CN101788296B - SINS/CNS deep integrated navigation system and realization method thereof - Google Patents

SINS/CNS deep integrated navigation system and realization method thereof Download PDF

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CN101788296B
CN101788296B CN2010101014291A CN201010101429A CN101788296B CN 101788296 B CN101788296 B CN 101788296B CN 2010101014291 A CN2010101014291 A CN 2010101014291A CN 201010101429 A CN201010101429 A CN 201010101429A CN 101788296 B CN101788296 B CN 101788296B
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navigation
sins
inertial
information
error
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CN2010101014291A
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CN101788296A (en
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王新龙
吴小娟
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北京航空航天大学
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Abstract

The invention discloses a SINS/CNS deep integrated navigation system and a realization method thereof. Wherein the navigation system comprises a strapdown inertial navigation system (SINS), a celestial navigation system (CNS), an integrated navigation filter, a inertial navigation posture measurement information structure unit; the realization unit comprises the following steps: 1. a large viewing field star sensor assists the strapdown inertial navigation system to obtain a high-precision mathematic horizontal reference; 2. CNS positioning is carried out based on the mathematic horizontal reference; 3. SINS/CNS deep integrated system model and measurement mode are established; 4. integrated navigation system information is fused; 5. the SINS and the CNS assists each other to realize high-precision positioning. In the invention, the star sensor high-precision posture information is employed to assist SINS to obtain high-precision SINS strapdown matrix which serves as the mathematic horizontal reference for CNS positioning, and on the basis, positions and postures of the CNS are employed to comprehensively calibrate the SINS, thus realizing SINS/CNS deep integration and finally achieving high-precision positioning and navigation.

Description

A kind of SINS/CNS deep integrated navigation system and its implementation
Technical field
The present invention relates to a kind of SINS/CNS deep integrated navigation system and its implementation, belong to the integrated navigation technical field.
Background technology
Celestial navigation system (CNS) good concealment, independence be strong, carrier positions and high-accuracy posture information can be provided, but output information is discontinuous, and can be subjected to the influence of external environment in some cases, when using separately, be difficult to satisfy high precision, high performance navigator fix requirement.
The bearing accuracy of tradition celestial navigation system depends primarily on the measuring accuracy of horizontal reference precision and heavenly body sensor.Because heavenly body sensor can more easily reach 1 rad observation precision at present, the precision of Horizon becomes the main factor that influences the celestial navigation precision.Inertial platform horizontal reference error is subjected to the error effect of its core component gyro at present, and list is extremely difficult from the precision that the improvement instrument design improves the inertia horizontal reference, and cost performance reduces; Direct responsive Horizon precision such as employing infrared horizon are lower; Employing based on starlight refraction the quadratic method precision is higher sensitively indirectly, but the uncertainty of atmospheric refraction model has directly influenced the precision and the reliability of horizontal reference.High level of accuracy benchmark technology has become the celestial navigation of realization high precision and has needed one of gordian technique of capturing badly.
And strapdown inertial navigation system (SINS) have navigational parameter comprehensively, output in time continuously, outstanding advantage such as good concealment, antijamming capability be strong, but because the existence of initial alignment error and inertia device error, the SINS error increases in time, be difficult to work alone for a long time, need to increase other metric informations, utilize the method for combination to improve the integrated navigation performance.Both have the characteristics of mutual supplement with each other's advantages SINS, CNS, and both combinations can be given full play to separately advantage, and along with the intensification of combined level, the overall performance of SINS/CNS combined system will be far superior to each autonomous system.
20th century, because detecting, celestial body is only limited to the detection of under the difference moment, different attitude, finishing different celestial bodies, astronomy/inertia combination can only be adopted system-level simple combination navigation mode, celestial navigation system receives the position and the attitude information of inertial navigation system output, regularly the drift of inertial navigation system is proofreaied and correct.This simple combination pattern can damping inertial navigation site error disperse, but its navigation accuracy is subjected to the restriction of the horizontal attitude precision of inertial navigation, and the horizontal attitude precision of inertial navigation depends on the precision of inertial sensor error, therefore this combined method corrective action and little.
To the middle and later periods nineties 20th century, because the development of big visual field celestial body Fast Detection Technique, celestial navigation system can be finished many stars synchronous detection in a certain moment, and the attitude information of under the outside initial information prerequisite of (comprising horizontal reference), determining carrier coordinate relative inertness coordinate (identical with the attitude information of gyro output).Therefore, be that the astronomy/inertia optimum combination pattern of core all can realize on theoretical and engineering fully with compensation inertial navigation gyroscopic drift, but this integrated mode is merely able to compensate the navigation error that causes because of gyroscopic drift in the inertial navigation.
At present, celestial navigation system adopts starlight to reflect the flat method of indirect geodetic and obtains the high-precision independent horizontal reference, can carry out high-precision independent and determine position (three-dimensional coordinate), course and attitude, thereby inertial navigation system is realized thoroughly optimum comprehensively the correction.This astronomy/inertia optimum combination pattern not only can compensate the random drift of gyro, and can the compensated acceleration meter etc. the error that causes of other factors, but the method can only be used on the high-altitude sail body more than the 30km at present, can not realize FR application.
In sum, utilize many stars synchro measure and instantaneous definite carrier inertia attitude principle, solve this gordian technique of celestial navigation high level of accuracy benchmark, and develop high precision astronomy applied widely/inertia deep integrated navigation pattern to replace traditional system-level, rough astronomy/inertia integrated mode, thereby realizing the hi-Fix navigation, is the main direction of studying of astronomy/inertia best of breed navigational system.
Summary of the invention
The objective of the invention is to determine the deficiency of scheme in order to overcome the existing level benchmark, utilize the characteristics of SINS/CNS combination sensor, a kind of SINS/CNS deep integrated navigation system and its implementation are proposed, this method has made full use of the navigation information of each subsystem, has improved the precision of SINS/CNS integrated navigation system.
