CN113448339A - Aircraft attack angle tracking control method based on virtual inversion - Google Patents
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Abstract
The invention relates to an aircraft attack angle tracking control method based on virtual inversion, and belongs to the technical field of aircraft overload control. Firstly, an attack angle sensor is adopted to measure an attack angle of an aircraft, and the attack angle sensor is compared with an attack angle instruction to obtain an attack angle error value; and then measuring pitch angle rate signals by adopting a gyroscope, constructing a virtual value of the pitch angle rate signals by adopting a method combining virtual control and inversion control, comparing the virtual value with the pitch angle rate signals to obtain an angle rate error signal, constructing a virtual secondary state, solving a secondary virtual uncertain signal, finally synthesizing the error signal and the virtual signal to form an error comprehensive signal, obtaining a final pitch angle comprehensive signal through nonlinear transformation, transmitting the final pitch angle comprehensive signal to an aircraft pitch rudder system, and controlling an aircraft attack angle to track a given attack angle signal. The invention has the advantages of high attack angle control precision, high speed, good stability and good dynamic performance.
Description
Technical Field
The invention relates to the technical field of aircraft control, in particular to a method for accurately tracking and controlling an aircraft attack angle by combining virtualization and inversion.
Background
The traditional aircraft attitude control has the characteristics of high stability margin and high reliability, and is widely applied since birth. Secondly, overload control has also found more use in smart aircraft in recent years due to its advantage of good maneuverability. The attack angle is related to the stress and the moment of the aircraft, so that the most close relationship is provided with the stability of the aircraft. It is also in direct relation to the maneuvering of the aircraft. Therefore, the angle of attack can be directly controlled and monitored, and extensive research is already carried out in foreign aircrafts. The main reason for the relatively late domestic research is that the difficulty of accurate measurement of the angle of attack is particularly high due to the insufficient angle of attack sensing technology. However, the accuracy of the angle of attack measurement in the future must be further improved, and various novel control technologies must also be applied to aircraft control. Compared with the traditional PID control, the inversion control has the advantages of strict theory due to layer-by-layer backward design, the inversion reasoning is tighter by the method of predicting the system uncertainty through virtual control, and based on the background reasons, the invention provides a method for stably tracking and controlling the attack angle of an aircraft by adopting the virtual inversion, and the method has the advantages of high response speed and high control precision, so that the method has high engineering application value.
It is to be noted that the information invented in the above background section is only for enhancing the understanding of the background of the present invention, and therefore, may include information that does not constitute prior art known to those of ordinary skill in the art.
Disclosure of Invention
The invention aims to provide an aircraft attack angle tracking control method based on virtual inversion, and further solves the problems of low attack angle tracking accuracy and poor dynamic characteristics caused by the limitations and defects of the related technology at least to a certain extent.
According to one aspect of the invention, an aircraft attack angle tracking control method based on virtual inversion is provided, and comprises the following steps:
and step S10, mounting an attack angle sensor on the aircraft, measuring the attack angle of the aircraft, comparing the measured attack angle with the attack angle instruction signal to obtain an attack angle error signal, and then mounting a rate gyroscope to measure the pitch angle rate of the aircraft.
And step S20, constructing a virtual value of the pitch angle rate signal according to the attack angle error signal, the attack angle signal, the pitch angle rate signal and the like.
And step S30, comparing the virtual value of the pitch angle rate signal with the pitch angle rate signal to obtain an error signal of the pitch angle rate, constructing a virtual secondary state, and solving a secondary virtual uncertain signal.
And step S40, performing linear synthesis according to the secondary virtual uncertainty item signal, the attack angle error signal and the pitch angle rate error signal to obtain an error comprehensive signal.
And step S50, according to the error comprehensive signal, carrying out nonlinear transformation to obtain a final pitch angle comprehensive signal, transmitting the final pitch angle comprehensive signal to a pitch rudder, and controlling an attack angle tracking set value of an aircraft pitch channel.
