CN111174779A - Inertia-astronomy combined navigation method for deep space exploration aircraft - Google Patents

Inertia-astronomy combined navigation method for deep space exploration aircraft Download PDF

Info

Publication number
CN111174779A
CN111174779A CN201911202097.3A CN201911202097A CN111174779A CN 111174779 A CN111174779 A CN 111174779A CN 201911202097 A CN201911202097 A CN 201911202097A CN 111174779 A CN111174779 A CN 111174779A
Authority
CN
China
Prior art keywords
aircraft
inertial
sun
error correction
relative
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
Application number
CN201911202097.3A
Other languages
Chinese (zh)
Other versions
CN111174779B (en
Inventor
王献忠
张肖
刘宇
张国柱
张丽敏
施常勇
刘赟
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Shanghai Aerospace Control Technology Institute
Original Assignee
Shanghai Aerospace Control Technology Institute
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Shanghai Aerospace Control Technology Institute filed Critical Shanghai Aerospace Control Technology Institute
Priority to CN201911202097.3A priority Critical patent/CN111174779B/en
Publication of CN111174779A publication Critical patent/CN111174779A/en
Application granted granted Critical
Publication of CN111174779B publication Critical patent/CN111174779B/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • 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
    • G01C21/165—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 combined with non-inertial navigation instruments
    • 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
    • G01C21/20—Instruments for performing navigational calculations
    • 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
    • G01C21/24—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 specially adapted for cosmonautical navigation

Landscapes

  • Engineering & Computer Science (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • Physics & Mathematics (AREA)
  • Automation & Control Theory (AREA)
  • General Physics & Mathematics (AREA)
  • Astronomy & Astrophysics (AREA)
  • Navigation (AREA)

Abstract

The invention provides an inertia-astronomical combined navigation method for a deep space exploration aircraft, which comprises the following steps: s1, establishing a centroid inertial coordinate system and a centroid orbit coordinate system, and calculating the direction vector of the aircraft in the centroid inertial system relative to the sun based on the angle obtained by the measurement of the optical sensor; s2, calculating the direction vector of the earth relative to the sun in the centroid inertia system; s3, resolving a direction vector of the aircraft relative to the earth center in a sun-center inertial system according to the inertial navigation measurement result; s4, calculating the direction vector of the sun relative to the aircraft in the centroid inertial system; s5, estimating a position/speed error correction quantity by PI filtering based on the direction vector error; and S6, and correcting the inertia-astronomical combined navigation based on the position/speed error. The deep space exploration aircraft inertia-astronomical combined navigation method can inhibit inertial navigation accumulated errors when the aircraft is in-orbit and navigated in real time.

