CN110687931A - Integrated maneuvering guiding method for switching azimuth attitude and preposed guidance - Google Patents
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Abstract
The invention mainly measures the relative distance, the line-of-sight angle and the line-of-sight angular rate of the aircraft and a target through a seeker, measures the yaw angle and the yaw angular rate of the aircraft by adopting an attitude and rate gyroscope, and then switches the guidance method according to the deviation between the yaw angle and the line-of-sight angle of the aircraft. In the initial stage, when the deviation is large, a nonlinear attitude and azimuth guidance method which takes a line-of-sight nonlinear signal as a main part is adopted; in the end section, when the deviation is small, a nonlinear preposed guiding method is adopted for guiding; when the guidance is interrupted, the guidance is realized through multiple times of switching, and meanwhile, the maneuvering is realized. The nonlinear preposed guidance and nonlinear azimuth attitude guidance method is greatly different from the traditional preposed guidance and attitude guidance method, and the combination and switching of the nonlinear preposed guidance and the nonlinear azimuth attitude guidance method not only keeps the high-precision advantage of the preposed guidance, but also enables the guidance law to have high efficiency initially, and can generate maneuver through switching to realize seamless connection of the maneuver and the guidance.
Description
Technical Field
The invention relates to the field of aircraft guidance, in particular to a composite guidance method for azimuth attitude guidance and prepositive proportional guidance switching.
Background
With the development of science and technology, the high-precision guidance technology has important value in military use and has more and more applications in the civil field. Such as automatic navigation of unmanned aircraft, automatic landing of unmanned aircraft, automated lane keeping by unmanned vehicles, and so forth. In the process, automatic guidance and guidance technologies can be adopted, particularly, automatic butt joint of aerospace vehicles, rail intersection and the like are adopted, and the automatic guidance technology is also widely adopted.
The conventional guidance methods include tracking, parallel approach, three-point, proportional guidance, and pre-guidance. The proportional pilot method and the pre-pilot method are two types which are applied more at present. Of course, with the development of computers, many new guidance methods for implementing complex algorithms through computer programming are currently available, but the core of the guidance methods is often refinement or extension of a certain traditional method. Although the proportional guidance method has the advantage of simple algorithm, the control quantity of the tail section is large, and the problem that the tail section turns too fast is easy to occur. The leading guidance algorithm leads the aircraft to start flying in the leading angle direction in advance after the leading angle is superposed, and solves the problem of excessively violent turning at the tail section to a certain extent. Therefore, in the conventional design, the problem of low initial segment guiding efficiency exists. Meanwhile, sometimes the aircraft has maneuverability requirement at the tail section, and the requirement and the guiding requirement have certain contradiction. Based on the reasons, the invention provides a mode of switching between azimuth attitude guidance and forward angle guidance, which not only solves the problem of low guidance effect in the initial guidance stage, but also realizes the requirements of guidance and maneuvering to a certain extent.
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 integrated maneuvering guiding method for switching azimuth attitude and preposed guidance, and further solves the problem that the guiding efficiency of an aircraft at the initial guiding stage is not high due to the limitations and defects of the related art at least to a certain extent.
According to one aspect of the invention, an integrated maneuvering guiding method for switching azimuth attitude and preposition guiding is provided, which comprises the following steps:
step S10, mounting a seeker device on the aircraft, and measuring the distance and the sight angle between the seeker device and the target; installing an angle measurement gyroscope on the aircraft, and measuring the yaw angle of the aircraft; and installing a rate gyroscope and measuring the yaw rate of the aircraft.
Step S20, according to the deviation between the yaw angle and the sight angle and the size of the switching threshold value, the judgment is carried out, and the nonlinear azimuth attitude guidance and the nonlinear preposed guidance are switched
And step S30, performing nonlinear transformation on the line-of-sight angle to obtain a nonlinear transformation signal, performing nonlinear integration on the line-of-sight angle to obtain a nonlinear integration signal, performing nonlinear transformation on the line-of-sight angular rate to obtain a nonlinear differential signal of the line-of-sight angle, and finally superposing the three types of nonlinear signals to calculate a nonlinear azimuth attitude guide signal.
