CN112445235A - Roll stability control method and system applied to high-dynamic aircraft - Google Patents

Roll stability control method and system applied to high-dynamic aircraft Download PDF

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Publication number
CN112445235A
CN112445235A CN201910797406.XA CN201910797406A CN112445235A CN 112445235 A CN112445235 A CN 112445235A CN 201910797406 A CN201910797406 A CN 201910797406A CN 112445235 A CN112445235 A CN 112445235A
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aircraft
roll
rudder deflection
deflection angle
angle
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CN112445235B (en
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王雨辰
王伟
林德福
郭永仓
南宇翔
纪毅
师兴伟
赵健廷
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Northwest Industrial Group Co ltd
Beijing Institute of Technology BIT
China North Industries Corp
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Northwest Industrial Group Co ltd
Beijing Institute of Technology BIT
China North Industries Corp
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    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
    • G05D1/10—Simultaneous control of position or course in three dimensions
    • G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft

Abstract

The invention discloses a roll stability control method and system applied to a high-dynamic aircraft, wherein the method comprises the steps of measuring roll angle information of the high-dynamic aircraft in real time, calculating in real time to obtain an instruction rudder deflection angle so as to control the roll angle rate of the aircraft, and eliminating dynamic lag of an executing mechanism and obtaining a compensated instruction rudder deflection angle after the aircraft control system transmits the instruction rudder deflection angle to the executing mechanism and before the aircraft actually steers. The method can effectively control the rolling stability of the aircraft, eliminate the adverse effect caused by the dynamic lag of the actuating mechanism, obviously improve the reliability of the aircraft and achieve the aim of accurate control. The system has strong universality, is suitable for both rolling aircrafts and non-rolling aircrafts, and can quickly converge to an expected value.

Description

Roll stability control method and system applied to high-dynamic aircraft
Technical Field
The invention relates to the field of roll stability control of aircrafts, in particular to a roll stability control method and a roll stability control system applied to a high-dynamic aircraft.
Background
The precision guidance aircraft can be roughly divided into a rotary aircraft and a non-rotary aircraft, and in order to realize stable control of the aircraft, the rotary aircraft is required to have stable rotation angular velocity, and the non-rotary aircraft is required to keep the rotation angular velocity to be zero. With the improvement of the requirement of modern war forms on guided munitions, the conventional aircraft with a non-rotating system cannot meet the high-dynamic and intelligent battlefield requirement, so that the rotary aircraft is widely applied.
High-speed rotating aircraft, while capable of improving the control efficiency of the aircraft, also have a number of problems, such as: pitch, coupling of yaw channels, high speed spinning leads to aircraft control instability, actuator hysteresis leads to reduced control accuracy, and the like. The existing controller generally ignores the nonlinear characteristics of the dynamic characteristics of the actuating mechanism and the pneumatic disturbance, only considers the application condition under the condition of a small attack angle, and is easy to cause the problem of control misalignment, and because the dynamic characteristics of the actuating mechanism are not considered, the dynamic lag of the actuating mechanism often exists, the control system clock is difficult to match, and the problem of control misalignment is further caused.
Therefore, it is necessary to provide a roll stability control method and system for a high dynamic aircraft, which can effectively control the roll stability of the aircraft, eliminate the adverse effect caused by the dynamic lag of the actuator, provide the reliability of the aircraft, and achieve the goal of precise control.
Disclosure of Invention
In order to overcome the problems, the inventor of the invention makes a keen study and designs a roll stability control method and a roll stability control system applied to a high-dynamic aircraft, the method measures roll angle information of the high-dynamic aircraft in real time, calculates the command rudder deflection angle in real time to control the roll angle rate of the aircraft, eliminates the kinetic lag of an executing mechanism after the command rudder deflection angle is transmitted to the executing mechanism by an aircraft control system and before the aircraft actually steers, obtains the compensated command rudder deflection angle, can effectively control the roll stability of the aircraft, eliminates the adverse effect caused by the kinetic lag of the executing mechanism, remarkably improves the reliability of the aircraft, and achieves the aim of accurate control, thereby completing the invention.
Specifically, the present invention aims to provide the following:
in a first aspect, a roll stability control method applied to a high dynamic aircraft is provided, the method comprising the following steps:
step 1, measuring in real time to obtain the roll angle information of a high-dynamic aircraft;
step 2, the aircraft control system calculates in real time according to the roll angle information obtained in the step 1 to obtain an instruction rudder deflection angle, and transmits the instruction rudder deflection angle to an aircraft executing mechanism so as to control the roll angle rate of the aircraft to be converged to a desired value;
in the step 2, after the aircraft control system transmits the command rudder deflection angle to the execution mechanism and before the aircraft actually steers, the method further comprises the step of eliminating the dynamic lag of the execution mechanism in real time and obtaining the compensated command rudder deflection angle.
