CN113759718A - Self-adaptive control method for airplane wing damage - Google Patents

Self-adaptive control method for airplane wing damage Download PDF

Info

Publication number
CN113759718A
CN113759718A CN202110962610.XA CN202110962610A CN113759718A CN 113759718 A CN113759718 A CN 113759718A CN 202110962610 A CN202110962610 A CN 202110962610A CN 113759718 A CN113759718 A CN 113759718A
Authority
CN
China
Prior art keywords
control
aircraft
airplane
determining
self
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
CN202110962610.XA
Other languages
Chinese (zh)
Other versions
CN113759718B (en
Inventor
卢正人
李佳
牛尔卓
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.)
Xian Flight Automatic Control Research Institute of AVIC
Original Assignee
Xian Flight Automatic Control Research Institute of AVIC
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 Xian Flight Automatic Control Research Institute of AVIC filed Critical Xian Flight Automatic Control Research Institute of AVIC
Priority to CN202110962610.XA priority Critical patent/CN113759718B/en
Publication of CN113759718A publication Critical patent/CN113759718A/en
Application granted granted Critical
Publication of CN113759718B publication Critical patent/CN113759718B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • 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

Landscapes

  • Engineering & Computer Science (AREA)
  • Health & Medical Sciences (AREA)
  • Artificial Intelligence (AREA)
  • Computer Vision & Pattern Recognition (AREA)
  • Evolutionary Computation (AREA)
  • Medical Informatics (AREA)
  • Software Systems (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Automation & Control Theory (AREA)
  • Feedback Control In General (AREA)
  • Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)

Abstract

The invention relates to the technical field of aviation, in particular to a self-adaptive control method for airplane wing damage. The method comprises the following steps: receiving a triaxial angular velocity signal, and determining a triaxial approximate angular acceleration signal according to a tracking differentiator; determining a reference model of the airplane according to the aerodynamic and control effect data, and determining a reference angular acceleration by taking the airplane state and the control surface deflection as input reference models; receiving the determined approximate angular acceleration and the determined reference angular acceleration, and determining the advanced control quantity of the airplane control surface instruction according to an advanced control module; receiving an aircraft attitude control instruction and a dynamics related state, determining an L1 self-adaptive control law and determining an L1 self-adaptive control quantity of an aircraft control surface instruction; and determining a control surface skewness instruction according to the determined advance control quantity and the determined L1 self-adaptive control quantity. The disturbance torque of the airplane is quickly compensated through advanced control, the airplane transient state after wing damage is restrained, meanwhile, the L1 self-adaptive control is adopted to carry out self-adaptive compensation on the residual disturbance, and the stable and accurate control of the airplane is realized.

