CN111766776B - Pre-guiding method adopting nonlinear proportional and integral type compensation predictor - Google Patents
Pre-guiding method adopting nonlinear proportional and integral type compensation predictor Download PDFInfo
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Abstract
The invention relates to a preposed guidance method adopting a nonlinear proportional and integral compensation predictor, belonging to the field of aircraft guidance and control. Firstly, a guide head is adopted to measure the line-of-sight angle and the line-of-sight angular rate of an aircraft and a target, a leading angle is obtained according to a leading condition, a yaw angle is measured through a gyroscope, and a yaw angle deviation signal and a leading deviation signal are obtained through comparison. And secondly, constructing an integral compensation pre-estimation observer with system uncertainty according to the line-of-sight angular rate signal, the angle error signal and the distance error signal to obtain an uncertainty observation signal. And finally, synthesizing the compensation observation signal, the angle error signal, the position error signal and the like to form guide comprehensive information, conveying the guide comprehensive information to an attitude stabilization system, resolving a yaw rudder deflection signal, and controlling the aircraft to accurately hit a target. The method has the advantages that the method can comprehensively consider angle and position error signals, estimate the uncertainty of the guidance system and compensate to improve the guidance precision.
Description
Technical Field
The invention belongs to the field of guidance and control of unmanned aerial vehicles, and particularly relates to a pre-guidance method adopting a nonlinear proportional and integral type compensation predictor.
Background
At present, conventional guidance methods, including proportional guidance, pre-guidance, etc., often perform processing such as proportional integral limiting on an individual angle signal or distance deviation signal. The method has the advantages of simple algorithm, convenient parameter debugging, economical measurement components and high computer calculation speed in the early stage. However, with the development of computer technology, the computing power of the control system on the aircraft has already broken through the traditional limitations. Meanwhile, with the development of measurement technology, measurement components are cheaper and more miniaturized, so that the current aircraft guidance is developed towards the trend of comprehensive measurement and comprehensive guidance of multiple components.
Based on the reasons, the invention provides a comprehensive preposed guiding method formed by comprehensive angle measurement information, such as a yaw deviation angle, a preposed deviation angle, distance error information, line-of-sight angular rate information and the like. Meanwhile, a novel composite nonlinear integral type pre-estimation observer is constructed by considering the inevitable uncertainty of a guidance and control system, and the uncertainty of the system is estimated through the observed value of the observer, so that a guidance signal is compensated, and the guidance precision is improved. The final case implementation also shows that the invention has high engineering practical value.
It is to be noted that the information disclosed in the above background section is only for enhancement of understanding of the background of the present disclosure, and thus may include information that does not constitute prior art known to those of ordinary skill in the art.
Disclosure of Invention
The present disclosure is directed to a pre-pilot method using a non-linear proportional and integral compensation predictor, and further to overcome, at least to some extent, the problem that the system uncertainty cannot be estimated and compensated due to insufficient comprehensive measurement information using the conventional pre-pilot method due to the limitations and defects of the related art.
The invention provides a preposed guidance method adopting a nonlinear proportional and integral type compensation predictor, which comprises the following steps:
and step S10, adopting a seeker device to measure the line-of-sight angle, the line-of-sight angular rate and the lateral error between the aircraft and the target, and simultaneously installing an attitude gyroscope on the aircraft to measure the yaw angle signal of the aircraft.
Step S20, extracting a leading angle signal according to the yaw angle signal and a lateral displacement deviation condition; comparing the yaw angle with a preposed angle signal to obtain a preposed deviation signal; and comparing the yaw angle signal with the sight angle signal to obtain a yaw deviation signal.
And step S30, performing linear superposition according to the pre-deviation signal, the yaw angle deviation signal, the lateral displacement deviation signal, the line-of-sight angle signal and the line-of-sight angular rate signal to form a guiding primary signal.
And step S40, constructing a nonlinear proportional integral compensation pre-estimation observer according to the pre-deviation signal, the yaw angle deviation signal, the lateral displacement deviation signal, the line-of-sight angle signal and the line-of-sight angular rate signal to obtain a system uncertainty observation signal.
And step S50, superimposing the system uncertainty observation signal on the guidance primary signal to obtain a final guidance comprehensive signal, and transmitting the final guidance comprehensive signal to the aircraft yaw channel attitude stabilization system after parameter debugging, so as to solve a rudder signal and control the aircraft to accurately hit a course target.
