CN102494686B - A kind of satellite attitude orbit determines system and method - Google Patents

Abstract

The invention provides a kind of satellite attitude orbit and determine system and method, the attitude track that on star, subsystem comprises determines that module obtains pseudorange from the satellite navigation signals from attitude track acquisition module, pseudorange rates and carrier phase data, utilize pseudorange, pseudorange rates and carrier phase data, generate relative position information between star, relative velocity and three dimensional local information between star, three dimensional local information and the satellite starlight elevation angle are carried out Federated Filtering, generate positional information and the velocity information of satellite relative inertness system of satellite relative inertness system, utilize satellite navigation signals, attitude of satellite angle information and the satellite starlight elevation angle, carry out combining determining appearance, generate three shaft angle degree, and three shaft angle degree are carried out optimal estimation, obtain three-axis attitude angle and the three-axis attitude angular velocity of satellite；According to attitude track, Attitude and orbit control computer determines that the information that module exports carries out becoming rail and posture adjustment controls.Use the system and method for the present invention, it is possible to increase determine the precision of appearance and real-time that track determines.

Description

Satellite attitude orbit determination system and method

Technical Field

The invention relates to the field of aircraft design, in particular to a satellite attitude orbit determination system and a satellite attitude orbit determination method.

Background

Fig. 1 is a schematic structural diagram of a conventional satellite attitude and orbit determination system. Referring now to fig. 1, a conventional satellite attitude and orbit determination system is described, which includes the following details:

the existing satellite attitude orbit determination system comprises an onboard subsystem 10 and a ground subsystem 11. The satellite-borne subsystem 10 is used for acquiring data for determining satellite orbits and satellite attitudes, finishing satellite attitude determination according to the data for determining the satellite attitudes, and transmitting the data for determining the satellite orbits to the ground subsystem 11 in a downlink manner; the ground subsystem 11 performs orbit determination operations according to the existing flight dynamics model and the data for determining the satellite orbit.

The satellite subsystem 10 comprises an inertial device 101, an Attitude and Orbit Control Computer (AOCC)102, a GPS receiver 103, a USB transponder 104 and a laser ranging system 105; the ground subsystem 11 includes a ground ranging station 111 and a first computer 112. The inertial device 101 determines an equation according to the existing attitude of inertial navigation, and determines the attitude of the satellite where the inertial device is located; the attitude and orbit control computer 102 acquires data of the inertial device 101, acquires orbit determination data according to the established original kinetic equation and the acquired data, and sends the orbit determination data to the USB responder 104 according to a preset period; a Global Positioning System (GPS) receiver 103 sends acquired GPS position data to a USB transponder 104 according to a preset period; when the satellite where the satellite-borne subsystem 10 is located passes through the measurement and control range of the ground subsystem 11, the USB transponder 104 downloads orbit determination data and GPS position data to the ground ranging station 111 of the ground subsystem 11 through a wireless radio frequency channel; the laser ranging system 105 establishes a communication link with the ground ranging station 111 so that the ground ranging station 111 obtains a laser ranging value; the ground ranging station 111 sends orbit determination data, GPS position data and laser ranging values to the first computer 112; the first computer 112 performs orbit determination based on the flight dynamics model and the received data. The orbit determination performed by the first computer 112 is a mathematical calculation process of continuous fitting iteration, and the longer the operation period is, the higher the accuracy of the obtained orbit determination result is.

In summary, in the existing satellite attitude orbit determination system, the orbit determination result needs to depend on the distance between the satellite and the ground, and the orbit determination operation process is a complex mathematical operation process completed by the ground subsystem, so that a long operation period needs to be consumed in order to improve the orbit determination accuracy; the precision of posture determination by using the inertial device is not high, and needs to be further improved; once the ground subsystem participates in the satellite orbit determination, the real-time performance of orbit determination cannot be guaranteed.

Disclosure of Invention

In view of the above, the present invention provides a satellite attitude orbit determination system, which can improve the accuracy of attitude determination and realize the real-time performance of orbit determination.

The invention aims to provide a satellite attitude orbit determination method which can improve the accuracy of attitude determination and realize the real-time performance of orbit determination.

In order to achieve the purpose, the technical scheme of the invention is realized as follows:

a satellite attitude orbit determination system, comprising an on-board subsystem, said on-board subsystem comprising:

the attitude orbit signal acquisition module acquires satellite navigation signals, satellite attitude angle signals and satellite starlight elevation angles of various navigation satellite constellations and outputs the satellite navigation signals, the satellite attitude angle signals and the satellite starlight elevation angles to the attitude orbit determination module;

the attitude orbit determination module is used for obtaining pseudo-range, pseudo-range rate and carrier phase data from satellite navigation signals, carrying out error correction and differential operation on the pseudo-range, pseudo-range rate and carrier phase data, obtaining inter-satellite relative position information and inter-satellite relative speed information, and outputting the inter-satellite relative position information and the inter-satellite relative speed information to an attitude orbit control computer;

the attitude orbit determination module generates three-dimensional position information according to pseudo range, pseudo range rate and carrier phase data by using a least square optimal estimation algorithm and a space-time unified algorithm, performs a federal filtering algorithm on the three-dimensional position information and satellite starlight elevation angle by using a flight dynamics model and a federal filtering algorithm, generates position information of a satellite relative to an inertial system and speed information of the satellite relative to the inertial system, and outputs the position information and the speed information to an attitude and orbit control computer;

the attitude orbit determination module performs combined attitude determination by using a satellite navigation signal, satellite attitude angle information and a satellite starlight elevation angle to generate a three-axis angle, performs optimal estimation on the three-axis angle by using a federal filter algorithm to obtain the three-axis attitude angle of the satellite and the three-axis attitude angular velocity of the satellite, and outputs the three-axis attitude angle and the three-axis attitude angular velocity to an attitude orbit control computer;

the attitude and orbit control computer outputs an orbit change instruction to the execution module according to the received inter-satellite relative position information, inter-satellite relative speed information, position information of the satellite relative inertia system and speed information of the satellite relative inertia system, and outputs an attitude adjustment instruction to the execution module according to a three-axis attitude angle of the satellite and a three-axis attitude angular speed of the satellite;

and the execution module adjusts the orbit of the satellite where the system is located according to the orbital transfer instruction and adjusts the attitude of the satellite where the system is located according to the attitude adjusting instruction.

Preferably, the satellite subsystem further comprises:

the inter-satellite communication module is used for establishing a communication link between the satellite and the satellite for cooperation, sending inter-satellite relative position information and inter-satellite relative speed information, receiving position information and speed information of another satellite and forwarding the position information and the speed information to the attitude orbit determination module;

the attitude orbit determination module further calculates and obtains inter-satellite relative position information and inter-satellite relative speed information according to the position information and the speed information of the other satellite, and outputs the inter-satellite relative position information and the inter-satellite relative speed information to the inter-satellite communication module.

Preferably, the system further comprises:

the ground subsystem generates and outputs a satellite in-orbit flight picture by utilizing the satellite attitude information and the satellite orbit information received through the wireless communication link;

the attitude and orbit control computer of the satellite subsystem further generates and outputs attitude information and satellite orbit information of the satellite according to the orbital transfer instruction and the attitude adjusting instruction;

the satellite subsystem further comprises:

the satellite platform bus is used for establishing a data transmission channel between the attitude and orbit control computer and the remote measuring module;

and the telemetry module receives the satellite attitude information and the satellite orbit information from the attitude and orbit control computer through the platform bus and outputs a satellite attitude signal and a satellite orbit signal through a wireless link.

In the above system, the attitude trajectory signal acquisition module includes:

the multi-mode satellite navigation GNSS receiver is used for collecting satellite navigation information of various navigation satellite constellations and outputting the satellite navigation information to the attitude orbit determination module;

the inertial device is used for acquiring satellite attitude angle signals of the satellite and outputting the satellite attitude angle signals to the attitude orbit determination module;

the star sensor collects the satellite starlight elevation angle and the quaternion attitude angle of the satellite and outputs the satellite starlight elevation angle and the quaternion attitude angle to the attitude orbit determination module;

the earth sensor is used for acquiring satellite attitude angle information of a satellite and outputting the satellite attitude angle information to the attitude orbit determination module;

the plurality of navigation satellite constellations at least comprise a Beidou satellite and a Global Positioning System (GPS) satellite.

