CN103090866B - Method for restraining speed errors of single-shaft rotation optical fiber gyro strapdown inertial navigation system - Google Patents

Method for restraining speed errors of single-shaft rotation optical fiber gyro strapdown inertial navigation system Download PDF

Info

Publication number
CN103090866B
CN103090866B CN201310005528.3A CN201310005528A CN103090866B CN 103090866 B CN103090866 B CN 103090866B CN 201310005528 A CN201310005528 A CN 201310005528A CN 103090866 B CN103090866 B CN 103090866B
Authority
CN
China
Prior art keywords
amp
omega
phi
represent
carrier
Prior art date
Application number
CN201310005528.3A
Other languages
Chinese (zh)
Other versions
CN103090866A (en
Inventor
王秋滢
齐昭
高峰
高伟
孙枫
Original Assignee
哈尔滨工程大学
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
Priority to CN201210432160 priority Critical
Priority to CN2012104321604 priority
Priority to CN201210432160.4 priority
Application filed by 哈尔滨工程大学 filed Critical 哈尔滨工程大学
Priority to CN201310005528.3A priority patent/CN103090866B/en
Publication of CN103090866A publication Critical patent/CN103090866A/en
Application granted granted Critical
Publication of CN103090866B publication Critical patent/CN103090866B/en

Links

Abstract

The invention discloses a method for restraining speed errors of a single-shaft rotation optical fiber gyro strapdown inertial navigation system. The method comprises the following steps of: 1. acquiring carrier position information through a global positioning system (GPS); 2. acquiring data output by an optical fiber gyro and a quartz accelerometer; 3. driving an inertial component to perform single-shaft forward and backward rotation and stop motion by a rotary mechanism, and adopting a rotation scheme that eight rotation and stop sequences are used as a rotation period; 4. acquiring information of linear velocity and angular velocity of the motion of measurement carriers of the optical fiber gyro and the quartz accelerometer, and performing navigation solution to obtain navigation information; and 5. constructing a Butterworth band elimination filter, and performing Butterworth filter processing on the carrier velocity obtained under a navigation system. According to the method, the Butterworth band elimination filter is designed according to the rotation angular velocity, oscillation error items related to the rotation angular velocity under the navigation system are filtered and removed, and the velocity information accuracy is improved.

Description

A kind of single-shaft-rotation fiber-optic gyroscope strapdown inertial navigation system velocity error suppressing method

Technical field

The present invention relates to a kind of fertile hereby based on Butterworth(Bart) the single-shaft-rotation fiber-optic gyroscope strapdown inertial navigation system velocity error suppressing method of digital filtering, belong in field of inertia technology the suppressing method reducing navigation information error.

Background technology

Strapdown inertial navitation system (SINS) (SINS) is the full autonomous navigation system of one be directly installed on by inertia assembly (gyroscope and accelerometer) on carrier.It is without any need for external information, utilizes and gathers line motion and the angular motion information that carrier movement measured by inertia assembly, exports the speed of carrier, position and attitude information continuously through navigation calculation.Due to SINS have volume little, lightweight, be easy to safeguard, high reliability is widely used in the field such as Aeronautics and Astronautics, navigation.But due to the existence of inertia assembly constant value deviation, causing system positioning error to be dispersed in time and constantly increasing is one of restriction SINS key factor of navigating for a long time.

In order to improve system accuracy, inertance element precision can be improved on the one hand, but due to the restriction by process technology level, unconfined raising components accuracy is difficult to realize; Be exactly the error suppression technology taking strapdown inertial navigation system on the other hand, the error of automotive resistance inertia device is on the impact of system accuracy.The inertance element so just can applying existing precision forms the strapdown inertial navigation system of degree of precision.

Rotation modulation technology is a kind of inertia device deviation method of self compensation, and the method drives the regular rotation of inertia assembly by rotating mechanism, offsets the impact of this error term on system to the modulation of inertia device constant value deviation, and then improves system accuracy.Although rotation modulation can suppress positioning error to be dispersed effectively, bring new error to navigation informations such as speed again simultaneously, constrain the availability of velocity information.Strapdown inertial navitation system (SINS) for being operated in single shaft rotating and stopping rotation approach: in rotation process, system is resolved in velocity error, has occurred the new oscillation error relevant with rotary rpm; In off-position process, behind rotary motion conversion IMU off-position position, device constant value deviation changes to some extent in the projection of navigation system, thus makes to introduce new disturbance in system, causes velocity error to encourage Schuler, earth periodic oscillation again.That is, often convert an off-position position, within the off-position time period, will excited oscillation error again.

In CNKI storehouse, open report has: 1. " system-level dual-axis rotation modulation inertial navigation error analysis and calibration ", this article mainly proposes a kind of system-level dual-axis rotation modulation system inertial navigation Project Realization scheme, have found the error source of navigation accuracy when influential system is long to navigate.2. " analysis of rotary laser gyro inertial navigation system error Propagation Property ", this article mainly conducts in-depth research the error Propagation Property of single shaft rotary laser gyro inertial navigation system, stop rotation approach for single shaft rotating, each error term modulation effect is analyzed.3. " Rotating Inertial Navigation System shaft direction is on the impact of system modulation precision ", herein Main Analysis single-shaft-rotation inertial navigation system intermediate station rotor shaft direction on the impact of system accuracy.Above document is all to suppress divergence expression positioning error, does not mention that rotation modulation resolves the impact of velocity information precision and applicability to system.

Summary of the invention

The object of the invention is to solve the problem, a kind of single-shaft-rotation fiber-optic gyroscope strapdown inertial navigation system velocity error suppressing method is proposed, based on Butterworth digital filter, using rotation modulation inertial navigation system computing speed information as input, by being descend the oscillation error item relevant with angular velocity of rotation according to navigation system in the Butterworth digital filter filtering velocity error that design with angular velocity of rotation, improve velocity accuracy, enhancing system resolves the applicability of velocity information.

