CN110316368B - Distributed power tilt rotor unmanned aerial vehicle and control method thereof - Google Patents

Distributed power tilt rotor unmanned aerial vehicle and control method thereof Download PDF

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CN110316368B
CN110316368B CN201910269423.6A CN201910269423A CN110316368B CN 110316368 B CN110316368 B CN 110316368B CN 201910269423 A CN201910269423 A CN 201910269423A CN 110316368 B CN110316368 B CN 110316368B
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rotor
control
mode
wing
rotors
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CN110316368A (en
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李康一
曾旭
王新华
岳凤玉
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Nanjing University of Aeronautics and Astronautics
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Nanjing University of Aeronautics and Astronautics
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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C19/00—Aircraft control not otherwise provided for
    • 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
    • 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
    • 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

Abstract

The invention discloses a distributed power tilt rotor unmanned aerial vehicle and a control method thereof. The distributed power tilt rotor unmanned aerial vehicle comprises a first rotor system arranged on a wing and a second rotor system arranged at a V-shaped empennage; the first rotor system comprises 4 tilting rotors, and the 4 tilting rotors are arranged on the left side wing and the right side wing in pairs; the second rotor system comprises 2 rotors, which are arranged on both sides of the V-shaped wing. The control method of the unmanned aerial vehicle comprises a rotor wing mode control law, a fixed wing mode control law and a transition mode control strategy. The invention can realize the optimal power matching of the power system and provide robust and reliable control for each state of the rotorcraft.

Description

Distributed power tilt rotor unmanned aerial vehicle and control method thereof
Technical Field
The invention belongs to the technical field of unmanned aerial vehicles, and particularly relates to a tilt rotor unmanned aerial vehicle.
Background
The tilt rotor aircraft can take off and land vertically and hover in the air like a helicopter, can fly with the fixed wing quickly like the fixed wing, and has the characteristics of high efficiency, high speed and long voyage. Therefore, the tilt rotor machine has wide application prospect.
The tilt rotor aircraft is always a hotspot of domestic and foreign research, and faces the technical problems of complex aerodynamic characteristics, difficult control and the like: the difference between the power requirements of the engine in the vertical takeoff and fixed wing flight modes is large, and the optimal power matching is difficult to realize; simultaneously, the aerodynamic instability that synchronous verts caused also makes the gyroplane difficult to control.
Therefore, there is still a need in the art to develop new tiltrotor aircraft that are structurally simple and provide reliable solutions for controlling various states of the rotorcraft.
Disclosure of Invention
In order to solve the technical problems in the background art, the invention aims to provide a distributed power tilt rotor unmanned aerial vehicle and a control method thereof.
In order to achieve the technical purpose, the technical scheme of the invention is as follows:
a distributed power tilt rotor unmanned aerial vehicle comprises a fuselage, wings and a V-shaped empennage, and further comprises a first rotor system arranged on the wings and a second rotor system arranged at the V-shaped empennage; the first rotor system comprises 4 tilting rotors, and the 4 tilting rotors are arranged on the left side wing and the right side wing in pairs; the second rotor system contains 2 rotors, and these 2 rotors set up respectively in the both sides of V-arrangement wing.
Further, the rotation axis of rotor that verts among the first rotor system each coincides with unmanned aerial vehicle's center of gravity axle, and at the in-process that verts, unmanned aerial vehicle's focus is approximate unchangeable.
Further, every rotor that verts has stop device to the angle of verting that will vert the rotor limits at 0 ~ 90 within range.
Further, the propellers of the rotors in the second rotor system are variable pitch propellers.
