CN103697911A - Initial attitude determination method for strapdown inertial navigation system under circumstance of unknown latitude - Google Patents

Initial attitude determination method for strapdown inertial navigation system under circumstance of unknown latitude Download PDF

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CN103697911A
CN103697911A CN201310694208.3A CN201310694208A CN103697911A CN 103697911 A CN103697911 A CN 103697911A CN 201310694208 A CN201310694208 A CN 201310694208A CN 103697911 A CN103697911 A CN 103697911A
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attitude
latitude
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孙枫
夏健钟
奔粤阳
杨晓龙
李敬春
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Harbin Engineering University
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    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C25/00—Manufacturing, calibrating, cleaning, or repairing instruments or devices referred to in the other groups of this subclass
    • G—PHYSICS
    • G01—MEASURING; TESTING
    • G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
    • G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
    • G01C21/10—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration
    • G01C21/12—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning
    • G01C21/16—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by using measurements of speed or acceleration executed aboard the object being navigated; Dead reckoning by integrating acceleration or speed, i.e. inertial navigation

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Abstract

The invention discloses an initial attitude determination method for a strapdown inertial navigation system under the circumstance of an unknown latitude. The initial attitude determination method is composed of the following steps: step 1, starting the inertial measurement assembly of the strapdown inertial navigation system with a fibre-optic gyroscope, adequately preheating, and then continuously acquiring the output data of the fibre-optic gyroscope and an accelerometer by virtue of an FPGA (field programmable gate array); step 2, running two compass alignment programs with different parameter settings simultaneously by virtue of the acquired output data of the fibre-optic gyroscope and the accelerometer, so as to acquire two groups of carrier attitude values; step 3, evaluating the steady-state error of an azimuth misalignment angle by virtue of the course angles output by the two alignment programs; step 4, evaluating a latitude value and the steady-state error of an east level misalignment angle by virtue of the course angles output by the two alignment programs; step 5, compensating the attitude by virtue of the evaluated steady-state error of the azimuth misalignment angle and the steady-state error of the east level misalignment angle, so as to finish initial alignment. The initial attitude determination method disclosed by the invention has the beneficial effect that accurate positioning can also be realized under the circumstance of the unknown latitude.

