CN103471593B - A kind of inertial navigation system measuring error modification method based on GPS information - Google Patents

A kind of inertial navigation system measuring error modification method based on GPS information Download PDF

Info

Publication number
CN103471593B
CN103471593B CN201310403697.2A CN201310403697A CN103471593B CN 103471593 B CN103471593 B CN 103471593B CN 201310403697 A CN201310403697 A CN 201310403697A CN 103471593 B CN103471593 B CN 103471593B
Authority
CN
China
Prior art keywords
longitude
latitude
error
revised
inertial navigation
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
CN201310403697.2A
Other languages
Chinese (zh)
Other versions
CN103471593A (en
Inventor
魏宗康
刘生炳
赵龙
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
China Aerospace Times Electronics Corp
Original Assignee
China Aerospace Times Electronics Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by China Aerospace Times Electronics Corp filed Critical China Aerospace Times Electronics Corp
Priority to CN201310403697.2A priority Critical patent/CN103471593B/en
Publication of CN103471593A publication Critical patent/CN103471593A/en
Application granted granted Critical
Publication of CN103471593B publication Critical patent/CN103471593B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Abstract

The invention discloses a kind of inertial navigation system measuring error modification method based on GPS information, the present invention proposes the method adopting series connection to control, introduce external GPS positional information, velocity information is to revise inertial navigation error, filtering to input signal noise can be completed by designing suitable controller, completion system exports the tracking to system input, namely the correction to inertial navigation system speed and site error is completed, the method does not rely on the accuracy of SYSTEM ERROR MODEL, fast operation, time overhead is short, also the correction to inertial navigation system speed and errors in position measurement can be completed when unit discharging frequency is higher, the error model of system only need can be re-established by change control device parameter, workload is little.

