CN114281092A - Hypersonic aircraft coordination attitude control method based on sliding mode disturbance observer - Google Patents
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Abstract
The invention relates to a hypersonic aircraft coordination attitude control method based on a sliding mode disturbance observer, which aims at the hypersonic aircraft attitude control problem under unknown nonlinear dynamics, pneumatic parameter uncertainty and disturbance, firstly, establishing a hypersonic aircraft attitude dynamics model containing the pneumatic parameter uncertainty and disturbance, calling the unknown nonlinear dynamics in the model as unknown dynamics, and calling the pneumatic parameter uncertainty and disturbance as total disturbance to obtain an equivalent model; secondly, designing a sliding mode disturbance observer, and respectively estimating unknown dynamic and total disturbance; and finally, designing the hypersonic aircraft coordination attitude controller based on the sliding mode disturbance observer by utilizing the unknown dynamic estimated value and the total disturbance estimated value and combining a backstepping method. The attitude control method realizes the attitude control of the flat flight section and the pressing section of the hypersonic aircraft, has the characteristics of strong robustness and rapidity, and is suitable for various hypersonic flight systems needing unpowered reentry.
Description
Technical Field
The invention relates to a hypersonic aircraft coordination attitude control method based on a sliding mode disturbance observer, and solves the problem of hypersonic aircraft attitude control including unknown nonlinear dynamics, pneumatic parameter uncertainty and disturbance.
Background
Hypersonic aircraft refers to aircraft having a flight speed greater than mach five and typically operating in the vicinity of 20km to 100km from the ground. Due to the characteristics of high flying speed and special working range, the hypersonic aircraft has various characteristics different from the traditional aerospace vehicle: strong uncertainty, interference, strong coupling, strong non-linearity, fast time-varying, multiple constraints. Firstly, the problem of rapidity is solved, the speed of the hypersonic aircraft is extremely high and can even reach 7000m/s in a reentry stage, and the speed variation range is very large, so that the control algorithm is required to have the performances of high precision, rapidity and the like, and the target instruction can be tracked in a short time; secondly, the problem of interference in the flying process is solved, the extremely high flying speed can enable the body of the hypersonic aircraft to be subjected to the real gas effect and the viscous effect, and can also cause serious pneumatic heating, so that the structure and the inherent mode of the hypersonic aircraft are changed, meanwhile, the atmospheric environment is complex and changeable, and the hypersonic aircraft can also be inevitably influenced by airflow disturbance in the flying environment to influence the control performance; then, the problem of strong coupling exists, strong coupling exists among channels of the hypersonic aerocraft, and the hypersonic aerocraft structure and the aeroelasticity, and the generated coupling torque is sometimes even larger than the control torque of the channel; finally, hypersonic aircraft also have the problem of strong uncertainty. Because a flight airspace covers the atmosphere to near space and exceeds the effective range of classical aerodynamics, the current flight test, wind tunnel simulation and computational fluid mechanics are not enough to carry out accurate analysis on the hypersonic aerocraft, and on the other hand, the parameter uncertainty of the hypersonic aerocraft is caused by the interaction between the fuselage/propulsion/structural mechanics of the hypersonic aerocraft. The factors bring many problems and challenges to the design of the attitude control system of the hypersonic aircraft, and the high-precision coordination attitude control method of the hypersonic aircraft has wide application prospect.
At present, research on attitude control of hypersonic flight vehicles mainly focuses on adaptive control and robust control. The patent number CN201910333259.0 proposes a hypersonic aircraft coordination control method based on dynamic coupling analysis, which applies a dynamic coupling analysis method to a longitudinal system model of a hypersonic aircraft to obtain a dynamic coupling relationship matrix between system variables and control input variables, and designs a longitudinal system coordination controller of the hypersonic aircraft. The patent No. CN201810950223.2 provides a hypersonic aircraft neural network learning control method based on lumped composite estimation, which comprises the steps of decoupling a longitudinal channel model of a hypersonic aircraft into a speed subsystem and an altitude subsystem, estimating by adopting a neural network aiming at the uncertainty of dynamics existing in the system, estimating by adopting a nonlinear observer aiming at coupling brought by an elastic mode, giving an altitude and a speed controller based on information of two estimators to realize tracking of altitude and speed, and the patent No. CN201710789243.1 provides a hypersonic aircraft neural network composite learning control method based on robust design, transforming a strict feedback form of an attitude subsystem to obtain an output feedback form, estimating a newly defined variable by using a high-gain observer, approximating the lumped uncertainty of the system by using the neural network, and having better robustness, however, the three methods are based on the simplified longitudinal model, neglect part of nonlinear dynamics in the hypersonic aircraft model, have certain difference with the actual application of the hypersonic aircraft, and are difficult to be directly applied in engineering. The document terminal sliding mode attitude control of the hypersonic aircraft designs a PID control law for a slow loop of a hypersonic aircraft attitude control system, designs a terminal sliding mode control law for a fast loop, and reduces the sensitivity of a controller to system parameters.