The control method based on the distributed power tilt rotor unmanned aerial vehicle comprises the following control laws:
(1) rotor mode control law:
with two-stage PID control, first according to the desired roll angle
Figure GDA0002728031920000021
Pitch angle thetadYaw angle psidAngle of roll to actual
Figure GDA0002728031920000022
And constructing a first-level PID control, namely an attitude angle closed-loop control law, according to the error amount of the pitch angle theta and the yaw angle psi:
Figure GDA0002728031920000023
in the above formula, pd、qd、rdFor desired roll rate, desired pitch rateDegrees and desired yaw rate;
Figure GDA0002728031920000024
eθ＝θd-θ，eψ＝ψd-ψ；
Figure GDA0002728031920000025
kθp、kψpthe roll angle, the pitch angle and the yaw angle are proportional parameters;
according to desired angular velocity (p)d,qd,rd) And (3) constructing a second-stage PID control, namely an attitude angular velocity closed-loop control law, according to the error amount with the actual angular velocity (p, q, r):
Figure GDA0002728031920000026
in the above formula,. DELTA.p、Δq、ΔrIs the PID output of the three attitude angular velocities; e.g. of the typep＝pd-p，eq＝qd-q，er＝rd-r；kpp、kqp、krpIs a proportional parameter of the three attitude angular velocities; k is a radical ofpd、kqd、krdAre differential parameters of the three attitude angular velocities;
(2) fixed-wing modal control law:
the method comprises a longitudinal channel control law and a transverse channel control law, wherein the longitudinal channel control law:
Figure GDA0002728031920000031
in the above formula, the first and second carbon atoms are,eoutputting a control quantity for the elevator; Δ H ═ Hc-H，HcH is the actual height; kHpProportional Ring coefficient of height Ring, KHiIs a height loop integral link coefficient;
Figure GDA0002728031920000035
being angular rate ringsCoefficient of proportionality, KωyDifferential link coefficients of the angular rate loop;
Figure GDA0002728031920000036
is a pitch angle, ωyIs the pitch angle rate; t is a sampling period, and k is the number of sampling periods;
α is the switching coefficient of the integral term:
Figure GDA0002728031920000032
wherein, Δ H (t) is the Δ H value at the current time t, and is a set height difference judgment threshold;
transverse channel control law:
when the sideslip distance D > M:
a＝KD1·D+Kψ·Δψ+Kωx·ωx
Figure GDA0002728031920000033
when the side offset distance D is less than or equal to M:
Figure GDA0002728031920000034
wherein the content of the first and second substances,aoutputs a controlled variable for the ailerons,rOutputting a control quantity for the rudder; omegax、ωzRespectively roll angular velocity and yaw angular velocity; delta psi ═ psid- ψ; beta is the switching coefficient of the integral term; kD1、KD2Respectively are lateral offset proportional coefficients during sectional processing; kψ、KγThe proportional link coefficient of the angular rate ring; kωz、KωxDifferential link coefficients of the angular rate loop; and M is a set lateral offset threshold value.
Further, during the tilting process, switching between a rotor mode and a fixed wing mode at any time is required, wherein the switching includes a transition mode from the rotor mode to the fixed wing mode and a transition mode from the fixed wing mode to the rotary wing mode, and a control strategy of the transition mode is as follows:
the rotor wing control system and the fixed wing control system act simultaneously, and the control weight changes along with the square of the airspeed until the mode is completely converted.
Further, the transition mode process from the rotor mode to the fixed wing mode is divided into 2 steps:
step 1: the two tilting rotors on the outermost sides of the wings on the two sides tilt forwards, in the tilting process, the rotors and the fixed wings adopt a mixed control mode, the control weight changes along with the square of the airspeed, the larger the airspeed, the higher the control weight of the fixed wings, and the rotor control at the stage adopts six-rotor control; when two tilting rotors at the outermost side tilt and the airspeed reaches V1The rotor control adopts a four-rotor mode control, and two tilting rotors at the outermost sides are accelerated to provide forward speed and course control;
step 2: when unmanned aerial vehicle reachs the safe speed and the height of setting for, two middle rotor wings that vert forward, and control weight continues to fly the square change of airspeed along with the front, is greater than T when the process time that verts1And the speed reaches V2Then the tilting is deemed complete.
Further, in the process of converting the rotor mode to the fixed wing mode, if the mode is switched back to the rotor mode, all tilting rotors tilt back rapidly, six-rotor control is called by the rotor control, and the weight of the rotor mode is set to be 100%.
