CN102494686A  Satellite attitude orbit determining system and method  Google Patents
Satellite attitude orbit determining system and method Download PDFInfo
 Publication number
 CN102494686A CN102494686A CN2011103151033A CN201110315103A CN102494686A CN 102494686 A CN102494686 A CN 102494686A CN 2011103151033 A CN2011103151033 A CN 2011103151033A CN 201110315103 A CN201110315103 A CN 201110315103A CN 102494686 A CN102494686 A CN 102494686A
 Authority
 CN
 China
 Prior art keywords
 satellite
 information
 attitude
 star
 pseudorange
 Prior art date
 Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
 Granted
Links
 210000004279 Orbit Anatomy 0.000 title claims abstract description 80
 239000000969 carrier Substances 0.000 claims abstract description 70
 238000001914 filtration Methods 0.000 claims abstract description 58
 230000001360 synchronised Effects 0.000 claims description 24
 238000004891 communication Methods 0.000 claims description 19
 238000004364 calculation method Methods 0.000 claims description 13
 238000006243 chemical reaction Methods 0.000 claims description 11
 239000005433 ionosphere Substances 0.000 claims description 11
 DMBHHRLKUKUOEGUHFFFAOYSAN Diphenylamine Chemical compound C=1C=CC=CC=1NC1=CC=CC=C1 DMBHHRLKUKUOEGUHFFFAOYSAN 0.000 claims description 8
 238000011156 evaluation Methods 0.000 claims description 8
 241001260012 Bursa Species 0.000 claims description 7
 230000005540 biological transmission Effects 0.000 claims description 7
 239000011159 matrix material Substances 0.000 description 13
 238000000034 method Methods 0.000 description 6
 230000000875 corresponding Effects 0.000 description 4
 238000009499 grossing Methods 0.000 description 4
 230000001276 controlling effect Effects 0.000 description 3
 230000004927 fusion Effects 0.000 description 3
 230000001737 promoting Effects 0.000 description 3
 239000002356 single layer Substances 0.000 description 3
 230000015572 biosynthetic process Effects 0.000 description 2
 230000001186 cumulative Effects 0.000 description 2
 238000000354 decomposition reaction Methods 0.000 description 2
 238000010586 diagram Methods 0.000 description 2
 230000000694 effects Effects 0.000 description 2
 238000005516 engineering process Methods 0.000 description 2
 238000005755 formation reaction Methods 0.000 description 2
 238000004989 laser desorption mass spectroscopy Methods 0.000 description 2
 238000005259 measurement Methods 0.000 description 2
 230000003068 static Effects 0.000 description 2
 206010057190 Respiratory tract infection Diseases 0.000 description 1
 230000002159 abnormal effect Effects 0.000 description 1
 229920001940 conductive polymer Polymers 0.000 description 1
 238000001514 detection method Methods 0.000 description 1
 238000010790 dilution Methods 0.000 description 1
 239000000284 extract Substances 0.000 description 1
 230000003993 interaction Effects 0.000 description 1
 201000002161 intrahepatic cholestasis of pregnancy Diseases 0.000 description 1
 239000010410 layer Substances 0.000 description 1
 230000004048 modification Effects 0.000 description 1
 238000006011 modification reaction Methods 0.000 description 1
 230000002123 temporal effect Effects 0.000 description 1
 230000001052 transient Effects 0.000 description 1
 230000017105 transposition Effects 0.000 description 1
Abstract
The invention provides a satellite attitude orbit determining system and a satellite attitude orbit determining method. The method comprises the following steps of: acquiring a pseudorange, a pseudorange rate and a carrier phase data from satellite navigation signals from an attitude orbit acquisition module by an attitude orbit determining module in an onsatellite subsystem, generating intersatellite relative position information, intersatellite relative velocity information and threedimensional position information by using the pseudorange, the pseudorange rate and the carrier phase data, performing a federal filtering algorithm on the threedimensional position information and a satellite starlight elevation to generate position information and velocity information of a relative satellite inertia system, performing joint attitude determining by using the satellite navigation signals, the satellite attitude angle information and the satellite starlight elevation to generate a triaxial angle, and performing optimal estimation on the triaxial angle to acquire a triaxial attitude angle and a triaxial attitude angular velocity of a satellite; and performing orbital transfer and attitude adjusting control by an attitude orbit control computer according to the information output by the attitude orbit determining module. By adopting the system and the method, the attitude determining precision and the orbit determining instantaneity can be improved.
Description
Technical field
The present invention relates to the Flight Vehicle Design field, particularly a kind of attitude of satellite track is confirmed system and method.
Background technology
Fig. 1 confirms the structural representation of system for existing attitude of satellite track.Combine Fig. 1 at present, existing attitude of satellite track is confirmed that system describes, specific as follows:
Existing attitude of satellite track confirms that system comprises subsystem 10 and ground subsystem 11 on the star.Wherein, subsystem 10 and is accomplished the appearance of deciding of satellite according to the data of confirming the attitude of satellite in order to obtaining the data of confirming satellite orbit and definite attitude of satellite on the star, and the data downstream of confirming satellite orbit is transferred to ground subsystem 11; Ground subsystem 11 is carried out the orbit determination computing according to the data of existing flight dynamics model and definite satellite orbit.
Subsystem 10 comprises inertia device 101, rail control computing machine (AOCC) 102, GPS receiver 103, USB answering machine 104 and LDMS 105 on the star; Ground subsystem 11 comprises the ground distance finding station 111 and first computing machine 112.Wherein, inertia device 101 is confirmed equation according to the attitude of existing inertial navigation, and the satellite at its place is carried out deciding appearance; The data that rail control computing machine 102 is gathered inertia device 101 according to the data acquisition orbit determination data that the original power of setting up is learned equation and collected, are sent the orbit determination data to USB answering machine 104 according to preset period; GPS (Global Positioning System, GPS) receiver 103 sends to USB answering machine 104 with the GPS position data that collects according to preset period; In the time of in the satellite process observing and controlling scope of ground subsystem 11 that subsystem 10 belongs on star, USB answering machine 104 is through the less radiofrequency passage, with the ground distance finding station 111 that reaches ground subsystem 11 under orbit determination data and the GPS position data; LDMS 105 is set up communication link with ground distance finding station 111, so that ground distance finding station 111 obtains the laser ranging value; Ground distance finding station 111 sends to first computing machine 112 with orbit determination data, GPS position data and laser ranging value; First computing machine 112 carries out the orbit determination computing according to flight dynamics model and the data that receive.Wherein, it is the mathematical computations process of a continuous match iteration that the orbit determination that first computing machine 112 is carried out is calculated, and execution cycle is long more, and the orbit determination result's who obtains precision is high more.
In sum; Existing attitude of satellite track confirms in the system that the orbit determination result need rely on the distance between satellite and ground, and the orbit determination calculating process is that ground subsystem is accomplished the complex mathematical calculating process; In order to improve orbit determination accuracy, need expend long execution cycle; It is not high that simple interest is accomplished the precision of deciding appearance with inertia device, is still waiting further raising; Confirm in case ground subsystem is participated in satellite orbit, can't guarantee the realtime of orbit determination.
Summary of the invention
In view of this, the object of the present invention is to provide a kind of attitude of satellite track to confirm system, this system can improve the precision of deciding appearance, realizes the realtime that track is confirmed.
The object of the present invention is to provide a kind of attitude of satellite track to confirm method, this method can improve the precision of deciding appearance, realizes the realtime that track is confirmed.
