CN115686036A - Variable-profile aircraft multi-dimensional composite control method based on preset performance - Google Patents
Variable-profile aircraft multi-dimensional composite control method based on preset performance Download PDFInfo
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Abstract
The invention discloses a variable-profile aircraft multi-dimensional composite control method based on preset performance, which comprises the following steps of: step (1), defining a common coordinate system and a motion variable of an aircraft; step (2), establishing a multi-rigid-body six-degree-of-freedom model of the variable-profile aircraft based on the Newton-Euler method; step (3), designing performance constraints of the aircraft attitude response process according to the aircraft flight environment and the flight mission, and completing constraint simplification; and (4) combining the backstepping sliding mode control with the extended state observer on the basis of the three steps to complete the design of the control law of the aircraft. The method provided by the invention can establish direct connection between the control effect and the control law, and simplify the design process of the controller, thereby completing the stable and rapid tracking of the aircraft attitude to the command signal and realizing the flight attitude control.
Description
Technical Field
The invention relates to the technical field of aircraft control, in particular to a variable-profile aircraft multi-dimensional composite control method based on preset performance.
Background
With the increasing diversity and complexity of flying scenes and flying situations, the performance requirements on the aircraft are more and more stringent, and the aircraft with the deformation mechanism is applied. A variable profile aircraft is an aircraft that can be varied in profile in response to changes in flight environment and mission. The variable-profile technology is applied to the aircraft, so that the environmental adaptability and the task adaptability of the aircraft are greatly enhanced, the application value is important, and the method is an important direction for the development of the aircraft in the future.
The aircraft can stably fly, and the accurate task completion depends on the performance of the controller, however, the problem in the design process of the controller of the variable-profile aircraft is more complicated due to the existence of the deformation mechanism. In addition, in the existing control scheme for the aircraft, direct connection between each control parameter and the control effect is lacked, the control parameters need to be continuously debugged to achieve ideal control performance, and the design difficulty of the controller is greatly increased.
(1) The aircraft model has a complex structure
The aircraft is a time-varying, coupling and nonlinear complex model, and the variable-profile aircraft introduces a deformation mechanism on the basis of the time-varying, coupling and nonlinear complex model, so that the dynamic modeling of the variable-profile aircraft as a single rigid body cannot be realized like the conventional aircraft. At the same time, the contouring process not only causes a change in the position of the aircraft's center of mass, but also produces additional forces and moments, which further exacerbate the uncertainty and coupling.
(2) Task working condition and environment are various and complicated
The difference of aerodynamic characteristics of the aircraft under different aerodynamic configurations is obvious, an unsteady effect and other complex aerodynamic characteristics can be generated in the deformation transition process, the uncertainty of the deformation process is further enhanced, and the design difficulty of the controller is increased.
(3) The controller is in theoretical connection with the control effect
A variable profile aircraft is a system with unknown non-linearities, and one important consideration for designing the controller is the transient and steady state performance of the system. In the design process of the traditional controller, each key performance index of the attitude response of the variable-profile aircraft lacks a direct mapping relation with the controller design, and each parameter cannot be reasonably selected in the design process of the controller, so that the response process meets a comprehensive performance evaluation system. Therefore, a control method is needed to establish a direct connection between a performance evaluation system and controller design, to clarify the corresponding relationship between each parameter and control effect in the controller design process, and to finally realize the multi-dimensional attitude composite control of the variable-profile aircraft.
In summary, the variable profile aircraft meets more complex and diversified flight tasks, and meanwhile, the characteristics of strong coupling, fast time variation, strong nonlinearity and the like also provide great challenges for the design of the controller. The traditional controller cannot meet the complex and diverse task requirements of the deformable aircraft and cannot ensure that the attitude response process conforms to an expected comprehensive performance evaluation system. Therefore, how to design a controller ensures that a response curve is kept within a performance envelope on the basis of ensuring accurate, rapid and stable tracking command signals of the aircraft, so that a comprehensive performance evaluation system is met, and the technical problem which needs to be solved urgently in attitude control of the aircraft at present becomes.