And step S40, solving an indirect nonlinear azimuth deviation signal according to the measured sight angle and the measured yaw angle, and recording the yaw angle of the aircraft at the moment as a leading angle when the indirect nonlinear deviation signal is close to a certain small value for the first time. And solving a nonlinear error signal of the yaw angle relative to the lead angle according to the yaw angle information and the lead angle information. And calculating a nonlinear azimuth deviation signal according to the azimuth deviation amount of the line-of-sight angle and the yaw angle. And finally, integrating the upper indirect nonlinear azimuth deviation signal, the nonlinear error signal of the yaw angle relative to the front angle and the nonlinear azimuth deviation signal of the line-of-sight angle and the yaw angle to obtain an error integrated signal, then obtaining a nonlinear integral signal of the error integrated signal, and finally superposing the error integrated signal and the nonlinear integral signal of the error integrated signal to obtain a nonlinear front guidance signal.
And step S50, after switching between the nonlinear azimuth attitude and the nonlinear preposed guidance according to the switching judgment principle, inputting the final input signal of the aircraft attitude stabilization control system to the aircraft attitude control system, so that the yaw angle of the aircraft, namely the yaw angle tracks the input signal, and finally transmitting the finally formed attitude stabilization control signal to an executing mechanism of the course control of the aircraft, such as a course steering engine, so that the target of the guidance law can be realized and the aircraft is guided to fly to the preset target.
In an exemplary embodiment of the present invention, the determining of the switching between the nonlinear azimuth attitude guidance and the nonlinear front guidance according to the line-of-sight angle and the yaw angle includes:
e1＝qh-ψ
wherein psirFor the guiding law of final use, #1Pilot signal, psi, for non-linear azimuthal guidance2Is a pilot signal for nonlinear pilot. e.g. of the type1Deviation of the line of sight angle from the yaw angle, e1＝qh-ψ，qhIs the line of sight angle and psi is the yaw angle. Epsilon is a switching judgment threshold value. d is the distance between the aircraft and the target. | e1If | ≧ epsilon, performing nonlinear azimuth attitude guidance; if | e1If | < ε, then perform nonlinear pre-steering.
In an exemplary embodiment of the present invention, the nonlinear transformation according to the line-of-sight angle to obtain the nonlinear transformation signal comprises:
wherein q ishFor line-of-sight angle signals, q1For the non-linear transformation of the signal at line-of-sight angle, k1、λ1Is a positive parameter that can be adjusted.
In an exemplary embodiment of the present invention, the non-linearly integrating from the line-of-sight angle signal to obtain a non-linearly integrated signal comprises:
wherein q issIs a line-of-sight non-linear integral signal.
In an exemplary embodiment of the present invention, the obtaining of the nonlinear differential signal of the line of sight angle by performing the nonlinear transformation on the line of sight angle differential signal includes:
wherein q isdFor line-of-sight non-linear differential signals, λ2Is a positive parameter that can be adjusted.
In an exemplary embodiment of the present invention, the obtaining the nonlinear azimuth attitude guidance signal by superimposing three types of nonlinear signals, namely, a line-of-sight nonlinear integral signal, and a line-of-sight differential signal, includes:
ψ1＝k2qh+k3q1+k4qs+k5qd
wherein psi1Steering the signals for non-linear azimuthal attitude, where k2、k3、k4And k is5Are parameters that can be adjusted.
In an exemplary embodiment of the present invention, resolving the non-linear error signal thereof based on the yaw and lead angle information comprises:
wherein e2fIs a non-linear error signal of yaw angle relative to lead angle, e2＝ψ-ψ0Phi is the yaw angle, also phi0Is a leading angle, k6、k7、k8And λ3Are parameters that can be adjusted.