For achieving the above object, technical scheme of the present invention specifically is achieved in that
A kind of attitude of satellite track is confirmed system, and this system comprises subsystem on the star, and subsystem comprises on the said star:
Attitude track signal acquisition module, satellite navigation signals, attitude of satellite angle signal and the satellite starlight elevation angle of gathering multiple Navsat constellation, output satellite navigation signals, attitude of satellite angle signal and the satellite starlight elevation angle to attitude track determination module;
Attitude track determination module; From satellite navigation signals, obtain pseudorange, pseudorange rates and carrier phase data; Pseudorange, pseudorange rates and carrier phase data are carried out error correction and calculus of differences, relative velocity between relative position information and star between the acquisition star, and export the rail control computing machine to;
Said attitude track determination module utilizes least square optimal estimation algorithm and spacetime to unify algorithm; Advance to generate three dimensional local information according to pseudorange, pseudorange rates and carrier phase data; Utilize flight dynamics model and federal filtering algorithm; The three dimensional local information and the satellite starlight elevation angle are carried out federal filtering algorithm, generate the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system, and export the rail control computing machine to;
Said attitude track determination module utilizes satellite navigation signals, attitude of satellite angle information and the satellite starlight elevation angle; Unite and decide appearance; Generate three shaft angle degree, utilize federal filtering algorithm, three shaft angle degree are carried out optimal estimation; Obtain the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite, and export the rail control computing machine to;
Said rail control computer based is in the positional information of relative velocity between relative position information, star, satellite relative inertness system between the star that receives and the velocity information of satellite relative inertness system; Output becomes rail and instructs to Executive Module; The threeaxis attitude angular speed of satellitebased threeaxis attitude angle and satellite, the output posture adjustment is instructed to Executive Module;
Execution module is according to the track that becomes rail instruction Adjustment System place satellite, according to the attitude of posture adjustment instruction Adjustment System place satellite.
Preferably, subsystem also comprises on the said star:
Communication module between star, in order to set up and the intersatellite communication link of cooperating, relative velocity between relative position information and star between the transmission star receives positional information and its velocity information of another satellite and is transmitted to attitude track determination module;
Said attitude track determination module is further according to the positional information of another satellite and its velocity information, calculates to obtain between star relative velocity between relative position information and star, and exports communication module between star to.
Preferably, this system also comprises:
Ground subsystem is utilized the attitude of satellite information and the satelliteorbit information that receive through wireless communication link, generates satellite in rail flight picture and output;
The rail control computer of subsystem generates attitude information and the satelliteorbit information and the output of satellite further based on becoming rail instruction and posture adjustment instruction on the said star;
Subsystem further comprises on the said star:
The satellite platform bus is set up the data transmission channel between rail control computing machine and telemetry module;
Telemetry module through platform bus, receives attitude of satellite information and satelliteorbit information from the rail control computing machine, and through Radio Link output satellite attitude signal and satellite orbit signal.
In the said system, said attitude track signal acquisition module comprises:
Multimodal satellite navigation GNSS receiver is gathered the satellite navigation information of multiple Navsat constellation, and exports attitude track determination module to;
Inertia device is gathered the attitude of satellite angle signal of satellite, and is exported attitude track determination module to;
Star sensor is gathered the satellite starlight elevation angle and the hypercomplex number attitude angle of satellite, and is exported attitude track determination module to;
Earth sensor is gathered the attitude of satellite angle information of satellite, and is exported attitude track determination module to;
Said multiple Navsat constellation comprises bigdipper satellite and global position system GPS satellite at least.
In the said system, said attitude track determination module comprises:
Data generating unit, the satellite navigation information according to the multiple Navsat constellation of GNSS receiver output generates pseudorange, pseudorange rates and the carrier phase data of multiple Navsat constellation, and exports amending unit and spacetime uniformity unit to;
Amending unit; The pseudorange that receives multiple aeronautical satellite constellation and carrier phase data are carried out ephemeris error correction, ionosphere timedelay are proofreaied and correct, integer ambiguity is found the solution and cycle slip detects, export the pseudorange and the carrier phase data of revised multiple aeronautical satellite constellation to the calculus of differences unit;
The calculus of differences unit utilizes the model of relative orbit key element, and the pseudorange and the carrier phase data of revised multiple Navsat constellation are carried out calculus of differences, obtains between star relative velocity between relative position information and star, exports the rail control computing machine to;
The spacetime uniformity unit; Pseudorange, pseudorange rates and carrier phase data according to multiple Navsat constellation; Utilize seven parameter Bursa conversion formulas to accomplish the unification of coordinate system, the time unification of pseudorange, pseudorange rates and the carrier phase data of multiple Navsat constellation to Coordinated Universal Time(UTC) UTC, is utilized integrated positioning equation and least square method successive iteration; Calculate and obtain three dimensional local information, export the first federal filter unit to;
The Model Calculation unit generates the state equation of satellite orbit kinetic model, and exports the first federal filter unit to;
The first federal filter unit; Utilize three dimensional local information, attitude of satellite angle information, the state equation of satellite orbit kinetic model, the state equation of GNSS range finding and the state equation of star sensor angular observation from star sensor; Carry out finding the solution of state equation; Obtain the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system; Utilize linear minimum variance algorithm and prediction algorithm, the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system are carried out filtering, and export the rail control computing machine to;
The speed integral unit carries out the speed integration to the attitude of satellite angle signal of inertia device output, obtains the one or three shaft angle degree information, and exports the second federal filter unit, first to and unite and decide appearance unit and second and unite and decide the appearance unit;
First unites and decides the appearance unit, utilizes the satellite starlight elevation angle and the hypercomplex number attitude angle of star sensor output, and the error of the one or three shaft angle degree information of speed integral unit output is proofreaied and correct, and obtains the two or three shaft angle degree information and exports the second federal wave filter to;
Second unites and decides the appearance unit, utilizes the attitude of satellite angle information of earth sensor output, and the error of the one or three shaft angle degree information of speed integral unit output is proofreaied and correct, and obtains the three or three shaft angle degree information and exports the second federal wave filter to;
The second federal filter unit; Utilize the main filtering equations of federal filtering algorithm that first gain, second gain and the 3rd gain are adjusted; Utilize first gain that the one or three shaft angle degree information is carried out filtering, utilize second gain that the two or three shaft angle degree information is carried out filtering, utilize the 3rd gain that the three or three shaft angle degree information is carried out filtering; Filtered three information are carried out unrulyvalue rejecting; Information after rejecting is carried out the optimal estimation computing, obtain the threeaxis attitude angular velocity of satellite threeaxis attitude angle and satellite, and export the rail control computing machine to.
Preferably, said attitude track determination module also comprises:
First predicting unit; Utilize preset prediction algorithm; Uniting the two or the three shaft angle degree information of deciding the output of appearance unit to first predicts; The two or three shaft angle degree information to the first of the prediction of output is united and is decided the appearance unit, so that first unites and decide the appearance unit and revise calculating the two or the three shaft angle degree information that obtains;
Second predicting unit; Utilize preset prediction algorithm; Uniting the three or the three shaft angle degree information of deciding the output of appearance unit to second predicts; The three or three shaft angle degree information to the second of the prediction of output is united and is decided the appearance unit, so that second unites and decide the appearance unit and revise calculating the three or the three shaft angle degree information that obtains.
In the said system, the said second federal filter unit comprises:
Three subtracters; The one or three shaft angle degree information, the two or three shaft angle degree information or the three or three shaft angle degree information that each subtracter will receive; Carry out subtraction with the satellite threeaxis attitude angle of optimal estimation subelement output, difference to the first gain subelement, the second gain subelement or the 3rd gain subelement that obtains calculated in output;
The first gain subelement according to first gain, carries out filtering to the one or three shaft angle degree information with from the difference of the satellite threeaxis attitude angle of optimal estimation subelement, exports filtered synchronous subelement of the one or three shaft angle degree information to the very first time;
The second gain subelement according to second gain, carries out filtering to the two or three shaft angle degree information with from the difference of the satellite threeaxis attitude angle of optimal estimation subelement, exports filtered synchronous subelement of the two or three shaft angle degree information to the very first time;
The 3rd gain subelement according to the 3rd gain, carries out filtering to the three or three shaft angle degree information with from the difference of the satellite threeaxis attitude angle of optimal estimation subelement, exports filtered synchronous subelement of the three or three shaft angle degree information to the very first time;
Synchronous subelement of the very first time; To receive filtered the one or three shaft angle degree information, the two or three shaft angle degree information and the three or three shaft angle degree information and export the preliminary treatment subelement to; And according to the main filtering equations of federal wave filter; Calculate first error, second sum of errors the 3rd error; Utilize first error that first gain of the first gain subelement is set; Utilize second error that second gain of the second gain subelement is set, utilize the 3rd error that the 3rd gain of the 3rd gain subelement is set;
The preservice subelement according to preset threshold value, carries out unrulyvalue rejecting to receiving filtered the one or three shaft angle degree information, the two or three shaft angle degree information and the three or three shaft angle degree information, exports the information after rejecting to the optimal estimation subelement;
The optimal estimation subelement according to the optimal estimation algorithm, carries out computing to the information after rejecting, and obtains the threeaxis attitude angular velocity of satellite threeaxis attitude angle and satellite, and exports the rail control computing machine to.
