CN111045440B - Hypersonic aircraft nose-down section rapid rolling control method - Google Patents

Hypersonic aircraft nose-down section rapid rolling control method Download PDF

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CN111045440B
CN111045440B CN201911294155.XA CN201911294155A CN111045440B CN 111045440 B CN111045440 B CN 111045440B CN 201911294155 A CN201911294155 A CN 201911294155A CN 111045440 B CN111045440 B CN 111045440B
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coupling
attitude
aircraft
angle
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CN111045440A (en
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胡庆雷
曹瑞浩
董宏洋
郑建英
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Beihang University
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    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
    • G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
    • G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft

Abstract

The invention discloses a hypersonic aircraft nose-down section fast roll control method, which comprises the following steps: establishing a control-oriented hypersonic aircraft attitude coupling model; establishing a three-channel attitude error dynamic model, and defining a coupling profit-and-loss judgment criterion based on control performance; designing a three-channel attitude coupling controller by using the defined coupling profit and disadvantage judgment criterion; considering the strong interference characteristics brought by unmodeled dynamics of the aircraft, external disturbance and the like, designing a gain self-adaptive method with anti-interference capability; and finally, obtaining a high-precision robust control strategy based on favorable coupling driving. The robust control strategy based on favorable coupling drive is adopted, the model does not need to be linearized, a channel-based controller does not need to be designed, the known interference upper bound information is not needed, the robust fast response attitude control method is suitable for the hypersonic aircraft nose-down section robust fast response attitude control task, and has higher reliability and practical value.

Description

Hypersonic aircraft nose-down section rapid rolling control method
Technical Field
The invention belongs to the field of design of a control system of a hypersonic aircraft, and particularly relates to a method for controlling rapid rolling of the hypersonic aircraft in a nose-down section.
Background
The hypersonic aircraft generally adopts a mode of bank turning in a diving section, so that a main lifting surface of the hypersonic aircraft can be quickly aligned to an overload demand direction to realize quick pressing and quick lateral maneuvering, and therefore enough large-range maneuvering capacity and striking capacity are obtained. The control mode brings higher requirements for high-performance tracking of the attitude system, namely fast tracking of an attack angle, fast zeroing of a sideslip angle and fast overturning of a roll angle. However, in the nose-down section, the high dynamics of the roll channel introduce strong cross-coupling effects between the pitch, yaw and roll channels, which present challenges for coupling control. In the actual control, the control mode is very easy to cause divergence of the attitude system, so that the reentry task fails. Therefore, it is necessary to study how to achieve high-performance attitude tracking capability while ensuring fast maneuvers.
Aiming at the problem of strong coupling, the current processing idea is mainly to adopt decoupling processing or regard a coupling item as a non-matched bounded disturbance item, and adopt sliding mode control and H∞Robust control methods such as control or control based on interference observation suppress the influence of the robust control methods on the system. The ideas are to directly offset and inhibit the coupling of the system, are essentially passive processing of the coupling, and lack deep analysis of influence relation between coupling information of the hypersonic vehicle and system stability and performance, so that the ideas have great conservation and do not fully exploit high-performance control capability of the system.
Disclosure of Invention
Therefore, the hypersonic aircraft coupling control problem is analyzed more comprehensively and scientifically, a coupling profit-and-disadvantage judgment criterion based on control performance is given, a three-channel attitude coupling controller is designed by utilizing the criterion, and then a control gain self-adaption method aiming at interference suppression is designed, so that a high-precision robust control strategy based on favorable coupling driving is obtained. Different from the existing method, the method can not only realize the rapid tracking of the attitude command of the dive section of the hypersonic aircraft while overcoming the existing conservative property, but also does not need to linearize the model, does not need to design a controller by channels, does not need to know the interference upper bound information, is suitable for the strong robust and rapid response attitude control task of the dive section of the hypersonic aircraft, and has higher reliability and practical value.