Description

Inertia-astronomy combined navigation method for deep space exploration aircraft
Technical Field
The invention relates to the technical field of deep space astronomical navigation, in particular to an inertia-astronomical combined navigation method of a deep space exploration aircraft based on a star sensor and an optical sensor, which is used for the on-orbit real-time navigation of the aircraft.
Background
The existing earth satellite and moon exploration aircraft are close to the earth, so that the earth measurement and control delay is short, more sensors are available for establishing an earth measurement and control link, and the reliability of the earth measurement and control rail is high. The deep space exploration aircraft is far away from the earth, the ground measurement and control delay is thirty minutes, a reliable ground measurement and control link is needed for a ground measurement and control rail, and the deep space exploration aircraft has certain autonomous navigation capability and is beneficial to establishment of the ground measurement and control link.
The method comprises the following steps that during the in-orbit operation of the aircraft, not only attitude control but also orbit control is carried out, and orbit parameters are required for the orbit control; the star sensor measurement gives the attitude relative to the J2000 system, while the target attitude is generally relative to the orbital system, and the satellite sensor calculation of the attitude relative to the orbital system requires orbital parameters.
The method is directly based on the inertial navigation of a gyro and an adding table, navigation errors are accumulated along with time, and the method cannot be applied for a long time. Inertial navigation is generally combined with GNSS compatible machines, magnetometers, pulsar sensors, planetary sensors, satellite sensors and the like. However, GNSS compatible machines, magnetometers are not available for deep space exploration; pulsar-based navigation requires the configuration of complex and expensive pulsar sensors. The deep space exploration aircraft orbit determination is generally based on a ground orbit determination, and the autonomous navigation on the aircraft generally adopts inertia-astronomy combined navigation.
Disclosure of Invention
The invention aims to provide an inertia-astronomical combined navigation method for a deep space exploration aircraft, which is used for carrying out the inertia-astronomical combined navigation on the deep space exploration aircraft based on a star sensor and an optical sensor, can inhibit the accumulated error of the inertial navigation, is simple in PI filtering combined navigation algorithm, is based on a conventional sensor and is easy to realize in engineering.
In order to achieve the above object, the present invention provides a deep space exploration aircraft inertia-astronomical combined navigation method, comprising the steps of:
s1, establishing a centroid inertial coordinate system and a centroid orbit coordinate system; calculating a direction vector of the aircraft relative to the sun in a sun center inertial system based on an angle obtained by measurement of the optical sensor;
s2, calculating the direction vector of the earth relative to the sun in the centroid inertia system;
s3, resolving a direction vector of the aircraft relative to the earth center in a sun-center inertial system according to the inertial navigation measurement result;
s4, calculating the direction vector of the sun relative to the aircraft in the centroid inertial system;
s5, estimating a position/speed error correction quantity by PI filtering based on the direction vector error;
and S6, and correcting the inertia-astronomical combined navigation based on the position/speed error.
In step S1, the method for calculating the direction vector of the aircraft in the centroid inertial system relative to the sun based on the angle measured by the optical sensor specifically includes the steps of:
s11, measuring the lower roll angle of the sensor system according to the optical sensor
Figure BDA0002296130730000021
Pitch angle thetasCalculating the rolling angle of the sun vector of the aircraft on the aircraft body
Figure BDA0002296130730000022
A pitch angle theta; wherein psiHThe optical sensor is arranged at an angle in a horizontal plane;
Figure BDA0002296130730000023
s12, rolling angle of the aircraft body according to the aircraft sun vector
Figure BDA0002296130730000024
The pitch angle theta is used for calculating the direction vector of the aircraft body relative to the sun in the system
Figure BDA0002296130730000025
Figure BDA0002296130730000026
S13, solving the direction vector of the aircraft body relative to the sun in the centroid inertia system
Figure BDA0002296130730000027
Figure BDA0002296130730000028
Wherein, CJ2000←bIs a rotation matrix of the aircraft body to the J2000 inertia system, which is obtained by calculating the quaternion of the body relative to the inertia system measured by a star sensor,
Figure BDA0002296130730000029
in step S2, the specific method for calculating the direction vector of the earth relative to the sun in the centroid inertial system is as follows:
Figure BDA00022961307300000210
wherein