In the above system, the attitude trajectory determination module includes:

the data generation unit is used for generating pseudo-range, pseudo-range rate and carrier phase data of various navigation satellite constellations according to the satellite navigation information of the various navigation satellite constellations output by the GNSS receiver and outputting the pseudo-range, pseudo-range rate and carrier phase data to the correction unit and the space-time unified unit;

the correction unit is used for performing ephemeris error correction, ionosphere delay correction, integer ambiguity solution and cycle slip detection on the received pseudo ranges and carrier phase data of various navigation satellite constellations and outputting the corrected pseudo ranges and carrier phase data of the various navigation satellite constellations to the differential operation unit;

the differential operation unit is used for carrying out differential operation on the pseudo ranges and carrier phase data of the corrected various navigation satellite constellations by utilizing the model of the relative orbit element to obtain inter-satellite relative position information and inter-satellite relative speed information, and outputting the inter-satellite relative position information and the inter-satellite relative speed information to the attitude and orbit control computer;

the time-space unification unit is used for completing unification of a coordinate system by using a seven-parameter Boolean conversion formula according to pseudo-range, pseudo-range rate and carrier phase data of various navigation satellite constellations, unifying the time of the pseudo-range, the pseudo-range rate and the carrier phase data of the various navigation satellite constellations to coordinated Universal Time Coordinated (UTC), performing successive iteration by using a combined positioning equation and a least square method, calculating to obtain three-dimensional position information, and outputting the three-dimensional position information to the first federal filtering unit;

the model calculation unit is used for generating a state equation of the satellite orbit dynamics model and outputting the state equation to the first federal filtering unit;

the first federal filtering unit is used for solving a state equation by utilizing three-dimensional position information, satellite attitude angle information from a star sensor, a state equation of a satellite orbit dynamics model, a state equation of GNSS distance measurement and a state equation of star sensor angle observation to obtain position information of a satellite relative to an inertial system and speed information of the satellite relative to the inertial system, filtering the position information of the satellite relative to the inertial system and the speed information of the satellite relative to the inertial system by utilizing a linear minimum variance algorithm and a prediction algorithm, and outputting the filtered position information and the speed information to an attitude and orbit control computer;

the velocity integration unit is used for carrying out velocity integration on the satellite attitude angle signals output by the inertial device to obtain first triaxial angle information and outputting the first triaxial angle information to the second joint filtering unit, the first joint attitude determination unit and the second joint attitude determination unit;

the first combined attitude determination unit corrects errors of the first triaxial angle information output by the rate integration unit by using the satellite starlight elevation angle and the quaternion attitude angle output by the star sensor to obtain second triaxial angle information and outputs the second triaxial angle information to the second combined filter;

the second joint attitude determination unit corrects the error of the first triaxial angle information output by the rate integration unit by using the satellite attitude angle information output by the earth sensor to obtain third triaxial angle information and outputs the third triaxial angle information to a second joint filter;

and the second federated filtering unit is used for adjusting the first gain, the second gain and the third gain by using a main filtering equation of a federated filtering algorithm, filtering the first triaxial angle information by using the first gain, filtering the second triaxial angle information by using the second gain, filtering the third triaxial angle information by using the third gain, performing wild value elimination on the three filtered information, performing optimal estimation operation on the eliminated information, obtaining a satellite triaxial attitude angle and a satellite triaxial attitude angular velocity, and outputting the satellite attitude angular velocity to an attitude and orbit control computer.

Preferably, the attitude trajectory determination module further comprises:

the first prediction unit predicts the second triaxial angle information output by the first combined attitude determination unit by using a preset prediction algorithm and outputs the predicted second triaxial angle information to the first combined attitude determination unit so that the first combined attitude determination unit corrects the calculated second triaxial angle information;

and the second prediction unit predicts the third triaxial angle information output by the second combined attitude determination unit by using a preset prediction algorithm and outputs the predicted third triaxial angle information to the second combined attitude determination unit so that the second combined attitude determination unit corrects the calculated third triaxial angle information.

In the above system, the second joint filtering unit includes:

each subtracter performs subtraction calculation on the received first triaxial angle information, second triaxial angle information or third triaxial angle information and the satellite triaxial attitude angle output by the optimal estimation subunit, and outputs the calculated difference to the first gain subunit, the second gain subunit or the third gain subunit;

the first gain subunit is used for filtering the difference value between the first triaxial angle information and the satellite triaxial attitude angle from the optimal estimation subunit according to the first gain and outputting the filtered first triaxial angle information to the first time synchronization subunit;

the second gain subunit is used for filtering the difference value between the second triaxial angle information and the satellite triaxial attitude angle from the optimal estimation subunit according to second gain and outputting the filtered second triaxial angle information to the first time synchronization subunit;

the third gain subunit is used for filtering the difference value between the third triaxial angle information and the satellite triaxial attitude angle from the optimal estimation subunit according to third gain and outputting the filtered third triaxial angle information to the first time synchronization subunit;

the first time synchronization subunit outputs the received filtered first triaxial angle information, second triaxial angle information and third triaxial angle information to the preprocessing subunit, calculates a first error, a second error and a third error according to a main filtering equation of the federal filter, sets a first gain of the first gain subunit by using the first error, sets a second gain of the second gain subunit by using the second error, and sets a third gain of the third gain subunit by using the third error;

the preprocessing subunit is used for performing wild value elimination on the first triaxial angle information, the second triaxial angle information and the third triaxial angle information which are received and filtered according to a preset threshold value, and outputting the eliminated information to the optimal estimation subunit;

and the optimal estimation subunit calculates the eliminated information according to an optimal estimation algorithm to obtain a satellite three-axis attitude angle and a satellite three-axis attitude angular velocity, and outputs the satellite three-axis attitude angular velocity to an attitude and orbit control computer.

In the above system, the first federal filter unit includes:

the information distribution subunit distributes weights to the first sub-filter and the second sub-filter according to an information conservation principle and an error evaluation result output by the main filter;

the first sub-filter is used for solving the state equation according to the distributed weight value, the state equation of the satellite orbit dynamic model, the state equation of the GNSS distance measurement, the three-dimensional position information and the pseudo range to obtain a first result, correcting the first result by using the predicted value from the main filter, and outputting the corrected first result to the main filter;

the second sub-filter is used for solving the state equation according to the distributed weight, the state equation of the satellite orbit dynamics model and the state equation observed by the star sensor angle to obtain a second result, correcting the second result by using the predicted value from the main filter, and outputting the corrected second result to the main filter;

and the main filter obtains a predicted value according to the three-dimensional position information, the corrected first result and the corrected second result by using a preset prediction algorithm and outputs the predicted value to the first sub-filter and the second sub-filter, carries out error evaluation on the corrected first result and the corrected second result, outputs an error evaluation result to the information distribution sub-unit, calculates and obtains position information of the satellite relative to an inertial system and speed information of the satellite relative to the inertial system by using a linear minimum variance algorithm, and outputs the position information and the speed information to the attitude and orbit control computer.

In the above system, the ground subsystem comprises:

the receiving module is used for establishing a wireless communication link with the satellite-borne subsystem and outputting the received satellite attitude information and the satellite orbit information to the data acquisition module;

the data acquisition module analyzes the satellite attitude information and the satellite orbit information to obtain inter-satellite relative position information, inter-satellite relative speed information, position information of the satellite relative inertial system, speed information of the satellite relative inertial system, a three-axis attitude angle of the satellite and a three-axis attitude angular speed of the satellite, and outputs the information to the display driving module;

the display driving module generates and outputs a satellite in-orbit flight picture by using the picture generation tool and the received data and information;

and the time synchronization module outputs the timestamp to the data acquisition module and the display driving module, and synchronizes the data received by the data acquisition module and the display driving module.

A method of satellite attitude orbit determination, the method comprising:

A. the satellite subsystem collects satellite navigation signals, satellite attitude angle signals and satellite starlight elevation angles of various navigation satellite constellations;

B. the satellite subsystem obtains pseudo range, pseudo range rate and carrier phase data from satellite navigation signals, and error correction and differential operation are carried out on the pseudo range, the pseudo range rate and the carrier phase data to obtain inter-satellite relative position information and inter-satellite relative speed information;

the satellite subsystem generates three-dimensional position information according to pseudo range, pseudo range rate and carrier phase data by using a least square optimal estimation algorithm and a space-time unified algorithm, and performs a federal filtering algorithm on the three-dimensional position information and satellite starlight elevation angle by using a flight dynamics model and a federal filtering algorithm to generate position information of a satellite relative to an inertial system and speed information of the satellite relative to the inertial system;

the satellite-borne subsystem performs combined attitude determination by using a satellite navigation signal, satellite attitude angle information and a satellite starlight elevation angle to generate a three-axis angle, and performs optimal estimation on the three-axis angle by using a federal filter algorithm to obtain the three-axis attitude angle of the satellite and the three-axis attitude angular velocity of the satellite;

C. the satellite-borne subsystem adjusts the orbit of the sub-satellite and the attitude of the satellite by utilizing the inter-satellite relative position information, the inter-satellite relative velocity information, the position information of the satellite relative to the inertial system, the velocity information of the satellite relative to the inertial system, the three-axis attitude angle of the satellite and the three-axis attitude angular velocity of the satellite.

In the above method, the performing error correction and difference operation on the pseudo-range, the pseudo-range rate, and the carrier phase data in step B includes:

b1, performing ephemeris error correction, ionosphere delay correction, integer ambiguity resolution and cycle slip detection on pseudo-ranges and carrier phase data of various navigation satellite constellations to obtain the corrected pseudo-ranges and carrier phase data of the various navigation satellite constellations;

and B2, carrying out differential operation on the pseudo ranges and the carrier phase data of the corrected multiple navigation satellite constellations by using the model of the relative orbit element to obtain the inter-satellite relative position information and the inter-satellite relative speed information.

In the method, the step B of generating three-dimensional position information according to the pseudo-range, the pseudo-range rate and the carrier phase data by using a least square optimal estimation algorithm and a space-time unified algorithm includes:

b3, according to pseudo-range, pseudo-range rate and carrier phase data of various navigation satellite constellations, unifying a coordinate system by using a seven-parameter Boolean Sasa conversion formula;

b4, unifying pseudo ranges, pseudo range rates and time of carrier phase data of various navigation satellite constellations to coordinated universal time UTC;

and B5, calculating to obtain three-dimensional position information by utilizing the combination positioning equation and the successive iteration of the least square method.

In the above method, the joint attitude determination using the satellite navigation signal, the satellite attitude angle information, and the satellite starlight elevation angle in step B, and generating the three-axis angle includes:

b6, performing rate integration on the satellite attitude angle signal output by the inertial device to obtain first triaxial angle information;

b7, correcting the error of the first triaxial angle information by using the satellite starlight elevation angle and the quaternion attitude angle output by the star sensor to obtain second triaxial angle information;

and B8, correcting the error of the first triaxial angle information by using the satellite attitude angle information output by the earth sensor to obtain third triaxial angle information.

