CN108759589A - A kind of uncoupled method and device of rotary missile Guidance and control - Google Patents
A kind of uncoupled method and device of rotary missile Guidance and control Download PDFInfo
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- CN108759589A CN108759589A CN201810779176.XA CN201810779176A CN108759589A CN 108759589 A CN108759589 A CN 108759589A CN 201810779176 A CN201810779176 A CN 201810779176A CN 108759589 A CN108759589 A CN 108759589A
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- 238000000034 method Methods 0.000 title claims abstract description 14
- 230000006340 racemization Effects 0.000 claims abstract description 22
- 238000005259 measurement Methods 0.000 claims abstract description 14
- 230000008878 coupling Effects 0.000 claims abstract description 9
- 238000010168 coupling process Methods 0.000 claims abstract description 9
- 238000005859 coupling reaction Methods 0.000 claims abstract description 9
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- 239000011159 matrix material Substances 0.000 claims description 12
- 230000000694 effects Effects 0.000 claims description 3
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- 238000005096 rolling process Methods 0.000 abstract description 2
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- 230000009466 transformation Effects 0.000 description 1
- 201000009482 yaws Diseases 0.000 description 1
Classifications
-
- 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
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/20—Design optimisation, verification or simulation
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- Chemical & Material Sciences (AREA)
- Aviation & Aerospace Engineering (AREA)
- Combustion & Propulsion (AREA)
- Computer Hardware Design (AREA)
- Evolutionary Computation (AREA)
- Geometry (AREA)
- General Physics & Mathematics (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
A kind of method of the decoupling Guidance and control of rotary missile, steps are as follows：1) racemization bay section is set between the preceding bay section of rotary missile and rear bay section；2) inertial measurement component is installed in preceding bay section, the sensitive axes of the accelerometer and gyro that make inertial measurement component are parallel with each axis of bay section before rotary missile；3) the three axis angular rate data and acceleration information exported to inertial measurement component by control panel are acquired and calculate, and obtain the roll angle angle value for being coupled to front deck end；4) by the difference of bay section roll angle before calculating and setting roll angle, the motor in racemization bay section is controlled so that the coupling between yaw, pitching and rolling triple channel is reduced to minimum；Racemization bay section includes shell, and enclosure is equipped with motor and bearing, axis is housed in motor and Bearing inner, the inside of axis is equipped with slip ring, and one end of axis is equipped with the structure interface of rear bay section；Solve the problems, such as yaw and pitch channel close coupling, control difficulty in existing rotary missile guidance.
Description
Technical field
The invention belongs to rotary missile Guidance and control technical fields, and in particular to a kind of decoupling Guidance and control of rotary missile
Method and device.
Background technology
With the transformation of modern war pattern, the core requirement of weaponry is had turned to smart possessed by weapon
True striking capabilities.How to improve its attack precision and has become the forward position studied both at home and abroad and hot spot.Rotating missile is as a kind of ratio
It is more suited to conventional ammunition low cost, high-precision modification scheme.Rotary missile has unique advantage using the scheme of rotation,
For example the influence for starting asymmetry, structure asymmetry, motor power bias etc. to trajectory can be reduced, it is conducive to body structure cloth
Office and Miniaturization Design reduce cost production.Also there are some inevitable problems simultaneously, rotary missile is because of gyroscopic effect
The inertia of induction is crosslinked, and the control crosslinking of dynamics delay induction, the close coupling between pitching and jaw channel, body is to external loop
The response of guidance system instruction is poor, influences guidance precision.Due to the unique kinetic characteristics of rotary missile, just to its controlling party
Method proposes certain particular/special requirement, and the decoupling problem of pitching and jaw channel control becomes study now main and asks
Topic.Based on this, it is necessary to a kind of uncoupled method of Guidance and control is invented, to solve above-mentioned existing for existing rotary missile ask
Topic, and retain advantage possessed by rotary missile.
Invention content
To overcome above-mentioned the deficiencies in the prior art, the object of the present invention is to provide a kind of decoupling Guidance and controls of rotary missile
Method and device, solve in existing rotary missile Guidance and control technology that yaw and pitch channel close coupling, control is difficult asks
Topic.