The dark combination implementing method of a kind of SINS/CNS specifically may further comprise the steps:
Step 1: the auxiliary SINS of star sensor obtains high-precision mathematics horizontal reference;
Step 2: carry out the CNS location based on the mathematics horizontal reference;
Step 3: set up dark combined system state equation of SINS/CNS and measurement equation;
A. make up the dark combined system state model of SINS/CNS;
B. make up the dark combined system of SINS/CNS and measure model;
Step 4: integrated navigation system information fusion;
The auxiliary mutually hi-Fix of realizing of step 5: SINS and CNS.
A kind of SINS/CNS deep integrated navigation system comprises strapdown inertial navitation system (SINS), celestial navigation system, integrated navigation wave filter, inertial navigation attitude measurement information tectonic element;
Strapdown inertial navigation system comprises inertial measurement cluster and navigation calculation unit; Inertial measurement cluster records angular velocity and the specific force of carrier with respect to inertial space, send the navigation calculation unit to the carrier angular velocity that obtains with than force information, the navigation calculation unit calculates the positional information latitude L of carrier in real time by mechanics layout algorithm according to the inertial measurement cluster information transmitted IAnd longitude λ I, speed and attitude; The navigation calculation unit is with the positional information L of carrier simultaneously I, λ IBe input in the integrated navigation wave filter, the SINS navigational parameter is input in the inertial navigation attitude measurement information tectonic element, described navigational parameter is the position L of current navigation time t, carrier I, λ IAnd attitude, the navigation calculation unit also inputs to celestial navigation system with SINS strapdown matrix as the mathematics horizontal reference and measures the elevation angle computing unit, with the positional information L of carrier I, λ IBe input to celestial navigation system analytic Height difference locating module;
Celestial navigation system comprises that big visual field star sensor, many vectors decide the appearance unit, measure the elevation angle computing unit and resolve the difference in height locating module;
Big visual field star sensor synchronization can be observed the starlight Vector Message that obtains three and three above fixed stars, and the observation information that obtains is offered many vectors respectively decide appearance unit and measurement elevation angle computing unit; Many vectors are decided the appearance unit starlight Vector Message that receives are handled, and obtain the attitude information in carrier relative inertness space And with the carrier inertia attitude information of determining Be input in the integrated navigation wave filter; Measure the elevation angle computing unit and utilize the starlight Vector Message of star sensor transmission and the mathematics horizontal reference that the navigation calculation unit provides, obtain the fixed star measurement elevation angle H on plane relatively 0, and the elevation angle measurement information inputed to analytic Height difference locating module; The carrier positions information L that analytic Height difference locating module provides according to the fixed star elevation angle information that measures the transmission of elevation angle computing unit and navigation calculation unit I, λ I, obtain the carrier latitude L of astronomical fixation C, longitude λ C, and with astronomical fixation L as a result C, λ CBe input to the integrated navigation wave filter as measurement information;
Inertial navigation attitude measurement information tectonic element is according to the navigational parameter of navigation calculation unit transmission, obtain inertial navigation attitude measurement information, described inertial navigation attitude measurement information be strapdown inertial navigation system determine be transformed into the direction cosine matrix of carrier coordinate system from equator, the earth's core inertial coordinates system Inertial navigation attitude measurement information tectonic element is with the inertial navigation attitude measurement information that obtains Offer the integrated navigation wave filter;
The integrated navigation wave filter is decided the SINS positional information L that appearance unit, analytic Height difference locating module, inertial navigation attitude measurement information tectonic element provide respectively according to navigation calculation unit, many vectors I, λ I, CNS attitude measurement information , CNS positioning result L C, λ CWith inertial navigation attitude measurement information Handle by Kalman filtering, the navigational parameter of strapdown inertial navigation system and the error of inertial measurement cluster are estimated, and it is fed back in the SINS navigation calculation unit, corresponding error is proofreaied and correct and compensated.
The invention has the advantages that:
(1) the present invention utilizes star sensor high-accuracy posture information to assist SINS, and real-time monitored also compensates mathematical platform misalignment and the gyroscopic drift of SINS, from obtaining high-precision SINS strapdown matrix as the mathematics horizontal reference, is used for the CNS location;
(2) the analytic Height difference method based on the mathematics horizontal reference is positioned error modeling, and survey the probabilistic influence of deduction mathematics horizontal reference in the equation, thereby eliminated the correlativity between CNS positioning error and the SINS attitude error in the position margin of error;
(3) utilize big visual field many stars of star sensor synchro measure and instantaneous definite carrier inertia attitude principle, auxiliary SINS obtains high-precision mathematics horizontal reference by star sensor, CNS can provide high precision position and attitude information simultaneously, SINS is realized comprehensively optimum the correction, further improved the precision of mathematics horizontal reference and astronomical fixation precision on this basis;
(4) SINS/CNS deep integrated navigation pattern of the present invention is auxiliary mutually by SINS, CNS, has given full play to the advantage of each subsystem, finally can realize high-precision location navigation.
Description of drawings
Fig. 1 is the structural representation of a kind of SINS/CNS deep integrated navigation system of the present invention;
Fig. 2 is a kind of process flow diagram of SINS/CNS deep integrated navigation system implementation method;
Fig. 3 is the synoptic diagram that the auxiliary SINS of star sensor of the present invention obtains high precision mathematics horizontal reference;
Among the figure:
1-strapdown inertial navigation system 2-celestial navigation system 3-integrated navigation wave filter 4-inertial navigation attitude measurement information
Tectonic element
The many vectors of 101-inertial measurement cluster 102-navigation calculation unit 201-big visual field star sensor 202-are decided the appearance unit
It is fixed that 203-measures elevation angle calculating 204-analytic Height difference
Position, unit module
Embodiment
The present invention is described in further detail below in conjunction with drawings and Examples.