In one example embodiment of the present invention, resolving the non-linear bearing deviation signal based on the yaw and line of sight angle information comprises:
wherein e1fNon-linear azimuth deviation signal being line-of-sight angle and yaw angle, e1＝qh-ψ，qhIs the line of sight angle, psi is the yaw angle, k9、k10、k11And λ4Are parameters that can be adjusted.
In an exemplary embodiment of the present invention, the obtaining the error integrated signal according to the integration of the indirect nonlinear azimuth deviation signal, the nonlinear error signal of the yaw angle relative to the lead angle, and the nonlinear azimuth deviation signal of the line of sight angle and the yaw angle, then obtaining the nonlinear integral signal of the error integrated signal, and performing superposition to obtain the nonlinear lead signal includes:
ez＝k12e3+k13e1f+k14e2f
ψ2＝k15ez+k16esz
wherein psi2For non-linear pre-pilot signals, ezFor error integration signal, eszIs a mistakeNon-linear integral signal of the difference synthesis signal, e3Is an indirect non-linear azimuth deviation signal, which is calculated as e3＝sin(qh)-sin(ψ)。e2fIs a non-linear error signal of yaw angle relative to lead angle, e1fIs a non-linear azimuth deviation signal of the line of sight angle and the yaw angle. Wherein k is12、k13、k14、k15、k16And λ5Are parameters that can be adjusted.
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 an integrated maneuver guidance method for switching azimuth attitude and pre-guidance provided by the present invention;
FIG. 2 is a relative motion trajectory of an aircraft and a target according to a method provided by an embodiment of the invention;
FIG. 3 is a graph illustrating the variation of the relative distance between an aircraft and a target according to the method of the present invention;
FIG. 4 is an enlarged graph of the variation of the relative distance between the aircraft and the target according to the method of the embodiment of the invention
FIG. 5 is a line-of-sight curve for an aircraft in accordance with a method provided by an embodiment of the invention;
FIG. 6 is a plot of aircraft yaw angle for a method provided by an embodiment of the present invention;
FIG. 7 is a plot of aircraft line-of-sight versus yaw angle using a method provided by an embodiment of the present invention;
fig. 8 is a signal curve of a switching timing in a method according to an embodiment of the present invention;
fig. 9 is a pilot signal curve of a method according to an embodiment of the present invention;
FIG. 10 is a plot of aircraft yaw error for a method provided by an embodiment of the present invention;
fig. 11 is an aircraft rudder deflection angle curve for a method provided by an embodiment of the 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 integrated design for maneuvering and guiding an aircraft by measuring the relative position and angle information of the aircraft and a target by adopting a seeker and a gyroscope device and switching between nonlinear azimuth attitude guiding and nonlinear preposed guiding. On one hand, compared with the traditional method, the method has higher efficiency and more definite directivity in the initial stage of guidance, and meanwhile, local maneuver in the guidance process is realized through switching, so that the method has high engineering application value.
The integrated maneuver guiding method for switching azimuth attitude and pre-guidance according to the present invention will be further explained and explained with reference to the drawings. Referring to fig. 1, the integrated maneuver guiding method for switching azimuth attitude and lead guidance may include the following steps:
step S10, measuring equipment installation and angle and distance measurement
Firstly, a seeker device is installed on an aircraft, the distance between the aircraft and a target is measured in real time and recorded as d, and the sight line angle between the aircraft and the target in a course plane is measured and recorded as qhMeasuring the angular rate of the line-of-sight angle between the aircraft and the target in the course plane, and recording as
。
Secondly, an angle measurement gyroscope is arranged on the aircraft, and the yaw angle of the aircraft is measured and recorded as psi.
