CN112666960A - L1-based augmented self-adaptive rotor craft control method - Google Patents

L1-based augmented self-adaptive rotor craft control method Download PDF

Info

Publication number
CN112666960A
CN112666960A CN202011466461.XA CN202011466461A CN112666960A CN 112666960 A CN112666960 A CN 112666960A CN 202011466461 A CN202011466461 A CN 202011466461A CN 112666960 A CN112666960 A CN 112666960A
Authority
CN
China
Prior art keywords
control
rate
representing
angular rate
loop
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
CN202011466461.XA
Other languages
Chinese (zh)
Other versions
CN112666960B (en
Inventor
程昊宇
常晓飞
符文星
黄汉桥
付斌
张通
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Xi'an Innno Aviation Technology Co ltd
Northwestern Polytechnical University
Original Assignee
Xi'an Innno Aviation Technology Co ltd
Northwestern Polytechnical University
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Xi'an Innno Aviation Technology Co ltd, Northwestern Polytechnical University filed Critical Xi'an Innno Aviation Technology Co ltd
Priority to CN202011466461.XA priority Critical patent/CN112666960B/en
Publication of CN112666960A publication Critical patent/CN112666960A/en
Application granted granted Critical
Publication of CN112666960B publication Critical patent/CN112666960B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation

Landscapes

  • Feedback Control In General (AREA)

Abstract

The invention relates to a rotorcraft control method based on L1 augmentation self-adaptation, which comprises the steps of firstly establishing a rotorcraft kinematics and dynamics model, designing cascade PID control rate and L1 augmentation self-adaptation control rate, and obtaining the total control input of the rotorcraft. The method can estimate the interference in real time on line, and can effectively realize the stable control of the rotor wing aircraft and the stable control of large disturbance by matching with the traditional cascade PID control. Practical engineering application shows that the method can be easily realized in engineering by combining L1-based augmentation self-adaptation with traditional cascade PID control, and compared with the traditional cascade PID control, the method has the advantages of better robustness, stronger disturbance resistance and better self-adaptation effect.

