CN102162731A - High-precision satellite independent navigation method based on pulse data of sun, earth and moon integrated sensor - Google Patents
High-precision satellite independent navigation method based on pulse data of sun, earth and moon integrated sensor Download PDFInfo
- Publication number
- CN102162731A CN102162731A CN2011100052394A CN201110005239A CN102162731A CN 102162731 A CN102162731 A CN 102162731A CN 2011100052394 A CN2011100052394 A CN 2011100052394A CN 201110005239 A CN201110005239 A CN 201110005239A CN 102162731 A CN102162731 A CN 102162731A
- Authority
- CN
- China
- Prior art keywords
- earth
- vector
- moon
- cos
- partiald
- 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.)
- Granted
Links
Images
Landscapes
- Navigation (AREA)
Abstract
The invention discloses to a high-precision satellite independent navigation method based on pulse data of a sun, earth and moon integrated sensor, and relates to the field of satellite navigation. The high-precision satellite independent navigation method is provided in order to solve the problem that the photocenter and the centroid of the moon cannot be overlapped because the pulse data comprising noises, the oblateness of the earth, the first quarter and the last quarter of the moon is used as the original data of a navigation system. The method comprises the following steps of: 1, providing the pulse data for a navigation computer; 2, determining directions of the three celestial bodies according to the pulse data; 3, calculating an inner product of a vector of the heliocentric direction and a vector of the geocentric direction, and an inner product of a vector of the selenocenter direction and a vector of the geocentric direction; 4, performing double-vector rough posture fixing; 5, initializing navigation; 6, performing rough navigation calculation so as to obtain a navigation result; 7, finishing refining of the direction of the moon; 8, performing double-vector fine posture fixing; 9, correcting the vectors of a geocentric distance and the geocentric direction; and 10, performing fine navigation calculation so as to obtain a final navigation result. By the high-precision satellite independent navigation method, the influence because the photocenter and the centroid of the moon cannot be overlapped according to the noises, the oblateness of the earth, the first quarter and the last quarter of the moon which are comprised in the pulse data is eliminated.
Description
Technical field
The present invention relates to the satellite navigation field, be specifically related to utilize day that day ground month integrated sensor measures,, month optics pulse data, position, speed and the attitude of satellite are carried out independently definite method.
Background technology
A day ground month autonomous navigation system is made up of navigation sensor and navigational computer.Navigational computer is handled the measured value of navigation sensor, through certain navigation algorithm, makes the position and the speed of satellite in real time, thereby realizes the independent navigation of satellite.Day ground month integrated sensor as depicted in figs. 1 and 2, scanning life sensor by the infrared double cone scanning type sensor of the earth and two fan-shaped slit visual fields is formed, the infrared double cone scanning type sensor of the earth has single optical scanning head, utilize mirror structure to obtain two infrared visual fields, the track of infrared visual field, scanning back is two coaxial circular cones, optical head scanning one circle, pyroelectric detector can detect four Horizons at most and pass through signal, the moment by the signal appearance can be determined the orientation of the earth's core direction vector with respect to satellite, and can try to achieve the distance of satellite to the earth's core.Two visible light sensors on the basis of the infrared double cone scanning type sensor of the earth, have been increased, in the scanning process of optical head, fan-shaped slit visual field is inswept spherical zone zone, utilize the si-photodiode detector can be responsive to the sun and the moon, according to the moment that the sun, the moon occur in fan-shaped slit visual field can be in the hope of the orientation of its direction vector with respect to satellite.Detecting device has a plurality of light intensity threshold values, can distinguish the sun and moon signal, and can reject earth signal.Day ground month navigational system is exactly to utilize such sensor to determine day orientation of the ground moon, thereby carries out the system of autonomous navigation of satellite.
At present to day ground month navigational system the research approach mainly be to be embodied in the angle information of utilizing, study navigation algorithm, do not consider measuring principle, these key factors that therefore yet can't embody compression of the earth, moon photocentre and barycenter do not overlap when going up lower edge and engineering is actual differs far away.
Pulse data is as the raw data of navigational system, the inside has comprised all information, and (noise, compression of the earth, last moon at the last quarter ball photocentre and barycenter do not overlap, or the like), the objective of the invention is to study day measuring principle of ground month integrated sensor, and research is based on the high-precision ground month navigation algorithm of subsisting of pulse data, this algorithm can compensate compression of the earth, also can the moon pulse during the last lower edge be compensated.
Summary of the invention
The present invention has wherein comprised noise, compression of the earth, has gone up the problem that moon at the last quarter ball photocentre and barycenter do not overlap in the inside as the raw data of navigational system in order to solve pulse data in the existing method, and has proposed the satellite high-precision autonomous navigation method based on day ground month integrated sensor pulse data.A measuring principle that the objective of the invention is to day ground month integrated sensor, and based on the high-precision ground month navigation algorithm of subsisting of pulse data, this algorithm can compensate compression of the earth, also can the moon pulse during the last lower edge be compensated.
