CN113759718B - Self-adaptive control method for aircraft wing damage - Google Patents
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Abstract
The invention relates to the technical field of aviation, in particular to a self-adaptive control method for aircraft wing damage. Comprising the following steps: receiving the triaxial angular velocity signal, and determining a triaxial approximate angular acceleration signal according to the tracking differentiator; determining a reference model of the aircraft according to the pneumatic and rudder performance data, and determining a reference angular acceleration by taking the aircraft 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 aircraft control surface command 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 deflection command according to the determined advanced control quantity and the determined L1 self-adaptive control quantity. The disturbance moment of the aircraft is quickly compensated through the advanced control, the transient state of the aircraft after the wing is damaged is restrained, meanwhile, the L1 adaptive control is adopted to carry out adaptive compensation on the residual disturbance, and the stability and the accurate control of the aircraft are realized.
Description
Technical Field
The invention relates to the technical field of aviation, in particular to an aircraft wing damage self-adaptive control method.
Background
After a unilateral wing is damaged due to an emergency in the flight of the aircraft, aerodynamic force and moment are changed greatly in a short time, the gravity center of the aircraft is deviated, the flight dynamics equation is changed severely, and the stability and self-adaptive control of the aircraft are difficult to realize by the existing control method. In the prior art, a passive error-tolerant control method based on neural network self-adaptive robust nonlinear model inverse exists. The method realizes the stability of the aircraft after the control surface of the aircraft is blocked or the area of a single-side wing is damaged by 40%, but the transient state of the aircraft is larger at the moment of wing damage due to the limitation of passive error-tolerant control, 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 control structure is adjusted after the faults occur. For sudden damage faults, the time delay of online observation can reduce the suppression performance of fault transient, and the high gain caused by the rapid self-adaption of the observer can cause high-frequency oscillation of the aircraft.
Disclosure of Invention
The purpose of the invention is that: the self-adaptive control method for the wing damage of the aircraft is provided, so that the transient state of the aircraft at the moment of wing damage is reduced, high-frequency oscillation is restrained, and the control performance is improved.
The technical scheme of the invention is as follows:
in a first aspect, an aircraft wing damage adaptive control method is provided, including: step S1: receiving an aircraft triaxial angular velocity signal p, and determining an aircraft triaxial approximate angular acceleration signal epsilon according to differential control of a tracking differentiator; step S2: determining a reference model of the aircraft according to the existing aerodynamic and rudder data of the aircraft, and setting the aircraft state tau and the control surface deflection as u s Determining a reference angular acceleration epsilon of an aircraft from a reference model of the aircraft c The method comprises the steps of carrying out a first treatment on the surface of the Step S3: receiving the approximate angular acceleration determined in S1 and the reference angular acceleration determined in S2, and determining the advanced control quantity u of the aircraft control surface command according to the advanced control performed by the advanced control module l The method comprises the steps of carrying out a first treatment on the surface of the Step S4: receiving an aircraft attitude control instruction, determining an L1 adaptive control law according to an aircraft dynamics related state, and performing adaptive control according to the L1 adaptive control law to determine an L1 adaptive control quantity u of an aircraft control surface instruction a The method comprises the steps of carrying out a first treatment on the surface of the Step S5: according to the advance control amount u determined in S3 l And the L1 adaptive control amount u determined in S4 a Determining control surface deflection command u s 。
Further, a tracking differentiator, in particularWherein z is 1 And z 2 For tracking the two states generated by the differentiator, z 1 (k) Z for k sampling instants 1 Value, z 2 (k) Z for k sampling instants 2 Value, z 1 And z 2 The initial value of (a) is 0, h is the sampling step length of a computer, epsilon (k) is epsilon value of k sampling moments, and r 0 Tracking factor, h, as a fhan function 0 Step size of fhan function, fhan (z 1 (k)-p(k),z 2 (k),r 0 ,h 0 ) The method comprises the following steps:
sign is a sign function, mu 1 ，μ 2 ，S z1 ，μ，S μ Intermediate transformations involved in calculating fhan functionsAmount of the components.
