CN109739250B - Self-adaptive finite time attitude control model acquisition method - Google Patents

Self-adaptive finite time attitude control model acquisition method Download PDF

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CN109739250B
CN109739250B CN201811513060.8A CN201811513060A CN109739250B CN 109739250 B CN109739250 B CN 109739250B CN 201811513060 A CN201811513060 A CN 201811513060A CN 109739250 B CN109739250 B CN 109739250B
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孙海滨
侯林林
宗广灯
杨东
齐文海
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Qufu Normal University
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Abstract

A self-adaptive finite time quantitative control method is provided for an aircraft system with quantitative input, and the self-adaptive finite time quantitative control method improves the control precision of the system and has the characteristic of high convergence speed, thereby realizing the control scheme of the aircraft to reduce communication traffic without influencing the stability and the control performance of the communication traffic.

Description

Self-adaptive finite time attitude control model acquisition method
Technical Field
The invention relates to a method and a system for acquiring a self-adaptive finite time attitude control model, a controller and a control method, in particular to a method and a system for acquiring a self-adaptive finite time attitude control model, a controller and a control method which are mainly applied to attitude control of a modular aircraft.
Background
In recent years, the modular spacecraft is paid attention, the modular spacecraft adopts a modular open network architecture, and various components of the spacecraft are modularized and assembled by utilizing the basic idea of a plug-and-play system, can reduce the cost of the spacecraft, accelerate the assembly speed, facilitate the easy disassembly and replacement of defective or outdated components, the core technology of the modularized spacecraft is wireless data communication and wireless power transmission, which is different from the traditional cable interconnection which is heavy, large and inflexible, under the technology, all functional components of the spacecraft are mutually independent and connected through a low-cost wireless network, however, the bandwidth and computational power of the wireless network responsible for data transfer between the actuator module and the control module is limited, therefore, it is important how to design a control scheme for a spacecraft to reduce traffic without affecting stability and control performance.
Disclosure of Invention
The object of the present invention is a method for obtaining an adaptive finite time attitude control model,
the object of the present invention is an adaptive finite time attitude control model acquisition system,
the object of the invention is an adaptive finite time attitude controller,
the object of the invention is a method for adaptive finite time attitude control.
In order to overcome the technical drawbacks mentioned above, it is an object of the present invention to provide an adaptive finite time attitude control model acquisition method and system, a controller and a control method, thus implementing a control scheme in an aircraft to reduce traffic without affecting stability and control performance.
In order to achieve the purpose, the invention adopts the technical scheme that:
a self-adaptive finite time attitude control model acquisition method comprises the following steps:
establishing an aircraft dynamics model and an attitude model containing external interference and quantitative input,
defining a state variable measurement error, obtaining a system model defined by the measurement state based on the model of the aircraft and the measurement error,
and (3) introducing the differential of the sliding mode differentiator estimation part signal by using an exponentiation integration method to obtain a design scheme of the self-adaptive finite time quantization controller of the aircraft.
Due to the design of the steps, the self-adaptive finite time quantization control method is provided for the aircraft system with the quantization input, the control precision of the system is improved, and the convergence speed is high, so that the control scheme of the aircraft is realized to reduce the communication traffic without influencing the stability and the control performance.
The invention designs an operation state equation set established in a CPU, which comprises the following contents:
establishing an aircraft dynamics model and an attitude model containing external interference and quantitative input:
Figure GDA0003463814490000021
wherein σ ═ σ1，σ2，σ3]Tω represents the attitude and angular velocity of the aircraft, respectively, and J ═ diag { J ═1，J2，J3Is the matrix of the inertia as well as the moment of inertia,
Figure GDA0003463814490000022
for external interference, | d (t) | ≦ κ, q (u (t)) q ═ q (u) and/or k1(t))，q(u2(t))，q(u3(t)]TIn order to quantize the input of the input,
Figure GDA0003463814490000023
I3is a unity matrix, s (σ) [0 σ ]3-σ2；-σ30σ1；σ2-σ10]. The quantizer operator is defined as
Figure GDA0003463814490000024
Wherein
Figure GDA0003463814490000025
Respectively representing the size of a quantizer dead zone and quantization density; q (t) represents the state at the previous time.
According to the measurement error, an aircraft system model represented by a measurement state is established, and the measured values and the true values of the attitude and the angular velocity meet the following relations:
Figure GDA0003463814490000026
wherein
Figure GDA0003463814490000027
Indicating the state of the measurement of the sensor,
Figure GDA0003463814490000028
