of three attitude angular velocities；
(2) fixed-wing Model control is restrained:
Fixed-wing Model control rule includes vertical passage control law and interconnection control law.
Longitudinally controlled channel control pitch angle and height are kept, and include 3 circuits: damping in pitch inner looping, pitch angle are protected
Hold circuit and height control loop.The ω that damping in pitch inner looping is exported by IMUyIt is fed back, is constituted in pitch angle damping
Ring, meanwhile, pitch angle is constituted according to the pitch angle feedback of integrated navigation output and controls external loop.
Since small drone has a constant value angle of attack in flat fly, therefore constant value trim is added in pitch loop
Pitch command, and corresponding tromming tab face angle is added in elevator.Height keeps circuit to be located at outermost layer circuit, passes through setting
Height value and the output height value of integrated navigation form height tolerance, to be converted into corresponding pitch command, the circuit
Using proportional plus integral control mode.Instruction clipping will be added in pitch command, prevent the excessive maneuver of unmanned plane.
As shown in figure 3, vertical passage control law:
In above formula, δeControl amount is exported for elevator；Δ H=Hc- H, HcFor altitude instruction, H is actual height；KHpFor height
Spend ring proportional component coefficient, KHiFor height ring integral element coefficient；For the proportional component coefficient of angular speed ring, KωyAngular speed
The differentiation element coefficient of ring；For pitch angle, ωyFor pitch rate；T is the sampling period, and k is sampling period number；
α is the switching coefficient of integral term:
Wherein, Δ H (t) is the Δ H value of current time t, and ε is the difference in height judgment threshold of setting；
As shown in figure 4, interconnection control law:
When lateral deviation is away from D > M:
δa=KD1·D+Kψ·Δψ+Kωx·ωx
When lateral deviation is away from D≤M:
Wherein, δaControl amount, δ are exported for aileronrControl amount is exported for rudder；ωx、ωzRespectively angular velocity in roll,
Yaw rate；Δ ψ=ψd-ψ；β is the switching coefficient of integral term；KD1、KD2Respectively segment processing when lateral deviation away from ratio system
Number；Kψ、KγFor the proportional component coefficient of angular speed ring；Kωz、KωxThe differentiation element coefficient of angular speed ring；M be setting lateral deviation away from
Threshold value, in the present embodiment, M are preferably 50 meters.
It needs to switch in rotor mode and fixed-wing mode at any time during verting, including from rotor mode to admittedly
The transition mode of wing mode conversion is determined and from fixed-wing mode to the transition mode of rotary wings mode conversion, as shown in figure 5, wherein
It (a) is from fixed-wing mode to the flow chart of rotary wings mode conversion, (b) for from rotor mode to the stream of fixed-wing mode conversion
Cheng Tu.
Rotor mode switchs to fixed-wing mode and is divided into 2 steps, and actual measurement inclining rotary mechanism completes needs from starting to be tilted to vert
10s, so mode of verting needs the general 20s time altogether.Two motors of outermost tilt forward on first step wing, are verting
In the process, rotor and fixed-wing use mixing mode, and control is with square variation of preceding winged air speed, and air speed is bigger, fixed-wing control
Weight processed is higher, and the rotor control in this stage is controlled using six rotors.When two motors of outermost vert completion and air speed reaches
To 12m/s, rotor control uses quadrotor scheme control, and two motors of outermost is allowed to accelerate to provide forward speed.?
On the basis of this, two motors of completion of verting also are used to provide Heading control.In other words, Heading control is controlled in original quadrotor
On the basis of, it is superimposed the control of two forward direction motors, can preferably be held position in this way.It is intermediate on wing that second step, which verts,
Two motors vert, and the premise verted is that aircraft has arrived at certain safe speed and height, and control weight continues to fly with preceding
Square variation of air speed.Think to vert when the process time that verts is greater than 20s and reaches safe speed 18m/s by time judgement
It completes.In rotor into fixed-wing mode conversion process, if mode switch switchback rotor mode, all motors return rapidly
Incline, the control of six rotors is called in rotor control, and rotor mode weight is set as 100%.
