CN103471593A - Method for correcting measurement errors of inertial navigation system based on global positioning system (GPS) information - Google Patents

Method for correcting measurement errors of inertial navigation system based on global positioning system (GPS) information Download PDF

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CN103471593A
CN103471593A CN2013104036972A CN201310403697A CN103471593A CN 103471593 A CN103471593 A CN 103471593A CN 2013104036972 A CN2013104036972 A CN 2013104036972A CN 201310403697 A CN201310403697 A CN 201310403697A CN 103471593 A CN103471593 A CN 103471593A
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longitude
latitude
revised
speed
error
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CN103471593B (en
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魏宗康
刘生炳
赵龙
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China Aerospace Times Electronics Corp
Beijing Aerospace Control Instrument Institute
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Abstract

The invention discloses a method for correcting measurement errors of an inertial navigation system based on global positioning system (GPS) information. A series control method is adopted; the external GPS position information and speed information are introduced to correct an inertial navigation error; by virtue of an appropriate controller, input signal noise can be filtered, system input can be tracked by system output, and speed and position errors of the inertial navigation system can be corrected. The method does not depend on the accuracy of a system error model and is high in operation speed and low in time cost; when the external measurement data frequency is high, the speed and position measurement errors of the inertial navigation system can be corrected; the system error model can be re-built by correcting the parameters of the controller; the workload is small.

