CN107478110A - A kind of rotating missile attitude angle computational methods based on state observer - Google Patents

A kind of rotating missile attitude angle computational methods based on state observer Download PDF

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Publication number
CN107478110A
CN107478110A CN201710631100.8A CN201710631100A CN107478110A CN 107478110 A CN107478110 A CN 107478110A CN 201710631100 A CN201710631100 A CN 201710631100A CN 107478110 A CN107478110 A CN 107478110A
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angle
attitude angle
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rotating missile
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CN107478110B (en
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魏宗康
江麒
黄云龙
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China Aerospace Times Electronics Corp
Beijing Aerospace Control Instrument Institute
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China Aerospace Times Electronics Corp
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B35/00—Testing or checking of ammunition
    • F42B35/02—Gauging, sorting, trimming or shortening cartridges or missiles
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/20—Instruments for performing navigational calculations

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  • Engineering & Computer Science (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • General Engineering & Computer Science (AREA)
  • Automation & Control Theory (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
  • Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)
  • Gyroscopes (AREA)

Abstract

The invention discloses a kind of rotating missile attitude angle computational methods based on state observer.This method includes：Angular speed of the rotating missile in current time Relative Navigation coordinate system is obtained, and determines component of the angular speed on each axle of missile coordinate system；Obtain roll angle correction component, rolling angle measurement component, the angle of pitch and yaw angle of the rotating missile at current time；Utilize the angular velocity component, roll angle correction component, rolling angle measurement component, the angle of pitch and yaw angle, structure attitude angle renewal matrix；Matrix is updated using attitude angle, calculates feedback gain matrix；Matrix, feedback gain matrix, and rotating missile are updated in the attitude angle at current time, attitude angle of the calculating rotating missile in subsequent time according to attitude angle.The present invention can realize the aerial alignment in real time of high speed rotating missile, improve the attack precision of high speed rotating missile.