Description

L1-based augmented self-adaptive rotor craft control method
Technical Field
The invention belongs to the technical field of unmanned aerial vehicle application, relates to a control method of an unmanned aerial vehicle, and particularly relates to a control method of a rotorcraft based on L1 augmentation self-adaption.
Background
With the development of scientific technology, the rotary wing aircraft has been applied to many military fields, such as the investigation of battlefield environment, the striking of small targets and the like. In the civil field, the rotor craft plays an important role in aerial photography, security, power inspection, oil and gas pipeline inspection and the like. The rotary-wing aircraft can achieve the expected position and attitude by controlling the rotating speed of the rotary-wing motor in the actual flying process, and the control of the rotary-wing aircraft is the core of the rotary-wing aircraft capable of achieving the flying, so the control is particularly important. The control quality of the rotor craft is directly related to the completion degree of the operation task, and is the core technical foundation of the operation. Because rotor craft's needs carry out the operation flow under the operating mode of complicacy, need rotor craft can adapt to different operating modes to can both have fine control quality under various different operating modes. Traditional PID control, because need not be based on the model, have extensive application, but because rotor flight needs to be in complicated abominable environment operation, traditional PID control is difficult to adapt to complicated operating mode environment, for example strong wind weather, or the condition such as sudden release of load, traditional PID control is under this kind of environment, it is difficult to have fine control quality or a set of controller parameter is difficult to adapt to a plurality of operating modes, this paper proposes traditional PID + L1 augmentative adaptive control's control rate design structure, can adapt to complicated operating mode, and need not set multiunit parameter, all have fine control quality under each complicated operating mode. The cascade PID + L1 augmented adaptive control structure is provided to enhance the adaptivity and robustness of a PID system so as to adapt to different working conditions, and the structure can adapt to the working conditions wider than a single cascade PID, does not need to re-adjust PID parameters to adapt to different working conditions like the single cascade PID, and can obtain better control quality.
Disclosure of Invention
Technical problem to be solved
In order to avoid the defects of the prior art, the invention provides a rotorcraft control method based on L1 augmentation self-adaption.
Technical scheme
A rotorcraft control method based on L1 augmented adaptive is characterized by comprising the following steps:
step 1, establishing a rotor craft kinematics and dynamics model:
rotorcraft kinematics model:
Figure BDA0002832663990000021
wherein: m represents the weight of the unmanned aerial vehicle, g represents the acceleration of gravity, U1Denotes lift, KdThe coefficient of resistance is expressed as a function of,
Figure BDA0002832663990000022
which represents the speed in three directions and,
Figure BDA0002832663990000023
three directional accelerations are shown.
Figure BDA0002832663990000024
The machine system is connected with the navigation system rotation matrix;
Figure BDA0002832663990000025
wherein: ψ represents a heading angle, θ represents a pitch angle, γ represents a roll angle,
Figure BDA0002832663990000026
a rotation matrix representing the machine system to the navigation system;
kinetic model of rotorcraft:
Figure BDA0002832663990000031
Figure BDA0002832663990000032
the derivative of the attitude angle is represented. cos denotes the cosine, sin denotes the sine,
Figure BDA0002832663990000033
Figure BDA0002832663990000034
Figure BDA0002832663990000035
Figure BDA0002832663990000036
Figure BDA0002832663990000037
the rate of acceleration of the three-axis angle,
Figure BDA0002832663990000038
representing the three-axis angular rate. I isxx,Iyy,IzzRepresenting the three-axis inertia, and L representing the distance from the center of the rotor to the center of mass of the aircraft; u shape1Representing lift, U2Torsion, U, representing the direction of roll3Torsion indicating pitch direction, U4A torque force representing a heading direction;
step 2, designing cascade PID control rate:
designing the control rate of a horizontal channel position loop:
vt＝Kp·(pt-p)
wherein: p is a radical oftIs a desired position; p is the current position, vtTo the desired speed, KpIs the position loop proportional gain;
designing the control rate of a horizontal channel speed loop:
at＝Kv·(vt-v)
wherein: v. oftA desired speed; v is the current speed, atFor desired acceleration,. psi.vProportional gain for the velocity loop;
Figure BDA0002832663990000041
θt＝arctan(-ax/g)
γt＝arctan(ay·cosθ/g)
wherein: thetatIs a desired pose; theta is the current attitude, ay，axRepresents a horizontal acceleration;
designing the control rate of an angle ring:
ωt＝Kθ·(αt-α)
ωtis the desired angular rate; ω is the current angular rate, KθFor attitude ring proportional gain, α represents the tilt angle, i.e., roll and pitch;
design of angular rate loop control rate:
Δω＝ωt-ω
Figure BDA0002832663990000042
wherein: kωFor angular rate loop proportional gain, KIFor angular rate loop integral gain, KDIn order to be the angular rate loop differential gain,
Figure BDA0002832663990000043
represents an integral operation, s represents a differential operation; Δ ω is the target angular rate ωtThe difference from the current angular rate co,
Figure BDA0002832663990000044
is an integral operation sign, and s is a differential operation sign;
step 3, L1 expands the adaptive control rate:
the state prediction equation of the L1 augmented adaptive is:
Figure BDA0002832663990000051
wherein: state matrix A, control matrix B, matrix ASPIs a 3 x 3 matrix with the determinant being negative, and T represents the period of controller operation;
Figure BDA0002832663990000052
wherein:
Figure BDA0002832663990000058
which is indicative of a state error,
Figure BDA0002832663990000054
representing the estimated state, x (t) representing the true state;
l1 adaptive adaptation rate:
Figure BDA0002832663990000055
output of adaptive controller:
Figure BDA0002832663990000056
wherein: the expression of C(s) is
Figure BDA0002832663990000057
And 4, inputting the total control of the rotor craft:
control output u of step 2b(t) and the output u of step 3a(t) adding to obtain the total control output u (t)
u(t)＝ub(t)+ua(t)。
Advantageous effects