In the said system, the said first federal filter unit comprises:
The information distribution subelement according to the error evaluation result of information conservation principle and senior filter output, is that first subfilter and second subfilter are distributed weights;
First subfilter; According to the state equation of the weights that distribute, satellite orbit kinetic model, state equation, three dimensional local information and the pseudorange of GNSS range finding; Carry out finding the solution of state equation, obtain first result, be used to predicted value from senior filter; First result is revised, export revised first result to senior filter;
Second subfilter; According to the weights that distribute, the state equation of satellite orbit kinetic model, the state equation of star sensor angular observation; Carry out finding the solution of state equation, obtain second result, be used to predicted value from senior filter; Second result is revised, export revised second result to senior filter;
Senior filter; According to three dimensional local information, revised first result and revised second result; Utilize preset prediction algorithm; Obtain predicted value and export first subfilter to and second subfilter, revised first result and revised second result are carried out error evaluation, output error assessment result to information distribution subelement; Utilize linear minimum variance algorithm computation to obtain the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system, and export the rail control computing machine to.
In the said system, said ground subsystem comprises:
Receiver module, the wireless communication link of subsystem exports the attitude of satellite information and the satelliteorbit information that receive to data acquisition module on foundation and the star;
Data acquisition module; Resolve attitude of satellite information and satelliteorbit information; Obtain the positional information that relative velocity, satellite relative inertness are between relative position information, star between star, velocity information, the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite of satellite relative inertness system, and export driver module to;
Driver module utilizes picture Core Generator, the data that receive and information, generates satellite in rail flight picture and output;
Time synchronized module, output time are stabbed to data acquisition module and driver module, the data that synchronous data collection module and driver module receive.
A kind of attitude of satellite track is confirmed method, and this method comprises:
Subsystem is gathered satellite navigation signals, attitude of satellite angle signal and the satellite starlight elevation angle of multiple Navsat constellation on A, the star;
Subsystem obtains pseudorange, pseudorange rates and carrier phase data on B, the star from satellite navigation signals, and pseudorange, pseudorange rates and carrier phase data are carried out error correction and calculus of differences, relative velocity between relative position information and star between the acquisition star;
Subsystem utilizes least square optimal estimation algorithm and spacetime to unify algorithm on the star; Advance to generate three dimensional local information based on pseudorange, pseudorange rates and carrier phase data; Utilize flight dynamics model and federal filtering algorithm; The three dimensional local information and the satellite starlight elevation angle are carried out federal filtering algorithm, generate the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system;
Subsystem utilizes satellite navigation signals, attitude of satellite angle information and the satellite starlight elevation angle on the star; Unite and decide appearance, generate three shaft angle degree, utilize federal filtering algorithm; Three shaft angle degree are carried out optimal estimation, obtain the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite;
Subsystem utilizes the positional information of relative velocity between relative position information between star, star, satellite relative inertness system, velocity information, the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite of satellite relative inertness system, the track of adjustment subsatellite and the attitude of satellite on C, the star.
In the said method, step B is said to carry out error correction and calculus of differences comprises to pseudorange, pseudorange rates and carrier phase data:
B1, the pseudorange of multiple Navsat constellation and carrier phase data are carried out that ephemeris error correction, ionosphere timedelay are proofreaied and correct, integer ambiguity is found the solution and jumped in week and detect, obtain the pseudorange and the carrier phase data of revised multiple Navsat constellation;
B2, utilize the model of relative orbit key element, the pseudorange and the carrier phase data of revised multiple Navsat constellation are carried out calculus of differences, obtain between star relative velocity between relative position information and star.
In the said method, said least square optimal estimation algorithm and the spacetime of utilizing of step B is unified algorithm, generates three dimensional local information according to pseudorange, pseudorange rates and carrier phase data and comprises:
B3, the pseudorange according to multiple Navsat constellation, pseudorange rates and carrier phase data utilize seven parameter Bursa conversion formulas to accomplish the unification of coordinate system;
The time unification of B4, the pseudorange with multiple Navsat constellation, pseudorange rates and carrier phase data is to Coordinated Universal Time(UTC) UTC;
B5, utilize integrated positioning equation and least square method successive iteration, calculate and obtain three dimensional local information.
In the said method, said satellite navigation signals, attitude of satellite angle information and the satellite starlight elevation angle of utilizing of step B united and decided appearance, generates three shaft angle degree and comprises:
B6, the attitude of satellite angle signal that inertia device is exported carry out the speed integration, obtain the one or three shaft angle degree information;
B7, the satellite starlight elevation angle and the hypercomplex number attitude angle of utilizing star sensor to export are proofreaied and correct the error of the one or three shaft angle degree information, obtain the two or three shaft angle degree information;
B8, the attitude of satellite angle information that utilizes earth sensor to export are proofreaied and correct the error of the one or three shaft angle degree information, obtain the three or three shaft angle degree information.
Visible by abovementioned technical scheme; The invention provides a kind of attitude of satellite track and confirm system and method; In the system and method for the present invention, rely on the GNSS receiver no longer fully and position, but introduced the information that inertia device, star sensor and earth sensor are gathered; With the assisted GNSS receiver; Realize associating independent navigation, having changed is that main satellite is decided appearance and orbit determination mode with ground, makes realtime high precision of can be under the situation of the disengaging land station autonomous realization of satellite and highly reliable orbit determination and decides appearance.
Description of drawings
Fig. 1 confirms the structural representation of system for existing attitude of satellite track.
Fig. 2 confirms the structural representation of system for attitude of satellite track of the present invention.
Fig. 3 is the structural representation of attitude track determination module of the present invention.
Fig. 4 confirms the process flow diagram of method for attitude track of the present invention.
Embodiment
For make the object of the invention, technical scheme, and advantage clearer, below with reference to the accompanying drawing embodiment that develops simultaneously, to further explain of the present invention.
Fig. 2 confirms the structural representation of system for attitude of satellite track of the present invention.Combine Fig. 2 at present, attitude of satellite track of the present invention is confirmed that system describes, specific as follows:
Attitude of satellite track of the present invention is confirmed the orbit determination of system and is decided appearance and all on subsystem on the star, realize.Subsystem comprises on the star: attitude track signal acquisition module 20, attitude track determination module 21, rail control computing machine 22 and execution module 23.
Attitude track signal acquisition module 20 is gathered satellite navigation signals, attitude of satellite angle signal and the satellite starlight elevation angle of multiple Navsat constellation, and exports attitude track determination module 21 to.Wherein, the multiple Navsat constellation that the present invention mentions is at least the Big Dipper (COMPASS) satellite and gps satellite, and satellite navigation signals also can be the navigation signal of multiple systems such as GALILEO, GLONASS, gives unnecessary details no longer one by one at this.Attitude track signal acquisition module 20 can pass through 422 interfaces (not shown among Fig. 2) and connect attitude track determination module 21.Wherein, the satellite navigation signals of attitude track signal acquisition module 20 collection acquisitions comprises pseudorange, pseudorange rates and carrier phase at least.
Attitude track determination module 21 obtains pseudorange, pseudorange rates and carrier phase data from satellite navigation signals; Pseudorange, pseudorange rates and carrier phase data are carried out error correction and calculus of differences; Relative velocity between relative position information and star between the acquisition star, and export rail control computing machine 22 to.
Attitude track determination module 21 utilizes least square optimal estimation algorithm and spacetime to unify algorithm; Advance to generate three dimensional local information according to pseudorange, pseudorange rates and carrier phase data; Utilize flight dynamics model and federal filtering algorithm; Carry out federal filtering algorithm to three dimensional local information with from the satellite starlight elevation angle of attitude track signal acquisition module 20, generate the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system, and export rail control computing machine 22 to.
Attitude track determination module 21 utilizes satellite navigation signals, attitude of satellite angle information and the satellite starlight elevation angle; Unite and decide appearance; Generate three shaft angle degree, utilize federal filtering algorithm, three shaft angle degree are carried out optimal estimation; Obtain the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite, and export rail control computing machine 22 to.