Finally, a rate gyroscope is installed on the aircraft, and the yaw rate omega of the aircraft is measuredy。
Step S20, determining and switching between orientation and attitude guidance and pre-guidance
By the measurement in step S10, the amount of azimuth deviation is first determined, and denoted as e1The calculation method is e1＝qh-ψ。
Next, a switching threshold is set, denoted as epsilon. The switch epsilon can be set in terms of distance, varying as the distance d of the aircraft from the target varies. It is set up as follows:
finally, the guidance switching judgment is carried out, and the judgment logic is as follows:
if | e1If | ≧ epsilon, performing nonlinear azimuth attitude guidance; if | e1If | < ε, then perform nonlinear pre-steering. The pilot signal for assuming non-linear azimuth attitude steering is denoted as psi1Pilot signal phi of nonlinear preamble pilot2And the input signal of the attitude stabilization control system of the aircraft is denoted by psirThen there is
The specific design methods of the non-linear azimuth-attitude guidance and the non-linear front guidance are described below.
Step S30, implementation of non-linear orientation attitude guidance
First, the viewing angle q obtained by the measurement of step S10hPerforming nonlinear transformation to calculate q after nonlinear transformation1：
Wherein k is1、λ1The specific design of the adjustable positive parameter adjusts the appropriate value selection according to the actual situation of the case.
Next, calculate qhIs a non-linear integral signal qsThe following were used:
again, the angular rate according to the line of sight angle measured in step S10
Calculating qhIs a non-linear differential signal qdThe following were used:
wherein λ2The specific design of the adjustable positive parameter adjusts the appropriate value selection according to the actual situation of the case.
Finally, the three types of signals are superposed to obtain the following azimuth attitude guidance signal which is written as psi1The superposition mode is as follows:
ψ1＝k2qh+k3q1+k4qs+k5qd
wherein k is2、k3、k4And k is5For adjustable parametersThe specific design adjusts the selection of the proper value according to the actual situation of the case.
Step S40, implementation of nonlinear pre-steering
First, the viewing angle q obtained by the measurement of step S10hSolving the non-linear azimuth deviation amount with the yaw angle psi, and recording the value as e3The calculation method is e3＝sin(qh) Sin (ψ). When | e is satisfied for the first time3When | < 0.1, recording the time as t1And recording the yaw angle psi of the aircraft at that moment0And sets the angle as a lead angle.
Secondly, the error of the yaw angle relative to the lead angle is solved and recorded as e2The calculation method is e2＝ψ-ψ0. On the basis, solving the nonlinear error quantity, and recording the quantity as e2fThe calculation method is as follows:
wherein k is6、k7、k8And λ3The specific design of the adjustable parameters is selected according to the actual situation of the case by adjusting the appropriate values.
Next, the azimuth deviation amount e obtained in step S201The calculation method is e1＝qhPsi, solving for its non-linear azimuthal deviation, denoted as e1fThe calculation method is as follows:
wherein k is9、k10、k11And λ4The specific design of the adjustable parameters is selected according to the actual situation of the case by adjusting the appropriate values.
Then, the error signals are combined to obtain an error combined signal, which is recorded as ezThe calculation method is as follows: e.g. of the typez＝k12e3+k13e1f+k14e2fThen find itNon-linear integral signal, denoted as eszThe calculation method is as follows:
wherein k is12、k13、k14And λ5The specific design of the adjustable parameters is selected according to the actual situation of the case by adjusting the appropriate values.
Finally, the final nonlinear preamble pilot signal psi is obtained2The calculation method is as follows:
ψ2＝k15ez+k16esz
wherein k is15And k is16The specific design of the adjustable parameters is selected according to the actual situation of the case by adjusting the appropriate values.
Step S50, implementation of guidance law
According to the step S20, after the nonlinear azimuth attitude and the nonlinear preposed guidance are switched, the final input signal psi of the aircraft attitude stabilization control system is obtainedrInput to the aircraft attitude control system so that the aircraft yaw angle, i.e. yaw angle psi, tracks the input signal psirThe guiding law can be realized. Since the design methods of the attitude control system are many and the invention is not focused on the discussion and protection, the following process is exemplified by the generally widely adopted PID control.