In conclusion, under the conditions of unknown nonlinear dynamics, uncertainty of pneumatic parameters and interference, the conventional method is lack of a high-precision coordination attitude control method for the hypersonic aircraft, and a coordinated attitude control method for the hypersonic aircraft based on a sliding mode interference observer is urgently needed.
Disclosure of Invention
The technical problem solved by the invention is as follows: aiming at the control problems of a horizontal flight section and a downward pressing section of a hypersonic aircraft containing unknown nonlinear dynamics, uncertain pneumatic parameters and interference, the defect of the prior art is overcome, the coordinated attitude control method of the hypersonic aircraft based on the sliding mode interference observer is provided, the estimation and compensation of the unknown dynamics and the total interference are realized, a set of control strategies are adopted to enable the attitude of the hypersonic aircraft to track a given expected attitude instruction, certain controller dynamic performance is ensured, and the autonomy, the accuracy and the anti-interference capability of the hypersonic aircraft in the control process are improved.
The technical solution of the invention is as follows: a hypersonic aircraft coordination attitude control method based on a sliding mode disturbance observer is characterized by firstly establishing a hypersonic aircraft attitude dynamics model containing pneumatic parameter uncertainty and disturbance according to the hypersonic aircraft attitude dynamics characteristics, calling unknown nonlinear dynamics in which attack angle, sideslip angle and speed roll angle channels are coupled with mass center motion states in the attitude dynamics model as unknown dynamics, calling the pneumatic parameter uncertainty and disturbance of roll, yaw and pitch channels as total disturbance, and expressing related state variables in the attitude dynamics model in the form of a differential equation set to obtain an equivalent model; secondly, designing a sliding mode disturbance observer, and respectively estimating the unknown dynamic and the total disturbance to obtain estimated values of the unknown dynamic and the total disturbance, wherein the estimation errors of the unknown dynamic and the total disturbance can be converged to zero within a limited time, and the specific numerical value of the limited time is determined by the initial state of the aircraft and the design parameters of the sliding mode disturbance observer; finally, designing a hypersonic aircraft coordination attitude controller based on the sliding mode disturbance observer by utilizing the unknown dynamic estimated value and the total disturbance estimated value and combining a backstepping method, and finishing the hypersonic aircraft coordination attitude control method based on the sliding mode disturbance observer, wherein the method comprises the following concrete implementation steps:
firstly, establishing a hypersonic speed aircraft attitude dynamics model containing pneumatic parameter uncertainty and interference according to the hypersonic speed aircraft attitude dynamics characteristics; unknown nonlinear dynamics of the attitude dynamics model, in which the attack angle, the sideslip angle and the speed roll angle channels are coupled with the mass center motion state, are called unknown dynamics; the three channels represent a roll channel, a yaw channel and a pitch channel, pneumatic parameter uncertainty and interference of the three channels in the attitude dynamics model are called total interference, relevant state variables in the attitude dynamics model are expressed in a form of a differential equation set, and an equivalent model of the attitude dynamics model of the hypersonic aircraft is obtained, and the method specifically comprises the following steps:
(1) establishing a hypersonic aircraft attitude dynamics model containing pneumatic parameter uncertainty and interference according to the hypersonic aircraft attitude dynamics characteristics;
wherein alpha is the attack angle of the hypersonic aerocraft, beta is the sideslip angle of the hypersonic aerocraft, and gamma isvFor a hypersonic aircraft speed pitch angle,
are respectively alpha, beta and gammavFirst derivative of, omegam、ωn、ωlRespectively pitch, yaw and roll rotation angular speeds,
are respectively omegam、ωn、ωlFirst derivative of (I)m、In、IlRespectively pitch, yaw and roll axis moments of inertia, r1、r2、r3Respectively representing unknown nonlinear dynamics of the coupling of the attack angle, sideslip angle, speed roll angle channels and the mass center motion state in the attitude dynamics model, d1、d2、d3Respectively representing the uncertainty and interference of pneumatic parameters of three channels in the attitude dynamics model, wherein V is the speed of the hypersonic aerocraft, theta is the track inclination angle, m is the mass of the hypersonic aerocraft, g is the gravity acceleration, T is the thrust of an engine, zero is obtained in the unpowered reentry process, D, L, Z, Mm、Mn、MlRespectively is resistance, lift, lateral force, pitching moment, yawing moment, roll moment, and the expression is as follows:
