CN109976364B - Attitude decoupling control method for six-rotor aircraft - Google Patents

Attitude decoupling control method for six-rotor aircraft Download PDF

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CN109976364B
CN109976364B CN201910232466.7A CN201910232466A CN109976364B CN 109976364 B CN109976364 B CN 109976364B CN 201910232466 A CN201910232466 A CN 201910232466A CN 109976364 B CN109976364 B CN 109976364B
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disturbance
aircraft
attitude
angle
rotor aircraft
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CN109976364A (en
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李涵雄
王雅昕
张沛毅
谭芳
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Central South University
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    • 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
    • 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
    • G05D1/10—Simultaneous control of position or course in three dimensions
    • G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft

Abstract

The embodiment of the invention provides a six-rotor aircraft attitude decoupling control method, which treats unnecessary coupling terms in each loop as output disturbance added to the loop, designs a proper disturbance compensator to compensate the output disturbance and the external disturbance together, and solves the adverse effects caused by strong coupling and external disturbance of an aircraft system. The disturbance observer in the embodiment of the invention can be used for decoupling control of the multi-input multi-output system, so that the disturbance suppression performance and the robust stability of the disturbance compensator are ensured, the decoupling and disturbance suppression are good, and the good tracking performance and the stability of the six-rotor aircraft can be ensured.

Description

Attitude decoupling control method for six-rotor aircraft
Technical Field
The invention belongs to the technical field of aircraft control, and particularly relates to a six-rotor aircraft attitude decoupling control method.
Background
The six-rotor aircraft is a miniature aircraft taking six motors as power, consists of six symmetrically distributed rotors, has the characteristics of small volume, simple mechanical structure, compact structure, capability of vertically taking off and landing and perfect flight control algorithm, and is widely used in civil and police fields of entertainment, aerial photography, investigation, early warning and the like. The aircraft can make the system present different work modes of under-drive or full drive by changing the tissue form of six rotors. Compare in general many rotor crafts and need through control attitude angle change and then control horizontal motion, six rotor crafts organisms are equipped with six connecting rods, evenly distributed is around the organism, and a rotor is connected to the outer end of every connecting rod. The rotary planes of the six rotors and the plane where the six connecting rods are located form six dip angles, the dip angle of the rotary plane of each rotor is equal to that of the rotary plane of the third rotor which is separated by one rotor, and the dip angles of the rotary planes of each rotor and the adjacent rotors and the plane where the six connecting rods are located are complementary angles. The six degrees of freedom of the aircraft are controlled by changing the rotating speed of each rotor wing, the high controllability and the high maneuverability are achieved, and the actions of flying forwards, hovering, flying forwards in any posture and the like can be realized.
The key point for ensuring the stable performance of the six-rotor aircraft is to overcome the influence caused by internal uncertainty and external disturbance of the system.
The internal uncertainty of the system is mainly due to the aircraft's own structure. The six-rotor aircraft has the characteristic of multivariable strong coupling, namely, the change of the state of each degree of freedom is influenced by the rotating speeds of a plurality of rotors, and the change of the rotating speed of any one rotor can also cause the change of the state of a plurality of degrees of freedom, and belongs to a Multiple-Input Multiple-Output (MIMO) system. The mutual coupling between the multiple control loops of the MIMO system makes the control of the MIMO system difficult. Generally, decoupling control needs to be performed on the MIMO system, that is, interaction between control loops in the system is eliminated by using a certain structure and finding a proper control law, so that each input controls only one corresponding output.
The external disturbance is mainly caused by external factors such as airflow and the like received by the aircraft during flight. Feedback control enables suppression of system disturbances when they are measurable, but accurate system models and disturbance dynamics are often difficult to obtain.
In conclusion, because the six-rotor aircraft system has the characteristics of multivariable strong coupling and susceptibility to disturbance, no effective solution exists for realizing high-quality attitude control of the aircraft.
