CN114153226B - Unmanned aerial vehicle view field keeping and guiding method and system assisted by dynamic sight line information - Google Patents

Unmanned aerial vehicle view field keeping and guiding method and system assisted by dynamic sight line information Download PDF

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CN114153226B
CN114153226B CN202111250587.8A CN202111250587A CN114153226B CN 114153226 B CN114153226 B CN 114153226B CN 202111250587 A CN202111250587 A CN 202111250587A CN 114153226 B CN114153226 B CN 114153226B
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aerial vehicle
unmanned aerial
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CN114153226A (en
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范军芳
陈仕伟
王伟
刘宁
王海森
牟清东
李俊贤
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Beijing Information Science and Technology University
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    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
    • G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/10—Simultaneous control of position or course in three dimensions
    • G05D1/101—Simultaneous control of position or course in three dimensions specially adapted for aircraft
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Abstract

The invention discloses a dynamic sight line information assisted unmanned aerial vehicle view field maintaining and guiding method and system, wherein a view field maintaining and guiding law considering overload constraint is adopted in the method; specifically, for the yaw direction, the dynamic sight line is used as a datum line, and the track of the aircraft is adjusted in real time by adopting a proportional guidance law according to the mutual motion relation between the aircraft and the target; in the pitching direction, in order to ensure that the target is within the maximum field angle range of the target detector at a specific distance, the aircraft always locks the central point of the target area to fly, and the attitude angle is constrained to be a proper range on the basis that the dynamic sight line is the reference line.

Description

Unmanned aerial vehicle view field keeping and guiding method and system assisted by dynamic sight line information
Technical Field
The invention relates to the field of aircraft guidance control, in particular to a dynamic sight information assisted unmanned aircraft view field keeping guidance method and system.
Background
In recent years, it has been a hot topic that an unmanned aerial vehicle performs target area detection and lands in a specific state to meet, and typical examples include a fighter aircraft landing at an aircraft carrier, a search and rescue helicopter landing at a specified area, an unmanned aerial vehicle sending a courier to a specified area, an air ground aircraft attacking a specific target in a certain area, and the like.
The common technical problems include target area identification and locking, and the attitude needs to be controlled in the flight process so as to ensure that target information is not lost, and finally the target information falls into a designated area in a specific state. In particular, the ground target hitting scene of the air-drop type aircraft is more challenging due to the relatively short flight time. Aiming at the characteristic that the launching height of an unmanned aerial vehicle aircraft launched by an aircraft type carrier changes in a certain interval, rules and requirements of a flight path scheme made therein are comprehensively considered, and some scholars provide a universal design method for a novel air-ground aircraft pitching attitude scheme adaptive to a large launching field high range.
However, many defects need to be overcome, for example, because the field of view of an image-seeking guidance head is small, the image-seeking guidance head can only provide information through attitude measurement sensors such as a gyroscope and the like, and cannot guarantee that an air-ground aircraft captures a target with high probability threatening the target, after the aircraft flies for a period of time, the attitude stability of the aircraft may not be enough to guarantee that the target is always in the field of view under extreme conditions, some schemes need to change the guidance rate or guide in sections in the flight process, and buffeting may occur at a switching point; some aircraft have difficulty identifying a particular target in a small area; and the parameters of the other aircrafts are not easy to adjust, the guidance law has high calculation complexity, the real-time performance of the project can be reduced, and the final control effect can be reduced.
For the above reasons, the present inventors have made intensive studies on the existing control method, dynamic sight information and field of view of an aircraft, and thus, it is expected that a control method and a control system capable of solving the above problems will be devised.
Disclosure of Invention
In order to overcome the problems, the inventor carries out intensive research and designs a dynamic sight line information-assisted field-of-view keeping and guiding method and system for the unmanned aerial vehicle, wherein a field-of-view keeping and guiding law considering overload constraint is adopted in the method; specifically, for the yaw direction, the dynamic sight line is used as a datum line, and the track of the aircraft is adjusted in real time by adopting a proportional guidance law according to the mutual motion relation between the aircraft and the target; in the pitching direction, in order to ensure that the target is within the maximum field angle range of the target detector at a specific distance, the aircraft always locks the central point of the target area to fly, and the attitude angle is restricted to be a proper range on the basis that the dynamic sight line is the reference line, thereby completing the invention.