The positional information that relative velocity, satellite relative inertness are between relative position information, star between the rail control computing machine 22 basis stars that receive and the velocity information of satellite relative inertness system; Output becomes rail and instructs to execution module; According to the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite, the output posture adjustment is instructed to execution module 23.
Execution module 23 is according to the track that becomes rail instruction Adjustment System place satellite, according to the attitude of posture adjustment instruction Adjustment System place satellite.
In order to realize the high precision relative positioning between the satellite, subsystem also comprises on the star of the present invention: communication module 24 between star; The intersatellite communication link of communication module 24 in order to set up and to cooperate between star, relative velocity between relative position information and star between transmission and reception star, particularly, communication module 24 is obtained the position of another satellite between star through communication link; Attitude track determination module 22 further carries out calculus of differences with the position of selfposition and another satellite and finds the solution the relative position between the satellite, and can be further relative velocity exports communication module 24 between star between relative position information and star with calculating between the star that obtains.
In order to show the moving situation of satellite in rail flight more intuitively, attitude of satellite track of the present invention confirms that system also comprises a ground subsystem.Wherein, attitude of satellite information and satelliteorbit information that ground subsystem utilizes wireless communication link to receive generate satellite in rail flight picture and output; The rail control computing machine 22 of subsystem generate the attitude information and the satelliteorbit information of satellite, and platform bus exports telemetry module 25 to via satellite according to becoming rail instruction and posture adjustment instruction on the star; Telemetry module 25 receives attitude of satellite information and the satelliteorbit information from rail control computing machine 22, and through Radio Link output satellite attitude signal and satellite orbit signal to ground subsystem.
Wherein, Ground subsystem generates satellite in the required information of rail flight picture; At least comprise: time on the current star, based on the triaxial coordinate value of inertial system and rate value, based on the threeaxis attitude angle of orbital coordinate system and angular speed, satellite flight height, longitude, dimension etc.; These information all can be passed through telemetry module 25, obtain from rail control computing machine 22, at the concrete acquisition methods of this each information that repeats no more.After rail control computing machine 22 receives the data of attitude track determination module 21, be used as the input of attitude track control on the one hand, on the other hand, periodically with reaching ground subsystem 25 times through telemetry module after the packing data.
Satellite platform bus of the present invention can adopt the 1553B platform bus.
In order to reach optimum locating effect, attitude track signal acquisition module of the present invention 20 the is compatible receiver of multiple Navsat constellation signal has improved the number of nautical star, reduces the precision dilution gfactor, has improved the redundance and the reliability of system.Wherein, attitude track signal acquisition module 20 comprises: GPS (Global Navigation Satellite System, GNSS) receiver 201, star sensor 202, earth sensor 203 and inertia device 204.
GNSS receiver 201 is gathered the satellite navigation information of bigdipper satellite and gps satellite, and exports attitude track determination module 21 to.Wherein, the satellite navigation signals of bigdipper satellite is the COMPASS system, and the satellite navigation signals of gps satellite is the GPS system.The signal that the GNSS receiver 201 that attitude track signal acquisition module 20 comprises directly collects is a navigation message; Many information have been comprised in the text; Such as, the almanac data of each Navsat constellation, temporal information, pseudorange, pseudorange rates and carrier phase in the Navsat constellation, abovementioned information all are through encrypting; GNSS receiver 201 need be deciphered earlier and export attitude track determination module 21 again to after getting access to a frame navigation message.
Star sensor 202 is gathered the satellite starlight elevation angle and the hypercomplex number attitude angle of satellite, and exports attitude track determination module 21 to.Earth sensor 203 is gathered the attitude of satellite angle information of satellite, and exports attitude track determination module 22 to.Inertia device 204 is gathered the attitude of satellite angle signal of satellite, and exports attitude track determination module 22 to.Wherein, star sensor 202, earth sensor 203 and inertia device 204 can adopt existing components and parts, no longer the structure of abovementioned components and parts are given unnecessary details at this.
Fig. 3 is the structural representation of attitude track determination module of the present invention.Combine Fig. 3 at present, attitude track determination module of the present invention is described, specific as follows:
Attitude track determination module 21 of the present invention comprises: the federal filter unit of data generating unit 210, amending unit 211, calculus of differences unit 212, spacetime uniformity unit 213, Model Calculation unit 214, first 215, speed integral unit 216, first are united and are decided appearance unit 217, second and unite and decide the appearance unit 218 and the second federal wave filter 219.
Data generating unit 210 generates pseudorange, pseudorange rates and the carrier phase data of multiple Navsat constellation according to the satellite navigation information of the multiple Navsat constellation of GNSS receiver 201 outputs, and exports amending unit 211 and spacetime uniformity unit 213 to.Particularly; The agreement that data generating unit 210 is supported according to multiple Navsat constellation; Such as, the agreement that agreement that GPS supported and bigdipper satellite are supported extracts pseudorange, pseudorange rates and carrier phase in the text after the deciphering of GNSS receiver 201 outputs.
Pseudorange and the carrier phase data that 211 pairs of amending units receive multiple Navsat constellation are carried out that ephemeris error correction, ionosphere timedelay are proofreaied and correct, integer ambiguity is found the solution and are jumped in week and detect, and export the pseudorange and the carrier phase data of revised multiple Navsat constellation to calculus of differences unit 212.
Calculus of differences unit 212 utilizes the model of relative orbit key element, and the pseudorange and the carrier phase data of revised multiple Navsat constellation are carried out calculus of differences, and relative velocity between relative position information and star exports rail control computing machine 22 between the acquisition star.
Spacetime uniformity unit 213 is according to pseudorange, pseudorange rates and the carrier phase data of multiple Navsat constellation; Utilize seven parameter Bursa conversion formulas to accomplish the unification of coordinate system; With the time unification of pseudorange, pseudorange rates and the carrier phase data of multiple Navsat constellation to Coordinated Universal Time(UTC) UTC; Utilize integrated positioning equation and least square method successive iteration, calculate and obtain three dimensional local information, export the first federal filter unit 215 to.
Model Calculation unit 214 generates the state equation of satellite orbit kinetic model, and exports the first federal filter unit 215 to.Particularly; Model Calculation unit 214 utilizes the state equation that generates the satellite orbit kinetic model based on the sequential disposal route of promoting Kalman filtering; Such as; Sequential disposal route based on promoting Kalman filtering comprises: M M positional value (M＜10) constantly defines covariance matrix P0 as initial value before choosing; A preceding M value is done linear minimum variance handle, obtain the estimated value of current time position; According to the step of promoting Kalman filtering, obtain statetransition matrix and new covariance matrix P1 according to P0 and preceding M value constantly; By that analogy, M1 value and current estimated value before utilizing, the current state transition matrix can obtain next location estimate constantly; Can obtain new covariance matrix P2 and new statetransition matrix according to covariance matrix P1 with next location estimate constantly.In order to improve precision, can make abovementioned covariance progressively level off to zero.
The first federal filter unit 215 utilizes three dimensional local information, the attitude of satellite angle information from star sensor 202, the state equation of satellite orbit kinetic model, the state equation of GNSS range finding and the state equation of star sensor angular observation; Carry out finding the solution of state equation; Obtain the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system; Utilize linear minimum variance algorithm and prediction algorithm; The positional information of satellite relative inertness system and the velocity information of satellite relative inertness system are carried out filtering, and export rail control computing machine 22 to.
The attitude of satellite angle signal of the 216 pairs of inertia devices of speed integral unit output carries out the speed integration, obtains the one or three shaft angle degree information, and exports the second federal filter unit 219, first to and unite and decide appearance unit 217 and second and unite and decide appearance unit 218.
First unites and decides the satellite starlight elevation angle and the hypercomplex number attitude angle that appearance unit 217 utilizes star sensor 202 output; Error to the one or three shaft angle degree information of speed integral unit 216 output is proofreaied and correct, and obtains the two or three shaft angle degree information and exports the second federal wave filter 219 to.
