CN106595649A - Method for inertia initial datum deviation compensation in flight - Google Patents
Method for inertia initial datum deviation compensation in flight Download PDFInfo
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- CN106595649A CN106595649A CN201611048884.3A CN201611048884A CN106595649A CN 106595649 A CN106595649 A CN 106595649A CN 201611048884 A CN201611048884 A CN 201611048884A CN 106595649 A CN106595649 A CN 106595649A
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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 invention provides a method for inertia initial datum deviation compensation in flight. According to the method, a pitch angle deviation, a yaw angle deviation and a roll angle deviation of a current moment in flight are calculated by means of an inertia initial datum misalignment angle obtained through estimation, attitude-control correction capacity is used, and a posture deviation is compensated for through a linear rule. The mathematical simulation technology and carrying flight tests prove that the method for inertia initial datum deviation compensation in flight can effectively compensate for an inertia initial datum deviation, and the aim of improving precision and reliability of a navigation system is achieved.
Description
Technical field
The present invention relates to Guidance ＆ Navigation technical field, more particularly to a kind of in-flight inertia initial baseline deviation compensation side
Method.
Background technology
Some guided missiles adopt direct laying of fire or Transfer Alignment, can produce larger initial baseline deviation, can so cause used
Property navigation error accelerated accumulation, it is therefore necessary to take measures compensate initial baseline deviation, with suppress inertial navigation error dissipate
Speed.
The content of the invention
It is an object of the invention to overcome the deficiencies in the prior art, there is provided a kind of in-flight inertia based on GPS is initial
Datum drift method of estimation, the method calculate in-flight current time using the inertia initial baseline misalignment that obtains of estimation
Pitch angle deviation, driftage angular displacement and roll angle deviation, are carried out to attitude misalignment using linear rule with reference to attitude control capability for correcting
Compensation.
The above-mentioned purpose of the present invention is realized by below scheme：
A kind of in-flight inertia initial baseline deviation compensation method, comprises the following steps：
(1), the inertia initial baseline misalignment obtained according to estimation, calculates the in-flight pitch angle deviation of guided missile
Driftage angular displacement δ ψ and roll angle deviation δ γ, specific formula for calculation are as follows：
Wherein：γ is respectively the angle of pitch, the roll angle that inertial navigation system is provided；φx0、φy0、φz0Respectively estimate to obtain
X of the guided missile under inertial coodinate system to initial baseline misalignment, Y-direction initial baseline misalignment, Z-direction initial baseline misalignment；
(2), setting the amendment cycle in, to missile attitude deviation serial update N clap, wherein, n-th clap makeover process
It is as follows：
(2a), calculate revised orientation vector：
Wherein：ψn′、γn' it is respectively the angle of pitch, yaw angle and roll angle after the n-th bat amendment；ψn、γnRespectively
Inertial navigation system claps the output angle of pitch, 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-thExport to inertial navigation system, inertial navigation system is sweared using the attitude
AmountAfter carrying out recurrence calculation, the angle of pitch that output (n+1)th is clappedYaw 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), estimates to obtain as follows
Inertia initial baseline misalignment：
(1a), calculate the state vector predictive value of moment k State vector for moment k-1 is estimated
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), calculate the prediction covariance matrix P of 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, P0For 9 × 9 dimension matrixes of setting；
(1c), calculate the gain matrix K of 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), calculate the state vector estimated value of moment kWherein, ZkFor outer
The observation vector of the moment k that portion provides；In Kalman filtering, observation vector Z=[δ X δ Y δ Z δ V are definedx δVy δVz]T；
δ X, δ Y, δ Z be respectively X that inertial navigation system and GPS system measure to the difference of position, the difference of Y-direction position, Z-direction position difference；δVx、
δVy、δVzThe respectively X of inertial navigation system and GPS system measurement to the difference of speed, the difference of Y-direction speed, Z-direction speed difference；
(1e), calculate moment k estimate covariance matrix Pk=(I-KkH)Pk|k-1, I is unit matrix；
(1f), in the initial value datum drift cycle estimator of setting, repeat step (1a)～(1e), by Kalman filtering
It is calculated initial position error, initial velocity error and initial baseline misalignment.