Description

A kind of inertial navigation system measuring error modification method based on GPS information
Technical field
The present invention relates to a kind of inertial navigation system measuring error modification method, relate in particular to a kind of inertial navigation system measuring error modification method based on GPS information, can be used for having the occasion of high precision navigation accuracy, fields such as Aero-Space, mapping.
Background technology
Inertial navigation system is a voyage Estimation System based on the acceleration quadratic integral, and it relies on plant equipment fully and corresponding algorithm is automatic, the complete independently navigation task, and any light, electrical communication do not occur in the external world.So good concealment, working environment is not subject to the restriction of meteorological condition.The advantage that this is unique, make it become a kind of widely used main navigational system in space flight, aviation, navigational field.But because inertial navigation system itself exists site error, Initial Alignment Error, gravity anomaly etc., when inertial navigation system works long hours, navigation error is dispersed in time, the effective way that improves the inertial navigation system performance is to adopt integrated navigation technology, by two or more non-similar navigational system, same navigation information is done to measure and resolve with formation volume and measure, calculate the error of each navigational system and proofread and correct from these measurement amounts.Adopt the system of integrated navigation technology to be called integrated navigation system, involved each system is called subsystem.
Because inertial navigation system itself exists site error, Initial Alignment Error, gravity anomaly etc., when inertial navigation system works long hours, navigation speed and site error are dispersed in time.Another importance of integrated navigation system is information fusion algorithm, is about to the different navigation system and obtains navigation information and merges and obtain high-precision navigation and export.At present common information fusion algorithm is generally based on the optimal estimation theory, for example kalman filtering theory with and derivative algorithm.Need to set up the error model of system when using Kalman filtering algorithm to carry out Design of Integrated Navigation System, in addition, Kalman filtering algorithm is a kind of recursive algorithm, need the parameters initial value during starting algorithm, the rationality of the correctness of error model and parameter initial value will have a direct impact convergence, need to rational parameter initial value be set for different product during the engineering application, debugging work load is very large.In addition, along with the raising of metric information frequency, will be difficult to calculate integrated navigation after the employing Kalman filter in one-period and calculate.
Summary of the invention
Technology of the present invention is dealt with problems: overcome the deficiencies in the prior art, a kind of inertial navigation system measuring error modification method based on GPS information is provided, the correction of realization to inertial navigation system speed and position, improved robustness and the navigation accuracy of integrated navigation system.
Technical solution of the present invention: a kind of inertial navigation system measuring error modification method based on GPS information comprises north orientation channel measurement error correction step and east orientation channel measurement error correction step:
North orientation channel measurement error correction step:
(1) north orientation acceleration a inertial navigation system measured ncarry out integration and obtain north orientation speed
Figure BDA00003786523400000218
(2) by north orientation speed carrier true north speed v with the GPS acquisition ndo the poor north orientation velocity error that obtains
Figure BDA00003786523400000219
? ▿ v n = v ^ n - v n ;
(3) north orientation speed correction link controller is according to the north orientation velocity error
Figure BDA00003786523400000221
generate latitude controlled quentity controlled variable u vn;
(4) utilize latitude controlled quentity controlled variable u vnto north orientation acceleration a nrevised, through revised north orientation acceleration a nagain carry out integration and obtain revised north orientation speed;
(5) utilize
Figure BDA0000378652340000021
by revised north orientation speed
Figure BDA0000378652340000022
be converted into the latitude rate of change
Figure BDA00003786523400000223
Figure BDA0000378652340000023
(6) to the latitude rate of change
Figure BDA0000378652340000024
carry out integration and obtain latitude
Figure BDA0000378652340000025
by latitude
Figure BDA0000378652340000026
carrier true latitude with the GPS acquisition
Figure BDA0000378652340000027
do the poor latitude error that obtains
Figure BDA0000378652340000028
?
(7) latitude correction link controller is according to latitude error
Figure BDA00003786523400000210
generate controlled quentity controlled variable
Figure BDA00003786523400000211
(8) utilize controlled quentity controlled variable
Figure BDA00003786523400000212
to the latitude rate of change
Figure BDA00003786523400000213
revised, through revised latitude rate of change
Figure BDA00003786523400000214
again carry out integration and obtain revised latitude
Figure BDA00003786523400000215
revised latitude
Figure BDA00003786523400000216
north orientation passage as inertial navigation system is exported, thereby completes the correction of north orientation channel measurement error;
East orientation channel measurement error correction step:
(1) east orientation acceleration a inertial navigation system measured ecarry out integration and obtain north orientation speed
Figure BDA00003786523400000222
(2) by east orientation speed the true east orientation speed v of carrier with the GPS acquisition edo poor the east orientation velocity error that obtains ? ▿ v e = v ^ e - v e ;
(3) east orientation speed correction link controller is according to the east orientation velocity error
Figure BDA00003786523400000320
generate longitude controlled quentity controlled variable u ve;
(4) utilize longitude controlled quentity controlled variable u veto east orientation acceleration a erevised, through revised east orientation acceleration a eagain carry out integration and obtain revised east orientation speed;
(5) utilize
Figure BDA0000378652340000033
by revised east orientation speed
Figure BDA0000378652340000034
be converted into the longitude rate of change
Figure BDA0000378652340000035
Figure BDA0000378652340000036
(6) to the longitude rate of change
Figure BDA0000378652340000037
carry out integration and obtain longitude
Figure BDA0000378652340000038
by longitude
Figure BDA0000378652340000039