Disclosure of Invention
In order to solve the problem that a six-rotor aircraft system has the characteristics of multivariable strong coupling and high susceptibility to disturbance in the prior art, the embodiment of the invention aims to provide a six-rotor aircraft attitude decoupling control method.
In order to achieve the above purpose, the embodiment of the invention adopts the following technical scheme:
a six-rotor aircraft attitude decoupling control method comprises the following steps:
(1) selecting a navigational coordinate system OXeYeZeBody coordinate system OXbYbZbAnd aircraft attitude angle vector
Figure BDA0002007149510000021
Establishing a six-rotor aircraft dynamic model;
(2) selecting flight system input, neglecting air resistance, three-channel coupling and gyro effective stress moment, and simplifying an aircraft dynamics model;
(3) considering the external disturbance term and the coupling disturbance term in each loop as output disturbance added on the loop, and designing a disturbance compensator;
(4) and compensating by using a disturbance compensator to realize decoupling and disturbance compensation of the aircraft attitude system.
In the step (2), air resistance, three-channel coupling and gyro effective stress moment are ignored.
And (4) compensating by using a disturbance compensator, namely tracking and estimating a coupling term between attitude angles and external disturbance by using the disturbance compensator and compensating.
Preferably, the compensating of step (4) further comprises using a cascade PID controller in combination.
Preferably, the roll angle
Figure BDA0002007149510000022
Is the X axis of the body and waterThe included angle between the planes is defined as [ -180 DEG ]](ii) a The pitch angle theta is the included angle between the Y axis of the machine body and the horizontal plane, and the definition domain is [ -90 DEG ]](ii) a The yaw angle psi is the included angle between the projection direction of the Z axis of the body on the horizontal plane and the parameter line on the plane, and the defined domain is [ 0-360 DEG ]]。
Preferably, the six-rotor aircraft dynamics model is:
Figure BDA0002007149510000031
preferably, the flight system inputs are:
Figure BDA0002007149510000032
preferably, the flight system inputs include an altitude control input U1, a roll control input U2, a pitch control input U3, and a yaw control input U4.
Preferably, the six-rotor aircraft dynamics model simplified in step (3) is:
Figure BDA0002007149510000041
preferably, the nominal model in the disturbance compensator is:
Figure BDA0002007149510000042
the filter in the disturbance compensator is designed to:
Figure BDA0002007149510000043
the technical scheme of the invention has the following specific technical details:
the six-rotor aircraft is simple in mechanical structure, consists of six symmetrically distributed rotors, and the included angle between every two adjacent rotors is pi/3, and the structure of the six-rotor aircraft is shown in figure 1.
Each rotor providing edge
Figure BDA0002007149510000053
The lift in the negative direction enables the six-rotor aircraft to fly upwards. In order to stabilize the flight of a six-rotor aircraft, the directions of rotation of adjacent rotors are different. The rotary wings No. 1, No. 3 and No. 5 form a group, and the rotating direction is anticlockwise; no. 2, No. 4, No. 6 rotors are in one group, and the rotating direction is clockwise. The rotation speeds of the six rotors are respectively:
Figure BDA0002007149510000051
omega in formula (1)l>0，ΩlThe negative sign in the front indicates that the direction of rotation of the rotor # i is counterclockwise.
Each rotor of the six-rotor aircraft can be controlled independently, and the operation of the rolling, pitching, yawing and other postures of the aircraft can be realized by changing the rotating speed of the rotors. Where each rotor speed change causes the vehicle to move in 3 directions with 6 degrees of freedom, the six-rotor vehicle is a highly coupled system. According to the technical scheme, the disturbance compensator is used for tracking and estimating the coupling terms between the attitude angles and external disturbance, so that decoupling and disturbance compensation of the aircraft attitude system are achieved.
And compensation can be performed by combining a cascade PID controller in order to obtain high-quality attitude control.