A kind of SINS/CNS deep integrated navigation system of the present invention as shown in Figure 1, comprises strapdown inertial navitation system (SINS) 1, celestial navigation system 2, integrated navigation wave filter 3, inertial navigation attitude measurement information tectonic element 4;
Strapdown inertial navigation system (SINS) 1 comprises inertial measurement cluster (IMU) 101 and navigation calculation unit 102.Inertial measurement cluster 101 (IMU) records angular velocity and the specific force of carrier with respect to inertial space, send navigation calculation unit 102 to the carrier angular velocity that obtains with than force information, navigation calculation unit 102 calculates the positional information latitude L of carrier in real time by mechanics layout algorithm according to inertial measurement cluster 101 information transmitted IAnd longitude λ I, speed and attitude.Navigation calculation unit 102 is with the positional information L of carrier I, λ IBe input in the integrated navigation wave filter 3, the SINS navigational parameter is input in the inertial navigation attitude measurement information tectonic element 4, described navigational parameter is the position L of current navigation time t, carrier I, λ IAnd attitude, navigation calculation unit 102 also inputs to measurement elevation angle computing unit 203 in the celestial navigation system 2 with SINS strapdown matrix as the mathematics horizontal reference, with the positional information L of carrier I, λ IBe input to the analytic Height difference locating module 204 in the celestial navigation system;
Celestial navigation system (CNS) 2 comprises that big visual field star sensor 201, many vectors decide appearance unit 202, measure elevation angle computing unit 203 and resolve difference in height locating module 204;
Big visual field star sensor 201 synchronizations can be observed the starlight Vector Message that obtains three and three above fixed stars, and the observation information that obtains is offered many vectors respectively decide appearance unit 202 and measurement elevation angle computing unit 203.Many vectors are decided the starlight Vector Message that 202 pairs of appearance unit receive and are handled, and obtain the attitude information in carrier relative inertness space And with the carrier inertia attitude information of determining Be input in the integrated navigation wave filter 3.Measure elevation angle computing unit 203 and utilize the starlight Vector Message of star sensor 201 transmission and the mathematics horizontal reference that navigation calculation unit 102 provides, obtain the fixed star measurement elevation angle H on plane relatively 0, and the elevation angle measurement information inputed to analytic Height difference locating module 204.The carrier positions information L that analytic Height difference locating module 204 provides according to the fixed star elevation angle information that measures 203 transmission of elevation angle computing unit and navigation calculation unit 102 I, λ I, obtain the carrier latitude L of astronomical fixation C, longitude λ C, and with astronomical fixation L as a result C, λ CBe input to integrated navigation wave filter 3 as measurement information;
Inertial navigation attitude measurement information tectonic element 4 is according to the navigational parameter of navigation calculation unit 102 transmission, obtain inertial navigation attitude measurement information, described inertial navigation attitude measurement information be strapdown inertial navigation system 1 determine be transformed into the direction cosine matrix of carrier coordinate system from equator, the earth's core inertial coordinates system , inertial navigation attitude measurement information tectonic element 4 is with the inertial navigation attitude measurement information that obtains Offer integrated navigation wave filter 3;
Integrated navigation wave filter 3 is decided the SINS positional information L that appearance unit 202, analytic Height difference locating module 204, inertial navigation attitude measurement information tectonic element 4 provide respectively according to navigation calculation unit 102, many vectors I, λ I, CNS attitude measurement information CNS positioning result L C, λ CWith inertial navigation attitude measurement information Handle by Kalman filtering, the navigational parameter of strapdown inertial navigation system 1 and the error of inertial measurement cluster are estimated, and it is fed back in the SINS navigation calculation unit 102, corresponding error is proofreaied and correct and compensated;
In whole SINS/CNS deep integrated navigation system,, finally realize high-precision location navigation by assisting mutually between strapdown inertial navigation system 1, the celestial navigation system 2.
A kind of SINS/CNS deep integrated navigation system implementation method of the present invention, flow process specifically may further comprise the steps as shown in Figure 2:
Step 1: star sensor 201 auxiliary strapdown inertial navigation systems 1 in big visual field obtain high-precision mathematics horizontal reference;
Owing to utilize the strapdown matrix of SINS can realize that carrier coordinate system arrives the coordinate conversion of platform coordinate system, obtain the expression of measurement vector at platform coordinate system, therefore the strapdown matrix that resolves is equivalent to set up mathematical platform, and the angle of rotation speed of geographic coordinate system is input in the calculation procedure of " mathematical platform ", but the surface level of platform real-time follow-up carrier loca.But the horizontal attitude information of SINS is drifted about in time, directly utilizes the strapdown matrix as horizontal reference, can cause the astronomical fixation error to be dispersed.And star sensor is equivalent to not have the gyro of drift, therefore utilize star sensor observation celestial body orientation to proofread and correct the drift of mathematical platform, the pure SINS mathematics horizontal reference that overcomes long time continuous working improves the precision of mathematics horizontal reference owing to the error that gyroscopic drift etc. causes.
The synoptic diagram of big visual field star sensor 201 auxiliary SINS acquisition high precision mathematics horizontal references as shown in Figure 3, big visual field star sensor 201 obtains the multidimensional starlight Vector Message of three or three above fixed stars in synchronization observation, many then vectors are decided appearance unit 202 pairs of multidimensional starlight Vector Message and are handled, and obtain the direction cosine matrix of the relative the earth's core of measurement coordinate system s equator inertial system I of big visual field star sensor 201 In conjunction with the installation Matrix C of big visual field star sensor 201 on carrier b s, obtaining carrier is the direction cosine matrix of b with respect to equator, the earth's core inertial system I
It is the direction cosine matrix of the relative the earth's core of b equator inertial system I that the navigation information that inertial navigation attitude measurement information tectonic element 4 is exported by navigation calculation unit 102 constructs the definite carrier of SINS Attitude information with 201 outputs of big visual field star sensor Be complementary, described navigation information is the locating information L of current navigation time t, SINS I, λ IAnd mathematics horizontal reference
Described direction cosine matrix Concrete computing method be:
Utilize the locating information L of SINS I, λ IConstruct the location matrix of SINS Limit obtains being transformed into the be connected direction cosine matrix C of coordinate system e of the earth from equator, the earth's core inertial system I according to current navigation time t I eStrapdown matrix in conjunction with SINS And location matrix Obtain:
C ^ I b = C ^ n b C ^ e n C I e = ( C ^ b n ) T C ^ e n C I e - - - ( 1 )
Consider the influence of factors such as alignment error and gyroscopic drift, the SINS mathematical platform is that n ' and navigation coordinate are to have mathematical platform misalignment vector between the n φ E, φ N, φ UFor east, north, day to misalignment, thereby SINS strapdown matrix Error relevant with misalignment; And because the latitude error δ L of SINS location I, longitude error δ λ IExistence, to be nc do not overlap with the actual n of Department of Geography in the calculating of SINS, and position error vector δ P=[-δ L is arranged Iδ λ ICos L Iδ λ ISin L I] T, the navigation that must cause SINS to determine is n with respect to the be connected direction cosine matrix of coordinate system e of the earth Position deviation δ L with SINS I, δ λ IClose.Then there is following relation:
C ^ e n = ( I - [ δP × ] ) C e n
Wherein, C n b, C e nNavigation system, navigation are the be connected direction cosine matrix of coordinate system of the relative earth relatively to be respectively real carrier system.