the true longitude λ of carrier obtained with GPS does the poor longitude error that obtains
Figure BDA00003786523400000321
?
Figure BDA00003786523400000310
(7) longitude correction link controller is according to longitude error
Figure BDA00003786523400000322
generate longitude controlled quentity controlled variable u λ;
(8) utilize longitude controlled quentity controlled variable u λto the longitude rate of change revised, through revised longitude rate of change
Figure BDA00003786523400000312
again carry out integration and obtain revised longitude
Figure BDA00003786523400000313
revised longitude
Figure BDA00003786523400000314
east orientation passage as inertial navigation system is exported, thereby completes the correction of east orientation channel measurement error.
The transport function of described north orientation speed correction link controller, latitude correction link controller, east orientation speed correction link controller and longitude correction link controller is respectively: represent respectively north orientation speed correction link controller, latitude correction link controller, east orientation speed correction link controller and longitude correction link controller;
Wherein v is the integral element number, v=0 or 1; The bandwidth in handicapping Buddhist nun loop is ω c, when v=0, T 1 ≥ T 2 > 1 ω c , When v=1, T 2 > 1 ω c > T 1 ;
T 1for the inertial element time constant;
T 2for the first derivative element time constant;
K is enlargement factor.
Ultimate principle of the present invention: when the actual position information (longitude λ, the latitude that obtain carrier ), velocity information (east orientation speed v e, the north orientation speed v n) time, can be by the positional information (longitude of inertial navigation system output
Figure BDA0000378652340000041
latitude
Figure BDA0000378652340000042
) and velocity information (east orientation speed
Figure BDA0000378652340000043
north orientation speed
Figure BDA0000378652340000044
) with actual position information and velocity information, do poorly respectively, form error signal,
Figure BDA0000378652340000045
Figure BDA0000378652340000046
this error signal formation control amount after controller is applied to respectively position integral equation and the rate integrating equation of inertial navigation, realizes the damping to INS errors.
The present invention's advantage compared with prior art is as follows:
(1) existing Kalman filtering algorithm need to be set up the error model of system, algorithm stability depends critically upon correctness and the levels of precision of navigation error model, method of the present invention does not rely on the accuracy of SYSTEM ERROR MODEL, the method that adopts series connection to control is proposed, introduce external GPS positional information, velocity information so that the inertial navigation error is revised, can complete the filtering to the input signal noise by designing suitable controller, the tracking of completion system output to the system input, complete the correction to inertial navigation system speed and site error.
(2) existing Kalman filtering algorithm time overhead is larger, and the filtering cycle is longer, method fast operation of the present invention, and time overhead is short, when outer survey data frequency is higher, also can complete the correction to inertial navigation system speed and errors in position measurement.
(3) if existing Kalman filtering algorithm information source changes, need to re-establish the error model of system, workload is larger, and the present invention only need get final product by change control device parameter.
The accompanying drawing explanation
Fig. 1 is the schematic diagram of realizing of the present invention;
The damping circuit open-loop transfer function Bode diagram that Fig. 2 is adoption rate link controller of the present invention;
Fig. 3 is adoption rate-hysteresis-differentiation element controller damping circuit open-loop transfer function Bode diagram;
Fig. 4 is adoption rate-integration-leading-delay component controller damping circuit open-loop transfer function Bode diagram;
Fig. 5 is the Output rusults figure of the integrated navigation system of use the present invention design;
Fig. 6 is for being used the integrated navigation system position navigation curve map of the present invention's design.
Embodiment
Realization approach of the present invention is: the integrated navigation system damping circuit comprises speed damping circuit (north orientation speed damping circuit, east orientation speed damping circuit) and position damping loop (latitude damping circuit, longitude damping circuit).The ultimate principle of damping is actual position information (longitude λ, latitude when obtaining carrier
Figure BDA0000378652340000051
), velocity information (east orientation speed v e, the north orientation speed v n) time, can be by the positional information (longitude of inertial navigation system output
Figure BDA0000378652340000052
latitude
Figure BDA0000378652340000053
) and velocity information (east orientation speed
Figure BDA0000378652340000054
north orientation speed
Figure BDA0000378652340000055
) with actual position information and velocity information, do poorly respectively, form error signal,
Figure BDA0000378652340000056
this error signal formation control amount after controller is applied to respectively position integral equation and the rate integrating equation of inertial navigation, realizes the damping to inertial navigation system velocity error and site error.
As shown in Figure 1, modification method of the present invention comprises north orientation channel measurement error correction step and east orientation channel measurement error correction step to concrete implementation method:
North orientation channel measurement error correction step:
(1) north orientation acceleration a inertial navigation system measured ncarry out integration and obtain north orientation speed
Figure BDA0000378652340000058
(2) by north orientation speed
Figure BDA0000378652340000059
carrier true north speed v with the GPS acquisition ndo the poor north orientation velocity error that obtains
Figure BDA00003786523400000527
? ▿ v n = v ^ n - v n ;
(3) north orientation speed correction link controller is according to the north orientation velocity error
Figure BDA00003786523400000528
generate latitude controlled quentity controlled variable u vn;
(4) utilize latitude controlled quentity controlled variable u vnto north orientation acceleration a nrevised, through revised north orientation acceleration a nagain carry out integration and obtain revised north orientation speed;
(5) utilize
Figure BDA00003786523400000511
by revised north orientation speed be converted into the latitude rate of change
Figure BDA00003786523400000529
Figure BDA00003786523400000513
(6) to the latitude rate of change
Figure BDA00003786523400000514
carry out integration and obtain latitude