In a second aspect, a roll stability control system for a high dynamic aircraft is provided, preferably for implementing the method of the first aspect, the system comprising a roll angle measurement module, a command rudder deflection angle calculation module and a command rudder deflection angle compensation module, wherein,
the roll angle measuring module is used for measuring the roll angle of the aircraft in real time,
the command rudder deflection angle resolving module is used for resolving and calculating in real time according to the roll angle information to obtain a command rudder deflection angle;
the command rudder deflection angle compensation module is used for eliminating the dynamic lag of an aircraft executing mechanism so as to obtain a compensated command rudder deflection angle.
The invention has the advantages that:
(1) the rolling stability control method applied to the high-dynamic aircraft provided by the invention not only controls the stability of the aircraft rolling angle and the rolling angular speed in the rolling channel, but also eliminates the influence caused by the dynamic lag of the actuating mechanism and improves the control precision;
(2) compared with the control method in the prior art, the roll stability control method applied to the high-dynamic aircraft can enable the roll angle rate of the aircraft to be rapidly converged;
(3) the roll stability control system applied to the high-dynamic aircraft is insensitive to aerodynamic interference, can quickly stabilize the aircraft system under different attack angles, and has good convergence under the condition of a large attack angle;
(4) the roll stability control system applied to the high-dynamic aircraft provided by the invention has strong universality, is suitable for both a roll aircraft and a non-roll aircraft, and can quickly converge to an expected value.
Drawings
FIG. 1 illustrates a variation of a coefficient of disturbance of an aircraft over a roll channel as a function of angle of attack;
FIG. 2 shows control curves of methods A-C for the roll angle of a high dynamic aircraft in Experimental example 1 of the present invention;
FIG. 3 shows control curves of methods A-C for roll rate of a high dynamic aircraft according to experimental example 1 of the present invention;
FIG. 4-1 shows the rolling angle variation curves of the aircraft controlled by the method A under different angles of attack in the uniform rolling state of the aircraft in the experimental example 2 of the present invention;
FIG. 4-2 shows a partial enlarged view of FIG. 4-1;
FIG. 5-1 shows the variation curve of the roll rate of the aircraft controlled by the method A under different angles of attack in the uniform roll state of the aircraft in the experimental example 2 of the present invention;
FIG. 5-2 shows a partial enlarged view of FIG. 5-1;
FIG. 6-1 shows the rolling angle variation curves of the aircraft controlled by the method A under different angles of attack in the non-rolling state of the aircraft in the experimental example 2 of the invention;
fig. 6-2 shows a partial enlarged view of 6-1.
Detailed Description
The present invention will be described in further detail below with reference to the accompanying drawings and embodiments. The features and advantages of the present invention will become more apparent from the description. In which, although various aspects of the embodiments are shown in the drawings, the drawings are not necessarily drawn to scale unless specifically indicated.
The invention provides a roll stability control method applied to a high-dynamic aircraft, which comprises the following steps:
step 1, measuring in real time to obtain the roll angle information of a high-dynamic aircraft;
and 2, calculating in real time by the aircraft control system according to the roll angle information obtained in the step 1 to obtain an instruction rudder deflection angle, transmitting the instruction rudder deflection angle to an aircraft executing mechanism, and further controlling the roll angle speed of the aircraft to be converged to a desired value.
After the aircraft control system transmits the command rudder deflection angle to the actuating mechanism and before the aircraft actually steers the rudder, the method also comprises the step of eliminating the dynamic lag of the actuating mechanism in real time.
And the command for eliminating the dynamic lag of the actuating mechanism is a compensated command rudder deflection angle, namely a final rudder-turning command of the aircraft steering engine.
In the invention, the high dynamic state means that the aircraft can fly with large maneuvering and has large normal acceleration (generally, the flying condition with the normal acceleration of more than 10g is called large maneuvering and g represents gravity acceleration); the actuating mechanism of the aircraft comprises a steering engine.
In the invention, the high dynamic aircraft refers to a rotating aircraft with the rotating speed of more than 10 r/s.
The roll stability control method applied to the high-dynamic aircraft according to the present invention is further described below.
Step 1, measuring in real time to obtain the roll angle information of the high-dynamic aircraft.
In the invention, the roll angle information of the high-dynamic aircraft is preferably obtained by real-time measurement of a gyroscope, wherein the gyroscope can directly measure and obtain the roll angle of the aircraft.
And 2, resolving in real time according to the roll angle information obtained in the step 1 to obtain an instruction rudder deflection angle, transmitting the instruction rudder deflection angle to an aircraft executing mechanism, and further controlling the roll angle speed of the aircraft to converge to an expected value.
During the flight of the aircraft, the dynamics of the rolling channel can be expressed by a second-order differential equation of the following formula (one):
Figure BDA0002181351960000053
wherein phi represents a roll angle,
Figure BDA0002181351960000051
the roll-angle rate is shown as,
Figure BDA0002181351960000052
representing the roll angle plus rate, deltaaIndicating rudder deflection angle, Cla140 denotes the disturbance factor, KRRoll rate bandwidth, K, is expressed as 5δRepresenting rudder deflection angle coefficient, Clasin4 phi denotes the perturbation term.
