CN111174779B - Inertia-astronomy combined navigation method for deep space exploration aircraft - Google Patents

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

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CN111174779B
CN111174779B CN201911202097.3A CN201911202097A CN111174779B CN 111174779 B CN111174779 B CN 111174779B CN 201911202097 A CN201911202097 A CN 201911202097A CN 111174779 B CN111174779 B CN 111174779B
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inertial
sun
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CN111174779A (en
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王献忠
张肖
刘宇
张国柱
张丽敏
施常勇
刘赟
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Shanghai Aerospace Control Technology Institute
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    • 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

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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, carrying out inertia-astronomical combined navigation based on the position/speed error correction.
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 GDA0003166505990000021
Pitch angle thetasCalculating the rolling angle of the sun vector of the aircraft on the aircraft
Figure GDA0003166505990000022
A pitch angle theta; wherein psi is the installation angle of the optical sensor in the horizontal plane;
Figure GDA0003166505990000023
s12, rolling angle of the aircraft according to the sun vector of the aircraft
Figure GDA0003166505990000024
A pitch angle theta, calculating the direction vector of the aircraft relative to the sun in the system
Figure GDA0003166505990000025
Figure GDA0003166505990000026
S13, solving the direction vector of the aircraft relative to the sun in the centroid inertia system
Figure GDA0003166505990000027
Figure GDA0003166505990000028
Wherein, CJ2000←bIs a rotating matrix of the J2000 inertia system of the aircraft, which is obtained by calculating a quaternion of the body relative to the inertia system measured by a star sensor,
Figure GDA0003166505990000029
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 GDA00031665059900000210
wherein
Figure GDA00031665059900000211
Is the direction vector of the earth relative to the sun in the J2000 inertial system,
Figure GDA00031665059900000212
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 GDA0003166505990000031
wherein the content of the first and second substances,
Figure GDA0003166505990000032
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 GDA0003166505990000033
wherein the content of the first and second substances,
Figure GDA0003166505990000034
is the vector of the aircraft relative to the sun in the direction of the sun-center inertial system,
Figure GDA0003166505990000035
the direction vector of the earth relative to the sun in the centroid inertial system is shown.
Step S5 specifically includes:
s51, calculating the position error of the aircraft relative to the sun in the centroid inertial system
Figure GDA0003166505990000036
Wherein
Figure GDA0003166505990000037
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 GDA0003166505990000038
Figure GDA0003166505990000039
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 GDA00031665059900000310
Speed error correction dv of kth step calculated by inertial navigationi,kTo obtain
Figure GDA00031665059900000311
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-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.
Step S61 specifically includes:
s611, let the position vector of the aircraft in the J2000 inertial system 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 GDA0003166505990000041
Figure GDA0003166505990000042
Figure GDA0003166505990000043
wherein the content of the first and second substances,
Figure GDA0003166505990000044
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,i agy,i agz,i]TCalculating the acceleration a of the aircraft under the J2000 inertial systemiWherein 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 GDA0003166505990000061
Pitch angle thetasCalculating the rolling angle of the sun vector of the aircraft on the aircraft
Figure GDA0003166505990000062
A pitch angle theta; wherein psi is the installation angle of the optical sensor in the horizontal plane;
Figure GDA0003166505990000063
s12, rolling angle of the aircraft according to the sun vector of the aircraft
Figure GDA0003166505990000064
A pitch angle theta, calculating the direction vector of the aircraft relative to the sun in the system
Figure GDA0003166505990000065
Figure GDA0003166505990000066
S13, solving the direction vector of the aircraft relative to the sun in the centroid inertia system
Figure GDA0003166505990000067
Figure GDA0003166505990000068
Wherein: cJ2000←bThe rotation matrix of the aircraft to the J2000 inertia system is obtained by calculating a 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 GDA0003166505990000069
wherein
Figure GDA00031665059900000610
Is the direction vector of the earth relative to the sun in the J2000 inertial system,
Figure GDA00031665059900000611
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 GDA00031665059900000612
wherein the content of the first and second substances,
Figure GDA00031665059900000613
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 GDA00031665059900000614
wherein the content of the first and second substances,
Figure GDA00031665059900000615
is the vector of the aircraft relative to the sun in the direction of the sun-center inertial system,
Figure GDA00031665059900000616
the direction vector of the earth relative to the sun in the centroid inertial system is shown.
In the context of figure 2, it is shown,
Figure GDA0003166505990000071
the direction vector of the earth relative to the sun in the centroid inertial system;
Figure GDA0003166505990000072
aircraft is used to sun in the daytimeA linear direction vector;
Figure GDA0003166505990000073
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 position error of the aircraft relative to the sun in the centroid inertial system
Figure GDA0003166505990000074
Wherein
Figure GDA0003166505990000075
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 GDA0003166505990000076
Figure GDA0003166505990000077
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, and i represents three axes of x, y and z;
let dri,k＝[dxi,k dyi,k dzi,k]T、
Figure GDA0003166505990000078
The triaxial independent estimation position error correction is in the form of:
Figure GDA0003166505990000079
Figure GDA00031665059900000710
Figure GDA00031665059900000711
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 GDA00031665059900000712
Speed error correction dv of kth step calculated by inertial navigationi,kTo obtain
Figure GDA0003166505990000081
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,k dvyi,k dvzi,k]T，
Figure GDA0003166505990000082
The triaxial independent estimation speed error correction is in the form as follows:
Figure GDA0003166505990000083
Figure GDA0003166505990000084
Figure GDA0003166505990000085
wherein: k is a radical ofp,dx、kp,dy、kp,dzA scaling factor is estimated for the three axis velocity error correction.
And S6, carrying out inertia-astronomical combined navigation based on the position/speed error correction.
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 in the J2000 inertial system 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 GDA0003166505990000086
Figure GDA0003166505990000087
Figure GDA0003166505990000088
wherein the content of the first and second substances,
Figure GDA0003166505990000089
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,i agy,i agz,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 (5)