Figure BDA00003786523400000515
by latitude
Figure BDA00003786523400000516
carrier true latitude with the GPS acquisition
Figure BDA00003786523400000517
do the poor latitude error that obtains
Figure BDA00003786523400000518
?
Figure BDA00003786523400000519
(7) latitude correction link controller is according to latitude error
Figure BDA00003786523400000520
generate controlled quentity controlled variable
Figure BDA00003786523400000521
(8) utilize controlled quentity controlled variable
Figure BDA00003786523400000522
to the latitude rate of change
Figure BDA00003786523400000523
revised, through revised latitude rate of change
Figure BDA00003786523400000524
again carry out integration and obtain revised latitude
Figure BDA00003786523400000525
revised latitude
Figure BDA00003786523400000526
north orientation passage as inertial navigation system is exported, thereby completes the correction of north orientation channel measurement error;
East orientation channel measurement error correction step:
(1) east orientation acceleration a inertial navigation system measured ecarry out integration and obtain north orientation speed
Figure BDA0000378652340000061
(2) by east orientation speed
Figure BDA0000378652340000062
the true east orientation speed v of carrier with the GPS acquisition edo the poor east orientation velocity error that obtains
Figure BDA00003786523400000617
? ▿ v e = v ^ e - v e ;
(3) east orientation speed correction link controller is according to the east orientation velocity error
Figure BDA00003786523400000618
generate longitude controlled quentity controlled variable u ve;
(4) utilize longitude controlled quentity controlled variable u veto east orientation acceleration a erevised, through revised east orientation acceleration a eagain carry out integration and obtain revised east orientation speed;
(5) utilize
Figure BDA0000378652340000064
by revised east orientation speed
Figure BDA0000378652340000065
be converted into the longitude rate of change
Figure BDA0000378652340000066
Figure BDA0000378652340000067
(6) to the longitude rate of change
Figure BDA00003786523400000619
carry out integration and obtain longitude by longitude
Figure BDA0000378652340000069
the true longitude λ of carrier obtained with GPS does the poor longitude error that obtains
Figure BDA00003786523400000620
?
Figure BDA00003786523400000610
(7) longitude correction link controller is according to longitude error
Figure BDA00003786523400000621
generate longitude controlled quentity controlled variable u λ;
(8) utilize longitude controlled quentity controlled variable u λto the longitude rate of change
Figure BDA00003786523400000611
revised, through revised longitude rate of change
Figure BDA00003786523400000612
again carry out integration and obtain revised longitude
Figure BDA00003786523400000613
revised longitude
Figure BDA00003786523400000614
east orientation passage as inertial navigation system is exported, thereby completes the correction of east orientation channel measurement error.
The transport function of described north orientation speed correction link controller, latitude correction link controller, east orientation speed correction link controller and longitude correction link controller is respectively:
Figure BDA00003786523400000623
represent respectively north orientation speed correction link controller, latitude correction link controller, east orientation speed correction link controller and longitude correction link controller;
Wherein v is the integral element number, v=0 or 1; The bandwidth in handicapping Buddhist nun loop is ω c, when v=0, T 1 ≥ T 2 > 1 ω c , When v=1, T 2 > 1 ω c > T 1 ;
T 1for the inertial element time constant;
T 2for the first derivative element time constant;
K is enlargement factor.
Due to controlling unit C 1(s), C 2(s), C 3and C (s) 4(s) be continuous system, when practical application, need carry out the discretize processing, discretization method is Bilinear transformation method, even
Figure BDA0000378652340000071
carry out discretize, the time discretization that wherein Δ T is navigational system.
Example 1., when v=0, is established T 1=T 2, have
Figure BDA00003786523400000713
at first, determine the shearing frequency ω of system c, the open-loop transfer function of damping system is
Figure BDA0000378652340000073
wherein, ω c=K.Fig. 2 solid line has provided ω cthe open loop Bode diagram of damping system during=100rad/s, establishing sample frequency is 100Hz, the system open loop Bode diagram after adopting the bilinearity discretize as Fig. 2 in dotted line.
Example 2., when v=0, is established T 1 ≥ T 2 > 1 ω c , Have C i ( s ) = K T 2 s + 1 T 1 s + 1 , i = 1,2,3,4 . At first, determine the shearing frequency ω of system c, get
Figure BDA0000378652340000076
the open-loop transfer function of damping system is
Figure BDA0000378652340000077
fig. 3 solid line has provided ω c=100rad/s, T 1=1, T 2the damping system open loop Bode diagram of=0.1 o'clock.As can be seen from the figure, system is improved in the gain of low-frequency range, and its multiple is
Figure BDA0000378652340000078
if sample frequency is 100Hz, the system open loop Bode diagram after adopting the bilinearity discretize as Fig. 3 in dotted line.
Example 3., when v=1, is established T 2 > 1 ω c > T 1 , Have C i ( s ) = K T 2 s + 1 s ( T 1 s + 1 ) , i = 1,2,3,4 . At first, determine the shearing frequency ω of system c, get K=T 2ω c 2, the open-loop transfer function of damping system is
Figure BDA00003786523400000711
fig. 4 solid line has provided ω c=100rad/s, T 1=0.001, T 2the open loop Bode diagram of=0.1 damping system.As can be seen from Figure 4, system is improved in the gain of low-frequency range.If sample frequency is 100Hz, the system open loop Bode diagram after adopting the bilinearity discretize as Fig. 4 in dotted line.
Can complete the inertial navigation system damping method Design of Integrated Navigation System based on external position and velocity information by said method, Fig. 5 the first row is respectively the longitude curve, latitude curve and altitude curve, the second row is respectively the east orientation rate curve, north orientation rate curve and sky are to rate curve, the third line is respectively the angle of pitch, roll angle and course angle, in Fig. 5, solid line is the GPS navigation result, dotted line is the integrated navigation system navigation results, can find out in the situation that GPS does not have frame losing, article two, curve almost completely overlaps, the method of the invention has realized the correction of GPS to the inertial navigation system measuring error, and the noise of integrated navigation system is less, the navigation position curve that Fig. 6 is integrated navigation system, can find out in the situation that GPS does not have frame losing integrated navigation system position navigation results to conform to actual.
The present invention not detailed description is known to the skilled person technology.