Figure BDA00022961307300000211
Is the direction vector of the earth relative to the sun in the inertial system of the centroid,
Figure BDA00022961307300000212
is the direction vector of the sun vector in the J2000 inertial system.
In step S3, the specific method for calculating the direction vector of the inertial system of the aircraft relative to the geocentric in the heliocentric mode is as follows:
Figure BDA0002296130730000031
wherein the content of the first and second substances,
Figure BDA0002296130730000032
the vector of the aircraft relative to the earth center in the direction of the sun-center inertial system; r isb_J2000The position of the aircraft in a J2000 inertial system is obtained according to inertial navigation measurement.
In step S4, the specific method for calculating the direction vector of the aircraft relative to the sun in the centroid inertial system is as follows:
Figure BDA0002296130730000033
wherein the content of the first and second substances,
Figure BDA0002296130730000034
is the vector of the aircraft relative to the sun in the direction of the sun-center inertial system,
Figure BDA0002296130730000035
the direction vector of the earth relative to the sun in the centroid inertial system is shown.
Step S5 specifically includes:
s51, calculating the direction error of the aircraft relative to the sun in the sun center inertial system
Figure BDA0002296130730000036
Wherein
Figure BDA0002296130730000037
S52, position error correction is estimated by adopting PI filtering; limiting the position error delta r, and calculating the position error correction of the k step by inertial navigation
Figure BDA0002296130730000076
Figure BDA0002296130730000077
For the k-th step position error after clipping, kp,rEstimating a scaling factor for the speed error correction; k is a radical ofp,rThe position error correction quantity can be independently estimated for a 3 multiplied by 3 diagonal array; i represents the three axes x, y, z.
S53, estimating a speed error correction quantity based on PI filtering; correction dr by the k step position errori,kEstimating the k-th step speed error correction dvi,k(ii) a Pair dri,kLimiting to prevent integral saturation, setting dri,kThe k-th step position error correction after amplitude limiting is
Figure BDA00022961307300000310
Speed error correction dv of kth step calculated by inertial navigationi,kTo obtain
Figure BDA00022961307300000311
Wherein k isp,vEstimating a scaling factor, k, for a speed error correctionp,vThe speed error correction can be estimated independently for a 3 x 3 diagonal matrix.
Step S6 specifically includes:
s61, obtaining the position r of the aircraft based on the aircraft inertial navigation calculation algorithm of the J2000 inertial systemiVelocity vi(ii) a Wherein i represents the three axes x, y, z.
S62, based on the position/speed error correction quantity combined navigation, deducting the position/speed error correction quantity in the J2000 inertial system, and applying a simplified integral algorithm to carry out inertial navigation calculation to obtain the speed v of deducting the speed error correction quantity in the k stepi,kThe k-th step deducts the position r of the position error correctioni,k；
vi,k＝vi,k-1+[ai,k-1+(ai,k-ai,k-1)/2]·T-dvi,k-1
ri,k＝ri,k-1+[vi,k-1+(vi,k-vi,k-1)/2]·T-dri,k-1
Wherein the content of the first and second substances,
ai,k-1is as followsk-1 step acceleration;
ai,kacceleration of the kth step;
dvi,k-1the speed error correction amount estimated for the k-1 step;
vi,k-1deducting the speed of the speed error correction quantity for the k-1 step;
dri,k-1the position error correction amount estimated for the (k-1) th step;
ri,k-1deducting the position of the position error correction quantity for the k-1 step;
and T is a navigation resolving period.
Step S61 specifically includes:
s611, let the position vector of the aircraft under the inertia of J2000 be r ═ x y z]TThen, the three axial components of the sun gravity acceleration of the aircraft under the J2000 inertial system are respectively:
Figure BDA0002296130730000041
Figure BDA0002296130730000042
Figure BDA0002296130730000043
wherein the content of the first and second substances,
Figure BDA0002296130730000044
distance of aircraft to the centre of the day, μsIs the solar gravitational constant;
s612, enabling the attitude transformation matrix from the accelerometer coordinate system to the J2000 inertial system to be AiaTo obtain the non-inertial acceleration a of the aircraft under the J2000 inertial systema,i＝Aia·aa,a(ii) a Wherein, aa,aAcceleration under an accelerometer coordinate system measured by an accelerometer;
s613, order ag,i＝[agx,iagy,iagz,i]TTo obtain the acceleration of the aircraft under the J2000 inertial systemDegree aiWherein i represents the three axes x, y and z.
ai＝ag,i+aa,i；
S614, carrying out inertial navigation calculation on the J2000 inertial system to obtain the position r of the aircraftiVelocity vi，
Wherein v isi＝∫ai·dt，ri＝∫vi·dt。
Compared with the prior art, the deep space exploration aircraft inertia-astronomy combined navigation method can inhibit inertial navigation accumulated errors when the aircraft is in orbit to navigate in real time. The PI filtering integrated navigation algorithm adopted by the invention is simple, is based on a conventional sensor and is easy to realize in engineering.