Specifically, the invention aims to provide a visual field maintaining and guiding method of an unmanned aerial vehicle assisted by dynamic sight line information, which is characterized by comprising the following steps:
step 1, throwing out the unmanned aerial vehicle through a carrier,
step 2, the unmanned aerial vehicle performs inertial flight reservationTime tz0Then, adjusting the yaw direction deviation by adopting a proportional integral method, and restricting the pitch direction deviation by limiting the attitude angle through the field angle;
and 3, after the unmanned aerial vehicle enters a final guide section, adjusting the deviation of the yaw direction by adopting a proportional guide method, and restraining the deviation of the pitch direction through the attitude angle limited by the terminal angle.
Wherein, in step 2, the predetermined time tz0Is 0.3-0.5 s, and the unmanned aerial vehicle is thrown out for tz0After time, the start point is reached.
In the step 2, in the process of adjusting the deviation of the yaw direction by adopting a proportional integral method, the overload required for the yaw direction of the initial guide section is obtained in real time through the following formula (1):
uz1＝kpzeψ+kiz∫eψdt+kLz1eLz (1)
wherein u isz1Indicating that the yaw direction of the initial brake-guide section needs overload;
kpzrepresenting a yaw rate coefficient;
eψrepresenting a yaw angle real-time deviation;
kizrepresenting an integral term coefficient;
kLz1indicating a position deviation proportionality coefficient;
eLzand representing the real-time position deviation of the unmanned aerial vehicle in the Z-axis direction.
In the step 2, in the process of restricting the deviation of the pitching direction by limiting the attitude angle through the field angle, the overload required for the pitching direction of the initial guiding section is obtained in real time through the following formula (2):
uy1＝kPe+kI∫e dt+kLy1eLy1 (2)
wherein u isy1Indicating that the pitch direction of the initial brake and guide section needs overload;
kPrepresenting a real-time deviation proportionality coefficient of a track inclination angle;
e represents the track inclination error;
kIrepresents an integral term coefficient;
kLy1representing a first stage pitch direction position deviation proportionality coefficient;
eLy1and representing the real-time position deviation of the unmanned aerial vehicle of the initial guidance section in the Y-axis direction.
Wherein, in step 3, the UAV enters a terminal pilot segment when the UAV's position satisfies the following equation (3):
Ry≤RMT (3)
wherein R isyRepresenting the distance between the unmanned aerial vehicle and the target;
RMTrepresenting the distance between the starting point and the target in the terminal guidance phase; in the step 3, in the process of adjusting the deviation of the yaw direction by adopting a proportional guidance method, the overload required for the yaw direction of the final guide section is obtained in real time through the following formula (4):
Figure BDA0003322335530000041
wherein u isz2Indicating that the final guide section needs to be overloaded in the yaw direction;
Nzrepresents a scaling factor;
Vmzrepresenting the component of the velocity of the unmanned aerial vehicle in the Z-axis direction;
Figure BDA0003322335530000042
representing a dynamic line of sight azimuthal velocity;
kLz2expressing a position deviation proportionality coefficient of the last guide section in the pitching direction;
eLzrepresenting the real-time position deviation of the unmanned aerial vehicle in the Z-axis direction.
In step 3, in the process of constraining the deviation of the pitch direction through the attitude angle limited by the terminal angle, the overload required for the pitch direction of the final guide section is obtained in real time through the following formula (5):
Figure BDA0003322335530000043
wherein u isy2Indicating that the pitch direction of the final guide section needs overload;
ζ represents a damping term;
Np、Nq、NLy、Ngall represent a scaling factor;
Figure BDA0003322335530000044
representing dynamic view elevation angular velocity;
Figure BDA0003322335530000045
a derivative representing a distance between the UAV and the target;
qλrepresenting dynamic view elevation angles;
θTrepresenting a desired terminal angle;
θm2representing the inclination angle of the trajectory of the aircraft at the final guidance section;
eLy2and representing the real-time position deviation of the unmanned aerial vehicle at the final guide section in the Y-axis direction.
Wherein the track inclination error e is obtained by the following formula (6),
e＝θm-θn (6)
wherein, thetamRepresenting the track inclination angle of the aircraft;
θnan attitude angle representing a desired constraint is obtained by the following equation (7):
Figure BDA0003322335530000051
wherein (x)H,yH) Coordinates of a projection point of the control point in an XOY plane are represented;
(xF,yF) The coordinates of the projected point of the switching point into the final guide segment in the XOY plane are shown.