wherein rho is the atmospheric density, S is the hypersonic aircraft reference area, cD、cL、cZRespectively the drag, lift and lateral force coefficients, b is the wing span length, l is the wing average aerodynamic chord length, cm、cn、clRoll, yaw and pitch moment coefficients are respectively provided, and the model of the aerodynamic force and the aerodynamic moment coefficient is as follows:
where Ma is Mach number, deltam、δn、δlRespectively a pitching rudder deflection angle, a yawing rudder deflection angle and a rolling rudder deflection angle,
all are polynomial model coefficients, and the controlled quantity is hypersonic aircraft pitching rudder deflection angle deltamYaw rudder deflection angle deltanYaw angle delta of rolling rudderl；
(2) Expressing the relevant state variables in the attitude dynamics model in a form of a differential equation system to obtain an equivalent model:
wherein x is1＝[α,β,γv]TRepresenting the attitude angle, x, of a hypersonic aircraft2＝[ωm,ωn,ωl]TRepresenting the angular velocity of the hypersonic aircraft, u ═ deltam,δn,δl]TRepresenting a control input, D1＝[r1 r2 r3]TRepresenting unknown dynamics, D2＝[d1 d2d3]TRepresenting the total interference and G the control gain matrix.
Secondly, designing a sliding mode disturbance observer based on an equivalent model, and respectively estimating unknown dynamic and total disturbance to obtain estimated values of the unknown dynamic and the total disturbance, wherein the estimated errors of the unknown dynamic and the total disturbance can be converged to zero within a limited time, and the specific numerical value of the limited time is determined by the initial state of the aircraft and the design parameters of the sliding mode disturbance observer;
based on the equivalent model, the sliding mode disturbance observer is designed as follows:
wherein, [ f ]1,f2,f3]T＝F(x1)x2，[f4,f5,f6]T＝H(x2)+M+Gu，
Representing an estimate of the hypersonic aerial vehicle angle of attack alpha,
representing an estimate of the hypersonic aerial vehicle sideslip angle beta,
representing the velocity tilt angle gamma of a hypersonic aircraftvIs determined by the estimated value of (c),
representing the pitching rotation angular velocity omega of the hypersonic aerocraftmIs determined by the estimated value of (c),
representing the yaw rotation angular velocity omega of the hypersonic aerocraftnIs determined by the estimated value of (c),
representing the rolling rotation angular velocity omega of the hypersonic aircraftlIs determined by the estimated value of (c),
representing unknown dynamics r1Is determined by the estimated value of (c),
representing unknown dynamics r2Is determined by the estimated value of (c),
representing unknown dynamics r3Is determined by the estimated value of (c),
representing the total interference d1Is determined by the estimated value of (c),
indicates total stemDisturbance d2Is determined by the estimated value of (c),
representing the total interference d3Is determined by the estimated value of (c),
both represent intermediate variables in the design process,
L1、L2、L3、L4、L5、L6respectively representing design parameters of the sliding mode disturbance observer, and defining a symbolic function as follows: for real number y
And for column vectors
sign(Y)＝[sign(y1)sign(y2)sign(y3)]T
In the third step, an unknown dynamic estimated value and a total interference estimated value are utilized, a backstepping method is combined, a three-channel coordinated attitude controller of the hypersonic aircraft is designed, and the coordinated attitude control method of the hypersonic aircraft based on the sliding mode interference observer is completed:
(1) designing a stabilization function to enable the attitude of the hypersonic aircraft to track an expected instruction x1d＝[αd,βd,γvd]T，αdRepresenting desired angle of attack, betadIndicating a desired sideslip angle, γvdRepresenting the roll angle of the expected speed and defining the attitude tracking errorz1＝x1d-x1Designing a stabilizing function:
wherein E represents a matrix F (x)1) The inverse matrix of (c):
(2) designing control input to make the angular rate tracking stabilization function z of the hypersonic aerocraftdDefining the angular rate tracking error z2＝zd-x2Designing a coordinated attitude controller of the hypersonic aircraft:
wherein the content of the first and second substances,
wherein the matrix Φ represents:
wherein the content of the first and second substances,
the representation matrix F (x)1) The derivative of the inverse matrix E, the expression:
wherein the content of the first and second substances,
representing the total interference D2Is determined by the estimated value of (c),
a derivative of the desired pose command is represented,
second derivative, k, representing the desired attitude command1、κ2、κ3、κ4、κ5、c1、c2And representing the design parameters of the hypersonic aircraft coordinated attitude controller.