Drawings
In order to more clearly illustrate the technical solution of the present invention, the drawings used in the description will be briefly introduced, and it is obvious that the drawings in the following description are an embodiment of the present invention, and other drawings can be obtained by those skilled in the art without creative efforts according to the drawings:
FIG. 1 is a schematic view of a centroid trajectory system of the present invention;
FIG. 2 is a schematic diagram of the relationship between the aircraft, the sun and the earth in the present invention;
FIG. 3 is a flow chart of the inertia-astronomical combined navigation method of the deep space exploration aircraft according to the invention.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
The invention provides an inertia-astronomical combined navigation method for a deep space exploration aircraft, which comprises the following steps of:
s1, establishing a centroid inertial coordinate system, wherein the specific method comprises the following steps: the sun is used as an origin, the X axis points to the J2000 spring break point, the Z axis points to the geocentric, and the Y axis is determined according to the right-hand rule.
Establishing a centroid orbit coordinate system, as shown in fig. 1, the specific method is as follows: and taking the mass center of the aircraft as an origin, pointing the Z axis to the sun, taking the X axis as the flight direction of the aircraft, and determining the Y axis according to a right-hand rule.
The method comprises the following steps of calculating a direction vector of the aircraft relative to the sun in a sun-center inertial system based on an angle obtained by measurement of an optical sensor, and specifically comprises the following steps:
s11, measuring the lower roll angle of the sensor system according to the optical sensor
Figure BDA0002296130730000061
Pitch angle thetasCalculating the rolling angle of the sun vector of the aircraft on the aircraft body
Figure BDA0002296130730000062
A pitch angle theta; wherein psiHThe optical sensor is arranged at an angle in a horizontal plane;
Figure BDA0002296130730000063
s12, rolling angle of the aircraft body according to the aircraft sun vector
Figure BDA0002296130730000064
The pitch angle theta is used for calculating the direction vector of the aircraft body relative to the sun in the system
Figure BDA0002296130730000065
Figure BDA0002296130730000066
S13, solving the direction vector of the aircraft body relative to the sun in the centroid inertia system
Figure BDA0002296130730000067
Figure BDA0002296130730000068
Wherein: cJ2000←bThe rotation matrix of the aircraft body to the J2000 inertia system is obtained by calculating the quaternion of the body relative to the inertia system measured by the star sensor.
S2, calculating the direction vector of the earth relative to the sun in the centroid inertia system; the specific method comprises the following steps:
Figure BDA0002296130730000069
wherein
Figure BDA00022961307300000610
Is the direction vector of the earth relative to the sun in the inertial system of the centroid,
Figure BDA00022961307300000611
is the direction vector of the sun vector in the J2000 inertial system.
S3, obtaining the position r of the aircraft in the J2000 inertial system according to inertial navigation measurementb_J2000And resolving the direction vector of the aircraft relative to the earth center in the sun-center inertial system, wherein the specific method comprises the following steps:
Figure BDA00022961307300000612
wherein the content of the first and second substances,
Figure BDA00022961307300000613
the vector of the aircraft relative to the earth center in the direction of the sun-center inertial system; .
S4, calculating the direction vector of the aircraft relative to the sun in the sun-center inertial system, wherein the specific method comprises the following steps:
Figure BDA00022961307300000614
wherein the content of the first and second substances,
Figure BDA00022961307300000615
is the vector of the aircraft relative to the sun in the direction of the sun-center inertial system,
Figure BDA00022961307300000616
the vector of the sphere relative to the sun in the direction of the centroid inertial system.
In the context of figure 2, it is shown,
Figure BDA0002296130730000071
the direction vector of the earth relative to the sun in the centroid inertial system;
Figure BDA0002296130730000072
the vector of the aircraft relative to the sun in the direction of the sun-centered inertial system;
Figure BDA0002296130730000073
the vector of the aircraft relative to the earth center in the direction of the sun-center inertial system.
S5, estimating a position/speed error correction quantity by PI (proportional integral) filtering based on the direction vector error;
step S5 specifically includes:
s51, calculating the direction error of the aircraft relative to the sun in the sun center inertial system
Figure BDA0002296130730000074
Wherein
Figure BDA0002296130730000075
S52, position error correction is estimated by adopting PI filtering; limiting the position error delta r, and calculating the position error correction of the k step by inertial navigation
