CN106595649B - A kind of in-flight inertia initial baseline deviation compensation method - Google Patents
A kind of in-flight inertia initial baseline deviation compensation method Download PDFInfo
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
- G01C25/00—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass
- G01C25/005—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass initial alignment, calibration or starting-up of inertial devices
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- 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
- G01C21/10—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration
- G01C21/12—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning
- G01C21/16—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation
- G01C21/165—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation combined with non-inertial navigation instruments
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Abstract
The present invention provides a kind of in-flight inertia initial baseline deviation compensation methods, the inertia initial baseline misalignment that this method is obtained using estimation, the in-flight pitch angle deviation at current time, yaw angular displacement and roll angle deviation are calculated, attitude misalignment is compensated using linear rule in conjunction with attitude control capability for correcting.Through mathematical simulation technology and the verifying of carrying flight test, in-flight inertia initial baseline deviation compensation method can effectively compensate for inertial reference deviation, achieve the purpose that the precision and reliability that improve navigation system.
Description
Technical field
The present invention relates to Guidance & Navigation technical field, in particular to a kind of in-flight inertia initial baseline deviation compensation side
Method.
Background technique
Certain guided missiles use direct laying of fire or Transfer Alignment, can generate biggish initial baseline deviation, can cause to be used in this way
Property navigation error accelerated accumulation, it is therefore necessary to take measures to compensate initial baseline deviation, with inhibit inertial navigation error dissipate
Speed.
Summary of the invention
It is an object of the invention to overcome the deficiencies of the prior art and provide a kind of, and the in-flight inertia based on GPS is initial
Datum drift estimation method, this method calculate in-flight current time using the obtained inertia initial baseline misalignment of estimation
Pitch angle deviation, yaw angular displacement and roll angle deviation carry out attitude misalignment using linear rule in conjunction with attitude control capability for correcting
Compensation.
Above-mentioned purpose of the invention is realized by following scheme:
A kind of in-flight inertia initial baseline deviation compensation method, comprising the following steps:
(1), the inertia initial baseline misalignment obtained according to estimation calculates the in-flight pitch angle deviation of guided missile
It yaws angular displacement δ ψ and roll angle deviation δ γ, specific formula for calculation is as follows:
Wherein:γ is respectively pitch angle, the roll angle that inertial navigation system provides；φx0、φy0、φz0Respectively estimate
X of the guided missile arrived under inertial coodinate system is to initial baseline misalignment, Y-direction initial baseline misalignment, Z-direction initial baseline misalignment
Angle；
(2), within the amendment period of setting, missile attitude deviation serial update N is clapped, wherein the n-th makeover process clapped
It is as follows:
(2a), revised orientation vector is calculated:
Wherein:ψn′、γn' it is respectively pitch angle, yaw angle and roll angle after the n-th bat amendment；ψn、γnRespectively
Inertial navigation system claps output pitch angle, yaw angle and roll angle n-th；δαzFor setting
Pitch orientation attitude control maximum modified value, δ αyFor the yaw direction attitude control maximum modified value of setting, δ αxFor the rotating direction of setting
Attitude control maximum modified value；
(2b), revised orientation vector is clapped by n-thIt exports to inertial navigation system, inertial navigation system is sweared using the posture
AmountAfter carrying out recurrence calculation, the pitch angle of the (n+1)th bat of outputYaw angle ψn+1With roll angle γn+1；
Wherein, n=1,2 ..., N.
Above-mentioned in-flight inertia initial baseline deviation compensation method, in step (1), estimation is obtained as follows
Inertia initial baseline misalignment:
(1a), the state vector predicted value for calculating moment k Estimate for the state vector of moment k-1
Value, the initial value of the state vector estimated valueIt is set as null vector；Wherein, in Kalman filtering, definition status vector
X=[φx0 φy0 φz0 δX0 δY0 δZ0 δVx0 δVy0 δVz0]T；φx0、φy0、φz0Respectively guided missile is in inertial coodinate system
Under X to initial baseline misalignment, Y-direction initial baseline misalignment, Z-direction initial baseline misalignment；δX0、δY0、δZ0Respectively lead
X of the bullet under inertial coodinate system is to initial position error, Y-direction initial position error, Z-direction initial position error；δVx0、δVy0、δ
Vz0Respectively X of the guided missile under inertial coodinate system is to initial velocity error, Y-direction initial velocity error, Z-direction initial velocity error；
(1b), the prediction covariance matrix P for calculating moment kk|k-1=Pk-1；Pk-1For the estimate covariance matrix of moment k-1,
The initial value of the estimate covariance matrix is P0, P09 × 9 for setting tie up matrix；
(1c), the gain matrix K for calculating moment kk=Pk|k-1HT(H Pk|k-1HT+R)-1；Wherein, H is the observation square of setting
Battle array, R are the noise variance matrix of setting:
(1d), the state vector estimated value for calculating moment kWherein, ZkIt is outer
The observation vector of k at the time of portion provides；In Kalman filtering, observation vector Z=[δ X δ Y δ Z δ V is definedx δVy δVz]T；
δ X, δ Y, δ Z are respectively the X of inertial navigation system and GPS system measurement to the difference of position, the difference of Y-direction position, the difference of Z-direction position；δVx、
δVy、δVzRespectively inertial navigation system and the X of GPS system measurement is to the difference of speed, the difference of Y-direction speed, the difference of Z-direction speed；
(1e), moment k estimate covariance matrix P is calculatedk=(I-KkH)Pk|k-1, I is unit matrix；
(1f), in the initial value datum drift cycle estimator of setting, step (1a)~(1e) is repeated, Kalman filtering is passed through
Initial position error, initial velocity error and initial baseline misalignment is calculated.