The core problem of the mathematical model of the six-rotor aircraft is the flight dynamics problem of the six-rotor aircraft, and since the flight speed of the six-rotor aircraft is generally not fast and the flight altitude is also in a low altitude range, the following reasonable assumptions are provided for the aircraft and the environment where the aircraft is located when the dynamics model is analyzed:
1. the six-rotor aircraft body is rigid and symmetrical;
2. the quality of the aircraft is unchanged in the flying process;
3. the origin of the body coordinates coincides with the center of gravity and the center of mass of the aircraft;
4. the lift generated by each rotor is proportional to the square of the speed of rotation of the rotor.
Before modeling, selecting a navigational coordinate system OXeYeZeAnd a body coordinate system OXbYbZbAnd aircraft attitude angle vector
Figure BDA0002007149510000052
Wherein, the roll angle phi is the included angle between the X axis of the machine body and the horizontal plane, and the defined field is [ -180 DEG ]](ii) a The pitch angle theta is the included angle between the Y axis of the machine body and the horizontal plane, and the definition domain is [ -90 DEG ]](ii) a The yaw angle psi is the included angle between the projection direction of the Z axis of the body on the horizontal plane and the parameter line on the plane, and the defined domain is [ 0-360 DEG ]]。
The relationship between the navigation coordinate system and the body coordinate system is as follows:
Figure BDA0002007149510000061
wherein the content of the first and second substances,
Figure BDA0002007149510000062
wherein the content of the first and second substances,
Figure BDA0002007149510000063
the dynamics model of the six-rotor aircraft consists of a position subsystem model and a posture subsystem model. Six-rotor aircraft in navigation coordinate system OXeYeZeThe kinetic model of (1):
Figure BDA0002007149510000064
wherein
Figure BDA0002007149510000065
Aircraft system inputs were selected as follows:
Figure BDA0002007149510000066
in the formula: u shape1Indicating a height control input, U2Indicating a roll control input, U3Indicating pitch control input, U4Representing a yaw control input.
Definition of
Ω＝Ω1-Ω2+Ω3-Ω4+Ω5-Ω6 (6)。
When the aircraft is flying slowly and smoothly in a small amplitude, equation (4) can be expressed as:
Figure BDA0002007149510000071
in order to design the control law of the aircraft, the established nonlinear model needs to be simplified as follows:
a. in the six-rotor aircraft translational dynamic equation, the air resistance is very small compared to the lift generated by the rotor, so that the air resistance (D) with a small magnitude is ignoredi(i＝x,y,z))。
b. In the six-rotor aircraft rotation dynamic equation, the coupling influence of the 3 channels of roll, pitch and yaw among each other is smaller in magnitude compared with the selection moment generated by the rotor, in order to simplify the model, the coupling part with smaller magnitude is ignored, and similarly, the gyro effect in the equation is very small compared with the rotation moment generated by the rotor, so that the gyro effect moment with smaller magnitude is ignored.
Thus, the simplified six-rotor aircraft kinetic equation is:
Figure BDA0002007149510000072
the disturbance compensation controller is applied to attitude control, coupling terms between attitude angles and external disturbance are tracked and estimated by the disturbance compensator, decoupling and disturbance compensation of an aircraft attitude system are achieved, and then high-quality attitude control of the six-rotor aircraft is achieved by combining cascade PID.
According to the six-rotor aircraft dynamic model, the equation expression of the attitude subsystem model is as follows:
Figure BDA0002007149510000081
as can be seen, equation (7) is a nonlinear MIMO coupled system, i.e., the six-rotor aircraft attitude subsystem is a multiple-input multiple-output system, as shown in fig. 2, and needs to be controlled in a decoupling manner. The disturbance compensator designed for the multiple-input multiple-output system in chapter ii is selected herein to implement the aircraft attitude decoupling control.