Further, the transition mode process from the fixed wing mode to the rotor mode is divided into 2 steps:
step 1: two rotors that vert in the outside on the wing of both sides are verted backward, and the fixed wing and rotor adopt the mixed control mode this moment, and control weight is along with the square change of airspeed, and the airspeed is less, and rotor control weight is higher, and when verting the completion and the airspeed reaches V1The rotor calls four-rotor control;
step 2: two middle tilt rotors tilt backwards until the tilt is completed, and six rotor controls are called.
Further, in the process of converting the fixed-wing mode to the rotor mode, if the mode is switched back to the fixed-wing mode, all the tilt rotors tilt forward rapidly and return to the fixed-wing mode rapidly, and the weight of the fixed-wing mode is set to be 100%.
Adopt the beneficial effect that above-mentioned technical scheme brought:
the invention adopts distributed propulsion on power, thus solving the problem of power matching during vertical take-off and landing and cruising flight; adopt many rotors to vert step by step on the strategy of verting, can also utilize partial rotor that verts to realize motor-driven reinforcing and gust and relax direct force control simultaneously, the redundant backup effect of a plurality of rotors of at last performance realizes fault-tolerant control when partial motor trouble. Therefore, the invention has the advantages of high speed, long range, flexible task, capability of vertically lifting and the like, and greatly improves the flight safety and the flight quality.
Drawings
Fig. 1 is a schematic diagram of the overall structure of the unmanned aerial vehicle of the invention;
figure 2 is a schematic view of a tiltrotor rotor of the present invention;
FIG. 3 is a block diagram of the longitudinal channel control concept of the present invention;
FIG. 4 is a block diagram of the cross-lane control concept of the present invention;
fig. 5 is a flow chart of the transition mode in the present invention.
Detailed Description
The technical scheme of the invention is explained in detail in the following with the accompanying drawings.
As shown in fig. 1, a distributed power tilt rotor unmanned aerial vehicle includes a fuselage 1, wings 2 and a V-shaped empennage 3. And a first rotor system 4 arranged on the wing and a second rotor system 5 arranged at the tail wing.
As shown in fig. 2, the first rotor system includes four tiltrotors that can tilt and a tilting mechanism for tilting, which are provided on the wing, and includes a motor 5, a blade 6, and a tilting motor 7.
This embodiment adopts preferred technical scheme, and every mechanism that verts has stop device, and 0 ~ 90 within range, the left and right sides in figure 2 are tilting angle respectively for the schematic diagram of 0 and 90.
This embodiment adopts preferred technical scheme, with the rotation axis of rotor among the first rotor system and unmanned aerial vehicle's the coincidence of axis of gravity, at the in-process that verts, unmanned aerial vehicle's focus is approximately unchangeable.
This embodiment adopts preferred technical scheme, and the screw in the second rotor system uses the variable pitch screw, controls unmanned aerial vehicle's longitudinal stability under the rotor mode.
In the invention, the number of motors in the first rotor system is large, and certain redundancy is provided for the safety of the airplane.
The invention also designs a control method aiming at the distributed power tilt rotor unmanned aerial vehicle.
(1) Rotor mode control law:
with two-stage PID control, first according to the desired roll angle
Figure GDA0002728031920000061
Pitch angle thetadYaw angle psidAngle of roll to actual
Figure GDA0002728031920000062
And constructing a first-level PID control, namely an attitude angle closed-loop control law, according to the error amount of the pitch angle theta and the yaw angle psi:
Figure GDA0002728031920000063
in the above formula, pd、qd、rdDesired roll rate, desired pitch rate, and desired yaw rate;
Figure GDA0002728031920000064
eθ＝θd-θ，eψ＝ψd-ψ；
Figure GDA0002728031920000065
kθp、kψpthe roll angle, the pitch angle and the yaw angle are proportional parameters;
according to desired angular velocity (p)d,qd,rd) And (3) constructing a second-stage PID control, namely an attitude angular velocity closed-loop control law, according to the error amount with the actual angular velocity (p, q, r):
Figure GDA0002728031920000066
in the above formula,. DELTA.p、Δq、ΔrIs the PID output of the three attitude angular velocities; e.g. of the typep＝pd-p，eq＝qd-q，er＝rd-r；kpp、kqp、krpIs a proportional parameter of the three attitude angular velocities; k is a radical ofpd、kqd、krdAre differential parameters of the three attitude angular velocities;
(2) fixed-wing modal control law:
the fixed wing mode control law comprises a longitudinal channel control law and a transverse channel control law.