The inventors have found that, as shown in fig. 1, when the aircraft is flying at a small angle of attack, Clathe numerical value is small, the nonlinear term (disturbance term) can be ignored, and the controller in the formula (one) meets the use requirement; when the aircraft is flying at a large angle of attack, ClaLarge values and drastic changes, in which case Cl is present even at small roll anglesaIt will cause large disturbance to the system, and the non-linear term in equation (one) is not negligible. Therefore, the nonlinear term is considered in the process of solving the rudder deflection angle command in real time according to the measured roll angle information.
In the present invention, the calculating of the command rudder deflection angle includes the steps of:
and 2-1, selecting a proper system switching surface according to the roll angle information measured in real time after the starting and the control of the aircraft.
In the process of controlling the aircraft system, the controller is purposefully and continuously changed in a jump mode according to the current state of the system, and the system is forced to move according to a state track of a preset 'sliding mode'. The aircraft system changes the track motion state, and needs to select a proper switching surface, so that the determined sliding mode is asymptotically stable and has good quality.
The switching surface is a hypersurface which divides a state space into two parts, and the state quantity is stabilized on the hypersurface by designing the control of a controller on the state quantity.
According to a preferred embodiment of the present invention, the system switching plane is represented by the following formula (two):
s1＝e2+α1e1+α2β(e1) (II)
Wherein S is1Indicating a system switching plane (first switching plane), e1And e2Are all the variables of the state, and are,
e1＝φ-φ*phi denotes the roll angle, phi*To desired roll angle, e1For errors in the actual roll angle and the desired roll angle,
Figure BDA0002181351960000061
Figure BDA0002181351960000062
the roll-angle rate is shown as,
Figure BDA0002181351960000063
to desired roll rate, e2Error of actual and expected roll angular velocities;
α1and alpha2For design parameters, the value of α is generally chosen1Is 20 to 30, alpha 220 to 30, preferably, alpha1Is 25, alpha2Is 25;
β(e1) For the switching function, the following is expressed:
if it is not
Figure BDA0002181351960000064
Or
Figure BDA0002181351960000065
Then: beta (e)1)＝e1 p/q；
If it is not
Figure BDA0002181351960000066
Then:
Figure BDA0002181351960000067
wherein the content of the first and second substances,
Figure BDA0002181351960000068
values representing the switching surface, p, q being two odd integers, generally satisfying 1/2 < p/q < 1, preferably p being 3, q being 5, b1＝(2-p/q)μp/q-1， b2＝(p/q-1)μp/q-2，
Figure BDA0002181351960000069
Mu is a small constant, generally 0.003, sgn (·) represents a sign function.
The inventor researches and discovers that the selected system switching surface can enable the high-dynamic aircraft to approach smoothly and quickly.
And 2-2, calculating in real time to obtain an instruction rudder deflection angle, so that the roll angle of the aircraft reaches a stable state within a limited time.
According to a preferred embodiment of the present invention, the command rudder deflection angle is obtained by solving the following equation (three):
Figure BDA0002181351960000071
wherein u is1Indicating commanded rudder deflection angle, ClaRepresenting the coefficient of disturbance, ωRRRepresents the roll angular rate bandwidth, generally taking the value of 5;
e1and e2Are all state variables, e1＝φ-φ*Phi denotes the roll angle, phi*To desired roll angle, e1For errors in the actual and desired roll angles,
Figure BDA0002181351960000072
Figure BDA0002181351960000073
the roll rate is expressed in terms of the roll rate,
Figure BDA0002181351960000074
to desired roll rate, e2Error of the actual roll rate and the expected roll rate;
α1and alpha2For design parameters, the value of α is generally chosen1Is 20 to 30, alpha 220 to 30, preferably, alpha1Is 25, alpha2Is 25; s1Representing a system switching plane; k is a radical of1The value is generally 140 to 160, preferably 150, k2The value is generally 20-40, preferably 30, lambda1The value is generally 1-2, and preferably 0.75; kδThe rudder deflection angle coefficient is represented, and generally takes a value of 1200-2500, preferably 1400; sgn (·) denotes a sign function;
Figure BDA0002181351960000075
represents:
if it is not
Figure RE-GDA0002261344310000076
Or
Figure RE-GDA0002261344310000077
Then:
Figure RE-GDA0002261344310000078
if it is not
Figure RE-GDA0002261344310000079
Then:
Figure RE-GDA00022613443100000710
wherein the content of the first and second substances,
Figure BDA00021813519600000710
values representing the switching plane, p, q being two odd integers, generally satisfying 1/2 < p/q < 1, b1＝(2-p/q)μp/q-1，b2＝(p/q-1)μp/q-2，
Figure BDA00021813519600000711
Mu is a small constant, typically 0.003.
In the invention, the command rudder deflection angle calculation method is selected, so that the high-dynamic aircraft is insensitive to aerodynamic interference, the system can be quickly and stably under different attack angles, and the convergence condition is good under a large attack angle.