Above-mentioned in-flight inertia initial baseline deviation compensation method, in step (1c), observing matrix H and noise variance
Matrix R is set separately are as follows:
Wherein: Sx、Sy、SzRespectively inertial navigation system provide inertial system X to, Y-direction, Z-direction regard position；Wx、Wy、WzRespectively
Inertial navigation system provide inertial system X to, Y-direction, Z-direction apparent velocity； Respectively set inertial system under X to, Y-direction, Z
To position detection noise variance；Under the inertial system respectively set X to, Y-direction, Z-direction speed observation noise
Variance.
Compared with prior art, the present invention having the advantage that
Through mathematical simulation technology and the verifying of carrying flight test, in-flight inertia initial baseline deviation compensation method can have
Effect ground compensation inertial reference deviation, achievees the purpose that the precision and reliability that improve navigation system.
Detailed description of the invention
Fig. 1 is in-flight inertia initial baseline deviation compensation method flow chart of the invention.
Specific embodiment
The present invention is described in further detail with specific example with reference to the accompanying drawing:
A kind of in-flight inertia initial baseline deviation compensation method of the invention, the inertia initial baseline obtained using estimation
Misalignment calculates the in-flight pitch angle deviation at current time, yaw angular displacement and roll angle deviation, in conjunction with attitude control capability for correcting
Attitude misalignment is compensated using linear rule.As shown in Figure 1, of the invention, the specific implementation steps are as follows:
(1), the inertia initial baseline misalignment obtained according to estimation calculates the in-flight pitch angle deviation of guided missile
It yaws angular displacement δ ψ and roll angle deviation δ γ, specific formula for calculation is as follows:
Wherein:γ is respectively pitch angle, the roll angle that inertial navigation system provides；φx0、φy0、φz0Respectively estimate
X of the guided missile arrived under inertial coodinate system is to initial baseline misalignment, Y-direction initial baseline misalignment, Z-direction initial baseline misalignment
Angle；
(2), within the amendment period, missile attitude deviation serial update N is clapped, wherein the makeover process of the n-th bat is as follows:
(2a), revised orientation vector is calculated:
Wherein:ψn′、γn' it is respectively pitch angle, yaw angle and roll angle after the n-th bat amendment；ψn、γnRespectively
Inertial navigation system claps output pitch angle, yaw angle and roll angle n-th；δαzFor setting
Pitch orientation attitude control maximum modified value, δ αyFor the yaw direction attitude control maximum modified value of setting, δ αxFor the rotating direction of setting
Attitude control maximum modified value, these three values are set according to the control ability of servo-control system；Max () is to seek maximum value
Operator, int () are round numbers operator；
(2b), revised orientation vector is clapped by n-thIt exports to inertial navigation system, inertial navigation system is sweared using the posture
AmountAfter carrying out recurrence calculation, the pitch angle of the (n+1)th bat of outputYaw angle ψn+1With roll angle γn+1；
Wherein, n=1,2 ..., N.
During above compensation calculation, it is inclined inertia initial baseline can be established according to inertial navigation error propagation mechanism
Differential mode type estimates inertia initial baseline misalignment using standard Kalman filtering method using GPS information as observed quantity.
Firstly, establishing initial baseline estimation of deviation state equation:
Wherein: state vector X=[φx0 φy0 φz0 δX0 δY0 δZ0 δVx0 δVy0 δVz0]T；φx0、φy0、φz0
Respectively X of the guided missile under inertial coodinate system is to initial baseline misalignment, Y-direction initial baseline misalignment, Z-direction initial baseline misalignment
Angle；δX0、δY0、δZ0Respectively guided missile is at the beginning of the X under inertial coodinate system to initial position error, Y-direction initial position error, Z-direction
Beginning location error；δVx0、δVy0、δVz0Respectively X of the guided missile under inertial coodinate system is to initial velocity error, Y-direction initial velocity
Error, Z-direction initial velocity error；Sytem matrix F is 9 × 9 dimension null matrix.