Further, the reference model is I.epsilon. c ＝M1(τ)+M2(τ,u s ) Wherein I is the rotational inertia of the aircraft, M1 (τ) is the aircraft moment associated with the aircraft state τ when the control surface deflection is 0, M2 (τ, u) s ) The deflection degree of the control surface is u s Moment epsilon of airplane caused by time c Is the output of the reference model.
Further, control surface deflection command u s The initial value of (2) is the trim rudder deflection in the current state of the aircraft.
Further, the advance control amount is calculated as follows:
wherein k is l For the control gain of the advance control, T 1 (u d ) Representation pair u d Time constant of the process is T 1 Is a first order smoothing filter of (a).
Further, determining the L1 adaptive control law specifically includes: determining an aircraft state observer and estimating the residual disturbance of the aircraft, wherein a calculation formula of the state observer is as follows For observing the state of the aircraft A o System matrix related to expected motion mode characteristics of airplane, B o Is the rudder matrix of the control surface of the airplane, u a The control surface instruction output by the L1 self-adaptive control law is the trim rudder deflection of the aircraft in the current state as the initial value>Estimated value of unknown input gain for faults such as aircraft wing damage, < >>Estimated value of state-related uncertainty parameter for failure such as aircraft wing damage, ++>The disturbance estimated value is unknown constant value;
a fast-adaptation law is determined and,wherein Γ is the adaptive gain, proj is the projection operator, < >>Observation state for state observer +.>Difference from actual state x measured by sensors on board the aircraft, < >>Is->Is transposed of P is Lyapunov equation +.>Q satisfies q=q T >0；
A control amount calculation algorithm is determined,where r is the system reference input, k g For tracking gain, generally k g Take-1/(inv (A) o )·B o ) Wherein inv (A) o ) Representation matrix A o Is the inverse of F (w, u) p ) Representing the pair signal u p Low-pass filtering with bandwidth of w, k r To control the gain.
Further, step S5 specifically includes: advance control amount u l And L1 adaptive control amount u a Adding to obtain control surface deflection command u s 。
Further, determination of aircraft state observers, in particularComprising the following steps: for longitudinal attitude control, the relevant states are (v, α, ω) z θ), where v is aircraft speed, α is aircraft angle of attack, ω z The pitch angle speed of the aircraft, and the theta is the pitch angle of the aircraft; for the lateral attitude control, the relevant state is (β, ω) x ,ω y Gamma, ψ), where beta is the aircraft sideslip angle, ω x Is the roll angle speed omega of the airplane y Is the yaw rate of the aircraft, gamma is the roll angle of the aircraft, and ψ is the yaw angle of the aircraft.
The invention has the advantages that:
the invention provides the approximate angular acceleration signal of the aircraft according to the nonlinear tracking differentiator, and can rapidly identify the interference signal of the aircraft by integrating the approximate angular acceleration signal with the angular acceleration signal output by the reference model. The disturbance moment and unmodeled dynamics of the aircraft are rapidly compensated through the advanced control, so that the transient state of the aircraft after the wing is damaged is greatly restrained, meanwhile, the residual disturbance and dynamics are adaptively compensated through the L1 adaptive control, and the stable and accurate control after the wing is damaged is realized.
Description of the drawings:
fig. 1 is a signal flow diagram of an adaptive control method for aircraft wing damage provided by the invention.
FIG. 2 is a schematic diagram of the roll angle acceleration difference calculated by the tracking differentiator and the reference model in accordance with an embodiment of the invention;
FIG. 3 is a schematic diagram of the lead control amount calculated by the tracking differentiator and the reference model according to an embodiment of the invention;
FIG. 4 is a schematic diagram of L1 adaptive control amounts calculated by a tracking differentiator and a reference model according to an embodiment of the invention;
FIG. 5 is a graph of roll angle variation of an aircraft calculated by a tracking differentiator and a reference model according to an embodiment of the invention.
