CN111895864B - Accelerometer-free overload pilot construction method for satellite guidance ammunition - Google Patents

Accelerometer-free overload pilot construction method for satellite guidance ammunition Download PDF

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CN111895864B
CN111895864B CN202010782130.0A CN202010782130A CN111895864B CN 111895864 B CN111895864 B CN 111895864B CN 202010782130 A CN202010782130 A CN 202010782130A CN 111895864 B CN111895864 B CN 111895864B
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overload
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data
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CN111895864A (en
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韩丁丁
张勇为
王刚
袁毅
卢朝林
李兴国
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Xi'an Ruigao Measurement And Control Technology Co ltd
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
    • F42B15/01—Arrangements thereon for guidance or control
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F41—WEAPONS
    • F41G—WEAPON SIGHTS; AIMING
    • F41G7/00—Direction control systems for self-propelled missiles
    • F41G7/34—Direction control systems for self-propelled missiles based on predetermined target position data
    • F41G7/346—Direction control systems for self-propelled missiles based on predetermined target position data using global navigation satellite systems, e.g. GPS, GALILEO, GLONASS

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  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)

Abstract

The invention discloses a construction method of an overload pilot of a satellite-guided ammunition without an accelerometer, which comprises the steps of measuring the motion information of a projectile body through a satellite navigation system, calculating the overload of the projectile body through an overload identification algorithm by utilizing the motion information, and constructing an overload loop. The method comprises the steps of measuring the motion information of a projectile body through a satellite navigation system, calculating overload of the projectile body through an overload identification algorithm by utilizing the motion information, and constructing an overload loop. Obviously, the method has no attitude coupling problem, so that the overload feedback information is more accurate.