The step of the satellite high-precision autonomous navigation method of month integrated sensor pulse data is as follows with the present invention is based on day:
Step 1: the navigation sensor navigation computing machine of being made up of the scanning life sensor of the infrared double cone scanning type sensor of the earth and two fan-shaped slit visual fields provides pulse data;
Step 2: the orientation of carrying out three celestial bodies according to pulse data is determined: determine the coordinate E of the earth's core direction vector under measurement coordinate system
SE, day heart direction vector coordinate S under measurement coordinate system
S|SEWith the coordinate M of moon heart direction vector under measurement coordinate system
S|SE, and the earth's core is apart from r;
Step 3: by the coordinate E of the earth's core direction vector under measurement coordinate system
SE, day heart direction vector coordinate S under measurement coordinate system
S|SEWith the coordinate M of moon heart direction vector under measurement coordinate system
S|SECalculate the inner product of day heart direction vector and the earth's core direction vector, and the inner product of month heart direction vector and the earth's core direction vector:
Step 4: carry out two vectors and slightly decide appearance, the vector of three pairwise orthogonals of definition is as follows:
v
1:r
m×r
s/|r
m×r
s|
v
2:r
m/|r
m| (1)
v
3:v
1×v
2
Wherein, r
mBe star moon vector, r
sBe star day vector; Described star moon vector r
mWith star day vector r
sExpression under inertial coordinates system by the ground moon vector, day vector is approximate obtains, and a ground month vector and a ground day vector can be obtained by ephemeris; If v
1, v
2, v
3Under inertial coordinates system, be expressed as E
1, E
2, E
3, under measurement coordinate system, be expressed as e
1, e
2, e
3, then inertial coordinate is tied to the rotation matrix of measurement coordinate system:
R
SEI=[e
1,e
2,e
3][E
1,E
2,E
3]
-1 (2)
Step 5: the navigation initialization, the initial position guestimate is calculated with following formula:
r
0=-rR
ISEE
SE (3)
Wherein,
By the formula mistake! Do not find Reference source.Obtain; The initial velocity guestimate then has the position of two adjacent moment to do approximate difference to obtain;
Step 6: adopt the least square navigation algorithm to carry out thick navigation operations and obtain navigation results;
Step 7: utilize the navigation results of the thick navigation in the step 6, revise star moon vector r
mValue, thereby obtain revised star moon vector r
M repaiiesFinished becoming more meticulous of moon orientation;
Step 8: carry out two vectors essences and decide appearance, star moon vector r
mThe revised star moon vector r that expression under inertial coordinates system is produced by previous step
M repaiiesSubstitute star day vector r
sConstant, the vector of three pairwise orthogonals of definition is as follows:
v
1': r
M repaiies* r
s/ | r
M repaiies* r
s|
v
2': r
M repaiies/ | r
M repaiies|
v
3′:v
1×v
2
If v
1', v
2', v
3' under inertial system, be expressed as E
1', E
2', E
3', under body coordinate system, be expressed as e
1', e
2', e
3' also can ask; Then two vectors of Xiu Zhenging inertial coordinate of deciding appearance is tied to the rotation matrix of measurement coordinate system:
R
SEI′=[e
1′,e
2′,e
3′][E
1′,E
2′,E
3′]
-1 (4)
Step 9: consider that inertial coordinate that compression of the earth is decided appearance based on two vectors of revising in the step 8 is tied to the rotation matrix R of measurement coordinate system
SEI' revise the earth's core apart from r and the earth's core orientation vector E
SE, obtain revised the earth's core apart from r
RepairWith revised the earth's core orientation vector E
SE repaiies
Step 10: adopt the least square navigation algorithm to carry out smart navigation operations and obtain final navigation results.
The present invention be according to day,, the optical pulse information of month integrated sensor determines the method for position, speed and the attitude of aircraft.At first be day,, month the orientation determine, two kinds of methods are arranged for orientation, the earth's core: determine (being used for thick navigation) and determine (smart navigation, ellipticity correction) based on the orientation, the earth's core of the spherical hypothesis of the earth based on the orientation, the earth's core of compression of the earth; Determine for the moon heart orientation during the last lower edge, considered that photocentre and moon barycenter do not overlap the influence that brings, and have provided compensation scheme.Provide the least square navigation implementation algorithm of a cover then, carry out the simulation accuracy assessment at last based on compensation of pulse data ellipticity and last lower edge correction.
Sensor scanning rotating speed is ω
Rot=240r/min.Two semi-cone angle of circular cone Horizon sensor infrared horizon visual field circular cone are respectively γ
1=38 ° and γ
2=73 °.Scanning head is with respect to slit sensor plane of symmetry M
1M
2The hysteresis angle be B
R1=B
R2=0 °, the sensor measurement coordinate system is γ with respect to the setting angle of system
I=π/6, β
I=π/6, α
I=π/6.Fan-shaped life sensor is β with respect to the pitch angle with the scanning rotating shaft
s=16 ° of slit visual fields 1 are ahead of angular distance υ=4 ° of slit visual field 2 in the sensor equatorial plane.
The initial time of track is taken as 0: 0: 0 on the 6th May in 2012.Initial six key elements of true track are: major semi-axis is 6878km, and orbital eccentricity is 0, and orbit inclination is 92 °, and right ascension of ascending node is π/6, and argument of perigee is 0 °, and very near angle is 0 °.When simulated data produced, the attitude of supposing satellite was strict absolute orientation.The time of using the orbit measurement data is in since 0: 0: 0 on the 12nd May in 2012 5000 seconds.
Come analog pulse with the orbital data that STK produces, consider compression of the earth in the pulse generation mechanism, 3 σ of earth impulsive noise=6.5e-5s, 3 σ=0.1 of corresponding angle noise °, the earth sensor systematic error is that (the respective pulses noise is 2.3e-5s to 0.03 degree, 3 σ=the 3.5e-5s of the scanning sun (moon) impulsive noise, 3 σ=0.05 of corresponding angle noise °.Navigation model is still used the dynamics of orbits model among the MATLAB, carries out Monte Carlo simulation, and simulation times is 120 times.The Filtering Estimation value of track initial value with the position average of the error of true track initial value is
The position mean square deviation
The speed average is
The speed mean square deviation
3 σ ellipsoids such as Fig. 6 and Fig. 7 of site error and velocity error.