Claims (2)

1. the inertial navigation system measuring error modification method based on GPS information is characterized in that comprising north orientation channel measurement error correction step and east orientation channel measurement error correction step:
North orientation channel measurement error correction step:
(1) north orientation acceleration a inertial navigation system measured ncarry out integration and obtain north orientation speed
Figure FDA0000378652330000011
(2) by north orientation speed
Figure FDA0000378652330000012
carrier true north speed v with the GPS acquisition ndo the poor north orientation velocity error that obtains
Figure FDA00003786523300000123
? ▿ v n = v ^ n - v n ;
(3) north orientation speed correction link controller is according to the north orientation velocity error
Figure FDA00003786523300000124
generate latitude controlled quentity controlled variable u vn;
(4) utilize latitude controlled quentity controlled variable u vnto north orientation acceleration a nrevised, through revised north orientation acceleration a nagain carry out integration and obtain revised north orientation speed;
(5) utilize
Figure FDA0000378652330000014
by revised north orientation speed
Figure FDA0000378652330000015
be converted into the latitude rate of change
Figure FDA00003786523300000127
Figure FDA0000378652330000016
(6) to the latitude rate of change
Figure FDA0000378652330000017
carry out integration and obtain latitude
Figure FDA0000378652330000018
by latitude
Figure FDA0000378652330000019
carrier true latitude with the GPS acquisition do the poor latitude error that obtains
Figure FDA00003786523300000111
?
Figure FDA00003786523300000112
(7) latitude correction link controller is according to latitude error
Figure FDA00003786523300000113
generate controlled quentity controlled variable
Figure FDA00003786523300000114
(8) utilize controlled quentity controlled variable
Figure FDA00003786523300000115
to the latitude rate of change
Figure FDA00003786523300000116
revised, through revised latitude rate of change
Figure FDA00003786523300000117
again carry out integration and obtain revised latitude revised latitude
Figure FDA00003786523300000119
north orientation passage as inertial navigation system is exported, thereby completes the correction of north orientation channel measurement error;
East orientation channel measurement error correction step:
(1) east orientation acceleration a inertial navigation system measured ecarry out integration and obtain north orientation speed
Figure FDA00003786523300000120
(2) by east orientation speed
Figure FDA00003786523300000121
the true east orientation speed v of carrier with the GPS acquisition edo the poor east orientation velocity error that obtains
Figure FDA00003786523300000125
? ▿ v e = v ^ e - v e ;
(3) east orientation speed correction link controller is according to the east orientation velocity error
Figure FDA00003786523300000126
generate longitude controlled quentity controlled variable u ve;
(4) utilize longitude controlled quentity controlled variable u veto east orientation acceleration a erevised, through revised east orientation acceleration a eagain carry out integration and obtain revised east orientation speed;
(5) utilize
Figure FDA0000378652330000021
by revised east orientation speed
Figure FDA0000378652330000022
be converted into the longitude rate of change
Figure FDA0000378652330000024
(6) to the longitude rate of change carry out integration and obtain longitude
Figure FDA0000378652330000026
by longitude
Figure FDA0000378652330000027
the true longitude λ of carrier obtained with GPS does the poor longitude error that obtains ?
Figure FDA0000378652330000028
(7) longitude correction link controller is according to longitude error
Figure FDA00003786523300000217
generate longitude controlled quentity controlled variable u λ;
(8) utilize longitude controlled quentity controlled variable u λto the longitude rate of change
Figure FDA0000378652330000029
revised, through revised longitude rate of change
Figure FDA00003786523300000210
again carry out integration and obtain revised longitude
Figure FDA00003786523300000211
revised longitude
Figure FDA00003786523300000212
east orientation passage as inertial navigation system is exported, thereby completes the correction of east orientation channel measurement error.
2. a kind of inertial navigation system measuring error modification method based on GPS information according to claim 1, it is characterized in that: the transport function of described north orientation speed correction link controller, latitude correction link controller, east orientation speed correction link controller and longitude correction link controller is respectively:
Figure FDA00003786523300000218
represent respectively north orientation speed correction link controller, latitude correction link controller, east orientation speed correction link controller and longitude correction link controller;
Wherein v is the integral element number, v=0 or 1; The bandwidth in handicapping Buddhist nun loop is ω c, when v=0, T 1 ≥ T 2 > 1 ω c , When v=1, T 2 > 1 ω c > T 1 ;
T 1for the inertial element time constant;
T 2for the first derivative element time constant;
K is enlargement factor.
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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN105241319A (en) * 2015-08-27 2016-01-13 北京航天控制仪器研究所 High-speed self-rotation guided cartridge aerial real-time alignment method
CN106595709A (en) * 2016-12-07 2017-04-26 北京航天控制仪器研究所 Inertial navigation system measuring error correction method based on external measuring information
CN111457921A (en) * 2020-04-10 2020-07-28 江西驰宇光电科技发展有限公司 Tunnel structure safety monitoring device and method based on laser inertial navigation system