According to the technical scheme, the satellite attitude orbit determination system and the satellite attitude orbit determination method are not dependent on a GNSS receiver for positioning, but introduce information collected by an inertia device, a star sensor and an earth sensor to assist the GNSS receiver to realize combined autonomous navigation, change a satellite attitude determination and orbit determination mode taking the ground as a main mode, and enable a satellite to autonomously realize real-time high-precision and high-reliability orbit determination and attitude determination under the condition of being separated from the ground station.

Drawings

Fig. 1 is a schematic structural diagram of a conventional satellite attitude and orbit determination system.

Fig. 2 is a schematic structural diagram of a satellite attitude orbit determination system according to the present invention.

Fig. 3 is a schematic structural diagram of the attitude trajectory determination module according to the present invention.

FIG. 4 is a flow chart of a method of attitude trajectory determination in accordance with the present invention.

Detailed Description

In order to make the objects, technical solutions and advantages of the present invention more apparent, the present invention is further described in detail below with reference to the accompanying drawings and examples.

Fig. 2 is a schematic structural diagram of a satellite attitude orbit determination system according to the present invention. Referring to fig. 2, a satellite attitude orbit determination system of the present invention is described, specifically as follows:

the orbit determination and attitude determination of the satellite attitude orbit determination system are realized on the satellite subsystem. The satellite subsystem comprises: a posture orbit signal acquisition module 20, a posture orbit determination module 21, a posture orbit control computer 22 and an execution module 23.

The attitude orbit signal collection module 20 collects satellite navigation signals, satellite attitude angle signals and satellite starlight elevation angles of a plurality of navigation satellite constellations and outputs the signals to the attitude orbit determination module 21. The various navigation satellite constellations mentioned in the present invention are at least Beidou (COMPASS) satellites and GPS satellites, and the satellite navigation signals may also be navigation signals of various systems such as GALILEO, GLONASS, etc., which are not described herein again. The pose trajectory signal acquisition module 20 may be connected to the pose trajectory determination module 21 through a 422 interface (not shown in fig. 2). The satellite navigation signals acquired by the attitude orbit signal acquisition module 20 at least include a pseudo range, a pseudo range rate and a carrier phase.

The attitude orbit determination module 21 obtains pseudorange, pseudorange rate and carrier phase data from the satellite navigation signal, performs error correction and differential operation on the pseudorange, the pseudorange rate and the carrier phase data, obtains inter-satellite relative position information and inter-satellite relative speed information, and outputs the inter-satellite relative position information and the inter-satellite relative speed information to the attitude and orbit control computer 22.

The attitude orbit determination module 21 further generates three-dimensional position information according to pseudo-range, pseudo-range rate and carrier phase data by using a least square optimal estimation algorithm and a space-time unified algorithm, performs a federal filtering algorithm on the three-dimensional position information and satellite starlight elevation angles from the attitude orbit signal acquisition module 20 by using a flight dynamics model and a federal filtering algorithm, generates position information of a satellite relative to an inertial system and speed information of the satellite relative to the inertial system, and outputs the position information and the speed information to the attitude orbit control computer 22.

The attitude orbit determination module 21 performs joint attitude determination by using the satellite navigation signal, the satellite attitude angle information and the satellite starlight elevation angle to generate a three-axis angle, performs optimal estimation on the three-axis angle by using the federal filter algorithm to obtain the three-axis attitude angle of the satellite and the three-axis attitude angular velocity of the satellite, and outputs the three-axis attitude angle and the three-axis attitude angular velocity to the attitude orbit control computer 22.

The attitude and orbit control computer 22 outputs an orbit change instruction to the execution module according to the received inter-satellite relative position information, inter-satellite relative velocity information, position information of the satellite relative inertia system and velocity information of the satellite relative inertia system, and outputs an attitude adjustment instruction to the execution module 23 according to a three-axis attitude angle of the satellite and a three-axis attitude angular velocity of the satellite.

The execution module 23 adjusts the orbit of the satellite where the system is located according to the orbital transfer instruction, and adjusts the attitude of the satellite where the system is located according to the attitude adjustment instruction.

In order to realize high-precision relative positioning between satellites, the on-board subsystem of the invention further comprises: an inter-satellite communication module 24; the inter-satellite communication module 24 is configured to establish a communication link with a cooperating satellite, and send and receive inter-satellite relative position information and inter-satellite relative velocity information, and specifically, the inter-satellite communication module 24 acquires a position of another satellite through the communication link; the attitude orbit determination module 22 further performs a differential operation on the position of itself and the position of another satellite to solve the relative position between the satellites, and may further output the inter-satellite relative position information and the inter-satellite relative velocity information obtained by the calculation to the inter-satellite communication module 24.

In order to more intuitively display the motion condition of the satellite in orbit flight, the satellite attitude and orbit determination system further comprises a ground subsystem. The ground subsystem generates and outputs a satellite in-orbit flight picture by utilizing satellite attitude information and satellite orbit information received by the wireless communication link; an attitude and orbit control computer 22 of the satellite subsystem generates attitude information and satellite orbit information of the satellite according to the orbit changing instruction and the attitude adjusting instruction, and outputs the attitude information and the satellite orbit information to a telemetry module 25 through a satellite platform bus; telemetry module 25 receives satellite attitude information and satellite orbit information from attitude and orbit control computer 22 and outputs satellite attitude signals and satellite orbit signals to the ground subsystem via a wireless link.

The information required by the ground subsystem to generate the satellite in-orbit flight picture at least comprises the following information: the current satellite time, the three-axis coordinate values and velocity values based on the inertial system, the three-axis attitude angles and angular velocities based on the orbital coordinate system, the satellite flight altitude, the longitude, the latitude and the like can be obtained from the attitude and orbit control computer 22 through the telemetry module 25, and a specific obtaining method of each piece of information is not repeated herein. After receiving the data of the attitude trajectory determination module 21, the attitude and orbit control computer 22 is used as an input for attitude trajectory control, and on the other hand, periodically packages the data and downloads the data to the ground subsystem through the telemetry module 25.

The satellite platform bus of the invention can adopt a 1553B platform bus.

In order to achieve the optimal positioning effect, the attitude orbit signal acquisition module 20 of the invention is compatible with receivers of various navigation satellite constellation signals, thereby increasing the number of navigation satellites, reducing the precision dilution factor and improving the redundancy and reliability of the system. Wherein, the attitude orbit signal acquisition module 20 includes: a Global Navigation Satellite System (GNSS) receiver 201, a star sensor 202, an earth sensor 203, and an inertial device 204.

The GNSS receiver 201 collects satellite navigation information of the beidou satellite and the GPS satellite, and outputs the satellite navigation information to the attitude orbit determination module 21. The satellite navigation signal of the Beidou satellite is a COMPASS system, and the satellite navigation signal of the GPS satellite is a GPS system. The signal directly acquired by the GNSS receiver 201 included in the attitude orbit signal acquisition module 20 is a navigation message, which includes a lot of information, such as ephemeris data, time information, pseudo-range rate, and carrier phase of each navigation satellite constellation in the navigation satellite constellation, which are encrypted, and after acquiring a frame of navigation message, the GNSS receiver 201 needs to decrypt and output the frame of navigation message to the attitude orbit determination module 21.

The star sensor 202 collects the satellite starlight elevation angle and the quaternion attitude angle of the satellite and outputs the satellite starlight elevation angle and the quaternion attitude angle to the attitude orbit determination module 21. The earth sensor 203 collects satellite attitude angle information of the satellite and outputs the information to the attitude orbit determination module 22. The inertial device 204 collects satellite attitude angle signals of the satellites and outputs the satellite attitude angle signals to the attitude orbit determination module 22. The star sensor 202, the earth sensor 203 and the inertial device 204 may be conventional components, and the structures of the components are not described herein again.

Fig. 3 is a schematic structural diagram of the attitude trajectory determination module of the present invention. Now, referring to fig. 3, the attitude trajectory determination module of the present invention is described as follows:

the attitude trajectory determination module 21 of the present invention includes: the system comprises a data generation unit 210, a correction unit 211, a difference operation unit 212, a space-time unification unit 213, a model calculation unit 214, a first federated filtering unit 215, a rate integration unit 216, a first joint pose determination unit 217, a second joint pose determination unit 218 and a second joint filter 219.

The data generating unit 210 generates pseudo-range, pseudo-range rate, and carrier phase data of a plurality of navigation satellite constellations from the satellite navigation information of the plurality of navigation satellite constellations output from the GNSS receiver 201, and outputs the pseudo-range, pseudo-range rate, and carrier phase data to the correcting unit 211 and the space-time unifying unit 213. Specifically, the data generating unit 210 extracts the pseudo range, the pseudo range rate, and the carrier phase in the decrypted electronic text output by the GNSS receiver 201 according to protocols supported by a plurality of navigation satellite constellations, for example, a protocol supported by a GPS and a protocol supported by a beidou satellite.

The correcting unit 211 performs ephemeris error correction, ionosphere delay correction, integer ambiguity resolution, and cycle slip detection on the received pseudo-ranges and carrier phase data of multiple navigation satellite constellations, and outputs the corrected pseudo-ranges and carrier phase data of the multiple navigation satellite constellations to the differential operation unit 212.

The differential operation unit 212 performs differential operation on the pseudo ranges and carrier phase data of the plurality of corrected navigation satellite constellations by using the model of the relative orbit element to obtain inter-satellite relative position information and inter-satellite relative velocity information, and outputs the inter-satellite relative position information and the inter-satellite relative velocity information to the attitude and orbit control computer 22.

The space-time unification unit 213 completes unification of a coordinate system by using a seven-parameter boolean conversion formula according to pseudo-ranges, pseudo-range rates and carrier phase data of a plurality of navigation satellite constellations, unifies the time of the pseudo-ranges, the pseudo-range rates and the carrier phase data of the plurality of navigation satellite constellations to coordinated Universal Time Coordinated (UTC), performs successive iteration by using a combined positioning equation and a least square method, calculates and obtains three-dimensional position information, and outputs the three-dimensional position information to the first federal filtering unit 215.