indicates measurement error, and satisfies
Figure GDA0003463814490000029
Combining (1.1) and (1.2) gave a system as follows
Figure GDA00034638144900000210
Figure GDA00034638144900000211
Wherein
Figure GDA00034638144900000212
Figure GDA00034638144900000213
Figure GDA0003463814490000031
Designing a self-adaptive finite time attitude controller by utilizing an exponentiation integral method and a sliding mode differentiator:
first a new variable theta is introduced1，θ2，φ1，φ2
Whereinθ1＝||b1||2，b1＝[b11，b12，b13，b14，b15]T，
Figure GDA0003463814490000032
Figure GDA0003463814490000033
Figure GDA0003463814490000034
Figure GDA0003463814490000035
Redefining new variables
Figure GDA0003463814490000036
Wherein
Figure GDA0003463814490000037
Figure GDA0003463814490000038
Designing an adaptive controller and an adaptive update law based on the virtual control quantity
Figure GDA0003463814490000039
Figure GDA00034638144900000310
Figure GDA00034638144900000311
v20By such asThe following differential equation:
Figure GDA00034638144900000312
wherein
Figure GDA00034638144900000313
μ20＞0，μ21＞0，ρ20，ρ21，v20Is the state of the system.
The invention designs a self-adaptive finite time attitude control model acquisition system, which comprises the following contents:
according to the aircraft dynamics model and attitude model building unit 10 containing external disturbances and quantitative inputs,
from the defined state variable measurement errors, the aircraft-based model and the measurement errors a system model building unit 20 defined by the measurement states is derived,
and obtaining the self-adaptive finite time quantization controller design scheme establishing unit 30 of the aircraft according to the differential of the estimation part of the signals by using the power integration method and introducing a sliding mode differentiator.
The invention designs a self-adaptive finite time attitude controller, which comprises the following contents: a control model of an adaptive finite-time gesture based on a quantized input strategy is stored in the controller,
the invention designs that the control model of the self-adaptive finite time attitude based on the quantization input strategy is obtained according to the acquisition method of the self-adaptive finite time attitude control model based on the quantization input strategy,
the invention designs a self-adaptive finite time attitude control method based on a quantitative input strategy, which comprises the following contents:
and applying an adaptive finite time attitude controller based on a quantitative input strategy in the CPU for control.
The invention has the technical effects that: compared with the prior art, the invention has the advantages that: the invention designs a self-adaptive finite time attitude control method based on a quantization input strategy, which aims at a type of disturbed aircraft system containing measurement noise, designs a self-adaptive finite time quantization controller, eliminates the influence of the disturbance, the measurement noise and the quantization error on the system, and provides the control precision and the convergence speed of the system; the self-adaptive finite time quantization control method designed by the invention greatly improves the performance of the system under the influence of quantization errors and interference.
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In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to the drawings without creative efforts.
FIG. 1 is a flow chart of a method for obtaining an adaptive finite time attitude control model based on a quantized input strategy according to the present invention,
FIG. 2 is a schematic structural diagram of an adaptive finite time attitude controller based on a quantized input strategy according to the present invention.
Detailed Description
Terms such as "having," "including," and "comprising," as used with respect to the present invention, are to be understood as not specifying the presence or addition of one or more other elements or combinations thereof, in accordance with the examination guidelines.
In the description of the present invention, it should be noted that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc., indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings, and are only for convenience of description and simplicity of description, but do not indicate or imply that the device or element being referred to must have a particular orientation, be constructed and operated in a particular orientation, and thus, should not be construed as limiting the present invention. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.
In the description of the present invention, it should be noted that, unless otherwise explicitly specified or limited, the terms "mounted," "connected," and "connected" are to be construed broadly, e.g., as meaning either a fixed connection, a removable connection, or an integral connection; can be mechanically or electrically connected; they may be connected directly or indirectly through intervening media, or they may be interconnected between two elements. The specific meanings of the above terms in the present invention can be understood in specific cases to those skilled in the art.
In addition, the technical features involved in the different embodiments of the present invention described below may be combined with each other as long as they do not conflict with each other.
The invention is further described below with reference to the following examples, which are intended to illustrate the invention but not to limit it further.
An adaptive finite time attitude control model acquisition method according to a first embodiment of the present invention includes the steps of:
step 100: establishing an aircraft dynamics model and an attitude model containing external interference and quantitative input,
step 200: defining a state variable measurement error, obtaining a system model defined by the measurement state based on the model of the aircraft and the measurement error,
step 300: and (3) introducing the differential of the sliding mode differentiator estimation part signal by using an exponentiation integration method to obtain a design scheme of the self-adaptive finite time quantization controller of the aircraft.
In the present embodiment, the step 100 specifically includes the following contents:
under the action of gust interference and quantitative input, an aircraft dynamics model and an attitude model of the aircraft system are built:
Figure GDA0003463814490000051
wherein σ ═ σ1，σ2，σ3]Tω represents the attitude and angular velocity of the aircraft, respectively, and J ═ diag { J ═1，J2，J3}
Is a matrix of the moments of inertia,
Figure GDA0003463814490000052
for external interference, | d (t) | is less than or equal to k,
q(u(t))＝[q(u1(t))，q(u2(t))，q(u3(t)]Tin order to quantize the input of the input,
Figure GDA0003463814490000053
I3is a unity matrix, s (σ) [0 σ ]3-σ2；-σ30σ1；02-σ10]. The quantizer operator is defined as
Figure GDA0003463814490000054
Wherein
Figure GDA0003463814490000055
Respectively representing the size of a quantizer dead zone and quantization density; q (t) represents the state at the previous time.
In this embodiment, the step 200 specifically includes the following contents:
according to the measurement error, an aircraft system model represented by a measurement state is established, and the measured values and the true values of the attitude and the angular velocity meet the following relations:
Figure GDA0003463814490000056
wherein
Figure GDA0003463814490000057
Indicating the state of the measurement of the sensor,
Figure GDA0003463814490000058
indicates measurement error, and satisfies
Figure GDA0003463814490000061
Combining (1.1) and (1.2) gave a system as follows
Figure GDA0003463814490000062
Figure GDA0003463814490000063
Wherein
Figure GDA0003463814490000064
Figure GDA0003463814490000065
Figure GDA0003463814490000066
In this embodiment, the step 300 specifically includes the following contents:
designing a self-adaptive finite-time attitude controller by utilizing an exponentiation integral method and a sliding mode differentiator based on the system model established in the second step:
first a new variable theta is introduced1，θ2，φ1，φ2
Wherein 01＝||b1||2，b1＝[b11，b12，b13，b14，b15]T，
Figure GDA0003463814490000067
Figure GDA0003463814490000068
Figure GDA0003463814490000069
Figure GDA00034638144900000610
Redefining new variables
Figure GDA00034638144900000611
Wherein
Figure GDA00034638144900000612
Figure GDA00034638144900000613
Designing an adaptive controller and an adaptive update law based on the virtual control quantity
Figure GDA00034638144900000614
Figure GDA00034638144900000615
Figure GDA00034638144900000616
v20Obtained by the differential equation:
Figure GDA0003463814490000071
wherein
Figure GDA0003463814490000072
μ20＞0，μ21＞0，，ρ20，ρ21，v20Is the state of the system.
An adaptive finite time attitude control model acquisition system comprises the following contents:
according to the aircraft dynamics model and attitude model building unit 10 containing external disturbances and quantitative inputs, for obtaining a nonlinear system model,
a system model building unit 20 defined by the measurement states is derived from the defined state variable measurement errors, the aircraft-based model and the measurement errors, for obtaining a system model with measurement errors,
and obtaining an adaptive finite time quantization controller design scheme establishing unit 30 of the aircraft according to the differential of the estimation part signal of the sliding mode differentiator by using an exponentiation integration method, wherein the adaptive finite time quantization controller design scheme is used for acting on the established nonlinear system.
An adaptive finite time attitude controller, comprising the following: a control model of an adaptive finite-time gesture based on a quantized input strategy is stored in the controller,
in the present embodiment, the control model of the adaptive finite time orientation based on the quantized input strategy is obtained according to the above-mentioned adaptive finite time orientation control model obtaining method based on the quantized input strategy,
step 100: establishing an aircraft dynamics model and an attitude model containing external interference and quantitative input,
step 200: defining a state variable measurement error, obtaining a system model defined by the measurement state based on the model of the aircraft and the measurement error,
step 300: and (3) introducing the differential of the sliding mode differentiator estimation part signal by using an exponentiation integration method to obtain a design scheme of the self-adaptive finite time quantization controller of the aircraft.
An adaptive finite time attitude control method comprises the following contents:
and applying an adaptive finite time attitude controller based on a quantitative input strategy in the CPU for control.
The above embodiment is only one implementation form of the adaptive finite-time attitude control model acquisition method and system, the controller, and the control method provided by the present invention, and all that is included in the protection scope of the present invention is that components or steps in the adaptive finite-time attitude control model are increased or decreased according to other modifications of the scheme provided by the present invention, or the present invention is used in other technical fields close to the present invention.