Compared with the prior art, the invention has the advantages that: aiming at the defects of unknown nonlinear dynamics, pneumatic parameter uncertainty and interference, no utilization of favorable coupling among channels, easy saturation of control input and lack of high-precision control capability of the existing method, the invention designs a sliding mode interference observer to estimate total interference composed of pneumatic parameter uncertainty and interference in unknown dynamics and angular rate rings in an attitude ring respectively, designs a control law based on a back-stepping method to carry out coordination control on three-axis attitude and compensate for the unknown dynamics and the total interference, realizes a three-axis coordination attitude controller, the output of the controller can be obviously reduced by adjusting design parameters, the occurrence of the saturation condition of the control surface of the actuating mechanism is reduced, meanwhile, the attitude control system can realize high-precision robust control of the attitude of the hypersonic aircraft under the condition of the uncertainty and the interference of pneumatic parameters, and is suitable for attitude control systems of a flat flight section and a pressing section of the hypersonic aircraft.
Drawings
FIG. 1 is a flow chart of the design of the method of the present invention;
FIG. 2 is a control block diagram of the method of the present invention;
FIG. 3 is a diagram of a simulation result of an unknown dynamic estimation effect of the method of the present invention;
FIG. 4 is a diagram of a simulation result of the total interference estimation effect of the method of the present invention;
FIG. 5 is a diagram of a simulation result of the attitude tracking effect of the method of the present invention.
Detailed Description
The present invention will be described in detail below with reference to the accompanying drawings and examples.
As shown in FIG. 1, the invention relates to a hypersonic aircraft coordination attitude control method based on a sliding mode disturbance observer. Firstly, establishing a hypersonic speed aircraft attitude dynamics model containing pneumatic parameter uncertainty and interference according to the hypersonic speed aircraft attitude dynamics characteristics; unknown nonlinear dynamics of the attitude dynamics model, in which the attack angle, the sideslip angle and the speed roll angle channels are coupled with the mass center motion state, are called unknown dynamics; the three channels represent a rolling channel, a yawing channel and a pitching channel, and the uncertainty and the interference of the pneumatic parameters of the three channels in the attitude dynamics model are called as total interference. Expressing the relevant state variables in the attitude dynamics model in a differential equation set form to obtain an equivalent model; and secondly, designing a sliding mode disturbance observer based on an equivalent model, and estimating the unknown dynamic and the total disturbance respectively to obtain estimated values of the unknown dynamic and the total disturbance, wherein the estimated errors of the unknown dynamic and the total disturbance can be converged to zero within a limited time, and the specific numerical value of the limited time is determined by the initial state of the aircraft and the design parameters of the sliding mode disturbance observer. Thirdly, designing a three-channel coordinated attitude controller of the hypersonic aircraft by utilizing the dynamic estimated value and the total interference estimated value and combining a backstepping method, and finishing the coordinated attitude control method of the hypersonic aircraft based on the sliding mode interference observer; the control law is designed based on a backstepping method to carry out coordination control on the three-axis attitude and compensate unknown dynamic and total interference, the three-axis coordination attitude controller is realized, the output of the controller can be obviously reduced by adjusting design parameters, the occurrence of the saturation condition of the control surface of an actuating mechanism is reduced, meanwhile, the high-precision robust control on the attitude of the hypersonic aircraft can be realized under the condition of the uncertainty and the interference of pneumatic parameters, and the control law is suitable for the attitude control systems of the level flight section and the pressing section of the hypersonic aircraft.