The model calculation unit 214 generates a state equation of the satellite orbit dynamics model and outputs the state equation to the first federal filtering unit 215. Specifically, the model calculating unit 214 generates the state equation of the satellite orbit dynamics model by using a sequential processing method based on the extended kalman filter, for example, the sequential processing method based on the extended kalman filter includes: selecting M position values (M <10) at the previous M moments as initial values, and defining a covariance matrix P0; performing linear minimum variance processing on the previous M values to obtain an estimated value of the position at the current moment; according to the step of popularizing Kalman filtering, a state transition matrix and a new covariance matrix P1 are obtained according to P0 and the value of the previous M time; by analogy, the current state transition matrix can obtain the position estimation value of the next moment by utilizing the previous M-1 values and the current estimation value; from the covariance matrix P1 and the position estimate at the next time instant, a new covariance matrix P2 and a new state transition matrix may be obtained. To improve accuracy, the covariance may be gradually brought to zero.

The first federal filter unit 215 solves the state equation by using the three-dimensional position information, the satellite attitude angle information from the star sensor 202, the state equation of the satellite orbit dynamics model, the state equation of GNSS ranging, and the state equation of star sensor angle observation, obtains the position information of the satellite relative to the inertial system and the velocity information of the satellite relative to the inertial system, filters the position information of the satellite relative to the inertial system and the velocity information of the satellite relative to the inertial system by using a linear minimum variance algorithm and a prediction algorithm, and outputs the filtered information to the attitude and orbit control computer 22.

The rate integration unit 216 performs rate integration on the satellite attitude angle signal output by the inertial device to obtain first triaxial angle information, and outputs the first triaxial angle information to the second joint filtering unit 219, the first joint attitude determination unit 217, and the second joint attitude determination unit 218.

The first joint attitude determination unit 217 corrects the error of the first triaxial angle information output by the rate integration unit 216 by using the satellite starlight elevation angle and the quaternion attitude angle output by the star sensor 202, obtains the second triaxial angle information, and outputs the second triaxial angle information to the second joint filter 219.

The joint attitude determination is adopted, namely, the real-time property of the satellite attitude angle information output by the inertial device 204 and the low error of the angle information output by the star sensor 202 are mutually compensated, and the high error of the satellite attitude angle information acquired by the inertial device 204 is reduced under the condition of ensuring the real-time property, so that the problems of long period and poor real-time property of the star sensor 202 are solved, and the problem of large accumulated error of the inertial device 204 is solved.

The second joint attitude determination unit 218 corrects the error of the first triaxial angle information output by the rate integration unit 216 by using the satellite attitude angle information output by the earth sensor 203, obtains third triaxial angle information, and outputs the third triaxial angle information to the second joint filter 219.

The joint attitude determination is adopted, namely, the real-time property of the satellite attitude angle information output by the inertia device 204 and the low error of the angle information output by the earth sensor 203 are mutually compensated, and the high error of the satellite attitude angle information acquired by the inertia device 204 is reduced under the condition of ensuring the real-time property, so that the problems of long period and poor real-time property of the earth sensor 203 are solved, and the problem of large accumulated error of the inertia device 204 is solved.

The second federated filtering unit 219 adjusts the first gain, the second gain, and the third gain using a main filtering equation of the federated filtering algorithm, filters the first triaxial angle information using the first gain, filters the second triaxial angle information using the second gain, filters the third triaxial angle information using the third gain, performs outlier rejection on the three filtered information, performs optimal estimation operation on the rejected information, obtains the three-axis attitude angle of the satellite and the three-axis attitude angular velocity of the satellite, and outputs the three-axis attitude angular velocity to the attitude and orbit control computer 22.

To improve the accuracy of the pose determination, the pose trajectory determination module 21 further comprises: a first prediction unit 220 and a second prediction unit 221.

The first prediction unit 220 predicts the second triaxial angle information output by the first combined attitude determination unit 217 by using a preset prediction algorithm, and outputs the predicted second triaxial angle information to the first combined attitude determination unit 217, so that the first combined attitude determination unit 217 corrects the calculated second triaxial angle information. The correction using the predicted value is to determine whether the calculated value is within a threshold range of the predicted value, if so, output the calculated value to the second joint filter 219, otherwise, output no data.

The second prediction unit 221 predicts the third triaxial angle information output by the second coupled pose determining unit 218 by using a preset prediction algorithm, and outputs the predicted third triaxial angle information to the second coupled pose determining unit 218, so that the second coupled pose determining unit 218 corrects the calculated third triaxial angle information. The correction using the predicted value is to determine whether the calculated value is within a threshold range of the predicted value, if so, output the calculated value to the second joint filter 219, otherwise, output no data.

The second combined filter unit 219 of the present invention includes: three subtractors 2190, a first gain sub-unit 2191, a second gain sub-unit 2192, a third gain sub-unit 2193, a first temporal synchronization sub-unit 2194, a pre-processing sub-unit 2195 and an optimal estimation sub-unit 2196.

Each of the three subtractors 2190 subtracts the received first triaxial angle information, second triaxial angle information, or third triaxial angle information from the satellite triaxial attitude angle output by the optimal estimation subunit 2196, and outputs the calculated difference value to the first gain subunit 2191, the second gain subunit 2192, or the third gain subunit 2193.

The first gain subunit 2191 filters the difference between the first triaxial angle information and the satellite triaxial attitude angle from the optimal estimation subunit 2196 according to the first gain, and outputs the filtered first triaxial angle information to the first time synchronization subunit 2194.

The second gain subunit 2191 filters, according to the second gain, the difference between the second triaxial angle information and the satellite triaxial attitude angle from the optimal estimation subunit 2196, and outputs the filtered second triaxial angle information to the first time synchronization subunit 2194.

The third gain subunit 2193 filters, according to the third gain, the difference between the third triaxial angle information and the satellite triaxial attitude angle from the optimal estimation subunit 2196, and outputs the filtered third triaxial angle information to the first time synchronization subunit 2194.

The first time synchronization subunit 2194 outputs the received filtered first triaxial angle information, second triaxial angle information, and third triaxial angle information to the preprocessing subunit 2195, calculates a first error, a second error, and a third error according to a main filtering equation of the federal filter, sets a first gain of the first gain subunit 2191 using the first error, sets a second gain of the second gain subunit 2192 using the second error, and sets a third gain of the third gain subunit 2193 using the third error.

The preprocessing subunit 2195 performs outlier rejection on the first triaxial angle information, the second triaxial angle information, and the third triaxial angle information, which are received after filtering, according to a preset threshold, and outputs the rejected information to the optimal estimation subunit 2196.

The optimal estimation subunit 2196 calculates the eliminated information according to an optimal estimation algorithm to obtain a satellite three-axis attitude angle and a satellite three-axis attitude angular velocity, and outputs the three-axis attitude angular velocity to the attitude and orbit control computer 22. The optimal estimation subunit 2196 further obtains the satellite three-axis attitude angle and the satellite three-axis attitude angular velocity by calculation using the output result of the state equation of the satellite orbit dynamics model.

The first federal filter unit of the present invention comprises: the filter comprises an information distribution subunit, a first filter, a second filter and a main filter. The information distribution subunit is connected with the main filter, the first filter and the second filter; the first filter is connected with the space-time unified unit 213 and the model calculation unit 214; the second filter is connected with the model calculation unit 214 and the star sensor 202; the main filter is connected to the space-time unification unit 213, the first filter, the second filter, and the information distribution subunit.

And the information distribution subunit distributes the weight values to the first sub-filter and the second sub-filter according to the information conservation principle and the error evaluation result output by the main filter. Wherein, the weight value is a concept in the information fusion technology; the method comprises the following steps of obtaining a value with higher precision through the existing fusion calculation according to the similar information collected by a plurality of sources, such as attitude angles collected by a star sensor 202, an earth sensor 203 and an inertial device 204, and belonging to information fusion; in this case, the information of each source cannot be simply averaged, but the data of different sources are multiplied by coefficients with different sizes according to respective error levels, and then added, where the coefficients with different sizes are weights, and the weights are not fixed values. After the calculation is completed each time, the main filter carries out error evaluation on data of various sources, an error evaluation result is sent to the information distribution subunit, and the information distribution subunit distributes a large weight to a source with a small error and a small weight to a source with a large error according to the error evaluation result. The data of various sources is the modified first result and the modified second result received by the main filter.

The first sub-filter performs solution of the state equation according to the assigned weight, the state equation of the satellite orbit dynamics model from the model calculation unit 214, the state equation of the GNSS ranging from the model calculation unit 214, the three-dimensional position information from the space-time unification unit 213, and the pseudo-range to obtain a first result, corrects the first result by using the predicted value from the main filter, and outputs the corrected first result to the main filter. The model calculating unit 214 may establish a state equation of GNSS ranging according to the state equation of the satellite orbit dynamics model and the ranging information of the GNSS receiver.

The second sub-filter solves the state equation according to the distributed weight, the state equation of the satellite orbit dynamics model from the model calculation unit 214, and the state equation of the star sensor angle observation from the model calculation unit 214 to obtain a second result, corrects the second result by using the predicted value from the main filter, and outputs the corrected second result to the main filter. The model calculation unit 214 may establish a state equation of the star sensor angle observation according to the state equation of the satellite orbit dynamics model and the angle information of the star sensor 202.

The main filter obtains a predicted value according to the three-dimensional position information, the corrected first result and the corrected second result by using a preset prediction algorithm and outputs the predicted value to the first sub-filter and the second sub-filter, and calculates position information of the satellite relative to the inertial system and speed information of the satellite relative to the inertial system by using a linear minimum variance algorithm and outputs the position information and the speed information to the attitude and orbit control computer 22.