To achieve the above object, the technical solution adopted by the present invention is：A kind of side of the decoupling Guidance and control of rotary missile
Method includes the following steps：
Step 1, a racemization bay section is added between the preceding bay section of rotary missile and rear bay section, racemization bay section includes outer
Shell, enclosure are equipped with motor and bearing, and axis is housed in motor and Bearing inner, and the inside of axis is equipped with slip ring, and the one of axis
Structure interface of the end equipped with rear bay section；The effect of the slip ring is to ensure the electrical connection of preceding bay section and rear bay section；
Step 2, inertial measurement component is installed in preceding bay section, makes the accelerometer of inertial measurement component and the sensitive axes of gyro
It is parallel with each axis of bay section before rotary missile；
Step 3, racemization bay section is modeled and is emulated, obtain ratio, product used in the PID controller of meet demand
Divide, differential parameter；
Step 4, the three axis angular rate data and acceleration information exported to inertial measurement component by control panel are adopted
Collect and be calculated the roll angle angle value that guided missile is coupled to front deck end, specific calculating process is as follows：
A, autoregistration establishes navigational coordinate system before MISSILE LAUNCHING first with IMU output informations, and autoregistration algorithm is such as
Shown in lower：
Analytic coarse alignment is carried out using accelerometer and gyro output information, and inertia device information is according to 50ms (i.e. 10
Period) smoothing processing to inhibit meter noise to influence,
φ=0
Wherein θ, γ, φ is respectively pitch angle, roll angle and yaw angle, ax, ay, and az indicates the x of missile coordinate system respectively
Axis, y-axis, the acceleration of z-axis.
Relationship between quaternary number and attitude angle is：
B, strapdown resolves, and carries out attitude angle update,
Due to the rotational angular velocity of missile coordinate system relative inertness coordinate systemFor：
WhereinFor gyro output angle speed,The projection for being rotational-angular velocity of the earth in navigational coordinate system,
For navigational coordinate system n to the attitude matrix of missile coordinate system b, since carrier initial heading is unknown, rotational-angular velocity of the earth suddenly,
Therefore：
According to the angular speed increment of previous moment quaternary number and current period, quaternary number is updated, not using attitude matrix
Same representation calculates current attitude angle, realizes solving of attitude,
Quaternary number recurrence formula：
Wherein
Missile coordinate system b is to navigational coordinate system n attitude matrixsWith quaternary number q0、q1、q2、q3Relationship it is as follows：
Attitude matrixIt is as follows with the relationship of current pitching angle theta, roll angle γ and yaw angle φ：
Compare the above attitude matrixCurrent pitching angle theta, roll angle γ and yaw angle φ can be obtained
Step 5, by the difference of bay section roll angle before calculating and setting roll angle, PID controller is used, controls racemization bay section
Interior motor, the rotating speed of bay section is maintained within the scope of setting value ± 2.0 ° before making so that will be between yaw, pitching and roll channel
Coupling be reduced to minimum.
A kind of device of the decoupling Guidance and control of rotary missile includes shell, and enclosure is equipped with motor and bearing,
Axis is housed in motor and Bearing inner, the inside of axis is equipped with slip ring, and one end of axis is equipped with the structure interface of rear bay section.
The present invention has the following advantages：
1) present invention is 0 ° due to controlling the roll angle of rotary missile, solves pitching and jaw channel controls coupled problem；
2) present invention due to control rotary missile roll angle be 0 °, rudder control instruction can guided missile pitch angle and partially
Directly in response to so improving response of the body to external loop guidance system to instruction on boat angle；
3) present invention is 0 ° due to controlling the roll angle of rotary missile, avoids pitching rudder and yaws the coupling in rudder control
It closes, the Guidance and control of guided missile is improved, so trajectory control accuracy can be improved.