Fixed-wing mode turns to rotor mode conversion and is equally divided into 2 steps.Conversion process is that rotor mode turns fixed-wing mode
Inverse process, the first step is that two motors vert backward on the outside of wing, and fixed-wing and rotor use mixing mode, control at this time
With air speed square variation, air speed is smaller, rotor control weight it is higher, when vert complete when and air speed reach 12m/s, rotor
Call quadrotor control.Second step is that the motor of inside two verts backward, until completion of verting is converted into the control of six rotors.Solid
Determine flapwise rotor to vert under mode, if pattern switching, to fixed-wing mode, all motors lean forward rapidly, goes to fixed-wing rapidly
Mode, fixed-wing mode weight are set as 100%.
Embodiment is merely illustrative of the invention's technical idea, and this does not limit the scope of protection of the present invention, it is all according to
Technical idea proposed by the present invention, any changes made on the basis of the technical scheme are fallen within the scope of the present invention.
Claims (10)
1. a kind of distributed-power tilting rotor wing unmanned aerial vehicle, including fuselage, wing and rudder-vator, it is characterised in that: further include setting
The second rotor system setting the first rotor system on wing and being arranged at rudder-vator；First rotor system includes 4
A tilting rotor, this 4 tilting rotors are arranged in two-by-two on left side wing and right side wing；Second rotor system includes 2
A rotor, this 2 rotors are separately positioned on the two sides of V-arrangement wing.
2. distributed-power tilting rotor wing unmanned aerial vehicle according to claim 1, it is characterised in that: respectively incline in the first rotor system
The rotary shaft of switch rotor and the center of gravity overlapping of axles of unmanned plane, during verting, the center of gravity approximation of unmanned plane is constant.
3. distributed-power tilting rotor wing unmanned aerial vehicle according to claim 1, it is characterised in that: each tilting rotor has limit
Position device, so that the tilt angle of tilting rotor is limited within the scope of 0~90 °.
4. distributed-power tilting rotor wing unmanned aerial vehicle according to claim 1, it is characterised in that: the rotation in the second rotor system
The propeller of the wing uses variable-pitch propeller.
5. the control method based on distributed-power tilting rotor wing unmanned aerial vehicle described in claim 1, which is characterized in that including following
Control law:
(1) rotor Model control is restrained:
Using two-stage PID control, first according to desired roll anglePitching angle thetad, yaw angle ψdWith actual roll angle
Pitching angle theta, the margin of error of yaw angle ψ construct first order PID control, i.e. attitude angle closed-loop control is restrained:
In above formula, pd、qd、rdFor expectation angular velocity in roll, expectation rate of pitch and expectation yaw rate；eθ
=θd- θ, eψ=ψd-ψ；kθp、kψpIt is roll angle, pitch angle and the scale parameter for yawing angle error；
According to desired angular speed (pd,qd,rd) with the margin of error of actual angular speed (p, q, r), construct second level PID control, i.e. appearance
State angular speed closed-loop control rule:
In above formula, Δ δp、Δδq、ΔδrIt is the PID output of three attitude angular velocities；ep=pd- p, eq=qd- q, er=rd-r；
kpp、kqp、krpIt is the scale parameter of three attitude angular velocities；, kpd、kqd、krdIt is the differential parameter of three attitude angular velocities；
(2) fixed-wing Model control is restrained:
Including vertical passage control law and interconnection control law, wherein vertical passage control law:
In above formula, δeControl amount is exported for elevator；Δ H=Hc- H, HcFor altitude instruction, H is actual height；KHpFor height ring
Proportional component coefficient, KHiFor height ring integral element coefficient；For the proportional component coefficient of angular speed ring, KωyAngular speed ring
Differentiation element coefficient；For pitch angle, ωyFor pitch rate；T is the sampling period, and k is sampling period number；
α is the switching coefficient of integral term:
Wherein, Δ H (k) is the Δ H value of current time t, and ε is the difference in height judgment threshold of setting；
Interconnection control law:
When lateral deviation is away from D > M:
δa=KD1·D+Kψ·Δψ+Kωx·ωx
When lateral deviation is away from D≤M:
Wherein, δaControl amount, δ are exported for aileronrControl amount is exported for rudder；ωx、ωzRespectively angular velocity in roll, yaw angle
Speed；Δ ψ=ψd-ψ；β is the switching coefficient of integral term；KD1、KD2Respectively segment processing when lateral deviation away from proportionality coefficient；Kψ、
KγFor the proportional component coefficient of angular speed ring；Kωz、KωxThe differentiation element coefficient of angular speed ring；M is the lateral deviation of setting away from threshold value.