In an exemplary embodiment of the present invention, extracting a lead angle signal according to a lateral displacement deviation condition according to the yaw angle signal, comparing the yaw angle signal with the lead angle signal to obtain a lead deviation signal, and comparing the yaw angle signal with a line of sight angle signal to obtain a yaw deviation signal includes:
e0＝ψ-ψa；
e1＝ψ-q；
wherein psi is an aircraft yaw angle signal; psiaAt a leading angle, taThe yaw angle signal value at the moment. And t isaIs a condition | ez|≤|0.5ez0Moment of first satisfaction, ezIs the aircraft and target lateral displacement deviation signal, ez0Is ezIs started. e.g. of the type0As yaw deviation signal, e1Is a pre-bias signal.
In an exemplary embodiment of the present invention, the linearly superimposing according to the pre-bias signal, the yaw angle bias signal, the lateral displacement bias signal, the line-of-sight angle signal, and the line-of-sight angular rate signal to form the pilot primary signal includes:
wherein k is1、k2、k3、k4、k5The detailed settings are described in the following examples. e.g. of the type1As yaw deviation signal, e0For a pre-offset signal, ezA lateral displacement deviation signal, a q-line-of-sight angle signal,
a line of sight angular rate signal. u. of1I.e. the pilot primary signal.
In an exemplary embodiment of the present invention, constructing a non-linear proportional integral compensation pre-estimation observer according to the pre-deviation signal, the yaw angle deviation signal, the lateral displacement deviation signal, the line-of-sight angle signal, and the line-of-sight angular rate signal, and obtaining the system uncertainty observation signal includes:
p(n+1)＝p(n)+T1pd1；
w＝e1+q+p；
where p is the state of the observer, with an initial value set to 0. p is a radical ofdUpdating variables, p, for observer statesd1For non-linear updating of variables, kd1、kd2、kd3、kd4、kd5、kd6、kd7、kd8、kd9、kd10、kd11、kd12、kd13、ε1Is a constant parameter. e.g. of the type1As yaw deviation signal, e0For a pre-offset signal, ezFor lateral displacementThe deviation signal, the q line-of-sight angle signal,
a line of sight angular rate signal. u. ofzFor the yaw path synthesis signal, where p (n +1) is the n +1 th data of the observer state p, T1The time interval between data can be selected as T10.001. w is the final sought system uncertainty observation signal.
In an exemplary embodiment of the present invention, the obtaining the final pilot combined signal by superimposing the pilot primary signal on the system uncertainty observation signal comprises:
uz＝u1-kww；
wherein u iszTo guide the combined signal u1For pilot primary signals, w is the system uncertainty observation signal, kwThe detailed settings are described in the following examples.
The invention relates to a preposed guidance method adopting a nonlinear proportional and integral type compensation predictor, which can be used for measuring the relative motion situation information of an aircraft and a target by integrating various components and carrying out information fusion processing on the relative motion situation information. The information comprises information of two aspects of position and angle, and the angle information comprises three aspects of preposition information, yaw deviation information, line-of-sight angular speed information and the like. On the other hand, the method for constructing the uncertain information integral of the system to estimate observation through the sight line angle, the lead angle deviation, the yaw angle deviation information and the sight line angular speed information is provided, so that the delicate relation between the miss distance and the measurement information is revealed. And the final optimal guidance parameters are obtained through parameter adjustment, so that higher guidance precision is ensured.
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 disclosure.
Drawings
The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the present disclosure and together with the description, serve to explain the principles of the disclosure. It is to be understood that the drawings in the following description are merely exemplary of the disclosure, and that other drawings may be derived from those drawings by one of ordinary skill in the art without the exercise of inventive faculty.
FIG. 1 is a flow chart of a pre-pilot method using a non-linear proportional and integral compensation estimator according to the present invention.
