CN110764534B - Nonlinear conversion-based method for guiding preposed guidance and attitude stabilization matching - Google Patents
Nonlinear conversion-based method for guiding preposed guidance and attitude stabilization matching Download PDFInfo
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Abstract
The invention relates to a nonlinear conversion-based method for guiding in a preposed guidance and attitude stabilization matching mode, which belongs to the technical field of aircraft guidance and comprises the following steps: measuring the line-of-sight angular rate of the aircraft moving relative to the target, and measuring the yaw angle and the yaw angular rate of the aircraft; constructing a line-of-sight angle signal according to the line-of-sight angle rate, and constructing an azimuth error signal and a lead angle error signal according to the line-of-sight angle signal and the yaw angle; constructing an amplitude limiting nonlinear signal according to the azimuth error signal, and constructing a pre-pilot law according to the amplitude limiting nonlinear signal and the pre-angle error signal; and constructing an output signal with stable posture according to a preposed guidance law, and obtaining an expected course angle of the aircraft according to the output signal with stable posture, so that the yaw angle can stably track the expected course angle. The method solves the problems that the output of the guidance law is rapidly increased and the miss distance of the missile is overlarge.
Description
Technical Field
The invention relates to the technical field of aircraft guidance, in particular to a nonlinear conversion-based method for guiding prepositive guidance and attitude stabilization matching.
Background
The precise guidance algorithm of the aircraft has higher application value, and can be applied to automatic landing of the unmanned aircraft, precise navigation and route planning of the unmanned aircraft, precise guidance striking of the unmanned aircraft and the like, rendezvous and docking of aerospace aircraft and the like. And has therefore been a significant problem for many years of research by engineers in the field of aircraft design. In particular to the high precision requirement of the butt joint of the aerospace vehicle and the precision guidance requirement of a weapon system, and the success or failure of task execution is directly related to the precision of guidance and guidance. Therefore, precision is the core in the research of the guiding algorithm problem.
The traditional proportional guidance has the main problems that the initial section is not sensitive to the target motion, and the guidance law output is too large due to the fact that the line-of-sight angular rate is changed too fast at the final section, so that the tracking pressure of a missile control system is too large, and a large miss distance exists. Other conventional guidance methods, such as pre-guidance, also inevitably suffer from the above-mentioned problem that the control distribution of the aircraft is not evenly distributed throughout the guidance process, resulting in an excessive amount of end-off-target.
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 invention and therefore may include information that does not constitute prior art known to a person of ordinary skill in the art.
Disclosure of Invention
The invention aims to provide a method for guiding preposed guidance and attitude stabilization matching based on nonlinear conversion, and further solves the problem that the missile miss distance is too large due to the rapid increase of guidance law output caused by the limitations and defects of the related technology at least to a certain extent.
The invention provides a nonlinear conversion-based method for guiding prepositive guidance and attitude stabilization matching, which comprises the following steps of:
measuring the line-of-sight angular rate of the aircraft moving relative to the target, and measuring the yaw angle and the yaw angular rate of the aircraft;
constructing a line-of-sight angle signal according to the line-of-sight angle rate, and constructing an azimuth error signal and a lead angle error signal according to the line-of-sight angle signal and the yaw angle;
constructing an amplitude limiting nonlinear signal according to the azimuth error signal, and constructing a pre-pilot law according to the amplitude limiting nonlinear signal and the pre-angle error signal;
and constructing an output signal with stable posture according to the preposed guidance law, and obtaining the expected course angle of the aircraft according to the output signal with stable posture, so that the yaw angle can stably track the expected course angle.
In an example embodiment of the present invention, constructing a line-of-sight angular signal from the line-of-sight angular rate comprises:
In an example embodiment of the present invention, constructing the azimuth error signal and the lead angle error signal from the line of sight angle signal and the yaw angle comprises:
e1＝u1-ψc；
e2＝u1-ψc0；
wherein e is1Is an azimuth error signal; u. of1Is a line-of-sight angle signal; psicA yaw angle of the aircraft; e.g. of the type2Is a lead angle error signal; psic0For the aircraft to fly to t0Yaw angle at a moment.
In an example embodiment of the present invention, constructing a clipped nonlinear signal from the azimuth error signal comprises:
wherein u is2Is a clipping non-linear signal; f. of1For non-linear transformation, k1、k2、ε1And epsilon2Is a constant parameter; and, | u2|＜|k1|+|k2|。
In an example embodiment of the present invention, constructing a preamble law from the clipped nonlinear signal and the preamble angle error signal comprises:
ua＝k3e2+u2(ii) a Wherein u isaIs the leading rule; k is a radical of3Is a constant parameter.