The inventor researches and discovers that the high-dynamic aircraft carries out real-time calculation on the measured roll angle information, and the control system of the high-dynamic aircraft obtains the command rudder deflection angle in real time through the above formula (II) and formula (III), so that the roll angle rate can be converged to the expected value quickly and stably. However, considering the time lag characteristics of the actuator dynamics, the actuator of the aircraft is difficult to match with the control system clock, resulting in control lag and reduced control accuracy.
Therefore, in the present invention, it is preferable to eliminate the time lag of the actuator dynamics after the aircraft control system transmits the command rudder deflection angle to the actuator and before the aircraft actually steers, so as to obtain the compensated command rudder deflection angle (i.e. the final steering command of the aircraft steering engine).
Wherein, the first-order lag model of the aircraft actuator is shown as the following formula (four):
Figure BDA0002181351960000081
wherein, deltaaThe rudder deflection angle is indicated and indicated,
Figure BDA0002181351960000082
representing the rudder deflection angle rate, deltacDenotes the command rudder deflection angle, and τ denotes the time lag constant, which is typically 0.4.
According to a preferred embodiment of the present invention, the compensated commanded rudder deflection angle is obtained by the following equation (five):
Figure BDA0002181351960000083
wherein u is2＝δcIndicating the commanded rudder deflection angle (post-compensation commanded rudder deflection angle), s2＝e3-u1Showing a second switching plane, e3＝δaIs a state variable, δaIndicating rudder deflection angle, u1As shown in formula (III); k is a radical of3、k4、λ2To design the parameter, k3The value is generally 20-40, preferably 30, k4The value is generally 20-40, preferably 30, lambda2The value is generally 0-1, preferably 0.75; τ represents the time lag constant, typically 0.4; sgn (·) denotes a sign function;
Figure BDA0002181351960000091
represents u1The first differential of the time domain is,
Figure BDA0002181351960000092
wherein the content of the first and second substances,
Figure BDA0002181351960000093
is represented as follows:
if it is not
Figure BDA0002181351960000094
Or
Figure BDA0002181351960000095
Then:
Figure BDA0002181351960000096
if it is not
Figure BDA0002181351960000097
Then:
Figure BDA0002181351960000098
wherein p and q are two odd integers, which satisfy 1/2 < p/q < 1, b1＝(2-p/q)μp/q-1，b2＝(p/q-1)μp/q-2，
Figure BDA0002181351960000099
Mu is a smaller constant, generally 0.003;
Figure BDA00021813519600000910
denotes e2First differential of (a).
The invention also provides a roll stability control system applied to the high-dynamic aircraft, which is preferably used for implementing the roll stability control method applied to the high-dynamic aircraft, and the system comprises a roll angle measuring module, an instruction rudder deflection angle resolving module and an instruction rudder deflection angle compensating module, wherein,
the roll angle measuring module is used for measuring the roll angle of the aircraft in real time,
the command rudder deflection angle resolving module is used for resolving and calculating in real time according to the roll angle information to obtain a command rudder deflection angle;
the command rudder deflection angle compensation module is used for eliminating the dynamic lag of an aircraft executing mechanism so as to obtain a compensated command rudder deflection angle.
According to a preferred embodiment of the invention, the roll angle measuring module is a gyroscope, and the roll angle information of the aircraft can be directly measured and obtained.
According to a preferred embodiment of the present invention, the command rudder deflection angle calculating module includes a switching plane selecting submodule and a real-time calculating submodule, wherein,
the switching surface selection submodule is used for selecting a system switching surface according to the roll angle information measured in real time after the starting and the control of the aircraft;
the real-time resolving submodule is used for resolving in real time to obtain an instruction rudder deflection angle, so that the rolling angle of the aircraft reaches a stable state within a limited time.
In a further preferred embodiment, the switching plane of the system selected by the switching plane selection sub-module is represented by formula (two):
s1＝e2+α1e1+α2β(e1) (II)
Wherein S is1Indicating a system switching plane (first switching plane), e1And e2Are all the variables of the state, and are,
e1＝φ-φ*phi denotes the roll angle, phi*To desired roll angle, e1For errors in the actual roll angle and the desired roll angle,
Figure BDA0002181351960000101
Figure BDA0002181351960000102
the roll-angle rate is shown as,
Figure BDA0002181351960000103
to desired roll rate, e2Error of actual and expected roll angular velocities;
α1and alpha2For design parameters, the value of α is generally chosen1Is 20 to 30, alpha 220 to 30, preferably, alpha1Is 25, alpha2Is 25;
β(e1) For the switching function, the following is expressed:
if it is not
Figure BDA0002181351960000104
Or
Figure BDA0002181351960000105
Then: beta (e)1)＝e1 p/q；
If it is not
Figure BDA0002181351960000106
Then:
Figure BDA0002181351960000107
wherein the content of the first and second substances,
Figure BDA0002181351960000108
values representing the switching plane, p, q being two odd integers, generally satisfying 1/2 < p/q < 1, preferably p being 3, q being 5, b1＝(2-p/q)μp/q-1， b2＝(p/q-1)μp/q-2，
Figure BDA0002181351960000109
Mu is a small constant, generally 0.003, sgn (·) represents a sign function.