The invention provides a hypersonic aircraft nose-down section fast roll control method, which comprises the following steps:
s1: establishing a control-oriented hypersonic aircraft attitude coupling model;
s2: obtaining a three-channel attitude error dynamic model, and giving a coupling profit-and-defect judgment criterion based on control performance;
s3: designing a three-channel attitude coupling controller by using the defined coupling profit and disadvantage judgment criterion;
s4: considering the strong interference characteristics brought by unmodeled dynamics of the aircraft, external disturbance and the like, designing a gain self-adaptive method with anti-interference capability;
s5: and obtaining a high-precision robust control strategy based on favorable coupling driving.
Further, the attitude coupling model of the control-oriented hypersonic aircraft established in step S1 is:
Figure BDA0002320035230000021
in the formula, x1＝[γv,β,α]T,γvRepresents roll angle, β represents slip angle, α represents angle of attack; x is the number of2＝[ωx,ωy,ωz]T,ωx,ωy,ωzRespectively representing the rolling rate, the yawing rate and the pitching angle rate of the aircraft rotating around the three axes of the body coordinate system; u ═ 2x,y,z]，x,y,zRespectively representing the roll, yaw and pitch rudder deflection of the aircraft; the interference experienced by the system is denoted as D1＝[D11,D12,D13]T,D2＝[D21,D22,D23]T,D11,D12,D13,D21,D22,D23Respectively represents the uncertainty of the six channels caused by unmodeled dynamics, the uncertainty of the aerodynamic coefficient and the external disturbance, and
Figure BDA0002320035230000022
Figure BDA0002320035230000023
is a scalar quantity, the size is unknown; matrix F11,F12,F21,F22And B is respectively represented as:
Figure BDA0002320035230000024
Figure BDA0002320035230000025
Figure BDA0002320035230000031
wherein q is 0.5 ρ V2Representing dynamic pressure, ρ representing atmospheric density, and V representing aircraft speed; srefRepresenting a reference area of the aircraft; l represents a reference length; j. the design is a squarex,Jy,JzRepresenting the rotational inertia of the aircraft relative to three axes of a body coordinate system; θ represents a ballistic dip angle; m represents the mass of the aircraft; g represents the gravitational acceleration; cLAnd CZRespectively representing a lift coefficient and a lateral force coefficient; m isx、myAnd mzRespectively representing a roll moment coefficient, a yaw moment coefficient and a pitch moment coefficient; cL0Represents a zero lift coefficient;
Figure BDA0002320035230000032
respectively, the pairs of lift coefficients a are indicated,zpartial derivatives of (d);
Figure BDA0002320035230000033
respectively, represent the lateral force coefficient pair beta,x,ypartial derivatives of (d);
Figure BDA0002320035230000034
respectively representing the roll torque coefficients with respect to beta,x,ypartial derivatives of (d);
Figure BDA0002320035230000035
respectively, the yaw moment coefficient pair beta is represented,x,ypartial derivatives of (d);
Figure BDA0002320035230000036
respectively, represent the pitch moment coefficient pair a,zpartial derivatives of (a).
Further, the three-channel attitude error dynamic model established in step S2 includes the following specific procedures:
let the attitude angle tracking command be x1c＝[γvc,βc,αc]T，γvc,βc,αcRespectively representing a roll angle command value, a sideslip angle command value and an attack angle command value; virtual control x is introduced based on thought of backstepping method2d∈R3×1And defining a tracking error e1＝x1-x1c,e2＝x2-x2dFor tracking error e1And e2And (5) obtaining a derivative and combining the formula (4) to obtain an error state equation:
Figure BDA0002320035230000037
wherein, F12e2,F22e1∈R3×1Is the coupling term between the attitude angle subsystem and the angular rate subsystem.
Further, the coupled pros and cons determination criteria based on the control performance defined in step S2 are:
defining coupled profit-fraud decision factor ζij＝eij(Fi2e3-i)j∈R，eijFor tracking error e1And e21,2, j 1,2,3, when ζij＝eij(Fi2e3-i)jIf < 0, the coupled term is determined (F)i2e3-i)jThe influence on the system is favorable, otherwise the coupling term (F) is determinedi2e3-i)jThe impact on the system is detrimental.