Description

Accelerometer-free overload pilot construction method for satellite guidance ammunition
Technical Field
The invention belongs to the field of missile control, and relates to a construction method of an overload pilot of satellite-guided ammunition without an accelerometer.
Background
Most of guided ammunitions have under-damping characteristics, and when the flying environment is severe and external interference factors are large, the ammunition is easy to destabilize and drop due to violent oscillation of an ammunition body; due to the requirement of a modern battlefield on speed increasing and range increasing of the guided munition, the conditions of bandwidth reduction, slow response, poor static stability and poor maneuvering capability of the guided munition under a high-airspace low-density environment must be improved. The design of the automatic pilot can effectively improve the dynamic response characteristic of the guided ammunition. The control system consisting of the servomechanism, control surface or thrust vector elements, the projectile and sensors and the corresponding feedback circuit elements is commonly referred to as an autopilot. An overload pilot is a kind of automatic pilot, which includes a longitudinal overload feedback loop of the projectile body, and the response speed and stability margin can be optimized by configuring the control loop pole of the projectile body. In the current field of engineering applications, simple damping loop pilots have been equipped on some models of guided munitions, but the field of guided munitions that overload pilots has less application. The reason is that: guided munitions generally require low cost and miniaturization, and limit the application of a high-precision inertial navigation system, while low-cost MEMS accelerometers (hereinafter referred to as accelerometers) cannot isolate elastic vibration of a projectile body, have high output noise and are difficult to filter, and cannot be used in a control system; the guided munition usually has a spin characteristic, the output of the accelerometer is spin-modulated into a sinusoidal signal, if an overload signal with sufficient precision needs to be constructed, a roll angle signal with higher precision is needed, if the roll angle signal has a constant value deviation, an overload pilot channel is coupled, the hit precision of the guided munition is influenced if the roll angle signal is light, and the guided munition is off-target if the roll angle signal is heavy. The above are reasons why it is difficult to widely apply the overload pilot to the conventional guided munition.
In the research background, the patent provides a method for constructing an overload pilot of a satellite-guided ammunition without an accelerometer, and the overload pilot can be constructed by generating an overload feedback signal by using satellite navigation information only by carrying a satellite navigation system on the ammunition.
Disclosure of Invention
The invention aims to overcome the defects of the prior art and provides a construction method of an overload pilot of a satellite-guided ammunition without an accelerometer. Obviously, the method has no attitude coupling problem, so that the overload feedback information is more accurate.
The implementation of the patent scheme is detailed below.
(1) Defining a transmission coordinate system:
origin o-emission point;
x-axis-emission direction in the horizontal plane;
y-axis-in the vertical plane containing the emission direction in the horizontal plane, perpendicular to the direction in the x-axis;
the z axis, the x axis and the y axis form a right-hand coordinate system;
(2) defining a speed coordinate system:
origin o' -center of mass of the projectile;
the x' axis — coincides with the velocity vector direction;
y 'axis-lying within the vertical plane containing the velocity vector, perpendicular to the direction in the x' axis;
the z ' axis, x ' and y ' axes form a right-hand coordinate system;
(3) orientation in missile-borne computers
Figure GDA0003562171060000038
Figure GDA0003562171060000037
The included angle between the projection of the emission direction on the horizontal plane and the positive north direction takes the north bias west as the positive.
(4) The satellite navigation system outputs the motion information of the projectile body in real time after the satellite is captured and positioned, and the satellite navigation system mainly comprises positioning information: longitude LongMLatitude latMHeight altMAnd speed measurement information: east speed VE, north speed VN and sky speed VV. The signal frequency is typically 10 Hz.
(5) The missile-borne computer performs the following steps:
(5.1) performing the following operation once per updating of the satellite navigation signal:
Figure GDA0003562171060000031
VYg＝VV
Figure GDA0003562171060000032
Figure GDA0003562171060000033
Figure GDA0003562171060000034
Figure GDA0003562171060000035
VXg,VYg,VZg-x, y, z velocity of the ammunition in the firing coordinate system
Theta-inclination of trajectory, angle between projectile velocity and plane of launch coordinate system xoz
ψVBallistic declination, the angle between the projection of projectile velocity on the plane of launch coordinate system xoz and the x-axis of the launch coordinate system