Ignore the above error term of secondary and secondary, then the definite carrier of SINS is the direction cosine matrix of equator, relative the earth's core inertial system Can be expressed as
C ^ I b = C I b + C n b [ φ × ] C e n C I e - C n b [ δP × ] C e n C I e
Suppose that desirable free from error carrier is that the direction cosine battle array of the relative the earth's core of b equator inertial system I is C I bBecause the measuring accuracy of star sensor is very high, can think that the carrier system of star sensor output is with respect to the direction cosine matrix of equator, the earth's core inertial system Be real direction cosine matrix C I bMeasure white noise acoustic matrix V with star sensor sStack, that is:
C ~ I b = C I b + V s - - - ( 5 )
With the carrier of being determined by SINS, star sensor respectively is the direction cosine matrix of equator, relative the earth's core inertial system Between difference note make Z s, then have
Z s = C ^ I b - C ~ I b = C n b [ φ × ] C e n C I e - C n b [ δP × ] C e n C I e - V s - - - ( 6 )
Because platform misalignment and gyroscope constant value drift have coupled relation, be the direction cosine battle array of equator, relative the earth's core inertial system with star sensor with the carrier that SINS determines respectively by integrated navigation wave filter 3 Carry out information fusion, can estimate SINS mathematical platform misalignment and gyroscopic drift in real time, then SINS mathematical platform misalignment and gyroscopic drift are revised, to improve the mathematics horizontal reference Precision, concrete grammar is as follows:
Gyroscope among the IMU101 and accelerometer are exported the specific force of carrier in real time And angular velocity And send navigation calculation unit 102 to.In navigation calculation unit 102, the ratio force information of carrier After the mathematics horizontal reference is handled, be directly inputted into the northern bit platform formula inertial reference calculation process of finger and carry out navigation calculation, obtain the position L of carrier I, λ I, navigation information such as speed, attitude; Utilize the estimated value of misalignment Can calculate the attitude error rectification Matrix C N ' n, pass through formula To the strapdown matrix Carry out the misalignment correction, can improve the mathematics horizontal reference Precision; And from gyrostatic output In the error delta of deduction gyroscopic drift in real time ω Ib b, carry out gyroscopic drift error compensation, can access the angular velocity vector information in carrier relative inertness space more accurately ω ~ ib b = ω ^ ib b - δ ω ib b , And then in conjunction with the mathematics horizontal reference and refer to that it is the angular velocity of equator, relative the earth's core inertial system that northern bit platform formula inertial reference calculation obtains navigating Calculate the attitude speed of relative navigation system of high-precision carrier system And the direction of passage cosine matrix differential equation C · b n = C b n Ω nb n Carry out the strapdown matrix update, can further improve the precision of strapdown matrix (mathematics horizontal reference), wherein Ω Nb bBe angular velocity vector Multiplication cross matrix in carrier coordinate system.
Last strapdown inertial navigation system 1 obtains high-precision strapdown matrix It is the mathematics horizontal reference.
Step 2: carry out the CNS location based on the mathematics horizontal reference;
Measure the high precision mathematics horizontal reference that elevation angle computing unit 203 utilizes the auxiliary SINS of star sensor to obtain In conjunction with the multidimensional starlight Vector Message that star sensor 201 observations in big visual field obtain, can obtain the elevation angle H of many observation astrologies for ground level 0, realize the celestial navigation location by analytic Height difference locating module 204 then.Concrete steps are as follows:
A. utilize star sensor 201 auxiliary strapdown inertial navitation system (SINS) 1 to obtain high-precision SINS strapdown matrix It is transferred to celestial navigation system 2 as the mathematics horizontal reference is used for the CNS location; And navigation calculation unit 102 output positional information L I, λ IThe latitude initial value Lat of iteration is provided for analytic Height difference method location AP, longitude initial value Lon AP
B. big visual field star sensor 201 synchronizations can be observed the expression of starlight vector in star sensor measurement coordinate system s and equator, the earth's core inertial system I of three or three above fixed stars that obtain.Consider star sensor measurement noise V i sExistence, actual measurement obtains the position vector of fixed star i in star sensor measurement coordinate system s system and is:
X ~ i s = X i s + V i s - - - ( 7 )
In the formula, X i sBe fixed star i real position vector in star sensor measurement coordinate system s.
At navigation time t, the auxiliary SINS of star sensor obtains high-precision mathematics horizontal reference Installation Matrix C in conjunction with star sensor b s, the starlight vector that can obtain fixed star i is expression among the n (ENU Department of Geography) at navigation coordinate
X ~ i n = C ~ b n C s b X ~ i s = C ~ b n ( C b s ) T X ~ i s - - - ( 8 )
With the starlight position vector that calculates With being defined in navigation is celestial body elevation angle among the n And position angle Expression can be described as:
X ~ i n = cos h ~ i n sin A ~ i n cos h ~ i n cos A ~ i n sin h ~ i n T = p 1 p 2 p 3 T - - - ( 9 )
The starlight vector that then can obtain fixed star i is a elevation angle among the n in navigation And position angle
h ~ i n = arcsin ( p 3 ) - - - ( 10 )
A ~ i n = arctan ( p 1 / p 2 ) - - - ( 11 )
Because to choose sky, northeast Department of Geography is n as navigation coordinate, the fixed star i starlight vector that obtains is a elevation angle among the n in navigation The measurement elevation angle H of the relative surface level of fixed star just 0, can be directly used in analytic Height difference method location.