According to the L1-based augmented self-adaptive rotor aircraft control method, interference can be estimated on line in real time, and the method is matched with the traditional cascade PID control, so that not only can stable control over the rotor aircraft be effectively realized, but also stable control over large disturbance can be realized. Practical engineering application shows that the method can be easily realized in engineering by combining L1-based augmentation self-adaptation with traditional cascade PID control, and compared with the traditional cascade PID control, the method has the advantages of better robustness, stronger disturbance resistance and better self-adaptation effect.
The invention has the beneficial effects that:
(1) compared with cascade PID, the anti-interference capability is stronger;
(2) compared with cascade PID, the method has stronger adaptability;
(3) compared with cascade PID, the robustness is stronger;
(4) compared with other modern control theories, the method is easier to realize in engineering, and a system model does not need to be finely identified;
the L1 augmented adaptive algorithm can enhance the adaptability and robustness of the system by combining with cascade PID, so that the control performance of the rotorcraft is greatly improved.
Drawings
FIG. 1 is a horizontal channel control structure;
FIG. 2 is a height control structure;
fig. 3L 1 expands the adaptive control architecture;
fig. 4 basic control rate + L1 augments the adaptive control structure.
Detailed Description
The invention will now be further described with reference to the following examples and drawings:
firstly, aiming at a controlled object which is a rotor aircraft, a dynamics and kinematics model of the rotor aircraft is given; in the second step, aiming at the model of the rotorcraft, a control strategy of a traditional cascade PID controller is given, and the output of an angular rate loop in the PID controller is used as a basic controller ub(output of base controller); thirdly, a concrete implementation method of the L1 controller is given, and the output of the L1 self-adaptive controller is recorded as ua(adaptive controller output); fourth, the control input for the model of the rotorcraft is utotalThe concrete expression method is as follows
utotal＝ua+ub (1)
The above steps can be written as
(1) Establishing a six-degree-of-freedom model of the rotor craft;
(2) aiming at a six-degree-of-freedom model of a rotor craft, a traditional cascade PID control strategy is provided;
(3) and aiming at a six-degree-of-freedom model of the rotor craft, a specific implementation mode of L1 augmented adaptive control is provided.
(4) The total control input of the rotorcraft is calculated by combining the control output of the cascade PID and the augmented adaptive control output of L1.
The various implementation steps are described in detail below.
Step 1, establishing a rotorcraft kinematics and dynamics model
The patent of the method only explains the specific control method of the traditional PID combined with the L1 augmentation self-adaptation, does not make complicated and fine model description on the rotorcraft system, and only adopts general system description, for example, engineering researchers want to know the system model of the rotorcraft more deeply and can refer to relevant literature data or books for further research.
Taking a four-rotor aircraft as an example, before deriving a six-degree-of-freedom model of the rotor aircraft, some variables to be used subsequently need to be subjected to variable description. The variables used subsequently are defined in
Shown in table 1.
Table 1 physical variable definitions
Figure BDA0002832663990000071
Figure BDA0002832663990000081
Establishing a rotorcraft kinematic equation
Rotorcraft are subjected mainly to several forces: gravity; a lifting force; resistance force.
According to the aerodynamic theory of rotorcraft, lift U1Three directional torsion (U)2，U3，U4) The specific calculation method of (2) is related to the rotating speed of the rotorcraft, and the expression is as follows:
Figure BDA0002832663990000082
Figure BDA0002832663990000083
Figure BDA0002832663990000084
Figure BDA0002832663990000085
the meanings of the letters in the formula are shown in Table 1. According to the stress condition of the rotor craft, applying Newton's second law to obtain the kinematic equation of the rotor craft:
Figure BDA0002832663990000091
the third term on the right of the middle sign in the above formula is a resistance term which is proportional to the speed of the aircraft and has opposite signs.
Wherein the rotation matrix
Figure BDA0002832663990000092
Expression (c):
Figure BDA0002832663990000093
ψ in equation (7) represents a heading angle, θ represents a pitch angle, and γ represents a roll angle.
Second, establish the dynamic equation of the rotor craft
The conversion relationship between the angular rate of a rotorcraft and the differential of the euler angle is as follows:
Figure BDA0002832663990000094
the following equation holds true according to the law of conservation of angular momentum:
Figure BDA0002832663990000095
the above formula is developed in detail as follows:
Figure BDA0002832663990000096
then the equations of dynamics for the rotorcraft can be derived.
Figure BDA0002832663990000101
Equations (7) and (11) are kinematic and dynamic model equations of a rotorcraft that describe translational and rotational motion of the rotorcraft in space.
Step 2, designing cascade PID control rate
For a dynamics and kinematics model of a rotorcraft, the rotorcraft can be controlled by using a traditional cascade PID control mode, and the following sections are introduced to the traditional cascade PID:
horizontal position channel control structure
The control structure of the horizontal position channel is shown in fig. 1, and the variables used in the figure are as follows:
(1)pta desired position; p represents the current position and is calculated by an airborne integrated navigation system;
(2)vta desired speed; v represents the current speed and is calculated by an airborne integrated navigation system;
(3)ata desired acceleration; a represents the current acceleration and can be calculated by an airborne integrated navigation system; see FIG. 1;
(4)θta desired pose; theta represents the current attitude and can be obtained by calculation of an airborne integrated navigation system;
(5)ωta desired angular rate; ω represents the current angular rate, which can be obtained by the onboard integrated navigation system.
Desired position ptObtaining the expected speed v through the position proportional controller by making difference with the current position pt(ii) a Desired velocity vtObtaining the expected acceleration a through a speed ratio controller by making difference with the current speed vt. The position loop and velocity loop control rates described above can be described by the following equations:
designing the control rate of a horizontal channel position loop:
vt＝Kp·(pt-p) (12)
wherein KpIs the position loop proportional gain.
Designing the control rate of a horizontal channel speed loop:
at＝Kv·(vt-v) (13)
wherein KvProportional gain for velocity loop
Desired acceleration atObtaining the expected attitude theta through theoretical calculation of small perturbation hypothesistCalculation of the desired attitude from the desired accelerationThe formula is as follows, the condition for the small disturbance assumption is that the rotorcraft has no movement in the altitude direction at equilibrium, the airframe has no movement in the heading direction and the heading angle ψ ≈ 0 is zero.
Figure BDA0002832663990000111
Is finished to obtain
Figure BDA0002832663990000112
And (5) finishing the formula (15) again to obtain a calculation formula from the acceleration to the inclination angle:
Figure BDA0002832663990000113
equation (16) is the equation from acceleration to pitch (roll and pitch). According to (16), the desired acceleration a can be passedtConversion to a desired inclination angle thetat. Desired inclination angle thetatObtaining an error angle by making a difference with the current inclination angle, and obtaining an expected angular rate omega by the error angle through an attitude proportional controllert. The attitude loop control rate design can be expressed by the following expression:
ωt＝Kθ·(αt-α) (17)
wherein KθFor the attitude ring proportional gain, α here denotes the tilt angle (roll and pitch).
Desired angular rate ωtObtaining angular output of the angular rate controller recorded as u through the angular rate PID controller after making difference with the current measured angular rateb(ubReferred to as the base controller output, which will be used in subsequent steps), the horizontal position control structure shown in fig. 1 describes how to transfer to the attitude control through the control structure thinking of the cascade PID, i.e., the outer ring is the position ring and the inner ring is the attitude ring. Thus the output u of the basic controllerbThe calculation formula of (a) is as follows:
Figure BDA0002832663990000121
wherein KωFor angular rate loop proportional gain, KIFor angular rate loop integral gain, KDIn order to be the angular rate loop differential gain,
Figure BDA0002832663990000122
denotes an integral operation, and s denotes a differential operation. Δ ω is the target angular rate ωtThe difference from the current angular rate co,
Figure BDA0002832663990000123
is the sign of the integral operation and s is the sign of the derivative operation.
Step 3, L1 augmented adaptive control rate design
A cascade PID + L1 augmented adaptive control structure is provided by combining a traditional cascade PID control structure and an L1 adaptive structure, so that the adaptivity and robustness of a PID system are enhanced, different working conditions are adapted, the structure can be adapted to the working conditions wider than a single cascade PID, and the PID parameters do not need to be re-adjusted to adapt to different working conditions like the single cascade PID.
This control method is described below in an attitude control loop:
the output is controlled by the angular velocity PID shown in fig. 1 as the basic controller control signal in fig. 3, and thus fig. 3 may become as shown in fig. 4. The controller combined by the cascaded PID control is now referred to as the base controller. For the attitude inner ring controller, the angular rate of a three-axis machine body is selected as a state variable x, and the output of an angular rate PID controller of a cascade PID is recorded as ub：
ubAngular rate PID controller output (19)
Namely, it is
Figure BDA0002832663990000131
Then the state prediction equation for the L1 augmented adaptation is
Figure BDA0002832663990000132
The parameter interpretation in the formula (21) is described in the section of L1 augmented adaptive, and when the practical engineering is used, the system matrixes A and B need to be determined by a system identification method, and the matrix ASPThe matrix is a 3 x 3 matrix with the determinant being negative, and can be adjusted in the actual process.
Figure BDA0002832663990000133
x (t) represents the true value and may be replaced with actual measured gyroscope data of the integrated navigation system.
L1 adaptive adaptation rate calculation:
Figure BDA0002832663990000134
t represents the running period of the controller, and the total uncertainty of the system can be estimated in real time when the running frequency of the adaptive control rate is higher in the actual use process
Figure BDA0002832663990000135
After obtaining the uncertainty estimate of the system, the adaptive control output u of L1 can be calculated by the following equationadI.e. the output of the adaptive controller:
Figure BDA0002832663990000136
the discrete form of C(s) actually used by the authors is:
Figure BDA0002832663990000137
the above equation is bilinear as:
Figure BDA0002832663990000138
order:
Figure BDA0002832663990000139
fcis the filter cut-off frequency, fsFor data sampling frequency, when the author actually uses fs＝1kHz，fcDifferent systems may differ slightly by 5 Hz.
Redefined and arranged as:
Figure BDA0002832663990000141
Figure BDA0002832663990000142
then use in real time
Figure BDA0002832663990000143
Performing second order low pass filtering
Figure BDA0002832663990000144
Wherein
Figure BDA0002832663990000145
The intermediate variable is wn, wn-1 is the value of wn in the previous cycle, and wn-2 is the value of wn in the previous cycle.
Step 4, calculating the total control input of the rotorcraft
The controller output calculated in step 2 and step 3 is used as the control input u (t) of the actuating mechanism of the rotorcraft, and the calculation formula is
u(t)＝ub(t)+ua(t) (32)
Wherein u isb(t) the calculation project of the output of the cascade PID angular rate controller is shown in a step 2, a formula (18), ua(t) the calculation process of the output of the augmented adaptive controller of L1 is shown in (24).
When the method is actually used, the operation cycle of the angular rate loop is 500HZ, the operation rate of the L1 self-adaptive control self-adaptive rate is 1000Hz, the higher the operation frequency of the self-adaptive controller is, the faster the estimation on the uncertainty of the system is, and the control effect is better.
The same applies to the height channel. By using the control method for the horizontal channel inner ring and the vertical direction height inner ring, the wind resistance, the posture stability and the overall stability of the rotor craft after the load is put on are greatly improved.