Definition H1＝l/Ix，H2＝l/Iy，H3＝l/Iz，fi(i ═ 1,2,3) is the coupling part of the system. The compound is obtained by arranging the formula (7):
Figure BDA0002007149510000082
in the formula:
Figure BDA0002007149510000083
the attitude subsystem of the six-rotor aircraft is a multi-input multi-output system, a disturbance compensator is used for estimating the coupling part of the attitude subsystem model in real time, and the coupling part is compensated in real time in a control law, so that the nonlinear MIMO system is converted into mutually independent linear SISO subsystems. A six-rotor aircraft attitude decoupling control block diagram 3 based on disturbance compensators is shown. The disturbance compensator is shown in fig. 4.
According to equation (9), the open-loop control transfer function of angular velocity in the control system can be obtained as:
Figure BDA0002007149510000091
it can be known that the disturbance compensator is located in the angular velocity inner loop of the control system, and the nominal model in the disturbance compensator is equivalent to the angular velocity open loop transfer function:
Figure BDA0002007149510000092
the filter in the disturbance compensator is designed as:
Figure BDA0002007149510000093
the control thinking of three attitude angles of the six-rotor aircraft is the same, so the attitude roll angle is selected
Figure BDA0002007149510000094
As a specific design object. The attitude roll angle control system based on the disturbance compensator is shown in fig. 5:
d in FIG. 51(s) is:
D1(s)＝d1(s)+f1(s) (13)
in the formula (d)1(s) is an external disturbance term, f1(s) is a coupled perturbation term.
The nominal model of the controlled object is:
Figure BDA0002007149510000095
the equivalent disturbances of the system are:
Figure BDA0002007149510000101
when H is present1(s)＝H1n(s), then the formula (15) is changed to
Figure BDA0002007149510000102
The closed loop transfer function of the roll angle inner loop control system is as follows:
Figure BDA0002007149510000103
the disturbance closed loop transfer function of the roll angle inner loop control system is as follows:
Figure BDA0002007149510000104
filter Q1(s) is a low-pass filter whose bandwidth ensures as far as possible that the system disturbances pass, i.e.:
Figure BDA0002007149510000105
from equation (18), the disturbance closed-loop transfer function Hd(s) ≈ 0, meaning that external disturbances and coupling terms of the system can be suppressed, thereby achieving attitude decoupling control of the six-rotor aircraft.
The closed loop transfer function of the roll angle outer loop control system is as follows:
Figure BDA0002007149510000106
the embodiment of the invention has the beneficial effects
The embodiment of the invention provides a six-rotor aircraft attitude decoupling control method, which treats unnecessary coupling terms in each loop as output disturbance added to the loop, designs a proper disturbance compensator to compensate the output disturbance and the external disturbance together, and solves the adverse effects caused by strong coupling and external disturbance of an aircraft system.
The disturbance observer in the embodiment of the invention can be used for decoupling control of the multi-input multi-output system, so that the disturbance suppression performance and the robust stability of the disturbance compensator are ensured, the decoupling and disturbance suppression are good, and the good tracking performance and the stability of the six-rotor aircraft can be ensured.
Drawings
Figure 1 is a schematic view of a six-rotor aircraft structure and its spatial coordinate system.
Figure 2 is a block diagram of a six-rotor aircraft attitude angle control system based on a disturbance compensation controller.
FIG. 3 is a block diagram of a disturbance compensator-based attitude decoupling control for a six-rotor aircraft.
FIG. 4 is a schematic diagram of a disturbance compensator.
FIG. 5 is a diagram of a roll angle attitude control system based on a disturbance compensator.
FIG. 6 is a simulation diagram of attitude control based on a cascade PID.
FIG. 7 is a simulation diagram of attitude decoupling control based on a disturbance compensator.
Figure 8 is a diagram of a simulation model of a six-rotor aircraft.
FIG. 9 is a diagram of a simulation result of attitude decoupling control based on a disturbance compensator.
FIG. 10 is a schematic diagram of attitude control simulation results based on a cascade PID.
FIG. 11 is a schematic diagram of a simulation result of attitude decoupling control based on a disturbance compensator under step disturbance.