C. analytic Height difference locating module 204 receives the carrier positions information L that SINS determines I, λ IAs iteration initial value Lat AP, Lon AP, and in conjunction with the elevation angle H that measures many relative surface levels of fixed star that elevation angle computing unit 203 provides 0, use analytic Height difference method can directly obtain the carrier longitude and latitude that is similar to, can rapidly converge to higher precision by the iterative resolution altitude difference method, finally obtain the latitude information L of CNS location C, longitude information λ C
Step 3: set up the dark combined system state model of SINS/CNS and measure model;
Measure the mathematics horizontal reference that elevation angle computing unit 203 provides according to SINS navigation calculation unit 102 Determine the measurement elevation angle H of observation fixed star 0, and offer analytic Height difference locating module 204; And analytic Height difference locating module 204 utilizes the fixed star elevation angle measurement information H that obtains 0, and navigation resolves the carrier positions L of unit 102 transmission I, λ IAs iterative initial value, it is definite that utilization analytic Height difference method is carried out the CNS position, thereby the CNS positioning error is relevant with SINS horizontal attitude error.If ignored the relation between measurement information and the state variable, can influence the estimated accuracy of Kalman filter, and may cause system's instability.Therefore set up CNS Model of locating error based on the mathematics horizontal reference, and the influence of elimination of level fiducial error in the measurement equation of the position of integrated navigation wave filter 3.
The error model of SINS/CNS deep integrated navigation system is made up of SINS, CNS error model.
A. make up the dark combined system state model of SINS/CNS;
Dark combined system state equation is taken as the error equation of SINS, and as the error state variable, the state equation of model is with the error of zero of platform misalignment, velocity error, site error and inertia device:
X · = FX + GW - - - ( 12 )
Wherein, X is the error state vector of SINS; The error state of SINS comprises that east, north, sky are to misalignment φ E, φ N, φ U, velocity error δ V E, δ V N, δ V U, latitude, longitude and height error δ L I, δ λ I, δ h, gyro zero drift ε Bx, ε By, ε BzAccelerometer zero-bit biasing ▽ Bx, ▽ By, ▽ BzF is the system state matrix, W=[ω Gx, ω Gy, ω Gz, ω Dx, ω Dy, ω Dz] TBe system noise sequence, ω Gi(i=x, y, z), ω Dt(i=x, y z) are respectively gyroscope, accelerometer random white noise, C b nBe SINS strapdown matrix, G is a noise matrix, F NState matrix for SINS.F, F SAnd G (t) is respectively
B. make up the dark combined system of SINS/CNS and measure model;
1) will be the direction cosine matrix of the relative the earth's core of b equator inertial system I by SINS with the definite carrier of star sensor respectively Between the difference note measurement amount Z that gestures s, then obtain formula (6).
With Z S (3 * 3)Be launched into column vector Z 1 (9 * 1), in conjunction with the state vector X of integrated navigation system, can be listed as and write out measurement equation and be:
Z 1=H 1X+V 1 (13)
Wherein, H 1For measuring matrix, V 1Measurement white noise sequence for star sensor.
2) the carrier latitude L that SINS is resolved I, longitude λ I, the latitude L that resolves of CNS C, longitude λ CDifference as the position detection amount, obtain:
δLat = L I - L C = δL I - ( δ L C + N L ) δLon = λ I - λ C = δ λ I - ( δ λ C + N λ ) - - - ( 14 )
Wherein, δ Lat and δ Lon are respectively difference of latitude, the difference of longitude of SINS, CNS location, δ L IWith δ λ IBe the positioning error of SINS, δ L CWith δ λ CBe the CNS positioning error that mathematics horizontal reference error causes, N L, N λBe the white Gaussian noise component in the CNS positioning error.Obtain;
Z 2=H 2X+V 2 (15)
In the formula, the position detection vector Z 2 = δLat δLon = L I - L C λ I - λ C ; The white noise component of CNS positioning error V 2 = - N L N λ ; Measure matrix H 2=[H c0 2 * 3I 2 * 20 2 * 7], H wherein c=M (A TA) -1A TB, M = 1 0 0 1 / cos ( L I ) , Suppose that n (n 〉=3) the fixed star true azimuth that the star sensor observation of big visual field obtains is A XNi(i=1,2 ... n), the position vector in star sensor measurement coordinate system s is X i s=[a 1ia 2ia 3i] T, as the direction cosine matrix of mathematics horizontal reference be C ~ b n = T 11 T 12 T 13 T 21 T 22 T 23 T 31 T 32 T 33 , Then obtain easily
A = cos A zN 1 sin A zN 1 cos A zN 2 sin A zN 2 · · · · · · cos A zNn sin A zNn , B = d E 1 d N 1 0 d E 2 d N 2 0 · · · · · · · · · d En d Nn 0 nx 3
d Ei = T 11 a 1 i + T 12 a 2 i + T 13 a 3 i 1 - ( T 31 a 1 i + T 32 a 2 i + T 33 a 3 i ) 2 , d Ni = - ( T 21 a 1 i + T 22 a 2 i + T 23 a 3 i ) 1 - ( T 31 a 1 i + T 32 a 2 i + T 33 a 3 i ) 2
3) when dark combined system is started working, because SINS has certain accumulation of error, at first utilize the auxiliary SINS of star sensor to obtain high precision mathematics horizontal reference information, this moment, Z was measured in the observation of integrated navigation system SINS/CNS=Z 1, corresponding measurement model is formula (13).