Claims (1)

1. A rotorcraft control method based on L1 augmented adaptive is characterized by comprising the following steps:
step 1, establishing a rotor craft kinematics and dynamics model:
rotorcraft kinematics model:
Figure FDA0002832663980000011
wherein: m represents the weight of the unmanned aerial vehicle, g represents the acceleration of gravity, U1Denotes lift, KdThe coefficient of resistance is expressed as a function of,
Figure FDA0002832663980000012
which represents the speed in three directions and,
Figure FDA0002832663980000013
three directional accelerations are shown.
Figure FDA0002832663980000014
The machine system is connected with the navigation system rotation matrix;
Figure FDA0002832663980000015
wherein: ψ represents a heading angle, θ represents a pitch angle, γ represents a roll angle,
Figure FDA0002832663980000016
a rotation matrix representing the machine system to the navigation system;
kinetic model of rotorcraft:
Figure FDA0002832663980000017
Figure FDA0002832663980000021
the derivative of the attitude angle is represented. cos denotes the cosine, sin denotes the sine,
Figure FDA0002832663980000022
Figure FDA0002832663980000023
Figure FDA0002832663980000024
Figure FDA0002832663980000025
Figure FDA0002832663980000026
the rate of acceleration of the three-axis angle,
Figure FDA0002832663980000027
representing the three-axis angular rate. I isxx,Iyy,IzzRepresenting the three-axis inertia, and L representing the distance from the center of the rotor to the center of mass of the aircraft; u shape1Representing lift, U2Torsion, U, representing the direction of roll3Torsion indicating pitch direction, U4A torque force representing a heading direction;
step 2, designing cascade PID control rate:
designing the control rate of a horizontal channel position loop:
vt＝Kp·(pt-p)
wherein: p is a radical oftIs a desired position; p is the current position, vtTo the desired speed, KpIs the position loop proportional gain;
designing the control rate of a horizontal channel speed loop:
at＝Kv·(vt-v)
wherein: v. oftA desired speed; v is the current speed, atFor desired acceleration,. psi.vProportional gain for the velocity loop;
Figure FDA0002832663980000031
θt＝arctan(-ax/g)
γt＝arctan(ay·cosθ/g)
wherein: thetatIs a desired pose; theta is the current attitude, ay，axRepresents a horizontal acceleration;
designing the control rate of an angle ring:
ωt＝Kθ·(αt-α)
ωtis the desired angular rate; ω is the current angular rate, KθFor attitude ring proportional gain, α represents the tilt angle, i.e., roll and pitch;
design of angular rate loop control rate:
Δω＝ωt-ω
Figure FDA0002832663980000032
wherein: kωFor angular rate loop proportional gain, KIFor angular rate loop integral gain, KDIn order to be the angular rate loop differential gain,
Figure FDA0002832663980000033
represents an integral operation, s represents a differential operation; Δ ω is the target angular rate ωtThe difference from the current angular rate co,
Figure FDA0002832663980000034
is an integral operation sign, and s is a differential operation sign;
step 3, L1 expands the adaptive control rate:
the state prediction equation of the L1 augmented adaptive is:
Figure FDA0002832663980000035
wherein: state matrix A, control matrix B, matrix ASPIs a 3 x 3 matrix with the determinant being negative, and T represents the period of controller operation;
Figure FDA0002832663980000041
wherein:
Figure FDA0002832663980000042
which is indicative of a state error,
Figure FDA0002832663980000043
representing the estimated state, x (t) representing the true state;
l1 adaptive adaptation rate:
Figure FDA0002832663980000044
output of adaptive controller:
Figure FDA0002832663980000045
wherein: the expression of C(s) is
Figure FDA0002832663980000046
And 4, inputting the total control of the rotor craft:
control output u of step 2b(t) and the output u of step 3a(t) adding to obtain the total control output u (t)
u(t)＝ub(t)+ua(t)。
CN202011466461.XA 2020-12-12 2020-12-12 L1 augmentation self-adaption-based control method for rotary wing aircraft Active CN112666960B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202011466461.XA CN112666960B (en) 2020-12-12 2020-12-12 L1 augmentation self-adaption-based control method for rotary wing aircraft