First of all the aircraft yaw angle psi, measured in step S10, is compared with the input signal psirComparing, finding the error signal of yaw angle, and recording as eψThe comparison method is as follows: e.g. of the typeψ＝ψr-ψ。
Secondly, the integral of the yaw angle error signal is solved and recorded as es＝∫eψdt。
Finally the yaw rate ω measured by the superposition step S10yForming a final attitude stabilization control signal, denoted as deltayThe calculation method is as follows:
δy＝kpeψ+kies+kdωy
wherein k isp、kiAnd k isdThe specific design of the adjustable parameters of the attitude stabilization control loop is selected according to the actual situation of the case by adjusting the appropriate values.
Finally, the finally formed attitude stabilization control signal delta is usedyAnd the target of a guidance law can be realized by an actuating mechanism of course control, such as a course steering engine, which is transmitted to the aircraft, and the aircraft is guided to fly to a preset target.
Case implementation and computer simulation result analysis
In step S10, the longitudinal distance between the target position and the guidance starting point is set to 9700 m and the lateral distance is set to-550 m, the moving speed of the target is set to 22 m/S, and the moving direction is set to-22 degrees, as shown in fig. 2 below. In step S20, ∈ 4 is set. K is set in step S301＝8、λ1＝2、λ2＝2，k2＝2、k3＝1、k40.2 and k51.5. K is set in step S406＝3、k7＝3、k80.5 and λ3＝1，k9＝2、k10＝2、k111 and λ4＝1，k12＝1.2、k13＝0.2、k14＝0.2、k17＝3、k18＝1.5、k191.5 and λ5＝2、k151 and k163. K is selected in step S50p＝2、ki1.3 and kd0.8. The resulting relative motion curve of the aircraft and the target, as shown in FIG. 2 below, shows that the aircraft is ultimately able to hit the target.
The relative distance curve of the aircraft from the target is shown in fig. 3, and the enlarged view is shown in fig. 4. It can be seen that at 43.0758 seconds, the relative distance between the aircraft and the target is 0.68 m, namely the final miss distance is less than 0.7 m, and the method provided by the invention has higher precision.
While the aircraft's line of sight angle is shown below in fig. 5 and the yaw angle is shown below in fig. 6, in the same figure as in fig. 7, it can be seen that the yaw angle can be varied around the line of sight direction to achieve gross tracking. The switching curve of the nonlinear azimuth attitude guidance and the nonlinear front guidance is shown in fig. 8, and it can be seen that multiple times of switching are realized in the early stage of guidance, so that certain maneuverability is realized in guidance. The resulting pilot signal curve is shown in fig. 9 as output to the attitude stabilization system as an input, and the tracking error of the attitude stabilization system is shown in fig. 10. It can be seen that the desired input signal for yaw tracking is ultimately achieved most of the time on the lead. The rudder deflection angle curve of the aircraft is shown in fig. 11, and it can be seen that the rudder deflection angle of the aircraft does not exceed the available range during the guidance process.
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 (4)
1. An integrated maneuvering guiding method for switching azimuth attitude and preposition guiding is characterized by comprising the following steps:
step S10, mounting a seeker device on the aircraft, and measuring the distance and the sight angle between the seeker device and the target; installing an angle measurement gyroscope on the aircraft, and measuring the yaw angle of the aircraft; and installing a rate gyroscope and measuring the yaw rate of the aircraft.
Step S20, according to the deviation between the yaw angle and the sight angle and the size of the switching threshold value, the judgment is carried out, and the nonlinear azimuth attitude guidance and the nonlinear preposed guidance are switched
And step S30, performing nonlinear transformation on the view angle to respectively obtain three types of nonlinear signals, and then performing superposition to obtain nonlinear azimuth attitude guide signals.
And step S40, acquiring a leading angle according to the measured sight angle and the measured yaw angle. And solving three types of solution error information according to the yaw angle information, the lead angle information and the line-of-sight angle information, and then respectively carrying out nonlinear transformation, signal integration and superposition to finally obtain a nonlinear lead signal.