The longitudinal control channel controls pitch and altitude maintenance, and comprises 3 loops: a pitch damping inner loop, a pitch angle maintenance loop and a height control loop. Omega output by pitch damping inner loop through IMUyAnd feeding back to form a pitch angle damping inner ring, and forming a pitch angle control outer loop according to the pitch angle feedback output by the integrated navigation.
Because the small unmanned aerial vehicle has a constant attack angle when flying flatly, a constant trim pitch angle instruction is added in a pitch angle control loop, and a corresponding trim rudder surface angle is added in an elevator. The height maintaining loop is positioned in the outermost loop, height deviation is formed through a set height value and an output height value of the combined navigation, and accordingly the height deviation is converted into a corresponding pitch angle instruction, and the loop adopts a proportional-integral control mode. The pitch angle instruction is added with instruction amplitude limiting, and the phenomenon of overlarge maneuvering action of the unmanned aerial vehicle is prevented.
As shown in fig. 3, the longitudinal channel control law:
Figure GDA0002728031920000071
in the above formula, the first and second carbon atoms are,eoutputting a control quantity for the elevator; Δ H ═ Hc-H，HcH is the actual height; kHpProportional Ring coefficient of height Ring, KHiIs a height loop integral link coefficient;
Figure GDA0002728031920000075
is the proportional link coefficient of the angular rate loop, KωyDifferential link coefficients of the angular rate loop;
Figure GDA0002728031920000076
is a pitch angle, ωyIs the pitch angle rate; t is a sampling period, and k is the number of sampling periods;
α is the switching coefficient of the integral term:
Figure GDA0002728031920000072
wherein, Δ H (t) is the Δ H value at the current time t, and is a set height difference judgment threshold;
as shown in fig. 4, the lateral channel control law:
when the sideslip distance D > M:
a＝KD1·D+Kψ·Δψ+Kωx·ωx
Figure GDA0002728031920000073
when the side offset distance D is less than or equal to M:
Figure GDA0002728031920000074
wherein the content of the first and second substances,aoutputs a controlled variable for the ailerons,rOutputting a control quantity for the rudder; omegax、ωzRespectively roll angular velocity and yaw angular velocity; delta psi ═ psid- ψ; beta isThe switching coefficient of the integral term; kD1、KD2Respectively are lateral offset proportional coefficients during sectional processing; kψ、KγThe proportional link coefficient of the angular rate ring; kωz、KωxDifferential link coefficients of the angular rate loop; m is a set offset threshold, and M is preferably 50 meters in this embodiment.
During the tilting process, it is necessary to switch between the rotor mode and the fixed-wing mode at any time, including a transition mode from the rotor mode to the fixed-wing mode and a transition mode from the fixed-wing mode to the rotary-wing mode, as shown in fig. 5, where (a) is a flowchart of the transition from the fixed-wing mode to the rotary-wing mode, and (b) is a flowchart of the transition from the rotor mode to the fixed-wing mode.