The specific implementation steps are as follows:
firstly, establishing a hypersonic speed aircraft attitude dynamics model containing pneumatic parameter uncertainty and interference according to the hypersonic speed aircraft attitude dynamics characteristics; unknown nonlinear dynamics of the attitude dynamics model, in which the attack angle, the sideslip angle and the speed roll angle channels are coupled with the mass center motion state, are called unknown dynamics; the three channels represent a roll channel, a yaw channel and a pitch channel, pneumatic parameter uncertainty and interference of the three channels in the attitude dynamics model are called total interference, relevant state variables in the attitude dynamics model are expressed in a form of a differential equation set, and an equivalent model of the attitude dynamics model of the hypersonic aircraft is obtained, and the method specifically comprises the following steps:
(1) establishing a hypersonic aircraft attitude dynamics model containing pneumatic parameter uncertainty and interference according to the hypersonic aircraft attitude dynamics characteristics;
wherein alpha is the attack angle of the hypersonic aerocraft, the initial value is 0 radian, beta is the sideslip angle of the hypersonic aerocraft, the initial value is 0 radian, and gamma isvThe speed inclination angle of the hypersonic aerocraft is 0 radian at the initial value,
are respectively alpha, beta and gammavFirst derivative of, omegam、ωn、ωlThe rotation angular speeds of pitching, yawing and rolling are respectively, the initial values are all 0 radian/second,
are respectively omegam、ωn、ωlFirst derivative of (I)mIs the rotational inertia of a pitch shaft and has a value of 1800 kg.m2，InThe value of the rotational inertia of the yaw axis is 1800 kg.m2，IlIs the rotational inertia of the rolling shaft and has a value of 150 kg.m2，r1、r2、r3Respectively representing unknown nonlinear dynamics of the coupling of the attack angle, sideslip angle, speed roll angle channels and the mass center motion state in the attitude dynamics model, d1、d2、d3Respectively representing the uncertainty and interference of aerodynamic parameters of three channels in the attitude dynamics model, wherein V is the speed of the hypersonic aerocraft, the initial value is 2400 m/s, theta is the track inclination angle, the initial value is 0 radian, m is the mass of the hypersonic aerocraft, the value is 400kg, g is the gravity acceleration, the value is 9.8 m/second of square, T is the thrust of an engine, and is zero in the unpowered reentry process D, L, Z, Mm、Mn、MlRespectively is resistance, lift, lateral force, pitching moment, yawing moment, roll moment, and the expression is as follows:
wherein rho is the atmospheric density, S is the hypersonic aircraft reference area, the value is 1 square meter, cD、cL、cZRespectively the drag, lift and lateral force coefficients, b is the wing span length with a value of 4 m, l is the wing average aerodynamic chord length with a value of 1 m, cm、cn、clRoll, yaw and pitch moment coefficients are respectively provided, and the model of the aerodynamic force and the aerodynamic moment coefficient is as follows:
where Ma is Mach number, deltam、δn、δlRespectively a pitching rudder deflection angle, a yawing rudder deflection angle and a rolling rudder deflection angle,
all are polynomial model coefficients, and the controlled quantity is hypersonic aircraft pitching rudder deflection angle deltamYaw rudder deflection angle deltanYaw angle delta of rolling rudderl。
(2) Expressing the relevant state variables in the attitude dynamics model in a form of a differential equation system to obtain an equivalent model:
wherein x is1＝[α,β,γv]TRepresenting the attitude angle, x, of a hypersonic aircraft2＝[ωm,ωn,ωl]TRepresenting the angular velocity of the hypersonic aircraft, u ═ deltam,δn,δl]TRepresenting a control input, D1＝[r1 r2 r3]TRepresenting unknown dynamics, D2＝[d1 d2d3]TRepresenting the total interference and G the control gain matrix.