Description of drawings
Fig. 1 is day mounting structure of ground month integrated sensor; A is a spacecraft, and B is the rotating shaft of the infrared double cone scanning type sensor of first earth, and C is the rotating shaft of the infrared double cone scanning type sensor of second earth, and D is the infrared visual fields of 38 degree, and E is the infrared visual fields of 73 degree, and F is the fan-shaped visual field of visible light; Fig. 2 is the visual field of the scanning life sensor of two fan-shaped slit visual fields; L is fan-shaped visual field, and M is the scanning rotating shaft, and N is an infrared probe; Fig. 3 is that earth sensor is measured geometric graph; Fig. 4 is a day ground month orientation synoptic diagram; Fig. 5 is a day ground month orientation synoptic diagram; Fig. 6 is 3 σ ellipsoid synoptic diagram of site error; Fig. 7 is 3 σ ellipsoid synoptic diagram of velocity error.
Embodiment
Embodiment one: in conjunction with Fig. 1 to Fig. 5 present embodiment is described, the present embodiment concrete steps are as follows:
Step 1: the navigation sensor navigation computing machine of being made up of the scanning life sensor of the infrared double cone scanning type sensor of the earth and two fan-shaped slit visual fields provides pulse data;
Step 2: the orientation of carrying out three celestial bodies according to pulse data is determined: determine the coordinate E of the earth's core direction vector under measurement coordinate system
SE, day heart direction vector coordinate S under measurement coordinate system
S|SEWith the coordinate M of moon heart direction vector under measurement coordinate system
S|SE, and the earth's core is apart from r;
Step 3: by the coordinate E of the earth's core direction vector under measurement coordinate system
SE, day heart direction vector coordinate S under measurement coordinate system
S|SEWith the coordinate M of moon heart direction vector under measurement coordinate system
S|SECalculate the inner product of day heart direction vector and the earth's core direction vector, and the inner product of month heart direction vector and the earth's core direction vector:
Step 4: carry out two vectors and slightly decide appearance, the vector of three pairwise orthogonals of definition is as follows:
v
1:r
m×r
s/|r
m×r
s|
v
2:r
m/|r
m| (5)
v
3:v
1×v
2
Wherein, r
mBe star moon vector, r
sBe star day vector; Described star moon vector r
mWith star day vector r
sExpression under inertial coordinates system by the ground moon vector, day vector is approximate obtains, and a ground month vector and a ground day vector can be obtained by ephemeris; If v
1, v
2, v
3Under inertial coordinates system, be expressed as E
1, E
2, E
3, under measurement coordinate system, be expressed as e
1, e
2, e
3, then inertial coordinate is tied to the rotation matrix of measurement coordinate system:
R
SEI=[e
1,e
2,e
3][E
1,E
2,E
3]
-1 (6)
Step 5: the navigation initialization, the initial position guestimate is calculated with following formula:
r
0=-rR
ISEE
SE (7)
Wherein,
By the formula mistake! Do not find Reference source.Obtain; The initial velocity guestimate then has the position of two adjacent moment to do approximate difference to obtain;
Step 6: adopt the least square navigation algorithm to carry out thick navigation operations and obtain navigation results;
Step 7: utilize the navigation results of the thick navigation in the step 6, revise star moon vector r
mValue, thereby obtain revised star moon vector r
M repaiiesFinished becoming more meticulous of moon orientation;
Step 8: carry out two vectors essences and decide appearance, star moon vector r
mThe revised star moon vector r that expression under inertial coordinates system is produced by previous step
M repaiiesSubstitute star day vector r
sConstant, the vector of three pairwise orthogonals of definition is as follows:
v
1': r
M repaiies* r
s/ | r
M repaiies* r
s|
v
2': r
M repaiies/ | r
M repaiies|
v
3′:v
1×v
2
If v
1', v
2', v
3' under inertial system, be expressed as E
1', E
2', E
3', under body coordinate system, be expressed as e
1', e
2', e
3' also can ask; Then two vectors of Xiu Zhenging inertial coordinate of deciding appearance is tied to the rotation matrix of measurement coordinate system:
R
SEI′=[e
1′,e
2′,e
3′][E
1′,E
2′,E
3′]
-1 (8)
Step 9: consider that inertial coordinate that compression of the earth is decided appearance based on two vectors of revising in the step 8 is tied to the rotation matrix R of measurement coordinate system
SEI' revise the earth's core apart from r and the earth's core orientation vector E
SE, obtain revised the earth's core apart from r
RepairWith revised the earth's core orientation vector E
SE repaiies
Step 10: adopt the least square navigation algorithm to carry out smart navigation operations and obtain final navigation results.