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202011466461.XA CN112666960B (en) 2020-12-12 2020-12-12 L1 augmentation self-adaption-based control method for rotary wing aircraft

Publications (2)

Publication Number Publication Date
CN112666960A true CN112666960A (en) 2021-04-16
CN112666960B CN112666960B (en) 2023-06-23

Family

ID=75405691

Family Applications (1)

Application Number Title Priority Date Filing Date
CN202011466461.XA Active CN112666960B (en) 2020-12-12 2020-12-12 L1 augmentation self-adaption-based control method for rotary wing aircraft

Country Status (1)

Country Link
CN (1) CN112666960B (en)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114265435A (en) * 2021-12-27 2022-04-01 海兴东方新能源发电有限公司 Method, system and device for realizing accurate landing of rotor unmanned aerial vehicle in multi-airport
CN114355963A (en) * 2021-12-27 2022-04-15 海兴东方新能源发电有限公司 Method, system, terminal and medium for realizing height control based on rotor unmanned aerial vehicle
CN115236972A (en) * 2022-07-25 2022-10-25 中国安全生产科学研究院 Multi-robot formation control method based on double closed-loop self-adaptive PID

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN106292297A (en) * 2016-10-26 2017-01-04 成都市优艾维机器人科技有限公司 Based on PID controller and the attitude control method of L1 adaptive controller
CN106444826A (en) * 2016-09-07 2017-02-22 广西师范大学 Flight control method of QUAV (Quadrotor Unmanned Aerial Vehicle)
WO2017078823A2 (en) * 2015-08-13 2017-05-11 The Board Of Regents For Oklahoma State University Modular autopilot design and development featuring bayesian non-parametric adaptive control
CN106933104A (en) * 2017-04-21 2017-07-07 苏州工业职业技术学院 A kind of quadrotor attitude based on DIC PID and the mixing control method of position
CN110456636A (en) * 2019-07-11 2019-11-15 西北工业大学 Aircraft discrete sliding mode self-adaptation control method based on upper bound estimation
CN111487868A (en) * 2020-04-23 2020-08-04 中国空气动力研究与发展中心设备设计及测试技术研究所 Adaptive control system and method suitable for integral feedback amplification system L1