Figure BDA0002296130730000076
Figure BDA0002296130730000077
For the k-th step position error after clipping, kp,rFor speed error correction estimationCalculating a proportionality coefficient; k is a radical ofp,rThe position error correction quantity can be independently estimated for a 3 multiplied by 3 diagonal array, and i represents three axes of x, y and z;
let dri,k＝[dxi,kdyi,kdzi,k]T、
Figure BDA0002296130730000078
The triaxial independent estimation position error correction is in the form of:
Figure BDA0002296130730000079
Figure BDA00022961307300000710
Figure BDA00022961307300000711
wherein: k is a radical ofp,x、kp,y、kp,zScaling factors are estimated for the three axis position error correction.
S53, estimating a speed error correction quantity based on PI filtering; the correction amount dr is based on the position error in consideration of the fact that the position error also reflects the speed errori,kA speed error correction is estimated. Correction dr by the k step position errori,kEstimating the k-th step speed error correction dvi,k(ii) a Pair dri,kLimiting to prevent integral saturation, setting dri,kThe k-th step position error correction after amplitude limiting is
Figure BDA00022961307300000712
Speed error correction dv of kth step calculated by inertial navigationi,kTo obtain
Figure BDA0002296130730000081
Wherein k isp,vEstimating a scaling factor, k, for a speed error correctionp,vThe speed error correction can be estimated independently for a 3 x 3 diagonal matrix.
Let dvi,k＝[dvxi,kdvyi,kdvzi,k]T，
Figure BDA0002296130730000082
The triaxial independent estimation speed error correction is in the form as follows:
Figure BDA0002296130730000083
Figure BDA0002296130730000084
Figure BDA0002296130730000085
wherein: k is a radical ofp,dx、kp,dy、kp,dzA scaling factor is estimated for the three axis velocity error correction.
And S6, and correcting the inertia-astronomical combined navigation based on the position/speed error.
Step S6 specifically includes:
s61, obtaining the position r of the aircraft based on the aircraft inertial navigation calculation algorithm of the J2000 inertial systemiVelocity viI represents the three axes x, y and z;
step S61 specifically includes:
s611, let the position vector of the aircraft under the inertia of J2000 be r ═ x y z]TThen, the three axial components of the sun gravity acceleration of the aircraft under the J2000 inertial system are respectively:
Figure BDA0002296130730000086
Figure BDA0002296130730000087
Figure BDA0002296130730000088
wherein the content of the first and second substances,
Figure BDA0002296130730000089
distance of aircraft to the centre of the day, μsIs the solar gravitational constant;
s612, enabling the attitude transformation matrix from the accelerometer coordinate system to the J2000 inertial system to be AiaTo obtain the non-inertial acceleration a of the aircraft under the J2000 inertial systema,i＝Aia·aa,a(ii) a Wherein, aa,aAcceleration under an accelerometer coordinate system measured by an accelerometer;
s613, order ag,i＝[agx,iagy,iagz,i]TCalculating the acceleration a of the aircraft under the J2000 inertial systemiI denotes the three axes x, y, z, ai＝ag,i+aa,i；
S614, carrying out inertial navigation calculation on the J2000 inertial system to obtain the position r of the aircraftiVelocity vi，
Wherein v isi＝∫ai·dt，ri＝∫vi·dt。
S62, combining navigation based on the position/speed error correction quantity, combining with the inertial navigation to calculate the integral process, and gradually deducting the position/speed error correction quantity in the inertial navigation integral process, thereby ensuring the stability of error correction.
Deducting the position/speed error correction quantity in the J2000 inertial system, and applying a simplified integral algorithm to carry out inertial navigation calculation to obtain the speed v of deducting the speed error correction quantity in the k stepi,kThe k-th step deducts the position r of the position error correctioni,k；
vi,k＝vi,k-1+[ai,k-1+(ai,k-ai,k-1)/2]·T-dvi,k-1
ri,k＝ri,k-1+[vi,k-1+(vi,k-vi,k-1)/2]·T-dri,k-1
Wherein the content of the first and second substances,
ai,k-1acceleration of the k-1 step;
ai,kacceleration of the kth step;
dvi,k-1the speed error correction amount estimated for the k-1 step;
vi,k-1deducting the speed of the speed error correction quantity for the k-1 step;
dri,k-1the position error correction amount estimated for the (k-1) th step;
ri,k-1deducting the position of the position error correction quantity for the k-1 step;
and T is a navigation resolving period.
Compared with the prior art, the deep space exploration aircraft inertia-astronomy combined navigation method can inhibit inertial navigation accumulated errors when the aircraft is in orbit to navigate in real time. The PI filtering integrated navigation algorithm adopted by the invention is simple, is based on a conventional sensor and is easy to realize in engineering.
While the invention has been described with reference to specific embodiments, the invention is not limited thereto, and various equivalent modifications and substitutions can be easily made by those skilled in the art within the technical scope of the invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.