Description

A kind of rotating missile attitude angle computational methods based on state observer
Technical field
The present invention relates to a kind of high speed rotating missile real-time technique of alignment, more particularly to rotation based on state observer in the air Play attitude angle computational methods.
Background technology
High speed rotating missile is a kind of bullet arrow class aircraft rotated in flight course around its longitudinal axis continuous high speed, extensively Applied on all kinds of tactics and strategic missile.High speed rotating missile produces gyroscopic couple by high speed rotary motion, so as to weaken matter The adverse effect that the unfavorable conditions such as amount bias, thrust eccentric and pneumatic, structure asymmetry caused by processing are brought, ensure that winged The flight stability of row device, system is possessed higher antijamming capability, beneficial to the impact accuracy for improving armament systems, go back simultaneously The strike of Laser Interception weapon can effectively be weakened, improve its survival ability.
To improve its guidance precision, the Air launching of high speed rotating missile need to be realized, and defended more using inertial navigation and GPS Star guiding combination technology, INS errors are corrected by GPS, improve the ability of guided missile precision strike target.Wherein, it is used to Property navigation system be a voyage Estimation System based on acceleration quadratic integral, inertial navigation system work before, it is necessary to Complete initial alignment.During in airflight state, inertial navigation system position and speed can be provided by GPS system, but posture Angle can not individually be provided by GPS or inertial navigation system.
It is to resolve to obtain course angle and the pitching at corresponding moment using GPS information to carry out aerial self aligned effective way Angle, roll angle is estimated by suitable method.Conventional Initial Alignment Method uses Kalman filter, but is not suitable for turning The faster application conditions of speed.As the GPS exterior measuring cycles be 0.1 second, and the rotary speed of high speed rotating missile be more than 1500 °/s situation Under, then it can not meet that signal reproduction requires, cause evaluated error.Therefore, there is an urgent need to a kind of method can calculate rotating missile The situation of rotation at a high speed.
The content of the invention
Present invention solves the technical problem that it is：Compared to prior art, there is provided a kind of rotation based on state observer Attitude angle computational methods are played, the aerial alignment in real time of high speed rotating missile can be realized, improve the attack precision of high speed rotating missile.
The above-mentioned purpose of the present invention is achieved by the following technical programs：
The invention provides a kind of rotating missile attitude angle computational methods based on state observer, including：
Obtain angular speed of the rotating missile in current time Relative Navigation coordinate systemAnd determine the angular speedIn bullet Component on each axle of body coordinate systemWith
The roll angle that rotating missile is obtained at current time corrects componentRolling angle measurement component α0, pitching angle theta0And driftage Angle ψ0, and record attitude angle of the rotating missile at current time and be
Utilize the componentWithAnd the roll angle correction componentRolling angle measurement component α0、 Pitching angle theta0And yaw angle ψ0, structure attitude angle renewal matrix A；
Matrix A is updated using the attitude angle, calculates feedback gain matrix K；
Matrix A, the feedback gain matrix K and the attitude angle are updated according to the attitude angleCalculate rotation Turn attitude angle of the bullet in subsequent time.
Further, the attitude angle renewal matrix A is：
In formula, aij, i=1,2,3, j=1,2,3 are the element that attitude angle updates matrix A.
Further, the feedback gain matrix K is：
Further, matrix A, the feedback gain matrix K and the attitude angle are updated according to the attitude angleAttitude angle of the rotating missile in subsequent time is calculated, including：
In formula, x1Attitude angle of the rotating missile in subsequent time is represented,
The present invention has the advantages that compared with prior art：
The invention provides a kind of rotating missile attitude angle computational methods based on state observer, pass through satellite navigation system Assistance data and inertial navigation system gyroscope Output speed, the attitude angle of high speed rotating missile is resolved, solves height Fast rotating missile is in the air under maneuvering condition, it is impossible to individually carries out self aligned problem by inertial navigation system, can realize height The aerial alignment in real time of fast rotating missile, and improve the attack precision of high speed rotating missile.
Brief description of the drawings
Fig. 1 is a kind of flow of rotating missile attitude angle computational methods based on state observer in the embodiment of the present invention Figure；
Fig. 2 is the error convergence curve of each attitude angle in the embodiment of the present invention；
Fig. 3 is roll angle convergence curve corresponding to rolling initial deviation difference in the embodiment of the present invention.
Embodiment
The present invention is described in further detail with reference to the accompanying drawings and examples.It is it is understood that described herein Specific embodiment be used only for explaining the present invention, rather than limitation of the invention.It also should be noted that for the ease of Describe, part related to the present invention rather than entire infrastructure are illustrate only in accompanying drawing.
Fig. 1 is a kind of flow of rotating missile attitude angle computational methods based on state observer in the embodiment of the present invention Figure, with reference to figure 1, the rotating missile attitude angle computational methods provided by the invention based on state observer can specifically include following step Suddenly：
S110, obtain angular speed of the rotating missile in current time Relative Navigation coordinate systemAnd determine the angular speedComponent on each axle of missile coordinate systemWith
Specifically, navigational coordinate system (g) local geographic coordinate system corresponding with launch point overlaps, its origin is launch point, Three reference axis point to geographical east, north, day direction；Missile coordinate system (b) is connected with carrier, and origin is body barycenter；XbAxle is The carrier longitudinal axis, point to immediately ahead of carrier；ZbAxle position is in the longitudinally asymmetric face of carrier, with XbAxle is vertical and points to；YbAxle is according to the right side Hand rule determines.Thus, angular speed of the rotating missile in current time Relative Navigation coordinate system is obtainedAnd determine the angle speed DegreeComponent on each axle of missile coordinate systemWith
S120, obtain roll angle correction component of the rotating missile at current timeRolling angle measurement component α0, pitching angle theta0 And yaw angle ψ0, and record attitude angle of the rotating missile at current time and be
Specifically, using satellite navigation system, the roll angle for obtaining rotating missile at current time corrects componentRoll angle Measure component α0, pitching angle theta0And yaw angle ψ0, and record attitude angle of the rotating missile at current time and be
S130, utilize the componentWithAnd the roll angle correction componentRolling angle measurement Component α0, pitching angle theta0And yaw angle ψ0, structure attitude angle renewal matrix A.
Optionally, the attitude angle renewal matrix A is：
In formula, aij, i=1,2,3, j=1,2,3 are the element that attitude angle updates matrix A.
S140, matrix A, calculating feedback gain matrix K are updated using the attitude angle.
Optionally, the feedback gain matrix K is：
S150, matrix A, the feedback gain matrix K and the attitude angle updated according to the attitude angle Calculate attitude angle of the rotating missile in subsequent time.
Optionally, matrix A, the feedback gain matrix K and the attitude angle are updated according to the attitude angleAttitude angle of the rotating missile in subsequent time is calculated, including：
In formula, x1Attitude angle of the rotating missile in subsequent time is represented,
Using the rotating missile attitude angle computational methods provided by the invention based on state observer, alignment resolving is carried out, is solved Calculate result as shown in Figures 2 and 3.
Include roll angle error delta γ, angle of pitch error delta θ and yaw angle error delta ψ convergence curve in fig. 2.Can from Fig. 2 To find out, each attitude angle can finally converge to true value.
Fig. 3 is roll angle γ initial deviation δ γ0 Respectively 0 °, -10 °, 10 °, -20 ° and At 20 °, the change during error delta γ Air launchings, it can be seen that roll angle γ gradually converges on true value, complete The Air launching of high speed rotating missile, fast convergence rate and convergence error is small.
The technical scheme of the present embodiment provides a kind of rotating missile attitude angle computational methods based on state observer, passes through The assistance data of satellite navigation system and the gyroscope Output speed of inertial navigation system, are solved to the attitude angle of high speed rotating missile Calculate, solve high speed rotating missile in the air under maneuvering condition, it is impossible to self aligned problem is individually carried out by inertial navigation system, The aerial alignment in real time of high speed rotating missile can be realized, and improves the attack precision of high speed rotating missile；Algorithm is simple, convergence rate It hurry up, while be applied to carrier low-speed running situation.
Pay attention to, above are only presently preferred embodiments of the present invention and institute's application technology principle.It will be appreciated by those skilled in the art that The invention is not restricted to specific embodiment described here, can carry out for a person skilled in the art various obvious changes, Readjust and substitute without departing from protection scope of the present invention.Therefore, although being carried out by above example to the present invention It is described in further detail, but the present invention is not limited only to above example, without departing from the inventive concept, also Other more equivalent embodiments can be included, and the scope of the present invention is determined by scope of the appended claims.