Embodiment two: present embodiment and embodiment one difference are the coordinate E of the earth's core direction vector under measurement coordinate system in the step 2
SEDefinite process, and the earth's core is as follows apart from definite process of r:
Two infrared probes on the infrared double cone scanning type sensor of the earth sweep terrestrial time:
α
1-in=-((ω
rott
ref+B
R1)-ω
rott
1-in)
(9)
α
2-in=-((ω
rott
ref+B
R2)-ω
rott
2-in)
ω wherein
RotBe the rotating speed of scanning sensor, B
R1, B
R2Be the drag angle of two infrared probes on the initial time double cone scanning type sensor with respect to benchmark x axle, t
RefFor the slit sensor plane of symmetry passes benchmark x axle to deserved pulse constantly, t
1-in, t
2-inThe pulse that forms when sweeping Horizon for two infrared probes on the double cone scanning type sensor;
Two infrared probes on the infrared double cone scanning type sensor of the earth scan out terrestrial time:
α
1-out=ω
rott
1-out-(ω
rott
ref+B
R1)
(10)
α
2-out=ω
rott
2-out-(ω
rott
ref+B
R2)
T wherein
1-out, t
2-outThe pulse that forms when scanning out Horizon for two infrared probes on the double cone scanning type sensor;
So the earth's core direction vector with respect to the position angle of measurement coordinate system is:
(11)
If do not consider compression of the earth, φ arranged
E1=φ
E2, and use φ
eExpression;
Thereby the string of the earth that the infrared double cone scanning type sensor of the earth obtains is wide be:
Ω
E1=-α
1-in+α
1-out
(12)
Ω
E2=-α
2-in+α
2-out
By the spherical trigonometry formula, have
Wherein ρ is the half angle with respect to the infrared radiation disk of the earth of satellite, γ
1, γ
2Be the semi-cone angle of earth sensor double cone, η is the angle between the earth's core direction and the sensor turning axle;
So
If do not consider compression of the earth, the coordinate of the earth's core direction vector under measurement coordinate system try to achieve into:
The earth's core is apart from being calculated by following formula:
r=R
E/sinρ (15)
Wherein, R
EBe the infrared radiation radius of a ball of the earth, ρ can be tried to achieve by the cosine law of spherical triangle:
Other step is identical with embodiment one.
Embodiment three: present embodiment and embodiment one difference are the coordinate S of day heart direction vector under measurement coordinate system in the step 2
S|SEDefinite process as follows:
Position angle and the elevation angle of day heart orientation under the measurement body coordinate system is:
δ
s=cot
-1(sinσ
scosβ
s)
Wherein: t
1sun, t
2sunBe the responsive pulse that forms to sunshine time the in the first fan-shaped slit visual field, the second fan-shaped slit visual field, σ
s=(ω
Rott
2sun-ω
Rott
1sun-υ)/2, υ is the scanning life sensor of the first fan-shaped slit visual field is ahead of the scanning life sensor of the second fan-shaped slit visual field on the sensor equatorial plane a focal length, β
sBe the angle of slit visual field with respect to the scanning axes of rotation skew;
The coordinate of day heart direction vector under measurement coordinate system is:
Other step is identical with embodiment one.
Embodiment four: present embodiment and embodiment one difference are the coordinate M of moon heart direction vector under measurement coordinate system in the step 2
S|SEDefinite process as follows:
Similar in the time of formula during the full moon the responsive moon time and responsive sunshine, position angle and the elevation angle of moon heart orientation under measurement coordinate system during the full moon is:
δ
m=tan
-1(sinσ
Mcosβ
s)
Wherein: t
1moon, t
2moonBe the responsive pulse that forms to the moon time in the first fan-shaped slit visual field, the second fan-shaped slit visual field, σ
M=(ω
Rott
2moon-ω
Rott
1moon-υ)/2;
Thereby the coordinate of moon heart direction vector under measurement coordinate system during the full moon is:
Moon heart orientation during the last lower edge determine and full moon during a month heart orientation determine that formula is the same, but individual impulse compensation module is arranged; Obtain the attitude information of satellite by thick navigation, can obtain sensor scan axis and x
mThe angle theta of (on the normal society face vertical axle) with ground moon direction
x, add a Δ t on the basis of the pulse that the pulse after the compensation should be original
Wherein, Δ t>0 in the time of the moon at the first quarter, Δ t<0 in the time of the moon at the last quarter; Again by the formula mistake! Do not find Reference source.Calculate φ
m, δ
m, by the formula mistake! Do not find Reference source.Moon heart orientation during the last lower edge after determining to compensate.Other step is identical with embodiment one.
Embodiment five: present embodiment and embodiment one difference are step 7 correction star moon vector r
m, because the distance of the sun and the earth is far, so star day vector r
sWith the direction vector R that points to the sun from the earth's core
sCan regard as parallel.And the moon is not far with respect to the distance of the earth, star moon vector r
mWith ground day vector R
mCan not regard as parallel.So star moon vector r
mUnder inertial coordinates system expression month vector R practicably not
mReplace, need carry out the limited distance correction.
The preliminary definite position r of satellite under inertial coordinates system from thick navigation
b, and moon inertial position is tried to achieve by ephemeris, thus determine star moon vector r
mExpression under inertial coordinates system, correction formula is as follows:
r
M repaiies=R
m-r
b
Embodiment six: in conjunction with Fig. 3 present embodiment is described, present embodiment and embodiment one difference are in the step 9 to determine to utilize based on the orientation, the earth's core of ellipticity correction the attitude information of step 8, to the earth's core apart from r and the earth's core orientation vector E
SERecomputate specific as follows:
By measuring geometric analysis, the earth's core is to the vector R of earth surface as can be known
EFor:
Wherein:
Be direction vector (expression under measurement coordinate system) from the earth's core sensing satellite,
With the unit vector (being projected under the measurement coordinate system) that scans sight line, make pulse t constantly respectively
1in, t
1out, t
2in, t
2outThe unit vector of corresponding scanning sight line is
If do not consider to measure noise, this tittle is accurately known, is respectively:
(22)
The earth's core is to the vector on earth infrared radiation ball surface
Be projected under the inertial coordinates system.
The ellipsoid equation is:
P=-1+1/ (1-ee) wherein
2.
By getting in the formula (24):
(25)
Wherein: L
3Be R
ISEThe third line.