A kind of single-shaft-rotation fiber-optic gyroscope strapdown inertial navigation system velocity error suppressing method, comprises the following steps:

Step one: gather carrier positions information by global location gps system, and bookbinding is in navigational computer;

Step 2: rotating mechanism is turned to the position that IMU system overlaps with carrier system, have wherein b represents carrier coordinate system, and s represents IMU coordinate system, represent that s is tied to b system transition matrix, I representation unit battle array; After fiber-optic gyroscope strapdown inertial navigation system is carried out abundant preheating, gather the data of fibre optic gyroscope and quartz accelerometer output; Obtain angular motion information and the line movable information of carrier; Angular motion information comprises magnitude of angular velocity, and line movable information comprises and compares force value;

Step 3: it is dynamic that rotating mechanism drives inertia assembly to carry out single shaft rotating stoppage in transit with ω; Adopt eight to turn and stop the rotation approach that order is a swing circle;

Step 4: Real-time Collection fibre optic gyroscope and quartz accelerometer measure linear velocity and the angular velocity information of carrier movement, and navigation calculation obtains navigation information;

Step 5: structure Butterworth rejection filter, the bearer rate obtained under navigation system is carried out Butterworth filter process, the lower oscillation error item relevant with angular velocity of rotation of filtering navigation system, filtered speed is as final navigation calculation output information.

The invention has the advantages that:

The present invention is directed to modulation type strapdown inertial navitation system (SINS) and resolve navigation information medium velocity errors of form, after the concrete form drawing modulated process medium velocity error, Butterworth rejection filter is devised by angular velocity of rotation, the bearer rate calculated under navigation system is carried out rejection filter process, the lower oscillation error item relevant with angular velocity of rotation of filtering navigation system, filtered speed, as final navigation calculation output information, improves velocity information precision, enhances the applicability of this information.

Accompanying drawing explanation

Fig. 1 is method flow diagram of the present invention;

Fig. 2 is IMU four-position rotation and stop schematic diagram in step 3 of the present invention;

Fig. 2 a is 1. ~ 4. rotary course; Fig. 2 b is 5. ~ 8. rotary course;

Fig. 3 is the Butterworth wave filter amplitude-frequency response of step 3 of the present invention structure;

Fig. 4 utilizes Visual C++ to emulate to obtain velocity error comparison curves before and after filtering in the embodiment of the present invention 1;

Fig. 5 utilizes single axle table to test to obtain velocity error comparison curves before and after filtering in the embodiment of the present invention 2.

Embodiment

Below in conjunction with drawings and Examples, the present invention is described in further detail.

The present invention is a kind of single-shaft-rotation fiber-optic gyroscope strapdown inertial navigation system velocity error suppressing method, and method flow as shown in Figure 1, comprises the following steps:

Step one: gather carrier positions information by global location gps system, and bookbinding is in navigational computer;

Navigation initial time, gather initial time carrier positions information, velocity information, and bookbinding is in navigational computer by global location gps system.Carrier positions information comprises longitude, the latitude information of carrier position.

In navigation procedure, utilize this initial information to upgrade, obtain the speed of any time carrier, position.

Step 2: rotating mechanism is turned to the position that IMU system overlaps with carrier system, have wherein b represents carrier coordinate system, and s represents IMU coordinate system, represent that s is tied to b system transition matrix, I representation unit battle array.After fiber-optic gyroscope strapdown inertial navigation system is carried out abundant preheating, gather the data of fibre optic gyroscope and quartz accelerometer output,

Obtain angular motion information and the line movable information of carrier.Angular motion information comprises magnitude of angular velocity, and line movable information comprises and compares force value.

The magnitude of angular velocity that the ratio force value exported according to accelerometer and acceleration of gravity relation and gyroscope export and rotational-angular velocity of the earth relation, determine attitude of carrier angle, completion system initial alignment, set up the initial strap-down matrix of inertial navigation system

C s n = C b n C s b = cos φ y 0 cos φ z 0 - sin φ x 0 sin φ y 0 sin φ z 0 - cos φ x 0 sin φ z 0 sin φ y 0 cos φ z 0 + sin φ z 0 sin φ x 0 cos φ y 0 sin φ z 0 cos φ y 0 + sin φ x 0 sin φ y 0 cos φ z 0 cos φ x 0 cos φ z 0 sin φ z 0 sin φ y 0 + cos φ z 0 sin φ x 0 cos φ y 0 - cos φ x 0 sin φ y 0 sin φ x 0 cos φ x 0 cos φ y 0 - - - ( 1 )

Wherein, φ x0, φ y0, φ z0represent the initial time carrier angle of pitch, roll angle, course angle respectively.

Step 3: it is dynamic that rotating mechanism drives inertia assembly to carry out single shaft rotating stoppage in transit with ω.Wherein, ω=6 °/s can be got.Adopt eight to turn and stop the rotation approach that order is a swing circle;

Described inertia assembly (Inertial Measurement Unit is called for short IMU) rotation process adopts eight turns to stop the transposition scheme that order is a swing circle, is specially:

As shown in Figure 2, wherein (a) transposition schematic diagram that is order 1 to order 4, in figure, 1. ~ and 4. representing front 4 rotary courses, A, B, C, D represent four stop places, x b, y brepresent the transverse axis of carrier coordinate system, and require that rotating initial time IMU overlaps completely with carrier coordinate system.Order 1, IMU rotates clockwise 180 ° of in-position C from A point, stand-by time T r; Order 2, IMU rotates counterclockwise 90 ° of in-position B from C point, stand-by time T r; Order 3, IMU rotates clockwise 180 ° of in-position D from B point, stand-by time T r; Order 4, IMU rotates counterclockwise 270 ° from D point, in-position A, stand-by time T r; As shown in Figure 2, wherein (b) transposition schematic diagram that is order 5 to order 8, in figure, 5. ~ and 8. representing rear 4 rotary courses, A, B, C, D represent four stop places, x b, y brepresent the transverse axis of carrier coordinate system.Order 5, IMU rotates counterclockwise 180 ° of in-position C from A point, stand-by time T r; Order 6, IMU rotates clockwise 90 ° of in-position D from C point, stand-by time T r; Order 7, IMU rotates counterclockwise 180 ° of in-position B from D point, stand-by time T r; Order 8, IMU rotates clockwise 270 ° from B point, in-position A, stand-by time T r; IMU rotates sequential loop according to this to carry out.

Wherein, time T can be got r=300s.

Step 4: Real-time Collection fibre optic gyroscope and quartz accelerometer measure linear velocity and the angular velocity information of carrier movement, and navigation calculation obtains navigation information;

By the gyro data gathered under rotation modulation state, upgrade strap-down matrix be specially:

Angular velocity upgrades:

ω ns s = ω is s - ( C s n ) T ( ω ie s + ω en s ) - - - ( 2 )

Wherein, i represents Earth central inertial system, and e represents terrestrial coordinate system, and s represents inertia assembly coordinate system, and n represents navigational coordinate system, adopts local geographic coordinate system here; represent that s is tied to n system transition matrix; represent that relative m system of p system angular velocity of rotation projects in q system.