Claims (4)

  1. A kind of 1. rotating missile attitude angle computational methods based on state observer, it is characterised in that including：
    Obtain angular speed of the rotating missile in current time Relative Navigation coordinate systemAnd determine the angular speedSat in body Component on each axle of mark systemWith
    The roll angle that rotating missile is obtained at current time corrects componentRolling angle measurement component α0, pitching angle theta0And yaw angle ψ0, and record attitude angle of the rotating missile at current time and be
    Utilize the componentWithAnd the roll angle correction componentRolling angle measurement component α0, pitching Angle θ0And yaw angle ψ0, structure attitude angle renewal matrix A；
    Matrix A is updated using the attitude angle, calculates feedback gain matrix K；
    Matrix A, the feedback gain matrix K and the attitude angle are updated according to the attitude angleCalculate rotating missile In the attitude angle of subsequent time.
  2. 2. rotating missile attitude angle computational methods according to claim 1, it is characterised in that the attitude angle updates matrix A For：
    In formula, aij, i=1,2,3, j=1,2,3 are the element that attitude angle updates matrix A.
  3. 3. rotating missile attitude angle computational methods according to claim 2, it is characterised in that the feedback gain matrix K is：
    <mrow> <mi>K</mi> <mo>=</mo> <mfenced open = "[" close = "]"> <mtable> <mtr> <mtd> <msub> <mi>a</mi> <mn>12</mn> </msub> </mtd> <mtd> <mrow> <msup> <mrow> <mo>(</mo> <msub> <mi>a</mi> <mn>12</mn> </msub> <mo>+</mo> <msub> <mi>a</mi> <mn>11</mn> </msub> <mo>)</mo> </mrow> <mn>2</mn> </msup> <mo>/</mo> <msub> <mi>a</mi> <mn>31</mn> </msub> </mrow> </mtd> </mtr> <mtr> <mtd> <msub> <mi>a</mi> <mn>12</mn> </msub> </mtd> <mtd> <mn>0</mn> </mtd> </mtr> <mtr> <mtd> <msub> <mi>a</mi> <mn>32</mn> </msub> </mtd> <mtd> <mrow> <mn>2</mn> <msub> <mi>a</mi> <mn>12</mn> </msub> <mo>+</mo> <msub> <mi>a</mi> <mn>11</mn> </msub> </mrow> </mtd> </mtr> </mtable> </mfenced> </mrow>
  4. 4. rotating missile attitude angle computational methods according to claim 3, it is characterised in that square is updated according to the attitude angle Battle array A, the feedback gain matrix K and the attitude angleAttitude angle of the rotating missile in subsequent time is calculated, including：
    In formula, x1Attitude angle of the rotating missile in subsequent time is represented,
CN201710631100.8A 2017-07-28 2017-07-28 Rotating elastic attitude angle calculation method based on state observer Active CN107478110B (en)