Above-mentioned in-flight inertia initial baseline deviation compensation method, in step (1c), observing matrix H and noise variance
Matrix R is respectively set as：
Wherein：Sx、Sy、SzRespectively inertial navigation system provide inertial system X to, Y-direction, Z-direction apparent place；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 for respectively setting, X is to, Y-direction, Z-direction speed observation noise side
Difference.
The present invention compared with prior art, with advantages below：
In-flight inertia initial baseline deviation compensation method can have for Jing mathematical simulations technology and the checking of carrying flight test
Effect ground compensation inertial reference deviation, reaches the purpose of the precision and reliability that improve navigation system.
Description of the drawings
Fig. 1 is the in-flight inertia initial baseline deviation compensation method flow chart of the present invention.
Specific embodiment
The present invention is described in further detail with instantiation below in conjunction with the accompanying drawings：
One kind in-flight inertia initial baseline deviation compensation method of the present invention, the inertia initial baseline obtained using estimation
Misalignment, calculates the in-flight pitch angle deviation at current time, driftage angular displacement and roll angle deviation, with reference to attitude control capability for correcting
Attitude misalignment is compensated using linear rule.As shown in figure 1, the present invention to implement step as follows：
(1), the inertia initial baseline misalignment obtained according to estimation, calculates the in-flight pitch angle deviation of guided missile
Driftage angular displacement δ ψ and roll angle deviation δ γ, specific formula for calculation are as follows：
Wherein：γ is respectively the angle of pitch, the roll angle that inertial navigation system is provided；φx0、φy0、φz0Respectively estimate to obtain
X of the guided missile under inertial coodinate system to initial baseline misalignment, Y-direction initial baseline misalignment, Z-direction initial baseline misalignment；
(2), in repairing positive period, missile attitude deviation serial update N is clapped, wherein, the n-th makeover process clapped is as follows：
(2a), calculate revised orientation vector：
Wherein：ψn′、γn' it is respectively the angle of pitch, yaw angle and roll angle after the n-th bat amendment；ψn、γnRespectively
Inertial navigation system claps the output angle of pitch, 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 ask for maximum
Operator, int () are round numbers operator；
(2b), revised orientation vector is clapped by n-thExport to inertial navigation system, inertial navigation system is sweared using the attitude
AmountAfter carrying out recurrence calculation, the angle of pitch that output (n+1)th is clappedYaw angle ψn+1With roll angle γn+1；
Wherein, n=1,2 ..., N.
During compensation calculation more than, inertia initial baseline can be set up according to inertial navigation error propagation mechanism inclined
Differential mode type, by the use of GPS information as observed quantity, estimates inertia initial baseline misalignment using standard Kalman filtering method.
First, set up 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、δZ0At the beginning of respectively X of the guided missile under inertial coodinate system to initial position error, Y-direction initial position error, Z-direction
Beginning site 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, set up the observational equation of inertia initial baseline estimation of deviation：
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 systems
The X of unified test amount to the difference of position, the difference of Y-direction position, Z-direction position difference；δVx、δVy、δVzRespectively inertial navigation system and GPS system
The X of measurement to the difference of speed, the difference of Y-direction speed, Z-direction speed difference；V is observation noise vector；H is observing matrix, at this
Specifically it is set as in bright：
Wherein：Sx、Sy、SzRespectively inertial navigation system provide inertial system X to, Y-direction, Z-direction apparent place；Wx、Wy、WzRespectively
Inertial navigation system provide inertial system X to, Y-direction, Z-direction apparent velocity.
Kalman filter equation based on more than, the estimation procedure of inertia initial baseline misalignment are as follows：
(1), in moment k, calculate state vector predictive value State vector for moment k-1 is estimated
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, predict covariance matrix Pk|k-1=Pk-1；Pk-1It is for the estimate covariance matrix of moment k-1, described
The initial value of estimate covariance matrix is P0, P0For 9 × 9 dimension matrixes of setting；
(3), in moment k, calculate gain matrix 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：
Wherein：Sx、Sy、SzRespectively inertial navigation system provide inertial system X to, Y-direction, Z-direction apparent place；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 detection noise variance；Under the inertial system for respectively setting, X is to, Y-direction, Z-direction speed observation noise side
Difference；
(4), calculate the state vector estimated value of moment kWherein, ZkFor outside
The observation vector of the moment k of offer；In Kalman filtering, observation vector Z=[δ X δ Y δ Z δ V are definedx δVy δVz]T；δ
X, δ Y, δ Z be respectively X that inertial navigation system and GPS system measure to the difference of position, the difference of Y-direction position, Z-direction position difference；δVx、δ
Vy、δVzThe respectively X of inertial navigation system and GPS system measurement to the difference of speed, the difference of Y-direction speed, Z-direction speed difference；
(5), calculate moment k estimate covariance matrix Pk=(I-KkH)Pk|k-1, I is unit matrix；
(6), in the initial value datum drift cycle estimator of setting, repeat step (1)～(5), by Kalman filtering meter
Calculation obtains initial baseline misalignment.