Disclosure of Invention
The invention aims to provide a variable-profile aircraft multi-dimensional composite control method based on preset performance so as to solve the problems.
The technical scheme adopted by the invention for solving the technical problem is as follows:
a variable-profile aircraft multi-dimensional composite control method based on preset performance comprises the following steps:
step (1), defining a common coordinate system and a motion variable of an aircraft;
step (2), establishing a multi-rigid-body six-degree-of-freedom model of the variable-profile aircraft based on a Newton Euler method;
step (3), designing performance constraints of the aircraft attitude response process according to the flight environment and the flight mission of the aircraft, and completing constraint simplification;
and (4) combining the backstepping sliding mode control with the extended state observer on the basis of the three steps to complete the multi-dimensional composite attitude control design of the aircraft.
Further, the method for defining the common coordinate system of the aircraft in the step (1) comprises the following steps: (1) selecting an aircraft (mass center) emission point A as an origin in a ground coordinate system Axyz, wherein Ax is an intersection line of a trajectory plane and a horizontal plane, the direction pointing to a target point is positive, ay is vertical to Ax and is upward in a plumb surface containing Ax, and Az is determined according to a right-hand rule (the invention is based on a threo coordinate system); (2) body coordinate system Ox 1 y 1 z 1 Selecting the centroid O of the aircraft as the origin of coordinates, ox 1 Is the longitudinal axis of the body, pointing to the head as positive, oy 1 Axis in longitudinal symmetry of the body, with Ox 1 Vertical, with upward positive, oz 1 Determining according to a right-hand rule; (3) ballistic coordinate system Ox 2 y 2 z 2 The origin of the coordinate coefficient is the mass center O, ox of the aircraft 2 Velocity vector coincidence of axis with center of mass of body, oy 2 The axis being in the vertical plane containing the velocity vector and perpendicular to Ox 2 Upward is positive, oz 2 Determining according to a right-hand rule; (4) velocity coordinate system Ox 3 y 3 z 3 The origin of the coordinate system is the centroid O, ox of the aircraft 3 Velocity vector coincidence of axis with center of mass of body, oy 3 In the longitudinal symmetry plane of the body and perpendicular to Ox 3 Upward is positive, oz 3 Determined by the right-hand rule and called the air flow coordinate system.
Further, the step (1) of defining the commonly used motion variables of the aircraft comprises the following steps:
(1) the attitude angle includes: pitch angle of the wing
Yaw angle psi and roll angle gamma;
(2) the ballistic angle includes: ballistic inclination angle theta and ballistic declination angle psi v ；
(3) The airflow angle includes: angle of attack α and sideslip angle β;
(4) speed ramp angle gamma v 。
Further, in step (2), the established variable-profile aircraft multi-rigid-body six-degree-of-freedom model comprises:
(1) an atmospheric density model;
(2) a gravitational acceleration model;
(3) a wind disturbance model.
Further, step (3), selecting key performance indexes in a response process according to the flight environment and the flight mission, establishing corresponding constraints, constructing a homomorphic mapping function according to the performance constraints, and realizing the transformation of unconstrained errors;
wherein the performance indicators include: (1) overshoot; (2) a response time; (3) a steady state error; (4) the response process completes the envelope curve.
Further, step (4), designing the multi-dimensional composite attitude control of the aircraft, comprising:
(1) a differential tracker: arranging a proper transition process for the command signal to avoid overlarge initial impact, calculating to obtain a differential signal of the command, and having a certain effect of inhibiting noise interference;
(2) backstepping sliding mode controller: the backstepping method decomposes the design problem of a high-order complex system into a plurality of low-order subsystems, and then designs a sliding mode surface, so that the overall Lyapunov stability of the system is ensured;
(3) extended State Observer (ESO): the "sum disturbance" of the system (including modeling errors, internal disturbances, and parameter uncertainties, etc.) is observed and feedback compensated into the control.