Description

Self-adaptive control method for airplane wing damage
Technical Field
The invention relates to the technical field of aviation, in particular to an adaptive control method for airplane wing damage.
Background
After a unilateral wing is damaged due to an emergency in the flight of the airplane, the aerodynamic force and the moment are greatly changed in a short time, the gravity center of the weight of the airplane is deviated, the flight dynamics equation is changed violently, and the stability and the self-adaptive control of the airplane are difficult to realize by the conventional control method. In the prior art, a passive fault-tolerant control method based on a neural network adaptive robust nonlinear model inverse exists. The method realizes the stability of the airplane after the control surface of the airplane is blocked or the area of the single-side wing is damaged by 40 percent, but due to the limitation of passive fault-tolerant control, the airplane attitude transient state is larger at the moment of wing damage, and the control performance needs to be improved. The active fault-tolerant control is based on means such as online observation and isolation of unknown faults or disturbances, and the control gain or the control structure is adjusted after the faults occur. For sudden damage faults, the suppression performance of fault transients is degraded by the time delay of online observation, and the high gain caused by the fast adaptation of the observer causes high-frequency oscillation of the aircraft.
Disclosure of Invention
The purpose of the invention is as follows: the self-adaptive control method for the damage of the wings of the airplane is provided to reduce the transient state of the airplane at the moment of damage of the wings, inhibit high-frequency oscillation and improve the control performance.
The technical scheme of the invention is as follows:
in a first aspect, an adaptive control method for aircraft wing damage is provided, which includes: step S1: receiving a three-axis angular velocity signal p of the airplane, and carrying out differential control according to a tracking differentiator to determine an approximate three-axis angular acceleration signal epsilon of the airplane; step S2: determining a reference model of the airplane according to the existing aerodynamic and control effect data of the airplane, and setting the state tau of the airplane and the deflection of a control plane as usDetermining a reference angular acceleration epsilon of the aircraft by inputting a reference model of the aircraftc(ii) a Step S3: receiving the approximate angular acceleration determined in the S1 and the reference angular acceleration determined in the S2, and determining an advanced control quantity u of the airplane control surface instruction according to the advanced control module for advanced controll(ii) a Step S4: receiving an aircraft attitude control command and an aircraft dynamics related state, determining an L1 self-adaptive control law, and performing self-adaptive control according to the L1 self-adaptive control law to determine an L1 self-adaptive control quantity u of an aircraft control surface commanda(ii) a Step S5: according to the advance control amount u determined in S3lAnd L1 adaptive control amount u determined in S4aDetermining control plane skewness instruction us。
Further, a tracking differentiator, in particular
Figure RE-GDA0003347170290000021
Wherein z is1And z2To track the two states generated by the differentiator, z1(k) Z for k sampling instants1Value z2(k) Z for k sampling instants2Value z1And z2Is 0, h is the sampling step length of the computer, epsilon (k) is the epsilon value of k sampling moments, r0Is a tracking factor, h, of the fhan function0As step size of fhan function, fhan (z)1(k)-p(k),z2(k),r0,h0) Comprises the following steps:
Figure RE-GDA0003347170290000022
sign is a sign function, μ1，μ2，Sz1，μ，SμTo calculate the intermediate variables involved in the fhan function.
Further, the reference model is I.epsilonc＝M1(τ)+M2(τ,us) Wherein I is the moment of inertia of the aircraft, M1 (tau) is the aircraft moment related to the aircraft state tau when the control plane deflection is 0, and M2 (tau, us) The deflection of the control surface is usMoment of flight, epsilon, induced by the flightcIs the output of the reference model.
Further, a control plane deflection instruction usThe initial value of (a) is trim rudder deflection in the current state of the aircraft.
Further, the advance control amount is calculated as follows:
Figure RE-GDA0003347170290000023
wherein k islControl gain for advanced control, T1(ud) Represents a pair udWith an ongoing time constant of T1To first order smoothing filtering.
Further, determining an L1 adaptive control law specifically includes: determining an aircraft state observer and estimating the residual disturbance of the aircraft, wherein the state observer has a calculation formula of
Figure RE-GDA0003347170290000031
Figure RE-GDA0003347170290000032
For observing the state of the aircraft, AoSystem matrix relating to desired aircraft movement modal characteristics, BoIs a control effect matrix of the aircraft control surfaces, uaThe control plane instruction output by the L1 self-adaptive control law has an initial value of trim rudder deflection under the current state of the airplane,
Figure RE-GDA0003347170290000033
for the estimation value of unknown input gain caused by faults such as airplane wing damage,
Figure RE-GDA0003347170290000034
is an estimated value of uncertain parameters related to states caused by faults such as airplane wing damage and the like,
Figure RE-GDA0003347170290000035
is an unknown constant disturbance estimation value;
the fast adaptation law is determined and,
Figure RE-GDA0003347170290000036
where Γ is the adaptive gain, and Proj is the projection operator,
Figure RE-GDA0003347170290000037
being the observed state of a state observer
Figure RE-GDA0003347170290000038
The difference from the actual state x measured by the sensors on board the aircraft,