Claims (8)

1. An inertia-astronomical combined navigation method for a deep space exploration aircraft is characterized by comprising the following steps:
s1, establishing a centroid inertial coordinate system and a centroid orbit coordinate system; calculating a direction vector of the aircraft relative to the sun in a sun center inertial system based on an angle obtained by measurement of the optical sensor;
s2, calculating the direction vector of the earth relative to the sun in the centroid inertia system;
s3, resolving a direction vector of the aircraft relative to the earth center in a sun-center inertial system according to the inertial navigation measurement result;
s4, calculating the direction vector of the sun relative to the aircraft in the centroid inertial system;
s5, estimating a position/speed error correction quantity by PI filtering based on the direction vector error;
and S6, and correcting the inertia-astronomical combined navigation based on the position/speed error.
2. The method for integrated inertia-astronomical navigation of a deep space exploration vehicle as claimed in claim 1, wherein said step S1 comprises the steps of:
s11, measuring the lower roll angle of the sensor system according to the optical sensor
Figure FDA0002296130720000011
Pitch angle thetasCalculating the rolling angle of the sun vector of the aircraft on the aircraft body
Figure FDA0002296130720000012
A pitch angle theta; wherein psiHThe optical sensor is arranged at an angle in a horizontal plane;
Figure FDA0002296130720000013
s12, rolling angle of the aircraft body according to the aircraft sun vector
Figure FDA0002296130720000014
The pitch angle theta is used for calculating the direction vector of the aircraft body relative to the sun in the system
Figure FDA0002296130720000015
Figure FDA0002296130720000016
S13, solving the direction vector of the aircraft body relative to the sun in the centroid inertia system
Figure FDA0002296130720000017
Figure FDA0002296130720000018
Wherein, CJ2000←bIs a rotation matrix of the aircraft body to the J2000 inertia system, which is obtained by calculating the quaternion of the body relative to the inertia system measured by a star sensor,
Figure FDA0002296130720000019
3. the combined inertial-astronomical navigation method for deep space exploration vehicles according to claim 1, wherein the specific method for calculating the direction vector of the earth relative to the sun in the inertial system of the day center in step S2 is as follows:
Figure FDA0002296130720000021
wherein
Figure FDA0002296130720000022
Is the direction vector of the earth relative to the sun in the inertial system of the centroid,
Figure FDA0002296130720000023
is the direction vector of the sun vector in the J2000 inertial system.
4. The combined inertial-astronomical navigation method for deep space exploration vehicles according to claim 1, wherein in step S3, the specific method for calculating the direction vector of the vehicle relative to the earth center in the inertial system of the sun center is as follows:
Figure FDA0002296130720000024
wherein the content of the first and second substances,
Figure FDA0002296130720000025
the vector of the aircraft relative to the earth center in the direction of the sun-center inertial system; r isb_J2000The position of the aircraft in a J2000 inertial system is obtained according to inertial navigation measurement.
5. The combined inertial-astronomical navigation method for a deep space exploration vehicle as claimed in claim 4, wherein in step S4, the specific method for calculating the direction vector of the vehicle in the centrosolar inertial system relative to the sun is as follows:
Figure FDA0002296130720000026
wherein the content of the first and second substances,
Figure FDA0002296130720000027
is the vector of the aircraft relative to the sun in the direction of the sun-center inertial system,
Figure FDA0002296130720000028
the direction vector of the earth relative to the sun in the centroid inertial system is shown.
6. The combined inertial-astronomical navigation method for deep space exploration vehicles according to claim 5, wherein step S5 specifically comprises:
s51, calculating the direction error of the aircraft relative to the sun in the sun center inertial system
Figure FDA0002296130720000029
Wherein
Figure FDA00022961307200000210
S52, position error correction is estimated by adopting PI filtering; limiting the position error delta r, and calculating the position error correction of the k step by inertial navigation
Figure DEST_PATH_BDA0002296130730000076
Figure DEST_PATH_BDA0002296130730000077
For the k-th step after clippingError, kp,rEstimating a scaling factor for the speed error correction; k is a radical ofp,rThe position error correction quantity can be independently estimated for a 3 multiplied by 3 diagonal array; i represents the three axes x, y and z;
s53, estimating a speed error correction quantity based on PI filtering; correction dr by the k step position errori,kEstimating the k-th step speed error correction dvi,k(ii) a Pair dri,kLimiting to prevent integral saturation, setting dri,kThe k-th step position error correction after amplitude limiting is
Figure FDA0002296130720000031
Speed error correction dv of kth step calculated by inertial navigationi,kTo obtain
Figure FDA0002296130720000032
Wherein k isp,vEstimating a scaling factor, k, for a speed error correctionp,vThe speed error correction can be estimated independently for a 3 x 3 diagonal matrix.
7. The combined inertial-astronomical navigation method for deep space exploration vehicles according to claim 5, wherein step S6 specifically comprises:
s61, obtaining the position r of the aircraft based on the aircraft inertial navigation calculation algorithm of the J2000 inertial systemiVelocity vi(ii) a Wherein i represents the three axes x, y, z;
s62, based on the position/speed error correction quantity combined navigation, deducting the position/speed error correction quantity in the J2000 inertial system, and applying a simplified integral algorithm to carry out inertial navigation calculation to obtain the speed v of deducting the speed error correction quantity in the k stepi,kThe k-th step deducts the position r of the position error correctioni,k；
vi,k＝vi,k-1+[ai,k-1+(ai,k-ai,k-1)/2]·T-dvi,k-1
ri,k＝ri,k-1+[vi,k-1+(vi,k-vi,k-1)/2]·T-dri,k-1
Wherein the content of the first and second substances,
ai,k-1acceleration of the k-1 step;
ai,kacceleration of the kth step;
dvi,k-1the speed error correction amount estimated for the k-1 step;
vi,k-1deducting the speed of the speed error correction quantity for the k-1 step;
dri,k-1the position error correction amount estimated for the (k-1) th step;
ri,k-1deducting the position of the position error correction quantity for the k-1 step;
and T is a navigation resolving period.
8. The combined inertial-astronomical navigation method for deep space exploration vehicles according to claim 6, wherein step S61 specifically comprises:
s611, let the position vector of the aircraft under the inertia of J2000 be r ═ x y z]TThen, the three axial components of the sun gravity acceleration of the aircraft under the J2000 inertial system are respectively:
Figure FDA0002296130720000041
Figure FDA0002296130720000042
Figure FDA0002296130720000043
wherein the content of the first and second substances,
Figure FDA0002296130720000044
distance of aircraft to the centre of the day, μsIs the solar gravitational constant;
s612, enabling the attitude transformation matrix from the accelerometer coordinate system to the J2000 inertial system to be AiaTo obtain the non-inertial acceleration a of the aircraft under the J2000 inertial systema,i＝Aia·aa,a(ii) a Wherein, aa,aAcceleration measured for an accelerometerMeasuring the acceleration under a coordinate system;
s613, order ag,i＝[agx,iagy,iagz,i]TCalculating the acceleration a of the aircraft under the J2000 inertial systemiWherein i represents the three axes x, y, z;
ai＝ag,i+aa,i；
s614, carrying out inertial navigation calculation on the J2000 inertial system to obtain the position r of the aircraftiVelocity vi，
Wherein v isi＝∫ai·dt，ri＝∫vi·dt。
CN201911202097.3A 2019-11-29 2019-11-29 Inertia-astronomy combined navigation method for deep space exploration aircraft Active CN111174779B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN201911202097.3A CN111174779B (en) 2019-11-29 2019-11-29 Inertia-astronomy combined navigation method for deep space exploration aircraft