2. The integrated maneuver guiding method for switching between azimuth attitude and pre-guidance according to claim 1, wherein the switching between guidance methods is determined according to the deviation between yaw angle and line-of-sight angle, comprising
e1＝qh-ψ
Wherein psirFor the guiding law of final use, #1Pilot signal, psi, for non-linear azimuthal guidance2Is a pilot signal for nonlinear pilot. e.g. of the type1Deviation of the line of sight angle from the yaw angle, e1＝qh-ψ，qhIs the line of sight angle and psi is the yaw angle. Epsilon is a switching judgment threshold value. d is the distance between the aircraft and the target. | e1If | ≧ epsilon, performing nonlinear azimuth attitude guidance; if | e1If | < ε, then perform nonlinear pre-steering.
3. The integrated maneuver guiding method for switching azimuth attitude and pre-guidance according to claim 2, wherein the non-linear transformation of the line-of-sight angle respectively obtains three types of non-linear signals, so as to superimpose the three types of non-linear signals into the guiding signal of the non-linear azimuth attitude comprises
ψ1＝k2qh+k3q1+k4qs+k5qd
Wherein psi1Is a non-linear azimuth attitude guidance signal, wherein qh is a line-of-sight angle signal,
for line-of-sight angular rate signals, q1For non-linear transformation of the signal at line-of-sight angle, qdFor line-of-sight non-linear differential signals, k1、k2、k3、k4、k5、λ1And λ2Is a positive parameter that can be adjusted.
4. The integrated maneuver guiding method for switching between azimuth attitude and pre-guidance according to claim 2, wherein the step of solving the error information according to the yaw angle information, the pre-angle information and the line-of-sight angle information, performing the non-linear transformation, the signal integration and the superposition, and finally obtaining the non-linear pre-guidance signal comprises
e2＝ψ-ψ0
e1＝qh-ψ
e3＝sin(qh)-sin(ψ)
ez＝k12e3+k13e1f+k14e2f
ψ2＝k15ez+k16esz
Where ψ is the yaw angle, ψ0Is a leading angle, qhAngle of view, e1Deviation of the line of sight angle from the yaw angle, e2Deviation of yaw angle from lead angle, e3As an indirect non-linear azimuth deviation signal, e1fNon-linear azimuth deviation signal being line-of-sight angle and yaw angle, e2fIs a non-linear error signal of yaw angle relative to lead angle, ezFor error integration signal, eszNon-linearly integrated signal psi for error integration signal2The final nonlinear pilot signal. Wherein k is6、k7、k8、k9、k10、k11、k12、k13、k14、k15、k16And λ3、λ4、λ5Are parameters that can be adjusted.
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| CN111399529A (en) * | 2020-04-02 | 2020-07-10 | 上海交通大学 | Aircraft composite guiding method based on nonlinear sliding mode and preposition |
| CN111708382A (en) * | 2020-07-06 | 2020-09-25 | 中国人民解放军海军航空大学 | Aircraft guiding method based on non-linear proportional integral |
| CN111766776A (en) * | 2020-07-06 | 2020-10-13 | 中国人民解放军海军航空大学 | Pre-guiding method adopting nonlinear proportional and integral type compensation predictor |
| CN112034889A (en) * | 2020-09-10 | 2020-12-04 | 中国人民解放军海军航空大学 | Aircraft overload control method adopting nonlinear advanced network |
| CN112668626A (en) * | 2020-12-23 | 2021-04-16 | 中国人民解放军海军航空大学 | Reliability prediction method and system for radar seeker |
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| CN111708382B (en) * | 2020-07-06 | 2022-02-15 | 中国人民解放军海军航空大学 | Aircraft guiding method based on non-linear proportional integral |
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| CN112034889A (en) * | 2020-09-10 | 2020-12-04 | 中国人民解放军海军航空大学 | Aircraft overload control method adopting nonlinear advanced network |
| CN112034889B (en) * | 2020-09-10 | 2022-09-30 | 中国人民解放军海军航空大学 | Aircraft overload control method adopting nonlinear advanced network |
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