Further, the specific process of designing the three-channel attitude coupling controller in step S3 is as follows:
to ensure tracking error e2When 0, the control law u is designed as follows:
u＝ueq+usw (6)
wherein u iseq,uswRespectively representing an equivalent control law and a discontinuous control law;
in the formula (5)
Figure BDA0002320035230000038
Obtain equivalent control
Figure BDA0002320035230000039
Figure BDA00023200352300000310
In conjunction with the coupled prosperity determination factor ζ defined in step S2ijWill equivalently control
Figure BDA0002320035230000041
The improvement is as follows:
Figure BDA0002320035230000042
wherein S is2＝diag{sgn(ζ21)，sgn(ζ22)，sgn(ζ23) }, sgn (·) denotes a sign function;
meanwhile, designing a discontinuous control law uswThe form is as follows:
Figure BDA0002320035230000043
wherein, K2To control the gain;
the control law u is obtained by substituting the formula (8) and the formula (9) into the formula (6)
Figure BDA0002320035230000044
To ensure tracking error e1Design the virtual control law x as 02dIs composed of
Figure BDA0002320035230000045
Wherein, K1To control the gain;
the designed three-channel attitude coupling controller comprises the following components:
Figure BDA0002320035230000046
further, the gain adaptive method with interference rejection capability designed in step S4 is:
Figure BDA0002320035230000047
wherein, i is 1,2,
Figure BDA0002320035230000048
and K0iIs a positive constant, KBiThe form of (A) is as follows:
Figure BDA0002320035230000049
where ρ isi,ηi(i ═ 1,2) as a design parameter, ρiGreater than 0 is a predetermined small constant, etaiIs constant and ηiIs greater than 1. Then, with the equations (13) and (14), when the tracking error of the system is large, i.e., | | ei||＞ρi/ηiAdaptive gain K of the controlleriIs equal to KAiWith KAiThe tracking error of the system is gradually reduced due to the continuous increase of the tracking error; until tracking error eiThe | | | is reduced to satisfy | | | ei||≤ρi/ηiSwitching is performed, when adaptive gain K is appliediIs equal to KBiAt | | | ei||∈[0，ρi/ηi]In the region of (A) KBiAlong with the system tracking error eiThe reduction of | decreases.
The robust control strategy based on the advantageous coupling driving obtained in step S5 is:
Figure BDA0002320035230000051
Figure BDA0002320035230000052
Figure BDA0002320035230000053
the invention has the beneficial effects that:
(1) the robust control strategy designed by the invention can realize the strong robust quick response attitude control of the hypersonic aerocraft in the dive section.
(2) The invention adopts a robust control strategy based on favorable coupling drive, does not need to linearize the model, does not need to design a controller by channels, does not need to know the interference upper bound information, and has higher reliability and practical value.
Drawings
FIG. 1 is a flow chart of a hypersonic aircraft nose-down section fast roll control method of the present invention;
FIG. 2 is a schematic diagram illustrating simulation of control effect in a nominal state by using the control method of the present invention;
FIG. 3 is a schematic view of rudder deflection angle at a nominal state using the control method of the present invention;
FIG. 4 is a comparison graph of the control effect of the control method of the present invention and the control method of the prior art respectively under the condition of pneumatic parameter bias;
FIG. 5 shows the adaptive gain K under the condition of the deviation of the pneumatic parameter by using the control method of the present invention1,K2A schematic diagram of (a);
fig. 6 is a schematic view of rudder deflection angle under the condition of pneumatic parameter deflection by using the control method of the invention.
Detailed Description
The technical solutions of the present invention will be described clearly and completely with reference to the accompanying drawings and embodiments, and it is to be understood that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments that can be derived by one of ordinary skill in the art from the embodiments given herein are intended to be within the scope of the present invention.