Has the advantages that:
according to the method, a Newton Euler method based on the theorem that a rigid body rotates around a fixed point and a parallel axis is adopted, so that a dynamic model of the aircraft is perfected, and the dynamic characteristics generated during the appearance transformation can be more accurately reflected; setting different performance envelopes for the aircraft attitude response process based on multi-dimensional composite control with preset performance, describing the transient state (such as approach rate, overshoot and the like) and the steady state performance (such as control precision and the like) of a controlled system through the convergence characteristic of a performance envelope function, designing homomorphic mapping to complete constraint simplification, and establishing direct relation between controller parameters and control effects; in the design process of a backstepping sliding mode control law, the introduction of a differential tracker and an extended state observer greatly improves the self-adaptive capacity of the controller.
Drawings
FIG. 1 is a diagram of a multi-dimensional composite control scheme for a variable profile aircraft based on preset performance according to the present invention;
FIG. 2 is a parameter schematic diagram of a variable sweep aircraft of the present invention;
FIG. 3 is a block diagram of a control system flow based on default performance according to the present invention;
FIG. 4 is a graph of the S (ε) function of the present invention;
FIG. 5 is a graph of the response of the control simulation angle of attack of the present invention;
FIG. 6 is a graph of the control simulation sideslip angle response of the present invention;
FIG. 7 is a graph of a control simulated roll angle response of the present invention;
FIG. 8 is a graph of a simulation of the predetermined performance angle of attack error in accordance with the present invention;
FIG. 9 is a graph of a preset performance sideslip angle error simulation of the present invention;
FIG. 10 is a graph of a predetermined performance roll angle error simulation according to the present invention.
Detailed Description
In order to make the objects, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are some, but not all, embodiments of the present invention. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
Referring to the attached figure 1, the invention discloses a variable-profile aircraft multi-dimensional composite control method based on preset performance, which comprises the following steps:
step (1), defining a common coordinate system and a motion variable of an aircraft;
step (2), establishing a multi-rigid-body six-degree-of-freedom model of the variable-profile aircraft based on a Newton Euler method;
step (3), designing performance constraints of the aircraft attitude response process according to the flight environment and the flight mission of the aircraft, and completing constraint simplification;
and (4) combining the backstepping sliding mode control with the extended state observer on the basis of the three steps to complete the multi-dimensional composite attitude control design of the aircraft.
The method for defining the common coordinate system of the aircraft in the step (1) comprises the following steps: (1) selecting an aircraft launching point A as an origin point in a ground coordinate system Axyz, wherein Ax is an intersection line of a trajectory plane and a horizontal plane, the direction pointing to a target point is positive, ay is in a plumb surface containing Ax, the Ax is vertical upwards, and Az is determined according to a right-hand rule; (2) body coordinate system Ox 1 y 1 z 1 Selecting the centroid O of the aircraft as the origin of coordinates, ox 1 Is the longitudinal axis of the body, pointing to the head as positive, oy 1 Axis in longitudinal symmetry of the body, with Ox 1 Vertical, with upward positive, oz 1 Determining according to a right-hand rule; (3) ballistic coordinate system Ox 2 y 2 z 2 The origin of the coordinate coefficient is the mass center O, ox of the aircraft 2 Velocity vector coincidence of axis with center of mass of body, oy 2 The axis being in the vertical plane containing the velocity vector and perpendicular to Ox 2 Upward is positive, oz 2 Determining according to a right-hand rule; (4) velocity coordinate system Ox 3 y 3 z 3 The origin of the coordinate system is the centroid O, ox of the aircraft 3 Velocity vector coincidence of axis with center of mass of body, oy 3 In the longitudinal symmetry plane of the body and perpendicular to Ox 3 Upward is positive, oz 3 Determined according to the right hand rule, called the air flow coordinate system.
The aircraft coordinate system defined in the step (1) adopts a threo coordinate system, and the related motion variables are defined as follows:
(1) attitude angle (included angle between ground coordinate system and body coordinate system):
(2) ballistic angle (angle between the ground coordinate system and the ballistic coordinate system):
(3) the airflow angle (the included angle between the speed coordinate system and the body coordinate system) comprises:
Velocity inclination angle (included angle between ballistic coordinate system and velocity coordinate system): aircraft Oy 3 Axis and vertical plane Ox containing velocity vector 2 y 2 The included angle of (a).