FIG. 12 is a schematic diagram of attitude control simulation results based on cascade PID under step disturbance.
FIG. 13 is a schematic diagram of a simulation result of attitude decoupling control based on a disturbance compensator under periodic disturbance.
FIG. 14 is a schematic diagram of attitude control simulation results based on cascade PID under periodic disturbance.
FIG. 15 is a schematic diagram of a disturbance compensator-based attitude control system in a roll angle tracking experiment under hovering.
FIG. 16 is a schematic diagram of a cascade PID-based attitude control system in a roll angle tracking experiment under hovering.
FIG. 17 is a schematic diagram of a disturbance compensator-based attitude control system in a roll angle step response and disturbance experiment.
FIG. 18 is a schematic diagram of a cascade PID-based attitude control system in a roll angle step response and disturbance experiment.
FIG. 19 is a schematic diagram of a disturbance compensator based attitude control system in a hovering pitch angle tracking experiment.
FIG. 20 is a schematic diagram of a cascade PID-based attitude control system in a hovering pitch angle tracking experiment.
FIG. 21 is a schematic diagram of a disturbance compensator-based attitude control system in a step response and disturbance experiment of a pitch angle.
FIG. 22 is a schematic diagram of a cascade PID-based attitude control system in a pitch angle step response and disturbance experiment.
Detailed Description
The following are specific examples of the present invention, and the technical solutions of the present invention will be further described with reference to the examples, but the present invention is not limited to the examples.
Example 1
The embodiment provides a six-rotor aircraft attitude decoupling control method, which comprises the following steps:
(1) selecting a navigational coordinate system OXeYeZeBody coordinate system OXbYbZbAnd aircraft attitude angle vector
Figure BDA0002007149510000111
Establishing a six-rotor aircraft dynamic model;
(2) selecting flight system input, neglecting air resistance, three-channel coupling and gyro effective stress moment, and simplifying an aircraft dynamics model;
(3) considering the external disturbance term and the coupling disturbance term in each loop as output disturbance added on the loop, and designing a disturbance compensator;
(4) and compensating by using a disturbance compensator to realize decoupling and disturbance compensation of the aircraft attitude system.
In the step (2), air resistance, three-channel coupling and gyro effective stress moment are ignored.
And (4) compensating by using a disturbance compensator, namely tracking and estimating a coupling term between attitude angles and external disturbance by using the disturbance compensator and compensating.
Preferably, the compensating of step (4) further comprises using a cascade PID controller in combination.
Preferably, the roll angle
Figure BDA0002007149510000122
Is the included angle between the X axis of the machine body and the horizontal plane, and the defined domain is [ -180 DEG to 180 DEG ]](ii) a The pitch angle theta is the included angle between the Y axis of the machine body and the horizontal plane, and the definition domain is [ -90 DEG ]](ii) a The yaw angle psi is the included angle between the projection direction of the Z axis of the body on the horizontal plane and the parameter line on the plane, and the defined domain is [ 0-360 DEG ]]。
Preferably, the six-rotor aircraft dynamics model is:
Figure BDA0002007149510000121
preferably, the flight system inputs are:
Figure BDA0002007149510000131
preferably, the flight system inputs include an altitude control input U1, a roll control input U2, a pitch control input U3, and a yaw control input U4.
Preferably, the six-rotor aircraft dynamics model simplified in step (3) is:
Figure BDA0002007149510000132
preferably, the nominal model in the disturbance compensator is:
Figure BDA0002007149510000133
the filter in the disturbance compensator is designed to:
Figure BDA0002007149510000141
example 2
The example compares the cascade PID and disturbance compensator based aircraft attitude control through simulation experiments.
Simulation test:
in Simulink software, attitude control simulation block diagrams of a six-rotor aircraft based on cascade PID and attitude decoupling control simulation block diagrams based on a disturbance compensator are respectively set up, and the simulation block diagrams are shown in fig. 6 and 7.