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, particularly relate to a kind of inertial navigation system measuring error modification method based on GPS information, can be used for the occasion with high precision navigation accuracy, such as the field such as Aero-Space, mapping.
Background technology
Inertial navigation system is a voyage Estimation System based on acceleration quadratic integral, and it relies on plant equipment completely and corresponding algorithm is automatic, complete independently navigation task, and any optical, electrical contact does not occur in the external world.Therefore good concealment, working environment is not by the restriction of meteorological condition.The advantage of this uniqueness, becomes a kind of widely used prime navaid 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 improving inertial navigation system performance adopts integrated navigation technology, namely by two or more non-similar navigational system same navigation information measured and resolved and measure with formation volume, from these measurement amounts, calculate the error of each navigational system and correct.Adopt the system of integrated navigation technology to be called integrated navigation system, each system involved 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, carries out fusion obtain high-precision navigation output by different navigation system acquisition navigation information.Information fusion algorithm common is at present generally theoretical based on optimal estimation, such as kalman filtering theory and its derivative algorithm.The error model of system is needed to set up when using Kalman filtering algorithm to carry out Design of Integrated Navigation System, in addition, Kalman filtering algorithm is a kind of recursive algorithm, parameters initial value is needed during starting algorithm, the correctness of error model and the rationality of initial parameter values will have a direct impact convergence, need during engineer applied to arrange rational initial parameter values for different product, debugging work load is very large.In addition, along with the raising of metric information frequency, will be difficult to have calculated integrated navigation after adopting Kalman filter and calculate in one-period.
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, realize the correction to inertial navigation system speed and position, improve 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) to the north orientation acceleration a that inertial navigation system is measured ncarry out integration and obtain north orientation speed
(2) by north orientation speed the carrier true north speed v obtained with GPS ndo difference and obtain north orientation velocity error namely ▿ v n = v ^ n - v n ;
(3) north orientation speed correction link controller is according to north orientation velocity error generate latitude controlled quentity controlled variable u vn;
(4) latitude controlled quentity controlled variable u is utilized vnto north orientation acceleration a nrevise, through revised north orientation acceleration a nagain carry out integration and obtain revised north orientation speed;
(5) utilize by revised north orientation speed be converted into latitude rate of change
(6) to latitude rate of change carry out integration and obtain latitude by latitude the carrier true latitude obtained with GPS do difference and obtain latitude error namely
(7) latitude correction link controller is according to latitude error generate controlled quentity controlled variable
(8) controlled quentity controlled variable is utilized to latitude rate of change revise, through revised latitude rate of change again carry out integration and obtain revised latitude revised latitude north orientation passage as inertial navigation system exports, thus completes the correction of north orientation channel measurement error;
East orientation channel measurement error correction step:
(1) to the east orientation acceleration a that inertial navigation system is measured ecarry out integration and obtain north orientation speed
(2) by east orientation speed the true east orientation speed v of the carrier obtained with GPS emake it so poor that to obtain east orientation velocity error namely ▿ v e = v ^ e - v e ;
(3) east orientation speed correction link controller is according to east orientation velocity error generate longitude controlled quentity controlled variable u ve;
(4) longitude controlled quentity controlled variable u is utilized veto east orientation acceleration a erevise, through revised east orientation acceleration a eagain carry out integration and obtain revised east orientation speed;
(5) utilize by revised east orientation speed be converted into longitude rate of change
(6) to longitude rate of change carry out integration and obtain longitude by longitude the true longitude λ of the carrier obtained with GPS does difference and obtains longitude error namely
(7) longitude correction link controller is according to longitude error generate longitude controlled quentity controlled variable u λ;
(8) longitude controlled quentity controlled variable u is utilized λto longitude rate of change revise, through revised longitude rate of change again carry out integration and obtain revised longitude revised longitude east orientation passage as inertial navigation system exports, thus 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 north orientation speed correction link controller, latitude correction link controller, east orientation speed correction link controller and longitude correction link controller respectively;
Wherein v is integral element number, v=0 or 1; If the bandwidth of damping circuit is ω c, as v=0, T 1 ≥ T 2 > 1 ω c , As v=1, T 2 > 1 ω c > T 1 ;
T 1for inertial element time constant;
T 2for first derivative element time constant;
K is enlargement factor.
Ultimate principle of the present invention: when obtaining actual position information (longitude λ, the latitude of carrier ), velocity information (east orientation speed v e, north orientation speed v n) time, the positional information (longitude that inertial navigation system can be exported latitude ) and velocity information (east orientation speed north orientation speed ) do difference with actual position information and velocity information respectively, form error signal, namely this error signal formation control amount after controller is applied to position integral equation and the rate integrating equation of inertial navigation respectively, realizes the damping to INS errors.
The present invention's advantage is compared with prior art as follows:
(1) existing Kalman filtering algorithm needs the error model setting up 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 adopting series connection to control is proposed, introduce external GPS positional information, velocity information to revise inertial navigation error, filtering to input signal noise can be completed by designing suitable controller, completion system exports the tracking to system input, namely completes the correction to inertial navigation system speed and site error.
(2) existing Kalman filtering algorithm time overhead is comparatively large, and the filtering cycle is longer, method fast operation of the present invention, and time overhead is short, also can complete the correction to inertial navigation system speed and errors in position measurement when unit discharging frequency is higher.