And (2) establishing a multi-rigid-body six-degree-of-freedom dynamic model of the variable-shape aircraft. Theoretically, whatever the method is used, the finally obtained kinetic equation should be equivalent, but different methods have different advantages and application scenarios, and some methods are more suitable for the dynamic modeling of a multi-rigid system. For example, when using newton euler method for modeling, the active force and constraint force of each interaction must be considered, and a balanced type of force and moment is established for each partial rigid body, so as to obtain a multi-rigid-body kinematic model. And the structural parameters and the pneumatic parameters of the variable-profile aircraft continuously change along with time, compared with the traditional single rigid body fixed configuration aircraft, the dynamic model has the characteristics of strong coupling and high-order strong nonlinearity due to the fact that more time-varying factors exist in the dynamic characteristics. The invention mainly aims at attitude control of an aircraft in a shape changing process, and brings stress, structural parameter variation and aerodynamic change generated by interaction between rigid bodies into a dynamic model, but does not consider the influence of dynamic response of control equipment and an actuating mechanism on the model.
The variable-backswept-wing aircraft is used as an explanatory object of the invention, a dynamic model of the variable-backswept-wing aircraft is established by adopting a Newton Euler method based on the parallel axis theorem, and the following assumptions are made on the premise of ensuring the reasonability:
(1) Neglecting energy consumption in the flight process, and regarding the wings of the airplane body as a rigid body with fixed quality;
(2) The distance between the centroid position and the connecting point of the wing and the fuselage is constant;
(3) The left wing and the right wing deform synchronously, the overall structure of the aircraft is bilaterally symmetrical, and the wings of the aircraft body always move in the same plane;
the variable sweep aircraft is shown schematically in FIG. 2 in terms of the structural configuration with the fuselage centroid O 1 A coordinate system is established for the origin, and the parameters are defined as follows:
the dynamic equation of the center of mass of the aircraft is established as follows:
the controller designed by the invention is mainly used for ensuring that the attitude angle can quickly and accurately track the command signal, so that an attitude control system model is established as follows:
wherein c is 1 ～c 9 The expression of (c) is as follows:
c 1 ＝I by +2I 1y +2I 2y -2m 1 (h 2 +i 2 +l 2 -2hl sinη+2il cosη)
c 3 ＝I bxy
c 5 ＝I bx +2I 1x +2I 2x -2m 1 (i+l cosη) 2
c 9 ＝I bz +2I 1z +2I 2z -2m 1 (h-l sinη) 2
in the formula, α, β, μ represent an attack angle, a sideslip angle and a roll angle, ω x 、ω y 、ω z Respectively representing the angular velocities of rotation in three directions around the body coordinate system. The foot mark b represents the fuselage, 1 represents the left wing, 2 represents the right wing, I represents the moment of inertia of the aircraft, F represents the resultant external force of the aircraft, M is the resultant external moment, M b 、m 1 Respectively, fuselage mass and wing mass, c1-c9 have no specific meaning and are set to facilitate writing a parametric form of the aircraft model definition.
The flight environment model established in the step (2) is as follows:
(1) atmospheric density model:
the most used atmospheric density model in the world is the american standard atmospheric model (SA 76), and with the atmospheric density value at sea level as a reference value, the calculation formula of atmospheric density at different altitudes is as follows:
h = 20.01631-32.1619 km:
T＝211.552(K)
ρ/ρ SL ＝0.032722W
h =32.1619 to 47.3501 km:
T＝250.654(K)
ρ/ρ SL ＝3.2618×10 -3 W -13.2011
where ρ is the atmospheric density, ρ SL The standard atmospheric density of sea level is 1.2250kg/m 3 (ii) a T represents temperature in K; y represents the earth radius, and W represents the terrain height.
(2) A gravity acceleration model:
for the variable-profile aircraft, due to wide flight space and large height difference, the gravity acceleration needs to be accurately modeled, and the relationship between the gravity acceleration and the height can be obtained according to a universal gravitation formula as follows:
wherein g is 0 The sea level acceleration is taken to be 9.80665m/s 2 ；R 0 The average radius of the earth is taken to be 6371.393km, and H is the altitude of the aircraft.