The ground subsystem of the present invention comprises: a receiving module 26, a data acquisition module 27, a display driving module 28 and a time synchronization module 29; the modules realize data interaction through RTI interfaces. The data acquisition module 27 and the display driving module 28 establish a connection function through a TCP/IP protocol, and a SOCKET channel of a WINDOWS system may be used as a channel for data transmission.

The receiving module 26 establishes a wireless communication link with the onboard subsystem and outputs the received satellite attitude information and satellite orbit information to the data acquisition module 27. The receiving module 26 may receive data from the onboard subsystem via a wireless data transmission channel or gateway.

The data acquisition module 27 analyzes the satellite attitude information and the satellite orbit information to obtain inter-satellite relative position information, inter-satellite relative velocity information, position information of the satellite relative inertial system, velocity information of the satellite relative inertial system, a three-axis attitude angle of the satellite, and a three-axis attitude angular velocity of the satellite, and outputs the information to the display drive module 28.

The display driver module 28 generates and outputs a satellite in-orbit flight picture using the picture generation tool and the received data and information. The display driving module 28 drives the animation scene update in real time by using a CONNECT programming interface provided by the STK tool, and displays the latest on-orbit flight condition of the satellite in the form of a near-view three-dimensional picture, a far-view three-dimensional picture, a two-dimensional picture, a dynamic parameter list and other windows.

The time synchronization module 29 outputs the time stamp to the data acquisition module 27 and the display driver module 28, and synchronizes the data received by the data acquisition module 27 and the display driver module 28. Specifically, the time synchronization module 29 controls the start and the step length of the display driver module 28 to ensure that the displayed data is the current latest flight parameter, and in the case that the animation display is delayed due to the network abnormality, controls the data acquisition module 27 to discard the data in the abnormal time, and reacquires the latest data to be provided to the display driver module 28. At regular intervals, the time synchronization module 29 broadcasts the current time to each module in the form of a timestamp, and determines whether each module lags behind as a basis for performing synchronization processing.

In order to ensure the fluency of the displayed picture, the ground subsystem of the invention also comprises a forecasting module, the forecasting module predicts and outputs the satellite orbit position in the time interval of receiving the two times of data to the display driving module 28 between the two times of data receiving according to the orbit data of the last time and the current environmental parameters, and ensures that the picture is fluent and uninterrupted; the time synchronization module 29 is further connected to the forecasting module, and controls the on/off of the track forecasting function of the forecasting module in the two data acquisition intervals.

FIG. 4 is a flow chart of a method for determining a satellite attitude orbit according to the present invention. Now, referring to fig. 4, a method for determining a satellite attitude orbit according to the present invention is described, specifically as follows:

step 401: the satellite subsystem collects satellite navigation signals, satellite attitude angle signals and satellite starlight elevation angles of various navigation satellite constellations;

the attitude orbit signal acquisition module 20 of the satellite subsystem acquires satellite navigation signals, satellite attitude angle signals and satellite starlight elevation angles of a plurality of navigation satellite constellations.

Step 402: the on-board subsystem calculates and generates inter-satellite relative position information, inter-satellite relative speed information, position information of the satellite relative to an inertial system, speed information of the satellite relative to the inertial system, a three-axis attitude angle of the satellite and a three-axis attitude angular speed of the satellite;

the satellite subsystem performs error correction and differential operation on the pseudo range, the pseudo range rate and the carrier phase data to obtain inter-satellite relative position information and inter-satellite relative speed information. The method specifically comprises the following steps:

step 4021, performing ephemeris error correction, ionosphere delay correction, integer ambiguity resolution and cycle slip detection on pseudo ranges and carrier phase data of various navigation satellite constellations to obtain the corrected pseudo ranges and carrier phase data of the various navigation satellite constellations; step 4022, using the model of the relative orbit element to perform smoothing and differential operation on the pseudo-range and carrier phase data of the corrected multiple navigation satellite constellations, and obtaining inter-satellite relative position information and inter-satellite relative speed information.

The ionosphere delay correction can adopt a model correction method, the model adopts a single-layer model, and the default ionosphere is distributed at a distance H from the ground_ionoOn an infinite thin layer. The delay of the propagation path of an electromagnetic wave as it passes through the ionosphere due to the rate of change of refraction is:

${\mathrm{\&delta;}}_p=-40.28\frac{1}{f^2}\&Integral;Nds---\left(1\right)$

let N_ΣWhere ═ Nds is the total number of electrons in the electromagnetic wave propagation path, the phase delay is:

the time delay is:

assuming that the included angle between the propagation path and the radial direction of the earth center of the receiver is theta, the integral along the propagation path in (1) is changed into the integral along the radial direction of the earth center, and according to the characteristics of the single-layer model, the following steps are provided:

${\mathrm{\&delta;}}_p=-40.28\frac{1}{f^2}\underset{H}{\&Integral;}\frac{N}{cos\mathrm{\&theta;}}dh=-40.28\frac{N_{\mathrm{\&Sigma;}}}{f^2{\mathrm{cos\&theta;}}^{\&prime;}}---\left(4\right)$

wherein theta' is the zenith distance of the observation satellite at the intersection point of the propagation path and the ionized layer, and satisfies

${\mathrm{sin\&theta;}}^{\&prime;}=\frac{R_e+H_{iono}}{R_e}sin(\mathrm{\&pi;}-{\mathrm{\&theta;}}_0)---\left(5\right)$

In the single-layer model, the total electron amount of the ionosphere in the daytime can be expressed as a cosine function of the local time, and is approximated as a constant at night, that is:

$N_{\mathrm{\&Sigma;}}=\left\{\begin{array}{c}{N}_0+Ncos\left(\frac{t-14}{12}\mathrm{\&pi;}\right),8:00<t<20:00\\ N_0\end{array}\right.---\left(6\right)$

wherein the integer ambiguity resolution specifically comprises: establishing all double-difference equation sets of the observed satellites in n epochs, namely:

$\left[\begin{array}{cccccc}{A}_1& 0& L& & 0& E\\ 0& A_2& & & & E\\ M& & O& & & M\\ & & & A_{n-1}& & \\ 0& L& & 0& A_n& E\end{array}\right]\left[\begin{array}{c}{X}_1\\ X_2\\ M\\ X_n\\ N_m\end{array}\right]=\left[\begin{array}{c}{L}_1\\ L_2\\ M\\ L_N\end{array}\right]---\left(7\right)$

in the formula A_nA coefficient matrix of a double difference equation set for the nth epoch; x_nObtaining a three-dimensional position coordinate vector for the nth epoch; n is a radical of_mAn integer ambiguity vector for the carrier phase; e is an identity matrix.

To reduce the complexity of the operation, the above equation needs to be simplified. For coefficient matrix A_iCarrying out QR decomposition A_iQ (i) r (i), transposing the sub-matrix of q (i) obtained by decomposition, and obtaining a left-hand product (7):

[Q(i)²]^T·L_i＝[Q(i)²]^T·λN+[Q(i)²]^T·_i(8)

the floating solution of the ambiguity is then obtained as:

$\hat{N}=\frac{1}{\mathrm{\&lambda;}}{\&lsqb;q{(1;m)}^T{Q}_q^{-1}q(1;m)\&rsqb;}^{-1}\&lsqb;q{(1;m)}^T{Q}_q^{-1}l(1;m)\&rsqb;---\left(9\right)$

the corresponding co-factors are:

$Q_{\hat{N}}={\&lsqb;q{(1;m)}^T{Q}_q^{-1}q(1;m)\&rsqb;}^{-1}---\left(10\right)$

on the basis of the solution, the ambiguity is searched by using LAMBDA, and the steps are as follows: variance covariance matrix Z transform (decorrelating transform); floating point decomposition decomposes an integer part and a decimal part; the floating point solution decimal part is transformed; setting the number of candidate solutions provided by searching; calculating the size of a search range; solving a corresponding fixed solution after the floating solution decimal part is subjected to Z transformation; and solving the fixed solution of the floating-point solution. The ambiguity search can be performed by using the existing search method, and the detailed description of the method is omitted here.

The cycle slip detection method specifically comprises the following steps: and performing cycle slip detection by using a Doppler frequency shift method, namely the carrier phase change rate. The following polynomial model is established for the carrier phase:

in the formula (12), the equation not including Δ N is a calculation formula before cycle slip, and the equation including Δ N is a calculation formula after cycle slip.

Selecting carrier phase observed values of 5 epochs and the change rate thereof:assuming that the carrier phase value of the first 4 epochs has no cycle slip, the carrier phase value is used for detecting whether the cycle slip occurs in the carrier phase of the 5 th epoch, and the following error equation is established: f ═ AX + v (12)

Wherein X is ═ a₀,a₁,a₂,a₃,ΔN]^T，

$A=\left[\begin{array}{ccccc}1& t_1& t_1^2& t_1^3& 0\\ 1& t_2& t_2^2& t_2^3& 0\\ 1& t_3& t_3^2& t_3^3& 0\\ 1& t_4& t_4^2& t_4^3& 0\\ 1& t_5& t_5^2& t_5^3& 0\\ 0& 1& 2t_1& 3t_1^2& 0\\ 0& 1& 2t_2& 3t_2^2& 0\\ 0& 1& 2t_3& 3t_3^2& 0\\ 0& 1& 2t_4& 3t_4^2& 0\\ 0& 1& 2t_5& 3t_5^2& 0\end{array}\right]$

According to the least squares algorithm, X ═ a^TA)^-1A^TAnd F, determining a threshold, and if | delta N | >, indicating that cycle slip exists in the carrier phase observed value of the 5 th epoch, wherein the cycle slip is estimated to be delta N.