(3) if existing Kalman filtering algorithm information source changes, need the error model re-establishing system, workload is comparatively large, and the present invention only need by change control device parameter.
Accompanying drawing explanation
Fig. 1 of the present inventionly realizes schematic diagram;
Fig. 2 is the damping circuit open-loop transfer function Bode diagram of adoption rate Absent measures device of the present invention;
Fig. 3 is adoption rate-delayed-differentiation element controller damping circuit open-loop transfer function Bode diagram;
Fig. 4 is adoption rate-integration-advanced-delay component controller damping circuit open-loop transfer function Bode diagram;
Fig. 5 is the Output rusults figure of the integrated navigation system using the present invention's design;
Fig. 6 is the integrated navigation system position navigation curve map using the present invention's design.
Embodiment
Realization approach of the present invention is: 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 ), velocity information (east orientation speed v e, north orientation speed v n) time, the positional information (longitude that inertial navigation system can be exported latitude ) and velocity information (east orientation speed north orientation speed ) do difference with actual position information and velocity information respectively, form error signal, namely this error signal formation control amount after controller is applied to position integral equation and the rate integrating equation of inertial navigation respectively, 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) to the north orientation acceleration a that inertial navigation system is measured ncarry out integration and obtain north orientation speed
(2) by north orientation speed the carrier true north speed v obtained with GPS ndo difference and obtain north orientation velocity error namely ▿ v n = v ^ n - v n ;
(3) north orientation speed correction link controller is according to north orientation velocity error generate latitude controlled quentity controlled variable u vn;
(4) latitude controlled quentity controlled variable u is utilized vnto north orientation acceleration a nrevise, through revised north orientation acceleration a nagain carry out integration and obtain revised north orientation speed;
(5) utilize by revised north orientation speed be converted into latitude rate of change
(6) to latitude rate of change carry out integration and obtain latitude by latitude the carrier true latitude obtained with GPS do difference and obtain latitude error namely
(7) latitude correction link controller is according to latitude error generate controlled quentity controlled variable
(8) controlled quentity controlled variable is utilized to latitude rate of change revise, through revised latitude rate of change again carry out integration and obtain revised latitude revised latitude north orientation passage as inertial navigation system exports, thus completes the correction of north orientation channel measurement error;
East orientation channel measurement error correction step:
(1) to the east orientation acceleration a that inertial navigation system is measured ecarry out integration and obtain north orientation speed
(2) by east orientation speed the true east orientation speed v of the carrier obtained with GPS edo difference and obtain east orientation velocity error namely ▿ v e = v ^ e - v e ;
(3) east orientation speed correction link controller is according to east orientation velocity error generate longitude controlled quentity controlled variable u ve;
(4) longitude controlled quentity controlled variable u is utilized veto east orientation acceleration a erevise, through revised east orientation acceleration a eagain carry out integration and obtain revised east orientation speed;
(5) utilize by revised east orientation speed be converted into longitude rate of change
(6) to longitude rate of change carry out integration and obtain longitude by longitude the true longitude λ of the carrier obtained with GPS does difference and obtains longitude error namely
(7) longitude correction link controller is according to longitude error generate longitude controlled quentity controlled variable u λ;
(8) longitude controlled quentity controlled variable u is utilized λto longitude rate of change revise, through revised longitude rate of change again carry out integration and obtain revised longitude revised longitude east orientation passage as inertial navigation system exports, thus 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 north orientation speed correction link controller, latitude correction link controller, east orientation speed correction link controller and longitude correction link controller respectively;
Wherein v is integral element number, v=0 or 1; If the bandwidth of damping circuit is ω c, as v=0, T 1 ≥ T 2 > 1 ω c , As v=1, T 2 > 1 ω c > T 1 ;
T 1for inertial element time constant;
T 2for first derivative element time constant;
K is enlargement factor.
Due to controlling unit C 1(s), C 2(s), C 3(s) and C 4s () is continuous system, need carry out sliding-model control when practical application, and discretization method is Bilinear transformation method, even carry out discretize, wherein Δ T is the time discretization of navigational system.
Example 1. is as v=0, if T 1=T 2, then have first, the shearing frequency ω of certainty annuity c, then the open-loop transfer function of damping system is wherein, ω c=K.Fig. 2 solid line gives ω cthe open loop Bode diagram of damping system during=100rad/s, if sample frequency is 100Hz, the system open loop Bode diagram after adopting bilinearity discretize as Fig. 2 in dotted line.
Example 2. as v=0, if T 1 ≥ T 2 > 1 ω c , Then have C i ( s ) = K T 2 s + 1 T 1 s + 1 , i = 1,2,3,4 . First, the shearing frequency ω of certainty annuity c, get then the open-loop transfer function of damping system is fig. 3 solid line gives ω c=100rad/s, T 1=1, T 2damping system open loop Bode diagram when=0.1.As can be seen from the figure, system is improved in the gain of low-frequency range, and its multiple is if sample frequency is 100Hz, the system open loop Bode diagram after adopting bilinearity discretize as Fig. 3 in dotted line.
Example 3. as v=1, if T 2 > 1 ω c > T 1 , Then have C i ( s ) = K T 2 s + 1 s ( T 1 s + 1 ) , i = 1,2,3,4 . First, the shearing frequency ω of certainty annuity c, get K=T 2ω c 2, then the open-loop transfer function of damping system is fig. 4 solid line gives 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 bilinearity discretize as Fig. 4 in dotted line.
The inertial navigation system damping method Design of Integrated Navigation System based on external position and velocity information can be completed by said method, Fig. 5 the first row is respectively longitude curve, latitude curve and altitude curve, second row is respectively 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 GPS navigation result, dotted line is integrated navigation system navigation results, can find out when GPS does not have frame losing, article two, curve almost overlaps completely, the method of the invention achieves the correction of GPS to inertial navigation system measuring error, and the noise of integrated navigation system is less, Fig. 6 is the navigation position curve of integrated navigation system, can find out that integrated navigation system position navigation results conforms to actual when GPS does not have frame losing.
The non-detailed description of the present invention is known to the skilled person technology.