Auxiliary SINS obtains on the basis of high level of accuracy benchmark at star sensor, can carry out the celestial navigation location, obtains astronomical fixation latitude and longitude information L C, λ CThis moment, the observation of SINS/CNS deep integrated navigation system was measured:
Z SINS / CNS = Z 1 Z 2
This moment, corresponding measurement model was
Z 1 Z 2 = H 1 H 2 X + V 1 V 2
Step 4: integrated navigation system information fusion;
The carrier positions L that integrated navigation wave filter 3 utilizes strapdown inertial navitation system (SINS) 1 and celestial navigation system 2 to determine respectively I, λ I, L C, λ C, the attitude measurement information Error state to SINS is estimated, obtains the estimation of error information of combined system navigational parameter and inertia device, and these control informations are fed back in the navigation calculation unit 102, and navigational parameter and component error are proofreaied and correct.
The auxiliary mutually hi-Fix of realizing of step 5: SINS and CNS;
Navigation calculation unit 102 obtains high-precision SINS strapdown matrix according to the SINS navigational parameter after proofreading and correct And it is offered as the mathematics horizontal reference measure elevation angle computing unit 203, obtain the more accurate fixed star elevation angle measurement information on plane relatively, be transferred to analytic Height difference locating module 204 then and be used for CNS and locate, can improve the CNS bearing accuracy.Integrated navigation wave filter 3 receives the more accurate position quantity measurement information that analytic Height difference locating module 204 provides, realization is estimated more accurately to the SINS error state, and these error estimates are fed back to navigation calculation unit 102, can further improve the navigation accuracy of SINS, finally make up deeply and realize the precise navigation location by SINS/CNS.

Claims (2)

1. a SINS/CNS deep integrated navigation system is characterized in that, comprises strapdown inertial navitation system (SINS), celestial navigation system, integrated navigation wave filter, inertial navigation attitude measurement information tectonic element;
Strapdown inertial navigation system comprises inertial measurement cluster and navigation calculation unit; Inertial measurement cluster records angular velocity and the specific force of carrier with respect to inertial space, send the navigation calculation unit to the carrier angular velocity that obtains with than force information, the navigation calculation unit calculates the positional information latitude L of carrier in real time by mechanics layout algorithm according to the inertial measurement cluster information transmitted IAnd longitude λ I, speed and attitude; The navigation calculation unit is with the positional information L of carrier I, λ IBe input in the integrated navigation wave filter, with the position L of current navigation time t, carrier I, λ IBe input in the inertial navigation attitude measurement information tectonic element with attitude, the navigation calculation unit also inputs to measurement elevation angle computing unit in the celestial navigation system with SINS strapdown matrix as the mathematics horizontal reference, with the positional information L of carrier I, λ IBe input to the analytic Height difference locating module in the celestial navigation system;
Celestial navigation system comprises that big visual field star sensor, many vectors decide the appearance unit, measure the elevation angle computing unit and resolve the difference in height locating module;
Big visual field star sensor synchronization can be observed the starlight Vector Message that obtains three and three above fixed stars, and the observation information that obtains is offered many vectors respectively decide appearance unit and measurement elevation angle computing unit; Many vectors are decided the appearance unit starlight Vector Message that receives are handled, and obtain the attitude information in carrier relative inertness space And with the carrier inertia attitude information of determining Be input in the integrated navigation wave filter; Measure the elevation angle computing unit and utilize the starlight Vector Message of star sensor transmission and the mathematics horizontal reference that the navigation calculation unit provides, obtain the fixed star measurement elevation angle H on plane relatively 0, and the elevation angle measurement information inputed to analytic Height difference locating module; The carrier positions information L that analytic Height difference locating module provides according to the fixed star elevation angle information that measures the transmission of elevation angle computing unit and navigation calculation unit I, λ I, obtain the carrier latitude L of astronomical fixation C, longitude λ C, and with astronomical fixation L as a result C, λ CBe input to the integrated navigation wave filter as measurement information;
Inertial navigation attitude measurement information tectonic element is according to the navigational parameter of navigation calculation unit transmission, obtain inertial navigation attitude measurement information, described inertial navigation attitude measurement information be strapdown inertial navigation system determine be transformed into the direction cosine matrix of carrier coordinate system from equator, the earth's core inertial coordinates system Inertial navigation attitude measurement information tectonic element is with the inertial navigation attitude measurement information that obtains Offer the integrated navigation wave filter;
The integrated navigation wave filter is decided the SINS positional information L that appearance unit, analytic Height difference locating module, inertial navigation attitude measurement information tectonic element provide respectively according to navigation calculation unit, many vectors I, λ I, CNS attitude measurement information CNS positioning result L C, λ CWith inertial navigation attitude measurement information Handle by Kalman filtering, the navigational parameter of strapdown inertial navigation system and the error of inertial measurement cluster are estimated, and it is fed back in the SINS navigation calculation unit, corresponding error is proofreaied and correct and compensated.