As shown in FIG. 1, the hypersonic aircraft nose-down section fast roll control method of the invention comprises the following steps:
s1: establishing control-oriented hypersonic aircraft attitude coupling model
Firstly, establishing an attitude motion equation of the hypersonic aircraft:
Figure BDA0002320035230000061
wherein, γvBeta, alpha are roll angle, sideslip angle and angle of attack, respectively; omegax,ωy,ωzRoll, yaw and pitch rates of the aircraft rotating around the three axes of the body coordinate system respectively; g is the acceleration of gravity; θ is the ballistic dip; d11,D12,D13And D21,D22,D23Representing disturbances (including external disturbances, unmodeled dynamics of the system, perturbation of aircraft parameters, etc.); l and Z are the lift and lateral forces to which the aircraft is subjected, respectively; mx,My,MzRespectively representing the roll, yaw and pitch moments to which the aircraft is subjected; j. the design is a squarex,Jy,JzRepresenting the rotational inertia of the aircraft relative to three axes of a body coordinate system; m represents the mass of the aircraft; v denotes the speed of the aircraft.
The expression aerodynamic and aerodynamic moments can be written as:
Figure BDA0002320035230000062
Figure BDA0002320035230000063
wherein q is 0.5 ρ V2Represents a dynamic pressure; srefRepresenting a reference area of the aircraft; l represents a reference length; cLAnd CZRespectively representing a lift coefficient and a lateral force coefficient; m isx、myAnd mzRespectively representing a roll moment coefficient, a yaw moment coefficient and a pitch moment coefficient;x、yandzrespectively representing the roll, yaw and pitch rudder deflection of the aircraft; cL0Represents a zero lift coefficient;
Figure BDA0002320035230000064
respectively, the pairs of lift coefficients a are indicated,zpartial derivatives of (d);
Figure BDA0002320035230000071
respectively, represent the lateral force coefficient pair beta,x,ypartial derivatives of (d);
Figure BDA0002320035230000072
respectively representing the roll torque coefficients with respect to beta,x,ypartial derivatives of (d);
Figure BDA0002320035230000073
respectively, the yaw moment coefficient pair beta is represented,x,ypartial derivatives of (d);
Figure BDA0002320035230000074
respectively, represent the pitch moment coefficient pair a,zpartial derivatives of (a).
By combining the formula (1), the formula (2) and the formula (3), the attitude coupling model of the hypersonic aerocraft facing the control can be obtained as
Figure BDA0002320035230000075
In the formula, x1＝[γv,β,α]T,x2＝[ωx,ωy,ωz]T,u＝[x,y,z]The interference experienced by the system is denoted as D1＝[D11,D12,D13]T,D2＝[D21,D22,D23]TAnd is and
Figure BDA0002320035230000076
is unknown. Matrix F11,F12,F21,F22And B is respectively represented as:
Figure BDA0002320035230000077
Figure BDA0002320035230000078
Figure BDA0002320035230000079
s2: based on the established control-oriented hypersonic aircraft attitude coupling model, a three-channel attitude error dynamic model is established, and a coupling profit-and-disadvantage judgment criterion based on control performance is defined, which is specifically as follows:
let the attitude angle tracking command be x1c＝[γvc,βc,αc]T，γvc,βc,αcCommand values representing a roll angle, a sideslip angle, and an angle of attack, respectively; and a virtual control law x is introduced based on the thought of a backstepping method2d∈R3×1And defining a tracking error e1＝x1-x1c,e2＝x2-x2dTo e is aligned with1And e2By taking the derivative in conjunction with equation (4), the following error state equation can be obtained:
Figure BDA00023200352300000710
wherein, F12e2,F22e1For the coupling term between the two subsystems, the coupling term of the system will have an advantageous or detrimental effect on the stability and dynamics of the system.
Defining the coupled prosperity judgment criterion based on the control performance as follows:
set ζij＝eij(Fi2e3-i)j∈R，eijFor tracking error e1And e21,2, j 1,2,3, when ζij＝eij(Fi2e3-i)jIf < 0, the coupled term is determined (F)i2e3-i)jThe influence on the system is favorable at this time, otherwise the coupling term (F) is determinedi2e3-i)jWhere the effect on the system is detrimental.