In a further preferred embodiment, the real-time solution submodule obtains the command rudder deflection angle in real time by the following formula (three):
Figure BDA00021813519600001010
wherein u is1Indicating commanded rudder deflection angle, ClaRepresenting the coefficient of disturbance, ωRRRepresents the roll angular rate bandwidth, generally taking the value of 5;
e1and e2Are all state variables, e1＝φ-φ*Phi denotes the roll angle, phi*To desired roll angle, e1For errors in the actual and desired roll angles,
Figure BDA0002181351960000111
Figure BDA0002181351960000112
the roll rate is expressed in terms of the roll rate,
Figure BDA0002181351960000113
to desired roll rate, e2Error of the actual roll rate and the expected roll rate;
α1and alpha2For design parameters, the value of α is generally chosen1Is 20 to 30, alpha 220 to 30, preferably, alpha1Is 25, alpha2Is 25; s1Representing a system switching plane; k is a radical of1The value is generally 140 to 160, preferably 150, k2The value is generally 20-40, preferably 30, lambda1The value is generally 1-2, and preferably 0.75; kδThe rudder deflection angle coefficient is represented, and generally takes a value of 1200-2500, preferably 1400; sgn (·) denotes a sign function;
Figure BDA0002181351960000114
represents:
if it is not
Figure BDA0002181351960000115
Or
Figure BDA0002181351960000116
Then:
Figure BDA0002181351960000117
if it is not
Figure BDA0002181351960000118
Then:
Figure BDA0002181351960000119
wherein the content of the first and second substances,
Figure BDA00021813519600001110
values representing the switching plane, p, q being two odd integers, generally satisfying 1/2 < p/q < 1, preferably p being 3, q being 5, b1＝(2-p/q)μp/q-1， b2＝(p/q-1)μp/q-2，
Figure BDA00021813519600001111
Mu is a small constant, typically 0.003.
According to a preferred embodiment of the present invention, the command rudder deflection angle compensating module obtains the compensated command rudder deflection angle by the following equation (five):
Figure BDA00021813519600001112
wherein u is2＝δcIndicating the commanded rudder deflection angle (post-compensation commanded rudder deflection angle), s2＝e3-u1Showing a second switching plane, e3＝δaIs a state variable, δaIndicating rudder deflection angle, u1As shown in formula (III); k is a radical of3、k4、λ2To design the parameter, k3The value is generally 20-40, preferably 30, k4The value is generally 20-40, preferably 30, lambda2The value is generally 0-1, preferably 0.75; τ represents the time lag constant, typically 0.4; sgn (·) denotes a sign function;
Figure BDA0002181351960000121
to representu1The first differential of the time domain is,
Figure BDA0002181351960000122
wherein the content of the first and second substances,
Figure BDA0002181351960000123
is represented as follows:
if it is not
Figure BDA0002181351960000124
Or
Figure BDA0002181351960000125
Then:
Figure BDA0002181351960000126
if it is not
Figure BDA0002181351960000127
Then:
Figure BDA0002181351960000128
wherein p and q are two odd integers, generally satisfying 1/2 < p/q < 1, preferably p is 3, q is 5, b1＝(2-p/q)μp/q-1，b2＝(p/q-1)μp/q-2，
Figure BDA0002181351960000129
Mu is a smaller constant, generally 0.003;
Figure BDA00021813519600001210
denotes e2First differential of (a).
In the invention, the system for controlling the rolling stability can eliminate the uncertainty and nonlinearity of the aerodynamic parameters, can quickly stabilize the aircraft system under different attack angles, and has good convergence under the condition of large attack angle; and the universal property is strong, the aircraft is suitable for both rolling aircrafts and non-rolling aircrafts, and both can be quickly converged to expected values.
Examples of the experiments
Experimental example 1
The rolling stability control simulation true experiment of the high dynamic aircraft is carried out through a computer, and the simulation conditions are as follows: high dynamic aircraft roll steadily at a speed of 10r/s, i.e.
Figure BDA00021813519600001211
The control conditions of the three control methods (method A, method B and method C) on the roll angle and the roll angle rate of the high-dynamic aircraft are simulated respectively, the roll angle change result when the attack angle is 16 degrees is shown in figure 2, and the roll angle rate change result when the attack angle is 16 degrees is shown in figure 3.
The method A is a roll stability control method applied to the high-dynamic aircraft, and specifically comprises the following steps:
(1) after the starting and control of the aircraft, selecting a proper system switching surface according to the roll angle information measured in real time, wherein the system switching surface is shown as the following formula (II):
s1＝e2+α1e1+α2β(e1) (II)
Wherein S is1Indicating a system switching plane (first switching plane), e1And e2Are all the variables of the state, and are,
e1＝φ-φ*phi denotes the roll angle, phi*To desired roll angle, e1For errors in the actual roll angle and the desired roll angle,
Figure BDA0002181351960000131
Figure BDA0002181351960000132
the roll-angle rate is shown as,
Figure BDA0002181351960000133
to desired roll rate, e2Is a reality ofError in roll rate and desired roll rate;
α1and alpha2To design the parameter, α1Is 25, alpha2Is 25;
β(e1) For the switching function, the following is expressed:
if it is not
Figure BDA0002181351960000134
Or
Figure BDA0002181351960000135
Then: beta (e)1)＝e1 p/q；
If it is not
Figure BDA0002181351960000136
Then:
Figure BDA0002181351960000137
wherein the content of the first and second substances,
Figure BDA0002181351960000138
values representing the switching plane, p is 3, q is 5, b1＝(2-p/q)μp/q-1， b2＝(p/q-1)μp/q-2，
Figure BDA0002181351960000139
μ is 0.003, sgn (. cndot.) represents a sign function.