2. the implementation method of a SINS/CNS deep integrated navigation system is characterized in that, comprises following step:
Step 1: star sensor auxiliary strapdown inertial navigation system in big visual field obtains high-precision mathematics horizontal reference;
Big visual field star sensor obtains the multidimensional starlight Vector Message of three or three above fixed stars in synchronization observation, many then vectors are decided the appearance unit multidimensional starlight Vector Message are handled, and obtain the direction cosine matrix of the relative the earth's core of measurement coordinate system s equator inertial system I of big visual field star sensor In conjunction with the big installation matrix of visual field star sensor on carrier Obtaining carrier is the direction cosine matrix of b with respect to equator, the earth's core inertial system I
It is the direction cosine matrix of the relative the earth's core of b equator inertial system I that the navigation information that inertial navigation attitude measurement information tectonic element is exported by the navigation calculation unit constructs the definite carrier of SINS Attitude information with the output of big visual field star sensor Be complementary, described navigation information is the locating information L of current navigation time t, SINS I, λ IAnd mathematics horizontal reference
Described direction cosine matrix Concrete computing method be:
Utilize the locating information L of SINS I, λ IConstruct the location matrix of SINS Obtain being transformed into the be connected direction cosine matrix of coordinate system e of the earth according to current navigation time t from equator, the earth's core inertial system I Strapdown matrix in conjunction with SINS And location matrix Obtain:
C ^ I b = C ^ n b C ^ e n C I e = ( C ^ b n ) T C ^ e n C I e - - - ( 1 )
The SINS mathematical platform is that n ' and navigation coordinate are to have mathematical platform misalignment vector between the n φ E, φ N, φ UFor east, north, day to misalignment, SINS strapdown matrix Error relevant with misalignment; Because the latitude error δ L of SINS location I, longitude error δ λ IExistence, to be nc do not overlap with the actual n of Department of Geography in the calculating of SINS, and position error vector δ P=[-δ L is arranged Iδ λ ICosL Iδ λ ISinL I] T, the navigation coordinate that causes SINS to determine is n with respect to the be connected direction cosine matrix of coordinate system e of the earth Position deviation δ L with SINS I, δ λ IRelevant, then there is following relation:
C ^ e n = ( I - [ δP × ] ) C e n - - - ( 3 )
Wherein, Navigation system, navigation are the be connected direction cosine matrix of coordinate system of the relative earth relatively to be respectively real carrier system;
Ignore the above error term of secondary and secondary, then the definite carrier of SINS is the direction cosine matrix of equator, relative the earth's core inertial system Be expressed as
Suppose that desirable free from error carrier is that the direction cosine battle array of the relative the earth's core of b equator inertial system I is Think that the carrier system of star sensor output is with respect to the direction cosine matrix of equator, the earth's core inertial system It is real direction cosine matrix Measure white noise acoustic matrix V with star sensor sStack, that is:
C ~ I b = C I b + V s - - - ( 5 )
With the carrier of being determined by SINS, star sensor respectively is the direction cosine matrix of equator, relative the earth's core inertial system Between difference note make Z s, then have
Platform misalignment and gyroscope constant value drift have coupled relation, are the direction cosine battle array of relative the earth's core equator inertial system with star sensor with the carrier that SINS determines respectively by the integrated navigation wave filter Carry out information fusion, estimate SINS mathematical platform misalignment and gyroscopic drift in real time, then SINS mathematical platform misalignment and gyroscopic drift are revised, to improve the mathematics horizontal reference Precision, concrete grammar is as follows:
Gyroscope in the inertial measurement cluster and accelerometer are exported the specific force of carrier in real time And angular velocity Send the navigation calculation unit to; In the navigation calculation unit, the ratio force information of carrier After the mathematics horizontal reference is handled, be directly inputted into the northern bit platform formula inertial reference calculation process of finger and carry out navigation calculation, obtain the position L of carrier I, λ I, speed, attitude navigation information; Utilize the estimated value of misalignment Calculate the attitude error rectification matrix Pass through formula To the strapdown matrix Carry out the misalignment correction, improve the mathematics horizontal reference Precision; From gyrostatic output In the error of deduction gyroscopic drift in real time Carry out gyroscopic drift error compensation, obtain the angular velocity vector information in carrier relative inertness space accurately In conjunction with the mathematics horizontal reference and refer to that it is the angular velocity of equator, relative the earth's core inertial system that northern bit platform formula inertial reference calculation obtains navigating Calculate the attitude speed of relative navigation system of high-precision carrier system And the direction of passage cosine matrix differential equation Carry out the strapdown matrix update, wherein Be angular velocity vector Multiplication cross matrix in carrier coordinate system;
Last strapdown inertial navigation system obtains high-precision strapdown matrix It is the mathematics horizontal reference;
Step 2: carry out the CNS location based on the mathematics horizontal reference;
A. utilize the auxiliary strapdown inertial navitation system (SINS) of star sensor to obtain high-precision SINS strapdown matrix It is transferred to celestial navigation system as the mathematics horizontal reference; And navigation calculation unit output positional information L I, λ IThe latitude initial value Lat of iteration is provided for analytic Height difference method location AP, longitude initial value Lon AP
The expression of starlight vector in star sensor measurement coordinate system s and equator, the earth's core inertial system I of three or three the above fixed stars that the star sensor synchronization observation of b. big visual field obtains; Because star sensor measurement noise Existence, actual measurement obtains the position vector of fixed star i in star sensor measurement coordinate system s system and is:
X ~ i s = X i s + V i s - - - ( 7 )
In the formula, Be fixed star i real position vector in star sensor measurement coordinate system s;
At navigation time t, the auxiliary SINS of star sensor obtains high-precision mathematics horizontal reference Installation matrix in conjunction with star sensor The starlight vector that obtains fixed star i is expression among the n at navigation coordinate
X ~ i n = C ~ b n C s b X ~ i s = C ~ b n ( C b s ) T X ~ i s - - - ( 8 )