(3) A wind interference model:
the influence of wind interference on an aircraft can be divided into three parts: actual flying speed V' due to wind speed; additional angle of attack Δ α due to wind speed; additional slip angle Δ β due to wind speed.
Normally, only horizontal wind disturbances are considered. In the local vertical coordinate system, the wind vector is considered to be in the local vertical coordinate system Ox u Direction of wind theta at right of axis W Is positive. Wherein, the wind vector V W Decomposition into a longitudinal wind W in a local vertical coordinate system x And cross wind W z 。
Wind vector V W In the local vertical coordinate system, the following is expressed:
the actual flying speed V' (the speed derivative is used for the speed derivative) finally obtained in case of wind disturbance
Expressed), additional angle of attack Δ α, additional angle of sideslip Δ β are as follows:
and (3) designing performance constraint of the aircraft attitude response process according to the aircraft flight environment and the flight task, and completing constraint simplification. The method comprises the steps of firstly, comprehensively considering the transient performance and the steady-state performance requirements of an aircraft attitude response curve, evaluating the control performance, designing an expected control performance envelope curve according to actual requirements, and converting an error curve in the response process according to a designed constraint curve to obtain an unconstrained error curve, so that the design complexity of a controller is reduced.
Referring to fig. 3, the invention discloses a specific control flow chart of a multidimensional composite controller based on preset performance designed for the attitude angle of an aircraft. According to the established dynamics model of the variable sweepback aircraft, the aircraft is a multi-input multi-output strong coupling system, and decoupling control is difficult to realize. In the design process of the preset performance controller, a performance constraint function needs to be set for a response curve, the function has the characteristics of smoothness and monotone decreasing, and the control performances such as response time, overshoot, steady-state error and the like are limited. Without loss of generality, the following two forms of constrained performance function p may be chosen i (t)：
p i (t)＝(p 0 -p ∞ )e -st +p ∞
In the formula: p is a radical of 0 ＞p ∞ ＞0，s，s 1 ，s 2 ＞0，p 0 And s 2 Limiting the maximum overshoot of the response, p ∞ Limiting the steady state error, s and s 1 The response time is controlled.
According to the positive and negative errors of the error initial value, the upper and lower bounds of the error track are defined as follows:
-δ i p i (t)＜e i (t)＜p i (t)if e i (0)＞0
-p i (t)＜e i (t)＜δ i p i (t)if e i (0)＜0
wherein 0 is larger than or equal to delta i Less than or equal to 1, constraint is applied to the track boundary, e i (t) denotes error, e i (0) Indicating the initial error. And for any variable i and any time t, the above formula is always ensured to be established, namely the preset performance control is realized. The further processing of the above formula is simplified as follows:
due to the introduction of the performance function, additional constraint is added to a control system, and the design difficulty of the controller is increased, so that after the design of the preset performance function is completed, a second step of converting a constrained control problem into an unconstrained control problem through homomorphic mapping is needed, and space peer-to-peer mapping is completed. Defining:
e i (t)＝S i (ε i )p i (t)
wherein epsilon i Representing unconstrained errors after isoblastoid mapping, S i (ε i ) To be a mapping function, it needs to have the following three properties:
1)S i (ε i ) Is a smooth and strictly monotonically increasing function
2)
3)
Mapping function S satisfying the above properties i (ε i ) It can be deduced that:
and further ensuring that the error trajectory is within the range of the preset performance function. Due to the nature of the mapping function and p i (t)≥p i∞ > 0, the inverse mapping can be obtained:
the homomorphic mapping function mainly comprises a logarithm mapping function and a tangent mapping function, wherein the logarithm mapping function S i (ε i ) The following:
referring to FIG. 4, it can be seen that when the error response curve is within the preset performance constraints, the resulting ε i (t) epsilon (— infinity, and + ∞), compared with the original error control system, the converted system is a constraint-free control system, thereby reducing the complexity of the design process of the controller. Because the original system and the new system are homomorphic mapping, if epsilon can be guaranteed i (t) bounded, it can be guaranteed
And because e i (t)＝S i (ε i )p i (t) and thus the error trajectory e can be guaranteed over the entire time domain i (t) satisfies a preset performance constraint, i.e.