(2) And (3) calculating the command rudder deflection angle in real time according to the following formula (III):
Figure BDA00021813519600001310
wherein u is1Indicating commanded rudder deflection angle, ClaRepresenting the coefficient of disturbance, ωRRThe bandwidth of the roll angular rate is represented and is 5;
e1and e2Are all state variables, e1＝φ-φ*Phi denotes the roll angle, phi*To a desired roll angle，e1For errors in the actual and desired roll angles,
Figure BDA0002181351960000141
Figure BDA0002181351960000142
the roll rate is expressed in terms of the roll rate,
Figure BDA0002181351960000143
to desired roll rate, e2Error of the actual roll rate and the expected roll rate;
α1and alpha2To design the parameter, α1Is 25, alpha2Is 25; s1Representing a system switching plane; k is a radical of1、k2、λ1To design the parameter, k1Is 150, k2Is 30, λ1Is 0.75; kδThe rudder deflection angle coefficient is 1400; sgn (·) denotes a sign function;
Figure BDA0002181351960000144
represents:
if it is not
Figure BDA0002181351960000145
Or
Figure BDA0002181351960000146
Then:
Figure BDA00021813519600001418
if it is not
Figure BDA0002181351960000147
Then:
Figure BDA0002181351960000148
wherein the content of the first and second substances,
Figure BDA0002181351960000149
values representing the switching plane, p is 3, q is 5, b1＝(2-p/q)μp/q-1， b2＝(p/q-1)μp/q-2，
Figure BDA00021813519600001410
μ is 0.003.
(3) Obtaining a compensated command rudder deflection angle according to the following formula (five):
Figure BDA00021813519600001411
wherein u is2＝δcIndicating the commanded rudder deflection angle (post-compensation commanded rudder deflection angle), s2＝e3-u1Showing a second switching plane, e3＝δaIs a state variable, δaIndicating rudder deflection angle, u1As shown in formula (III); k is a radical of3、k4、λ2To design the parameter, k3Is 30, k4Is 30, λ2Is 0.75; τ represents a time lag constant of 0.4; sgn (·) denotes a sign function;
Figure BDA00021813519600001412
represents u1The first differential of the time domain is,
Figure BDA00021813519600001413
wherein the content of the first and second substances,
Figure BDA00021813519600001414
is represented as follows:
if it is not
Figure BDA00021813519600001415
Or
Figure BDA00021813519600001416
Then:
Figure BDA00021813519600001417
if it is not
Figure BDA0002181351960000151
Then:
Figure BDA0002181351960000152
wherein p is 3, q is 5, b1＝(2-p/q)μp/q-1，
Figure BDA0002181351960000153
Figure BDA0002181351960000154
Mu is 0.003;
Figure BDA0002181351960000155
denotes e2First differential of (a).
The method B is a method for controlling the roll stability of the high-dynamic aircraft based on the design of a linear sliding mode control theory (LSM), wherein a rudder deflection angle instruction of the aircraft is obtained by the following formula (seven):
Figure BDA0002181351960000156
wherein s is3Showing a switching plane e2Indicating the roll angle, e1Representing the roll angular velocity, rho representing the design parameter, and taking the value of 1, k5Represents design parameters with a value of 100, omegaRRRepresents the roll rate bandwidth with the value of 5, KδIndicating a rudder deflection coefficient of 1400.
The method C is a method for controlling the roll stability of the high-dynamic aircraft based on a terminal sliding mode control Theory (TSM), wherein a rudder deflection angle instruction of the aircraft is obtained by the following formula (eight):
Figure BDA0002181351960000157
wherein s is4Indicating a rudder deflection angle command, e2Indicating the roll angle, e1Showing the roll angular velocity, rho showing the design parameter, with the value of 1, a showing the design parameter, with the value of 3, b showing the design parameter, with the value of 5, k6Represents design parameters with a value of 100, omegaRRRepresents the roll rate bandwidth with the value of 5, KδIndicating a rudder deflection coefficient of 1400.
As can be seen from FIGS. 2 and 3, the roll angle curves controlled by methods B and C have a tendency to converge in an infinite time, while the roll angle controlled by method A can converge in a finite time and at a fast rate of convergence.