S3: and designing a three-channel attitude coupling controller by using the designed coupling profit and disadvantage judgment criterion. The specific process is as follows:
taking the attitude angular rate subsystem as an example, to ensure the tracking error e2And (5) designing a three-channel attitude coupling control law u based on the coupling prosperity judgment criterion as follows:
u＝ueq+usw (6)
wherein u iseq,uswRespectively representing an equivalent control law and a discontinuous control law.
First order in formula (5)
Figure BDA0002320035230000081
Obtain equivalent control
Figure BDA0002320035230000082
Figure BDA0002320035230000083
Wherein D is2Representing interference terms, which are not considered when designing a nominal controller; f22e1For the coupling item between two subsystems, considering that the influence of the coupling item on the stability and the dynamic performance of the system is not necessarily adverse, the judgment on the influence of the coupling item is needed, and the coupling item F is judged by utilizing the defined coupling prosperity judgment factor22e1And (4) processing. Will equivalently control
Figure BDA0002320035230000084
The improvement is as follows:
Figure BDA0002320035230000085
wherein S is2＝diag{sgn(ζ21)，sgn(ζ22)，sgn(ζ23)}∈R3×3。
Meanwhile, designing a discontinuous guidance law uswThe form is as follows:
Figure BDA0002320035230000086
wherein, K2To control the gain.
The formula (8) and the formula (9) are substituted into the formula (6), and the three-channel attitude coupling control law u based on the coupling prosperity judgment criterion is obtained
Figure BDA0002320035230000087
Similarly, for attitude angle system, to ensure tracking error e1Designing the three-channel attitude coupling virtual control law x based on the coupling prosperity judgment criterion2dIs composed of
Figure BDA0002320035230000088
In the formula, S1＝diag{sgn(ζ11)，sgn(ζ12)，sgn(ζ13)}∈R3×3(ii) a Therein, ζ1j(j ═ 1,2,3) is the defined coupling prosperity decision factor; k1To control the gain.
Thus, the three-channel attitude coupling controller can be obtained as follows:
Figure BDA0002320035230000091
s4: considering the strong interference characteristics brought by unmodeled dynamics of an aircraft, external disturbance and the like, the gain self-adaptive method with the anti-interference capability is designed as follows:
Figure BDA0002320035230000092
Figure BDA0002320035230000093
and K0iIs a positive constant, KBiThe form of (A) is as follows:
Figure BDA0002320035230000094
where ρ isi,ηi(i ═ 1,2) as a design parameter, ρiGreater than 0 is a predetermined small constant, etaiIs constant and ηiIs more than 1; then, with the equations (13) and (14), when the tracking error of the system is large, i.e., | | ei||＞ρi/ηiAdaptive gain K of the controlleriIs equal to KAiWith KAiThe tracking error of the system is gradually reduced due to the continuous increase of the tracking error; until tracking error eiThe | | | is reduced to satisfy | | | ei||≤ρi/ηiSwitching is performed, when adaptive gain K is appliediIs equal to KBiAt | | | ei||∈[0，ρi/ηi]In the region of (A) KBiAlong with the system tracking error eiThe reduction of | decreases.
And finally, obtaining a high-precision robust control strategy based on favorable coupling driving as follows:
Figure BDA0002320035230000095
Figure BDA0002320035230000096
Figure BDA0002320035230000097
wherein the content of the first and second substances,
Figure BDA0002320035230000098
K0i＞0,ρi＞0,ηiand more than 1(i is 1 and 2) is a design parameter.