In an example embodiment of the present invention, constructing an attitude-stabilized output signal according to the pre-steering law comprises:
u＝k4∫u3dt; wherein, the u posture is stable output signals; k4 is a constant parameter; integral sign dt; u. of3For the pre-pilot law uaAnd performing the processed guidance law.
The nonlinear conversion-based method for the pre-guidance and attitude stabilization matching guidance enables a guidance law initial stage to have larger output signals, and the output signals at the tail end are not increased remarkably. Therefore, the problem that the guidance law is insensitive to the movement of the target in the initial stage and the guidance law output is rapidly increased and the miss distance of the missile is too large due to the extremely rapid increase of the line angle far away from the target in the final stage is solved. Meanwhile, the nonlinear guidance law has the advantages that the output is relatively balanced, and the high-precision target hitting is realized by the back-and-forth fluctuation. Therefore, the nonlinear conversion pre-guidance method provided by the invention has the advantages of novel method and high precision, and has very high engineering application value. Meanwhile, the method can also be applied to the navigation of unmanned aerial vehicles such as unmanned aerial vehicles, and has higher economic 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.
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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 a method for guiding preposition guidance and attitude stabilization matching based on nonlinear transformation provided by the invention.
FIG. 2 is a graph of the relative motion of the aircraft and the target in a course plane (in meters) according to the method of the present invention.
FIG. 3 is a miss-measure curve (in meters) for a method provided by an embodiment of the invention.
FIG. 4 is a plot of off-target magnification (in meters) for a method provided by an embodiment of the invention.
FIG. 5 is a graph of actual yaw angle versus desired yaw angle (in degrees) for a method provided by an embodiment of the present invention.
FIG. 6 shows the output (in degrees/per second) of the new pre-pilot law in accordance with the method of the present invention.
Fig. 7 is a lead angle (unit: degree) of a method provided by an embodiment of the present invention.
Fig. 8 shows a non-linear pilot signal (unit: degree/second) according to 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 discloses a method for guiding prepositive guidance and stable attitude matching based on nonlinear conversion, which can measure the yaw angle of an aircraft according to an inertial navigation combination device or a gyroscope according to the line-of-sight angular rate measured by an aircraft guide head, then carry out nonlinear conversion according to the error between the line-of-sight angle and the yaw angle, give out a nonlinear guidance signal, superimpose the error signal between the attitude angle and the prepositive angle, finally carry out integration, and send the integrated signal to an attitude stabilizing loop as a driving signal to drive the aircraft to strike a target with high precision.
The following describes a method for guiding a pre-navigation system and a pose stabilization system based on nonlinear transformation according to an embodiment of the present invention. Referring to fig. 1, the nonlinear transformation based method for guiding the preposition and the pose stabilization matching can comprise the following steps:
and step S10, measuring the line-of-sight angular rate of the aircraft relative to the target movement, and measuring the yaw angle and the yaw angular rate of the aircraft.
First, the line-of-sight angular rate of movement of the aircraft relative to the target is measured using a seeker and recorded as
Which mathematically represents the viewing angle qεOf (a) wherein q isεIs defined as:
Δx＝xT-x
Δy＝yT-y；
Δz＝zT-z
wherein x, y and z are position coordinates of the aircraft on a space three-dimensional coordinate system; x is the number ofT、yT、zTIs aimed atPosition coordinates on a three-dimensional coordinate system.
Secondly, an attitude angle signal of the aircraft is measured by adopting an inertial navigation combination, and a course channel and a course plane guide are taken as research objects below. Assuming that the yaw angle of the aircraft course channel is measured by adopting an inertial navigation combination component and is recorded as psic(ii) a Simultaneously measuring the yaw rate of the aircraft as omegay(ii) a An attitude angle measuring gyroscope and a rate gyroscope can also be arranged on the aircraft to respectively measure the yaw angle psicWith yaw rate omegay。
And step S20, constructing a line-of-sight angle signal according to the line-of-sight angle rate of the relative target motion, and constructing an azimuth error signal and a lead angle error signal according to the line-of-sight angle signal and the yaw angle.