Attitude control simulation parameters based on cascade PID: the parameter value of the inner ring PID controller of the roll angle and pitch angle channel is Kp＝6，Ki＝2，KpThe value of the outer ring proportional controller is K-5; the parameter value of an inner ring PID controller of a yaw angle channel is Kp＝2.7，Ki＝1，KpThe value of the outer ring proportional controller is K10 which is 0.01.
Attitude decoupling control simulation parameters based on the disturbance compensator: the filter parameter values of the roll angle, the pitch angle and the yaw angle are Q(s) < 1/(0.001s +1), and the parameter values of the inner ring PID controller and the outer ring proportional controller are the same as above.
(1) System response without external disturbance
In order to verify the decoupling effect of the attitude decoupling control method based on the disturbance compensator, the attitude control simulation test of the six-rotor aircraft is carried out without external disturbance. A simulation model diagram of a six-rotor aircraft is shown in fig. 8. In the simulation process, the height holding value of the six-rotor aircraft is set to be 1m, and the attitude angle
Figure BDA0002007149510000142
Is (0 °,0 °,0 °). When t is equal toAt 2s, the roll angle command is
Figure BDA0002007149510000143
When t is 5s, the pitch angle command is thetarEqual to 9 °; when t is 7s, the yaw angle command is psirSin (pi t) (unit (°)). The simulation time was 10 s.
The attitude control simulation results of the six-rotor aircraft are shown in fig. 9 and 10. FIG. 9 is a simulation curve of attitude decoupling control based on a disturbance compensator. When t is 2s, the roll angle command is
Figure BDA0002007149510000144
The pitch angle and yaw angle remain almost zero and are not affected. Similarly, when t is 5s, the pitch angle command is θrThe roll angle and the yaw angle are not affected by the coupling action and are still respectively kept at 10 degrees and 0 degrees; when t is 7s, the yaw angle command is psirSin (tt), roll and pitch angles remain unaffected by the coupling. FIG. 10 is a simulation plot of attitude control based on a cascade PID. When t is 5s, the pitch angle command is thetarThe roll angle and the yaw angle are affected by the coupling of the pitch angle and fluctuate as 9 degrees. Similarly, when t is 7s, the yaw angle command is ψrSin (tt), the roll and pitch angles are also significantly affected by the coupling of the yaw angle, resulting in large fluctuations.
In conclusion, the control method based on the disturbance compensator, which is provided by the invention, can effectively realize attitude decoupling control of the six-rotor aircraft and eliminate the coupling effect between attitude angles.
(2) System response under external disturbances
In order to verify the disturbance suppression effect of the attitude decoupling control method based on the disturbance compensator, external disturbance is added in the attitude control simulation test. In the simulation process, the height holding value of the six-rotor aircraft is set to be 1m, and the attitude angle
Figure BDA0002007149510000151
Is (0 °,0 °,0 °). When t is 2s, the roll angle command is
Figure BDA0002007149510000152
When t is 5s, the pitch angle command is thetarEqual to 9 °; the yaw angle command is psi r0 deg.. The simulation time was 20 s.
The external disturbance can be classified into a step disturbance and a periodic disturbance, and therefore, in the attitude control simulation test 2, simulation under the two conditions of the step disturbance and the periodic disturbance is performed respectively.
a. Step disturbance:
when t is 10s, adding step disturbance d to the pitch angle channelθ＝2π/9；
When t is 15s, adding step disturbance d to the yaw angle channelψ＝π/5。
And (3) simulation result analysis:
fig. 11 is a simulation curve of attitude decoupling control of a six-rotor aircraft based on a disturbance compensator, and it can be seen that when t is 10s, the pitch angle has slight fluctuation under the action of external step disturbance, and the roll angle and the pitch angle are not affected; similarly, when t is 15s, the yaw angle is affected by the step disturbance, and a small amount of change is generated, while the roll angle and the pitch angle are not changed.