The present invention completes the control to preceding bay section roll angle by inertial measurement component, control circuit and racemization bay section,
The roll angle of bay section not only ensure that advantage possessed by rotary missile itself within the scope of setting value ± 2.0 ° before ensureing, and
And the coupled relation between pitching and jaw channel is reduced, reach to the uncoupled purpose of rotary missile Guidance and control, improves
Trajectory control parameter.
Description of the drawings
Fig. 1 is the structural schematic diagram of racemization bay section of the present invention.
Fig. 2 is racemization bay section simulation model of the present invention.
Fig. 3 is the workflow of racemization bay section of the present invention.
Fig. 4 is the photo in kind of racemization bay section of the present invention.
Specific implementation mode
Invention is further described in detail with reference to the accompanying drawings and examples.
Embodiment one
A kind of method of the decoupling Guidance and control of rotary missile, specific implementation include the following steps：
Step 1, a racemization bay section, racemization bay section, pictorial diagram are added between the preceding bay section of rotary missile and rear bay section
As shown in Figure 4；
Step 2, bay section installs inertial measurement component before rotary missile, makes the accelerometer and gyro of inertial measurement component
Sensitive axes it is parallel with each axis of rotary missile；
Step 3, racemization bay section is modeled and is emulated, obtain ratio, product used in the PID controller of meet demand
Divide, differential parameter.Fig. 2 is the Controlling model that racemization bay section is established in Fig. 4, and model uses unity negative feedback model, in model
The transmission function of motor torque and voltage is G1, and the transmission function between motor angular velocity and electric current is G2, electric moter voltage and PID
Transmission function between controller output is G3, and the relationship between counter electromotive force of motor and motor angular velocity is G4, angle and angle
The transmission function of speed is G5, because there are voltage amplitude limits in specific implement, therefore an amplitude limit model, Fig. 4 is added behind G3
Proportionality coefficient is 153.8 used in middle racemization bay section, and proportionality coefficient is 253.2；Integral coefficient is 9.6；
G3=12
G4=0.086
Wherein：T (s) is the torque that motor provides, and U (s) is motor both end voltage, and W (s) is the angular speed of motor, Tf(s)
For the moment of friction of system；
Step 4, the three axis angular rate data and acceleration information exported to inertial measurement component by control panel are adopted
Collect and be calculated the roll angle angle value that guided missile is coupled to front deck end, specific calculating process is as follows：
A, autoregistration resolves, and navigational coordinate system is established first with IMU output informations before MISSILE LAUNCHING, and autoregistration is calculated
Method is as follows：
Analytic coarse alignment is carried out using accelerometer and gyro output information, and inertia device information is according to 50ms (i.e. 10
Period) smoothing processing to be to inhibit meter noise to influence.
φ=0
Wherein θ, γ, φ is respectively pitch angle, roll angle and yaw angle, ax, ay, and az indicates the x of missile coordinate system respectively
Axis, y-axis, the acceleration of z-axis,
Relationship between quaternary number and attitude angle is：
B, strapdown resolves, and carries out attitude angle update,
Due to being in the rotational angular velocity of body coordinate system relative inertness coordinate system：
WhereinFor gyro output angle speed,The projection for being rotational-angular velocity of the earth in navigational coordinate system,
For navigational coordinate system n to the attitude matrix of missile coordinate system b, since carrier initial heading is unknown, rotational-angular velocity of the earth suddenly,
Therefore：
According to the angular speed increment of previous moment quaternary number and current period, quaternary number is updated, not using attitude matrix
Same representation calculates current attitude angle, realizes solving of attitude,
Quaternary number recurrence formula：
Wherein
Missile coordinate system b is to navigational coordinate system n attitude matrixsWith quaternary number q0、q1、q2、q3Relationship it is as follows：
Attitude matrixIt is as follows with the relationship of current pitching angle theta, roll angle γ and yaw angle φ：
The above attitude matrix is compared, current pitching angle theta, roll angle γ and yaw angle φ are obtained
Step 5, by the error between bay section roll angle before calculating and setting roll angle, disappeared using PID controller control
The motor in bay section is revolved,
Step 6, the parameter of PID controller is adjusted on the basis of the pid parameter that racemization model emulation goes out so that in rear deck
In the case of Duan Xuanzhuan, the rotating speed of preceding bay section is maintained within the scope of setting value ± 2.0 ° so that by yaw, pitching and roll channel
Between coupling be reduced to minimum.