FIG. 2 is a line-of-sight angle signal (in degrees) for a method provided by an embodiment of the invention;
FIG. 3 is a plot of the line of sight angular rate signal (in degrees per second) for a method provided by an embodiment of the present invention;
FIG. 4 is a side error signal (in meters) for a method provided by an embodiment of the present invention;
FIG. 5 is a plot of yaw angle signals (in degrees) for a method provided by an embodiment of the present invention;
FIG. 6 shows a pre-offset signal (in degrees) for a method according to an embodiment of the present invention;
FIG. 7 is a yaw bias signal angle (in degrees) for a method provided by an embodiment of the present invention;
FIG. 8 is a system uncertainty observation signal (unitless) of a method provided by an embodiment of the invention;
fig. 9 is a pilot integrated signal (unitless) of a method provided by an embodiment of the invention;
FIG. 10 is a plot of yaw rudder deflection angle (in degrees) for a method provided by an embodiment of the present invention;
FIG. 11 is a graph of the side slip angle (in degrees) for a method provided by an embodiment of the present invention;
FIG. 12 is a graph of mesh distance (in meters) for a method provided by an embodiment of the invention;
FIG. 13 shows the trajectory of the aircraft and the target (in meters) according to the method 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 give a thorough understanding of embodiments of the disclosure. One skilled in the relevant art will recognize, however, that the subject matter of the present disclosure can be practiced without one or more of the specific details, or with other methods, components, devices, steps, and the like. In other instances, well-known technical solutions have not been shown or described in detail to avoid obscuring aspects of the present disclosure.
The invention discloses a preposed guidance method adopting a nonlinear proportional and integral compensation predictor, which comprises the steps of firstly adopting measuring components to measure the sight angle information, the yaw angle information, the sight angle rate and the lateral position deviation information of an aircraft, extracting a preposed angle according to a preposed angle extraction condition, and obtaining a preposed error angle and a yaw error angle. And describing a subtle relationship between the system miss distance and the measurement information by a system uncertainty integral type pre-estimation observer, and observing the system uncertainty by the pre-estimation observer so as to compensate a guidance initial signal consisting of the angle information and the position information and obtain a final high-precision guidance rule. Therefore, the preposed guidance method adopting the nonlinear proportional and integral type compensation predictor provided by the invention has the advantages of fusion and synthesis of information measured by various measurement components, and observation and prediction of system uncertainty, thereby having high engineering application value.
The following will further explain and explain a pre-pilot method using a non-linear proportional and integral type compensation estimator according to the present invention with reference to the accompanying drawings. Referring to fig. 1, the method for pre-steering using a non-linear proportional and integral type compensation estimator includes the following steps:
and step S10, adopting a seeker device to measure the line-of-sight angle, the line-of-sight angular rate and the lateral error between the aircraft and the target, and simultaneously installing an attitude gyroscope on the aircraft to measure the yaw angle signal of the aircraft.
Specifically, first, a seeker device is used to measure the projection of the line of sight angle between the aircraft and the target on a horizontal plane, denoted as q. It should be noted that since the present invention is discussed primarily with respect to course navigation, the line-of-sight angle is not specifically illustrated here and in the following, but is the projected component of the value line-of-sight signal in the horizontal plane. The guiding design of the pitch channel can be performed with reference to the yaw channel, and therefore, the description is not repeated here.
Further, the angular rate of line of sight signal between the aircraft and the target is measured by the seeker and recorded as
The line-of-sight angular rate here is also a rate signal of a projection signal of the three-dimensional space line-of-sight angular signal on a horizontal plane.
Then, the lateral displacement deviation of the aircraft and the target is provided by a seeker or an inertial navigation component or a ground measuring device, and is recorded as ez. The information can be obtained by calculating the difference of the position measurement information or indirectly by measuring the distance signal and the angle information and then converting. The method of measurement, which is not the subject matter of the present invention, will not be described in detail herein.
Finally, an attitude gyroscope is mounted on the aircraft, and a yaw angle signal of the aircraft is measured and recorded as psi. It should be noted that, when attitude stabilization is performed on an aircraft, an attitude control system of an attitude gyroscope is installed and measures an attitude angle signal, so that the attitude gyroscope does not need to be installed again. In the guidance design, a yaw angle signal measured by the attitude control system can be directly used.
Step S20, extracting a leading angle signal according to the yaw angle signal and a lateral displacement deviation condition; comparing the yaw angle with a preposed angle signal to obtain a preposed deviation signal; and comparing the yaw angle signal with the sight angle signal to obtain a yaw deviation signal.
Specifically, firstly, according to the lateral displacement deviation signal ezExtracting its initial value, and recording it as ez0。
Secondly, the leading angle extraction conditions are set as follows: | ez|≤|0.5ez0When the initial lead angle extraction condition is satisfied for the first time, the value is recorded as ta. Then record taYaw angle signals at moments denoted psiaThis is taken as the lead angle.