Fig. 12 is a simulation curve of attitude control of a six-rotor aircraft based on cascade PID, and it can be seen from the graph that when t is 10s, the pitch angle fluctuates under the influence of step disturbance, and the roll angle and the yaw angle are also affected by coupling. Similarly, when t is 15s, the yaw angle fluctuates due to the step disturbance, and the roll angle and the pitch angle fluctuate.
Therefore, the attitude decoupling control method based on the disturbance compensator can estimate the externally-added step disturbance in real time, and adds corresponding compensation in the control input quantity, so that disturbance suppression of the six-rotor aircraft attitude system is realized. The comparison of simulation results can obtain: the control method based on the disturbance compensator is superior to the attitude control method based on the cascade PID in disturbance suppression performance.
b. Periodic disturbance:
when 10s<t<At 13s, pitch channel is addedIn-period disturbance dθ＝0.5πsin(2πt)；
When 15s<t<At 18s, adding periodic disturbance d into the yaw angle channelψ＝0.25πsin(πt)。
And (3) simulation result analysis:
fig. 13 shows that the attitude decoupling control method based on the disturbance compensator provided herein can effectively suppress periodic disturbance and realize accurate control.
FIG. 14 shows that the attitude control method based on cascade PID is greatly influenced by periodic disturbance, and coupling action exists between attitude angles.
Compared with simulation results, the attitude decoupling control method based on the disturbance compensator provided by the invention has better periodic disturbance inhibition performance than the attitude control method based on the cascade PID.
Example 3
In the embodiment, an attitude control experiment is performed on a self-built six-rotor aircraft platform so as to further verify the disturbance suppression performance of the attitude decoupling control method of the six-rotor aircraft in the embodiment 1 of the invention. In the experiment, three attitude angles of the six-rotor aircraft need to be controlled independently, so that a single-channel mooring experiment method is adopted. The single-channel mooring experiment mainly aims to carry out the angle control experiment on the remaining one channel on the premise of limiting the freedom degrees of other channels.
The attitude roll (pitch) angle control experiment of the six-rotor aircraft is carried out on the built experiment platform, attitude control methods based on a disturbance compensator and a cascade PID are respectively adopted, the attitude of the aircraft is controlled by using a remote controller in the experiment, attitude information is measured by a gyroscope and an accelerometer, is transmitted to an upper computer in real time through a serial port, and finally MATLAB is used for processing experiment data. Wherein:
(1) roll angle control experiment
In the hovering experiment, the roll angle command is 0 degrees, the results of two groups of comparison experiments are shown in fig. 15 and fig. 16, the steady-state error of the control system based on the disturbance compensator is within +/-2 degrees, the control system is smaller than the attitude control based on the cascade PID, and the steady-state precision is higher.
In the disturbance suppression experiment, after the aircraft is stabilized at a horizontal position, a roll angle instruction of approximately 0 degree is made through a remote controller, and external disturbance in the roll angle direction is artificially increased to deviate from a steady-state value in 23 seconds. Fig. 17 and 18 show that the disturbance compensator-based control system rapidly returns to the home position under the action of external force disturbance, so that the six-rotor aircraft can effectively resist the disturbance of wind power when flying outdoors. In addition, the simulation results are analyzed by adopting the control performance indexes ISE, IAE and ITAE, and the data in the table 1 shows that the control method provided by the invention is superior to the attitude control based on the cascade PID in performance.
TABLE 1 Performance index of control method
Figure BDA0002007149510000171
(2) Pitch angle control experiment
In the hovering experiment, the pitch angle command is 0 °, the experiment result is shown in fig. 19 and 20, the attitude control steady-state error based on the disturbance compensator is within ± 1 degree, and the steady-state precision is higher than that of the method based on the cascade PID.