Why employing is united is decided appearance; Utilize the low error of the angle information that realtime and the star sensor 202 of the attitude of satellite angle information of inertia device 204 output export to remedy each other exactly; Under the situation that guarantees realtime; The high level error of the attitude of satellite angle information that reduction inertia device 204 collects had both solved the cycle length of star sensor 202, the problem of realtime difference, had solved the big problem of cumulative errors of inertia device 204 again.
Second unites and decides the attitude of satellite angle information that appearance unit 218 utilizes earth sensor 203 output, and the error of the one or three shaft angle degree information of speed integral unit 216 outputs is proofreaied and correct, and obtains the three or three shaft angle degree information and exports the second federal wave filter 219 to.
Why employing is united is decided appearance; Utilize the low error of the angle information that realtime and the earth sensor 203 of the attitude of satellite angle information of inertia device 204 output export to remedy each other exactly; Under the situation that guarantees realtime; The high level error of the attitude of satellite angle information that reduction inertia device 204 collects had both solved the cycle length of earth sensor 203, the problem of realtime difference, had solved the big problem of cumulative errors of inertia device 204 again.
The second federal filter unit 219 utilizes the main filtering equations of federal filtering algorithm that first gain, second gain and the 3rd gain are adjusted; Utilize first gain that the one or three shaft angle degree information is carried out filtering; Utilize second gain that the two or three shaft angle degree information is carried out filtering; Utilize the 3rd gain that the three or three shaft angle degree information is carried out filtering; Filtered three information are carried out unrulyvalue rejecting, the information after rejecting is carried out the optimal estimation computing, obtain the threeaxis attitude angular velocity of satellite threeaxis attitude angle and satellite and export rail control computing machine 22 to.
In order to improve the precision that attitude is confirmed, attitude track determination module 21 also comprises: first predicting unit 220 and second predicting unit 221.
First predicting unit 220 is utilized preset prediction algorithm; Uniting the two or the three shaft angle degree information of deciding 217 outputs of appearance unit to first predicts; The two or three shaft angle degree information to the first of the prediction of output is united and is decided appearance unit 217, so that first unites and decide 217 pairs of appearance unit and calculate the two or the three shaft angle degree information that obtain and revise.Wherein, utilizing predicted value correction is exactly whether the numerical value that judge to calculate obtains is in a threshold range of predicted value, if then export the federal wave filter 219 of calculated value to the second, otherwise do not export any data.
Second predicting unit 221 is utilized preset prediction algorithm; Uniting the three or the three shaft angle degree information of deciding 218 outputs of appearance unit to second predicts; The three or three shaft angle degree information to the second of the prediction of output is united and is decided appearance unit 218, so that second unites and decide 218 pairs of appearance unit and calculate the three or the three shaft angle degree information that obtain and revise.Wherein, utilizing predicted value correction is exactly whether the numerical value that judge to calculate obtains is in a threshold range of predicted value, if then export the federal wave filter 219 of calculated value to the second, otherwise do not export any data.
The of the present invention second federal filter unit 219 comprises: synchronous subelement of three subtracters 2190, the first gain subelement 2191, the second gain subelement 2192, the 3rd gain subelement 2193, the very first time 2194, preservice subelement 2195 and optimal estimation subelement 2196.
The one or three shaft angle degree information, the two or three shaft angle degree information or the three or three shaft angle degree information that each subtracter in three subtracters 2190 will receive; Carry out subtraction with the satellite threeaxis attitude angle of optimal estimation subelement 2196 outputs, difference to the first gain subelement 2191, the second gain subelement 2192 or the 3rd gain subelement 2193 that obtains calculated in output.
The first gain subelement 2191 is according to first gain; Carry out filtering to the one or three shaft angle degree information with from the difference of the satellite threeaxis attitude angle of optimal estimation subelement 2196, export filtered synchronous subelement 2194 of the one or three shaft angle degree information to the very first time.
The second gain subelement 2191 is according to second gain; Carry out filtering to the two or three shaft angle degree information with from the difference of the satellite threeaxis attitude angle of optimal estimation subelement 2196, export filtered synchronous subelement 2194 of the two or three shaft angle degree information to the very first time.
The 3rd gain subelement 2193 is according to the 3rd gain; Carry out filtering to the three or three shaft angle degree information with from the difference of the satellite threeaxis attitude angle of optimal estimation subelement 2196, export filtered synchronous subelement 2194 of the three or three shaft angle degree information to the very first time.
Synchronous subelement 2194 of the very first time will receive filtered the one or three shaft angle degree information, the two or three shaft angle degree information and the three or three shaft angle degree information and export preservice subelement 2195 to; And according to the main filtering equations of federal wave filter; Calculate first error, second sum of errors the 3rd error; Utilize first error that first gain of the first gain subelement 2191 is set; Utilize second error that second gain of the second gain subelement 2192 is set, utilize the 3rd error that the 3rd gain of the 3rd gain subelement 2193 is set.
Preservice subelement 2195 carries out unrulyvalue rejecting according to preset threshold value to receiving filtered the one or three shaft angle degree information, the two or three shaft angle degree information and the three or three shaft angle degree information, exports the information after rejecting to optimal estimation subelement 2196.
Optimal estimation subelement 2196 carries out computing according to the optimal estimation algorithm to the information after rejecting, and obtains the threeaxis attitude angular velocity of satellite threeaxis attitude angle and satellite, and exports rail control computing machine 22 to.Optimal estimation subelement 2196 also utilizes the output result of the state equation of satellite orbit kinetic model, calculates the threeaxis attitude angular velocity that obtains satellite threeaxis attitude angle and satellite.
The of the present invention first federal filter unit comprises: information distribution subelement, first wave filter, second wave filter and senior filter.Wherein, the information distribution subelement connects senior filter, first wave filter and second wave filter; First wave filter connects spacetime uniformity unit 213 and Model Calculation unit 214; The second wave filter link model computing unit 214 and star sensor 202; Senior filter connects spacetime uniformity unit 213, first wave filter, second wave filter and information distribution subelement.
The information distribution subelement is that first subfilter and second subfilter are distributed weights according to the error evaluation result of information conservation principle and senior filter output.Wherein, weights are notions in the information fusion technology; Same category information with multiple source is gathered like the attitude angle that collects from star sensor 202, earth sensor 203 and inertia device 204, calculates a value that precision is higher through existing fusion, just belongs to information fusion; In this case, can not simply ask on average the information in each source, but will be according to error level separately; Letting not, the data of homology multiply by the coefficient that varies in size respectively; And then addition, this coefficient that varies in size is exactly weights, and weights are not fixed values.Senior filter can carry out error evaluation to the data of each provenance after each completion is calculated; The error evaluation result is sent to the information distribution subelement; The information distribution subelement is again according to the error evaluation result; Distribute big weights for the little source of error, distribute little weights for the big source of error.The data of each provenance are revised first result and revised second result that senior filter receives.
First subfilter is according to the weights that distribute, from the state equation of the satellite orbit kinetic model of Model Calculation unit 214, from the state equation of the GNSS range finding of Model Calculation unit 214, from the three dimensional local information and the pseudorange of spacetime uniformity unit 213; Carry out finding the solution of state equation; Obtain first result; Be used to predicted value, first result is revised, export revised first result to senior filter from senior filter.Wherein, the state equation of GNSS range finding can be set up according to the ranging information of the state equation and the GNSS receiver of satellite orbit kinetic model in Model Calculation unit 214.
Second subfilter is according to the weights that distribute, from the state equation of the satellite orbit kinetic model of Model Calculation unit 214, from the state equation of the star sensor angular observation of Model Calculation unit 214; Carry out finding the solution of state equation; Obtain second result; Be used to predicted value, second result is revised, export revised second result to senior filter from senior filter.Wherein, the state equation of star sensor angular observation can be set up according to the state equation of satellite orbit kinetic model and the angle information of star sensor 202 in Model Calculation unit 214.
Senior filter is according to three dimensional local information, revised first result and revised second result; Utilize preset prediction algorithm; Obtain predicted value and export first subfilter to and second subfilter; Utilize linear minimum variance algorithm computation to obtain the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system, and export rail control computing machine 22 to.
Ground subsystem of the present invention comprises: receiver module 26, data acquisition module 27, driver module 28 and time synchronized module 29; Wherein, abovementioned module realizes data interaction through the RTI interface.Data acquisition module 27 and driver module 28 can adopt the SOCKET passage of WINDOWS system through the ICP/IP protocol function that connects in order to the passage that carries out data transmission.