(5.2) with the time reference point of 1s before the starting control time, the missile-borne computer allocates a memory for storing the following data: (t, θ, ψ)V) Where t represents the computer time. Continuously storing 10 sets of data from a time reference point (t)i,θi,ψVi) (i is 1 to 10). When enough 10 sets of data are stored, every time the data are updated, the new data are updatedV) And data, discarding the first packet of data in the stored data, and adding new data to the end of the data packet to ensure that 10 groups of data in the stored data are up-to-date.
(5.3) reaching the start-up time, the missile-borne computer executing the following algorithm:
(5.3.1) solving ballistic inclination Rate
Figure GDA0003562171060000041
Rate of change of ballistic declination
Figure GDA0003562171060000042
Figure GDA0003562171060000043
(5.3.2) constructing a virtual overload signal:
Figure GDA0003562171060000044
Figure GDA0003562171060000045
wherein:
ay＇voverload in the x 'o' y 'plane of the speed coordinate system, also called y' direction overload
az＇vOverload in the x 'o' z 'plane of the speed coordinate system, also called z' direction overload
A is toy＇v、az＇vThe overload feedback loop can be constructed.
Concept definition
Figure GDA0003562171060000046
Figure GDA0003562171060000051
Two symbol list
Figure GDA0003562171060000052
Has the advantages that:
the invention provides a construction method of an overload pilot of a satellite-guided ammunition without an accelerometer, which can generate an overload feedback signal by using satellite navigation information only by carrying a satellite navigation system on the ammunition to construct the overload pilot.
Drawings
FIG. 1 is a schematic diagram of a two-circuit overload pilot;
FIG. 2 is a schematic diagram of a two-overload pilot configuration using the disclosed technique;
FIG. 3 is a graph comparing virtual overload and real overload identified by the disclosed scheme;
figure 4 is a graph comparing overload to true overload to which an accelerometer is sensitive.
Detailed Description
The invention is described in further detail below with reference to the accompanying drawings:
the common overload pilot comprises a two-loop overload pilot and a three-loop overload pilot, and the construction idea of the overload pilot is described by taking the two-loop overload pilot as an example. The two-circuit overload pilot structure is shown in fig. 1.
In fig. 1, a feedback loop constructed by the angular rate information of the projectile measured by the angular rate gyro is called a damping loop, and the damping loop can improve the under-damping characteristic of the projectile; a feedback loop constructed by the overload information measured by the accelerometer is called an overload loop, and the overload loop can improve the frequency characteristics of the control system as a whole. Since the accelerometer is not generally installed at the center of mass of the projectile, the output signal of the accelerometer contains the second order differential of the attitude information of the projectile, which is called attitude coupling, and the amplitude of the second order differential is positively correlated with the distance from the installation position of the accelerometer to the center of mass. The patent provides a method for identifying projectile overload and constructing an overload loop by using data measured by a satellite navigation system, an accelerometer is not used, and the improved structure is as follows:
as shown in fig. 2, the motion information of the projectile body is measured by a satellite navigation system, and the overload of the projectile body is calculated by an overload identification algorithm using the motion information to construct an overload loop. Obviously, the method has no attitude coupling problem, so that the overload feedback information is more accurate. The implementation of the patent scheme is detailed below.
(1) Defining a transmission coordinate system:
origin o-emission point;
x-axis-emission direction in the horizontal plane;
y-axis-in the vertical plane containing the emission direction in the horizontal plane, perpendicular to the direction in the x-axis;
the z axis, the x axis and the y axis form a right-hand coordinate system;
(2) defining a speed coordinate system:
origin o' -center of mass of the projectile;
the x' axis — coincides with the velocity vector direction;
y 'axis-lying within the vertical plane containing the velocity vector, perpendicular to the direction in the x' axis;
the z ' axis, x ' and y ' axes form a right-hand coordinate system;
(3) orientation in missile-borne computers
Figure GDA0003562171060000071
Figure GDA0003562171060000072
The included angle between the projection of the emission direction on the horizontal plane and the positive north direction takes the north bias west as the positive.
(4) The satellite navigation system outputs the motion information of the projectile body in real time after the satellite is captured and positioned, and the satellite navigation system mainly comprises positioning information: longitude LongMLatitude latMHeight altMAnd speed measurement information: east VE, north VN and sky VV, the signal frequency is typically 10 Hz.
(5) The missile-borne computer performs the following steps:
(5.1) performing the following operation once per updating of the satellite navigation signal:
Figure GDA0003562171060000081
VYg＝VV
Figure GDA0003562171060000082
Figure GDA0003562171060000083
Figure GDA0003562171060000084
Figure GDA0003562171060000085
VXg,VYg,VZg-x, y, z velocity of the ammunition in the firing coordinate system