Specifically, first, according to the line-of-sight angular rate
Constructional line-of-sight angle signal u1Specifically, the following may be mentioned:
integral dt represents the sign of the integral.
It should be noted that u is now1The signal is integrated by the measuring signal, and the measuring signal inevitably carries delay and noise, so u1With ideal line-of-sight angle signal qεIn contrast, delay and noise are mainly involved.
Then, based on the line-of-sight angle signal u1With yaw angle psicComparing to obtain an azimuth error signal e1It is defined as follows:
e1＝u1-ψc。
further, selecting an aircraft to fly to a certain time t0A handle t0Yaw angle at a moment is denoted by psic0。
Defining a pre-error signal e2The following were used:
e2＝u1-ψc0。
and step S30, constructing a limiting nonlinear signal according to the azimuth error signal, and constructing a pre-pilot law according to the limiting nonlinear signal and the pre-angle error signal.
Specifically, first, at e1Based on the following nonlinear transformation f1Constructing a limiting non-linear signal u of the line of sight angle2The method comprises the following steps:
wherein k is1、k2、ε1、ε2The specific selection of the constant parameter is described in the following examples. Further, since | u2|＜|k1|+|k2Therefore, the device can play a role in amplitude limiting, and avoids the phenomenon that the control quantity is larger to cause the aircraft to turn sharply in the guiding process, so that the end miss distance is increased. It should be noted that the nonlinear transformation can avoid the problem of the increase of the miss distance due to the over-size of the ratio pilot signal at the end. By selecting epsilon at the same time1In addition, in the initial stage of guidance, although the deviation between the azimuth angle and the attitude angle is small, the input of the control quantity is still large, so that the problems that other guidance laws are insensitive in the initial stage and turn too slowly can be avoided.
Then, a novel preposed guidance law u provided by the invention is constructedaSpecifically, the following may be mentioned:
ua＝k3e2+u2(ii) a Wherein k is3Is a constant parameter.
Further, considering that the seeker measures the blind area of the component, the guidance law can be improved and processed as follows:
wherein u isa0Is R ═ daTime uaA value of (d); r is the target and aircraftAnd da is the target distance.
And step S40, constructing an output signal with stable posture according to the preposed guidance law, and obtaining the expected course angle of the aircraft according to the output signal with stable posture, so that the yaw angle can stably track the expected course angle.
Specifically, first, the pre-steering law constructs an output signal u with a stable posture, which may be specifically as follows:
u＝k4∫u3dt; wherein k is4Being a constant parameter, ^ dt represents the sign of the integral.
Then, the output signal u of the pre-pilot signal is used as the input signal of the attitude stabilization loop of the aircraft, namely the expected heading angle of the aircraft
So that the yaw angle psi of the aircraftcCan stably track expected course angle
I.e. the output u of the new type of pre-pilot law, can be set
Further, the following describes a design process of an attitude stabilization loop tracking law and a process of aircraft attitude stabilization tracking flight by taking a conventional attitude stabilization loop as an example.
dyc＝ka1e+ka2∫edt+ka3wy；
Wherein, deltaycThe control quantity of the aircraft yaw channel is a yaw rudder deflection command signal; e is an error signal defined as:
Finally, the invention relates to target simulation and parameter adjustment.
Specifically, first, the amount of off-target is defined as a simple form as follows:
when Δ x ═ xTAnd when x is less than 0, the distance R between the aircraft and the target is the miss distance.
And selecting target motion parameters to be attacked, simulating a scene in which maneuvering targets with different speeds and different initial position conditions meet the tail section of the aircraft, and designing the parameters of the preposed guidance rule. And judging the miss distance under different guidance law parameters through the motion simulation of various targets, so as to select the final guidance law parameters, complete the guidance control parameter design of the aircraft for the front guidance and stable attitude tracking matching, and realize the accurate guidance of the targets.
Case implementation and computer simulation result analysis
Firstly, selecting the moment when the aircraft flies to 3km away from a target, and recording the attitude angle of the aircraft at the moment as a leading angle.
Next, select k1＝1.6、ε1＝0.5、k3＝0.32、ε2＝0.2、k3＝0.3、da＝40、k4＝16、ka1＝1.8、ka20.6 and ka3＝0.4。
Further, taking a certain type of moving object as an example, the process of case implementation is described. Assume an initial target position of xT(0)＝7000、yT(0)＝1、zT(0) At 400, the target moves at a constant speed, the speed is 10m/s, and the direction of the target forms an included angle of 30 degrees with the positive direction of the x axis. Other initial positions and target speed,In the case of the speed direction, the principle of parameter selection is the same, and is not described here. Simulation curves implemented according to the above-described case are shown in fig. 2 to 8.