The specific embodiment is as follows:
therefore, in order to overcome the defects in the prior art, the invention provides an aircraft wing damage self-adaptive control method, which is used for reducing the aircraft transient at the moment of wing damage, inhibiting high-frequency oscillation and improving control performance.
The present invention will be further described in detail below with reference to the specific embodiments and with reference to fig. 1, in order to make the objects, technical solutions and advantages of the present invention more apparent.
Step S1: receiving an airplane triaxial angular velocity signal p, and designing a tracking differentiator to determine an airplane triaxial approximate angular acceleration signal epsilon;
the tracking differentiator calculation formula is as follows:
wherein z is 1 And z 2 For tracking the two states generated by the differentiator, z 1 (k) Z for k sampling instants 1 Value, z 2 (k) Z for k sampling instants 2 Value, z 1 And z 2 The initial value of (2) is 0.h is the sampling step length of the computer. Epsilon (k) is the epsilon value for k sample instants. r is (r) 0 Tracking factor, h, as a fhan function 0 Step size of fhan function, fhan (z 1 (k)-p(k),z 2 (k),r 0 ,h 0 ) The calculation formula of (c) is as follows,
wherein sign is a sign function, mu 1 ，μ 2 ，S z1 ，μ，S μ To calculate the intermediate variables involved in the fhan function, the signal output by the tracking differentiator is less sensitive to the sampling time and is smoother than conventional differentiation.
Step S2: determining a reference model of the aircraft according to the existing aerodynamic and rudder data of the aircraft, and setting the aircraft state tau and the control surface deflection as u s Determining a reference angular acceleration epsilon of an aircraft from a reference model of the aircraft c ；
The reference model calculation formula:
I·ε c ＝M1(τ)+M2(τ,u s )
wherein I is the rotational inertia of the aircraft, M1 (tau) is the moment of the aircraft relative to the aircraft state tau when the steering angle is 0, M2(τ,u s ) The deflection degree of the control surface is u s Moment epsilon of airplane caused by time c For the output of the reference model, the reference angular acceleration represents the normal state of the aircraft, wherein the control surface deflection command u s The initial value of (2) is the trim rudder deflection in the current state of the aircraft.
Step S3: receiving the approximate angular acceleration determined in the step S1 and the reference angular acceleration determined in the step S2, and determining the advanced control quantity of the aircraft control surface command according to the advanced control module;
the advance control amount calculation formula is as follows:
wherein k is l Control gain for lead control, ε - ε c Representing the disturbance moment of the damage of the aircraft wing, T 1 (u d ) Representation pair u d Time constant of the process is T 1 First order smoothing filtering of T 1 The magnitude depends on the response bandwidth of the aircraft angular speed, so that the direct and rapid compensation of the interference moment can be realized, the damage transient state of the aircraft wing is reduced, the high-frequency noise brought by the differentiator is restrained by the smoothing filter, and the control instruction is smoothed.
Step S4: receiving an aircraft attitude control instruction, designing an L1 self-adaptive control law in an aircraft dynamics related state, 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 ofIn->For the observed aircraft state, for longitudinal attitude control, the relevant states are (v, α, ω) z θ), where v is aircraft speed, α is aircraft angle of attack, ω z The pitch angle speed of the aircraft, the pitch angle theta of the aircraft and the transverse heading postureControl, the relevant state is (beta, omega) x ,ω y Gamma, ψ), where beta is the aircraft sideslip angle, ω x Is the roll angle speed omega of the airplane y Is the yaw rate of the aircraft, gamma is the roll angle of the aircraft, ψ is the yaw rate of the aircraft, A o System matrix related to expected motion modal characteristics of aircraft, system stability, B o Is the rudder matrix of the control surface of the airplane, u a The control surface instruction output by the L1 self-adaptive control law is the trim rudder deflection of the aircraft in the current state as the initial value>Estimated value of unknown input gain for faults such as aircraft wing damage, < >>Estimated value of state-related uncertainty parameter for failure such as aircraft wing damage, ++>Is an unknown constant disturbance estimated value.