Description

Strapdown inertial navitation system (SINS) initial attitude under the unknown situation of a kind of latitude is determined method
Technical field
The invention belongs to navigational system technical field, the strapdown inertial navitation system (SINS) initial attitude relating under the unknown situation of a kind of latitude is determined method.
Background technology
Initial alignment is the prerequisite of strapdown inertial navigation system navigation work, and Initial Alignment Error is one of main error source of strapdown inertial navigation system, the accuracy affects navigation accuracy of aligning.Therefore, before navigation, obtain the accurate initial attitude of carrier particularly important.Initial alignment is divided and can be divided into static-base alignment and moving alignment by the motion state of pedestal, by the demand of punctual externally information is divided and can be divided into autonomous alignment and non-autonomous alignment.That which kind of alignment so is all to carry out under the prerequisite of position ten-four conventionally, yet in actual use because the restriction of some particular surroundings causes carrier may face the problem starting under positional information unknown situation.Patent publication No. is in the file of CN103217159A, " a kind of SINS/GPS/ polarized light integrated navigation system modeling and initial alignment on moving base method " disclosed, utilizing speed that GPS provides and positional information to set up error equation estimates attitude of carrier, because gps signal is easily disturbed, cannot meet the navigation task under some specific environments, so the method antijamming capability is weak and do not have a generality.Alignment algorithm under a kind of positional information unknown situation has been proposed in document " SINS Initial Alignment Analysis Under Geographic Latitude Uncertainty ", utilize the mode rough calculation of vector product to go out latitude, therefore under quiet pedestal condition, effect is fine, yet in the situation of swaying base, precision can be greatly affected.
Summary of the invention
Object of the present invention is providing the strapdown inertial navitation system (SINS) initial attitude under the unknown situation of a kind of latitude to determine method, has solved existing inertial navigation system not high problem of alignment precision under the unknown situation of latitude.
The technical solution adopted in the present invention is comprised of following steps:
Step 1: start the inertial measurement cluster of fiber optic gyro strapdown inertial navigation system, carry out after abundant preheating, utilize the output data of FPGA continuous acquisition fibre optic gyroscope and accelerometer;
Step 2: the fibre optic gyroscope that utilization collects and accelerometer output data are moved two cover parameters simultaneously different compass alignment procedures is set, and obtain two groups of attitude of carrier values;
Step 3: utilize the course angle of two cover alignment procedure outputs to ask for orientation misalignment steady-state error;
Step 4: utilize the course angle of two cover program outputs to calculate latitude value and the horizontal misalignment steady-state error of east orientation;
Step 5: utilize orientation misalignment steady-state error and the horizontal misalignment steady-state error of east orientation of trying to achieve to compensate attitude, complete initial alignment.
It is as follows that step 2 is obtained the algorithm of two groups of attitude of carrier values:
Strapdown attitude matrix
Figure BSA0000099096330000021
update algorithm as follows:
C · b n = C b n Ω nb b
ω nb b = ω ib b - C n b ( ω en n + ω ie n ) - C n b ω c n
The computation process of every angular velocity is as follows:
Figure BSA0000099096330000024
Wherein,
Figure BSA0000099096330000026
represent that carrier system (b system) is to the strapdown attitude matrix of navigation system (n system), represent
Figure BSA0000099096330000028
antisymmetric matrix,
Figure BSA0000099096330000029
the projection of the angular velocity that expression b system with respect to n is under b system,
Figure BSA0000099096330000031
represent fibre optic gyroscope output data,
Figure BSA0000099096330000032
represent the projection under n is with respect to the angular velocity of earth system (e system) of n system, ω ierepresent rotational-angular velocity of the earth,
Figure BSA0000099096330000033
represent the projection of rotational-angular velocity of the earth under n system, represent the projection of pilot angle speed under n system, R represents earth radius,
Figure BSA0000099096330000035
represent local latitude, V eand V nrepresent that respectively bearer rate is in the projection of east orientation and north orientation;
By above process, can obtain attitude of carrier matrix attitude matrix is expressed as to following form:
C b n = C 11 C 12 C 13 C 21 C 22 C 23 C 31 C 32 C 33
Wherein, C ij(i=1,2,3; J=1,2,3) expression attitude matrix in the value of the capable j of i row;
Basis
Figure BSA0000099096330000039
can obtain attitude of carrier, the main value of pitch angle θ, roll angle γ and course angle ψ is as follows:
θ main=sin -1(C 32)
Figure BSA00000990963300000310
Figure BSA00000990963300000311
According to the field of definition restriction of pitch angle (90 °, 90 °), roll angle (90 °, 90 °) and course angle (180 °, 180 °), by main value, determined that the formula of true value is as follows:
θ=θ main
γ=γ main
The attitude that compass alignment procedure 1 obtains is respectively: pitch angle θ 1, roll angle γ 1with course angle ψ 1;
The attitude that compass alignment procedure 2 obtains is respectively: pitch angle θ 2, roll angle γ 2with course angle ψ 2.
The algorithm of step 3 is as follows:
According to relational expression:
ψ 1=ψ 0+φ z1
ψ 2=ψ 0+φ z2
Figure BSA0000099096330000042
The orientation misalignment steady-state error that obtains two cover programs is respectively:
Figure BSA0000099096330000043
Figure BSA0000099096330000044
Figure BSA0000099096330000045
Wherein, ψ 1and ψ 2represent respectively the course angle output of synchronization two cover programs, ψ 0the true course angle that represents this moment, φ z1and φ z2the orientation misalignment that represents respectively two cover programs, ε eand ε urepresent that respectively equivalent east orientation and sky are to gyroscopic drift, φ ' z1and φ ' z2the orientation misalignment steady-state error that represents respectively two cover programs.
The algorithm of step 4 is as follows:
Figure BSA0000099096330000046
When latitude is non-vanishing, due to
Figure BSA0000099096330000047
?
Figure BSA0000099096330000048
therefore, can ignore part during practical application
Figure BSA0000099096330000049
calculate approx local latitude value;
Figure BSA00000990963300000410
When latitude value is determined, can calculate the horizontal misalignment offset of east orientation in two cover programs:
Figure BSA0000099096330000052
Wherein, φ ' x1and φ ' x2the horizontal misalignment steady-state error of east orientation that represents respectively two cover programs.
The invention has the beneficial effects as follows under the unknown situation of latitude and also can accurately aim at.
Accompanying drawing explanation
Fig. 1 is schematic diagram of the present invention;
Fig. 2 is inertial navigation compass method alignment principles figure;
Fig. 3 is for revising angular speed calculation block diagram;
Fig. 4 is the latitude calculated value that dual program compass method is aimed at;
Fig. 5 is the attitude error curve of compass program 1 and compass program 2;
Fig. 6 is the attitude error curve that dual program compass method is aimed at.
Embodiment
The present invention comprises the following steps as shown in Figure 1:
Step 1: start the inertial measurement cluster of fiber optic gyro strapdown inertial navigation system, carry out after abundant preheating, utilize FPGA continuous acquisition fibre optic gyroscope and accelerometer output data;
Step 2: the fibre optic gyroscope that utilization collects and accelerometer output data are moved two cover parameters simultaneously different compass alignment procedures is set, and obtain two groups of attitude of carrier values;
For strapdown inertial navitation system (SINS) compass method is aimed at, its process utilizes compass principle to ask for correction angle speed exactly, then strapdown attitude matrix is upgraded, thereby obtains attitude of carrier.
Strapdown attitude matrix
Figure BSA0000099096330000053
update algorithm as follows:
C · b n = C b n Ω nb b
ω nb b = ω ib b - C n b ( ω en n + ω ie n ) - C n b ω c n
The computation process of every angular velocity is as follows:
Figure BSA0000099096330000061
Figure BSA0000099096330000062
Wherein,
Figure BSA0000099096330000063
represent that carrier system (b system) is to the strapdown attitude matrix of navigation system (n system),
Figure BSA0000099096330000064
represent
Figure BSA0000099096330000065
antisymmetric matrix,
Figure BSA0000099096330000066
the projection of the angular velocity that expression b system with respect to n is under b system,
Figure BSA0000099096330000067
represent fibre optic gyroscope output data,
Figure BSA0000099096330000068
represent the projection under n is with respect to the angular velocity of earth system (e system) of n system, ω ierepresent rotational-angular velocity of the earth, represent the projection of rotational-angular velocity of the earth under n system, represent the projection of pilot angle speed under n system, R represents earth radius,
Figure BSA00000990963300000611
represent local latitude, V eand V nrepresent that respectively bearer rate is in the projection of east orientation and north orientation.
By above process, can obtain attitude of carrier matrix
Figure BSA00000990963300000612
attitude matrix is expressed as to following form:
C b n = C 11 C 12 C 13 C 21 C 22 C 23 C 31 C 32 C 33
Wherein, C ij(i=1,2,3; J=1,2,3) expression attitude matrix in the value of the capable j of i row.
Basis
Figure BSA00000990963300000615
can obtain attitude of carrier, the main value of pitch angle θ, roll angle γ and course angle ψ is as follows:
θ main=sin -1(C 32)
Figure BSA00000990963300000616
Figure BSA00000990963300000617
According to the field of definition restriction of pitch angle (90 °, 90 °), roll angle (90 °, 90 °) and course angle (180 °, 180 °), by main value, determined that the formula of true value is as follows:
θ=θ main
γ=γ main
Figure BSA0000099096330000071
The attitude that compass alignment procedure 1 obtains is respectively: pitch angle θ 1, roll angle γ 1with course angle ψ 1.
The attitude that compass alignment procedure 2 obtains is respectively: pitch angle θ 2, roll angle γ 2with course angle ψ 2.
The parameter designing of two cover compass alignment procedures is as follows:
The parameter of compass program 1:
Figure BSA0000099096330000072
The parameter of compass program 2:
k 1 ′ = k 1 k 2 ′ = k 2 k E ′ = 1.5 k E K N ′ = 1.5 k N k U ′ = k U
Wherein, the damping ratio that ξ is system, ω nfor the undamped oscillation frequency of system, g is local gravitational acceleration, k 1, k 2, k e, k n, k ufor the gain term in compass alignment procedure 1, k ' 1, k ' 2, k ' e, k ' n, k ' ufor the gain term in compass alignment procedure 2.Fig. 2 is inertial navigation compass method alignment principles figure; By accelerometer output, ask for pilot angle speed
Figure BSA0000099096330000074
process can be obtained by Figure of description 3.
Step 3: utilize the course angle of two cover alignment procedure outputs to ask for orientation misalignment steady-state error;
According to relational expression:
ψ 1=ψ 0+φ z1
ψ 2=ψ 0+φ z2
Figure BSA0000099096330000081
Figure BSA0000099096330000082
The orientation misalignment steady-state error that obtains two cover programs is respectively:
Figure BSA0000099096330000083
Figure BSA0000099096330000084
Wherein, ψ 1and ψ 2represent respectively the course angle output of synchronization two cover programs, ψ 0the true course angle that represents this moment, φ z1and φ z2the orientation misalignment that represents respectively two cover programs, ε eand ε urepresent that respectively equivalent east orientation and sky are to gyroscopic drift, φ ' z1and φ ' z2the orientation misalignment steady-state error that represents respectively two cover programs.
Step 4: utilize the course angle of two cover program outputs to calculate latitude value and the horizontal misalignment steady-state error of east orientation;
According to the latitude value of gained in step 3 and course angle relation:
Figure BSA0000099096330000086
When latitude is non-vanishing, due to
Figure BSA0000099096330000087
? therefore, can ignore part during practical application
Figure BSA0000099096330000089
calculate approx local latitude value.
Figure BSA00000990963300000810
When latitude value is determined, can calculate the horizontal misalignment offset of east orientation in two cover programs:
Figure BSA0000099096330000091
Figure BSA0000099096330000092
Wherein, φ ' x1and φ ' x2the horizontal misalignment steady-state error of east orientation that represents respectively two cover programs.
Step 5: utilize orientation misalignment steady-state error and the horizontal misalignment steady-state error of east orientation of trying to achieve to compensate attitude, complete initial alignment.
Due to the horizontal misalignment of north orientation, to be that roll angle γ steady-state error is affected by latitude less, do not need compensation, again because the compensation principle of two systems is identical, in order to save calculated amount, only pitch angle and the course angle of 1 output of compass alignment procedure compensated:
θ′ 1=θ 1-φ′ x1
ψ′ 1=ψ 1-φ′ z1
Wherein, θ ' 1represent the pitch angle output after 1 compensation of compass program, ψ ' 1represent the course angle output after 1 compensation of compass program.
Method tool of the present invention has the following advantages: (1), under Geographic Latitude Uncertainty, determines strapdown inertial navitation system (SINS) initial attitude value rapidly and accurately.(2) utilize a set of inertial measurement cluster to support two cover compass program parallelizations to calculate simultaneously, under the prerequisite that does not increase cost, improved the precision of initial alignment.
With specific embodiment, the present invention will be described below.
Embodiment 1:
In order to verify the strapdown inertial navitation system (SINS) compass method alignment methods under the unknown situation of this latitude, utilize one group of actual measurement experimental data (latitude
Figure BSA0000099096330000094
longitude λ=126.599 °) carry out data simulation experiment, the time is 1 hour.
Setting latitude is unknown, sets up two compass program alignment algorithms, and the damping ratio of system 1 and undamped oscillation frequency are got ξ=0.707 and ω n=0.015, the parameter of system 1 is:
k 1 = k 2 = 0.0283 k E = k N = 6.12 × 10 - 5 k U = 1.12 × 10 - 4
The parameter of system 2 is:
k 1 ′ = k 2 ′ = 0.0283 k E ′ = k N ′ = 9.18 × 10 - 5 k U ′ = 1.12 × 10 - 4
Be illustrated in figure 4 the latitude calculated value that dual program compass method is aimed at; Be illustrated in figure 5 the attitude error curve of compass program 1 and compass program 2; The attitude error curve that the method for dual program compass shown in comparison diagram 6 is aimed at, obtains the latitude calculated value of dual program compass method aligning and the attitude error curve that dual program compass method is aimed at by emulation experiment.Result shows after adjustment after a while, latitude value can be estimated in real time, and latitude value is 45.7119 degree in stable state part average, and mean square deviation is 0.2414 degree, compare with true latitude 45.776 degree, the latitude value that reaches this precision can be thought on misalignment compensation value calculation without impact.The attitude error that dual program compass method is aimed at and the contrast of separate payment attitude error, can adopt dual program compass method can effectively compensate the unknown caused constant error of latitude.The azimuthal error average of system after compensation in steady-state process is 0.0586 degree, and mean square deviation is 0.0263 degree, adopts as seen the method can effectively realize the accurate aligning under latitude unknown situation.