A kind of device of the decoupling Guidance and control of rotary missile, includes shell 1, inside shell (1) equipped with motor (3) with
And bearing (4), axis (5) is housed inside motor (3) and bearing (4), the inside of axis (5) is equipped with slip ring (2), and the one of axis (5)
Structure interface (6) of the end equipped with rear bay section.
The present invention operation principle be：
The operation principle of the present invention is theoretical using closed-loop control, and three axis at front deck end are acquired using inertial measurement component
Angular speed and acceleration value calculate the roll angle at front deck end by navigation algorithm, then according to the roll angle at front deck end with set
The error set between value reaches the roll angle at front deck end using PID controller control motor to control the roll angle at front deck end
For the purpose within the scope of setting value ± 2.0 °, to solve the problems, such as that yaw and pitch channel couples, to improve rotary missile
Control accuracy.
The algorithm flow chart of the present invention is as shown in Figure 3.It is initialized first, completion hardware initialization, parameter initialization,
Then the roll angle that front deck end is calculated using autoregistration, is since then resolved using strapdown, the rolling at update guided missile front deck end
Angle is rotated according to the error between the roll angle and setting value of current rotary missile using PID controller control motor so that
The roll angle at front deck end is within the scope of setting value ± 2.0 °, thus will be between yaw and pitch channel under bay section rotational case afterwards
Coupling is reduced to minimum.
Claims (2)
1. a kind of method of the decoupling Guidance and control of rotary missile, which is characterized in that include the following steps：
Step 1, a racemization bay section is added between the preceding bay section of rotary missile and rear bay section, racemization bay section includes shell, outside
Motor (3) and bearing (4) are housed inside shell (1), axis (5), the inside of axis (5) are housed inside motor (3) and bearing (4)
Equipped with slip ring (2), one end of axis (5) is equipped with the structure interface (6) of rear bay section；The effect of the slip ring be ensure before bay section with
The electrical connection of bay section afterwards；
Step 2, inertial measurement component is installed in preceding bay section, makes the accelerometer of inertial measurement component and the sensitive axes of gyro and rotation
Each axis of bay section is parallel before transduction bullet；
Step 3, racemization bay section is modeled and is emulated, obtain ratio, integral used in the PID controller of meet demand,
Differential parameter；
Step 4, the three axis angular rate data and acceleration information exported to inertial measurement component by control panel are acquired simultaneously
The roll angle angle value that guided missile is coupled to front deck end is calculated, specific calculating process is as follows：
A, autoregistration
Navigational coordinate system is established using IMU output informations, autoregistration algorithm is as follows：
Analytic coarse alignment is carried out using accelerometer and gyro output information, inertia device information is according to 50ms (i.e. 10 periods)
Smoothing processing to inhibit meter noise to influence,
φ=0
Wherein θ, γ, φ is respectively pitch angle, roll angle and yaw angle, ax, ay, and az indicates missile coordinate system coordinate system respectively
X-axis, y-axis, the acceleration of z-axis,
Relationship between quaternary number and attitude angle is：
B, strapdown resolves, and carries out attitude angle update,
Since the rotational angular velocity of missile coordinate system relative inertness coordinate system is：
WhereinFor gyro output angle speed,The projection for being rotational-angular velocity of the earth in navigational coordinate system, due to carrier
Initial heading is unknown, suddenly rotational-angular velocity of the earth, therefore：
According to the angular speed increment of previous moment quaternary number and current period, quaternary number is updated, the different tables of attitude matrix are utilized
Show mode, calculate current attitude angle, realizes solving of attitude,
Quaternary number recurrence formula：
Wherein
The relationship of attitude matrix and quaternary number is as follows：
Attitude matrix and the relationship of current pitching angle theta, roll angle γ and yaw angle φ are as follows：
The above attitude matrix is compared, current pitching angle theta, roll angle γ and yaw angle φ can be obtained
Step 5, it is controlled in racemization bay section using PID controller by the difference of bay section roll angle before calculating and setting roll angle
Motor, the rotating speed of bay section is maintained within the scope of setting value ± 2.0 ° before making so that by the coupling between yaw, pitching and roll channel
Conjunction is reduced to minimum.