In an exemplary embodiment of the invention, a rate gyroscope is mounted to measure the pitch rate of the aircraft, and the resulting pitch rate signal is designated ωzInstalling attack angle sensor on the aircraft, measuring the attack angle of the aircraft, recording the measured attack angle as alpha, setting the command signal of the attack angle as alphadAnd comparing the angle of attack signal with the angle of attack command signal to obtain an angle of attack error signal, and recording the angle of attack error signal as eαThe calculation method is as follows: e.g. of the typeα＝αd-α。
In an exemplary embodiment of the invention, constructing the virtual value of the pitch rate signal from the angle of attack error signal, the angle of attack signal, the pitch rate signal, and the virtual value of the pitch rate signal comprises:
z(n+1)＝z(n)+zdT；
where α is the angle of attack signal, eαIs an attack angle error signal, omega is a pitch angle rate signal, z is a virtual first-order state,
is a primary virtual uncertainty, zdIncrease of primary state, z (n) is the nth data of virtual primary state, which is initialThe value is chosen to be 0, i.e. z (1) ═ 0, T, k1、k2、k3、k4、ε1、ka1The detailed settings are described in the following examples. OmegadI.e. the virtual value of the pitch angle rate sought.
In an exemplary embodiment of the present invention, comparing the virtual value of the pitch angle rate signal with the pitch angle rate signal to obtain an error signal of the pitch angle rate, and constructing a virtual secondary state, wherein solving the secondary virtual uncertain signal includes:
eω＝ω-ωd；
s(n+1)＝s(n)+sdT1；
y(n+1)＝y(n)+ydT1；
wherein ω isdFor a virtual value of the pitch angle rate signal, eωError signal for pitch angle rate, eαIs an angle of attack error signal, omega is a pitch angle rate signal, y is a virtual two-stage state, ydIncrease of secondary state, sdFor the two-stage error increment, s (n) is the nth data of the two-stage integral quantity s, y (n) is the nth data of the virtual two-stage state, and the initial value is 0, i.e. y (1) is 0, T1、k6、k7、k8、k9、k10、ε2、ε3The detailed settings are described in the following examples.
I.e. the second level virtual uncertainty term sought.
In an exemplary embodiment of the invention, linearly synthesizing the secondary virtual uncertainty signal, the angle of attack error signal, and the pitch angle rate error signal to obtain an error synthesized signal includes:
where u is the desired error sum signal, eαIs an angle of attack error signal, omega is a pitch angle rate signal, eωIs the error signal for the pitch angle rate,
is a secondary virtual uncertainty. k is a radical of11、k12、k13、k14、k15The detailed settings are described in the following examples.
In an exemplary embodiment of the present invention, performing a nonlinear transformation on the error synthesized signal to obtain a final pitch angle synthesized signal includes:
where u is the error integration signal, ufFor the final pitch integrated signal notation, kf1,kf2,kf3The detailed design of the control parameter is described in the following examples.
On the basis, the comprehensive signal of the pitching channel is transmitted to the pitching rudder of the aircraft, so that the aircraft attack angle can track a given attack angle instruction, and the aircraft control task of the pitching channel is completed. The invention provides a method for stably tracking and controlling the attack angle of an aircraft by adopting attack angle measurement and angular velocity feedback, and particularly combines virtual control and inversion control, so that the rapidity and the precision of the attack angle control are greatly improved, and the method has high theoretical innovation value and engineering application value.
It is to be understood that both the foregoing general description and the following detailed description are exemplary and explanatory only and are not restrictive of the invention, as claimed.
Drawings
The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the invention and together with the description, serve to explain the principles of the invention. It is obvious that the drawings in the following description are only some embodiments of the invention, and that for a person skilled in the art, other drawings can be derived from them without inventive effort.