With mistake! Do not find Reference source.A substitution mistake! Do not find Reference source.To obtaining quadratic equation:
(a+Δa)l
2+(b+Δb)l+(c+Δc)=0 (26)
Wherein:
(b+Δb)
2-4(a+Δa)(c+Δc)=0 (27)
Also promptly:
(28)
With in the navigation slightly each constantly
Can from last equation, solve revised as initial value
It should be noted that
Also should satisfy constraint condition
4 unknown numbers, 5 equations approach with least-squares estimation
Order
x
0Be the iterative initial value of x, with F at x
0Point single order Taylor expansion has:
Wherein:
And:
The form of being write as iteration has:
x
kBeing exactly the optimal estimation of x, also is revised the earth's core distance and the earth's core direction vector
Embodiment seven: present embodiment and embodiment one difference are that the least square navigation algorithm in step 6 and the step 10 is: suppose N observation t constantly
1, t
2..., t
NObserved quantity be
Described observed quantity is angle between month ground and the angle between day ground, is obtained by the arc cosine of step 3 inner product, and its true value is
Make residual error be:
Set up the least square index:
Wherein:
To go up in the equation substitution functional index and can get:
Following formula can be remembered and does:
Wherein:
Thereby can in the hope of:
Corresponding iterative formula can be write:
X
k+1=X
k-A
-1B
Content of the present invention is not limited only to the content of the respective embodiments described above, and the combination of one of them or several embodiments equally also can realize the purpose of inventing.
Claims (7)
1. based on the satellite high-precision autonomous navigation method of day ground month integrated sensor pulse data, it is characterized in that its step is as follows:
Step 1: the navigation sensor navigation computing machine of being made up of the scanning life sensor of the infrared double cone scanning type sensor of the earth and two fan-shaped slit visual fields provides pulse data;
Step 2: the orientation of carrying out three celestial bodies according to pulse data is determined: determine the coordinate E of the earth's core direction vector under measurement coordinate system
SE, day heart direction vector coordinate S under measurement coordinate system
S|SEWith the coordinate M of moon heart direction vector under measurement coordinate system
S|SE, and the earth's core is apart from r;
Step 3: by the coordinate E of the earth's core direction vector under measurement coordinate system
SE, day heart direction vector coordinate S under measurement coordinate system
S|SEWith the coordinate M of moon heart direction vector under measurement coordinate system
S|SECalculate the inner product of day heart direction vector and the earth's core direction vector, and the inner product of month heart direction vector and the earth's core direction vector:
Step 4: carry out two vectors and slightly decide appearance, the vector of three pairwise orthogonals of definition is as follows:
v
1:r
m×r
s/|r
m×r
s|
v
2:r
m/|r
m| (1)
v
3:v
1×v
2
Wherein, r
mBe star moon vector, r
sBe star day vector; Described star moon vector r
mWith star day vector r
sExpression under inertial coordinates system by the ground moon vector, day vector is approximate obtains, and a ground month vector and a ground day vector can be obtained by ephemeris; If v
1, v
2, v
3Under inertial coordinates system, be expressed as E
1, E
2, E
3, under measurement coordinate system, be expressed as e
1, e
2, e
3, then inertial coordinate is tied to the rotation matrix of measurement coordinate system:
R
SEI=[e
1,e
2,e
3][E
1,E
2,E
3]
-1 (2)
Step 5: the navigation initialization, the initial position guestimate is calculated with following formula:
r
0=-rR
ISEE
SE (3)
Wherein,
Obtain by formula (2); The initial velocity guestimate then has the position of two adjacent moment to do approximate difference to obtain;
Step 6: adopt the least square navigation algorithm to carry out thick navigation operations and obtain navigation results;
Step 7: utilize the navigation results of the thick navigation in the step 6, revise star moon vector r
mValue, thereby obtain revised star moon vector r
M repaiiesFinished becoming more meticulous of moon orientation;
Step 8: carry out two vectors essences and decide appearance, star moon vector r
mThe revised star moon vector r that expression under inertial coordinates system is produced by previous step
M repaiiesSubstitute star day vector r
sConstant, the vector of three pairwise orthogonals of definition is as follows:
v
1': r
M repaiies* r
s/ | r
M repaiies* r
s|
v
2': r
M repaiies/ | r
M repaiies|
v
3′:v
1×v
2
If v
1', v
2', v
3' under inertial system, be expressed as E
1', E
2', E
3', under body coordinate system, be expressed as e
1', e
2', e
3' also can ask; Then two vectors of Xiu Zhenging inertial coordinate of deciding appearance is tied to the rotation matrix of measurement coordinate system:
R
SEI′=[e
1′,e
2′,e
3′][E
1′,E
2′,E
3′]
-1 (4)
Step 9: consider that inertial coordinate that compression of the earth is decided appearance based on two vectors of revising in the step 8 is tied to the rotation matrix R of measurement coordinate system
SEI' revise the earth's core apart from r and the earth's core orientation vector E
SE, obtain revised the earth's core apart from r
RepairWith revised the earth's core orientation vector E
SE repaiies
Step 10: adopt the least square navigation algorithm to carry out smart navigation operations and obtain final navigation results.