The wireless communication link of subsystem exports attitude of satellite information that receives and satelliteorbit information to data acquisition module 27 on receiver module 26 foundation and the star.Receiver module 26 can pass through the wireless data sending passage or gateway receives the data from subsystem on the star.
Data acquisition module 27 is resolved attitude of satellite information and satelliteorbit information; Obtain the positional information that relative velocity, satellite relative inertness are between relative position information, star between star, velocity information, the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite of satellite relative inertness system, and export driver module 28 to.
Driver module 28 utilizes picture Core Generator, the data that receive and information, generates satellite in rail flight picture and output.Wherein, Driver module 28 utilizes the CONNECT DLL Real Time Drive cartoon scene that the STK instrument provides to upgrade, with the form of windows such as close shot threedimensional picture, distant view threedimensional picture, twodimensional picture, kinetic parameter tabulation show satellite uptodate in the rail flight condition.
Time synchronized module 29 output times stab to data acquisition module 27 and driver module 28, the data that synchronous data collection module 27 and driver module 28 receive.Particularly; The initial sum steplength of time synchronized module 29 control driver modules 28; Guarantee that its data presented is current uptodate flight parameter; Take place to cause unusually under the situation of animation display hysteresis at network, control data acquisition module 27 abandons the data in the abnormal time, gathers uptodate data again and offers driver module 28.Every at a distance from some cycles, time synchronized module 29 is broadcasted current time with the form of timestamp to each module, and judges whether each module lags behind, as the foundation of carrying out synchronous processing.
In order to guarantee the fluency of display frame; Ground subsystem of the present invention also comprises a forecast module; The forecast module is between twice Data Receiving; According to the orbital data of last time and current environmental parameter, predict and export satellite orbital position to the driver module 28 in twice Data Receiving time interval, guarantee that picture is smooth uninterrupted; Time synchronized module 29 also further connects the forecast module, in twice data acquisition opening and closing of the orbit prediction function of inner control forecast module at interval.
Fig. 4 confirms the process flow diagram of method for attitude of satellite track of the present invention.Combine Fig. 4 at present, attitude of satellite track of the present invention is confirmed that method describes, specific as follows:
Step 401: subsystem is gathered satellite navigation signals, attitude of satellite angle signal and the satellite starlight elevation angle of multiple Navsat constellation on the star;
The attitude track signal acquisition module 20 of subsystem is gathered satellite navigation signals, attitude of satellite angle signal and the satellite starlight elevation angle that obtains multiple Navsat constellation on the star.
Step 402: subsystem calculates and generates the positional information that relative velocity, satellite relative inertness are between relative position information, star between star, velocity information, the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite of satellite relative inertness system on the star;
Wherein, subsystem carries out error correction and calculus of differences to pseudorange, pseudorange rates and carrier phase data on the star, relative velocity between relative position information and star between the acquisition star.Be specially:
Step 4021 is carried out that ephemeris error correction, ionosphere timedelay are proofreaied and correct, integer ambiguity is found the solution and is jumped in week and detect the pseudorange of multiple Navsat constellation and carrier phase data, obtains the pseudorange and the carrier phase data of revised multiple Navsat constellation; Step 4022 is utilized the model of relative orbit key element, and the pseudorange and the carrier phase data of revised multiple Navsat constellation are carried out level and smooth and calculus of differences, obtains between star relative velocity between relative position information and star.
Wherein, the ionosphere timedelay is revised and can be adopted the model revised law, and model adopts singlelayer model, and acquiescence ionosphere is distributed in apart from ground H
_{Iono}Infinitely thin layer on.Electromagnetic wave during through ionosphere since the delay of the travel path that the rate of change of refraction causes be:
Make N
_{∑}=∫ Nds is the total electron amount on the electromagnetic wave propagation path, and then its phase delay is:
Time delay is:
The earth's core radius vector angle of supposing travel path and receiver is θ, and with the integration that changes the radius vector along the earth's core in (1) along the integration of travel path into, the characteristic according to singlelayer model has:
Wherein, θ ' is the zenith distance of travel path and ionosphere intersection point place observation satellite, satisfies
In the abovementioned singlelayer model, the cosine function in the LZT that daytime, total amount of electrons can being expressed as in ionosphere is approximately a constant night, that is:
Wherein, integer ambiguity is found the solution and is specially: set up all two eikonal equation groups of institute's observation satellite in n epoch, that is:
A in the formula
_{n}Be two eikonal equation group matrix of coefficients of n epoch; X
_{n}It is the threedimensional location coordinates vector that n obtains epoch; N is the integer ambiguity vector of carrier phase; E is a unit matrix.
In order to reduce computational complexity, need following formula is simplified.To coefficient matrices A
_{i}Carry out QR and decompose A
_{i}=Q (i) R (i), the transposition conversion done in the submatrix of the Q (i) that decomposition is obtained, and premultiplication formula (7) obtains:
[Q(i)
^{2}]
^{T}·L
_{i}＝[Q(i)
^{2}]
^{T}·λN+[Q(i)
^{2}]
^{T}·ε
_{i} (8)
And then the floatingpoint that obtains blur level separate into:
Its corresponding association factor is:
On abovementioned basis of separating, utilize LAMBDA that blur level is searched for, step is following: variance and covariance battle array transform (falling correlating transforms); Floatingpoint is separated integer decomposition part and fraction part; Floatingpoint is separated fraction part and is carried out conversion; The candidate solution number that search provides is set; Calculate the size of hunting zone; Finding the solution floatingpoint separates fraction part and carries out static solution corresponding behind the transform; Find the solution the static solution that floatingpoint is separated.Can adopt existing searching method to carry out the blur level search, no longer concrete grammar given unnecessary details at this.
Wherein, week jumps detection method and is specially: adopt doppler shift method, promptly the carrier phase rate of change carries out jumping in week and detects.Set up following multinomial model about carrier phase:
Wherein, the equality that does not comprise Δ N in the formula (12) is the computing formula before jumping in week, and the equality that comprises Δ N is the computing formula after jumping in week.
Choose carrier phase observation data and the rate of change thereof of 5 epoch:
suppose that the carrier phase value of preceding 4 epoch does not have week to jump; Be used for surveying the carrier phase of the 5th epoch and whether take place to jump in week, set up following error equation: F=AX+v (12)
Wherein, X=[a
_{0}, a
_{1}, a
_{2}, a
_{3}, Δ N]
^{T},
According to leastsquares algorithm, X=(A
^{T}A)
^{1}A
^{T}F confirms a threshold epsilon, if  Δ N＞ε, explain that then the carrier phase observation data of the 5th epoch exists week to jump, it is Δ N that week is jumped valuation.
Step 4021 adopts said method, pseudorange and carrier phase data is revised, to improve the precision of orbit determination.
The method for building up of the model of the relative orbit key element in the step 4022 is following:
At first; Definition relative position and relative velocity are the relative orbit state vector; The orbital elements that are designated as
primary are designated as e (t), and through the Taylor series expansion, the orbital elements that can obtain primary and target star are poor; Be the relative orbit key element, be designated as δ e (t).Obtain the relation between relative orbit state vector and the relative orbit key element through coordinate conversion matrix.
The angular velocity of primary is expressed as in orbital coordinate system:
Ω＝Ω
_{r}e
_{x}+Ω
_{t}e
_{y}+Ω
_{n}e
_{z} (13)
Wherein, Ω r, Ω t, Ω n be respectively radial angle speed, the speed that cuts angle, secondary normal angle speed.
The position of primary and target star and the speed component array in the primary orbital coordinate system is:
(r
_{c})
_{o}＝Re
_{x}
(r
_{a})
_{o}＝(r
_{c}+r)
_{o}＝(R+x)e
_{x}+ye
_{y}+ze
_{z}
Wherein, rc and ra represent the earth's core of primary and target star respectively apart from vector, and Vc and Va represent the velocity of primary and target star respectively; R representes the position vector of the relative primary of target star; R representes the earth's core distance of central satellite, and Vr, Vt and Vn represent radial velocity, tangential velocity and secondary normal velocity, x respectively; Y; Z, the component array of
expression relative position and relative velocity, i.e. state vector.