In order to analyze the stability of the control methods proposed by the equations (12) to (14), still taking the attitude angular rate subsystem as an example, the Lyapunov function is selected as follows:
Figure BDA0002320035230000101
derivation of formula (15) to obtain
Figure BDA0002320035230000102
Substituting formula (10) for a second formula of formula (5) to obtain
Figure BDA0002320035230000103
Order to
Figure BDA0002320035230000104
Is provided with
Figure BDA0002320035230000105
Order to
Figure BDA0002320035230000106
Is provided with
Figure BDA0002320035230000107
Wherein the content of the first and second substances,
Figure BDA0002320035230000108
indicating the disturbance to which the system is subjected D2The upper bound of (a), which exists but is unknown in size. K2Is an adaptive gain, and is known from equation (13):
Figure BDA0002320035230000109
wherein the content of the first and second substances,
Figure BDA00023200352300001010
K02＞0,ρ2＞0,η2> 1 is a design parameter.
As is apparent from formula (18), when e2When not equal to 0, if
Figure BDA00023200352300001011
Then there is
Figure BDA00023200352300001012
And then have
Figure BDA00023200352300001013
e2Will converge to 0, however in practice
Figure BDA00023200352300001014
Is unknown.
As can be seen from the formula (19), when | | | e2||＞ρ2/η2，K2＝KA2，KA2Is monotonically increasing, it is inevitable to make it after a certain time
Figure BDA00023200352300001015
Is satisfied, then
Figure BDA00023200352300001016
This is true.
When | | | e2||≤ρ2/η2When, K2＝KB2A Lyapunov function including adaptive gain is selected as
Figure BDA00023200352300001017
The derivation of the above formula can be obtained,
Figure BDA00023200352300001018
then, from the formula (16), the formula (17), the formula (18) and the formula (19), it is possible to obtain
Figure BDA00023200352300001019
Wherein, beta0＝ρ2/(ρ2-||e2||)2＞0,
Figure BDA00023200352300001020
Is provided with
Figure BDA00023200352300001021
And taking K into accountB2In | | | e2||∈[0,ρ2) Above is monotonically increasing, then when p2＞||e2||＞ρs2Is provided with
Figure BDA00023200352300001022
Namely, it is
Figure BDA00023200352300001023
Thus is provided with
Figure BDA00023200352300001024
Wherein
Figure BDA0002320035230000111
Therefore, when | | e2||＞ρs2When the temperature of the water is higher than the set temperature,
Figure BDA0002320035230000112
is negatively determined, i.e. norm of attitude angular rate tracking error2I must be limited to E2||≤ρs2In this region, i.e. the tracking error of attitude angular rate | | | e2The | l must converge to a predefined area | e2||≤ρ2. In the same way, it can be verified that the tracking error e of the attitude angle system1The | l also converges to a predefined area | e1||≤ρ1In (1). Therefore, the designed control law (12) and the adaptive law (13) thereof can realize attitude stabilization control under the condition that the upper bound of system interference is unknown.
It should be understood that the addition of a point to a variable is a derivative of the variable unless the derivative of the variable has an actual physical meaning.
Examples
The effectiveness of the method provided by the invention is described below by taking an attitude angle instruction of a certain hypersonic aircraft for tracking a dive section under the condition of pneumatic parameter deviation as an example. The Mach number of the flight speed of the aircraft is 6, the flight height is 20Km, and the initial attack angle, the sideslip angle and the roll angle are 0 degree, 4 degrees and 0 degree respectively. The amplitude limit of the steering engine is [ -20 DEG, 20 DEG ]]The limiting rate is [ -300 °/s, 300 °/s]. Controller parameters
Figure BDA0002320035230000113
K0i,ρi,ηi(i-1, 2) are each independently
Figure BDA0002320035230000114
K01＝0.1,ρ1＝0.1,η1＝10,
Figure BDA0002320035230000115