As can be seen from fig. 2, the aircraft and the target can approach each other in the heading plane; as can be seen from fig. 3, the relative distance between the aircraft and the target is gradually reduced; as can be seen from fig. 4, the final miss distance is 0.5 m, which is fully satisfactory for small targets with a size greater than 1 m, such as vehicle tanks and the like. As can be seen from fig. 5, the actual yaw angle and the desired yaw angle are consistent in shape, where the blue curve is the desired yaw angle curve and the black curve is the actual yaw angle curve of the aircraft. The back-and-forth oscillation mode of the guidance output realizes the final high-precision guidance target hitting, so that the miss distance is smaller than 1 m. As can be seen from FIG. 6, the output value of the guidance law is larger in the initial segment and smaller in the final segment, the miss distance of the novel guidance law is small due to the fact that the end command is smaller, and the miss distance is large due to the fact that the end line-of-sight angle of the traditional proportional guidance law is larger and the command is larger. Therefore, the novel preposed guidance law has the advantage of high precision. Fig. 7 shows the lead angle signal, which is reasonable as shown in the figure, with the end stabilized around 4 degrees. Fig. 8 shows a nonlinear conversion signal, and it can be seen from the graph that although there are more oscillating switches, the output amplitude does not change significantly, so that the energy of the signal is relatively uniform in the whole guiding process, and the disadvantages of insensitive initial section and sharp turning of final section in the conventional proportional guidance are avoided. This is also the reason why the novel lead laws provided by the invention have a small miss distance.
On the basis, the transformation of different target motion speeds or the change of the initial position of the target is considered, the guide law parameters are finely adjusted, and the parameters of the whole set of novel preposed guide law are finally determined, so that the matching design of novel preposed guide and attitude tracking is completed.
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 (1)
1. A method for guiding preposition guidance and attitude stabilization matching based on nonlinear transformation is characterized by comprising the following steps:
measuring the line-of-sight angular rate of the aircraft moving relative to the target, and measuring the yaw angle and the yaw angular rate of the aircraft;
constructing a line-of-sight angle signal according to the line-of-sight angle rate, and constructing an azimuth error signal and a lead angle error signal according to the line-of-sight angle signal and the yaw angle;
constructing an amplitude limiting nonlinear signal according to the azimuth error signal, and constructing a pre-pilot law according to the amplitude limiting nonlinear signal and the pre-angle error signal;
constructing an output signal with stable posture according to the pre-guidance law, and obtaining an expected course angle of the aircraft according to the output signal with stable posture, so that the yaw angle can stably track the expected course angle;
wherein constructing a line-of-sight angle signal from the line-of-sight angular rate comprises:
constructing an azimuth error signal and a lead angle error signal from the line of sight angle signal and the yaw angle comprises:
e1＝u1-ψc；
e2＝u1-ψc0；
wherein e is1Is an azimuth error signal; u. of1Is a line-of-sight angle signal; psicA yaw angle of the aircraft; e.g. of the type2Is a lead angle error signal; psic0For the aircraft to fly to t0Yaw angle at a moment;
constructing a clipped nonlinear signal from the azimuth error signal comprises:
wherein u is2Is a clipping non-linear signal; f. of1For non-linear transformation, k1、k2、e1And e2Is a constant parameter; and, | u2|<|k1|+|k2|；
Constructing a pre-steering law according to the amplitude limiting nonlinear signal and the pre-angle error signal comprises:
ua＝k3e2+u2(ii) a Wherein u isaIs the leading rule; k is a radical of3Is a constant parameter;
constructing an output signal with stable posture according to the pre-pilot law comprises the following steps:
u＝k4∫u3dt; wherein, the u posture is stable output signals; k is a radical of4Is a constant parameter; integral sign dt; u. of3For the pre-pilot law uaAnd performing the processed guidance law.
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| CN111766776B (en) * | 2020-07-06 | 2022-02-15 | 中国人民解放军海军航空大学 | Pre-guiding method adopting nonlinear proportional and integral type compensation predictor |
| CN113359856B (en) * | 2021-07-14 | 2022-11-18 | 中国人民解放军海军航空大学 | Unmanned aerial vehicle designated course target point guiding method and system |
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