2. the satellite high-precision autonomous navigation method based on day ground month integrated sensor pulse data according to claim 1 is characterized in that the coordinate E of the earth's core direction vector under measurement coordinate system in the step 2
SEDefinite process, and the earth's core is as follows apart from definite process of r:
Two infrared probes on the infrared double cone scanning type sensor of the earth sweep terrestrial time:
α
1-in=-((ω
rott
ref+B
R1)-ω
rott
1-in)
(5)
α
2-in=-((ω
rott
ref+B
R2)-ω
rott
2-in)
ω wherein
RotBe the rotating speed of scanning sensor, B
R1, B
R2Be the drag angle of two infrared probes on the initial time double cone scanning type sensor with respect to benchmark x axle, t
RefFor the slit sensor plane of symmetry passes benchmark x axle to deserved pulse constantly, t
1-in, t
2-inThe pulse that forms when sweeping Horizon for two infrared probes on the double cone scanning type sensor;
Two infrared probes on the infrared double cone scanning type sensor of the earth scan out terrestrial time:
α
1-out=ω
rott
1-out-(ω
rott
ref+B
R1)
(6)
α
2-out=ω
rott
2-out-(ω
rott
ref+B
R2)
T wherein
1-out, t
2-outThe pulse that forms when scanning out Horizon for two infrared probes on the double cone scanning type sensor;
So the earth's core direction vector with respect to the position angle of measurement coordinate system is:
(7)
If do not consider compression of the earth, φ arranged
E1=φ
E2, and use φ
eExpression;
Thereby the string of the earth that the infrared double cone scanning type sensor of the earth obtains is wide be:
Ω
E1=-α
1-in+α
1-out
(8)
Ω
E2=-α
2-in+α
2-out
By the spherical trigonometry formula, have
Wherein ρ is the half angle with respect to the infrared radiation disk of the earth of satellite, γ
1, γ
2Be the semi-cone angle of earth sensor double cone, η is the angle between the earth's core direction and the sensor turning axle;
So
If do not consider compression of the earth, the coordinate of the earth's core direction vector under measurement coordinate system try to achieve into:
The earth's core is apart from being calculated by following formula:
r=R
E/sinρ (11)
Wherein, R
EBe the infrared radiation radius of a ball of the earth, ρ can be tried to achieve by the cosine law of spherical triangle:
。
3. the satellite high-precision autonomous navigation method based on day ground month integrated sensor pulse data according to claim 1 is characterized in that the coordinate S of day heart direction vector under measurement coordinate system in the step 2
S|SEDefinite process as follows:
Position angle and the elevation angle of day heart orientation under the measurement body coordinate system is:
δ
s=cot
-1(sinσ
scosβ
s)
Wherein: t
1sun, t
2sunBe the responsive pulse that forms to sunshine time the in the first fan-shaped slit visual field, the second fan-shaped slit visual field, σ
s=(ω
Rott
2sun-ω
Rott
1sun-υ)/2, υ is the scanning life sensor of the first fan-shaped slit visual field is ahead of the scanning life sensor of the second fan-shaped slit visual field on the sensor equatorial plane a focal length, β
sBe the angle of slit visual field with respect to the scanning axes of rotation skew;
The coordinate of day heart direction vector under measurement coordinate system is:
4. the satellite high-precision autonomous navigation method based on day ground month integrated sensor pulse data according to claim 1 is characterized in that the coordinate M of moon heart direction vector under measurement coordinate system in the step 2
S|SEDefinite process as follows:
Position angle and the elevation angle of moon heart orientation under measurement coordinate system during the full moon is:
δ
m=tan
-1(sinσ
Mcosβ
s)
Wherein: t
1moon, t
2moonBe the responsive pulse that forms to the moon time in the first fan-shaped slit visual field, the second fan-shaped slit visual field, σ
M=(ω
Rott
2moon-ω
Rott
1moon-υ)/2;
Thereby the coordinate of moon heart direction vector under measurement coordinate system during the full moon is:
The coordinate M of moon heart direction vector under measurement coordinate system during the last lower edge
S|SEDefinite process as follows:
Obtain the attitude information of satellite by thick navigation, obtain sensor scan axis and x
mAngle theta
x, on the basis of original pulse, add a Δ t then;
Wherein, Δ t>0 in the time of the moon at the first quarter, Δ t<0 in the time of the moon at the last quarter; Calculate φ by formula (15) again
m, δ
m, by the coordinate of moon heart direction vector under measurement coordinate system during the last lower edge after the definite compensation of formula (16).
5. the satellite high-precision autonomous navigation method based on day ground month integrated sensor pulse data according to claim 1 is characterized in that step 7 correction star moon vector r
m, the preliminary definite position r of satellite under inertial coordinates system from thick navigation
b, and moon inertial position is tried to achieve by ephemeris, thus determine star moon vector r
mExpression under inertial coordinates system, correction formula is as follows:
r
M repaiies=R
m-r
b
6. the satellite high-precision autonomous navigation method based on day ground month integrated sensor pulse data according to claim 1, it is characterized in that step 9 to the earth's core apart from r and the earth's core orientation vector E
SERecomputate specific as follows:
By measuring geometric analysis, the earth's core is to the vector R of earth surface as can be known
EFor:
Wherein:
Be direction vector from the earth's core sensing satellite,
With the unit vector that scans sight line, make pulse t constantly respectively
1in, t
1out, t
2in, t
2outThe unit vector of corresponding scanning sight line is
If do not consider to measure noise, this tittle is accurately known, is respectively:
(18)
The earth's core is to the vector on earth infrared radiation ball surface
Be projected under the inertial coordinates system;
The ellipsoid equation is:
P=-1+1/ (1-ee) wherein
2.