Experimental example 2
Under different attack angle conditions (2 degrees, 6 degrees, 10 degrees, 14 degrees and 16 degrees), the changes of the roll angle and the roll angle rate controlled by the method A are simulated, the constant-speed roll state (10r/s) and the no-roll state of the aircraft are respectively considered,
the simulation results of the uniform rolling state of the aircraft are shown in FIGS. 4-1, 4-2, 5-1 and 5-2; the simulation results for the no-roll state are shown in FIGS. 6-1 and 6-2.
As can be seen from fig. 4-1, in the uniform rolling state of the aircraft, the rolling angle of the aircraft changes linearly after being stabilized, which indicates that the projectile is in the stable uniform rolling state, and the changing trajectories are substantially consistent under different attack angles; as can be seen from fig. 4-2 (enlarged view), the trajectories of the projectile change at the attack angles of 14 ° and 16 ° coincide.
As can be seen from FIG. 5-1, in the uniform roll state of the aircraft, the roll rate of the projectile reaches the desired roll rate at about 0.5s, the convergence rate is fast, and the convergence trajectories of the roll rates are almost the same under different attack angles.
As can be seen from fig. 5-2, method a (the roll stability control method of the present invention applied to a high dynamic aircraft) can accommodate severe changes in aerodynamic parameters; the convergence trajectories of the roll rate of the projectile at angles of attack of 10 °, 14 ° and 16 ° coincide.
As can be seen from FIG. 6-1, in the non-rolling state of the aircraft, the method A can control the aircraft to converge the rolling angular velocity to zero within a limited time under different attack angle flight conditions, i.e., the aircraft does not roll; as can be seen from fig. 6-2, the trajectories of the projectile changes at angles of attack of 10 °, 14 ° and 16 ° coincide.
The present invention has been described above in connection with preferred embodiments, which are merely exemplary and illustrative. On the basis of the above, the invention can be subjected to various substitutions and modifications, and the substitutions and the modifications are all within the protection scope of the invention.

Claims (10)

1. A roll stability control method for a high dynamic aircraft, the method comprising the steps of:
step 1, measuring in real time to obtain the roll angle information of a high-dynamic aircraft;
step 2, the aircraft control system calculates in real time according to the roll angle information obtained in the step 1 to obtain an instruction rudder deflection angle, and transmits the instruction rudder deflection angle to an aircraft executing mechanism so as to control the roll angle rate of the aircraft to converge to a desired value;
in the step 2, after the aircraft control system transmits the command rudder deflection angle to the actuating mechanism and before the aircraft actually steers, the method further comprises the steps of eliminating the dynamic lag of the actuating mechanism in real time and obtaining the compensated command rudder deflection angle.
2. The method of claim 1, wherein in step 2, the calculating of the commanded rudder deflection angle comprises the steps of:
step 2-1, after the starting and control of the aircraft, selecting a proper system switching surface according to the roll angle information measured in real time;
and 2-2, calculating in real time to obtain an instruction rudder deflection angle, so that the aircraft approaches to a switching surface within a limited time.
3. The method according to claim 2, wherein in step 2-1, the selected system switching plane is represented by the following formula (two):
s1＝e2+α1e1+α2β(e1) (II)
Wherein S is1Indicating system switching plane, e1And e2Are all state variables; e.g. of the type1＝φ-φ*Phi denotes the roll angle, phi*To desired roll angle, e1For errors in the actual and desired roll angles,
Figure FDA0002181351950000011
Figure FDA0002181351950000012
the roll-angle rate is shown as,
Figure FDA0002181351950000013
to desired roll rate, e2Error of the actual roll rate and the expected roll rate;
α1and alpha2For design parameters, the value of α is generally chosen1Is 20 to 30, alpha220 to 30;
β(e1) For the switching function, the following is expressed:
if it is not
Figure FDA0002181351950000016
Or
Figure FDA0002181351950000017
|e1If | ≧ μ:
Figure FDA0002181351950000014
if it is not
Figure FDA0002181351950000018
|e1If | < mu, then:
Figure FDA0002181351950000015
4. the method according to claim 2, wherein in step 2-2, the commanded rudder deflection angle is obtained by solving the following equation (three):
Figure FDA0002181351950000021
wherein u is1Indicating commanded rudder deflection angle, ClaRepresenting the coefficient of disturbance, ωRRThe roll rate bandwidth is represented and generally takes a value of 5;
e1and e2Are all state variables, e1＝φ-φ*Phi denotes the roll angle, phi*To desired roll angle, e1For errors in the actual and desired roll angles,
Figure FDA0002181351950000022
Figure FDA0002181351950000023
the roll-angle rate is shown as,
Figure FDA0002181351950000024
to desired roll rate, e2Error of the actual roll rate and the expected roll rate;
α1and alpha2For design parameters, the value of α is generally chosen1Is 20 to 30, alpha220 to 30; k is a radical of1、k2、λ1To design the parameter, k1The general value of k is 140-1602The general value is 20-40 lambda1The general value is 1-2, KδThe rudder deflection angle coefficient is represented, and the value is generally 1200-2500; sgn (·) represents a sign function.