Step 4021 uses the above method to correct the pseudorange and carrier phase data to improve the accuracy of orbit determination.

The method for establishing the model of the relative orbit element in step 4022 is as follows:

first, the relative position and relative velocity are defined as relative orbit state vectors, which are recorded asThe orbit element of the main star is marked as e (t), and the orbit element difference between the main star and the target star, namely the relative orbit element, can be obtained through Taylor series expansion, and is marked as e (t). And obtaining the relation between the relative orbit state vector and the relative orbit element through the coordinate transformation matrix.

The angular velocity of the primary star is expressed in the orbital coordinate system as:

Ω＝Ω_re_x+Ω_te_y+Ω_ne_z(13)

wherein Ω r, Ω t, and Ω n are radial angular velocity, tangential angular velocity, and sub-normal angular velocity, respectively.

The component array of the positions and the speeds of the main star and the target star in the orbit coordinate system of the main star is as follows:

(r_c)_o＝Re_x

${\left(V_c\right)}_o=\left(\stackrel{\&CenterDot;}{R}\right)e_x+\left({\mathrm{R\&Omega;}}_n\right)e_y+(-{\mathrm{R\&Omega;}}_t)e_z=\left(V_r\right)e_x+\left(V_t\right)e_y+\left(V_n\right)e_z$

(r_a)_o＝(r_c+r)_o＝(R+x)e_x+ye_y+ze_z

${\left(V_a\right)}_o=(V_r+\stackrel{\&CenterDot;}{x}-{\mathrm{y\&Omega;}}_n+{\mathrm{z\&Omega;}}_t)e_x+(V_t+\stackrel{\&CenterDot;}{y}+{\mathrm{x\&Omega;}}_n-{\mathrm{z\&Omega;}}_r)e_y+(V_n+\stackrel{\&CenterDot;}{z}-{\mathrm{x\&Omega;}}_t+{\mathrm{y\&Omega;}}_r)e_z$

where rc and ra respectively represent the earth center distance vectors of the main satellite and the target satellite, Vc and Va respectively represent the velocity vectors of the main satellite and the target satellite, R represents the position vector of the target satellite relative to the main satellite, R represents the earth center distance of the central satellite, Vr, Vt and Vn respectively represent the radial velocity, the tangential velocity and the sub-normal velocity,an array of components representing relative position and relative velocity, i.e. a state vector.

Making the mean orbit elements of the target and the main stars different, i.e. relative mean orbit elementsFor the system state vector, the relative motion equation is

Wherein,for the transformation matrix of relative average orbit elements, the relative distance (r) between stars and the azimuth angle are adoptedAnd a pitch angle (θ) as a measurement, the measurement being related to the instantaneous relative position (x, y, z) between the satellites as follows:

according to the relation between the instantaneous relative position between the satellites and the relative average orbit element, the relation between the measured value and the state vector can be obtained, and then a measurement equation is obtained, so that a model of the relative orbit element is obtained, and the model is written in the following form:

where ω (t) and v (t) are the system noise and the measurement noise, respectively; Σ (t) is a transition matrix from the average orbit element to the motion state, and d (t) is a transition matrix from the average orbit element to the instantaneous orbit element. As can be seen from the model of relative orbit elements, the state vector is a relative average orbit element, corresponding to the control performed to maintain the formation configuration; the measured values are the relative distance and the orientation between the satellites and correspond to the instantaneous measurement in the relative orbit determination, so that the model can unify the relative average orbit state required by control and the relative instantaneous orbit state obtained by measurement, and the contradiction between control and measurement in the relative orbit determination of the formation flight of the satellites under the influence of perturbation is solved.

The differential operation in step 4022 is to smooth the pseudorange and the carrier phase. The pseudo range is smoothed by adopting a carrier phase, and the smoothing method comprises the following steps:

where ρ is_s,kAnd (5) smoothing the pseudo range of the carrier phase at the moment k, wherein M is a smoothing time constant and is 20-100.

The satellite subsystem generates three-dimensional position information according to pseudo range, pseudo range rate and carrier phase data by using a least square optimal estimation algorithm and a space-time unified algorithm, and performs a federal filtering algorithm on the three-dimensional position information and satellite starlight elevation angle by using a flight dynamics model and a federal filtering algorithm to generate position information of a satellite relative to an inertial system and speed information of the satellite relative to the inertial system. The method specifically comprises the following steps:

4023, unifying coordinate systems by using a seven-parameter Boolean Sasa conversion formula according to pseudo ranges, pseudo range rates and carrier phase data of various navigation satellite constellations; step 4024, unifying pseudo ranges, pseudo range rates and time of carrier phase data of various navigation satellite constellations to UTC; step 4025, calculating to obtain three-dimensional position information by utilizing a combined positioning equation and a least square method for successive iteration; and 4026, performing a federal filtering algorithm on the three-dimensional position information and the satellite starlight elevation angle by using a flight dynamics model and the federal filtering algorithm to generate position information of the satellite relative to the inertial system and speed information of the satellite relative to the inertial system.

The method for unifying the coordinate systems in step 4023 includes:

the observation equation is established at epoch time as follows:

$\left\{\begin{array}{c}{\mathrm{\&rho;}}_{GPS}^i=\sqrt{{(X_{GPS}^i-x)}^2+{(Y_{GPS}^i-y)}^2+{(Z_{GPS}^i-z)}^2}+{\mathrm{c\&Delta;t}}_{GPS}\\ {\mathrm{\&rho;}}_{COMPASS}^i=\sqrt{{(X_{COMPASS}^i-x)}^2+{(Y_{COMPASS}^i-y)}^2+{(Z_{COMPASS}^i-z)}^2}+{\mathrm{c\&Delta;t}}_{COMPASS}\end{array}\right.---\left(18\right)$

wherein, (x, y, z)^TIn order to obtain the three-dimensional position information,three-dimensional position information of the ith GPS satellite in a WGS-84 coordinate system,is the three-dimensional position information of the jth Beidou satellite in the CGS2000 coordinate system, c delta t_GPS、cΔt_COMPASSThe parameters on the left side of the equation represent the actual observed distance, respectively the distance error caused by clock skew.

Because the coordinate system is not uniform, the origin translation and the coordinate axis rotation of the rectangular coordinate system are completed by using a seven-parameter Boolean's conversion formula, so that the uniformity of the coordinate system is realized. The conversion formula is as follows:

$\left[\begin{array}{c}{X}^{\&prime;}\\ Y^{\&prime;}\\ Z^{\&prime;}\end{array}\right]=\left[\begin{array}{c}\mathrm{\&Delta;}X\\ \mathrm{\&Delta;}Y\\ \mathrm{\&Delta;}Z\end{array}\right]+(1+k)\left[\begin{array}{ccc}1& {\mathrm{\&Omega;}}_z& -{\mathrm{\&Omega;}}_Y\\ -{\mathrm{\&Omega;}}_z& 1& -{\mathrm{\&Omega;}}_X\\ {\mathrm{\&Omega;}}_Y& {\mathrm{\&Omega;}}_X& 1\end{array}\right]\left[\begin{array}{c}X\\ Y\\ Z\end{array}\right]---\left(19\right)$

in step 4024, although the time of the two signals is unified to the UTC coordinated world, the clock frequencies of the systems also have a deviation, and the transmission of the two receivers also has a delay deviation, so in step 4025, the three-dimensional position information can be solved by using a combined positioning algorithm and a least square iterative algorithm, which are specifically as follows:

through the space-time unification processing, the observation equation can be unified into the following form:

${\mathrm{\&rho;}}_i^j=R_i^j+{\mathrm{c\&Delta;t}}_i=\sqrt{{(X_i^j-x)}^2+{(Y_i^j-y)}^2+{(Z_i^j-z)}^2}+{\mathrm{c\&Delta;t}}_i---\left(20\right)$

the unknowns required to be solved include 5 unknowns including three-dimensional position information and clock bias of GPS and COMPASS systems, and therefore, more than 5 navigation satellites must be observed at the same time to realize the final solution.

Assuming that m GPS satellites and n beidou satellites are observed, and m and n are natural numbers, the combined positioning equation can be described as follows:

AX＝B (21)

wherein,X＝(Δx,Δy,Δz,cΔt_GPS,cΔt_COMPASS)^T，

and (C) successively iterating by using a least square method until the error amount is smaller than a preset range, and obtaining the current three-dimensional position information X ═ (A)^TA)^-1A^TB。

The federal filtering algorithm performed in step 4026 specifically includes:

step 40261, establishing a satellite orbit dynamics model state equation, specifically:

$\stackrel{\&CenterDot;}{X}\left(t\right)=f_1\left(X\right(t),t)+W\left(t\right)---\left(22\right)$

wherein,w (t) is state noise; f (X), (t), t ═ v_x,v_y,v_z,f_x(X(t),t),f_y(X(t),t),f_z(X(t),t)]；

R_mThe method is the global non-spherical perturbation force, and specifically comprises the following steps:

step 40262, establishing a GNSS ranging state equation, specifically:

the ranging information provided by the ith satellite is:

$S_{1i}=\sqrt{{(X_{B1}-x)}^2+{(Y_{B1}-y)}^2+{(Z_{B1}-z)}^2}+{\mathrm{\&epsiv;}}_i---\left(23\right)$

the above equation is expressed as a state equation:

S₁(t_k)＝f₂(X(t_k),t_k)+V₁(t_k) (24)

step 40263, establishing a state equation of the star sensor angle observation information, specifically:

the starlight elevation angle of the satellite at the current moment can be obtained through the star direction measured by the star sensor and the horizon direction measured by the earth sensor; according to the model, the following results are obtained:

$\mathrm{\&gamma;}\left(t_k\right)=arccos(-\frac{r\left(t_k\right)g\hat{h}}{\left|r\left(t_k\right)\right|})-arcsin\left(\frac{R_{me}+{\mathrm{\&Delta;R}}_{me}}{\left|r\left(t_k\right)\right|}\right)+\mathrm{\&epsiv;}\left(t_k\right)---\left(25\right)$

the above equation is expressed as a state equation:

S₂(t_k)＝f₃(X(t_k),t_k)+V₁(t_k) (26)

step 40264, establishing a federated Kalman filtering model,

the method specifically comprises the following steps: establishing a Federal Kalman filtering model according to the formulas (22), (24) and (26); distributing weight values for filtering according to an information conservation principle; solving the state equation according to the distributed weight, the state equation of the satellite orbit dynamics model, the state equation of the GNSS distance measurement, the three-dimensional position information and the pseudo range to obtain a first result, and correcting the first result by using a predicted value calculated at the previous moment to obtain a current corrected first result; solving the state equation according to the distributed weight, the state equation of the satellite orbit dynamics model and the state equation observed by the star sensor angle to obtain a second result, and correcting the second result by utilizing the predicted value calculated at the previous moment to obtain the second result after current correction; obtaining a predicted value of the next moment by using a preset prediction algorithm according to the three-dimensional position information, the corrected first result and the corrected second result; and calculating to obtain the position information of the satellite relative to the inertial system and the speed information of the satellite relative to the inertial system by using a linear minimum variance algorithm and the corrected first calculation result, second calculation result and three-dimensional position information of the current moment.

Wherein, the system state noise and the observation noise are zero-mean white noise which are not correlated with each other.

The satellite-borne subsystem performs combined attitude determination by using an inertial device attitude determination signal, satellite attitude angle information and a satellite starlight elevation angle to generate a three-axis angle, and performs optimal estimation on the three-axis angle by using a federal filter algorithm to obtain the three-axis attitude angle of the satellite and the three-axis attitude angular velocity of the satellite. The method specifically comprises the following steps:

step 4027, performing rate integration on the satellite attitude angle signal output by the inertial device to obtain first triaxial angle information; 4028, correcting errors of the first triaxial angle information by using the satellite starlight elevation angle and the quaternion attitude angle output by the star sensor to obtain second triaxial angle information; step 4029, correcting the error of the first triaxial angle information by using the satellite attitude angle information output by the earth sensor to obtain third triaxial angle information.

Step 403: the satellite-borne subsystem adjusts the orbit of the satellite and the attitude of the satellite by utilizing the inter-satellite relative position information, the inter-satellite relative velocity information, the position information of the satellite relative to the inertial system, the velocity information of the satellite relative to the inertial system, the three-axis attitude angle of the satellite and the three-axis attitude angular velocity of the satellite. Specifically, the orbit of the satellite and the attitude of the satellite can be adjusted by using the existing attitude orbit control algorithm, and the method is not described herein again.

Step 403 may be followed by: and generating and outputting the satellite in-orbit flight picture by utilizing the inter-satellite relative position information, the inter-satellite relative speed information, the satellite relative inertial system position information, the satellite relative inertial system speed information, the satellite three-axis attitude angle and the satellite three-axis attitude angular speed.

In the preferred embodiment of the invention, in order to solve the problem of positioning accuracy reduction caused by completely depending on a GNSS receiver for positioning, the invention introduces information collected by an inertia device, a star sensor and an earth sensor during orbit determination and attitude determination calculation so as to assist the GNSS receiver and realize combined autonomous navigation, thereby changing the conventional attitude determination and orbit determination modes of a satellite mainly based on the ground and enabling the satellite to autonomously realize real-time high-accuracy and high-reliability orbit determination and attitude determination under the condition of being separated from a ground station; the method has the advantages that accurate position information is provided for a satellite to complete a space task, meanwhile, the operation cost of the ground subsystem is greatly reduced, the task burden of the ground subsystem is reduced, the concealment and the safety of the satellite task are improved, and the normal operation of the satellite for orbit determination and attitude determination can be still maintained when the ground subsystem is blocked or even damaged.

In summary, the above is a preferred embodiment of the present invention, and is not intended to limit the scope of the present invention. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.

Claims (10)

1. A satellite attitude orbit determination system, comprising an on-board subsystem, wherein said on-board subsystem comprises:

the attitude orbit signal acquisition module acquires satellite navigation signals, satellite attitude angle signals and satellite starlight elevation angles of various navigation satellite constellations and outputs the satellite navigation signals, the satellite attitude angle signals and the satellite starlight elevation angles to the attitude orbit determination module;

the attitude orbit determination module is used for obtaining pseudo-range, pseudo-range rate and carrier phase data from satellite navigation signals, carrying out error correction and differential operation on the pseudo-range, pseudo-range rate and carrier phase data, obtaining inter-satellite relative position information and inter-satellite relative speed information, and outputting the inter-satellite relative position information and the inter-satellite relative speed information to an attitude orbit control computer;

the attitude orbit determination module generates three-dimensional position information according to pseudo range, pseudo range rate and carrier phase data by using a least square optimal estimation algorithm and a space-time unified algorithm, performs a federal filtering algorithm on the three-dimensional position information and satellite starlight elevation angle by using a flight dynamics model and a federal filtering algorithm, generates position information of a satellite relative to an inertial system and speed information of the satellite relative to the inertial system, and outputs the position information and the speed information to an attitude and orbit control computer;

the attitude orbit determination module performs combined attitude determination by using a satellite navigation signal, satellite attitude angle information and a satellite starlight elevation angle to generate a three-axis angle, performs optimal estimation on the three-axis angle by using a federal filter algorithm to obtain the three-axis attitude angle of the satellite and the three-axis attitude angular velocity of the satellite, and outputs the three-axis attitude angle and the three-axis attitude angular velocity to an attitude orbit control computer;

the attitude and orbit control computer outputs an orbit change instruction to the execution module according to the received inter-satellite relative position information, inter-satellite relative speed information, position information of the satellite relative inertia system and speed information of the satellite relative inertia system, and outputs an attitude adjustment instruction to the execution module according to a three-axis attitude angle of the satellite and a three-axis attitude angular speed of the satellite;

the execution module adjusts the orbit of the satellite where the system is located according to the orbital transfer instruction and adjusts the attitude of the satellite where the system is located according to the attitude adjusting instruction;

the attitude orbit signal acquisition module comprises:

the multi-mode satellite navigation GNSS receiver is used for collecting satellite navigation information of various navigation satellite constellations and outputting the satellite navigation information to the attitude orbit determination module;

the inertial device is used for acquiring satellite attitude angle signals of the satellite and outputting the satellite attitude angle signals to the attitude orbit determination module;

the star sensor collects the satellite starlight elevation angle and the quaternion attitude angle of the satellite and outputs the satellite starlight elevation angle and the quaternion attitude angle to the attitude orbit determination module;

the earth sensor is used for acquiring satellite attitude angle information of a satellite and outputting the satellite attitude angle information to the attitude orbit determination module;

the multiple navigation satellite constellations at least comprise a Beidou satellite and a Global Positioning System (GPS) satellite;

the attitude trajectory determination module includes:

the data generation unit is used for generating pseudo-range, pseudo-range rate and carrier phase data of various navigation satellite constellations according to the satellite navigation information of the various navigation satellite constellations output by the GNSS receiver and outputting the pseudo-range, pseudo-range rate and carrier phase data to the correction unit and the space-time unified unit;

the correction unit is used for performing ephemeris error correction, ionosphere delay correction, integer ambiguity solution and cycle slip detection on the received pseudo ranges and carrier phase data of various navigation satellite constellations and outputting the corrected pseudo ranges and carrier phase data of the various navigation satellite constellations to the differential operation unit;

the differential operation unit is used for carrying out differential operation on the pseudo ranges and carrier phase data of the corrected various navigation satellite constellations by utilizing the model of the relative orbit element to obtain inter-satellite relative position information and inter-satellite relative speed information, and outputting the inter-satellite relative position information and the inter-satellite relative speed information to the attitude and orbit control computer;

the time-space unification unit is used for completing unification of a coordinate system by using a seven-parameter Boolean conversion formula according to pseudo-range, pseudo-range rate and carrier phase data of various navigation satellite constellations, unifying the time of the pseudo-range, the pseudo-range rate and the carrier phase data of the various navigation satellite constellations to coordinated Universal Time Coordinated (UTC), performing successive iteration by using a combined positioning equation and a least square method, calculating to obtain three-dimensional position information, and outputting the three-dimensional position information to the first federal filtering unit;

the model calculation unit is used for generating a state equation of the satellite orbit dynamics model and outputting the state equation to the first federal filtering unit;

the first federal filtering unit is used for solving a state equation by utilizing three-dimensional position information, satellite attitude angle information from a star sensor, a state equation of a satellite orbit dynamics model, a state equation of GNSS distance measurement and a state equation of star sensor angle observation to obtain position information of a satellite relative to an inertial system and speed information of the satellite relative to the inertial system, filtering the position information of the satellite relative to the inertial system and the speed information of the satellite relative to the inertial system by utilizing a linear minimum variance algorithm and a prediction algorithm, and outputting the filtered position information and the speed information to an attitude and orbit control computer;

the velocity integration unit is used for carrying out velocity integration on the satellite attitude angle signals output by the inertial device to obtain first triaxial angle information and outputting the first triaxial angle information to the second joint filtering unit, the first joint attitude determination unit and the second joint attitude determination unit;

the first combined attitude determination unit corrects errors of the first triaxial angle information output by the rate integration unit by using the satellite starlight elevation angle and the quaternion attitude angle output by the star sensor to obtain second triaxial angle information and outputs the second triaxial angle information to the second combined filter;

the second joint attitude determination unit corrects the error of the first triaxial angle information output by the rate integration unit by using the satellite attitude angle information output by the earth sensor to obtain third triaxial angle information and outputs the third triaxial angle information to a second joint filter;

and the second federated filtering unit is used for adjusting the first gain, the second gain and the third gain by using a main filtering equation of a federated filtering algorithm, filtering the first triaxial angle information by using the first gain, filtering the second triaxial angle information by using the second gain, filtering the third triaxial angle information by using the third gain, performing wild value elimination on the three filtered information, performing optimal estimation operation on the eliminated information, obtaining a satellite triaxial attitude angle and a satellite triaxial attitude angular velocity, and outputting the satellite attitude angular velocity to an attitude and orbit control computer.