Figure RE-GDA0003347170290000039
is composed of
Figure RE-GDA00033471702900000310
Is the transposition of (1), P is the Lyapunov equation
Figure RE-GDA00033471702900000311
Q satisfies Q ═ QT>0；
A control quantity calculation algorithm is determined and,
Figure RE-GDA00033471702900000312
where r is the system reference input, kgFor tracking gain, usually kgTake-1/(inv (A)o)·Bo) Wherein inv (A)o) Representation matrix AoInverse of (d), F (w, u)p) Represents a pair signal upLow-pass filtering is carried out, and the bandwidth of the filter is w, krTo control the gain.
Further, step S5 is specifically: advanced control quantity ulAnd L1 adaptive control quantity uaAdding to obtain a control surface skewness instruction us。
Further, determining the aircraft state observer specifically includes: for longitudinal attitude control, the relevant states are (v, α, ω)zθ), where v is the aircraft velocity, α is the aircraft angle of attack, ωzThe speed is the pitching angle of the airplane, and theta is the pitching angle of the airplane; for yaw-rate attitude control, the relevant state is (β, ω)x,ωyγ, ψ), where β is the aircraft sideslip angle, ωxIs the aircraft roll angular velocity, omegayThe yaw rate of the aircraft, gamma the roll angle of the aircraft, and psi the yaw angle of the aircraft.
The invention has the advantages that:
the method gives the approximate angular acceleration signal of the airplane according to the nonlinear tracking differentiator, and the approximate angular acceleration signal is synthesized with the angular acceleration signal output by the reference model, so that the interference signal of the airplane can be quickly identified. The disturbance moment and unmodeled dynamics of the airplane are quickly compensated through advanced control, so that the airplane transient state after wing damage is greatly restrained, and meanwhile, the L1 adaptive control is adopted to perform adaptive compensation on the residual disturbance and dynamics, so that the stability and the accuracy of the airplane after wing damage are realized.
Description of the drawings:
fig. 1 is a signal flow diagram of an adaptive control method for aircraft wing damage according to the present invention.
FIG. 2 is a roll angle acceleration differential plot calculated by the tracking differentiator and the reference model in accordance with an embodiment of the present invention;
FIG. 3 is a schematic of the tracking differentiator and the look-ahead control calculated by the reference model in accordance with an embodiment of the present invention;
FIG. 4 is a schematic diagram of the L1 adaptive control quantities calculated by the tracking differentiator and the reference model in accordance with an embodiment of the present invention;
FIG. 5 is a schematic diagram of the change in roll angle of an aircraft calculated by a tracking differentiator and a reference model in accordance with an embodiment of the invention.
The specific implementation mode is as follows:
therefore, in order to overcome the defects of the prior art, the invention provides an adaptive control method for the wing damage of an aircraft, which is used for reducing the transient state of the aircraft at the moment of wing damage, inhibiting high-frequency oscillation and improving the control performance.
In order to make the objects, technical solutions and advantages of the present invention more apparent, the present invention is described in further detail below with reference to the accompanying drawings 1 in conjunction with specific embodiments.
Step S1: receiving a three-axis angular velocity signal p of the airplane, and designing a tracking differentiator to determine an approximate three-axis angular acceleration signal epsilon of the airplane;
the tracking differentiator calculation formula is as follows:
Figure RE-GDA0003347170290000041
in the formula, z1And z2To track the two states generated by the differentiator, z1(k) Z for k sampling instants1Value z2(k) Z for k sampling instants2Value z1And z2Has an initial value of 0. h is the sampling step of the computer. ε (k) is the value of ε for k sampling instants. r is0Is a tracking factor, h, of the fhan function0As step size of fhan function, fhan (z)1(k)-p(k),z2(k),r0,h0) The calculation formula of (a) is as follows,
Figure RE-GDA0003347170290000051
where sign is a sign function, μ1，μ2，Sz1，μ，SμCompared with the traditional differentiation, the tracking differentiator outputs a signal which is less sensitive to the sampling time and smoother in order to calculate the intermediate variable related to the fhan function.
Step S2: determining a reference model of the airplane according to the existing aerodynamic and control effect data of the airplane, and setting the state tau of the airplane and the deflection of a control plane as usDetermining a reference angular acceleration epsilon of the aircraft by inputting a reference model of the aircraftc；
Reference model calculation formula:
I·εc＝M1(τ)+M2(τ,us)
wherein I is the moment of inertia of the aircraft, M1 (tau) is the aircraft moment related to the aircraft state tau when the control plane deflection is 0, M2 (tau, us) The deflection of the control surface is usMoment of flight, epsilon, induced by the flightcIs output of the reference model and represents the reference angular acceleration of the normal state of the aircraft, wherein the control plane deflection instruction usThe initial value of (a) is trim rudder deflection in the current state of the aircraft.
Step S3: receiving the approximate angular acceleration determined in the S1 and the reference angular acceleration determined in the S2, and determining an advance control quantity of an airplane control surface instruction according to an advance control module;
the calculation formula of the advance control amount is as follows:
Figure RE-GDA0003347170290000052