At the same time due to p i (t) monotonically decreasing property, error trajectory e i (t) eventually converges to within the defined limited set.
And (4) combining the backstepping sliding mode control with the extended state observer on the basis of the three steps to complete the design of the control law of the aircraft. For the convenience of controller design, the conversion is performed for the established aircraft model. Let y = [ alpha beta mu ]] T ,ω＝[ω x ω y ω z ] T ，u＝[M x M y M z ] T Establishing a relationship between the torque and the rudder deflection function as follows:
wherein q represents dynamic pressure, S a Representing the aircraft reference area, b representing the aircraft characteristic length, m β 、m α Representing the corresponding sideslip angle and angle of attack coefficients;
and
the rudder deflection coefficients of a rolling channel, a yawing channel and a pitching channel are obtained; delta x 、δ y And delta z Respectively representing the rudder deflection of a rolling channel, a yawing channel and a pitching channel; w is a Mx 、w My And w Mz Errors in the fitting process of the moments and the rudder deflection in the three directions are represented; the aircraft model is converted into the following form:
wherein f is 1 、f 2 、g 1 、g 2 As system parameters, the expression is:
d 1 、d 2 for external interference, the system parameter f is not accurate during the actual flight process, so 1 、f 2 、g 1 、g 2 The exact value cannot be obtained in real time, so it is written as:
in the formula: f. of 10 、f 20 、g 10 、g 20 Respectively known system parameters f 1 、f 2 、g 1 、g 2 A nominal value of (d); delta of 1 、Δ 2 For the nonlinear generalized uncertainty term of the system, the expression is as follows:
and (4) establishing a controller combining a backstepping method and a sliding mode, compensating error items caused by interference including pneumatic parameters, a deformation process, an external environment and the like by introducing an extended state observer, and finally limiting a conversion error in a bounded range. On the basis of the three steps, the method is combined with a backstepping sliding mode to control the converted error curve, wherein the outer ring controls the converted attitude error curve by introducing virtual control quantity angular speed, and the inner ring controls the angular speed by actual control quantity; and the design process of the controller seriously depends on the parameter precision of the aircraft, and the model uncertainty is compensated by introducing an extended state observer.
The backstepping method is based on a recursion idea, decomposes a high-order complex system into a plurality of low-order subsystems, and controls each subsystem to further realize the control of the high-order system; the sliding mode control completes the tracking of the instruction signal by leading the tracking error to tend to a designed sliding mode surface, and the control process of the sliding mode control has discontinuity, so that the sliding mode control is a special nonlinear control method; the extended state observer regards the disturbance as an additional state variable, dynamically observes the additional state variable, and completes compensation in a feedback mode, so that the extended state observer does not depend on a specific model and has strong universality.
The first step is to introduce a virtual control quantity [ omega ] x ,ω y ,ω z ]Attitude angle tracking error e = y-y d Y denotes the aircraft flight attitude, y d Representing an aircraft attitude instruction, and selecting a Lyapunov function form in order to realize that the error control after the homoembryo mapping is in a limited range:
the first derivative of the attitude angle error ε of the isoembryo map is:
wherein
The material is brought into the available state,
the angular velocity form is chosen as follows:
brought in to
The derivative of the Lyapunov function selected in the first step is
The second step updates the Lyapunov function due to the angular velocity [ omega ] x ,ω y ,ω z ] T Is a virtual control quantity, and further updates the lyapunov function as follows:
wherein
Indicating an attitude angular velocity error. And also
Will be provided with
Is brought into availability
At this time, the derivative of the new Lyapunov function is
The control law is selected as follows
Then
In said step (4), the system error is obtained by subtracting the command signal from the actual signal. Due to the inertia of the system, the response signal cannot jump with the jump of the command signal, which not only increases the amount of control, but also exacerbates the contradiction between overshoot and rapidity. Furthermore, since the initial error is too large, it is necessary to ensure that the preset initial error is larger than the actual error, which limits the design of the performance function. The invention designs a differential Tracker (TD), which reduces initial influence and inhibits noise interference by designing a transitional process of a command signal, and obtains a buffer signal and a differential signal thereof by tracking a differentiator. The differential tracker is formed as follows:
wherein x is 1 Representing the tracking signal, x 2 A differential signal, fhan (x), representing the tracking signal 1 ,x 2 And r, h) represents the fastest control comprehensive function of the system, and the concrete formula is as follows:
the value of r reflects the acceleration of the transition, the larger the value, the faster the tracking. x is a radical of a fluorine atom 2 The tracking signal is a differential signal of the tracking signal, h represents a sampling step length, and the anti-interference capability can be improved by selecting proper h.