The conversion of the rotor mode into the fixed wing mode is divided into 2 steps, and the actually measured tilting mechanism needs 10s from the beginning of tilting to the completion of tilting, so that the total time of the tilting mode needs about 20 s. Two motors in the outside incline forward on the wing of first step, at the in-process that verts, rotor and stationary vane adopt the mixed control mode, and the square change of airspeed is flown forward to the control weight, and the airspeed is big more, and stationary vane control weight is high more, and the rotor control at this stage adopts six rotor controls. When the two motors at the outermost side are completely tilted and the airspeed reaches V1At 12m/s, the rotor control uses a four-rotor mode control and accelerates the outermost two motors to provide forward speed. On the basis, the two motors which are tilted are also used for providing heading control. In other words, the heading control superposes the control of two forward motors on the basis of the original four-rotor control, so that the attitude can be stabilized better. The second tilting is that the two motors in the middle on the wing tilt, and the premise of tilting is that the airplane reaches a certain safe speed and height, and the control weight continues to change along with the square of the forward flying airspeed. Judging whether the tilting process time is more than T through the time120s and reaches a safe speed V2Tilting is considered complete at 18 m/s. In the process of converting the rotor to the fixed wing mode, if the mode switch is switched back to the rotor mode, all motors quickly tilt back, six rotor controls are called by the rotor control, and the weight of the rotor mode is set to be 100%.
The conversion of the fixed wing mode to the rotor mode is also divided into 2 steps. Conversion process is rotor mode changes the reverse process of stationary vane mode, and first step is that two motors in the wing outside vert backward, and stationary vane and rotor adopt the mixed accuse mode this moment, and the control right changes along with the square of airspeed, and the airspeed is less, and rotor control weight is higher, and the airspeed reaches 12m/s when verting the completion, and the rotor calls four rotor controls. The second step is that two motors in the inboard vert backward, accomplish to vert and convert six rotor controls into. Under the mode that the fixed wing tilts towards the rotor wing, if the mode is switched to the fixed wing mode, all motors tilt forwards rapidly and turn to the fixed wing mode rapidly, and the weight of the fixed wing mode is set to be 100%.
The embodiments are only for illustrating the technical idea of the present invention, and the technical idea of the present invention is not limited thereto, and any modifications made on the basis of the technical scheme according to the technical idea of the present invention fall within the scope of the present invention.

Claims (9)

1. Control method based on distributed power rotor unmanned aerial vehicle that verts, distributed power rotor unmanned aerial vehicle that verts includes fuselage, wing and V-arrangement fin, its characterized in that: the first rotor system is arranged on the wing, and the second rotor system is arranged at the V-shaped empennage; the first rotor system comprises 4 tilting rotors, and the 4 tilting rotors are arranged on the left side wing and the right side wing in pairs; the second rotor system comprises 2 rotors, and the 2 rotors are respectively arranged on two sides of the V-shaped wing;
the method is characterized by comprising the following control laws:
(1) rotor mode control law:
with two-stage PID control, first according to the desired roll angle
Figure FDA0002728031910000011
Pitch angle thetadYaw angle psidAngle of roll to actual
Figure FDA0002728031910000012
Pitch angle thetaAnd the error amount of the yaw angle psi, constructing a first-level PID control, namely an attitude angle closed-loop control law:
Figure FDA0002728031910000013
in the above formula, pd、qd、rdDesired roll rate, desired pitch rate, and desired yaw rate;
Figure FDA0002728031910000014
eθ＝θd-θ，eψ＝ψd-ψ；
Figure FDA0002728031910000015
kθp、kψpthe roll angle, the pitch angle and the yaw angle are proportional parameters;
according to desired angular velocity (p)d,qd,rd) And (3) constructing a second-stage PID control, namely an attitude angular velocity closed-loop control law, according to the error amount with the actual angular velocity (p, q, r):
Figure FDA0002728031910000016
in the above formula,. DELTA.p、Δq、ΔrIs the PID output of the three attitude angular velocities; e.g. of the typep＝pd-p，eq＝qd-q，er＝rd-r；kpp、kqp、krpIs a proportional parameter of the three attitude angular velocities; k is a radical ofpd、kqd、krdAre differential parameters of the three attitude angular velocities;
(2) fixed-wing modal control law:
the method comprises a longitudinal channel control law and a transverse channel control law, wherein the longitudinal channel control law:
Figure FDA0002728031910000021
in the above formula, the first and second carbon atoms are,eoutputting a control quantity for the elevator; Δ H ═ Hc-H，HcH is the actual height; kHpProportional Ring coefficient of height Ring, KHiIs a height loop integral link coefficient;
Figure FDA0002728031910000026
is the proportional link coefficient of the angular rate loop, KωyDifferential link coefficients of the angular rate loop;
Figure FDA0002728031910000025
is a pitch angle, ωyIs the pitch angle rate; t is a sampling period, and k is the number of sampling periods;
α is the switching coefficient of the integral term:
Figure FDA0002728031910000022
wherein, Δ H (k) is a Δ H value at the current time t, and is a set height difference judgment threshold;
transverse channel control law:
when the sideslip distance D > M:
a＝KD1·D+Kψ·Δψ+Kωx·ωx
Figure FDA0002728031910000023
when the side offset distance D is less than or equal to M:
Figure FDA0002728031910000024
wherein the content of the first and second substances,aoutputs a controlled variable for the ailerons,rOutputting a control quantity for the rudder; omegax、ωzRespectively roll angular velocity and yaw angular velocity; delta psi ═ psid- ψ; beta is the switching coefficient of the integral term; kD1、KD2Respectively are lateral offset proportional coefficients during sectional processing; kψ、KγThe proportional link coefficient of the angular rate ring; kωz、KωxDifferential link coefficients of the angular rate loop; and M is a set lateral offset threshold value.