in the formula, klControl gain, epsilon-epsilon, for advanced controlcDisturbance torque, T, representing damage to the wing of an aircraft1(ud) Represents a pair udWith an ongoing time constant of T1First order smoothing of (T)1The size of the response bandwidth depends on the angular speed of the airplane, the direct and rapid compensation of the interference torque can be realized, the damage transient state of the airplane wing is reduced, and the smooth filtering is realizedThe wave suppresses high frequency noise caused by the differentiator and smoothes the control command.
Step S4: receiving an aircraft attitude control instruction and an aircraft dynamics related state, designing an L1 self-adaptive control law, and determining an L1 self-adaptive control quantity of an aircraft control surface instruction;
(1) designing an aircraft state observer, and estimating the residual disturbance of the aircraft;
the state observer has a calculation formula of
Figure RE-GDA0003347170290000061
In the formula
Figure RE-GDA0003347170290000062
For the observed aircraft state, for longitudinal attitude control, the relevant state is (v, α, ω)zθ), where v is the aircraft velocity, α is the aircraft angle of attack, ωzFor aircraft pitch angular velocity, θ is aircraft pitch angle, and for yaw attitude control, the associated state is (β, ω)x,ωyγ, ψ), where β is the aircraft sideslip angle, ωxIs the aircraft roll angular velocity, omegayIs the aircraft yaw rate, gamma is the aircraft roll angle, psi is the aircraft yaw angle, AoSystem matrix, system stabilization, B, relating to desired aircraft motion modal characteristicsoIs a control effect matrix of the aircraft control surfaces, uaThe control plane instruction output by the L1 self-adaptive control law has an initial value of trim rudder deflection under the current state of the airplane,
Figure RE-GDA0003347170290000063
for the estimation value of unknown input gain caused by faults such as airplane wing damage,
Figure RE-GDA0003347170290000064
is an estimated value of uncertain parameters related to states caused by faults such as airplane wing damage and the like,
Figure RE-GDA0003347170290000065
is an unknown constant disturbance estimation value.
(2) Designing a fast self-adaptive law;
Figure RE-GDA0003347170290000066
Figure RE-GDA0003347170290000067
Figure RE-GDA0003347170290000068
where Γ is the adaptive gain, Proj is the projection operator,
Figure RE-GDA0003347170290000069
being the observed state of a state observer
Figure RE-GDA00033471702900000610
The difference from the actual state x measured by the sensors on board the aircraft,
Figure RE-GDA00033471702900000611
is composed of
Figure RE-GDA00033471702900000612
Is the transposition of (1), P is the Lyapunov equation
Figure RE-GDA00033471702900000613
Q satisfies Q ═ QT>0。
(3) And (3) resolving a control quantity:
Figure RE-GDA00033471702900000614
where r is the system reference input, kgFor tracking gain, usually kgTake-1/(inv (A)o)·Bo) Wherein inv (A)o) Representation matrix AoInverse of (d), F (w, u)p) Represents a pair signal upPerforming low-pass filteringWave, the bandwidth of the filter is w. w is not more than the bandwidth, k, of the airplane steering enginerTo control the gain. Control quantity upCan effectively compensate and stabilize the residual disturbance of the airplane, realize the accurate tracking of the input signal, and the low-pass filter makes the control quantity upThe disturbance is compensated within the bandwidth range of the controller, high-frequency oscillation caused by rapid self-adaption is restrained, and control performance is improved.
Step S5: according to the advance control amount u determined in S3lAnd L1 adaptive control amount u determined in S4aDetermining control plane skewness instruction usAnd controlling the attitude and movement of the aircraft.
us＝ul+ua
In the existing control research aiming at the damage of the wings of the airplane, the problems of small airplane transient state after the damage of the wings and avoidance of high-frequency oscillation caused by self-adaptation cannot be considered for both passive fault-tolerance and active fault-tolerance.
Example (b):
the method is characterized in that a plane with a certain conventional layout is taken as a research object, the self-adaptive control method provided by the invention is used for carrying out dynamic simulation of damage to one side wing of the plane, the simulation step length is set to be 0.01 second, the plane flies horizontally at the height of 5000 meters at the speed of Mach 0.6 at the height of 0 second, the roll angle keeps 0 degree, the right side wing is set to be 40% damage of the wing tip according to the span length at 5 seconds, FIG. 2 is the roll angle acceleration difference calculated by a tracking differentiator and a reference model and represents the interference moment in the roll direction of the plane, FIG. 3 is an advance control quantity, FIG. 4 is an L1 self-adaptive control quantity, FIG. 5 is a roll angle change curve of the plane, and under the action of the advance control, the transient state of the plane after wing damage is small, the plane roll angle is gradually recovered and a target command is kept under the self-adaptive control of L1, the control quantity and the plane response are smooth, and the engineering application is convenient.
Although the subject matter has been described in language specific to structural features and/or methodological acts, it is to be understood that the subject matter defined in the appended claims is not necessarily limited to the specific features or acts described above. Rather, the specific features and acts described above are disclosed as example forms of implementing the claims, and any modifications, equivalents, improvements, etc. that come within the spirit and scope of the inventions are intended to be included therein. The above description is only a preferred embodiment of the present invention, and it should be noted that, for those skilled in the art, several modifications and variations can be made without departing from the technical principle of the present invention, and these modifications and variations should also be regarded as the protection scope of the present invention.