FIG. 1 is a flow chart of an aircraft angle of attack tracking control method based on virtual inversion provided by the invention;
FIG. 2 is a graph of angle of attack error signals (in degrees) according to a method provided by an embodiment of the present invention;
FIG. 3 is a plot of the pitch rate signal (in degrees per second) for a method provided by an embodiment of the present invention;
FIG. 4 is a plot (without units) of virtual values of a pitch rate signal for a method provided by an embodiment of the present invention;
FIG. 5 is a two-level virtual indeterminate signal curve (unitless) for a method provided by an embodiment of the invention;
FIG. 6 is a graph of an error synthesized signal (without units) for a method provided by an embodiment of the present invention;
fig. 7 is a plot (without units) of a pitch angle integrated signal for a method provided by an embodiment of the present invention. FIG. 8 is a graph of angle of attack (in degrees) for a method provided by an embodiment of the invention;
FIG. 9 is a pitch rudder deflection angle curve (in degrees) for a method provided by an embodiment of the present invention;
Detailed Description
Example embodiments will now be described more fully with reference to the accompanying drawings. Example embodiments may, however, be embodied in many different forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the concept of example embodiments to those skilled in the art. The described features, structures, or characteristics may be combined in any suitable manner in one or more embodiments. In the following description, numerous specific details are provided to provide a thorough understanding of embodiments of the invention. One skilled in the relevant art will recognize, however, that the invention may be practiced without one or more of the specific details, or with other methods, components, devices, steps, and so forth. In other instances, well-known technical solutions have not been shown or described in detail to avoid obscuring aspects of the invention.
The invention provides an aircraft attack angle tracking control method based on virtual inversion, which comprises the steps of firstly measuring the attack angle and pitch angle rate of an aircraft, then designing a virtual value of the aircraft pitch angle rate by adopting a virtual control method, comparing the virtual value with a real value of the aircraft pitch angle rate to obtain a pitch angle rate error, designing an error comprehensive signal by means of backward estimation layer by the inversion control method, and finally carrying out nonlinear transformation to obtain a final stable tracking control law of the attack angle of a pitch channel. The method has the advantages of high control precision, high response speed and good dynamic performance, thereby having high engineering application value.
The following describes a virtual inversion-based aircraft angle of attack tracking control method according to the present invention with reference to the accompanying drawings. Referring to fig. 1, the method for tracking and controlling the attack angle of the aircraft based on the virtual inversion may include the following steps:
and step S10, mounting an attack angle sensor on the aircraft, measuring the attack angle of the aircraft, comparing the measured attack angle with the attack angle instruction signal to obtain an attack angle error signal, and then mounting a rate gyroscope to measure the pitch angle rate of the aircraft.
Specifically, firstly, an attack angle sensor is arranged on an aircraft to measure an attack angle, and the measured attack angle is recorded as alpha;
secondly, setting the attack angle according to the pitching channel control task of the aircraftLet the signal be recorded as alphad；
Then comparing the angle of attack with the angle of attack command signal to obtain an angle of attack error signal, and recording the angle of attack error signal as eαThe comparison method is as follows: e.g. of the typeα＝αd-α；
Finally, a rate gyroscope is installed on the aircraft, and the pitch angle rate of the aircraft is measured and recorded as omegaz。
And step S20, constructing a virtual value of the pitch angle rate signal according to the attack angle error signal, the attack angle signal, the pitch angle rate signal and the like.
Specifically, first, according to the attack angle signal α and the attack angle error signal eαAnd solving a primary virtual uncertainty term by a pitch angle rate signal omega as follows:
z(n+1)＝z(n)+zdT；
where z is a virtual primary state,
is a primary virtual uncertainty, zdZ (n) is the nth data of the virtual primary state, and its initial value is selected to be 0, i.e. z (1) ═ 0, T, k1、k2、k3、k4、ε1The detailed settings are described in the following examples.
Finally, a virtual value ω of the pitch angle rate is setdThe following were used:
wherein k isa1Is a constant parameter, and the detailed design is shown inThe following examples are carried out.
And step S30, comparing the virtual value of the pitch angle rate signal with the pitch angle rate signal to obtain an error signal of the pitch angle rate, constructing a virtual secondary state, and solving a secondary virtual uncertain signal.