Secondly, designing a sliding mode disturbance observer based on an equivalent model, and respectively estimating unknown dynamic and total disturbance to obtain estimated values of the unknown dynamic and the total disturbance, wherein the estimated errors of the unknown dynamic and the total disturbance can be converged to zero within a limited time, and the specific numerical value of the limited time is determined by the initial state of the aircraft and the design parameters of the sliding mode disturbance observer;
the sliding mode disturbance observer is designed as follows:
wherein, [ f ]1,f2,f3]T＝F(x1)x2，[f4,f5,f6]T＝H(x2)+M+Gu，
Representing an estimate of the hypersonic aerial vehicle angle of attack alpha,
representing an estimate of the hypersonic aerial vehicle sideslip angle beta,
representing the velocity tilt angle gamma of a hypersonic aircraftvIs determined by the estimated value of (c),
representing the pitching rotation angular velocity omega of the hypersonic aerocraftmIs determined by the estimated value of (c),
representing the yaw rotation angular velocity omega of the hypersonic aerocraftnIs determined by the estimated value of (c),
representing the rolling rotation angular velocity omega of the hypersonic aircraftlIs determined by the estimated value of (c),
representing unknown dynamics r1Is determined by the estimated value of (c),
representing unknown dynamics r2Is determined by the estimated value of (c),
representing unknown dynamics r3Is determined by the estimated value of (c),
representing the total interference d1Is determined by the estimated value of (c),
representing the total interference d2Is determined by the estimated value of (c),
representing the total interference d3Is determined by the estimated value of (c),
both represent intermediate variables in the design process,
L1、L2、L3、L4、L5、L6respectively represent the design parameters of the sliding mode disturbance observer, and can be taken as follows:
wherein the sign function is defined as: for real number y
And for column vectors
sign(Y)＝[sign(y1) sign(y2) sign(y3)]T
Thirdly, designing a three-channel coordinated attitude controller of the hypersonic aircraft by utilizing an unknown dynamic estimated value and a total interference estimated value and combining a backstepping method, and finishing the coordinated attitude control of the hypersonic aircraft based on the sliding mode interference observer:
(1) designing a stabilization function to enable the attitude of the hypersonic aircraft to track an expected instruction x1d＝[αd,βd,γvd]T，αdRepresents an expected angle of attack, and is a step signal with an initial value of 0 radian and a final value of 0.17 radian and a step response time of 0.5 seconds, betadRepresenting a desired sideslip angle, and having a value of zero, γvdRepresenting the expected speed roll angle, is a square wave signal with the amplitude of 0.087 radian and the period of 20 seconds, and defines an attitude tracking error z1＝x1d-x1Designing a stabilizing function:
wherein E represents a matrix F (x)1) The inverse matrix of (c):
(2) designing control input to make the angular rate tracking stabilization function z of the hypersonic aerocraftdDefining the angular rate tracking error z2＝zd-x2Designing a coordinated attitude controller of the hypersonic aircraft:
wherein the content of the first and second substances,
wherein the matrix Φ represents:
wherein the content of the first and second substances,
the representation matrix F (x)1) The derivative of the inverse matrix E, the expression:
wherein the content of the first and second substances,
representing the total interference D2Is determined by the estimated value of (c),
a derivative of the desired pose command is represented,
second derivative, k, representing the desired attitude command1、κ2、κ3、κ4、κ5、c1、c2The design parameters for representing the coordinated attitude controller of the hypersonic aircraft can be as follows:
c2＝1.3,κ2＝0.3,κ3＝0.3,κ4＝0.3,κ5＝0.3
the hypersonic speed aircraft coordination attitude control method based on the sliding mode disturbance observer and the parameters are adopted, and simulation software is used for testing. FIG. 3 shows: and (3) estimating the effect of the sliding mode disturbance observer on the unknown dynamic state. The solid line in the upper graph of FIG. 3 represents the unknown dynamics r1True value of (d), dashed line represents unknown dynamics r1The estimation error converges to zero over 6 seconds; the solid line in the graph in FIG. 3 represents the unknown dynamics r2True value of (d), dashed line represents unknown dynamics r2The estimation error converges to zero over 4 seconds; the solid line in the lower graph of FIG. 3 represents the unknown dynamics r3True value of (d), dashed line represents unknown dynamics r3The estimation error converges to zero over 4 seconds; fig. 3 illustrates that the estimation error of the sliding mode disturbance observer designed by the invention to the unknown dynamic can be converged to zero rapidly.
FIG. 4 shows: and (3) estimating the effect of the sliding mode disturbance observer on the total disturbance. The solid line in the upper graph of fig. 4 represents the total interference d1Is true ofThe value, dashed line, represents the total interference d1The estimation error converges to zero after 8 seconds; the solid line in the diagram in fig. 4 represents the total interference d2True value of (d), dashed line indicates total interference d2The estimation error converges to zero over 4 seconds; the solid line in the lower graph of fig. 4 represents the total disturbance d3True value of (d), dashed line indicates total interference d3The estimation error converges to zero over 4 seconds; fig. 4 illustrates that the estimation error of the sliding mode disturbance observer designed by the invention to the total disturbance can be converged to zero rapidly.