With the starlight position vector that calculates With being defined in navigation is celestial body elevation angle among the n And position angle Expression:
X ~ i n = cos h ~ i n sin A ~ i n cos h ~ i n cos A ~ i n sin h ~ i n T = p 1 p 2 p 3 T - - - ( 9 )
The starlight vector that then obtains fixed star i is a elevation angle among the n in navigation And position angle
h ~ i n = arcsin ( p 3 ) - - - ( 10 )
A ~ i n = arctan ( p 1 / p 2 ) - - - ( 11 )
Because to choose sky, northeast Department of Geography is n as navigation coordinate, the fixed star i starlight vector that obtains is a elevation angle among the n in navigation Measurement elevation angle H for the relative surface level of fixed star 0
C. analytic Height difference locating module receives the carrier positions information L that SINS determines I, λ IAs iteration initial value Lat AP, Lon AP, and in conjunction with the measuring height angle H that measures many relative surface levels of fixed star that the elevation angle computing unit provides 0, use analytic Height difference method directly to obtain the carrier longitude and latitude that is similar to, rapidly converge to higher precision by the iterative resolution altitude difference method, obtain the latitude information L of CNS location C, longitude information λ C
Step 3: set up the dark combined system state model of SINS/CNS and measure model;
The error model of SINS/CNS deep integrated navigation system comprises SINS, CNS error model;
A. make up the dark combined system state model of SINS/CNS;
Dark combined system state equation is taken as the error equation of SINS, and as the error state variable, the state equation of model is with the error of zero of platform misalignment, velocity error, site error and inertia device:
X · = FX + GW - - - ( 12 )
Wherein, X is the error state vector of SINS; The error state of SINS comprises that east, north, sky are to misalignment φ E, φ N, φ U, velocity error δ V E, δ V N, δ V U, latitude, longitude and height error δ L I, δ λ I, δ h, gyro zero drift ε Bx, ε By, ε BzThe biasing of accelerometer zero-bit F is the system state matrix, W=[ω Gx, ω Gy, ω Gz, ω Dx, ω Dy, ω Dz] TBe system noise sequence, ω Gi(i=x, y, z), ω Di(i=x, y z) are respectively gyroscope, accelerometer random white noise, Be SINS strapdown matrix, G is a noise matrix, F NState matrix for SINS; F, F SBe respectively with G
F = F N F S 0 6 × 9 0 6 × 6 15 × 15
B. make up the dark combined system of SINS/CNS and measure model;
1) will be the direction cosine matrix of the relative the earth's core of b equator inertial system I by SINS with the definite carrier of star sensor respectively Between the difference note measurement amount Z that gestures s, then obtain formula (6);
With Z S (3 * 3)Be launched into column vector Z 1 (9 * 1), in conjunction with the state vector X of integrated navigation system, row write out measurement equation and are:
Z 1=H 1X+V 1 (13)
Wherein, H 1For measuring matrix, V 1Measurement white noise sequence for star sensor;
2) the carrier latitude L that SINS is resolved I, longitude λ I, the latitude L that resolves of CNS C, longitude λ CDifference as the position detection amount, obtain:
δLat = L I - L C = δL I - ( δL C + N L ) δLon = λ I - λ C = δ λ I - ( δλ C + N λ ) - - - ( 14 )
Wherein, δ Lat and δ Lon are respectively difference of latitude, the difference of longitude of SINS, CNS location, δ L IWith δ λ IBe the positioning error of SINS, δ L CWith δ λ CBe the CNS positioning error that mathematics horizontal reference error causes, N L, N λBe the white Gaussian noise component in the CNS positioning error; Obtain;
Z 2=H 2X+V 2 (15)
In the formula, the position detection vector Z 2 = δLat δLon = L I - L C λ I - λ C ; The white noise component of CNS positioning error V 2 = - N L N λ ; Measure matrix H 2=[H c0 2 * 3I 2 * 20 2 * 7], H wherein c=M (A TA) -1A TB, M = 1 0 0 1 / cos ( L I ) , Suppose that n (n 〉=3) the fixed star true azimuth that the star sensor observation of big visual field obtains is A ZNi(i=1,2 ... n), the position vector in star sensor measurement coordinate system s is X i s = a 1 i a 2 i a 3 i T , Direction cosine matrix as the mathematics horizontal reference is C ~ b n = T 11 T 12 T 13 T 21 T 22 T 23 T 31 T 32 T 33 , Then obtain:
A = cos A zN 1 sin A zN 1 cos A zN 2 sin A zN 2 . . . . . . cos A zNn sin A zNn , B = d E 1 d N 1 0 d E 2 d N 2 0 . . . . . . . . . d En d Nn 0 n × 3
d Ei = T 11 a 1 i + T 12 a 2 i + T 13 a 3 i 1 - ( T 31 a 1 i + T 32 a 2 i + T 33 a 3 i ) 2 , d Ni = - ( T 21 a 1 i + T 22 a 2 i + T 23 a 3 i ) 1 - ( T 31 a 1 i + T 32 a 2 i + T 33 a 3 i ) 2
3) when dark combined system is started working, at first utilize the auxiliary SINS of star sensor to obtain high precision mathematics horizontal reference information, this moment, Z was measured in the observation of integrated navigation system SINS/CNS=Z 1, corresponding measurement model is formula (13);
Auxiliary SINS obtains on the basis of high level of accuracy benchmark at star sensor, carries out the celestial navigation location, obtains astronomical fixation latitude and longitude information L C, λ CThis moment, the observation of SINS/CNS deep integrated navigation system was measured:
Z SINS / CNS = Z 1 Z 2
This moment, corresponding measurement model was
Z 1 Z 2 = X 1 H 2 X + V 1 V 2
Step 4: integrated navigation system information fusion;
The carrier positions L that the integrated navigation wave filter utilizes strapdown inertial navitation system (SINS) and celestial navigation system to determine respectively I, λ I, L C, λ C, the attitude measurement information Error state to SINS is estimated, obtains the estimation of error information of combined system navigational parameter and inertia device, and these control informations are fed back in the navigation calculation unit, and navigational parameter and component error are proofreaied and correct;
The auxiliary mutually hi-Fix of realizing of step 5: SINS and CNS;
The navigation calculation unit obtains high-precision SINS strapdown matrix according to the SINS navigational parameter after proofreading and correct It is offered measurement elevation angle computing unit as the mathematics horizontal reference, obtain the more accurate fixed star elevation angle measurement information on plane relatively, be transferred to analytic Height difference locating module and be used for the CNS location; The integrated navigation wave filter receives the more accurate position quantity measurement information that analytic Height difference locating module provides, realization is estimated more accurately to the SINS error state, error estimate is fed back to the navigation calculation unit, further improve the navigation accuracy of SINS, finally make up deeply and realize the precise navigation location by SINS/CNS.
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Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN1869589A (en) * 2006-06-27 2006-11-29 北京航空航天大学 Strapdown intertial/celestial combined navigation semi-material emulation system

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN1869589A (en) * 2006-06-27 2006-11-29 北京航空航天大学 Strapdown intertial/celestial combined navigation semi-material emulation system

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
全伟等.SINS/CNS组合导航半实物仿真系统及其实验研究.《系统仿真学报》.2007,(第15期), *

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