Wherein the damping term ζ is obtained by the following formula (8):
Figure BDA0003322335530000052
wherein R isMT0Representing the distance between the control point and the target;
Ryrepresenting the distance between the unmanned aerial vehicle and the target;
κ represents a damping coefficient.
The invention also provides a dynamic sight information assisted unmanned aerial vehicle visual field maintaining and guiding system, which is used for executing the dynamic sight information assisted unmanned aerial vehicle visual field maintaining and guiding method;
preferably, the system comprises an initial guidance segment guidance module and a final guidance segment guidance module;
in the initial guidance section guidance module, a proportional integral method is adopted to adjust the deviation of the yaw direction, and the deviation of the pitch direction is restrained by limiting the attitude angle through the field angle;
in the final guide section guide module, a proportional guide method is adopted to adjust the deviation of the yaw direction, and the deviation of the pitch direction is restrained through the attitude angle limited by the terminal angle.
The invention has the advantages that:
(1) according to the improved dynamic sight line information-assisted unmanned aerial vehicle visual field maintaining and guiding method and system, the track and the attitude angle of the unmanned aerial vehicle are reasonably constrained by taking the dynamic sight line as a reference, the target area can be locked when the control rate is switched, and the final control guiding section adopts an overload-constrained final angle constraining and guiding method, so that the target information detector can be ensured to always lock the target area and hit the target at the expected attitude angle under available overload;
(2) according to the improved dynamic sight line information-assisted unmanned aerial vehicle view field maintaining and guiding method and system, the initial position deviation is corrected by utilizing the deviation between the position of the aerial vehicle and the initial dynamic sight line and the obtained sight line deviation angle to assist the flight of the initial guidance section; by restricting the yaw angle, when the conductivity intersection point is guaranteed to be manufactured, the target area is in the field range, and interception intersection of the specific target closest to the central point of the target area can be achieved.
Drawings
FIG. 1 is a logic diagram illustrating an overall method for maintaining guidance of field of view of an unmanned aerial vehicle assisted by dynamic eye gaze information, in accordance with a preferred embodiment of the present invention;
FIG. 2 illustrates a three-dimensional flight trajectory plot obtained in an embodiment of the present invention;
FIG. 3 illustrates a plot of the XOY aircraft trajectory obtained in an embodiment of the present invention;
FIG. 4 shows a plot of the XOZ vehicle trajectory obtained in an embodiment of the present invention;
FIG. 5 shows a velocity profile obtained in an embodiment of the invention;
FIG. 6 shows a time plot of angle of attack, sideslip angle obtained in an embodiment of the present invention;
FIG. 7 is a graph of time curves of track inclination and declination obtained in an embodiment of the present invention;
FIG. 8 shows a time plot of pitch angle and yaw angle obtained in an embodiment of the present invention;
FIG. 9 shows a Y-axis overload graph obtained in an embodiment of the present invention;
FIG. 10 shows a Z-axis overload graph obtained in an embodiment of the invention;
FIG. 11 is a time plot of elevation and azimuth of a line of sight obtained in an embodiment of the present invention;
fig. 12 shows a view of the angle of field required at the point of view obtained in the embodiment of the present invention.
Detailed Description
The invention is explained in more detail below with reference to the figures and examples. The features and advantages of the present invention will become more apparent from the description.
The word "exemplary" is used exclusively herein to mean "serving as an example, embodiment, or illustration. Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. While the various aspects of the embodiments are presented in drawings, the drawings are not necessarily drawn to scale unless specifically indicated.
According to the method for maintaining and guiding the visual field of the unmanned aerial vehicle assisted by the dynamic sight line information, as shown in FIG. 1, the method comprises the following steps:
step 1, throwing out the unmanned aerial vehicle through a carrier,
in this application the aerial craft means can be in the aerial plane of throwing in unmanned aerial vehicle aircraft, can be helicopter or fixed wing aircraft, and the unmanned aerial vehicle in this application can not set up power unit, only sets up the steering wheel, can rely on inertia to continue to fly along the flight direction of aerial craft after being thrown from the aerial craft.
Step 2, the unmanned aerial vehicle flies for a preset time t under inertiaz0And then, adjusting the yaw direction deviation by adopting a proportional integral method, and restricting the pitch direction deviation by limiting the attitude angle through the field angle.