2. The control method according to claim 1, wherein switching between the rotor mode and the fixed-wing mode is required at any time during the tilting process, and the switching comprises a transition mode from the rotor mode to the fixed-wing mode and a transition mode from the fixed-wing mode to the rotary-wing mode, and the control strategy of the transition mode is as follows:
the rotor wing control system and the fixed wing control system act simultaneously, and the control weight changes along with the square of the airspeed until the mode is completely converted.
3. The control method according to claim 2, wherein the transition mode from the rotor mode to the fixed wing mode is performed in 2 steps:
step 1: the two tilting rotors on the outermost sides of the wings on the two sides tilt forwards, in the tilting process, the rotors and the fixed wings adopt a mixed control mode, the control weight changes along with the square of the airspeed, the larger the airspeed, the higher the control weight of the fixed wings, and the rotor control at the stage adopts six-rotor control; when two tilting rotors at the outermost side tilt and the airspeed reaches V1The rotor control adopts a four-rotor mode control, and two tilting rotors at the outermost sides are accelerated to provide forward speed and course control;
step 2: when unmanned aerial vehicle reachs the safe speed and the height of setting for, two middle rotor wings that vert forward, and control weight continues to fly the square change of airspeed along with the front, is greater than T when the process time that verts1And the speed reaches V2Then the tilting is deemed complete.
4. The control method of claim 3, wherein during a transition from a rotor mode to a fixed wing mode, if the mode is switched back to the rotor mode, all tilt rotors tilt back rapidly, the rotor control invokes a six-rotor control, and the rotor mode weight is set to 100%.
5. The control method according to claim 2, wherein the transition mode from the fixed-wing mode to the rotor mode is performed in 2 steps:
step 1: two rotors that vert in the outside on the wing of both sides are verted backward, and the fixed wing and rotor adopt the mixed control mode this moment, and control weight is along with the square change of airspeed, and the airspeed is less, and rotor control weight is higher, and when verting the completion and the airspeed reaches V1The rotor calls four-rotor control;
step 2: two middle tilt rotors tilt backwards until the tilt is completed, and six rotor controls are called.
6. The control method of claim 5, wherein during a transition from fixed-wing mode to rotor mode, if the mode is switched back to fixed-wing mode, all tilt rotors tilt forward rapidly and return to fixed-wing mode rapidly, and fixed-wing mode weight is set to 100%.
7. The control method according to claim 1, characterized in that: the rotation axis of each rotor that verts in the first rotor system coincides with unmanned aerial vehicle's center of gravity axle, and at the in-process that verts, unmanned aerial vehicle's focus is approximate unchangeable.
8. The control method according to claim 1, characterized in that: every rotor that verts has stop device to the angle of verting that will vert the rotor limits at 0 ~ 90 within range.
9. The control method according to claim 1, characterized in that: the propellers of the rotors in the second rotor system adopt variable pitch propellers.
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