Finally, the yaw angle signal is compared with the lead angle signal to obtain a lead deviation signal, which is recorded as e0The calculation method is e0＝ψ-ψa；
Comparing the yaw angle signal with the line-of-sight angle signal to obtain a yaw deviation signal, and recording the yaw deviation signal as e1The calculation method is e1＝ψ-q。
And step S30, performing linear superposition according to the pre-deviation signal, the yaw angle deviation signal, the lateral displacement deviation signal, the line-of-sight angle signal and the line-of-sight angular rate signal to form a guiding primary signal.
In particular, the pilot primary signal is denoted u1The composition mode is as follows:
wherein k is1、k2、k3、k4、k5The detailed settings are described in the following examples. e.g. of the type1As yaw deviation signal, e0For a pre-offset signal, ezA lateral displacement deviation signal, a q-line-of-sight angle signal,
a line of sight angular rate signal.
And step S40, constructing a nonlinear proportional integral compensation pre-estimation observer according to the pre-deviation signal, the yaw angle deviation signal, the lateral displacement deviation signal, the line-of-sight angle signal and the line-of-sight angular rate signal to obtain a system uncertainty observation signal.
Specifically, the state of the observer is first defined as p, and its initial value is set to 0.
Next, an observer state update variable is calculated, denoted as pdIt is defined as:
wherein p isd1For non-linear updating of variables, kd1、kd2、kd3、kd4、kd5、kd6、kd7、kd8、kd9、kd10、kd11、kd12、kd13、ε1The detailed settings are described in the following examples. e.g. of the type1As yaw deviation signal, e0For a pre-offset signal, ezA lateral displacement deviation signal, a q-line-of-sight angle signal,
a line of sight angular rate signal. Where w is the system uncertainty observation signal, uzThe signal is synthesized for the yaw path, the detailed calculation of which is described in the following step.
Finally, observer state update is performed according to the following formula:
p(n+1)＝p(n)+T1pd1；
where p (n +1) is the n +1 th data of the observer state p, T1The time interval between data can be selected as T10.001. Meanwhile, resolving the system uncertainty observation signal according to the following formula:
w＝e1+q+p；
and step S50, superimposing the system uncertainty observation signal on the guidance primary signal to obtain a final guidance comprehensive signal, and transmitting the final guidance comprehensive signal to the aircraft yaw channel attitude stabilization system after parameter debugging, so as to solve a rudder signal and control the aircraft to accurately hit a course target.
Specifically, the pilot integrated signal is defined as uzLinear superposition is performed according to the following formula:
uz＝u1-kww；
wherein u is1For pilot primary signal, w is system uncertainty observation signal, kwThe detailed settings are described in the following examples.
And finally, obtaining a comprehensive guide signal, and transmitting the comprehensive guide signal to an attitude stabilization loop, so that the yaw angle of the aircraft tracks and guides the comprehensive signal, and the yaw rudder yaw angle signal is solved, and the aircraft is controlled to accurately hit a target. It should be noted that, because the design method of the attitude stabilization loop is different from aircraft to aircraft, the conventional method adopts the PID control of the attitude angle, and the implementation of the present invention also adopts the PID control to verify the correctness of the proposed method.
Case implementation and computer simulation result analysis
In order to verify the effectiveness of the method provided by the invention, case implementation simulation is firstly carried out. The aircraft model adopts a three-channel six-channel digital model and has high authenticity. The flying speed is accelerated by the driving of an engine oil supply rule at 0 meter per second.
Specifically, in step S10, the seeker device is used to measure the line-of-sight angle, the line-of-sight angular rate, and the lateral error between the aircraft and the target, the specific curves are shown in fig. 2, fig. 3, and fig. 4, respectively, and the yaw angle signal is obtained as shown in fig. 5.
In step S20, comparing the yaw angle with the lead angle signal to obtain a lead deviation signal as shown in fig. 6; the yaw deviation signal obtained by comparing the yaw angle signal with the line-of-sight angle signal is shown in fig. 7.
In step S3In 0, set k1＝-1、k2＝-0.5、k3＝0.1、k4＝0.2、k50.002. In step S40, k is selectedd1＝0.6、kd2＝0.1、kd3＝0.05、kd4＝0.001、kd5＝0.2、kd6＝0.3、kd7＝0.1、kd8＝0.1、kd9＝0.001、kd10＝1、kd11＝0.1、kd12＝0.1、kd13＝0.001、ε1A system uncertainty observation signal is obtained at 0.5 as shown in fig. 8.