5. The method according to claim 4, wherein in the formula (III),
Figure FDA0002181351950000025
represents:
if it is not
Figure FDA00021813519500000210
Or
Figure FDA00021813519500000211
|e1If | ≧ μ:
Figure FDA0002181351950000026
if it is not
Figure FDA00021813519500000212
|e1If | < mu, then:
Figure FDA0002181351950000027
wherein the content of the first and second substances,
Figure FDA00021813519500000213
representing the value of the switching surface, p, q are two odd integers satisfying 1/2 < p/q < 1, b1＝(2-p/q)μp/q-1b2＝(p/q-1)μp/q-2
Figure FDA0002181351950000028
Mu is a small constant, typically 0.003.
6. The method of claim 5, wherein the compensated commanded rudder deflection angle is obtained by the equation (V):
Figure FDA0002181351950000029
wherein u is2Indicating the compensated commanded rudder deflection angle, s2＝e3-u1Indicating a second switching plane，e3＝δaIs a state variable, δaIndicating rudder deflection angle, k3、k4、λ2To design the parameter, k3The general value is 20-40, k4The general value is 20-40 lambda2The value is generally 0-1; τ represents the time lag constant, typically 0.4; sgn (·) represents a sign function.
7. Roll stability control system for high dynamic aircraft, preferably for implementing the method according to one of claims 1 to 6, characterized in that the system comprises a roll angle measurement module, a commanded rudder deflection angle calculation module and a commanded rudder deflection angle compensation module, wherein,
the roll angle measuring module is used for measuring the roll angle of the aircraft in real time,
the command rudder deflection angle resolving module is used for resolving in real time according to the roll angle information to obtain a command rudder deflection angle;
the command rudder deflection angle compensation module is used for eliminating the dynamic lag of an aircraft executing mechanism so as to obtain a compensated command rudder deflection angle.
8. The system of claim 7 wherein the roll angle measurement module is a gyroscope to directly measure roll angle information for the aircraft.
9. The system of claim 7, wherein the commanded rudder deflection angle solution module comprises a switching plane selection submodule and a real-time solution submodule, wherein,
the switching surface selection submodule is used for selecting a system switching surface according to the roll angle information measured in real time after the starting and the control of the aircraft;
and the real-time calculating submodule is used for calculating in real time to obtain an instruction rudder deflection angle so that the aircraft approaches to a switching surface within a limited time.
10. The system of claim 7, wherein the commanded rudder deflection compensation module obtains the compensated commanded rudder deflection angle by the equation (five):
Figure FDA0002181351950000031
wherein u is2Indicating the compensated commanded rudder deflection angle, s2＝e3-u1Showing a second switching plane, e3＝δaIs a state variable, δaIndicating rudder deflection angle, k3、k4、λ2To design the parameter, k3The general value is 25-35, k4The general value is 25-35, lambda2The value is generally 0.5-1; τ represents the time lag constant, typically 0.4; sgn (·) represents a sign function.
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Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113791635A (en) * 2021-08-24 2021-12-14 中国航天科工集团第二研究院 Aircraft rolling control method and system and electronic equipment
CN115903875A (en) * 2023-01-05 2023-04-04 北京航天众信科技有限公司 Effective security control method for target attitude instability

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN106325291A (en) * 2016-10-10 2017-01-11 上海拓攻机器人有限公司 Four-rotor aircraft attitude control method and system based on sliding-mode control law and ESO
CN109596011A (en) * 2018-12-07 2019-04-09 上海机电工程研究所 The stable canard configuration guided missile overall architecture of rolling racemization
CN109992003A (en) * 2019-04-29 2019-07-09 北京理工大学 Robustness roll angle method of rate control and system
CN110109473A (en) * 2019-04-29 2019-08-09 北京理工大学 It can be applied to the non-linear roll stabilization control system and method for Rotary aircraft

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN106325291A (en) * 2016-10-10 2017-01-11 上海拓攻机器人有限公司 Four-rotor aircraft attitude control method and system based on sliding-mode control law and ESO
CN109596011A (en) * 2018-12-07 2019-04-09 上海机电工程研究所 The stable canard configuration guided missile overall architecture of rolling racemization
CN109992003A (en) * 2019-04-29 2019-07-09 北京理工大学 Robustness roll angle method of rate control and system
CN110109473A (en) * 2019-04-29 2019-08-09 北京理工大学 It can be applied to the non-linear roll stabilization control system and method for Rotary aircraft

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
王伟 等: "基于二阶滑模控制理论的新型滚转稳定控制器", 《控制与决策》 *

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113791635A (en) * 2021-08-24 2021-12-14 中国航天科工集团第二研究院 Aircraft rolling control method and system and electronic equipment
CN113791635B (en) * 2021-08-24 2023-07-25 中国航天科工集团第二研究院 Aircraft rolling control method, system and electronic equipment
CN115903875A (en) * 2023-01-05 2023-04-04 北京航天众信科技有限公司 Effective security control method for target attitude instability
CN115903875B (en) * 2023-01-05 2023-04-28 北京航天众信科技有限公司 Effective An Kongfang method for unstable target posture

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