(2) Designing a rapid self-adaptive law;
where Γ is the adaptive gain, proj is the projection operator,observation state for state observer +.>Difference from actual state x measured by sensors on board the aircraft, < >>Is->Is transposed of P is Lyapunov equation +.>Q satisfies q=q T >0。
(3) Control amount calculation:
where r is the system reference input, k g For tracking gain, generally k g Take-1/(inv (A) o )·B o ) Wherein inv (A) o ) Representation matrix A o Is the inverse of F (w, u) p ) Representing the pair signal u p Low pass filtering is performed, and the bandwidth of the filter is w. The w size is not larger than the bandwidth of an airplane steering engine, and k is not larger than the bandwidth of the airplane steering engine r To control the gain. Control amount u p Can effectively compensate and stabilize the residual disturbance of the aircraft, realize accurate tracking of input signals, and enable a low-pass filter to control the quantity u p The disturbance is compensated within the bandwidth range of the controller, high-frequency oscillation caused by rapid self-adaption is restrained, and the control performance is improved.
Step S5: according to the advance control amount u determined in S3 l And the L1 adaptive control amount u determined in S4 a Determining control surface deflection command u s Controlling the attitude and movement of the aircraft.
u s ＝u l +u a
In the existing control research aiming at the wing damage of the aircraft, both passive error tolerance and active fault tolerance cannot be considered, the problem of small transient state of the aircraft after the wing damage and high-frequency oscillation caused by self-adaption are avoided, the invention combines the advanced control based on the tracking differentiator and the L1 self-adaption control for the first time, and the disturbance moment of the aircraft is quickly compensated through the advanced control, so that the transient state of the aircraft after the wing damage is restrained, and meanwhile, the L1 self-adaption control is adopted to carry out self-adaption compensation on the residual disturbance and the dynamics, so that the stable and accurate control after the wing damage of the aircraft is realized.
Examples:
the self-adaptive control method provided by the invention is used for carrying out the simulation of the damage dynamics of the single-side wing of the aircraft by taking a certain conventional layout aircraft as a research object, the simulation step length is set to be 0.01 second, the aircraft flies horizontally at a Mach speed of 0.6 at a height of 5000 meters for 0 second, the roll angle is kept at 0 degree, 40% damage of the right-side wing according to the unfolding length from the wing tip is set for 5 seconds, the roll angle acceleration difference calculated by a tracking differentiator and a reference model represents the disturbance moment of the roll direction of the aircraft, the lead control quantity is shown in FIG. 3, the L1 self-adaptive control quantity is shown in FIG. 5, the transient state after the damage of the wing of the aircraft is smaller under the effect of the lead control, the roll angle of the aircraft is gradually recovered and kept as a target instruction under the effect of the L1 self-adaptive control, the control quantity and the response of the aircraft are smoother, and 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 exemplary forms of implementing the claims, and any modifications, equivalents, improvements or otherwise made within the spirit and principles of the invention are intended to be included within the scope of this invention. The foregoing is merely a preferred embodiment of the present invention, and it should be noted that it will be apparent to those skilled in the art that modifications and variations can be made without departing from the technical principles of the present invention, and these modifications and variations should also be regarded as the scope of the invention.