Claims (2)

1., based on an inertial navigation system measuring error modification method for GPS information, it 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) to the north orientation acceleration a that inertial navigation system is measured ncarry out integration and obtain north orientation speed
(2) by north orientation speed the carrier true north speed v obtained with GPS ndo difference and obtain north orientation velocity error namely
(3) north orientation speed correction link controller is according to north orientation velocity error generate latitude controlled quentity controlled variable u vn;
(4) latitude controlled quentity controlled variable u is utilized vnto north orientation acceleration a nrevise, through revised north orientation acceleration a nagain carry out integration and obtain revised north orientation speed;
(5) utilize will through the revised north orientation speed of step (4) be converted into latitude rate of change
(6) to latitude rate of change carry out integration and obtain latitude by latitude the carrier true latitude obtained with GPS do difference and obtain latitude error namely
(7) latitude correction link controller is according to latitude error generate controlled quentity controlled variable
(8) controlled quentity controlled variable is utilized to latitude rate of change revise, through revised latitude rate of change again carry out integration and obtain revised latitude revised latitude north orientation passage as inertial navigation system exports, thus completes the correction of north orientation channel measurement error;
East orientation channel measurement error correction step:
(1) to the east orientation acceleration a that inertial navigation system is measured ecarry out integration and obtain east orientation speed
(2) by east orientation speed the true east orientation speed v of the carrier obtained with GPS edo difference and obtain east orientation velocity error namely
(3) east orientation speed correction link controller is according to east orientation velocity error generate longitude controlled quentity controlled variable u ve;
(4) longitude controlled quentity controlled variable u is utilized veto east orientation acceleration a erevise, through revised east orientation acceleration a eagain carry out integration and obtain revised east orientation speed;
(5) utilize will through the revised east orientation speed of step (4) be converted into longitude rate of change
(6) to longitude rate of change carry out integration and obtain longitude by longitude the true longitude λ of the carrier obtained with GPS does difference and obtains longitude error namely
(7) longitude correction link controller is according to longitude error generate longitude controlled quentity controlled variable u λ;
(8) longitude controlled quentity controlled variable u is utilized λto longitude rate of change revise, through revised longitude rate of change again carry out integration and obtain revised longitude revised longitude east orientation passage as inertial navigation system exports, thus 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, 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: represent north orientation speed correction link controller, latitude correction link controller, east orientation speed correction link controller and longitude correction link controller respectively;
Wherein v is integral element number, v=0 or 1; If the bandwidth of damping circuit is ω c, as v=0, as v=1,
T 1for inertial element time constant;
T 2for first derivative element time constant;
K is enlargement factor.
CN201310403697.2A 2013-09-06 2013-09-06 A kind of inertial navigation system measuring error modification method based on GPS information Active CN103471593B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN201310403697.2A CN103471593B (en) 2013-09-06 2013-09-06 A kind of inertial navigation system measuring error modification method based on GPS information

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN201310403697.2A CN103471593B (en) 2013-09-06 2013-09-06 A kind of inertial navigation system measuring error modification method based on GPS information

Publications (2)

Publication Number Publication Date
CN103471593A CN103471593A (en) 2013-12-25
CN103471593B true CN103471593B (en) 2015-12-23