Claims (8)

1. An adaptive control method for aircraft wing damage, comprising:
step S1: receiving a three-axis angular velocity signal p of the airplane, and carrying out differential control according to a tracking differentiator to determine an approximate three-axis angular acceleration signal epsilon of the airplane;
step S2: determining a reference model of the airplane according to the existing aerodynamic and control effect data of the airplane, and setting the state tau of the airplane and the deflection of a control plane as usDetermining a reference angular acceleration epsilon of the aircraft by inputting a reference model of the aircraftc；
Step S3: receiving the approximate angular acceleration determined in the S1 and the reference angular acceleration determined in the S2, and determining an advanced control quantity u of the airplane control surface instruction according to the advanced control module for advanced controll；
Step S4: receiving an aircraft attitude control command and an aircraft dynamics related state, determining an L1 self-adaptive control law, and performing self-adaptive control according to the L1 self-adaptive control law to determine an L1 self-adaptive control quantity u of an aircraft control surface commanda；
Step S5: according to the advance control amount u determined in S3lAnd L1 adaptive control amount u determined in S4aDetermining control plane skewness instruction us。
2. Method according to claim 1, characterized in that a differentiator is tracked, in particular
Figure FDA0003222626120000011
Wherein z is1And z2To track the two states generated by the differentiator, z1(k) Z for k sampling instants1Value z2(k) Z for k sampling instants2Value z1And z2Is 0, h is the sampling step length of the computer, epsilon (k) is the epsilon value of k sampling moments, r0Is a tracking factor, h, of the fhan function0As step size of fhan function, fhan (z)1(k)-p(k),z2(k),r0,h0) Comprises the following steps:
Figure FDA0003222626120000021
sign is a sign function, μ1，μ2，
Figure FDA0003222626120000029
μ，SμTo calculate the intermediate variables involved in the fhan function.
3. The method of claim 1, wherein the reference model is I · sc＝M1(τ)+M2(τ,us) Wherein I is the moment of inertia of the aircraft, M1 (tau) is the aircraft moment related to the aircraft state tau when the control plane deflection is 0, and M2 (tau, us) The deflection of the control surface is usMoment of flight, epsilon, induced by the flightcIs the output of the reference model.
4. Method according to claim 3, characterized in that the control plane deflection command usThe initial value of (a) is trim rudder deflection in the current state of the aircraft.
5. The method of claim 1, wherein the look-ahead control amount is calculated as follows:
Figure FDA0003222626120000022
wherein k islControl gain for advanced control, T1(ud) Represents a pair udWith an ongoing time constant of T1To first order smoothing filtering.
6. The method according to claim 1, wherein determining the L1 adaptive control law specifically comprises:
determining an aircraft state observer and estimating the residual disturbance of the aircraft, wherein the state observer has a calculation formula of
Figure FDA0003222626120000023
Figure FDA0003222626120000024
For observing the state of the aircraft, AoSystem matrix relating to desired aircraft movement modal characteristics, BoIs a control effect matrix of the aircraft control surfaces, uaThe control plane instruction output by the L1 self-adaptive control law has an initial value of trim rudder deflection under the current state of the airplane,
Figure FDA0003222626120000025
for the estimation value of unknown input gain caused by faults such as airplane wing damage,
Figure FDA0003222626120000026
is an estimated value of uncertain parameters related to states caused by faults such as airplane wing damage and the like,
Figure FDA0003222626120000027
is an unknown constant disturbance estimation value;
the fast adaptation law is determined and,
Figure FDA0003222626120000028
where Γ is the adaptive gain, and Proj is the projection operator,
Figure FDA0003222626120000035
being the observed state of a state observer
Figure FDA0003222626120000031
The difference from the actual state x measured by the sensors on board the aircraft,
Figure FDA0003222626120000032
is composed of
Figure FDA0003222626120000033
Is the transposition of (A), P is the Lyapunov equation Ao TSolution of P + PA ═ Q, Q satisfying Q ═ QT>0；
A control quantity calculation algorithm is determined and,
Figure FDA0003222626120000034
where r is the system reference input, kgFor tracking gain, usually kgTake-1/(inv (A)o)·Bo) Wherein inv (A)o) Representation matrix AoInverse of (d), F (w, u)p) Represents a pair signal upLow-pass filtering is carried out, and the bandwidth of the filter is w, krTo control the gain.
7. The method according to claim 1, wherein step S5 is specifically:
advanced control quantity ulAnd L1 adaptive control quantity uaAdding to obtain a control surface skewness instruction us。
8. The method according to claim 1, characterized in that determining an aircraft state observer comprises in particular:
for longitudinal attitude control, the relevant states are (v, α, ω)zθ), where v is the aircraft velocity, α is the aircraft angle of attack, ωzFor aircraft pitch angle velocity, theta for flyA machine pitch angle;
for yaw-rate attitude control, the relevant state is (β, ω)x,ωyγ, ψ), where β is the aircraft sideslip angle, ωxIs the aircraft roll angular velocity, omegayThe yaw rate of the aircraft, gamma the roll angle of the aircraft, and psi the yaw angle of the aircraft.
CN202110962610.XA 2021-08-20 2021-08-20 Self-adaptive control method for aircraft wing damage Active CN113759718B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN202110962610.XA CN113759718B (en) 2021-08-20 2021-08-20 Self-adaptive control method for aircraft wing damage