Fig. 5 shows: under the control action of the hypersonic aircraft coordination attitude controller based on the sliding mode disturbance observer, the hypersonic aircraft performs the control on the expected instruction x1d＝[αd,βd,γvd]TThe tracking effect of (1). The solid line in the upper graph of fig. 5 represents the desired angle of attack αdThe dashed line represents the actual angle of attack α, and the tracking error converges to zero over 8 seconds; the solid line in the graph in fig. 5 represents the desired sideslip angle βdThe dashed line represents the actual sideslip angle β, with the tracking error converging to zero over 4 seconds; the solid line in the lower graph of fig. 5 represents the desired roll angle γvdThe dotted line represents the actual speed roll angle γvThe tracking error converges to zero over 4 seconds; fig. 5 illustrates that the coordinated attitude controller of the hypersonic flight vehicle designed by the invention has good control effect on the attitude of the hypersonic flight vehicle, and the tracking error of the attitude instruction can be rapidly converged to zero.
The attitude control method for the hypersonic aircraft can be adopted to control the attitude of the hypersonic aircraft, a set of attitude control strategies can be used in a horizontal flight section and a downward pressing section of the hypersonic aircraft, better dynamic performance of a controller is kept, and no steady-state error exists in unknown dynamic estimation, total interference estimation and attitude tracking. Meanwhile, compared with a single-channel controller with interference-free estimation and compensation, the control effect is relatively smaller in control input, and can bear larger pneumatic parameter uncertainty and system total interference, so that the requirements of high precision and strong robustness are met.
Those skilled in the art will appreciate that the invention may be practiced without these specific details.
Claims (4)
1. A hypersonic aircraft coordination attitude control method based on a sliding mode disturbance observer is characterized by comprising the following steps:
firstly, establishing a hypersonic speed aircraft attitude dynamics model containing pneumatic parameter uncertainty and interference according to the hypersonic speed aircraft attitude dynamics characteristics; unknown nonlinear dynamics of the attitude dynamics model, in which the attack angle, the sideslip angle and the speed roll angle channels are coupled with the mass center motion state, are called unknown dynamics; three channels represent a rolling channel, a yawing channel and a pitching channel, pneumatic parameter uncertainty and interference of the three channels in the attitude dynamics model are called total interference, and related state variables in the attitude dynamics model are expressed in the form of a differential equation set to obtain an equivalent model;
secondly, designing a sliding mode disturbance observer based on an equivalent model, and respectively estimating unknown dynamic and total disturbance to obtain estimated values of the unknown dynamic and the total disturbance, wherein the estimated errors of the unknown dynamic and the total disturbance can be converged to zero within a limited time, and the specific numerical value of the limited time is determined by the initial state of the aircraft and the design parameters of the sliding mode disturbance observer;
and thirdly, designing a hypersonic aircraft coordination attitude controller based on the sliding mode disturbance observer by utilizing the unknown dynamic estimated value and the total disturbance estimated value and combining a backstepping method, and finishing the hypersonic aircraft coordination attitude control method based on the sliding mode disturbance observer technology.
2. The method for controlling the coordinated attitude of the hypersonic aircraft based on the sliding-mode disturbance observer according to claim 1, is characterized in that: in the first step, a hypersonic speed aircraft attitude dynamics model containing pneumatic parameter uncertainty and interference is established according to the hypersonic speed aircraft attitude dynamics characteristics; unknown nonlinear dynamics of the attitude dynamics model, in which the attack angle, the sideslip angle and the speed roll angle channels are coupled with the mass center motion state, are called unknown dynamics; the three channels represent a roll channel, a yaw channel and a pitch channel, pneumatic parameter uncertainty and interference of the three channels in the attitude dynamics model are called total interference, related state variables in the attitude dynamics model are expressed in a form of a differential equation set to obtain an equivalent model, and the method specifically comprises the following steps:
(1) establishing a hypersonic aircraft attitude dynamics model containing pneumatic parameter uncertainty and interference according to the hypersonic aircraft attitude dynamics characteristics;
wherein alpha is the attack angle of the hypersonic aerocraft, beta is the sideslip angle of the hypersonic aerocraft, and gamma isvFor a hypersonic aircraft speed pitch angle,
are respectively alpha, beta and gammavFirst derivative of, omegam、ωn、ωlRespectively pitch, yaw and roll rotation angular speeds,