By getting in the formula (17):
(20)
Wherein: L
3Be R
ISEThe third line;
With (20) substitution (19) to obtaining quadratic equation:
(a+Δa)l
2+(b+Δb)l+(c+Δc)=0 (21)
Wherein:
(b+Δb)
2-4(a+Δa)(c+Δc)=0 (22)
Also promptly:
(23)
With in the navigation slightly each constantly
Can from last equation, solve revised as initial value
Also should satisfy constraint condition
4 unknown numbers, 5 equations approach with least-squares estimation
Order
x
0Be the iterative initial value of x, with F at x
0Point single order Taylor expansion has:
Wherein:
And:
The form of being write as iteration has:
7. the satellite high-precision autonomous navigation method based on day ground month integrated sensor pulse data according to claim 1 is characterized in that the least square navigation algorithm in step 6 and the step 10 is:
Suppose N observation t constantly
1, t
2..., t
NObserved quantity be
Described observed quantity is angle between month ground and the angle between day ground, is obtained by the arc cosine of step 3 inner product, and its true value is
Make residual error be:
If do not measure noise
Set up the least square index:
Wherein:
To go up in the equation substitution functional index and can get:
Following formula can be remembered and does:
Wherein:
Thereby can in the hope of:
Corresponding iterative formula can be write:
X
k+1=X
k-A
-1B
。
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
CN 201110005239 CN102162731B (en) | 2011-01-12 | 2011-01-12 | High-precision satellite independent navigation method based on pulse data of sun, earth and moon integrated sensor |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
CN 201110005239 CN102162731B (en) | 2011-01-12 | 2011-01-12 | High-precision satellite independent navigation method based on pulse data of sun, earth and moon integrated sensor |
Publications (2)
Publication Number | Publication Date |
---|---|
CN102162731A true CN102162731A (en) | 2011-08-24 |
CN102162731B CN102162731B (en) | 2012-12-12 |
Family
ID=44464079
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
CN 201110005239 Expired - Fee Related CN102162731B (en) | 2011-01-12 | 2011-01-12 | High-precision satellite independent navigation method based on pulse data of sun, earth and moon integrated sensor |
Country Status (1)
Country | Link |
---|---|
CN (1) | CN102162731B (en) |
Cited By (11)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN102519471A (en) * | 2011-12-08 | 2012-06-27 | 北京控制工程研究所 | Imaging type earth sensor earth oblateness compensation method based on trigonometric function fitting |
CN102519454A (en) * | 2011-12-08 | 2012-06-27 | 北京控制工程研究所 | Selenocentric direction correction method for sun-earth-moon navigation |
CN102538784A (en) * | 2011-12-23 | 2012-07-04 | 北京控制工程研究所 | Oblateness correcting method for geocenter direction of sun, earth and moon navigation |
CN103279127A (en) * | 2013-05-22 | 2013-09-04 | 上海新跃仪表厂 | Autonomous control method for GEO (Geosynchronous) orbit satellite by using angle information only |
CN103398711A (en) * | 2013-08-07 | 2013-11-20 | 清华大学 | Multi-view-field-separated Earth sensor |
CN104408279A (en) * | 2014-10-09 | 2015-03-11 | 北京宇航系统工程研究所 | Space external heat flux calculation method for carrier rocket |
CN106525026A (en) * | 2016-11-01 | 2017-03-22 | 李清林 | Celestial positioning analysis method achieved by projecting subject position to spring equinox equatorial coordinate system |
CN106767844A (en) * | 2017-01-05 | 2017-05-31 | 北京航天自动控制研究所 | A kind of method for improving earth sensor body the earth's core vector accuracy |
CN111209523A (en) * | 2020-01-06 | 2020-05-29 | 中国科学院紫金山天文台 | Rapid processing method suitable for precise calculation of dense ephemeris of large eccentricity orbit |
CN111460898A (en) * | 2020-03-04 | 2020-07-28 | 北京空间飞行器总体设计部 | Skyline acquisition method based on monocular camera image of lunar surface inspection tour device |
CN112179334A (en) * | 2020-09-15 | 2021-01-05 | 中国科学院微小卫星创新研究院 | Star navigation method and system based on two-step Kalman filtering |
Citations (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5109346A (en) * | 1990-02-01 | 1992-04-28 | Microcosm, Inc. | Autonomous spacecraft navigation system |
-
2011
- 2011-01-12 CN CN 201110005239 patent/CN102162731B/en not_active Expired - Fee Related
Patent Citations (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5109346A (en) * | 1990-02-01 | 1992-04-28 | Microcosm, Inc. | Autonomous spacecraft navigation system |
Non-Patent Citations (4)
Title |
---|
《哈尔滨工业大学学报》 20021030 黄翔宇等 基于"日-地-月"信息的卫星自主导航技术研究 第34卷, 第05期 * |
张燕等: "基于日地月方位信息的月球卫星自主导航", 《宇航学报》 * |
荆武兴: "基于日地月方位信息的近地轨道卫星自主导航", 《宇航学报》 * |
黄翔宇等: "基于"日-地-月"信息的卫星自主导航技术研究", 《哈尔滨工业大学学报》 * |
Cited By (21)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN102519454A (en) * | 2011-12-08 | 2012-06-27 | 北京控制工程研究所 | Selenocentric direction correction method for sun-earth-moon navigation |
CN102519454B (en) * | 2011-12-08 | 2014-10-08 | 北京控制工程研究所 | Selenocentric direction correction method for sun-earth-moon navigation |
CN102519471A (en) * | 2011-12-08 | 2012-06-27 | 北京控制工程研究所 | Imaging type earth sensor earth oblateness compensation method based on trigonometric function fitting |
CN102538784A (en) * | 2011-12-23 | 2012-07-04 | 北京控制工程研究所 | Oblateness correcting method for geocenter direction of sun, earth and moon navigation |
CN102538784B (en) * | 2011-12-23 | 2015-05-27 | 北京控制工程研究所 | Oblateness correcting method for geocenter direction of sun, earth and moon navigation |