Theta-inclination of trajectory, angle between projectile velocity and plane of launch coordinate system xoz
ψVBallistic declination, the angle between the projection of projectile velocity on the plane of launch coordinate system xoz and the x-axis of the launch coordinate system
(5.2) with the time reference point of 1s before the starting control time, the missile-borne computer allocates a memory for storing the following data: (t, θ, ψ)V) Where t represents the computer time. Continuously storing 10 sets of data from a time reference point (t)i,θi,ψVi) (i is 1 to 10). When enough 10 sets of data are stored, every time the data are updated, the new data are updatedV) And data, discarding the first packet of data in the stored data, and adding new data to the end of the data packet to ensure that 10 groups of data in the stored data are up-to-date.
(5.3) reaching the start-up time, the missile-borne computer executing the following algorithm:
(5.3.1) solving ballistic inclination Rate
Figure GDA0003562171060000086
Rate of change of ballistic declination
Figure GDA0003562171060000087
Figure GDA0003562171060000091
(5.3.2) constructing a virtual overload signal:
Figure GDA0003562171060000092
Figure GDA0003562171060000093
wherein:
ay＇voverload in the x 'o' y 'plane of the speed coordinate system, also called y' direction overload;
az＇voverload in the x 'o' z 'plane of the speed coordinate system, also called z' direction overload;
a is toy＇v、az＇vThe overload feedback loop can be constructed.
Three simulation cases
The simulation is carried out by taking a certain guided ammunition as an example, the speed is set to be 220m/s, the initial shooting direction is 0 degree, the initial projectile distance is 8000m, and the target is positioned in the direction of-20 degrees. The virtual overload of the patent scheme is compared with the accelerometer-sensitive overload through mathematical simulation. Setting the updating frequency of a satellite navigation system as 10 Hz; the positioning error is in accordance with normal distribution, and the standard deviation is 3.3 m; the speed measurement error is in accordance with normal distribution, and the standard deviation is 0.66 m/s; setting the fluctuation of the added amplitude caused by elastic vibration of the projectile body to be in accordance with normal distribution, wherein the standard deviation is 0.005 g; the accelerometer mounting position is 10cm away from the centroid. The comparison of virtual overload and real overload identified by the patent scheme is shown in figure 3, and the comparison of real overload and real overload sensitive by an accelerometer is shown in figure 4.
As can be known from simulation, the offset of less than 2.5 percent exists between the virtual overload and the real overload constructed by the method, and the fluctuation amplitude is less than 3 percent; while the overload measured directly using the accelerometer has an offset of at least 10% from the true overload and a fluctuation of up to 5% due to the attitude coupling in relation to the elastic vibration of the projectile. No matter which index is compared, the scheme of the patent is far superior to the scheme of the accelerometer. Especially for ammunition using a satellite guidance system, the separation from the accelerometer saves more cost because the satellite navigation system is originally equipped on the ammunition.
(1) Illustrating the difference between constructing an overload loop and not constructing an overload loop
The implementation of the overload loop allows the configuration of the poles of the projectile control system with a substantially improved response relative to the case where the overload loop is not implemented (no feedback loop or only a damping loop is used). This is proof of theory and is a common recognition in the industry. Reference can be made to reading relevant books and documents, such as tactical missile overload pilot design and guidance law analysis (Lindefu, Wanghui, Wangjiang, Van military Fang, etc.).
(2) Illustrating the advantages of this patent over the traditional use of accelerometers to construct the overload loop:
(2.1) compared with the traditional method, the overload loop has no attitude coupling, the overload signal is compared with the true value, the offset is reduced by about 70%, and the noise level is reduced by about 50%;
(2.2) when the method is applied to the spinning characteristic ammunition, compared with the traditional method, the method does not need to process the modulated acceleration signal, the accuracy of the rolling angle is not required, and the accuracy of the rolling angle of the spinning characteristic ammunition is generally +/-10 degrees, so the channel coupling degree is reduced by about 5 percent.
(2.3) due to the characteristics of (2.1) and (2.2), the patent can reduce the CEP value of satellite guidance ammunition using a traditional overload pilot by at least about 30 percent.
And (2.4) the system is separated from an accelerometer, the cost is reduced to a small extent, the space on the missile is saved, the miniaturization of the guided munition is more favorably realized, and the development trend of the modern guided munition is met.
Although the present invention has been described with reference to a preferred embodiment, it should be understood that various changes, substitutions and alterations can be made herein without departing from the spirit and scope of the invention as defined by the appended claims.