Hypercomplex number attitude matrix upgrades:

If any time, the rotation hypercomplex number of carrier coordinate system opposed platforms coordinate system was:

Q=q 0+q 1i b+q 2j b+q 3k b(3)

Wherein, Q is hypercomplex number; q 0, q 1, q 2, q 3for four real numbers of hypercomplex number; i b, j b, k brepresent IMU coordinate system ox respectively saxle, oy saxle, oz sunit direction vector on axle.

The timely correction of hypercomplex number Q:

q · 1 q · 1 q · 2 q · 3 = 1 2 0 - ω nsx s - ω nsy s - ω nsz s ω nsx s 0 ω nsz s - ω nsx s ω nsy s - ω nsz s 0 ω nsx s ω nsz s ω nsy s - ω nsx s 0 q 0 q 1 q 2 q 3 - - - ( 4 )

Wherein, represent that the angular velocity of satellite motion of rotating mechanism Relative Navigation system is at IMU coordinate system ox respectively saxle, oy saxle, oz scomponent on axle. represent q respectively 0, q 1, q 2, q 3rate of change;

According to the rotation hypercomplex number q of k moment carrier coordinate system opposed platforms coordinate system 0(k), q 1(k), q 2(k), q 3(k), asking for the rate of change that the k moment rotates hypercomplex number is:

q · 1 ( k ) q · 1 ( k ) q · 2 ( k ) q · 3 ( k ) = 1 2 0 - ω nsx s - ω nsy s - ω nsz s ω nsx s 0 ω nsz s - ω nsx s ω nsy s - ω nsz s 0 ω nsx s ω nsz s ω nsy s - ω nsx s 0 q 0 ( k ) q 1 ( k ) q 2 ( k ) q 3 ( k ) - - - ( 5 )

Be specially in the rotation hypercomplex number of k+1 moment carrier:

q 0 ( k + 1 ) = q 0 ( k ) + q · 0 ( k ) q 1 ( k + 1 ) = q 1 ( k ) + q · 1 ( k ) q 2 ( k + 1 ) = q 2 ( k ) + q · 2 ( k ) q 3 ( k + 1 ) = q 3 ( k ) + q · 3 ( k ) - - - ( 6 )

Utilize the q obtained 0(k+1), q 1(k+1), q 2(k+1), q 3(k+1), strap-down matrix is upgraded

C s n = q 0 2 + q 1 2 - q 2 2 - q 3 2 2 ( q 1 q 2 - q 0 q 3 ) 2 ( q 1 q 3 + q 0 q 2 ) 2 ( q 1 q 2 + q 0 q 3 ) q 0 2 - q 1 2 + q 2 2 - q 3 2 2 ( q 2 q 3 - q 0 q 1 ) 2 ( q 1 q 3 - q 0 q 2 ) 2 ( q 2 q 3 + q 0 q 1 ) q 0 2 - q 1 2 - q 2 2 + q 3 2 - - - ( 7 )

Wherein, the q in (7) formula i(i=1,2,3,4) are q in (6) formula i(k+1) (i=1,2,3,4), eliminate (k+1) in (7) formula.

Upgrade attitude of carrier information, be specially:

φ x = arcsin ( c 33 ) φ y = arctan ( c 32 / c 31 ) φ z = arctan ( c 13 / c 23 ) - - - ( 8 )

Wherein, c ij(i=1,2,3, j=1,2,3) represent in the i-th row jth column matrix element; φ x, φ y, φ zrepresent carrier pitch angle, roll angle, course angle.

The ratio force information that degree of will speed up meter is measured along IMU coordinate system, passes through strap-down matrix carry out projection transform:

f n = C s n f s - - - ( 9 )

Wherein, f n, f srepresent that accelerometer measures specific force is in n system and the projection of s system respectively;

Utilize following differential equation carrier movement speed:

v · x v · y v · z = f x n f y n f z n - 0 0 g + 0 2 ω iez n - ( 2 ω iey n + ω eny n ) - ω iez n 0 2 ω iex n + ω enx n 2 ω iey n + ω eny n - ( 2 ω iex n + ω enx n ) 0 v x v y v z - - - ( 10 )

Wherein, v x, v y, v zrepresent that resolving bearer rate in navigation is ox respectively naxle, oy naxle, oz ncomponent on axle; represent v x, v y, v zrate of change, namely carrier along navigation be ox naxle, oy naxle, oz nthe acceleration of motion of axle; represent that accelerometer measures specific force is ox in navigation respectively naxle, oy naxle, oz ncomponent on axle; G is acceleration of gravity. represent that rotational-angular velocity of the earth is ox in navigation respectively naxle, oy naxle, oz ncomponent on axle; representing respectively causes navigating due to carrier movement be the angular velocity of rotation of relative earth system change is ox in navigation naxle, oy naxle, oz ncomponent on axle.

According to the carrier east orientation horizontal velocity v in k moment x(k), north orientation horizontal velocity v yk () and sky are to speed v z(k), asking for k moment bearer rate rate of change is:

v · x ( k ) v · y ( k ) v · z ( k ) = f x n f y n f z n - 0 0 g + 0 2 ω iez n - ( 2 ω iey n + ω eny n ) - ω iez n 0 2 ω iex n + ω enx n 2 ω iey n + ω eny n - ( 2 ω iex n + ω enx n ) 0 v x ( k ) v y ( k ) v z ( k ) - - - ( 11 )

Be respectively in k+1 moment bearer rate and position:

v x ( k + 1 ) = v x ( k ) + v · x ( k ) v y ( k + 1 ) = v y ( k ) + v · y ( k ) v z ( k + 1 ) = v z ( k ) + v · z ( k ) - - - ( 12 )

Wherein, R represents earth radius; λ represents the latitude and longitude information that calculate carrier geographic location respectively, as k=1, and v x(1), v y(1), v z(1) the carrier initial velocity utilizing GPS to obtain for step is a kind of, the carrier initial position that λ (1) utilizes GPS to obtain for step is a kind of.

So far, the speed of carrier, position, attitude is obtained according to (8), (12), (13) formula.Then directly export attitude, position to navigational computer display, provide carrier every navigation information; Further filtering process in the pending step 5 of speed is optimized, to reduce its error.

Step 5: structure Butterworth rejection filter, the bearer rate obtained under navigation system is carried out Butterworth filter process, the lower oscillation error item relevant with angular velocity of rotation of filtering navigation system, filtered speed is as final navigation calculation output information.