Claims (4)

1. the strapdown inertial navitation system (SINS) initial attitude under the unknown situation of latitude is determined a method, it is characterized in that being comprised of following steps:
Step 1: start the inertial measurement cluster of fiber optic gyro strapdown inertial navigation system, carry out after abundant preheating, utilize the output data of FPGA continuous acquisition fibre optic gyroscope and accelerometer;
Step 2: the fibre optic gyroscope that utilization collects and accelerometer output data are moved two cover parameters simultaneously different compass alignment procedures is set, and obtain two groups of attitude of carrier values;
Step 3: utilize the course angle of two cover alignment procedure outputs to ask for orientation misalignment steady-state error;
Step 4: utilize the course angle of two cover program outputs to calculate latitude value and the horizontal misalignment steady-state error of east orientation;
Step 5: utilize orientation misalignment steady-state error and the horizontal misalignment steady-state error of east orientation of trying to achieve to compensate attitude, complete initial alignment.
2. according to the strapdown inertial navitation system (SINS) initial attitude under the unknown situation of a kind of latitude shown in claim 1, determine method, it is characterized in that: it is as follows that described step 2 is obtained the algorithm of two groups of attitude of carrier values:
Strapdown attitude matrix
Figure FSA0000099096320000011
update algorithm as follows:
C · b n = C b n Ω nb b
ω nb b = ω ib b - C n b ( ω en n + ω ie n ) - C n b ω c n
The computation process of every angular velocity is as follows:
Figure FSA0000099096320000016
Figure FSA0000099096320000015
Wherein,
Figure FSA0000099096320000021
represent that carrier system (b system) is to the strapdown attitude matrix of navigation system (n system),
Figure FSA0000099096320000022
represent
Figure FSA0000099096320000023
antisymmetric matrix,
Figure FSA0000099096320000024
the projection of the angular velocity that expression b system with respect to n is under b system,
Figure FSA0000099096320000025
represent fibre optic gyroscope output data,
Figure FSA0000099096320000026
represent the projection under n is with respect to the angular velocity of earth system (e system) of n system, ω ierepresent rotational-angular velocity of the earth,
Figure FSA0000099096320000027
represent the projection of rotational-angular velocity of the earth under n system,
Figure FSA0000099096320000028
represent the projection of pilot angle speed under n system, R represents earth radius,
Figure FSA0000099096320000029
represent local latitude, V eand V nrepresent that respectively bearer rate is in the projection of east orientation and north orientation;
By above process, can obtain attitude of carrier matrix
Figure FSA00000990963200000210
attitude matrix is expressed as to following form:
C b n = C 11 C 12 C 13 C 21 C 22 C 23 C 31 C 32 C 33
Wherein, C ij(i=1,2,3; J=1,2,3) expression attitude matrix
Figure FSA00000990963200000212
in the value of the capable j of i row;
Basis
Figure FSA00000990963200000213
can obtain attitude of carrier, the main value of pitch angle θ, roll angle γ and course angle ψ is as follows:
θ main=sin -1(C 32)
Figure FSA00000990963200000214
Figure FSA00000990963200000215
According to the field of definition restriction of pitch angle (90 °, 90 °), roll angle (90 °, 90 °) and course angle (180 °, 180 °), by main value, determined that the formula of true value is as follows:
θ=θ main
γ=γ main
Figure FSA00000990963200000216
The attitude that compass alignment procedure 1 obtains is respectively: pitch angle θ 1, roll angle γ 1with course angle ψ 1;
The attitude that compass alignment procedure 2 obtains is respectively: pitch angle θ 2, roll angle γ 2with course angle ψ 2.
3. according to the strapdown inertial navitation system (SINS) initial attitude under the unknown situation of a kind of latitude shown in claim 1, determine method, it is characterized in that: the algorithm of described step 3 is as follows:
According to relational expression:
ψ 1=ψ 0+φ z1
ψ 2=ψ 0+φ z2
Figure FSA0000099096320000031
Figure FSA0000099096320000032
The orientation misalignment steady-state error that obtains two cover programs is respectively:
Figure FSA0000099096320000033
Figure FSA0000099096320000034
Figure FSA0000099096320000035
Wherein, ψ 1and ψ 2represent respectively the course angle output of synchronization two cover programs, ψ 0the true course angle that represents this moment, φ z1and φ z2the orientation misalignment that represents respectively two cover programs, ε eand ε urepresent that respectively equivalent east orientation and sky are to gyroscopic drift, φ ' z1and φ ' z2the orientation misalignment steady-state error that represents respectively two cover programs.
4. according to the strapdown inertial navitation system (SINS) initial attitude under the unknown situation of a kind of latitude shown in claim 1, determine method, it is characterized in that: the algorithm of described step 4 is as follows:
Figure FSA0000099096320000036
When latitude is non-vanishing, due to
Figure FSA0000099096320000037
?
Figure FSA0000099096320000038
therefore, can ignore part during practical application calculate approx local latitude value;
Figure FSA00000990963200000310
When latitude value is determined, can calculate the horizontal misalignment offset of east orientation in two cover programs:
Figure FSA0000099096320000041
Figure FSA0000099096320000042
Wherein, φ ' x1and φ ' x2the horizontal misalignment steady-state error of east orientation that represents respectively two cover programs.
CN201310694208.3A 2013-12-18 2013-12-18 Initial attitude determination method for strapdown inertial navigation system under circumstance of unknown latitude Pending CN103697911A (en)

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CN106840208A (en) * 2017-02-13 2017-06-13 哈尔滨工业大学 A kind of dual system compass method static-base alignment method under latitude unknown situation
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Application publication date: 20140402