Claims (7)
1. An aircraft wing damage adaptive control method, comprising:
step S1: receiving an aircraft triaxial angular velocity signal p, and determining an aircraft triaxial approximate angular acceleration signal epsilon according to differential control of a tracking differentiator;
step S2: determining a reference model of the aircraft according to the existing aerodynamic and rudder data of the aircraft, and setting the aircraft state tau and the control surface deflection as u s Determining a reference angular acceleration epsilon of an aircraft from a reference model of the aircraft c ；
Step S3: receiving the approximate angular acceleration determined in S1 and the reference angular acceleration determined in S2, and determining the advanced control quantity u of the aircraft control surface command according to the advanced control performed by the advanced control module l ；
Step S4: receiving an aircraft attitude control instruction, determining an L1 adaptive control law according to an aircraft dynamics related state, and performing adaptive control according to the L1 adaptive control law to determine an L1 adaptive control quantity u of an aircraft control surface instruction a ；
Step S5: according to the advance control amount u determined in S3 l And the L1 adaptive control amount u determined in S4 a Determining control surface deflection command u s ，
Wherein the tracking differentiator is in particularz 1 And z 2 For tracking the two states generated by the differentiator, z 1 (k) Z for k sampling instants 1 Value, z 2 (k) Z for k sampling instants 2 Value, z 1 And z 2 The initial value of (a) is 0, h is the sampling step length of a computer, epsilon (k) is epsilon value of k sampling moments, and r 0 Tracking factor, h, as a fhan function 0 Step size of fhan function, fhan (z 1 (k)-p(k),z 2 (k),r 0 ,h 0 ) The method comprises the following steps:
sign is a sign function, mu 1 ，μ 2 ，S z1 ，μ，S μ To calculate the intermediate variables involved in the fhan function.
2. The method of claim 1, wherein the reference model is I ∈ c ＝M1(τ)+M2(τ,u s ) Wherein I is the rotational inertia of the aircraft, M1 (τ) is the aircraft moment associated with the aircraft state τ when the control surface deflection is 0, M2 (τ, u) s ) The deflection degree of the control surface is u s Moment epsilon of airplane caused by time c Is the output of the reference model.
3. The method of claim 2, wherein the control surface bias command u s The initial value of (2) is the trim rudder deflection in the current state of the aircraft.
4. The method of claim 1, wherein the advance control amount is calculated as follows:
wherein k is l For the control gain of the advance control, T 1 (u d ) Representation pair u d Time constant of the process is T 1 Is a first order smoothing filter of (a).
5. 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 a calculation formula of the state observer is as follows For observing the state of the aircraft A o System matrix related to expected motion mode characteristics of airplane, B o Is the rudder matrix of the control surface of the airplane, u a Rudder outputting L1 self-adaptive control lawA face instruction, wherein an initial value is trim rudder deflection in the current state of the aircraft, < ->Estimated value of unknown input gain for aircraft wing damage,/->Estimated value of state-related uncertainty parameter for aircraft wing damage +.>The disturbance estimated value is unknown constant value;
a fast-adaptation law is determined and,wherein Γ is the adaptive gain, proj is the projection operator, < >>Observation state for state observer +.>Difference from actual state x measured by sensors on board the aircraft, < >>Is->Is transposed of P is Lyapunov equation A o T Solution of p+pa= -Q, Q satisfying q=q T ＞0；
A control amount calculation algorithm is determined,where r is the system reference input, k g Take-1/(inv (A) o )·B o ) Wherein inv (A) o ) Representation matrix A o Is the inverse of F (w, u) p ) Representing the pair signal u p Low-pass filtering with bandwidth of w, k r To control the gain.
6. The method according to claim 1, wherein step S5 is specifically:
advance control amount u l And L1 adaptive control amount u a Adding to obtain control surface deflection command u s 。
7. The method according to claim 1, characterized in that determining an aircraft state observer, in particular comprises:
for longitudinal attitude control, the relevant states are (v, α, ω) z θ), where v is aircraft speed, α is aircraft angle of attack, ω z The pitch angle speed of the aircraft, and the theta is the pitch angle of the aircraft;
for the lateral attitude control, the relevant state is (β, ω) x ,ω y Gamma, ψ), where beta is the aircraft sideslip angle, ω x Is the roll angle speed omega of the airplane y Is the yaw rate of the aircraft, gamma is the roll angle of the aircraft, and ψ is the yaw angle of the aircraft.
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| 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 |
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