Wherein the predetermined time tz0Is 0.3-0.5 s, and the unmanned aerial vehicle is thrown out for tz0After time, the start point is reached. The coordinate position of the start control point passes through the coordinate position of the throwing point, the speed of the carrier and the preset time tz0And (4) resolving to obtain.
Per se, wherein the predetermined time tz0Set to 0.455s, the lateral distance of the aircraft can reach about 100m in the time, and the safety of the aircrafts and the unmanned aerial vehicles can be ensured.
All the coordinates of the throwing point coordinate, the starting point coordinate, the switching point coordinate entering the final guide section, the position of the unmanned aerial vehicle, the target position and the like are coordinates in a geodetic coordinate system, namely the coordinate system established by taking the reference ellipsoid as a datum plane in geodetic measurement.
In this application, after the aircraft arrived at the control point, the sensor spare of carrying on the aircraft began to work, and it needs the overload to obtain to control the flight attitude of aircraft in view of the above, the initial guidance section in this application be the guidance flight orbit between control point and the final guidance section of initiating.
In a preferred embodiment, in step 2, during the process of adjusting the deviation of the yaw direction by using a proportional-integral method, the yaw direction of the initial guidance segment is obtained in real time by the following formula (1):
uz1＝kpzeψ+kiz∫eψdt+kLz1eLz (1)
wherein u isz1Indicating that the yaw direction of the initial brake-guide section needs overload;
kpzthe yaw angle proportional coefficient is represented, and the value of the yaw angle proportional coefficient is 0.2;
eψrepresents a yaw angle real-time deviation, which is obtained by the following formula (1.1):
eψ＝ψm-arctan[Vz0tz0/(xH-xT)] (1.1)；
kizthe integral term coefficient is expressed, and the value of the integral term coefficient is 0.01;
kLz1the position deviation proportionality coefficient is expressed, and the value of the position deviation proportionality coefficient is 0.04;
eLzrepresenting the real-time position deviation of the unmanned aerial vehicle in the Z-axis direction, which is obtained by the following formula (1.2):
eLz＝zIMU-xIMU arctan qγ0 (1.2)
ψmthe trajectory drift angle of the aircraft can be replaced by the drift angle acquired by the IMU;
Vz0representing the component of the speed of the unmanned aerial vehicle in the Z-axis direction during starting and controlling;
zIMUthe Z-axis coordinate representing the mass center position of the unmanned aerial vehicle is obtained by real-time resolving through an inertial device carried on the unmanned aerial vehicle;
xIMUthe X-axis coordinate representing the mass center position of the unmanned aerial vehicle is obtained by real-time resolving through an inertial device carried on the unmanned aerial vehicle;
qγ0the dynamic sight line azimuth angle in the control starting is represented, and the angle is a fixed value and is obtained through the following formula (1.3):
Figure BDA0003322335530000091
after control is started, the dynamic sight azimuth angle passes through qγ0Acquiring a target detector;
the dynamic sight azimuth angle in the application refers to an angle of a connecting line between an aircraft and a target in a yaw direction, and the dynamic sight elevation angle in the application refers to an angle of a connecting line between the aircraft and the target in a pitch direction.
The dynamic sight line information includes a dynamic sight line azimuth angle and a dynamic sight line elevation angle.
In a preferred embodiment, in step 2, in the process of constraining the pitch direction deviation by limiting the attitude angle through the field angle, the initial guided segment pitch direction is obtained in real time by the following formula (2) to require overload:
uy1＝kPe+kI∫e dt+kLy1eLy1 (2)
wherein u isy1Indicating that the pitch direction of the initial guiding section needs overload;
kPrepresenting a real-time deviation proportionality coefficient of the track inclination angle, wherein the value of the proportionality coefficient is 5;
kIthe integral term coefficient is expressed, and the value of the integral term coefficient is 0.2;
kLy1the position deviation proportional coefficient of the pitching direction of the first stage is represented, and the value of the proportional coefficient is 0.3;
integral of the real-time deviation of the track inclination angle is denoted by e dt;
e represents a track pitch error, which is obtained by the following equation (6),
e＝θm-θn (6)
wherein, thetamThe track inclination angle of the aircraft can be replaced by a pitch angle acquired by an IMU;
θnan attitude angle representing the desired constraint is obtained by the following equation (7):
Figure BDA0003322335530000092
wherein (x)H,yH) Coordinates of a projection point of the control point in the XOY plane are represented;
(xF,yF) Coordinates representing the projected point of the switching point into the last leader segment in the XOY plane; the XOY is a plane formed by an X axis and a Y axis in a geodetic coordinate system.