In step S50, k is selectedwAnd (1) obtaining a final guidance comprehensive signal as shown in fig. 9, transmitting the final guidance comprehensive signal to an aircraft yaw channel attitude stabilization system, calculating a rudder signal as shown in fig. 10, and controlling the aircraft to accurately hit a heading target, wherein the sideslip angle of the aircraft is shown in fig. 11, the final eye distance curve of the aircraft is shown in fig. 12, and the motion situation curve of the aircraft and the target is shown in fig. 13.
As can be seen in fig. 8 and 9, the observer output signal curve is a large proportion of the final guidance complex, thus illustrating that the system uncertainty observed by the observer accounts for the compensation effect in the overall aircraft guidance. As can be seen from fig. 12 and 13, the final miss distance of the aircraft is less than 0.5 m, thus indicating that the present invention has a high hit accuracy. Therefore, the method has higher theoretical innovation value and engineering application value.
Other embodiments of the disclosure will be apparent to those skilled in the art from consideration of the specification and practice of the disclosure disclosed herein. This application is intended to cover any variations, uses, or adaptations of the disclosure following, in general, the principles of the disclosure and including such departures from the present disclosure as come within known or customary practice within the art to which the disclosure pertains. It is intended that the specification and examples be considered as exemplary only, with a true scope and spirit of the disclosure being indicated by the following claims.
Claims (2)
1. A pre-guiding method adopting a nonlinear proportional and integral type compensation predictor is characterized by comprising the following steps:
step S10, adopting a seeker device to measure the line-of-sight angle, the line-of-sight angular rate and the lateral error between the aircraft and the target, and meanwhile, installing an attitude gyroscope on the aircraft to measure the yaw angle signal of the aircraft;
step S20, extracting a leading angle signal according to the yaw angle signal and a lateral displacement deviation condition; comparing the yaw angle with a preposed angle signal to obtain a preposed deviation signal; and comparing the yaw angle signal with the line-of-sight angle signal to obtain a yaw deviation signal as follows:
e0＝ψ-ψa；
e1＝ψ-q；
wherein psi is an aircraft yaw angle signal; q is the line-of-sight angle signal, psiaIs a leading angle formed by cutting taA yaw angle signal value at a time; and t isaIs obtained by intercepting the condition | ez|≤|0.5ez0Moment of first satisfaction, ezIs the aircraft and target lateral displacement deviation signal, ez0Is ezAn initial value of (1); e.g. of the type1For a pre-offset signal, e0Is a yaw deviation signal;
step S30, according to the pre-deviation signal, the yaw angle deviation signal, the lateral displacement deviation signal, the line-of-sight angle signal and the line-of-sight angular rate signal, performing linear superposition to form a guiding primary signal as follows:
wherein k is1、k2、k3、k4、k5Is a constant parameter; e.g. of the type1As yaw deviation signal, e0For a pre-offset signal, ezIs a lateral displacement deviation signal, q is a line-of-sight angle signal,
a line-of-sight angular rate signal; u. of1I.e. the pilot primary signal;
step S40, constructing a nonlinear proportional integral type compensation pre-estimation observer according to the pre-deviation signal, the yaw angle deviation signal, the lateral displacement deviation signal, the line-of-sight angle signal and the line-of-sight angular rate signal to obtain a system uncertainty observation signal;
p(n+1)＝p(n)+T1pd1；
w＝e1+q+p；
where p is the state of the observer, with an initial value set to 0; p is a radical ofdUpdating variables, p, for observer statesd1For non-linear updating of variables, kd1、kd2、kd3、kd4、kd5、kd6、kd7、kd8、kd9、kd10、kd11、kd12、kd13、ε1Is a constant parameter; e.g. of the type1As yaw deviation signal, e0For a pre-offset signal, ezA lateral displacement deviation signal, a q-line-of-sight angle signal,
a line-of-sight angular rate signal; u. ofzFor the yaw path synthesis signal, where p (n +1) is the n +1 th data of the observer state p, T1The time interval between data can be selected as T10.001; w is the finally solved system uncertainty observation signal;
and step S50, superimposing the system uncertainty observation signal on the guidance primary signal to obtain a final guidance comprehensive signal, and transmitting the final guidance comprehensive signal to the aircraft yaw channel attitude stabilization system after parameter debugging, so as to solve a rudder signal and control the aircraft to accurately hit a course target.
2. The method of claim 1, wherein the step of superimposing the system uncertainty observation signal on the pilot primary signal to obtain the final pilot synthesis signal comprises:
uz＝u1-kww；
wherein u iszTo guide the combined signal u1For pilot primary signals, w is the system uncertainty observation signal, kwIs a constant parameter.
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