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;
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 includes the steps of:
s11, measuring the lower roll angle of the sensor system according to the optical sensor
Figure FDA0003166505980000011
Pitch angle thetasCalculating the rolling angle of the sun vector of the aircraft on the aircraft
Figure FDA0003166505980000012
A pitch angle theta; wherein psi is the installation angle of the optical sensor in the horizontal plane;
Figure FDA0003166505980000013
s12, rolling angle of the aircraft according to the sun vector of the aircraft
Figure FDA0003166505980000014
A pitch angle theta, calculating the direction vector of the aircraft relative to the sun in the system
Figure FDA0003166505980000015
Figure FDA0003166505980000016
S13, solving the direction vector of the aircraft relative to the sun in the centroid inertia system
Figure FDA0003166505980000017
Figure FDA0003166505980000018
Wherein, CJ2000←bIs a rotating matrix of the J2000 inertia system of the aircraft, which is obtained by calculating a quaternion of the body relative to the inertia system measured by a star sensor,
Figure FDA0003166505980000019
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;
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 FDA00031665059800000110
wherein the content of the first and second substances,
Figure FDA00031665059800000111
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;
s4, calculating the direction vector of the sun relative to the aircraft in the centroid inertial system;
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 FDA0003166505980000021
wherein the content of the first and second substances,
Figure FDA0003166505980000022
is the vector of the aircraft relative to the sun in the direction of the sun-center inertial system,
Figure FDA0003166505980000023
is the direction vector of the sun relative to the sun in the centroid inertial system
S5, estimating a position/speed error correction quantity by PI filtering based on the direction vector error;
and S6, carrying out inertia-astronomical combined navigation based on the position/speed error correction.
2. 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 FDA0003166505980000024
wherein
Figure FDA0003166505980000025
Is the direction vector of the earth relative to the sun in the J2000 inertial system,
Figure FDA0003166505980000026
is the direction vector of the sun vector in the J2000 inertial system.
3. The combined inertial-astronomical navigation method for deep space exploration vehicles according to claim 1, wherein step S5 specifically comprises:
s51, calculating the position error of the aircraft relative to the sun in the centroid inertial system
Figure FDA0003166505980000027
Wherein
Figure FDA0003166505980000028
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 FDA0003166505980000029
Figure FDA00031665059800000210
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 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 FDA00031665059800000211
Speed error correction dv of kth step calculated by inertial navigationi,kTo obtain
Figure FDA00031665059800000212
Wherein k isp,vEstimating a proportional system for a speed error correctionNumber, kp,vThe speed error correction can be estimated independently for a 3 x 3 diagonal matrix.
4. The combined inertial-astronomical navigation method for deep space exploration vehicles according to claim 1, 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.
5. The combined inertial-astronomical navigation method for deep space exploration vehicles according to claim 4, wherein step S61 specifically comprises:
s611, let the position vector of the aircraft in the J2000 inertial system 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 FDA0003166505980000031
Figure FDA0003166505980000032
Figure FDA0003166505980000041
wherein the content of the first and second substances,
Figure FDA0003166505980000042
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,i agy,i agz,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。
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