2. The system of claim 1, wherein the on-board subsystem further comprises:

the inter-satellite communication module is used for establishing a communication link between the satellite and the satellite for cooperation, sending inter-satellite relative position information and inter-satellite relative speed information, receiving position information and speed information of another satellite and forwarding the position information and the speed information to the attitude orbit determination module;

the attitude orbit determination module further calculates and obtains inter-satellite relative position information and inter-satellite relative speed information according to the position information and the speed information of the other satellite, and outputs the inter-satellite relative position information and the inter-satellite relative speed information to the inter-satellite communication module.

3. The system of claim 1, further comprising:

the ground subsystem generates and outputs a satellite in-orbit flight picture by utilizing the satellite attitude information and the satellite orbit information received through the wireless communication link;

the attitude and orbit control computer of the satellite subsystem further generates and outputs attitude information and satellite orbit information of the satellite according to the orbital transfer instruction and the attitude adjusting instruction;

the satellite subsystem further comprises:

the satellite platform bus is used for establishing a data transmission channel between the attitude and orbit control computer and the remote measuring module;

and the telemetry module receives the satellite attitude information and the satellite orbit information from the attitude and orbit control computer through the platform bus and outputs a satellite attitude signal and a satellite orbit signal through a wireless link.

4. The system of claim 1, wherein the pose trajectory determination module further comprises:

the first prediction unit predicts the second triaxial angle information output by the first combined attitude determination unit by using a preset prediction algorithm and outputs the predicted second triaxial angle information to the first combined attitude determination unit so that the first combined attitude determination unit corrects the calculated second triaxial angle information;

and the second prediction unit predicts the third triaxial angle information output by the second combined attitude determination unit by using a preset prediction algorithm and outputs the predicted third triaxial angle information to the second combined attitude determination unit so that the second combined attitude determination unit corrects the calculated third triaxial angle information.

5. The system of claim 4, wherein the second federated filtering unit comprises:

each subtracter performs subtraction calculation on the received first triaxial angle information, second triaxial angle information or third triaxial angle information and the satellite triaxial attitude angle output by the optimal estimation subunit, and outputs the calculated difference to the first gain subunit, the second gain subunit or the third gain subunit;

the first gain subunit is used for filtering the difference value between the first triaxial angle information and the satellite triaxial attitude angle from the optimal estimation subunit according to the first gain and outputting the filtered first triaxial angle information to the first time synchronization subunit;

the second gain subunit is used for filtering the difference value between the second triaxial angle information and the satellite triaxial attitude angle from the optimal estimation subunit according to second gain and outputting the filtered second triaxial angle information to the first time synchronization subunit;

the third gain subunit is used for filtering the difference value between the third triaxial angle information and the satellite triaxial attitude angle from the optimal estimation subunit according to third gain and outputting the filtered third triaxial angle information to the first time synchronization subunit;

the first time synchronization subunit outputs the received filtered first triaxial angle information, second triaxial angle information and third triaxial angle information to the preprocessing subunit, calculates a first error, a second error and a third error according to a main filtering equation of the federal filter, sets a first gain of the first gain subunit by using the first error, sets a second gain of the second gain subunit by using the second error, and sets a third gain of the third gain subunit by using the third error;

the preprocessing subunit is used for performing wild value elimination on the first triaxial angle information, the second triaxial angle information and the third triaxial angle information which are received and filtered according to a preset threshold value, and outputting the eliminated information to the optimal estimation subunit;

and the optimal estimation subunit calculates the eliminated information according to an optimal estimation algorithm to obtain a satellite three-axis attitude angle and a satellite three-axis attitude angular velocity, and outputs the satellite three-axis attitude angular velocity to an attitude and orbit control computer.

6. The system of claim 4, wherein the first federated filtering unit comprises:

the information distribution subunit distributes weights to the first sub-filter and the second sub-filter according to an information conservation principle and an error evaluation result output by the main filter;

the first sub-filter is used for solving the state equation according to the distributed weight value, the state equation of the satellite orbit dynamic model, the state equation of the GNSS distance measurement, the three-dimensional position information and the pseudo range to obtain a first result, correcting the first result by using the predicted value from the main filter, and outputting the corrected first result to the main filter;

the second sub-filter is used for solving the state equation according to the distributed weight, the state equation of the satellite orbit dynamics model and the state equation observed by the star sensor angle to obtain a second result, correcting the second result by using the predicted value from the main filter, and outputting the corrected second result to the main filter;

and the main filter obtains a predicted value according to the three-dimensional position information, the corrected first result and the corrected second result by using a preset prediction algorithm and outputs the predicted value to the first sub-filter and the second sub-filter, carries out error evaluation on the corrected first result and the corrected second result, outputs an error evaluation result to the information distribution sub-unit, calculates and obtains position information of the satellite relative to an inertial system and speed information of the satellite relative to the inertial system by using a linear minimum variance algorithm, and outputs the position information and the speed information to the attitude and orbit control computer.

7. The system of claim 3, wherein the ground subsystem comprises:

the receiving module is used for establishing a wireless communication link with the satellite-borne subsystem and outputting the received satellite attitude information and the satellite orbit information to the data acquisition module;

the data acquisition module analyzes the satellite attitude information and the satellite orbit information to obtain inter-satellite relative position information, inter-satellite relative speed information, position information of the satellite relative inertial system, speed information of the satellite relative inertial system, a three-axis attitude angle of the satellite and a three-axis attitude angular speed of the satellite, and outputs the information to the display driving module;

the display driving module generates and outputs a satellite in-orbit flight picture by using the picture generation tool and the received data and information;

and the time synchronization module outputs the timestamp to the data acquisition module and the display driving module, and synchronizes the data received by the data acquisition module and the display driving module.

8. A method for satellite attitude orbit determination, the method comprising:

A. the satellite subsystem collects satellite navigation signals, satellite attitude angle signals and satellite starlight elevation angles of various navigation satellite constellations;

B. the satellite subsystem obtains pseudo range, pseudo range rate and carrier phase data from satellite navigation signals, and error correction and differential operation are carried out on the pseudo range, the pseudo range rate and the carrier phase data to obtain inter-satellite relative position information and inter-satellite relative speed information;

the satellite subsystem generates three-dimensional position information according to pseudo range, pseudo range rate and carrier phase data by using a least square optimal estimation algorithm and a space-time unified algorithm, and performs a federal filtering algorithm on the three-dimensional position information and satellite starlight elevation angle by using a flight dynamics model and a federal filtering algorithm to generate position information of a satellite relative to an inertial system and speed information of the satellite relative to the inertial system;

the satellite-borne subsystem performs combined attitude determination by using a satellite navigation signal, satellite attitude angle information and a satellite starlight elevation angle to generate a three-axis angle, and performs optimal estimation on the three-axis angle by using a federal filter algorithm to obtain the three-axis attitude angle of the satellite and the three-axis attitude angular velocity of the satellite;

C. the satellite subsystem adjusts the orbit of the sub-satellite and the attitude of the satellite by utilizing the inter-satellite relative position information, the inter-satellite relative speed information, the position information of the satellite relative to the inertial system, the speed information of the satellite relative to the inertial system, the three-axis attitude angle of the satellite and the three-axis attitude angular speed of the satellite;

and B, the error correction and differential operation of the pseudo range, the pseudo range rate and the carrier phase data comprises the following steps:

b1, performing ephemeris error correction, ionosphere delay correction, integer ambiguity resolution and cycle slip detection on pseudo-ranges and carrier phase data of various navigation satellite constellations to obtain the corrected pseudo-ranges and carrier phase data of the various navigation satellite constellations;

and B2, carrying out differential operation on the pseudo ranges and the carrier phase data of the corrected multiple navigation satellite constellations by using the model of the relative orbit element to obtain the inter-satellite relative position information and the inter-satellite relative speed information.

9. The method of claim 8, wherein the step B of generating three-dimensional position information from the pseudorange, the pseudorange rate, and the carrier phase data using a least squares optimal estimation algorithm and a space-time unified algorithm comprises:

b3, according to pseudo-range, pseudo-range rate and carrier phase data of various navigation satellite constellations, unifying a coordinate system by using a seven-parameter Boolean Sasa conversion formula;

b4, unifying pseudo ranges, pseudo range rates and time of carrier phase data of various navigation satellite constellations to coordinated universal time UTC;

and B5, calculating to obtain three-dimensional position information by utilizing the combination positioning equation and the successive iteration of the least square method.

10. The method of claim 8, wherein the step B of jointly determining attitude using satellite navigation signals, satellite attitude angle information, and satellite starlight elevation angles, and generating three-axis angles comprises:

b6, performing rate integration on the satellite attitude angle signal output by the inertial device to obtain first triaxial angle information;

b7, correcting the error of the first triaxial angle information by using the satellite starlight elevation angle and the quaternion attitude angle output by the star sensor to obtain second triaxial angle information;

and B8, correcting the error of the first triaxial angle information by using the satellite attitude angle information output by the earth sensor to obtain third triaxial angle information.