(1) construct Butterworth rejection filter, be specially:

In step 3, rotating mechanism adopts angular velocity of rotation to be ω, and this angular velocity causes the oscillation error frequency of navigation information to be about f=1/ ω.Set the passband lower-cut-off frequency of Butterworth rejection filter as (0.59 ~ 0.61) f, upper cut-off frequency cutoff frequency is (1.67 ~ 1.69) f, stopband lower-cut-off frequency is (0.83 ~ 0.85) f, stopband upper cut-off frequency is (1.13 ~ 1.15) f, then have:

Ω 1 = ( 0.59 ~ 0.61 ) f Ω 3 = ( 1.67 ~ 1.69 ) f Ω sl = ( 0.83 ~ 0.85 ) f Ω sh = ( 1.13 ~ 1.15 ) f - - - ( 14 )

Wherein, Ω 1represent passband lower-cut-off frequency; Ω 3represent upper cut-off frequency cutoff frequency; Ω slrepresent stopband lower-cut-off frequency; Ω shrepresent stopband upper cut-off frequency.

Ω BW = Ω 3 - Ω 1 Ω 2 2 = Ω 1 Ω 3 - - - ( 15 )

Wherein, Ω bWrepresent pass band width, Ω 2represent stopband center frequency.

Frequency normalization, for:

η 1 = Ω 1 / Ω BW η 3 = Ω 3 / Ω BW η sl = Ω sl / Ω BW η sh = Ω sh / Ω BW η 2 2 = η 1 η 3 = Ω 1 Ω 3 / Ω BW 2 - - - ( 16 )

Wherein, η i(i=1,2,3, sl, sh) represents corresponding frequencies Ω respectively ithe normalized frequency that (i=1,2,3, sl, sh) is corresponding.

The frequency inverted of the corresponding low-pass filter of rejection filter, detailed process is:

λ p = η 3 - η 1 = 1 - λ s ′ = η sl / ( η sl 2 - η 2 2 ) λ s ′ = η sh / ( η sh 2 - η 2 2 ) - - - ( 17 )

Wherein, λ prepresent the cut-off frequecy of passband of low-pass filter; λ sthe stopband lower-cut-off frequency of ' expression low-pass filter;-λ sthe symmetrical frequency of the stopband lower-cut-off frequency of ' expression low-pass filter.

According to (17) formula, get λ swith-λ sone that absolute value is less is final stopband lower-cut-off frequency, i.e. λ s=min{ λ s' ,-λ s'.

Design low-pass filter G (p).Asking for filter order is:

N = lg 10 α s / 10 - 1 10 α p / 10 - 1 / lg λ s - - - ( 18 )

Wherein, N represents the order of wave filter, α srepresent the minimal attenuation that stopband should reach, α prepresent that passband allows maximum attenuation usually to adopt α s=14dB, α p=3dB.

Therefore,

G ( p ) = 1 p N + 1 - - - ( 19 )

Wherein, p represents wave filter parameter.

Parameter fixed relationship between low-pass filter and rejection filter is,

p = s ( Ω 3 - Ω 1 ) s 2 + Ω 1 Ω 3 | s = z - 1 z + 1 = ( z 2 - 1 ) ( Ω 3 - Ω 1 ) ( z - 1 ) 2 + Ω 1 Ω 3 ( z + 1 ) 2 - - - ( 20 )

(20) formula substitutes into (19) formula, constructs the transfer function of discrete type second order band resistance Butterworth wave filter,

H ( z ) = G ( p ) | p = ( z 2 - 1 ) ( Ω 3 - Ω 1 ) ( z - 1 ) 2 + Ω 1 Ω 3 ( z + 1 ) 2 = 1 [ ( z 2 - 1 ) ( Ω 3 - Ω 1 ) ( z - 1 ) 2 + Ω 1 Ω 3 ( z + 1 ) 2 ] N + 1 - - - ( 21 )

According in step 3, if ω=6 °/s, then f=0.1667Hz, the passband lower-cut-off frequency of setting Butterworth rejection filter is 0.01Hz, and upper cut-off frequency cutoff frequency is 0.028Hz, and stopband lower-cut-off frequency is 0.014Hz, stopband upper cut-off frequency is 0.019Hz, tries to achieve λ s=2.9993.Adopt α s=14dB, α pduring=3dB, ask for filter order N=1, the transfer function of final structure discrete type second order band resistance Butterworth wave filter is

H ( z ) = 0.98232 - 1.96355 z - 1 + 0.98177 z - 2 1.00000 - 1.96355 z - 1 + 0.96465 z - 2 - - - ( 22 )

Fig. 3 is the Butterworth wave filter amplitude-frequency response of (22) formula structure.As can be seen from the figure this wave filter can meet the demands, and the information that input information medium frequency is about 0.0167Hz is rejected.

(2) bearer rate under resolving the navigation system that obtains in step 4 is carried out rejection filter process, oscillation error item relevant with angular velocity of rotation in this velocity information of filtering, filtered speed, as final navigation calculation output information, is specially:

Single-shaft-rotation strapdown inertial navitation system (SINS) computing speed is modeled as:

V INS=V+δV INS(r)+δV INS(s)(23)

Wherein, V iNSrepresent the velocity information obtained through navigation calculation in step 4; V represents carrier movement speed; δ V iNSr () represents the oscillation error caused by rotation modulation; δ V iNSs () represents inertial reference calculation Schuler, earth oscillation error and constant error.

Due to place frequency range and other several error place frequency ranges differ greatly, right the wave filter using step (1) to build, obtains:

V′ INS=H(z)·V INS(24)

In formula, H (z) represents the rejection filter built according to angular velocity of rotation; V ' iNSfor filtered speed, wherein only comprise V, δ V iNS (s).

Like this, the velocity fluctuation error because rotation modulation causes just is eliminated.With filtered speed V ' iNSfinally resolve velocity information as system, and export the display device of navigational computer to, provide carrier underway motion velocity information.