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CN201911202097.3A CN111174779B (en) 2019-11-29 2019-11-29 Inertia-astronomy combined navigation method for deep space exploration aircraft

Publications (2)

Publication Number Publication Date
CN111174779A true CN111174779A (en) 2020-05-19
CN111174779B CN111174779B (en) 2021-11-05

Family

ID=70653775

Family Applications (1)

Application Number Title Priority Date Filing Date
CN201911202097.3A Active CN111174779B (en) 2019-11-29 2019-11-29 Inertia-astronomy combined navigation method for deep space exploration aircraft

Country Status (1)

Country Link
CN (1) CN111174779B (en)

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101788296A (en) * 2010-01-26 2010-07-28 北京航空航天大学 SINS/CNS deep integrated navigation system and realization method thereof
CN103063216A (en) * 2013-01-06 2013-04-24 南京航空航天大学 Inertial and celestial combined navigation method based on star coordinate modeling
CN103913169A (en) * 2014-03-12 2014-07-09 哈尔滨工程大学 Strap-down inertial/starlight refraction combined navigation method of aircrafts
EP3073223A1 (en) * 2015-03-24 2016-09-28 Honeywell International Inc. Tightly coupled celestial-intertial navigation system
CN106153051A (en) * 2016-06-29 2016-11-23 上海航天控制技术研究所 A kind of spacecraft shading device combined navigation methods
CN107179084A (en) * 2017-06-27 2017-09-19 上海航天控制技术研究所 GNSS compatibles pseudorange is with adding table pack navigation and drift estimate method
CN108592945A (en) * 2018-03-27 2018-09-28 中国人民解放军国防科技大学 Online calibration method for errors of inertia/astronomical combination system