are respectively omegam、ωn、ωlFirst derivative of (I)m、In、IlRespectively pitch, yaw and roll axis moments of inertia, r1、r2、r3Respectively representing unknown nonlinear dynamics of the coupling of the attack angle, sideslip angle, speed roll angle channels and the mass center motion state in the attitude dynamics model, d1、d2、d3Pneumatic parameter uncertainty respectively representing three channels in attitude dynamics modelSex and interference, V is hypersonic aircraft speed, theta is track inclination angle, m is hypersonic aircraft mass, g is gravitational acceleration, T is engine thrust, zero in unpowered reentry process, D, L, Z, Mm、Mn、MlRespectively is resistance, lift, lateral force, pitching moment, yawing moment, roll moment, and the expression is as follows:
wherein rho is the atmospheric density, S is the hypersonic aircraft reference area, cD、cL、cZRespectively the drag, lift and lateral force coefficients, b is the wing span length, l is the wing average aerodynamic chord length, cm、cn、clRoll, yaw and pitch moment coefficients are respectively provided, and the model of the aerodynamic force and the aerodynamic moment coefficient is as follows:
where Ma is Mach number, deltam、δn、δlRespectively a pitching rudder deflection angle, a yawing rudder deflection angle and a rolling rudder deflection angle,
all are polynomial model coefficients, and the controlled quantity is hypersonic aircraft pitching rudder deflection angle deltamYaw rudder deflection angle deltanYaw angle delta of rolling rudderl；
(2) Expressing the relevant state variables in the attitude dynamics model in a form of a differential equation system to obtain an equivalent model:
wherein x is1＝[α,β,γv]TRepresenting the attitude angle, x, of a hypersonic aircraft2＝[ωm,ωn,ωl]TRepresenting the angular velocity of the hypersonic aircraft, u ═ deltam,δn,δl]TRepresenting a control input, D1＝[r1 r2 r3]TRepresenting unknown dynamics, D2＝[d1 d2 d3]TRepresenting the total interference and G the control gain matrix.
3. The method for controlling the coordinated attitude of the hypersonic aircraft based on the sliding-mode disturbance observer according to claim 1, is characterized in that: in the second step, based on the equivalent model, the sliding mode disturbance observer is designed as follows:
wherein, [ f ]1,f2,f3]T＝F(x1)x2，[f4,f5,f6]T＝H(x2)+M+Gu，
Representing an estimate of the hypersonic aerial vehicle angle of attack alpha,
representing an estimate of the hypersonic aerial vehicle sideslip angle beta,
representing the velocity tilt angle gamma of a hypersonic aircraftvIs determined by the estimated value of (c),
representing the pitching rotation angular velocity omega of the hypersonic aerocraftmIs determined by the estimated value of (c),
representing the yaw rotation angular velocity omega of the hypersonic aerocraftnIs determined by the estimated value of (c),
representing the rolling rotation angular velocity omega of the hypersonic aircraftlIs determined by the estimated value of (c),
representing unknown dynamics r1Is determined by the estimated value of (c),
representing unknown dynamics r2Is determined by the estimated value of (c),
representing unknown dynamics r3Is determined by the estimated value of (c),
representing the total interference d1Is determined by the estimated value of (c),
representing the total interference d2Is determined by the estimated value of (c),
representing the total interference d3Is determined by the estimated value of (c),
both represent intermediate variables in the design process,
L1、L2、L3、L4、L5、L6respectively representing design parameters of the sliding mode disturbance observer, and defining a symbolic function as follows: for real number y
And for a column vector Y ═ Y1,y2,y3]T，
sign(Y)＝[sign(y1) sign(y2) sign(y3)]T
4. The method for controlling the coordinated attitude of the hypersonic aircraft based on the sliding-mode disturbance observer according to claim 1, is characterized in that: the third step is specifically realized as follows:
(1) designing a stabilization function to enable the attitude of the hypersonic aircraft to track an expected instruction x1d＝[αd,βd,γvd]T，αdRepresenting desired angle of attack, betadIndicating a desired sideslip angle, γvdRepresenting the desired roll angle of the velocity, defining the attitude tracking error z1＝x1d-x1Designing a stabilizing function:
wherein E represents a matrix F (x)1) The inverse matrix of (c):
(2) designing control input to make the angular rate tracking stabilization function z of the hypersonic aerocraftdDefining the angular rate tracking error z2＝zd-x2Designing a coordinated attitude controller of the hypersonic aircraft:
wherein the content of the first and second substances,
wherein the matrix Φ represents:
wherein the content of the first and second substances,
the representation matrix F (x)1) The derivative of the inverse matrix E, the expression:
wherein the content of the first and second substances,
representing the total interference D2Is determined by the estimated value of (c),
a derivative of the desired pose command is represented,
second derivative, k, representing the desired attitude command1、κ2、κ3、κ4、κ5、c1、c2And representing the design parameters of the hypersonic aircraft coordinated attitude controller.
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