Claims (1)

1. A self-adaptive finite time attitude control model acquisition method is characterized by comprising the following steps: the method comprises the following steps:
establishing an aircraft dynamics model and an attitude model containing external interference and quantitative input:
Figure FDA0003463814480000011
wherein σ ═ σ1,σ2,σ3]Tω represents the attitude and angular velocity of the aircraft, respectively, and J ═ diag { J ═1,J2,J3Is the matrix of the inertia as well as the moment of inertia,
Figure FDA0003463814480000012
d(t)＝[d1(t),d2(t),d3(t)]Tfor external interference, | d (t) | is less than or equal to k,
Figure FDA0003463814480000013
and κ is a normal number, q (u (t)) - [ q (u)) - ]1(t)),q(u2(t)),q(u3(t)]TIn order to quantize the input of the input,
Figure FDA0003463814480000014
I3is a matrix of the units,
s(σ)＝[0σ3-σ2；-σ3 0σ1；σ2-σ1 0]the quantizer operator is defined as
Figure FDA0003463814480000015
Wherein
Figure FDA0003463814480000016
Figure DEST_PATH_DEST_PATH_IMAGE002
Figure DEST_PATH_DEST_PATH_IMAGE004
Figure DEST_PATH_DEST_PATH_IMAGE006
Representing quantizer dead-zone size and quantization density, q, respectively-(t) represents the state at the last time instant,
based on the model of the aircraft and the measurement error, a system model defined by the measurement state is obtained, and the measured values and the true values of the attitude and the angular velocity satisfy the following relations:
Figure FDA0003463814480000021
wherein
Figure FDA0003463814480000022
Indicating the state of the measurement of the sensor,
Figure FDA0003463814480000023
indicates measurement error, and satisfies
Figure FDA0003463814480000024
Combining (0.1) and (0.2) gave a system as follows
Figure FDA0003463814480000025
Figure FDA0003463814480000026
Wherein
Figure FDA0003463814480000027
Figure FDA0003463814480000028
Figure FDA0003463814480000029
Utilizing an exponentiation integral method and a sliding mode differentiator, and designing a self-adaptive finite time attitude controller:
first a new variable theta is introduced1,θ2,φ1,φ2
Wherein theta is1＝||b1||2,b1＝[b11,b12,b13,b14,b15]T,
Figure FDA00034638144800000210
Figure FDA00034638144800000211
Figure FDA00034638144800000212
Figure FDA00034638144800000213
Redefining new variables
Figure FDA00034638144800000214
Wherein
Figure FDA00034638144800000215
Figure FDA00034638144800000216
Design adaptive finite time attitude controller and adaptive update law
Figure FDA0003463814480000031
Figure FDA0003463814480000032
δ＝diag{δ1,δ2,δ3},k2＞0,γ1＞0,ι1＞0,γ2＞0,ι2＞0,a1＞0,
Figure FDA0003463814480000033
ν20Obtained by the differential equation:
Figure FDA0003463814480000034
wherein mu20＞0,μ21＞0,ρ20,ρ21,ν20Is the state of the system.
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