In step (4), Δ 1 And Δ 2 Represents an uncertainty in the system that is compensated for by an established ESO-based control system compensator. Order:
i.e. h 1 And h 2 Representing the total disturbance Δ 1 And Δ 2 The ESO observer is designed in the form:
wherein u represents the attitude angular velocity of the virtual controlled variable in the outer ESO, and the actual controlled variable torque, a, beta, in the inner ESO 1 、β 2 And λ are all parameters to be designed, such that z 1 Is close to y, z 2 Approximating the expression Δ, fal (e, a, λ) as follows:
referring to fig. 5-7, the invention discloses the change of each attitude angle response curve under the condition that the sweep back angle is changed at a constant speed, the initial attack angle is 6 degrees, the attack angle instruction is given 13 degrees, the sideslip angle is kept unchanged at 0 degree, the initial roll angle is 0 degree, and the roll angle instruction is given 10 degrees. Referring to fig. 8-10, the present invention discloses a comparison of the attitude angle error curve with the pre-set performance constraint curve under the above conditions.
The invention designs a multi-dimensional composite controller based on preset performance aiming at a variable-profile aircraft. Compared with the defects of the traditional single rigid body six-degree-of-freedom modeling process in the aspect of dynamic characteristic description, the multi-rigid body modeling method used by the invention is more suitable for modeling research of the variable-profile aircraft. The method has the advantages that the relation between the control parameters and the response curve is determined, and the controlled object is converted into a non-constraint system from a constraint system by converting the expected performance index into the visual error constraint; the attitude angle based on the preset performance control has the advantages of high response speed, small overshoot and stable state error controlled within a preset error range, and the transient state error and the stable state error are ensured to meet the preset performance constraint. In addition, the extended state observer is introduced, so that the controller has strong robustness and self-adaptive capacity, and can keep a good control effect under the simultaneous action of different interferences.
According to the invention, the Newton-Euler method based on the theorem that the rigid body rotates around the fixed point and the parallel axis is adopted, so that the dynamic model of the aircraft is perfected, and the dynamic characteristics generated during the appearance transformation can be more accurately reflected; setting different performance envelopes for the aircraft attitude response process based on multi-dimensional composite control with preset performance, describing the transient state (such as approach rate, overshoot and the like) and the steady state performance (such as control precision and the like) of a controlled system through the convergence characteristic of a performance envelope function, designing homomorphic mapping to complete constraint simplification, and establishing direct relation between controller parameters and control effects; in the design process of a backstepping sliding mode control law, the introduction of a differential tracker and an extended state observer greatly improves the self-adaptive capacity of the controller.
Finally, it should be noted that: the above examples are only intended to illustrate the technical solution of the present invention, and not to limit it; although the present invention has been described in detail with reference to the foregoing embodiments, it will be understood by those of ordinary skill in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some technical features may be equivalently replaced; and such modifications or substitutions do not depart from the spirit and scope of the corresponding technical solutions of the embodiments of the present invention.