K02＝0.5,ρ2＝0.16,η22. Fig. 2 is a schematic diagram of a simulation of attitude control when the pneumatic parameter takes a nominal value, and a dotted line and a solid line in the diagram respectively represent an attitude angle command and an actual value of an attitude angle. As can be seen from FIG. 2, in the nominal state, the attitude angle can quickly and accurately track the attitude angle command, and the control effect is good. Fig. 3 is the control input in the nominal state, i.e. the rudder deflection angles of the elevators, the rudder and the ailerons. As can be seen from fig. 3, the change of the rudder deflection angle is continuous, and the maximum value does not exceed 20 degrees, so that the constraints on the amplitude limiting and the speed limiting rate of the actuator are met. The schematic diagram of attitude control simulation under the condition that the pneumatic parameters are biased by 30% is shown in FIG. 4, wherein the dotted line represents the attitude angle instruction of the expected three channels, and the solid line represents the attitude angle instruction of the three channels obtained by utilizing the robust control strategy of the inventionThe actual value of the attitude angle, the dotted line, represents the actual value of the three-channel attitude angle obtained by using the sliding mode control method in the prior document. As can be seen from fig. 4, the control method provided by the present invention can still realize accurate tracking of the desired attack angle, slip angle and roll angle under the condition of pneumatic parameter deviation. And as is apparent from fig. 4, compared with the sliding mode control method in the prior art, the method provided by the present invention has a better tracking effect. FIG. 5 is a graph of controller gain K using the control strategy of the present invention under a pneumatic parameter bias condition1,K2Adaptive change of (2). The schematic diagram of the change of the rudder deflection angle obtained by adopting the method provided by the invention under the condition that the pneumatic parameter deviates by 30% is shown in fig. 6, and the obtained change of the rudder deflection angle can be seen in the diagram to meet the restriction on the amplitude limit and the speed limit rate of an actuating mechanism.
In a word, the control method provided by the invention effectively realizes strong robust quick response attitude control of the hypersonic aircraft in a diving section. Meanwhile, the robust control method does not need to linearize the model, does not need to design a controller by channels, does not need to know the interference upper bound information, and has higher reliability and practical value.
The above description is only for the purpose of illustrating the preferred embodiments of the present invention and is not to be construed as limiting the invention, and any modifications, equivalents, improvements and the like that fall within the spirit and principle of the present invention are intended to be included therein.

Claims (1)

1. A hypersonic aircraft nose-down section fast roll control method is characterized by comprising the following steps:
s1: establishing a control-oriented hypersonic aircraft attitude coupling model;
the established control-oriented hypersonic aircraft attitude coupling model is as follows:
Figure FDA0002717658410000011
in the formula, x1＝[γv,β,α]T,γvRepresents roll angle, β represents slip angle, α represents angle of attack; x is the number of2＝[ωx,ωy,ωz]T,ωx,ωy,ωzRespectively representing the rolling rate, the yawing rate and the pitching angle rate of the aircraft rotating around the three axes of the body coordinate system; u ═ 2x,y,z]，x,y,zRespectively representing the roll, yaw and pitch rudder deflection of the aircraft; the interference experienced by the system is denoted as D1＝[D11,D12,D13]TAnd D2＝[D21,D22,D23]T,D11,D12,D13,D21,D22,D23Respectively represents the uncertainty of the six channels caused by unmodeled dynamics, the uncertainty of the aerodynamic coefficient and the external disturbance, and
Figure FDA0002717658410000012
Figure FDA0002717658410000013
is a scalar quantity, the size is unknown; superscript-represents the upper bound; superscript-representation derivation; superscript T represents a transpose matrix; matrix F11,F12,F21,F22And B is respectively represented as:
Figure FDA0002717658410000014
Figure FDA0002717658410000015
Figure FDA0002717658410000016
wherein q is 0.5 ρ V2Representing dynamic pressure, ρ representing atmospheric density, and V representing aircraft speed; srefRepresenting a reference area of the aircraft; l represents a reference length; j. the design is a squarex,Jy,JzRepresenting the rotational inertia of the aircraft relative to three axes of a body coordinate system; θ represents a ballistic dip angle; m represents the mass of the aircraft; g represents the gravitational acceleration; cLAnd CZRespectively representing a lift coefficient and a lateral force coefficient; m isx、myAnd mzRespectively representing a roll moment coefficient, a yaw moment coefficient and a pitch moment coefficient; cL0Represents a zero lift coefficient;