When concrete engineering is realized, the initial baseline misalignment method of estimation and initial baseline deviation compensation method of the present invention
Realize that during missile flight, Flight Control Software calls the method, to guided missile by the software being solidificated in missile-borne computer
Initial inertia datum drift is estimated and compensated, so as to improve the navigation accuracy and accuracy at target of guided missile.
The above, only one specific embodiment of the present invention, but protection scope of the present invention is not limited thereto, and appoints
What those familiar with the art the invention discloses technical scope in, the change or replacement that can be readily occurred in, all
Should be included within the scope of the present invention.
The content not being described in detail in description of the invention belongs to the known technology of professional and technical personnel in the field.
Claims (3)
1. one kind in-flight inertia initial baseline deviation compensation method, it is characterised in that comprise the following steps：
(1), the inertia initial baseline misalignment obtained according to estimation, calculates the in-flight pitch angle deviation of guided missileDriftage
Angular displacement δ ψ and roll angle deviation δ γ, specific formula for calculation are as follows：
Wherein：γ is respectively the angle of pitch, the roll angle that inertial navigation system is provided；φx0、φy0、φz0Respectively estimate that what is 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), setting the amendment cycle in, to missile attitude deviation serial update N clap, wherein, n-th clap makeover process such as
Under：
(2a), calculate revised orientation vector：
Wherein：ψn′、γn' it is respectively the angle of pitch, yaw angle and roll angle after the n-th bat amendment；ψn、γnRespectively inertial navigation
System claps the output angle of pitch, 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-thExport to inertial navigation system, inertial navigation system utilizes the orientation vectorAfter carrying out recurrence calculation, the angle of pitch that output (n+1)th is clappedYaw 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, estimate as follows to obtain inertia initial baseline misalignment：
(1a), calculate the state vector predictive value of moment kFor the state vector estimated value of moment k-1,
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 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), calculate the prediction covariance matrix P of moment kk|k-1=Pk-1；Pk-1It is for the estimate covariance matrix of moment k-1, described
The initial value of estimate covariance matrix is P0, P0For 9 × 9 dimension matrixes of setting；
(1c), calculate the gain matrix K of 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), calculate the state vector estimated value of moment kWherein, ZkThere is provided for outside
Moment k observation vector；In Kalman filtering, observation vector Z=[δ X δ Y δ Z δ V are definedx δVy δVz]T；δX、δY、
δ Z are respectively the difference of the X of inertial navigation system and GPS system measurement to position, the difference of Y-direction position, the difference of Z-direction position；δVx、δVy、δVz
The respectively X of inertial navigation system and GPS system measurement to the difference of speed, the difference of Y-direction speed, Z-direction speed difference；
(1e), calculate moment k estimate covariance matrix Pk=(I-KkH)Pk|k-1, I is unit matrix；
(1f), in the initial value datum drift cycle estimator of setting, repeat step (1a)～(1e) 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 bias estimation method based on GPS according to claim 2, its feature exist
In：In step (1c), observing matrix H and noise variance matrix R are respectively set as：
Wherein：Sx、Sy、SzRespectively inertial navigation system provide inertial system X to, Y-direction, Z-direction apparent place；Wx、Wy、WzRespectively inertial navigation
System provide inertial system X to, Y-direction, Z-direction apparent velocity； X is set under inertial system respectively to, Y-direction, Z-direction position
Put observation noise variance；Under the inertial system for respectively setting X to, Y-direction, Z-direction speed observation noise variance.
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