6. control method according to claim 5, which is characterized in that needed during verting at any time in rotor mode and solid
Determine wing mode to switch over, including the transition mode from rotor mode to fixed-wing mode conversion and from fixed-wing mode to rotation
The transition mode of wing mode conversion, the control strategy of transition mode are as follows:
Rotor control system works simultaneously with fixed-wing steerable system, and control weight changes with a square alternating for air speed, until
It is transformed into a certain mode completely.
7. control method according to claim 6, which is characterized in that from rotor mode to the stage die of fixed-wing mode conversion
The process of state is divided into 2 steps:
Step 1: outermost two tilting rotors tilt forward on the wing of two sides, and during verting, rotor and fixed-wing are adopted
With mixing mode, weight is controlled with square variation of air speed, air speed is bigger, and fixed-wing control weight is higher, the rotor in this stage
Control is controlled using six rotors；When two tilting rotors of outermost vert completion and air speed reaches V1, rotor control is using four
Rotor mode control, and two tilting rotors of outermost is allowed to accelerate to provide forward speed and Heading control；
Step 2: when unmanned plane reaches the safe speed and height of setting, intermediate two tilting rotors are tilted forward, and control weight
Continue square variation with preceding winged air speed, when the process time that verts is greater than T1And speed reaches V2, then it is assumed that completion of verting.
8. control method according to claim 7, which is characterized in that in rotor mode into fixed-wing mode conversion process,
If mode switches back to rotor mode, all tilting rotors, which return rapidly, to incline, and the control of six rotors, rotor mode are called in rotor control
Weight is set as 100%.
9. control method according to claim 6, which is characterized in that from fixed-wing mode to the stage die of rotor mode conversion
The process of state is divided into 2 steps:
Step 1: outermost two tilting rotors vert backward on the wing of two sides, and fixed-wing and rotor use mixing side at this time
Formula, control weight with air speed square variation, air speed is smaller, rotor control weight it is higher, when vert complete when and air speed reach
To V1, rotor calling quadrotor control；
Step 2: intermediate two tilting rotors vert backward, complete until verting, and call the control of six rotors.
10. control method according to claim 9, which is characterized in that in fixed-wing mode into rotor mode conversion process,
If mode switches back to fixed-wing mode, all tilting rotors lean forward rapidly, rapidly return back to fixed-wing mode, fixed-wing mode power
It resets and is set to 100%.
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| CN111026146A (en) * | 2019-12-24 | 2020-04-17 | 西北工业大学 | Attitude control method for composite wing vertical take-off and landing unmanned aerial vehicle |
| CN111196346A (en) * | 2019-11-20 | 2020-05-26 | 湖北航天飞行器研究所 | Distributed electric propulsion tilt rotor unmanned aerial vehicle |
| CN111258324A (en) * | 2020-01-19 | 2020-06-09 | 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 |
| CN112357060A (en) * | 2020-10-29 | 2021-02-12 | 天津腾云智航科技有限公司 | Fixed wing unmanned aerial vehicle wing flap elevation convergence control method |
| CN113075879A (en) * | 2020-01-03 | 2021-07-06 | 中国科学院沈阳自动化研究所 | Engine control system of tilt rotor unmanned aerial vehicle |
| CN113148135A (en) * | 2021-04-08 | 2021-07-23 | 南京航空航天大学 | Multi-vector thrust tilt rotor unmanned aerial vehicle and course control method thereof |
| CN116643578A (en) * | 2023-07-18 | 2023-08-25 | 北京航空航天大学 | Multimode unified control method for microminiature tailstock unmanned aerial vehicle |
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