Then, the observational equation of inertia initial baseline estimation of deviation is established:
Z=HX+V；
Wherein: observation vector Z=[δ X δ Y δ Z δ Vx δVy δVz]T；δ X, δ Y, δ Z are respectively inertial navigation system and GPS system
The X of unified test amount is to the difference of position, the difference of Y-direction position, the difference of Z-direction position；δVx、δVy、δVzRespectively inertial navigation system and GPS system
The X of measurement is to the difference of speed, the difference of Y-direction speed, the difference of Z-direction speed；V is observation noise vector；H is observing matrix, in this hair
It is specifically set in bright are as follows:
Wherein: Sx、Sy、SzRespectively inertial navigation system provide inertial system X to, Y-direction, Z-direction regard position；Wx、Wy、WzRespectively
Inertial navigation system provide inertial system X to, Y-direction, Z-direction apparent velocity.
Based on above Kalman filter equation, the estimation procedure of inertia initial baseline misalignment is as follows:
(1), in moment k, state vector predicted value is calculated Estimate for the state vector of moment k-1
Value, the initial value of the state vector estimated valueIt is set as null vector；Wherein, in Kalman filtering, definition status vector
X=[φx0 φy0 φz0 δX0 δY0 δZ0 δVx0 δVy0 δVz0]T；φx0、φy0、φz0Respectively guided missile is in inertial coodinate system
Under X to initial baseline misalignment, Y-direction initial baseline misalignment, Z-direction initial baseline misalignment；δX0、δY0、δZ0Respectively lead
X of the bullet under inertial coodinate system is to initial position error, Y-direction initial position error, Z-direction initial position error；δVx0、δVy0、δ
Vz0Respectively X of the guided missile under inertial coodinate system is to initial velocity error, Y-direction initial velocity error, Z-direction initial velocity error；
(2), in moment k, covariance matrix P is predictedk|k-1=Pk-1；Pk-1It is described for the estimate covariance matrix of moment k-1
The initial value of estimate covariance matrix is P0, P09 × 9 for setting tie up matrix；
(3), in moment k, gain matrix K is calculatedk=Pk|k-1HT(H Pk|k-1HT+R)-1；Wherein, H is the observation square of setting
Battle array, R are the noise variance matrix of setting:
Wherein: Sx、Sy、SzRespectively inertial navigation system provide inertial system X to, Y-direction, Z-direction regard position；Wx、Wy、WzRespectively
Inertial navigation system provide inertial system X to, Y-direction, Z-direction apparent velocity； Respectively set inertial system under X to, Y-direction, Z
To position detection noise variance；Under the inertial system respectively set X to, Y-direction, Z-direction speed observation noise
Variance；
(4), the state vector estimated value of moment k is calculatedWherein, ZkFor outside
The observation vector of k at the time of offer；In Kalman filtering, observation vector Z=[δ X δ Y δ Z δ V is definedx δVy δVz]T；δ
X, δ Y, δ Z are respectively the X of inertial navigation system and GPS system measurement to the difference of position, the difference of Y-direction position, the difference of Z-direction position；δVx、δ
Vy、δVzRespectively inertial navigation system and the X of GPS system measurement is to the difference of speed, the difference of Y-direction speed, the difference of Z-direction speed；
(5), moment k estimate covariance matrix P is calculatedk=(I-KkH)Pk|k-1, I is unit matrix；
(6), in the initial value datum drift cycle estimator of setting, step (1)~(5) is repeated, Kalman filtering meter is passed through
Calculation obtains initial baseline misalignment.
When concrete engineering is realized, initial baseline misalignment estimation method of the invention and initial baseline deviation compensation method
By the software realization being solidificated in missile-borne computer, during missile flight, Flight Control Software calls the method, to guided missile
Initial inertia datum drift is estimated and compensated, to improve the navigation accuracy and accuracy at target of guided missile.
The above, a specific embodiment only of the invention, but scope of protection of the present invention is not limited thereto, appoints
In the technical scope disclosed by the present invention, any changes or substitutions that can be easily thought of, all by what those familiar with the art
It is covered by the protection scope of the present invention.
The content that description in the present invention is not described in detail belongs to the well-known technique of professional and technical personnel in the field.