CN103279127A (en) * | 2013-05-22 | 2013-09-04 | 上海新跃仪表厂 | Autonomous control method for GEO (Geosynchronous) orbit satellite by using angle information only |
CN103279127B (en) * | 2013-05-22 | 2016-06-22 | 上海新跃仪表厂 | A kind of only by the GEO orbiter Autonomous Control method of angle information |
CN103398711A (en) * | 2013-08-07 | 2013-11-20 | 清华大学 | Multi-view-field-separated Earth sensor |
CN103398711B (en) * | 2013-08-07 | 2015-10-28 | 清华大学 | The earth sensor that many visual fields are separated |
CN104408279B (en) * | 2014-10-09 | 2017-07-28 | 北京宇航系统工程研究所 | A kind of carrier rocket space heat flux computational methods |
CN104408279A (en) * | 2014-10-09 | 2015-03-11 | 北京宇航系统工程研究所 | Space external heat flux calculation method for carrier rocket |
CN106525026A (en) * | 2016-11-01 | 2017-03-22 | 李清林 | Celestial positioning analysis method achieved by projecting subject position to spring equinox equatorial coordinate system |
CN106525026B (en) * | 2016-11-01 | 2019-09-03 | 李清林 | Project to the parsing astronomical positioning method of the first point of Aries equatorial system of coordinates |
CN106767844A (en) * | 2017-01-05 | 2017-05-31 | 北京航天自动控制研究所 | A kind of method for improving earth sensor body the earth's core vector accuracy |
CN106767844B (en) * | 2017-01-05 | 2019-05-28 | 北京航天自动控制研究所 | A method of improving earth sensor body geocentric vector precision |
CN111209523A (en) * | 2020-01-06 | 2020-05-29 | 中国科学院紫金山天文台 | Rapid processing method suitable for precise calculation of dense ephemeris of large eccentricity orbit |
CN111209523B (en) * | 2020-01-06 | 2020-12-29 | 中国科学院紫金山天文台 | Rapid processing method suitable for precise calculation of dense ephemeris of large eccentricity orbit |
US11319094B2 (en) | 2020-01-06 | 2022-05-03 | Purple Mountain Observatory, Chinese Academy Of Sciences | Method for accurately and efficiently calculating dense ephemeris of high-eccentricity orbit |
CN111460898A (en) * | 2020-03-04 | 2020-07-28 | 北京空间飞行器总体设计部 | Skyline acquisition method based on monocular camera image of lunar surface inspection tour device |
CN112179334A (en) * | 2020-09-15 | 2021-01-05 | 中国科学院微小卫星创新研究院 | Star navigation method and system based on two-step Kalman filtering |
CN112179334B (en) * | 2020-09-15 | 2023-03-14 | 中国科学院微小卫星创新研究院 | Star navigation method and system based on two-step Kalman filtering |
Also Published As
Publication number | Publication date |
---|---|
CN102162731B (en) | 2012-12-12 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
CN102162731B (en) | High-precision satellite independent navigation method based on pulse data of sun, earth and moon integrated sensor | |
CN101344391B (en) | Lunar vehicle posture self-confirming method based on full-function sun-compass | |
CN104655152B (en) | A kind of real-time Transfer Alignments of airborne distributed POS based on federated filter | |
CN103221788A (en) | Device and method of gyro sensor calibration | |
CN102175259B (en) | Autonomous navigation simulation test system based on earth-sun-moon integrated sensor | |
CN103913181B (en) | A kind of airborne distributed POS Transfer Alignments based on parameter identification | |
CN103900576B (en) | A kind of information fusion method of survey of deep space independent navigation | |
CN105371844B (en) | A kind of inertial navigation system initial method based on inertia/astronomical mutual assistance | |
CN106842271B (en) | Navigation positioning method and device | |
CN105160125B (en) | A kind of simulating analysis of star sensor quaternary number | |
CN105548976A (en) | Shipborne radar offshore precision identification method | |
CN103913180A (en) | Mounting angle calibration method for onboard large-view-field high-precision star sensor | |
CN102538819A (en) | Autonomous navigation semi-physical simulation test system based on biconical infrared and star sensors | |
CN109708663B (en) | Star sensor online calibration method based on aerospace plane SINS assistance | |
CN103968834B (en) | Autonomous celestial navigation method for deep space probe on near-earth parking orbit | |
CN102116634A (en) | Autonomous dimensionality reduction navigation method for deep sky object (DSO) landing detector | |
Gou et al. | INS/CNS integrated navigation based on corrected infrared earth measurement | |
CN103389099A (en) | Spacecraft attitude and position measurement system and method based on X-ray pulsar | |
CN102426025A (en) | Simulation analysis method for drift correction angle during remote sensing satellite attitude maneuver | |
JP5130965B2 (en) | Medium-altitude satellite acquisition method and apparatus | |
CN105737848B (en) | System-level star sensor star viewing system and star viewing method | |
Stark et al. | Mercury's rotational parameters from MESSENGER image and laser altimeter data: A feasibility study | |
CN102607563B (en) | System for performing relative navigation on spacecraft based on background astronomical information | |
CN102519454B (en) | Selenocentric direction correction method for sun-earth-moon navigation | |
CN109099911B (en) | Navigation positioning method and system for aviation system |
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 | ||
C17 | Cessation of patent right | ||
CF01 | Termination of patent right due to non-payment of annual fee |
Granted publication date: 20121212 Termination date: 20140112 |