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN202110962610.XA CN113759718B (en) 2021-08-20 2021-08-20 Self-adaptive control method for aircraft wing damage

Publications (2)

Publication Number Publication Date
CN113759718A true CN113759718A (en) 2021-12-07
CN113759718B CN113759718B (en) 2023-10-20

Family

ID=78790653

Family Applications (1)

Application Number Title Priority Date Filing Date
CN202110962610.XA Active CN113759718B (en) 2021-08-20 2021-08-20 Self-adaptive control method for aircraft wing damage

Country Status (1)

Country Link
CN (1) CN113759718B (en)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114415516A (en) * 2022-01-20 2022-04-29 中国空气动力研究与发展中心低速空气动力研究所 Fault-tolerant control method for damage fault of control surface of fixed-wing aircraft
CN115783247A (en) * 2022-11-11 2023-03-14 中国航空工业集团公司西安飞行自动控制研究所 Active control method for improving longitudinal riding quality

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20120138729A1 (en) * 2010-12-01 2012-06-07 Raytheon Company Flight-control system for canard-controlled flight vehicles and methods for adaptively limiting acceleration
US9296474B1 (en) * 2012-08-06 2016-03-29 The United States of America as represented by the Administrator of the National Aeronautics & Space Administration (NASA) Control systems with normalized and covariance adaptation by optimal control modification
CN109062237A (en) * 2018-09-17 2018-12-21 南京航空航天大学 A kind of unmanned tiltrotor active disturbance rejection attitude control method
US20200183342A1 (en) * 2018-12-11 2020-06-11 Dalian University Of Technology Steady state and transition state multi-variable control method of a turbofan engine based on an active disturbance rejection control theory (adrc)
CN112346470A (en) * 2020-10-29 2021-02-09 中国地质大学（武汉） Four-rotor attitude control method based on improved active disturbance rejection control