Specifically, firstly, according to the virtual value ω of the pitch angle rate signaldComparing the pitch angle rate signal with the difference to obtain an error signal e of the pitch angle rateωThe following were used:
eω＝ω-ωd；
secondly, according to the attack angle error signal eαPitch angle rate signal omega, error signal e of pitch angle rateωSolving for the secondary virtual uncertainty term is as follows:
s(n+1)＝s(n)+sdT1；
y(n+1)＝y(n)+ydT1；
where y is a virtual secondary state,
for a secondary virtual uncertainty, ydIncrease of secondary state, sdFor the two-stage error increment, s (n) is the nth data of the two-stage integral quantity s, y (n) is the nth data of the virtual two-stage state, and the initial value is 0, i.e. y (1) is 0, T1、k6、k7、k8、k9、k10、ε2、ε3The detailed settings are described in the following examples.
And step S40, performing linear synthesis according to the secondary virtual uncertainty item signal, the attack angle error signal and the pitch angle rate error signal to obtain an error comprehensive signal.
Specifically, the linear synthesis form is as follows:
where u is the error integration signal, eαIs an angle of attack error signal, omega is a pitch angle rate signal, eωIs the error signal for the pitch angle rate,
is a secondary virtual uncertainty. k is a radical of11、k12、k13、k14、k15The detailed settings are described in the following examples.
And step S50, according to the error comprehensive signal, carrying out nonlinear transformation to obtain a final pitch angle comprehensive signal, transmitting the final pitch angle comprehensive signal to a pitch rudder, and controlling an attack angle tracking set value of an aircraft pitch channel.
Specifically, according to the error comprehensive signal u, the following nonlinear transformation is carried out to obtain the final pitching comprehensive signal which is recorded as ufThe nonlinear transformation is as follows:
wherein k isf1,kf2,kf3The detailed design of the control parameter is described in the following examples.
On the basis, debugging and selecting of all control parameters are carried out, the selection is based on the tracking performance of the attack angle, namely, comprehensive judgment is carried out on rapidity, accuracy and stability according to the selection, appropriate parameters are selected, a final pitching channel attack angle tracking control system is formed, and the control task of a pitching channel is completed.
Case implementation and computer simulation result analysis
In order to verify the correctness of the method provided by the invention, the aircraft adopted in the case is a three-channel six-degree-of-freedom nonlinear full ballistic model, and the speed of the aircraft is accelerated from 0 meter per second.
In step S10, the desired angle of attack command signal is set to αdT > 4, 4. Before 4 seconds, the speed of the aircraft at the initial stage is relatively low, so that the attack angle instruction is not set. The resulting angle of attack error signal is shown in fig. 2, and the pitch rate signal of the aircraft is shown in fig. 3.
In step S20, T is set to 0.001, k1＝0.5、k2＝0.3、k3＝1.4、k4＝0.6、ε1＝0.5、k 15. The virtual values of the pitch angle rate signals constructed according to the attack angle error signals, the attack angle signals, the pitch angle rate signals are shown in figure 4.
In step S30, T is set1＝0.002、k6＝0.4、k7＝2.5、k8＝0.1、k9＝1.5、k10＝1.2、ε2＝2、ε31.5. The resulting secondary virtual uncertain signal is shown in fig. 5.
In step S40, k is set11＝0.5、k12＝4、k13＝0.2、k14＝1、k150.3. The resulting error integration signal is shown in fig. 6.
In step S50, k is setf1＝0.3,kf2＝0.4,kf3The final pitch angle integrated signal is obtained as shown in fig. 7, when the value is 0.3. The angle of attack of the resulting aircraft is shown in fig. 8 and the aircraft pitch rudder deflection angle is shown in fig. 9.
Before 4 seconds, the aircraft does not control the attack angle, mainly because the speed of the aircraft is small in the initial section, and therefore the attitude stability control is mainly performed. After 4s, the given angle of attack command for the aircraft is 4 degrees, and as can be seen from fig. 9, after 4 seconds the rudder deflection curve has stabilized to within 4 degrees, and the rudder deflection signal can be steered without exceeding the available range. As can be seen from fig. 8, the attack angle curve is stabilized at about 4 degrees after 4 seconds, and the overshoot is small, the overshoot is about 5%, and particularly the response speed of the attack angle is less than 0.2 s. Therefore, the method provided by the invention has a good control effect and a high engineering application value.