Make the Mean Orbit Elements of target star and primary poor; Be that relative Mean Orbit Elements
is the system state vector, then Equation of Relative Motion with Small does
Wherein,
is the transition matrix of relative Mean Orbit Elements; Relative distance (r), position angle
and the angle of pitch (θ) are as measured value between the employing star; Instantaneous relative position (x between measured value and star; Y, relation z) is following:
According to the relation of instantaneous relative position between star with relative Mean Orbit Elements, can obtain the relation between measured value and the state vector, also just draw the measurement equation, and then obtained the model of relative orbit key element, write as following form:
Wherein, ω (t) and v (t) are respectively system noise and measure noise; ∑ (t) is the transition matrix of Mean Orbit Elements to motion state, and D (t) is the transition matrix of Mean Orbit Elements to the instantaneous orbit key element.Can find out that from the model of relative orbit key element state vector is relative Mean Orbit Elements, corresponding the control of keeping the formation configuration and implementing; Measured value is relative distance and orientation between star; Transient measurement in corresponding the orbit determination relatively; Therefore; This model can be united with the relative instantaneous orbit state that measures controlling required relative mean orbit state, has solved the contradiction of under Perturbation Effect, controlling and measuring in the relative orbit determination of satellite formation flying.
Calculus of differences in the step 4022 is pseudorange and carrier phase is carried out smoothing processing.Wherein, adopt the carrier phase smoothing pseudo range, smoothing method is:
ρ wherein
_{S, k}Be k moment carrier phase smoothing pseudo range, M is the smoothingtime constant, gets 20～100.
Wherein, Subsystem utilizes least square optimal estimation algorithm and spacetime to unify algorithm on the star; Advance to generate three dimensional local information according to pseudorange, pseudorange rates and carrier phase data; Utilize flight dynamics model and federal filtering algorithm, the three dimensional local information and the satellite starlight elevation angle are carried out federal filtering algorithm, generate the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system.Be specially:
Step 4023 according to pseudorange, pseudorange rates and the carrier phase data of multiple Navsat constellation, utilizes seven parameter Bursa conversion formulas to accomplish the unification of coordinate system; Step 4024, with the time unification of pseudorange, pseudorange rates and the carrier phase data of multiple Navsat constellation to Coordinated Universal Time(UTC) UTC; Step 4025 is utilized integrated positioning equation and least square method successive iteration, calculates to obtain three dimensional local information; Step 4026 is utilized flight dynamics model and federal filtering algorithm, and the three dimensional local information and the satellite starlight elevation angle are carried out federal filtering algorithm, generates the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system.
The unified method of wherein, carrying out coordinate system in the step 4023 is:
It is following to set up observation equation constantly in epoch:
Wherein, (x, y, z)
^{T}The three dimensional local information of trying to achieve for needs,
Be the three dimensional local information of i gps satellite under the WGS84 coordinate system,
Be the three dimensional local information of j bigdipper satellite under the CGS2000 coordinate system, c Δ t
_{GPS}, c Δ t
_{COMPASS}Be respectively the distance error that clock jitter causes, the actual observed range of parametric representation in equality left side.
Because the coordinate system disunity need utilize seven parameter Bursa conversion formulas to accomplish the origin translation and the coordinate axis rotation of rectangular coordinate system, realizes the unification of coordinate system.Conversion formula is following:
In the step 4024, though with the time unification of two kinds of signals to the UTC Coordinated Universal Time(UTC), also have deviation between the clock frequency of each system; The transmission of two kinds of receivers simultaneously also has timedelay deviation; Therefore, in the step 4025, integrated positioning algorithm capable of using and least square iterative algorithm; Find the solution three dimensional local information, specific as follows:
Through the spacetime uniformity processing, observation equation can be unified into following form:
The required unknown quantity that resolves comprises that three dimensional local information and GPS, COMPASS connect the clock jitter of a system, and therefore totally 5 unknown quantitys, realize that final resolving must observe the nautical star more than 5 simultaneously.
Suppose to observe m gps satellite and n bigdipper satellite, m and n are natural number, and then the integrated positioning equation can be described as following form:
AX＝B (21)
Wherein,
$A=\left[\begin{array}{cccccc}{a}_{x}^{1}& {a}_{y}^{1}& {a}_{z}^{1}& 1& 0& 0\\ & & L& L& & \\ {a}_{x}^{m}& {a}_{y}^{m}& {a}_{z}^{m}& 1& 0& 0\\ {a}_{x}^{m+1}& {a}_{x}^{m+1}& {a}_{x}^{m+1}& 0& 1& 0\\ & & L& L& & \\ {a}_{x}^{m+n}& {a}_{x}^{m+n}& {a}_{x}^{m+n}& 0& 1& 0\end{array}\right],$ ${a}_{x}^{j}=\frac{x{X}_{i}^{j}}{{R}_{i}^{j}},$ ${a}_{y}^{j}=\frac{y{Y}_{i}^{j}}{{R}_{i}^{j}},$ ${a}_{x}^{j}=\frac{z{Z}_{i}^{j}}{{R}_{i}^{j}},$
X＝(Δx，Δy，Δz，cΔt
_{GPS}，cΔt
_{COMPASS})
^{T}，
$B=\left[\begin{array}{c}{\mathrm{\ρ}}_{\mathrm{GPS}}^{1}{R}_{\mathrm{GPS}}^{1}\\ L\\ {\mathrm{\ρ}}_{\mathrm{GPS}}^{m}{R}_{\mathrm{GPS}}^{m}\\ {\mathrm{\ρ}}_{\mathrm{COMPASS}}^{m+1}{R}_{\mathrm{GPS}}^{m+1}\\ L\\ {\mathrm{\ρ}}_{\mathrm{COMPASS}}^{m+n}{R}_{\mathrm{COMPASS}}^{m+n}\end{array}\right]$
Utilize the least square method successive iteration, less than predefined scope, can obtain current three dimensional local information X=(A until the margin of error
^{T}A)
^{1}A
^{T}B.
Carrying out federal filtering algorithm in the step 4026 is specially:
Step 40261 is set up satellite orbit kinetic model state equation, is specially:
Wherein,
$X\left(t\right)={(x,y,z,\stackrel{\·}{x},\stackrel{\·}{y},\stackrel{\·}{z})}^{T};$ W (t) is a statenoise; 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)];
${f}_{x}(X\left(t\right),t)=\mathrm{\μ}\frac{x}{{r}^{3}}+\frac{\∂R}{\∂x}+{f}_{\mathrm{Dx}};$ ${f}_{y}(X\left(t\right),t)=\mathrm{\μ}\frac{y}{{r}^{3}}+\frac{\∂R}{\∂y}+{f}_{\mathrm{Dy}};$ ${f}_{z}(X\left(t\right),t)=\mathrm{\μ}\frac{z}{{r}^{3}}+\frac{\∂R}{\∂z}+{f}_{\mathrm{Dz}};$
R is a perturbation of earths gravitational field power, is specially:
Step 40262 is set up GNSS distance measuring states equation, is specially:
The ranging information that i satellite provides is:
Following formula is expressed as state equation:
S
_{1}(t
_{k})＝f
_{2}(X(t
_{k})，t
_{k})+V
_{1}(t
_{k}) (24)
Step 40263 is set up the state equation of star sensor angular observation information, is specially:
The local horizon direction that fixed star direction that records through star sensor and earth sensor record can obtain the starlight elevation angle of current time satellite; According to abovementioned model, can get:
Following formula is expressed as state equation:
S
_{2}(t
_{k})＝f
_{3}(X(t
_{k})，t
_{k})+V
_{1}(t
_{k}) (26)
Step 40264 is set up federal Kalman Filtering Model,
Be specially: set up federal Kalman Filtering Model according to formula (22), (24), (26); Distribute in order to carry out the weights of filtering according to the information conservation principle; According to the state equation of the weights that distribute, satellite orbit kinetic model, state equation, three dimensional local information and the pseudorange of GNSS range finding; Carry out finding the solution of state equation; Obtain first result; The predicted value of utilizing previous moment to calculate is revised first result, obtains current revised first result; According to the weights that distribute, the state equation of satellite orbit kinetic model, the state equation of star sensor angular observation; Carry out finding the solution of state equation, obtain second result, the predicted value of utilizing previous moment to calculate; Second result is revised, obtain current revised second result; According to three dimensional local information, revised first result and revised second result, utilize preset prediction algorithm, obtain next predicted value constantly; Utilize revised first checkout result, second result of calculation and the three dimensional local information of linear minimum variance algorithm and current time, calculate the positional information of acquisition satellite relative inertness system and the velocity information of satellite relative inertness system.