Embodiment:

Beneficial effect of the present invention such as under type is verified:

(1) under Visual C++ simulated conditions, emulation experiment is carried out to the method:

Carrier does three-axis swinging motion.Carrier with the deviation from voyage route of sinusoidal rule to angle, pitch angle and roll angle wave, its mathematical model is:

ψ = ψ m sin ( 2 π / T ψ + φ ψ ) + k θ = θ m sin ( 2 π / T θ + φ θ ) γ = γ m sin ( 2 π / T γ + φ γ ) - - - ( 25 )

Wherein, ψ, θ, γ represent the swing angle variable around course angle, pitch angle and roll angle respectively; ψ m, θ m, γ mrepresent corresponding swing angle amplitude respectively, ψ mmm=5 ° of φ ψ, φ θ, φ γrepresent corresponding initial phase respectively, t ψ, T θ, T γrepresent the rolling period of corresponding swinging shaft respectively, T ψ=T θ=T γ=4s; K is true flight path, k=30 °; o

Carrier initial position: north latitude 45.7796 °, east longitude 126.6705 °;

Carrier at the uniform velocity sails through to motion, and movement velocity is v=15m/s;

Equatorial radius: R=6378393.0m;

By the available earth surface acceleration of gravity of universal gravitation: g=9.78049m/s 2;

Rotational-angular velocity of the earth: Ω=7.2921158 × 10 -5rad/s;

Constant: π=3.1415926535;

Fiber optic gyroscope constant drift: 0.01 °/h;

Optical fibre gyro white noise error: 0.005 °/h;

Optical fibre gyro scale factor error: 10ppm;

Optical fibre gyro alignment error: 1 × 10 -3rad;

Accelerometer bias: 10 -4g 0;

Accelerometer white noise error: 5 × 10 -5g 0;

Accelerometer scale factor error: 10ppm;

Accelerometer alignment error: 1 × 10 -3rad;

Simulation time: t=48h;

Sample frequency: Hn=0.01s;

The parameter of IMU four-position rotation and stop scheme:

The dead time of four positions: T r=300s;

Rotate the rotational angular velocity of 180 ° and 90 °: ω=6 °/s;

Rotate in the process of 180 ° and 90 °, speed that the angle in each transposition adds (subtracting): α=3 °/s 2;

Utilize the described method of invention, to obtain before and after filtering velocity error comparison curves as shown in Figure 4, wherein (a) and (b) figure represents the east orientation velocity error before filtering and north orientation velocity error respectively, and (c), (d) figure represents filtered east orientation velocity error and north orientation velocity error respectively.Result show the present invention can suppress modulation condition preferably under system resolve navigation error, improve navigation accuracy, strengthen velocity information availability.

(2) the single axle table test of gyroscope inertial navigation system

The gyroscope inertial navigation system structure pilot system adopting 920E type single shaft test table and develop voluntarily.

Gyroscope inertial navigation system the key technical indexes used is as follows:

Dynamic range: ± 100 °/s;

Bias instaility :≤0.005 °/h;

Random walk:

Scale factory non-linearity degree :≤5ppm.

920E type single axle table platform the key technical indexes is as follows:

Face diameter: 450mm;

Load request: weight 50kg;

Stage body rotating accuracy: ± 2 ' ';

Stage body angle range: unlimited continuously;

Positional precision: ± 3 ' ';

Position resolution power: 0.0001 °;

Speed range: 0.005-200 °/s;

Speed precision: 5 × 10 -5(360 ° average), 5 × 10 -4(10 ° average), 1 × 10 -2(1 ° average).

Utilize the described method of invention, to obtain before and after filtering velocity error comparison curves as shown in Figure 5, wherein (a) and (b) figure represents the east orientation velocity error before filtering and north orientation velocity error respectively, and (c), (d) figure represents filtered east orientation velocity error and north orientation velocity error respectively.Result shows that the present invention suppresses velocity error ability better, can practical requirement.

Claims (1)