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Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN109341717A (en) * 2018-09-13 2019-02-15 红色江山(湖北)导航技术有限公司 MEMS inertial navigation system horizontal attitude based on maneuvering condition judgement reviews one's lessons by oneself correction method
CN111780749A (en) * 2020-05-26 2020-10-16 北京航天控制仪器研究所 Attitude control method for full-attitude inertial navigation of orbital transfer maneuvering aircraft
CN112363195A (en) * 2020-09-30 2021-02-12 东南大学 Rotary missile air rapid coarse alignment method based on kinematic equation
CN113447024A (en) * 2021-06-28 2021-09-28 北京航天控制仪器研究所 Inertial navigation attitude angle resolving method and system based on extended Krafft angle
CN114894047A (en) * 2021-09-17 2022-08-12 中国人民解放军63875部队 Station distribution analysis method for optimizing target range key section attitude rendezvous measurement

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JPH042596A (en) * 1990-04-17 1992-01-07 Mitsubishi Electric Corp Flying body performance evaluating method and device therefor
CN105987695A (en) * 2015-01-29 2016-10-05 中北大学 Interval quartering lagrangian method used for high-speed spinning projectile attitude algorithm
CN106840195A (en) * 2016-12-19 2017-06-13 中北大学 A kind of rotary half strapdown micro-inertial measuring system error inhibition method

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPH042596A (en) * 1990-04-17 1992-01-07 Mitsubishi Electric Corp Flying body performance evaluating method and device therefor
CN105987695A (en) * 2015-01-29 2016-10-05 中北大学 Interval quartering lagrangian method used for high-speed spinning projectile attitude algorithm
CN106840195A (en) * 2016-12-19 2017-06-13 中北大学 A kind of rotary half strapdown micro-inertial measuring system error inhibition method

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN109341717A (en) * 2018-09-13 2019-02-15 红色江山(湖北)导航技术有限公司 MEMS inertial navigation system horizontal attitude based on maneuvering condition judgement reviews one's lessons by oneself correction method
CN111780749A (en) * 2020-05-26 2020-10-16 北京航天控制仪器研究所 Attitude control method for full-attitude inertial navigation of orbital transfer maneuvering aircraft
CN112363195A (en) * 2020-09-30 2021-02-12 东南大学 Rotary missile air rapid coarse alignment method based on kinematic equation
CN113447024A (en) * 2021-06-28 2021-09-28 北京航天控制仪器研究所 Inertial navigation attitude angle resolving method and system based on extended Krafft angle
CN113447024B (en) * 2021-06-28 2022-07-05 北京航天控制仪器研究所 Inertial navigation attitude angle resolving method and system based on extended Krafft angle
CN114894047A (en) * 2021-09-17 2022-08-12 中国人民解放军63875部队 Station distribution analysis method for optimizing target range key section attitude rendezvous measurement
CN114894047B (en) * 2021-09-17 2023-10-03 中国人民解放军63875部队 Optimization station distribution analysis method for intersection measurement of key section gestures of target range

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