Preferably, in the flight process of the initial guidance section, the distance between the unmanned aerial vehicle and the target is obtained by continuously calculating the real-time position of the aircraft through the IMU and combining the target position obtained by the target detector in real time or the target coordinate information filled in the aircraft, and then the distance is compared with the distance between the starting point and the target of the fixed value final guidance stage to judge whether the unmanned aerial vehicle enters the final guidance section, so that the control strategy switching is realized.
eLy1Representing the real-time position deviation of the unmanned aerial vehicle of the initial guidance section in the Y-axis direction, which is obtained by the following formula (2.1):
Figure BDA0003322335530000101
yIMUthe Y-axis coordinate representing the mass center position of the unmanned aerial vehicle is obtained by real-time resolving through an inertial device carried on the unmanned aerial vehicle;
qλ0the dynamic sight line elevation angle in the control starting is represented, the angle is a fixed value, and the angle is obtained through the following formula (2.2):
Figure BDA0003322335530000102
after control starting, the dynamic sight azimuth angle is obtained through a target detector; the target detector can be selected at will according to actual needs, can achieve the detection function, and can be set as a television seeker for example.
And 3, after the unmanned aerial vehicle enters a final guide section, adjusting the deviation of the yaw direction by adopting a proportional guide method, and restraining the deviation of the pitch direction through the attitude angle limited by the terminal angle.
Wherein the UAV enters a terminal pilot segment when the UAV's position satisfies the following equation (3):
Ry≤RMT＝800m
(3) wherein R isyRepresenting the distance between the unmanned aerial vehicle and the target;
RMTrepresenting the distance between the starting point of the last guidance phase and the target,
the value is 500-1500 m;
therein, the predetermined time distance RMTThe setting is 800m, so that the angle and overload constraint can be met;
Figure BDA0003322335530000111
and representing the included angle between the azimuth angle and the starting control point of the final guide section and the target position.
In a preferred embodiment, in step 3, during the adjustment of the yaw direction deviation by using the proportional guidance method, the real-time acquisition of the yaw direction of the last-guided segment by the following formula (4) requires overload:
Figure BDA0003322335530000112
wherein u isz2Indicating that the final guide section needs overload in the yaw direction;
Nzexpressing a proportionality coefficient, wherein the value of the proportionality coefficient is 2-6, and 3 is selected;
Vmzrepresenting the component of the speed of the unmanned aerial vehicle in the Z-axis direction, wherein the component is obtained through I MU real-time calculation;
Figure BDA0003322335530000113
the azimuth speed representing the dynamic sight line can be obtained by resolving through a target detector;
kLz2the position deviation proportionality coefficient is expressed, and the value of the position deviation proportionality coefficient is 0.2;
eLzindicates that there is noReal-time position deviation of the human aircraft in the Z-axis direction.
In a preferred embodiment, in step 3, in the process of constraining the pitch direction deviation by the attitude angle limited by the terminal angle, the overload required for the pitch direction of the final lead segment is obtained in real time by the following formula (5):
Figure BDA0003322335530000114
wherein u isy2Indicating that the pitch direction of the final guide section needs overload;
Np、Nq、NLy、Ngall represent a proportionality coefficient with the value Np＝1.5、 Nq＝1.1、NLy＝0.3、Ng＝0.4；
Figure BDA0003322335530000115
The high and low angular speeds of the dynamic sight line are represented and can be obtained through the solution of a target detector;
Figure BDA0003322335530000121
a derivative representing a distance between the unmanned aerial vehicle and the target, obtained in real time by the IMU;
qλthe height angle of the dynamic sight line is represented and can be obtained by resolving through a target detector;
θm2representing that the final guide section track inclination angle can be replaced by the pitch angle acquired by the IMU; specifically, the pitch angle obtained by the velocity arctangent function obtained by the IMU is equal to θm2；
θTRepresents the desired terminal angle, which takes the value-60 °; the best effect can be achieved by setting the terminal angle to-60 degrees;
eLy2representing the real-time position deviation of the final-stage unmanned aerial vehicle in the Y-axis direction, which is obtained by the following formula (5.1):
Figure BDA0003322335530000122
wherein the field angle range of the target information detector carried on the unmanned aerial vehicle is +/-qmax；
ζ represents a damping term obtained by the following formula (8):
Figure BDA0003322335530000123
wherein R isMTRepresenting the distance between the control and the target;
Ryrepresenting the distance between the unmanned aerial vehicle and the target, which is obtained in real time through the I MU;
kappa denotes the damping coefficient, which takes the value 0.75.