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2017078823A2 (en) * 2015-08-13 2017-05-11 The Board Of Regents For Oklahoma State University Modular autopilot design and development featuring bayesian non-parametric adaptive control
CN106444826A (en) * 2016-09-07 2017-02-22 广西师范大学 Flight control method of QUAV (Quadrotor Unmanned Aerial Vehicle)
CN106292297A (en) * 2016-10-26 2017-01-04 成都市优艾维机器人科技有限公司 Based on PID controller and the attitude control method of L1 adaptive controller
CN106933104A (en) * 2017-04-21 2017-07-07 苏州工业职业技术学院 A kind of quadrotor attitude based on DIC PID and the mixing control method of position
CN110456636A (en) * 2019-07-11 2019-11-15 西北工业大学 Aircraft discrete sliding mode self-adaptation control method based on upper bound estimation
CN111487868A (en) * 2020-04-23 2020-08-04 中国空气动力研究与发展中心设备设计及测试技术研究所 Adaptive control system and method suitable for integral feedback amplification system L1

Non-Patent Citations (4)

* Cited by examiner, † Cited by third party
Title
"Asynchronously finite-time HN control for morphing aircraft" *
吕蓉蓉;徐亮;薛辰;陆宇平;: "阵风扰动下大柔性飞行器姿态跟踪控制设计" *
王易南;陈康;符文星;闫杰;: "带有攻角约束的高超声速飞行器航迹倾角跟踪控制方法" *
薛静等: "基于L1自适应控制的无人机横侧向控制", 《西北工业大学学报》 *

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114265435A (en) * 2021-12-27 2022-04-01 海兴东方新能源发电有限公司 Method, system and device for realizing accurate landing of rotor unmanned aerial vehicle in multi-airport
CN114355963A (en) * 2021-12-27 2022-04-15 海兴东方新能源发电有限公司 Method, system, terminal and medium for realizing height control based on rotor unmanned aerial vehicle
CN115236972A (en) * 2022-07-25 2022-10-25 中国安全生产科学研究院 Multi-robot formation control method based on double closed-loop self-adaptive PID

Also Published As

Publication number Publication date
CN112666960B (en) 2023-06-23

Similar Documents

Publication Publication Date Title
CN111766899B (en) Interference observer-based quad-rotor unmanned aerial vehicle cluster anti-interference formation control method
CN112666960A (en) L1-based augmented self-adaptive rotor craft control method
Hamel et al. Dynamic modelling and configuration stabilization for an X4-flyer.
Luukkonen Modelling and control of quadcopter
Ahmed et al. Flight control of a rotary wing UAV using backstepping
CN112346470A (en) Four-rotor attitude control method based on improved active disturbance rejection control
CN110119089B (en) Immersion constant flow pattern self-adaptive quad-rotor control method based on integral sliding mode
CN102830622A (en) Auto-disturbance-rejection automatic flight control method for four-rotor aircraft
CN109885074B (en) Finite time convergence attitude control method for quad-rotor unmanned aerial vehicle
Suzuki et al. Attitude control of quad rotors QTW-UAV with tilt wing mechanism
CN110254703B (en) Tilting double-rotor wing automatic hovering T-shaped unmanned aerial vehicle system
CN111650951B (en) Full-loop composite dynamic inverse tracking control method for complex track of quad-rotor unmanned aerial vehicle
Tran Modeling and control of a quadrotor in a wind field
Yan et al. Robust adaptive backstepping control for unmanned autonomous helicopter with flapping dynamics
CN107678442B (en) Dual-model-based four-rotor autonomous landing control method
Qingtong et al. Backstepping-based attitude control for a quadrotor UAV using nonlinear disturbance observer
CN113031655B (en) Flying and lifting system accompanied with load lifting and control method thereof
Manai et al. Identification of a UAV and design of a hardware-in-the-loop system for nonlinear control purposes
CN108279562A (en) A kind of flight mechanical arm based on sliding formwork PID control
Alcan et al. Robust hovering control of a quadrotor using acceleration feedback
CN115343949B (en) Fixed wing unmanned plane tracking guidance law design method and verification platform
CN111857171A (en) Unmanned aerial vehicle control method based on quality self-adaptive multilayer neural network
Kumar et al. Design and implementation of fuzzy logic controller for quad rotor UAV
CN116679548A (en) Three-degree-of-freedom helicopter robust output feedback control method based on time-varying observer
Song et al. Research on attitude control of quadrotor uav based on active disturbance rejection control

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
GR01 Patent grant
GR01 Patent grant