In a step response experiment, after the aircraft is stabilized at a horizontal position, a pitch angle instruction of approximately 0 degree is made through a remote controller, and at 20s, external disturbance in the pitch angle direction is artificially increased to enable the aircraft to deviate from a steady state value. Fig. 21 and 22 show that the disturbance compensator-based control system can suppress external disturbances and restore the aircraft to a stable attitude. As can be seen from table 2, the control method proposed herein is slightly superior in performance index to the cascade PID based attitude control.
TABLE 2 Performance index of control method
Figure BDA0002007149510000172

Claims (5)

1. A six-rotor aircraft attitude decoupling control method is characterized by comprising the following steps:
(1) selecting a navigational coordinate system OXeYeZeBody coordinate system OXbYbZbAnd aircraft attitude angle vector
Figure FDA0002931447610000012
Establishing a six-rotor aircraft dynamics model, wherein,
Figure FDA0002931447610000013
roll angle, theta pitch angle and psi yaw angle;
(2) selecting flight system input, and simplifying an aircraft dynamics model;
(3) considering the external disturbance term and the coupling disturbance term in each loop as output disturbance added on the loop, and designing a disturbance compensator;
(4) compensating by using a disturbance compensator to realize decoupling and disturbance compensation of an aircraft attitude system;
the six-rotor aircraft dynamics model simplified in the step (3) is as follows:
Figure FDA0002931447610000011
u1 is height control input, U2 is roll control input, U3 is pitch control input, U4 is yaw control input, x is displacement in the x direction, y is displacement in the y direction, z is displacement in the z direction, l is the distance from the motor to the center of gravity, IxIs the moment of inertia of the aircraft about the x-axis, IyIs the moment of inertia of the aircraft about the y-axis, IzIs the moment of inertia, omega, of the aircraft about the z-axisiThe rotating speed of the ith motor is defined as m, the mass of the machine body is defined as g, the gravity acceleration is defined as b, the rolling torque coefficient is defined as d, and the reaction torque coefficient is defined as d;
in the disturbance compensator, a nominal model is as follows:
Figure FDA0002931447610000021
H1nas a nominal model of the rolling path, H2nFor a nominal model of the pitch channel, H3nA nominal model of the yaw channel;
in the disturbance compensator, the filter is designed as follows:
Figure FDA0002931447610000022
Q1filter transfer function, Q, for roll channel2Filter transfer function, Q, for the pitch channel3As a filter transfer function of the yaw path, τ1Is the sampling frequency, tau, of the cross-track filter2For the sampling frequency, tau, of the pitch channel filter3Is the filter sampling frequency of the yaw channel and s is the laplacian operator.
2. The hexa-rotor aircraft attitude decoupling control method of claim 1, wherein the compensating of step (4) further comprises using a cascaded PID controller in combination.
3. The hexa-rotor aircraft attitude decoupling control method of claim 1, wherein roll angle is
Figure FDA0002931447610000023
Is the included angle between the X axis of the machine body and the horizontal plane, and the defined domain is [ -180 DEG to 180 DEG ]](ii) a The pitch angle theta is the included angle between the Y axis of the machine body and the horizontal plane, and the definition domain is [ -90 DEG ]](ii) a The yaw angle psi is the included angle between the projection direction of the Z axis of the body on the horizontal plane and the parameter line on the plane, and the defined domain is [ 0-360 DEG ]]。
4. The six-rotor aircraft attitude decoupling control method of claim 1, wherein the six-rotor aircraft dynamics model is:
Figure FDA0002931447610000031
Ftis the sum of the lifting forces, DxAir resistance in the x-direction, DyAir resistance in the y direction, DzIs the air resistance in the z direction, τxFor roll moment, τyFor pitching moment, τzFor yaw moment, τgyroThe gyro effective stress moment is represented by p, the roll angular velocity, the pitch angular velocity, the yaw angular velocity and the Euler transformation matrix.
5. The hexa-rotor aircraft attitude decoupling control method of claim 1, wherein the flight system inputs are:
Figure FDA0002931447610000032
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