In the application, the damping term is set, so that an angle constraint instruction in a guidance instruction is reduced in a final guidance section, and a target is approached by a small acceleration; and gradually increasing the proportion of the angle restriction instruction in the guidance instruction as the aircraft approaches the target along with the progress of the guidance process. By introducing the damping item, the terminal line-of-sight angle constraint control method reduces the output value of the overall guidance instruction at the initial guidance stage, achieves the purpose of reducing the acceleration instruction, and meanwhile enables the unmanned aerial vehicle to meet the requirements on the final angle and the target miss distance.
The utility model provides an unmanned vehicles is last to be provided with the steering wheel, and what pitch and yaw direction were received simultaneously to the steering wheel needs to use the overload to receiving needs to do the fusion processing with the overload instruction, utilizes the result after fusing to beat the rudder control, so that satisfy pitch and yaw direction's control simultaneously.
The invention also provides a dynamic sight line information assisted unmanned aerial vehicle visual field maintaining and guiding system, which is used for executing the dynamic sight line information assisted unmanned aerial vehicle visual field maintaining and guiding method;
preferably, the system comprises an initial guidance segment guidance module and a final guidance segment guidance module;
in the initial guidance section guidance module, a proportional integral method is adopted to adjust the deviation of the yaw direction, and the deviation of the pitch direction is restrained by limiting the attitude angle through the field angle;
in the final guide section guide module, a yaw direction deviation is adjusted by adopting a proportional guide method, and a pitch direction deviation is restrained through an attitude angle limited by a terminal angle;
more preferably, the system is further provided with a terminal section judging module, which is used for judging whether the aircraft enters the terminal section in real time.
Example (b):
the flight state of the unmanned aerial vehicle after being thrown from the 500m high altitude is simulated in a mathematical simulation mode, and a dynamic sight line information assisted unmanned aerial vehicle view field maintaining and guiding method is filled in the unmanned aerial vehicle.
Wherein the initial takeoff point position of the unmanned aerial vehicle is Pm(0,500,0) m, angle of throw
Figure RE-GDA0003502140330000131
Aircraft initial velocity V0220 m/s; target position PtAfter 0.455s, the aircraft enters the start position, where the linear distance Δ R between the drone and the target is 800m, and the desired terminal angle θT＝-60°。
After starting control, the flight path of the aircraft enters an initial guiding section, at the moment, the yaw direction of the initial guiding section is obtained in real time through the following formula (1) and overload is needed, the pitch direction of the initial guiding section is obtained in real time through the following formula (2) and overload is needed, and the flight attitude is controlled through steering according to the overload,
uz1＝kpzeψ+kiz∫eψdt+kLz1eLz (1)
uy1＝kPe+kI∫e dt+kLy1eLy1 (2)
wherein k ispzIs taken to be 0.2, kizIs taken to be 0.01, kLz1Is taken to be 0.04, kPIs given a value of 5, kIHas a value of 0.2, kLy1Is 0.3;
eψobtained by the following formula (1.1):
eψ＝ψm-arctan[Vz0tz0/(xH-xT)] (1.1)；
eLzobtained by the following formula (1.2):
eLz＝zIMU-xIMU arctan qγ0 (1.2)；
ψmthe yaw angle obtained by the IMU can be used for replacing;
qγ0obtained by the following formula (1.3):
Figure BDA0003322335530000141
e is obtained by the following formula (6),
e＝θm-θn (6)
θmobtained by the IMU through the calculation of the IMU,
θnobtained by the following formula (7):
Figure BDA0003322335530000142
eLy1obtained by the following formula (2.1):
Figure BDA0003322335530000143
qλ0obtained by the following formula (2.2):
Figure BDA0003322335530000151
Vz0＝100/s，tz0＝0.455s，xH＝100m，yH＝500m，xF＝1800m， xT＝2500，RMT＝800m，xIMU、yIMUand zIMUThe method is obtained through real-time calculation of inertial devices carried on the unmanned aerial vehicle.