2. a kind of device of the decoupling Guidance and control of rotary missile, which is characterized in that include shell (1), filled inside shell (1)
There are motor (3) and bearing (4), axis (5) is housed inside motor (3) and bearing (4), the inside of axis (5) is equipped with slip ring
(2), one end of axis (5) is equipped with the structure interface (6) of rear bay section.
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| CN201810779176.XA CN108759589A (en) | 2018-07-16 | 2018-07-16 | A kind of uncoupled method and device of rotary missile Guidance and control |
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| CN201810779176.XA CN108759589A (en) | 2018-07-16 | 2018-07-16 | A kind of uncoupled method and device of rotary missile Guidance and control |
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Cited By (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109579617A (en) * | 2018-12-21 | 2019-04-05 | 上海机电工程研究所 | Rolling control method, system and the medium of canard aerodynamic arrangement guided missile |
| CN109596011A (en) * | 2018-12-07 | 2019-04-09 | 上海机电工程研究所 | The stable canard configuration guided missile overall architecture of rolling racemization |
| CN109669480A (en) * | 2019-01-03 | 2019-04-23 | 西安航天动力技术研究所 | A kind of guiding head controlling method of future position |
| CN110645843A (en) * | 2019-08-16 | 2020-01-03 | 北京理工大学 | High-dynamic compensation guidance control system and method for high-speed maneuvering target |
| CN113503773A (en) * | 2021-06-28 | 2021-10-15 | 山西华洋吉禄科技股份有限公司 | Rotating holder for PGK and control method thereof |
| CN113970276A (en) * | 2021-11-16 | 2022-01-25 | 天津爱思达新材料科技有限公司 | High-strength connecting assembly for carbon fiber composite material |
| CN115755838A (en) * | 2022-11-08 | 2023-03-07 | 湖南航天有限责任公司 | Precision analysis method of missile guidance control system |
-
2018
- 2018-07-16 CN CN201810779176.XA patent/CN108759589A/en active Pending
Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109596011A (en) * | 2018-12-07 | 2019-04-09 | 上海机电工程研究所 | The stable canard configuration guided missile overall architecture of rolling racemization |
| CN109596011B (en) * | 2018-12-07 | 2020-08-04 | 上海机电工程研究所 | Rolling despinning stable duck-type layout missile overall framework |
| CN109579617A (en) * | 2018-12-21 | 2019-04-05 | 上海机电工程研究所 | Rolling control method, system and the medium of canard aerodynamic arrangement guided missile |
| CN109669480A (en) * | 2019-01-03 | 2019-04-23 | 西安航天动力技术研究所 | A kind of guiding head controlling method of future position |
| CN109669480B (en) * | 2019-01-03 | 2021-11-09 | 西安航天动力技术研究所 | Seeker control method for predicting target position |
| CN110645843A (en) * | 2019-08-16 | 2020-01-03 | 北京理工大学 | High-dynamic compensation guidance control system and method for high-speed maneuvering target |
| CN113503773A (en) * | 2021-06-28 | 2021-10-15 | 山西华洋吉禄科技股份有限公司 | Rotating holder for PGK and control method thereof |
| CN113970276A (en) * | 2021-11-16 | 2022-01-25 | 天津爱思达新材料科技有限公司 | High-strength connecting assembly for carbon fiber composite material |
| CN115755838A (en) * | 2022-11-08 | 2023-03-07 | 湖南航天有限责任公司 | Precision analysis method of missile guidance control system |
| CN115755838B (en) * | 2022-11-08 | 2024-05-28 | 湖南航天有限责任公司 | Precision analysis method of missile guidance control system |
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Application publication date: 20181106 |