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20120138729A1 (en) * 2010-12-01 2012-06-07 Raytheon Company Flight-control system for canard-controlled flight vehicles and methods for adaptively limiting acceleration
US9296474B1 (en) * 2012-08-06 2016-03-29 The United States of America as represented by the Administrator of the National Aeronautics & Space Administration (NASA) Control systems with normalized and covariance adaptation by optimal control modification
CN109062237A (en) * 2018-09-17 2018-12-21 南京航空航天大学 A kind of unmanned tiltrotor active disturbance rejection attitude control method
US20200183342A1 (en) * 2018-12-11 2020-06-11 Dalian University Of Technology Steady state and transition state multi-variable control method of a turbofan engine based on an active disturbance rejection control theory (adrc)
CN112346470A (en) * 2020-10-29 2021-02-09 中国地质大学（武汉） Four-rotor attitude control method based on improved active disturbance rejection control

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114415516A (en) * 2022-01-20 2022-04-29 中国空气动力研究与发展中心低速空气动力研究所 Fault-tolerant control method for damage fault of control surface of fixed-wing aircraft
CN114415516B (en) * 2022-01-20 2023-04-07 中国空气动力研究与发展中心低速空气动力研究所 Fault-tolerant control method for damage fault of control surface of fixed-wing aircraft
CN115783247A (en) * 2022-11-11 2023-03-14 中国航空工业集团公司西安飞行自动控制研究所 Active control method for improving longitudinal riding quality

Also Published As

Publication number Publication date
CN113759718B (en) 2023-10-20

Similar Documents

Publication Publication Date Title
Mechali et al. Observer-based fixed-time continuous nonsingular terminal sliding mode control of quadrotor aircraft under uncertainties and disturbances for robust trajectory tracking: Theory and experiment
Yu et al. Safe control of trailing UAV in close formation flight against actuator fault and wake vortex effect
CN106773713B (en) High-precision nonlinear path tracking control method for under-actuated marine vehicle
CN111158383B (en) Unmanned ship track tracking control method based on interference observer and RBFNN
CN113759718B (en) Self-adaptive control method for aircraft wing damage
CN111367182A (en) Hypersonic aircraft anti-interference backstepping control method considering input limitation
CN112180965A (en) High-precision overload control method
CN114488803A (en) Unmanned ship/aircraft robust adaptive neural cooperative control method based on event triggering
CN106802570B (en) Method and device for tracking position of unmanned helicopter
Chang et al. Integrated guidance and control design for the hypersonic interceptor based on adaptive incremental backstepping technique
CN111610794A (en) Large-attack-angle dynamic inverse control method for fighter based on sliding mode disturbance observer
Li et al. Angular acceleration estimation-based incremental nonlinear dynamic inversion for robust flight control
CN113568419A (en) Fault-tolerant control method for variable-load quad-rotor unmanned aerial vehicle
CN110209192A (en) Fighter plane course augmentation control design method
CN112947518A (en) Four-rotor robust attitude control method based on disturbance observer
CN113219838A (en) High-precision control method of triaxial inertially stabilized platform based on adaptive time-varying bandwidth observer
Li et al. Anti-disturbance control for attitude and altitude systems of the helicopter under random disturbances
CN114721266B (en) Self-adaptive reconstruction control method under condition of structural failure of control surface of airplane
CN113128035A (en) Civil aircraft flight control sensor signal reconstruction fault-tolerant control method
Zhang et al. Integral barrier Lyapunov function-based three-dimensional low-order integrated guidance and control design with seeker's field-of-view constraint
Xiong et al. Disturbance Rejection in UAV's velocity and attitude control: Problems and solutions
CN113625732A (en) Incremental feedback inverse angular velocity control law design method based on angular acceleration estimation
CN113110538A (en) Fixed-time fault-tolerant control method for carrier-based aircraft landing based on backstepping control
CN115793449B (en) Helicopter flight attitude controller design method and device, equipment and storage medium
CN116755337A (en) Hypersonic aircraft fault-tolerant control method based on long-short-term memory neural network linear 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