In the flying process of the aircraft, whether the aircraft enters a terminal pilot segment is judged in real time through the following formula (3),
Ry≤RMT＝800m (3)
wherein, when the formula (3) is established, the final guide section is entered, then the overload required by the yaw direction of the final guide section is obtained in real time through the following formula (4), the overload required by the pitch direction of the final guide section is obtained in real time through the following formula (5), and the flying attitude is controlled according to the overload required by the pitch direction,
Figure BDA0003322335530000152
Figure BDA0003322335530000153
wherein, NzIs taken as 3, NpIs 1.5, NqIs 1.1, NLyIs taken to be 0.3, NgIs 0.4;
Vmzobtained through real-time solution of the IMU,
Figure BDA0003322335530000154
is obtained by the calculation of the target detector,
Figure BDA0003322335530000155
is obtained by the calculation of the target detector,
Figure BDA0003322335530000156
obtained through real-time solution of the IMU,
qλis obtained by the calculation of the target detector,
θm2obtained through real-time solution of the IMU,
g＝9.8m/s2；kLz2is 0.2;
eLy2obtained by the following formula (5.1):
Figure BDA0003322335530000157
qmaxis +/-30 degrees;
ζ is obtained by the following formula (8):
Figure BDA0003322335530000161
Ryobtained through real-time solution of the IMU,
kappa is 0.75.
The unmanned aerial vehicle is controlled through the control mode, the target hit by the unmanned aerial vehicle is finally obtained, the three-dimensional flight trajectory curve of the target hit is shown in fig. 2, the XOY flight trajectory curve is shown in fig. 3, the XOZ flight trajectory curve is shown in fig. 4, the speed curve is shown in fig. 5, the time curves of attack angle and sideslip angle are shown in fig. 6, the time curves of track inclination angle and yaw angle are shown in fig. 7, the time curves of pitch angle and yaw angle are shown in fig. 8, the overload curve of Y-axis is shown in fig. 9, the overload curve of Z-axis is shown in fig. 10, the time curves of high and low angle of sight and azimuth angle are shown in fig. 11, and the angle of view required at the start point is shown in fig. 12.
The required field angle is smaller than the limit range of the target detector through the image, the target information detector is ensured to be always locked in the target area and hit at the expected attitude angle under available overload, and the final angle is constrained to the preset value, so that the aircraft can reach the designated position in an ideal state.
The present invention has been described above in connection with preferred embodiments, which are intended to be exemplary only and illustrative only. On the basis of the above, the invention can be subjected to various substitutions and modifications, and the substitutions and the modifications are all within the protection scope of the invention.

Claims (6)

1. The unmanned aerial vehicle visual field maintaining and guiding method assisted by dynamic sight line information is characterized by comprising the following steps of:
step 1, throwing out the unmanned aerial vehicle through a carrier,
step 2, the unmanned aerial vehicle flies for a preset time t under inertiaz0Then, adjusting the yaw direction deviation by adopting a proportional integral method, and restricting the pitch direction deviation by limiting the attitude angle through the field angle;
step 3, after the unmanned aerial vehicle enters a final guiding section, adjusting yaw direction deviation by adopting a proportional guidance method, and restraining pitch direction deviation through an attitude angle limited by a terminal angle;
in the step 2, in the process of adjusting the deviation of the yaw direction by adopting a proportional integral method, the initial guiding section yaw direction is obtained in real time by the following formula (1) and overload is required:
uz1＝kpzeψ+kiz∫eψdt+kLz1eLz (1)
wherein u isz1Indicating that the yaw direction of the initial brake-guide section needs overload;
kpzrepresenting a yaw rate coefficient;
eψrepresenting real-time deviation of the yaw angle;
kizrepresenting an integral term coefficient;
kLz1indicating a position deviation proportionality coefficient;
eLzrepresenting the real-time position deviation of the unmanned aerial vehicle in the Z-axis direction;
in step 2, in the process of restricting the deviation of the pitch direction by limiting the attitude angle through the field angle, the initial guidance segment pitch direction is obtained in real time through the following formula (2) that the overload is required:
uy1＝kPe+kI∫e dt+kLy1eLy1 (2)
wherein u isy1Indicating that the pitch direction of the initial brake and guide section needs overload;
kPrepresenting real-time deviation of track inclinationA proportionality coefficient;