Claims (6)
1. A variable-profile aircraft multi-dimensional composite control method based on preset performance is characterized by comprising the following steps:
step (1), defining a common coordinate system and a motion variable of an aircraft;
step (2), establishing a multi-rigid-body six-degree-of-freedom model of the variable-profile aircraft based on a Newton Euler method;
step (3), designing performance constraints of the aircraft attitude response process according to the flight environment and the flight mission of the aircraft, and completing constraint simplification;
and (4) combining the backstepping sliding mode control with the extended state observer on the basis of the three steps to complete the multi-dimensional composite attitude control design of the aircraft.
2. The multi-dimensional compound control method for the variable-profile aircraft based on the preset performance as claimed in claim 1, wherein the method for defining the common coordinate system of the aircraft in the step (1) comprises the following steps: (1) selecting an aircraft launching point A as an origin point in a ground coordinate system Axyz, wherein Ax is an intersection line of a trajectory plane and a horizontal plane, the direction pointing to a target point is positive, ay is in a plumb surface containing Ax, the Ax is vertical upwards, and Az is determined according to a right-hand rule; (2) coordinate system Ox of body 1 y 1 z 1 Selecting the centroid O of the aircraft as the origin of coordinates, ox 1 Is the longitudinal axis of the body, pointing to the head as positive, oy 1 Axis in longitudinal symmetry of the body, with Ox 1 Vertical, upward is positive, oz 1 Determining according to a right-hand rule; (3) ballistic coordinate system Ox 2 y 2 z 2 The origin of the coordinate coefficient is the mass center O, ox of the aircraft 2 Velocity vector coincidence of axis with center of mass of body, oy 2 The axis being in the vertical plane containing the velocity vector and perpendicular to Ox 2 Upward is positive, oz 2 Determining according to a right-hand rule; (4) velocity coordinate system Ox 3 y 3 z 3 The origin of the coordinate system is the centroid O, ox of the aircraft 3 Velocity vector coincidence of axis with center of mass of body, oy 3 In the longitudinal symmetry plane of the body and perpendicular to Ox 3 Upward is positive, oz 3 Determined according to the right hand rule, called the air flow coordinate system.
3. The multi-dimensional compound control method for the variable-profile aircraft based on the preset performance according to claim 2, wherein the step (1) of defining the commonly used motion variables of the aircraft comprises the following steps:
(1) the attitude angle includes: pitch angle of the wing
Yaw angle psi and roll angle gamma;
(2) the ballistic angle includes: ballistic inclination angle theta and ballistic declination angle psi v ；
(3) The airflow angle includes: angle of attack α and sideslip angle β;
(4) velocity ramp angle gamma v 。
4. The multi-dimensional compound control method for the variable-profile aircraft based on the preset performance as claimed in claim 3, wherein the step (2) is implemented by establishing a multi-rigid-body six-degree-of-freedom model of the variable-profile aircraft, which comprises:
(1) an atmospheric density model;
(2) a gravitational acceleration model;
(3) a wind disturbance model.
5. The multi-dimensional composite control method for the variable-profile aircraft based on the preset performance is characterized in that in the step (3), key performance indexes in a response process are selected according to a flight environment and a flight task, corresponding constraints are established, a homoembryonic mapping function is constructed according to the performance constraints, and transformation without constraint errors is achieved;
wherein the performance indicators include: (1) overshoot; (2) a response time; (3) a steady state error; (4) the response process completes the envelope curve.
6. The variable-profile aircraft multi-dimensional compound control method based on the preset performance is characterized in that the step (4) of designing the aircraft multi-dimensional compound attitude control comprises the following steps:
(1) a differential tracker: arranging a proper transition process for the instruction signal to avoid overlarge initial impact, calculating to obtain a differential signal of the instruction, and having a certain effect of inhibiting noise interference;
(2) backstepping sliding mode controller: the backstepping method decomposes the design problem of a high-order complex system into a plurality of low-order subsystems, and then designs a sliding mode surface to ensure the global Lyapunov stability of the system;
(3) expanding the state observer: the "sum disturbance" of the system is observed and feedback compensated into the control.
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| CN117289598B (en) * | 2023-08-01 | 2024-06-11 | 北京理工大学重庆创新中心 | Method and system for controlling backstepping sliding mode of aircraft |
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