Figure FDA0002717658410000021
respectively, the pairs of lift coefficients a are indicated,zpartial derivatives of (d);
Figure FDA0002717658410000022
respectively, represent the lateral force coefficient pair beta,x,ypartial derivatives of (d);
Figure FDA0002717658410000023
respectively representing the roll torque coefficients with respect to beta,x,ypartial derivatives of (d);
Figure FDA0002717658410000024
respectively, the yaw moment coefficient pair beta is represented,x,ypartial derivatives of (d);
Figure FDA0002717658410000025
respectively, represent the pitch moment coefficient pair a,zpartial derivatives of (d);
s2: establishing a three-channel attitude error dynamic model based on the control-oriented hypersonic aircraft attitude coupling model established in the step S1, and defining a coupling profit-and-loss judgment criterion based on control performance;
the established three-channel attitude error dynamic model comprises the following specific processes:
let the attitude angle tracking command be x1c＝[γvc,βc,αc]T，γvc,βc,αcRespectively representing a roll angle command value, a sideslip angle command value and an attack angle command value; virtual control law x is introduced based on thought of backstepping method2d∈R3×1And defining a tracking error e1＝x1-x1c,e2＝x2-x2dTracking error e1And e2Respectively three-dimensional column vector, for tracking error e1And e2And (5) obtaining a derivative and combining the formula (4) to obtain an error state equation:
Figure FDA0002717658410000026
wherein, F12e2,F22e1∈R3×1Coupling terms between the attitude angle subsystem and the angular rate subsystem are respectively;
the defined coupled pros and cons judgment criteria based on the control performance are as follows:
defining coupled profit-fraud decision factor ζij＝eij(Fi2e3-i)j∈R，eijFor tracking error e1And e21,2, j 1,2,3, when ζij＝eij(Fi2e3-i)jIf < 0, the coupled term is determined (F)i2e3-i)jThe influence on the system is favorable, otherwise the coupling term (F) is determinedi2e3-i)jHarmful to the system;
s3: designing a three-channel attitude coupling controller by using the coupling prosperity judgment criterion defined in the step S2;
the designed three-channel attitude coupling controller comprises the following specific processes:
to ensure tracking error e2When 0, the control law u is designed as follows:
u＝ueq+usw (6)
wherein u iseq,uswRespectively representing an equivalent control law and a discontinuous control law;
in the formula (5)
Figure FDA0002717658410000027
Interference term D2Obtaining equivalent control by temporarily not considering in designing nominal controller
Figure FDA0002717658410000028
Figure FDA0002717658410000029
In conjunction with the coupled prosperity determination factor ζ defined in step S2ijWill equivalently control
Figure FDA0002717658410000031
The improvement is as follows:
Figure FDA0002717658410000032
wherein S is2＝diag{sgn(ζ21)，sgn(ζ22)，sgn(ζ23) }, sgn (·) denotes a sign function;
meanwhile, designing a discontinuous control law uswThe form is as follows:
Figure FDA0002717658410000033
wherein, K2To control the gain;
the control law u is obtained by substituting the formula (8) and the formula (9) into the formula (6)
Figure FDA0002717658410000034
To ensure tracking error e1Design the virtual control law x as 02dIs composed of
Figure FDA0002717658410000035
Wherein, K1To control the gain;
the designed three-channel attitude coupling controller comprises the following components:
Figure FDA0002717658410000036
s4: considering the strong interference characteristics brought by unmodeled dynamics of the aircraft, external disturbance and the like, designing a gain self-adaptive method with anti-interference capability;
the designed gain self-adaptive method with anti-interference capability comprises the following steps:
Figure FDA0002717658410000037
wherein, i is 1,2,
Figure FDA0002717658410000038
and K0iIs a positive constant, KBiThe form of (A) is as follows:
Figure FDA0002717658410000039
where ρ isi,ηiTo design the parameter, ρiGreater than 0 is a predetermined small constant, etaiIs constant and ηi＞1；
S5: and obtaining a high-precision robust control strategy based on favorable coupling driving.
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