1. a single-shaft-rotation fiber-optic gyroscope strapdown inertial navigation system velocity error suppressing method, is characterized in that, comprise the following steps:
Step one: gather carrier positions information by global location gps system, and bookbinding is in navigational computer;
Navigation initial time, gather initial time carrier positions information, velocity information, and bookbinding is in navigational computer by global location gps system; Carrier positions information comprises longitude, the latitude information of carrier position;
In navigation procedure, utilize this initial information to upgrade, obtain the speed of any time carrier, position;
Step 2: rotating mechanism is turned to the position that IMU system overlaps with carrier system, have wherein b represents carrier coordinate system, and s represents IMU coordinate system, represent that s is tied to b system transition matrix, I representation unit battle array; After fiber-optic gyroscope strapdown inertial navigation system is carried out abundant preheating, gather the data of fibre optic gyroscope and quartz accelerometer output; Obtain angular motion information and the line movable information of carrier; Angular motion information comprises magnitude of angular velocity, and line movable information comprises and compares force value;
The magnitude of angular velocity that the ratio force value exported according to accelerometer and acceleration of gravity relation and gyroscope export and rotational-angular velocity of the earth relation, determine attitude of carrier angle, completion system initial alignment, set up the initial strap-down matrix of inertial navigation system
C s n = C b n C s b = cos φ y 0 cos φ z 0 - sin φ x 0 sin φ y 0 sin φ z 0 - cos φ x 0 sin φ z 0 sin φ y 0 cos φ z 0 + sin φ z 0 sin φ x 0 cos φ y 0 sin φ z 0 cos φ y 0 + sin φ x 0 sin φ y 0 cos φ z 0 cos φ x 0 cos φ z 0 sin φ z 0 sin φ y 0 + cos φ z 0 sin φ x 0 cos φ y 0 - cos φ x 0 sin φ y 0 sin φ x 0 cos φ x 0 cos φ y 0 - - - ( 1 )
Wherein, φ x0, φ y0, φ z0represent the initial time carrier angle of pitch, roll angle, course angle respectively;
Step 3: it is dynamic that rotating mechanism drives inertia assembly to carry out single shaft rotating stoppage in transit with ω; Adopt eight to turn and stop the rotation approach that order is a swing circle;
Described inertia assembly is called for short IMU, IMU rotation process and adopts eight turns to stop the transposition scheme that order is a swing circle, is specially:
Order 1, IMU rotates clockwise 180 ° from A point, in-position C, stand-by time T r; Order 2, IMU rotates counterclockwise 90 ° from C point, in-position B, stand-by time T r; Order 3, IMU rotates clockwise 180 ° from B point, in-position D, stand-by time T r; Order 4, IMU rotates counterclockwise 270 ° from D point, in-position A, stand-by time T r; Order 5, IMU rotates counterclockwise 180 ° from A point, in-position C, stand-by time T r; Order 6, IMU rotates clockwise 90 ° from C point, in-position D, stand-by time T r; Order 7, IMU rotates counterclockwise 180 ° from D point, in-position B, stand-by time T r; Order 8, IMU rotates clockwise 270 ° from B point, in-position A, stand-by time T r; IMU rotates sequential loop according to this to carry out;
Step 4: Real-time Collection fibre optic gyroscope and quartz accelerometer measure linear velocity and the angular velocity information of carrier movement, and navigation calculation obtains navigation information;
By the gyro data gathered under rotation modulation state, upgrade strap-down matrix be specially:
Angular velocity upgrades:
ω ns s = ω is s - ( C s n ) T ( ω ie s + ω en s ) - - - ( 2 )
Wherein, i represents Earth central inertial system, and e represents terrestrial coordinate system, and s represents inertia assembly coordinate system, and n represents navigational coordinate system, adopts local geographic coordinate system here; represent that s is tied to n system transition matrix; represent that relative m system of p system angular velocity of rotation is in the projection of q system, m=n, i, e, p=s, e, n, q=s;
Hypercomplex number attitude matrix upgrades:
If any time, the rotation hypercomplex number of carrier coordinate system opposed platforms coordinate system was:
Q=q 0+q 1i b+q 2j b+q 3k b(3)
Wherein, Q is hypercomplex number; q 0, q 1, q 2, q 3for four real numbers of hypercomplex number; i b, j b, k brepresent IMU coordinate system ox respectively saxle, oy saxle, oz sunit direction vector on axle;
The timely correction of hypercomplex number Q:
q · 0 q · 1 q · 2 q · 3 = 1 2 0 - ω nsx s - ω nsy s - ω nsz s ω nsx s 0 ω nsz s - ω nsx s ω nsy s - ω nsz s 0 ω nsx s ω nsz s ω nsy s - ω nsx s 0 q 0 q 1 q 2 q 3 - - - ( 4 )
Wherein, represent that the angular velocity of satellite motion of rotating mechanism Relative Navigation system is at IMU coordinate system ox respectively saxle, oy saxle, oz scomponent on axle; represent q respectively 0, q 1, q 2, q 3rate of change;
According to the rotation hypercomplex number q of k moment carrier coordinate system opposed platforms coordinate system 0(k), q 1(k), q 2(k), q 3(k), asking for the rate of change that the k moment rotates hypercomplex number is:
q · 0 ( k ) q · 1 ( k ) q · 2 ( k ) q · 3 ( k ) = 1 2 0 - ω nsx s - ω nsy s - ω nsz s ω nsx s 0 ω nsz s - ω nsx s ω nsy s - ω nsz s 0 ω nsx s ω nsz s ω nsy s - ω nsx s 0 q 0 ( k ) q 1 ( k ) q 2 ( k ) q 3 ( k ) - - - ( 5 )
Be specially in the rotation hypercomplex number of k+1 moment carrier:
q 0 ( k + 1 ) = q 0 ( k ) + q · 0 ( k ) q 1 ( k + 1 ) = q 1 ( k ) + q · 1 ( k ) q 2 ( k + 1 ) = q 2 ( k ) + q · 2 ( k ) q 3 ( k + 1 ) = q 3 ( k ) + q · 3 ( k ) - - - ( 6 )
Utilize the q obtained 0(k+1), q 1(k+1), q 2(k+1), q 3(k+1), strap-down matrix is upgraded
C s n = q 0 2 + q 1 2 - q 2 2 - q 3 2 2 ( q 1 q 2 - q 0 q 3 ) 2 ( q 1 q 3 + q 0 q 2 ) 2 ( q 1 q 2 + q 0 q 3 ) q 0 2 - q 1 2 + q 2 2 - q 3 2 2 ( q 2 q 3 - q 0 q 1 ) 2 ( q 1 q 3 - q 0 q 2 ) 2 ( q 2 q 3 + q 0 q 1 ) q 0 2 - q 1 2 - q 2 2 + q 3 2 - - - ( 7 )
Wherein, the q in (7) formula ifor q in (6) formula i(k+1), i=1,2,3,4;
Upgrade attitude of carrier information, be specially:
φ x = arcsin ( c 33 ) φ y = arctan ( c 32 / c 31 ) φ z = arctan ( c 13 / c 23 ) - - - ( 8 )
Wherein, c ijrepresent in the i-th row jth column matrix element, i=1,2,3, j=1,2,3; φ x, φ y, φ zrepresent carrier pitch angle, roll angle, course angle;
The ratio force information that degree of will speed up meter is measured along IMU coordinate system, passes through strap-down matrix carry out projection transform:
f n = C s n f s - - - ( 9 )
Wherein, f n, f srepresent that accelerometer measures specific force is in n system and the projection of s system respectively;