Family

ID=49796544

Family Applications (1)

Application Number Title Priority Date Filing Date
CN201310403697.2A Active CN103471593B (en) 2013-09-06 2013-09-06 A kind of inertial navigation system measuring error modification method based on GPS information

Country Status (1)

Country Link
CN (1) CN103471593B (en)

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN105241319B (en) * 2015-08-27 2016-11-30 北京航天控制仪器研究所 A kind of guided cartridge of spin at a high speed real-time alignment methods in the air
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

Family Cites Families (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5432520A (en) * 1993-10-18 1995-07-11 Hughes Aircraft Company SAR/GPS inertial method of range measurement
JP3338304B2 (en) * 1996-10-01 2002-10-28 横河電子機器株式会社 Navigation equipment
US6427122B1 (en) * 2000-12-23 2002-07-30 American Gnc Corporation 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
CN100587641C (en) * 2007-08-06 2010-02-03 北京航空航天大学 A kind of attitude determination system that is applicable to the arbitrary motion mini system
CN101718560B (en) * 2009-11-20 2011-11-16 哈尔滨工程大学 Strapdown system error inhibition method based on uniaxial four-position rotation and stop scheme

Also Published As

Publication number Publication date
CN103471593A (en) 2013-12-25

Similar Documents

Publication Publication Date Title
CN102608596B (en) Information fusion method for airborne inertia/Doppler radar integrated navigation system
CN103344259B (en) A kind of INS/GPS integrated navigation system feedback correction method estimated based on lever arm
CN104457446B (en) A kind of aerial Alignment Method of the guided cartridge that spins
CN103017755B (en) A kind of underwater navigation attitude measurement method
CN103217699B (en) Integrated navigation system recursion optimizing initial-alignment method based on polarization information
CN104344836B (en) Posture observation-based redundant inertial navigation system fiber-optic gyroscope system level calibration method
CN103674059A (en) External measured speed information-based horizontal attitude error correction method for SINS (serial inertial navigation system)
CN101246012B (en) Combinated navigation method based on robust dissipation filtering
CN101858748A (en) Fault-tolerance autonomous navigation method of multi-sensor of high-altitude long-endurance unmanned plane
CN103196448A (en) Airborne distributed inertial attitude measurement system and transfer alignment method of airborne distributed inertial attitude measurement system
CN103630136A (en) Optimum navigational parameter fusion method based on three-level filtering under redundant sensor configuration
CN103900566B (en) A kind of eliminate the method that rotation modulation type SINS precision is affected by rotational-angular velocity of the earth
CN103900574A (en) Attitude estimation method based on iteration volume Kalman filter
RU2380656C1 (en) Integrated strapdown inertial and satellite navigation system on coarse sensors
CN103604430A (en) Marginalized cubature Kalman filter (CKF)-based gravity aided navigation method
CN103644911A (en) Gyroscope assisted positioning method
CN103900559A (en) High precision attitude resolving system based on interference estimation
CN103557864A (en) Initial alignment method for micro electro mechanical system (MEMS) strap-down inertial navigation adaptive square-root cubature Kalman filtering (SCKF)
CN103712598A (en) Attitude determination system and method of small unmanned aerial vehicle
CN102519485A (en) Gyro information-introduced double-position strapdown inertial navigation system initial alignment method
CN102393204B (en) Combined navigation information fusion method based on SINS (Ship's Inertial Navigation System)/CNS (Communication Network System)
CN106443062B (en) Unmanned plane speed measurement method, device and unmanned plane
CN103471593B (en) A kind of inertial navigation system measuring error modification method based on GPS information
CN101929862A (en) Method for determining initial attitude of inertial navigation system based on Kalman filtering
CN104597460A (en) Beidou satellite navigation receiver based carrier wave tracking loop crystal oscillator acceleration speed sensitivity coefficient calibration method

Legal Events

Date Code Title Description
C06 Publication
PB01 Publication
C10 Entry into substantive examination
SE01 Entry into force of request for substantive examination
C14 Grant of patent or utility model
GR01 Patent grant