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101788296A (en) * 2010-01-26 2010-07-28 北京航空航天大学 SINS/CNS deep integrated navigation system and realization method thereof
CN103063216A (en) * 2013-01-06 2013-04-24 南京航空航天大学 Inertial and celestial combined navigation method based on star coordinate modeling
CN103913169A (en) * 2014-03-12 2014-07-09 哈尔滨工程大学 Strap-down inertial/starlight refraction combined navigation method of aircrafts
EP3073223A1 (en) * 2015-03-24 2016-09-28 Honeywell International Inc. Tightly coupled celestial-intertial navigation system
US20160282123A1 (en) * 2015-03-24 2016-09-29 Honeywell International Inc. Tightly coupled celestial-intertial navigation system
CN106153051A (en) * 2016-06-29 2016-11-23 上海航天控制技术研究所 A kind of spacecraft shading device combined navigation methods
CN107179084A (en) * 2017-06-27 2017-09-19 上海航天控制技术研究所 GNSS compatibles pseudorange is with adding table pack navigation and drift estimate method
CN108592945A (en) * 2018-03-27 2018-09-28 中国人民解放军国防科技大学 Online calibration method for errors of inertia/astronomical combination system

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
YANQIANG YANG;CHUNXI ZHANG;JIAZHEN LU;HAO ZHANG: "Classification of Methods in the SINS/CNS Integration Navigation System", 《IEEE ACCESS》 *
张承，等: "直接敏感地平的空天飞行器惯性/天文组合方法", 《中国空间科学技术》 *

Also Published As

Publication number Publication date
CN111174779B (en) 2021-11-05

Similar Documents

Publication Publication Date Title
CN108051866B (en) Based on strap down inertial navigation/GPS combination subsidiary level angular movement isolation Gravimetric Method
Searcy et al. Magnetometer-only attitude determination using novel two-step Kalman filter approach
CN100356139C (en) Miniature assembled gesture measuring system for mini-satellite
CN109459065B (en) Gyro installation matrix calibration method based on satellite inertial space rotation attitude
CN108181916B (en) Control method and device for relative attitude of small satellite
CN109823571A (en) A kind of multistage attitude control method of remote sensing micro-nano satellite
CN105806369B (en) A kind of in-orbit aberration modification method of star sensor
JPH0577245B2 (en)
CN106153051B (en) A kind of spacecraft shading device combined navigation methods
Godard et al. Orbit determination of Rosetta around comet 67P/Churyumov-Gerasimenko
CN103278165B (en) Based on the magnetic survey of remanence calibration and the autonomous navigation method of starlight backup
CN106989761A (en) A kind of spacecraft Guidance instrumentation on-orbit calibration method based on adaptive-filtering
CN109460049B (en) Geosynchronous orbit satellite apogee orbit transfer method based on inertial pointing mode
CN110296719B (en) On-orbit calibration method
CN103630146A (en) Laser gyroscope IMU (inertial measurement unit) calibration method combining discrete analysis and Kalman filtration
CN103968844B (en) Big oval motor-driven Spacecraft Autonomous Navigation method based on low rail platform tracking measurement
CN104833375B (en) A kind of IMU Two position methods by star sensor
CN110285815A (en) It is a kind of can in-orbit whole-process application micro-nano satellite multi-source information attitude determination method
CN105606093B (en) Inertial navigation method and device based on gravity real-Time Compensation
CN107179084B (en) Pseudo-range and tabulation combined navigation and drift estimation method for GNSS compatible machine
CN109489661A (en) Gyro constant value drift estimation method when a kind of satellite is initially entered the orbit
CN110667892B (en) Satellite despinning control method based on geomagnetic measurement
CN111174779B (en) Inertia-astronomy combined navigation method for deep space exploration aircraft
CN111351490B (en) Method for quickly reconstructing inertial navigation reference in planet landing process
CN107132850B (en) Change rail posture based on angular speed tracking keeps control method

Legal Events

Date Code Title Description
PB01 Publication
PB01 Publication
SE01 Entry into force of request for substantive examination
SE01 Entry into force of request for substantive examination
GR01 Patent grant
GR01 Patent grant