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0649034A2 (en) * 1993-10-18 1995-04-19 Hughes Aircraft Company SAR/GPS inertial method of range measurement
JPH10104015A (en) * 1996-10-01 1998-04-24 Yokogawa Denshi Kiki Kk Navigation apparatus
US20020116126A1 (en) * 2000-12-23 2002-08-22 Ching-Fang Lin Positioning and data integrating method and system thereof
CN101046387A (en) * 2006-08-07 2007-10-03 南京航空航天大学 Scene matching method for raising navigation precision and simulating combined navigation system
CN101109959A (en) * 2007-08-06 2008-01-23 北京航空航天大学 Attitude determining system of mini system suitable for any motion
CN101718560A (en) * 2009-11-20 2010-06-02 哈尔滨工程大学 Strapdown system error inhibition method based on uniaxial four-position rotation and stop scheme

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0649034A2 (en) * 1993-10-18 1995-04-19 Hughes Aircraft Company SAR/GPS inertial method of range measurement
JPH10104015A (en) * 1996-10-01 1998-04-24 Yokogawa Denshi Kiki Kk Navigation apparatus
US20020116126A1 (en) * 2000-12-23 2002-08-22 Ching-Fang Lin Positioning and data integrating method and system thereof
CN101046387A (en) * 2006-08-07 2007-10-03 南京航空航天大学 Scene matching method for raising navigation precision and simulating combined navigation system
CN101109959A (en) * 2007-08-06 2008-01-23 北京航空航天大学 Attitude determining system of mini system suitable for any motion
CN101718560A (en) * 2009-11-20 2010-06-02 哈尔滨工程大学 Strapdown system error inhibition method based on uniaxial four-position rotation and stop scheme

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
张源等: "《GPS姿态测量系统对惯性导航系统误差修正能力分析》", 《 情报指挥控制系统与仿真技术 》 *
张源等: "《基于卫星导航技术惯性导航系统误差评估 》", 《 微计算机信息》 *

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN105241319A (en) * 2015-08-27 2016-01-13 北京航天控制仪器研究所 High-speed self-rotation guided cartridge aerial real-time alignment method
CN106595709A (en) * 2016-12-07 2017-04-26 北京航天控制仪器研究所 Inertial navigation system measuring error correction method based on external measuring information
CN106595709B (en) * 2016-12-07 2019-09-06 北京航天控制仪器研究所 A kind of inertial navigation system measurement error modification method based on metric information
CN111457921A (en) * 2020-04-10 2020-07-28 江西驰宇光电科技发展有限公司 Tunnel structure safety monitoring device and method based on laser inertial navigation system

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