Utilize following differential equation carrier movement speed:
v · x v · y v · z = f x n f y n f z n - 0 0 g + 0 2 ω iez n - ( 2 ω iey n + ω eny n ) - ω iez n 0 2 ω iex n + ω enx n 2 ω iey n + ω eny n - ( 2 ω iex n + ω enx n ) 0 v x v y v z - - - ( 10 )
Wherein, v x, v y, v zrepresent that resolving bearer rate in navigation is ox respectively naxle, oy naxle, oz ncomponent on axle; represent v x, v y, v zrate of change, namely carrier along navigation be ox naxle, oy naxle, oz nthe acceleration of motion of axle; represent that accelerometer measures specific force is ox in navigation respectively naxle, oy naxle, oz ncomponent on axle; G is acceleration of gravity; represent that rotational-angular velocity of the earth is ox in navigation respectively naxle, oy naxle, oz ncomponent on axle; representing respectively causes navigating due to carrier movement be the angular velocity of rotation of relative earth system change is ox in navigation naxle, oy naxle, oz ncomponent on axle;
According to the carrier east orientation horizontal velocity v in k moment x(k), north orientation horizontal velocity v yk () and sky are to speed v z(k), asking for k moment bearer rate rate of change is:
v · x ( k ) v · y ( k ) v · z ( k ) = f x n f y n f z n - 0 0 g + 0 2 ω iez n - ( 2 ω iey n + ω eny n ) - ω iez n 0 2 ω iex n + ω enx n 2 ω iey n + ω eny n - ( 2 ω iex n + ω enx n ) 0 v x ( k ) v y ( k ) v z ( k ) - - - ( 11 )
Be respectively in k+1 moment bearer rate and position:
v x ( k + 1 ) = v x ( k ) + v · x ( k ) v y ( k + 1 ) = v y ( k ) + v · y ( k ) v z ( k + 1 ) = v z ( k ) + v · z ( k ) - - - ( 12 )
Wherein, R represents earth radius; λ represents the latitude and longitude information that calculate carrier geographic location respectively, as k=1, and v x(1), v y(1), v z(1) the carrier initial velocity for utilizing GPS to obtain in step one, the carrier initial position of λ (1) for utilizing GPS to obtain in step one;
So far, the speed of carrier, position, attitude is obtained according to (8), (12), (13) formula; Then directly export attitude, position to navigational computer display, provide carrier every navigation information; Further filtering process in the pending step 5 of speed is optimized;
Step 5: structure Butterworth rejection filter, the bearer rate obtained under navigation system is carried out Butterworth filter process, the lower oscillation error item relevant with angular velocity of rotation of filtering navigation system, filtered speed is as final navigation calculation output information;
(1) construct Butterworth rejection filter, be specially:
In step 3, rotating mechanism adopts angular velocity of rotation to be ω, and this angular velocity causes the oscillation error frequency of navigation information to be about f=1/ ω; Set the passband lower-cut-off frequency of Butterworth rejection filter as (0.59 ~ 0.61) f, upper cut-off frequency cutoff frequency is (1.67 ~ 1.69) f, stopband lower-cut-off frequency is (0.83 ~ 0.85) f, stopband upper cut-off frequency is (1.13 ~ 1.15) f, then have:
Ω 1 = ( 0.59 ~ 0.61 ) f Ω 3 = ( 1.67 ~ 1.69 ) f Ω sl = ( 0.83 ~ 0.85 ) f Ω sh = ( 1.13 ~ 1.15 ) f - - - ( 14 )
Wherein, Ω 1represent passband lower-cut-off frequency; Ω 3represent upper cut-off frequency cutoff frequency; Ω slrepresent stopband lower-cut-off frequency; Ω shrepresent stopband upper cut-off frequency;
Ω BW = Ω 3 - Ω 1 Ω 2 2 = Ω 1 Ω 3 - - - ( 15 )
Wherein, Ω bWrepresent pass band width, Ω 2represent stopband center frequency;
Frequency normalization, for:
η 1 = Ω 1 / Ω BW η 3 = Ω 3 / Ω BW η sl = Ω sl / Ω BW η sh = Ω sh / Ω BW η 2 2 = η 1 η 3 = Ω 1 Ω 3 / Ω BW 2 - - - ( 16 )
Wherein, η irepresent corresponding frequencies Ω respectively icorresponding normalized frequency, i=1,2,3, sl, sh;
The frequency inverted of the corresponding low-pass filter of rejection filter, detailed process is:
λ p = η 3 - η 1 = 1 - λ s ′ = η sl / ( η sl 2 - η 2 2 ) λ s ′ = η sh / ( η sh 2 - η 2 2 ) - - - ( 17 )
Wherein, λ prepresent the cut-off frequecy of passband of low-pass filter; λ ' srepresent the stopband lower-cut-off frequency of low-pass filter;-λ ' srepresent the symmetrical frequency of the stopband lower-cut-off frequency of low-pass filter;
According to (17) formula, get λ s' and-λ s' absolute value is less one be final stopband lower-cut-off frequency, i.e. λ s=min{| λ ' s|, |-λ ' s|;
Design low-pass filter G (p); Asking for filter order is:
N = lg 10 α s / 10 - 1 10 α p / 10 - 1 / lg λ s - - - ( 18 )
Wherein, N represents the order of wave filter, α srepresent the minimal attenuation that stopband should reach, α prepresent that passband allows maximum attenuation usually to adopt α s=14dB, α p=3dB;
Therefore,
G ( p ) = 1 p N + 1 - - - ( 19 )
Wherein, p represents wave filter parameter;
Parameter fixed relationship between low-pass filter and rejection filter is,
p = s ( Ω 3 - Ω 1 ) s 2 + Ω 1 Ω 3 | s = z - 1 z + 1 = ( z 2 - 1 ) ( Ω 3 - Ω 1 ) ( z - 1 ) 2 + Ω 1 Ω 3 ( z + 1 ) 2 - - - ( 20 )
(20) formula substitutes into (19) formula, constructs the transfer function of discrete type second order band resistance Butterworth wave filter,
H ( z ) = G ( p ) | p = ( z 2 - 1 ) ( Ω 3 - Ω 1 ) ( z - 1 ) 2 + Ω 1 Ω 3 ( z + 1 ) 2 = 1 [ ( z 2 - 1 ) ( Ω 3 - Ω 1 ) ( z - 1 ) 2 + Ω 1 Ω 3 ( z + 1 ) 2 ] N + 1
(2) bearer rate under resolving the navigation system that obtains in step 4 is carried out rejection filter process, oscillation error item relevant with angular velocity of rotation in this velocity information of filtering, filtered speed, as final navigation calculation output information, is specially:
Single-shaft-rotation strapdown inertial navitation system (SINS) computing speed is modeled as:
V INS=V+δV INS(r)+δV INS(s)(22)
Wherein, V iNSrepresent the velocity information obtained through navigation calculation in step 4; V represents carrier movement speed; δ V iNS (r)represent the oscillation error caused by rotation modulation; δ V iNS (s)represent inertial reference calculation Schuler, earth oscillation error and constant error;
Due to place frequency range and other several error place frequency ranges differ greatly, right the wave filter using step (1) to build, obtains:
V′ INS=H(z)·V INS(23)
In formula, H (z) represents the rejection filter built according to angular velocity of rotation; V ' iNSfor filtered speed, wherein only comprise V, δ V iNS (s);
With filtered speed V ' iNSfinally resolve velocity information as system, and export the display device of navigational computer to, provide carrier underway motion velocity information.
CN201310005528.3A 2012-11-02 2013-01-08 Method for restraining speed errors of single-shaft rotation optical fiber gyro strapdown inertial navigation system CN103090866B (en)