Claims (3)
1. a kind of in-flight inertia initial baseline deviation compensation method, it is characterised in that the following steps are included:
(1), the inertia initial baseline misalignment obtained according to estimation calculates the in-flight pitch angle deviation of guided missileYaw
Angular displacement δ ψ and roll angle deviation δ γ, specific formula for calculation are as follows:
Wherein:γ is respectively pitch angle, the roll angle that inertial navigation system provides；φx0、φy0、φz0What respectively estimation obtained leads
X of the bullet under inertial coodinate system is to initial baseline misalignment, Y-direction initial baseline misalignment, Z-direction initial baseline misalignment；
(2), within the amendment period of setting, missile attitude deviation serial update N is clapped, wherein the n-th makeover process clapped is such as
Under:
(2a), revised orientation vector is calculated:
Wherein:ψ′n、γ′nRespectively n-th claps pitch angle, yaw angle and roll angle after amendment；ψn、γnRespectively inertial navigation
System claps output pitch angle, yaw angle and roll angle n-th；δαzFor bowing for setting
Face upward direction attitude control maximum modified value, δ αyFor the yaw direction attitude control maximum modified value of setting, δ αxFor the rotating direction attitude control of setting
Maximum modified value；
(2b), revised orientation vector is clapped by n-thIt exports to inertial navigation system, inertial navigation system utilizes the orientation vectorAfter carrying out recurrence calculation, the pitch angle of the (n+1)th bat of outputYaw angle ψn+1With roll angle γn+1；
Wherein, n=1,2 ..., N.
2. a kind of in-flight inertia initial baseline deviation compensation method according to claim 1, it is characterised in that: in step
(1) in, estimation obtains inertia initial baseline misalignment as follows:
(1a), the state vector predicted value for calculating moment k For the state vector estimated value of moment k-1, institute
State the initial value of state vector estimated valueIt is set as null vector；Wherein, in Kalman filtering, definition status vector X=
[φx0 φy0 φz0 δX0 δY0 δZ0 δVx0 δVy0 δVz0]T；φx0、φy0、φz0Respectively guided missile is under inertial coodinate system
X to initial baseline misalignment, Y-direction initial baseline misalignment, Z-direction initial baseline misalignment；δX0、δY0、δZ0Respectively guided missile
X under inertial coodinate system is to initial position error, Y-direction initial position error, Z-direction initial position error；δVx0、δVy0、δVz0
Respectively X of the guided missile under inertial coodinate system is to initial velocity error, Y-direction initial velocity error, Z-direction initial velocity error；
(1b), the prediction covariance matrix P for calculating moment kk|k-1=Pk-1；Pk-1It is described for the estimate covariance matrix of moment k-1
The initial value of estimate covariance matrix is P0, P09 × 9 for setting tie up matrix；
(1c), the gain matrix K for calculating moment kk=Pk|k-1HT(HPk|k-1HT+R)-1；Wherein, H is the observing matrix of setting, and R is
The noise variance matrix of setting:
(1d), the state vector estimated value for calculating moment kWherein, ZkIt is provided for outside
At the time of k observation vector；In Kalman filtering, observation vector Z=[δ X δ Y δ Z δ V is definedx δVy δVz]T；δX,δY,
δ Z is respectively the X of inertial navigation system and GPS system measurement to the difference of position, the difference of Y-direction position, the difference of Z-direction position；δVx、δVy、δVz
Respectively inertial navigation system and the X of GPS system measurement is to the difference of speed, the difference of Y-direction speed, the difference of Z-direction speed；
(1e), moment k estimate covariance matrix P is calculatedk=(I-KkH)Pk|k-1, I is unit matrix；
(1f), in the initial value datum drift cycle estimator of setting, step (1a)~(1e) is repeated, is calculated by Kalman filtering
Obtain initial position error, initial velocity error and initial baseline misalignment.
3. a kind of in-flight inertia initial baseline deviation compensation method according to claim 2, it is characterised in that: in step
In (1c), observing matrix H and noise variance matrix R are set separately are as follows:
Wherein: Sx、Sy、SzRespectively inertial navigation system provide inertial system X to, Y-direction, Z-direction regard position；Wx、Wy、WzRespectively inertial navigation
System provide inertial system X to, Y-direction, Z-direction apparent velocity； Respectively set inertial system under X to, Y-direction, Z-direction position
Set observation noise variance；Under the inertial system respectively set X to, Y-direction, Z-direction speed observation noise variance.
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| CN110186481B (en) * | 2019-06-10 | 2023-03-28 | 西安航天三沃机电设备有限责任公司 | Calibration system and calibration method suitable for inertial component on small-sized guided missile |
| CN115077520B (en) * | 2022-08-22 | 2022-11-01 | 中国船舶重工集团公司第七0七研究所 | Attitude compensation method based on resonant inertial navigation system |
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