e represents the track inclination error;
kIrepresenting an integral term coefficient;
kLy1representing a first stage pitch direction position deviation proportionality coefficient;
eLy1representing the real-time position deviation of the unmanned aerial vehicle of the initial guidance section in the Y-axis direction;
in step 3, in the process of adjusting the deviation of the yaw direction by adopting a proportional guidance method, the overload required by the yaw direction of the final guidance segment is obtained in real time through the following formula (4):
Figure FDA0003675366600000021
wherein u isz2Indicating that the final guide section needs overload in the yaw direction;
Nzrepresents a scaling factor;
Vmzrepresenting the component of the velocity of the unmanned aerial vehicle in the Z-axis direction;
Figure FDA0003675366600000022
representing a dynamic line of sight azimuthal velocity;
kLz2representing a position deviation proportionality coefficient of the final guide section in the pitching direction;
eLzrepresenting the real-time position deviation of the unmanned aerial vehicle in the Z-axis direction;
in step 3, in the process of constraining the deviation of the pitch direction through the attitude angle limited by the terminal angle, the overload required for the pitch direction of the final guide section is obtained in real time through the following formula (5):
Figure FDA0003675366600000023
wherein u isy2Indicating that the pitch direction of the final guide section needs overload;
ζ represents a damping term;
Np、Nq、NLy、Ngall represent a scaling factor;
Figure FDA0003675366600000024
representing dynamic view elevation angular velocity;
Figure FDA0003675366600000025
a derivative representing a distance between the UAV and the target;
qλrepresenting dynamic view elevation angles;
θTrepresenting a desired terminal angle;
θm2representing the track inclination angle of the final pilot segment aircraft;
eLy2and representing the real-time position deviation of the unmanned aerial vehicle at the final guide section in the Y-axis direction.
2. The dynamic gaze information assisted unmanned aerial vehicle field of view preservation guidance method of claim 1,
in step 2, the predetermined time tz0Is 0.3-0.5 s, and the unmanned aerial vehicle is thrown out for tz0After the time, the start point is reached.
3. The dynamic gaze information assisted unmanned aerial vehicle field of view preservation guidance method of claim 1,
in step 3, when the position of the unmanned aerial vehicle satisfies the following formula (3), the unmanned aerial vehicle enters a final guidance segment:
Ry≤RMT (3)
wherein R isyRepresenting the distance between the unmanned aerial vehicle and the target;
RMTrepresenting the distance between the starting point of the last guidance phase and the target.
4. The dynamic gaze information assisted unmanned aerial vehicle field of view preservation guidance method of claim 1,
the track inclination error e is obtained by the following formula (6),
e＝θm-θn (6)
wherein, thetamRepresenting the track inclination angle;
θnan attitude angle representing the desired constraint is obtained by the following equation (7):
Figure FDA0003675366600000031
wherein (x)H,yH) Coordinates of a projection point of the control point in an XOY plane are represented;
(xF,yF) The coordinates of the projected point of the switching point into the final leader segment in the XOY plane are shown.
5. The dynamic gaze information assisted unmanned aerial vehicle field of view preservation guidance method of claim 1,
the damping term ζ is obtained by the following formula (8):
Figure FDA0003675366600000041
wherein R isMT0Representing the distance between the point of departure and the target;
Ryrepresenting the distance between the unmanned aerial vehicle and the target;
κ represents a damping coefficient.
6. The visual field maintaining and guiding system of the unmanned aerial vehicle assisted by the dynamic sight line information is characterized in that,
the system is used for executing the visual field maintaining and guiding method of the unmanned aerial vehicle assisted by the dynamic sight line information according to one of claims 1 to 5;
the system comprises an initial guidance section guidance module and a final guidance section guidance module;
in the initial guidance section guidance module, a proportional integral method is adopted to adjust the deviation of the yaw direction, and the deviation of the pitch direction is restrained by limiting the attitude angle through the field angle;
in the final guide section guide module, a proportional guide method is adopted to adjust the deviation of the yaw direction, and the deviation of the pitch direction is restrained through the attitude angle limited by the terminal angle.
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