Priority Applications (4)

Application Number Priority Date Filing Date Title
CN201210432160 2012-11-02
CN2012104321604 2012-11-02
CN201210432160.4 2012-11-02
CN201310005528.3A CN103090866B (en) 2012-11-02 2013-01-08 Method for restraining speed errors of single-shaft rotation optical fiber gyro strapdown inertial navigation system

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN201310005528.3A CN103090866B (en) 2012-11-02 2013-01-08 Method for restraining speed errors of single-shaft rotation optical fiber gyro strapdown inertial navigation system

Publications (2)

Publication Number Publication Date
CN103090866A CN103090866A (en) 2013-05-08
CN103090866B true CN103090866B (en) 2015-05-27

Family

ID=48203732

Family Applications (1)

Application Number Title Priority Date Filing Date
CN201310005528.3A CN103090866B (en) 2012-11-02 2013-01-08 Method for restraining speed errors of single-shaft rotation optical fiber gyro strapdown inertial navigation system

Country Status (1)

Country Link
CN (1) CN103090866B (en)

Families Citing this family (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN103743413B (en) * 2013-12-27 2016-05-04 哈尔滨工程大学 Heeling condition modulated is sought northern instrument alignment error On-line Estimation and is sought northern error compensating method
CN103759731B (en) * 2014-01-16 2016-08-17 电子科技大学 Angular speed initial conditions places an order increment Research on Rotation Vector Attitude method
CN103940445B (en) * 2014-04-10 2016-08-17 哈尔滨工程大学 A kind of single-shaft-rotation inertial navigation system inertial device error compensation method
CN104978476B (en) * 2014-04-14 2019-06-25 李巍岳 Indoor map scene, which is carried out, using smart phone mends the method surveyed
CN105606124A (en) * 2015-12-21 2016-05-25 河北汉光重工有限责任公司 Dual feedback loop gyro modulation method
CN106919192B (en) * 2015-12-24 2019-11-15 北京自动化控制设备研究所 A kind of control method of rotating device
CN105628025B (en) * 2015-12-31 2018-06-29 中国人民解放军国防科学技术大学 A kind of constant speed offset frequency/machine laser gyroscope shaking inertial navigation system air navigation aid
CN106017507B (en) * 2016-05-13 2019-01-08 北京航空航天大学 A kind of used group quick calibrating method of the optical fiber of precision low used in
CN107741240A (en) * 2017-10-11 2018-02-27 成都国卫通信技术有限公司 A kind of combined inertial nevigation system self-adaption Initial Alignment Method suitable for communication in moving
CN107830873A (en) * 2017-11-01 2018-03-23 北京航空航天大学 A kind of high-precision vehicle positioning and orienting method combined based on uniaxial horizontal rotation inertial navigation with odometer
CN109211279A (en) * 2018-11-07 2019-01-15 中国兵器工业集团第二四研究所苏州研发中心 A kind of System and method for for MIMU gyroscope nonlinearity automatic Calibration

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101514899A (en) * 2009-04-08 2009-08-26 哈尔滨工程大学 Optical fibre gyro strapdown inertial navigation system error inhibiting method based on single-shaft rotation
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
CN101963512A (en) * 2010-09-03 2011-02-02 哈尔滨工程大学 Initial alignment method for marine rotary fiber-optic gyroscope strapdown inertial navigation system

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101514899A (en) * 2009-04-08 2009-08-26 哈尔滨工程大学 Optical fibre gyro strapdown inertial navigation system error inhibiting method based on single-shaft rotation
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
CN101963512A (en) * 2010-09-03 2011-02-02 哈尔滨工程大学 Initial alignment method for marine rotary fiber-optic gyroscope strapdown inertial navigation system

Also Published As

Publication number Publication date
CN103090866A (en) 2013-05-08

Similar Documents

Publication Publication Date Title
Jekeli Inertial navigation systems with geodetic applications
CN103675861B (en) Satellite autonomous orbit determination method based on satellite-borne GNSS multiple antennas
Titterton et al. Strapdown inertial navigation technology
CN102519460B (en) Non-linear alignment method of strapdown inertial navigation system
US6459990B1 (en) Self-contained positioning method and system thereof for water and land vehicles
CN102192741B (en) Stabilised estimation of the pitch angles of an aircraft
CN101907714B (en) GPS aided positioning system and method based on multi-sensor data fusion
CN101788296B (en) SINS/CNS deep integrated navigation system and realization method thereof
CN102621565B (en) Transfer aligning method of airborne distributed POS (Position and Orientation System)
CN103245360B (en) Carrier-borne aircraft rotation type strapdown inertial navigation system Alignment Method under swaying base
CN103196448B (en) A kind of airborne distributed inertia surveys appearance system and Transfer Alignment thereof
Sun et al. MEMS-based rotary strapdown inertial navigation system
CN101413800B (en) Navigating and steady aiming method of navigation / steady aiming integrated system
Curey et al. Proposed IEEE inertial systems terminology standard and other inertial sensor standards
CN101344391B (en) Lunar vehicle posture self-confirming method based on full-function sun-compass
CN101893440B (en) Celestial autonomous navigation method based on star sensors
CN103389092B (en) A kind of kite balloon airship attitude measuring and measuring method
CN102252673B (en) Correction method for on-track aberration of star sensor
CN104567931A (en) Course-drifting-error elimination method of indoor inertial navigation positioning
CN102980577B (en) Micro-strapdown altitude heading reference system and working method thereof
CN102927994B (en) A kind of quick calibrating method of oblique redundant strapdown inertial navigation system
CN103323026B (en) The attitude reference estimation of deviation of star sensor and useful load and modification method
CN100585602C (en) Inertial measuring system error model demonstration test method
CN103076015B (en) A kind of SINS/CNS integrated navigation system based on optimum correction comprehensively and air navigation aid thereof
CN104655131B (en) Inertial navigation Initial Alignment Method based on ISTSSRCKF

Legal Events

Date Code Title Description
PB01 Publication
C06 Publication
SE01 Entry into force of request for substantive examination
C10 Entry into substantive examination
CB03 Change of inventor or designer information

Inventor after: Wang Qiuying

Inventor after: Qi Zhao

Inventor after: Gao Feng

Inventor after: Gao Wei

Inventor after: Sun Feng

Inventor before: Sun Feng

Inventor before: Wang Qiuying

Inventor before: Qi Zhao

Inventor before: Gao Wei

Inventor before: Gao Feng

Inventor after: Wang Qiuying

Inventor after: Qi Zhao

Inventor after: Gao Feng

Inventor after: Gao Wei

Inventor after: Sun Feng

Inventor before: Sun Feng

Inventor before: Wang Qiuying

Inventor before: Qi Zhao

Inventor before: Gao Wei

Inventor before: Gao Feng

COR Change of bibliographic data

Free format text: CORRECT: INVENTOR; FROM: SUN FENG WANG QIUYING QI ZHAO GAO WEI GAO FENG TO: WANG QIUYING QI ZHAO GAO FENG GAO WEI SUN FENG

GR01 Patent grant
C14 Grant of patent or utility model
CF01 Termination of patent right due to non-payment of annual fee

Granted publication date: 20150527

Termination date: 20190108

CF01 Termination of patent right due to non-payment of annual fee