Claims (1)

1. A construction method of an accelerometer-free overload pilot of satellite-guided munitions is characterized by comprising the following steps: measuring the motion information of the projectile body through a satellite navigation system, calculating the overload of the projectile body through an overload identification algorithm by utilizing the motion information, and constructing an overload loop; the method comprises the following steps:
(1) defining a transmission coordinate system:
origin o-emission point;
x-axis-emission direction in the horizontal plane;
y-axis-in the vertical plane containing the emission direction in the horizontal plane, perpendicular to the direction in the x-axis;
the z axis, the x axis and the y axis form a right-hand coordinate system;
(2) defining a speed coordinate system:
origin o' -center of mass of the projectile;
the x' axis — coincides with the velocity vector direction;
y 'axis-lying within the vertical plane containing the velocity vector, perpendicular to the direction in the x' axis;
the z ' axis, x ' and y ' axes form a right-hand coordinate system;
(3) orientation in missile-borne computers
Figure FDA0003562171050000011
Figure FDA0003562171050000012
The included angle between the projection of the emission direction on the horizontal plane and the positive north direction takes the north-west as positive;
(4) the satellite navigation system outputs the motion information of the projectile body in real time after the satellite is captured and positioned, and the motion information comprises positioning information: longitude LongMLatitude latMHeight altMAnd speed measurement information: east speed VE, north speed VN, and sky speed VV; the signal frequency is typically 10 Hz;
(5) constructing a virtual overload signal by the missile-borne computer;
the step (5) comprises the following steps:
(5.1) performing the following operation once per updating of the satellite navigation signal:
Figure FDA0003562171050000021
VYg＝VV
Figure FDA0003562171050000022
Figure FDA0003562171050000023
Figure FDA0003562171050000024
Figure FDA0003562171050000025
VXg,VYg,VZg-x, y, z velocity of the ammunition in the firing coordinate system;
theta is the angle between the trajectory inclination angle and the projectile velocity and the plane of the launch coordinate system xoz;
ψVballistic declination, the angle between the projection of projectile velocity on the plane of launch coordinate system xoz and the x-axis of the launch coordinate system;
(5.2) with the first 1s of the start control time as a time reference point, the missile-borne computer allocates a memory for storing the following data: (t, θ, ψ)V) Wherein t represents a computer time; continuously storing 10 sets of data from a time reference point (t)i,θi,ψVi) (i is 1 to 10); when enough 10 sets of data are stored, every time the data are updated, the new data are updatedV) Data, namely discarding the first packet of data in the stored data, and adding new data to the tail of the data packet to ensure that 10 groups of data in the stored data are up-to-date;
(5.3) when the start control time is reached, the missile-borne computer executes the following algorithm to construct a virtual overload signal;
the step (5.3) comprises the following steps:
(5.3.1) solving the trajectoryRate of change of inclination
Figure FDA0003562171050000026
Rate of change of ballistic declination
Figure FDA0003562171050000027
Figure FDA0003562171050000031
(5.3.2) constructing a virtual overload signal:
Figure FDA0003562171050000032
Figure FDA0003562171050000033
wherein:
ay＇voverload in the x 'o' y 'plane of the speed coordinate system, also called y' direction overload
az＇vOverload in the x 'o' z 'plane of the speed coordinate system, also called z' direction overload
A is toy＇v、az＇vThe overload feedback loop can be constructed.
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CN102425980A (en) * 2011-09-15 2012-04-25 北京理工大学 Control method for realizing overload autopilot by using accelerometer
CN103076806A (en) * 2011-10-26 2013-05-01 北京航天长征飞行器研究所 Integrated analyzing and setting method for control parameters of three-loop automatic pilot
CN104792232A (en) * 2015-04-28 2015-07-22 北京理工大学 Minimum overload terminal guiding method with terminal angular constraint
CN111142371A (en) * 2019-12-25 2020-05-12 中国人民解放军海军航空大学 Aircraft overload loop design method for providing damping by adopting angular acceleration
CN111434586A (en) * 2019-01-14 2020-07-21 北京理工大学 Aircraft guidance control system

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FR2918168B1 (en) * 2007-06-27 2009-08-28 Nexter Munitions Sa METHOD FOR CONTROLLING THE RELEASE OF AN ATTACK MODULE AND DEVICE USING SUCH A METHOD
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CN102425980A (en) * 2011-09-15 2012-04-25 北京理工大学 Control method for realizing overload autopilot by using accelerometer
CN103076806A (en) * 2011-10-26 2013-05-01 北京航天长征飞行器研究所 Integrated analyzing and setting method for control parameters of three-loop automatic pilot
CN104792232A (en) * 2015-04-28 2015-07-22 北京理工大学 Minimum overload terminal guiding method with terminal angular constraint
CN111434586A (en) * 2019-01-14 2020-07-21 北京理工大学 Aircraft guidance control system
CN111142371A (en) * 2019-12-25 2020-05-12 中国人民解放军海军航空大学 Aircraft overload loop design method for providing damping by adopting angular acceleration

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