Other embodiments of the invention will be apparent to those skilled in the art from consideration of the specification and practice of the invention disclosed herein. This application is intended to cover any variations, uses, or adaptations of the invention following, in general, the principles of the invention and including such departures from the present disclosure as come within known or customary practice within the art to which the invention pertains. It is intended that the specification and examples be considered as exemplary only, with a true scope and spirit of the invention being indicated by the following claims.
Claims (5)
1. An aircraft attack angle tracking control method based on virtual inversion is characterized by comprising the following steps:
and step S10, mounting an attack angle sensor on the aircraft, measuring the attack angle of the aircraft, comparing the measured attack angle with the attack angle instruction signal to obtain an attack angle error signal, and then mounting a rate gyroscope to measure the pitch angle rate of the aircraft.
And step S20, constructing a virtual value of the pitch angle rate signal according to the attack angle error signal, the attack angle signal, the pitch angle rate signal and the like.
And step S30, comparing the virtual value of the pitch angle rate signal with the pitch angle rate signal to obtain an error signal of the pitch angle rate, constructing a virtual secondary state, and solving a secondary virtual uncertain signal.
And step S40, performing linear synthesis according to the secondary virtual uncertainty item signal, the attack angle error signal and the pitch angle rate error signal to obtain an error comprehensive signal.
And step S50, according to the error comprehensive signal, carrying out nonlinear transformation to obtain a final pitch angle comprehensive signal, transmitting the final pitch angle comprehensive signal to a pitch rudder, and controlling an attack angle tracking set value of an aircraft pitch channel.
2. The method for tracking and controlling the attack angle of the aircraft based on the virtual inversion according to claim 1, wherein constructing the virtual value of the pitch angle rate signal according to the attack angle error signal, the attack angle signal, the pitch angle rate signal and the virtual value of the pitch angle rate signal comprises:
eα＝αd-α
z(n+1)＝z(n)+zdT；
wherein alpha is an angle of attack signal, alphadFor angle of attack command signals, eαIs an attack angle error signal, omega is a pitch angle rate signal, z is a virtual first-order state,
is a primary virtual uncertainty, zdZ (n) is the nth data of the virtual primary state, and its initial value is selected to be 0, i.e. z (1) ═ 0, T, k1、k2、k3、k4、ε1、ka1Is a constant parameter. OmegadI.e. the virtual value of the pitch angle rate sought.
3. The aircraft attack angle tracking control method based on virtual inversion according to claim 1, wherein the virtual value of the pitch angle rate signal is compared with the pitch angle rate signal to obtain an error signal of the pitch angle rate, a virtual secondary state is constructed, and solving the secondary virtual uncertain signal comprises:
eω＝ω-ωd；
s(n+1)＝s(n)+sdT1；
y(n+1)＝y(n)+ydT1；
wherein ω isdFor a virtual value of the pitch angle rate signal, eωError signal for pitch angle rate, eαIs an angle of attack error signal, omega is a pitch angle rate signal, y is a virtual two-stage state, ydIncrease of secondary state, sdFor the two-stage error increment, s (n) is the nth data of the two-stage integral quantity s, y (n) is the nth data of the virtual two-stage state, and the initial value is 0, i.e. y (1) is 0, T1、k6、k7、k8、k9、k10、ε2、ε3Is a constant parameter.
I.e. the second level virtual uncertainty term sought.
4. The aircraft attack angle tracking control method based on virtual inversion according to claim 1, wherein the linear synthesis of the secondary virtual uncertainty signal, the attack angle error signal and the pitch angle rate error signal to obtain an error synthesis signal comprises:
where u is the desired error sum signal, eαIs an angle of attack error signal, omega is a pitch angle rate signal, eωIs the error signal for the pitch angle rate,
is a secondary virtual uncertainty. k is a radical of11、k12、k13、k14、k15Is a constant parameter.
5. The method for tracking and controlling the attack angle of the aircraft based on the virtual inversion as claimed in claim 1, wherein the step of performing the nonlinear transformation according to the error comprehensive signal to obtain the final pitch angle comprehensive signal comprises:
where u is the error integration signal, ufFor the final pitch integrated signal notation, kf1,kf2,kf3The parameter is controlled to be constant.
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