Wherein, system state noise and observation noise are mutual incoherent zeromean white noises.
Wherein, subsystem utilizes inertia device to decide appearance signal, attitude of satellite angle information and the satellite starlight elevation angle on the star, unites and decides appearance; Generate three shaft angle degree; Utilize federal filtering algorithm, three shaft angle degree are carried out optimal estimation, obtain the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite.Be specially:
Step 4027, to inertia device output attitude of satellite angle signal carry out the speed integration, obtain the one or three shaft angle degree information; Step 4028, the satellite starlight elevation angle and the hypercomplex number attitude angle of utilizing star sensor to export are proofreaied and correct the error of the one or three shaft angle degree information, obtain the two or three shaft angle degree information; Step 4029, the attitude of satellite angle information that utilizes earth sensor to export is proofreaied and correct the error of the one or three shaft angle degree information, obtains the three or three shaft angle degree information.
Step 403: subsystem utilizes the positional information of relative velocity between relative position information between star, star, satellite relative inertness system, velocity information, the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite of satellite relative inertness system, the track of adjustment satellite and the attitude of satellite on the star.Specifically can adopt existing attitude track control algolithm, the track of adjustment satellite and the attitude of satellite are no longer given unnecessary details this method at this.
Can further comprise after the step 403: utilize the positional information of relative velocity between relative position information between star, star, satellite relative inertness system, velocity information, the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite of satellite relative inertness system, generate satellite in rail flight picture and output.
In the abovementioned preferred embodiment of the present invention; Rely on the problem that the GNSS receiver positions the bearing accuracy decline that is caused fully in order to remedy; The present invention orbit determination with decide appearance when calculating; Introduced the information that inertia device, star sensor and earth sensor are gathered,, realized the associating independent navigation with the assisted GNSS receiver; Changed is that main satellite decide appearance and orbit determination mode with ground in the past, makes realtime high precision of can be under the situation of the disengaging land station autonomous realization of satellite and highly reliable orbit determination and decides appearance; When precise position information being provided for satellite completion space tasks; Greatly reduce the operating cost of ground subsystem; Alleviated the task burden of ground subsystem; Improved the disguise and the security of satellite task, promptly still can keep Satellite Orbit Determination in the ground subsystem cause clogging even when being destroyed and decide the normal operation of appearance.
In sum, be preferred embodiment of the present invention more than, be not to be used for limiting protection scope of the present invention.All within spirit of the present invention and principle, any modification of being done, be equal to replacement, improvement etc., all should be included within protection scope of the present invention.
Claims (13)
1. an attitude of satellite track is confirmed system, and this system comprises subsystem on the star, it is characterized in that, subsystem comprises on the said star:
Attitude track signal acquisition module, satellite navigation signals, attitude of satellite angle signal and the satellite starlight elevation angle of gathering multiple Navsat constellation, output satellite navigation signals, attitude of satellite angle signal and the satellite starlight elevation angle to attitude track determination module;
Attitude track determination module; From satellite navigation signals, obtain pseudorange, pseudorange rates and carrier phase data; Pseudorange, pseudorange rates and carrier phase data are carried out error correction and calculus of differences, relative velocity between relative position information and star between the acquisition star, and export the rail control computing machine to;
Said attitude track determination module utilizes least square optimal estimation algorithm and spacetime to unify algorithm; Advance to generate three dimensional local information according to pseudorange, pseudorange rates and carrier phase data; Utilize flight dynamics model and federal filtering algorithm; The three dimensional local information and the satellite starlight elevation angle are carried out federal filtering algorithm, generate the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system, and export the rail control computing machine to;
Said attitude track determination module utilizes satellite navigation signals, attitude of satellite angle information and the satellite starlight elevation angle; Unite and decide appearance; Generate three shaft angle degree, utilize federal filtering algorithm, three shaft angle degree are carried out optimal estimation; Obtain the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite, and export the rail control computing machine to;
Said rail control computer based is in the positional information of relative velocity between relative position information, star, satellite relative inertness system between the star that receives and the velocity information of satellite relative inertness system; Output becomes rail and instructs to Executive Module; The threeaxis attitude angular speed of satellitebased threeaxis attitude angle and satellite, the output posture adjustment is instructed to Executive Module;
Execution module is according to the track that becomes rail instruction Adjustment System place satellite, according to the attitude of posture adjustment instruction Adjustment System place satellite.
2. system according to claim 1 is characterized in that, subsystem also comprises on the said star:
Communication module between star, in order to set up and the intersatellite communication link of cooperating, relative velocity between relative position information and star between the transmission star receives positional information and its velocity information of another satellite and is transmitted to attitude track determination module;
Said attitude track determination module is further according to the positional information of another satellite and its velocity information, calculates to obtain between star relative velocity between relative position information and star, and exports communication module between star to.
3. system according to claim 1 is characterized in that, this system also comprises:
Ground subsystem is utilized the attitude of satellite information and the satelliteorbit information that receive through wireless communication link, generates satellite in rail flight picture and output;
The rail control computer of subsystem generates attitude information and the satelliteorbit information and the output of satellite further based on becoming rail instruction and posture adjustment instruction on the said star;
Subsystem further comprises on the said star:
The satellite platform bus is set up the data transmission channel between rail control computing machine and telemetry module;
Telemetry module through platform bus, receives attitude of satellite information and satelliteorbit information from the rail control computing machine, and through Radio Link output satellite attitude signal and satellite orbit signal.
4. according to claim 1,2 or 3 described systems, it is characterized in that said attitude track signal acquisition module comprises:
Multimodal satellite navigation GNSS receiver is gathered the satellite navigation information of multiple Navsat constellation, and exports attitude track determination module to;
Inertia device is gathered the attitude of satellite angle signal of satellite, and is exported attitude track determination module to;
Star sensor is gathered the satellite starlight elevation angle and the hypercomplex number attitude angle of satellite, and is exported attitude track determination module to;
Earth sensor is gathered the attitude of satellite angle information of satellite, and is exported attitude track determination module to;
Said multiple Navsat constellation comprises bigdipper satellite and global position system GPS satellite at least.
5. system according to claim 4 is characterized in that, said attitude track determination module comprises:
Data generating unit, the satellite navigation information according to the multiple Navsat constellation of GNSS receiver output generates pseudorange, pseudorange rates and the carrier phase data of multiple Navsat constellation, and exports amending unit and spacetime uniformity unit to;
Amending unit; The pseudorange that receives multiple aeronautical satellite constellation and carrier phase data are carried out ephemeris error correction, ionosphere timedelay are proofreaied and correct, integer ambiguity is found the solution and cycle slip detects, export the pseudorange and the carrier phase data of revised multiple aeronautical satellite constellation to the calculus of differences unit;
The calculus of differences unit utilizes the model of relative orbit key element, and the pseudorange and the carrier phase data of revised multiple Navsat constellation are carried out calculus of differences, obtains between star relative velocity between relative position information and star, exports the rail control computing machine to;
The spacetime uniformity unit; Pseudorange, pseudorange rates and carrier phase data according to multiple Navsat constellation; Utilize seven parameter Bursa conversion formulas to accomplish the unification of coordinate system, the time unification of pseudorange, pseudorange rates and the carrier phase data of multiple Navsat constellation to Coordinated Universal Time(UTC) UTC, is utilized integrated positioning equation and least square method successive iteration; Calculate and obtain three dimensional local information, export the first federal filter unit to;
The Model Calculation unit generates the state equation of satellite orbit kinetic model, and exports the first federal filter unit to;
The first federal filter unit; Utilize three dimensional local information, attitude of satellite angle information, the state equation of satellite orbit kinetic model, the state equation of GNSS range finding and the state equation of star sensor angular observation from star sensor; Carry out finding the solution of state equation; Obtain the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system; Utilize linear minimum variance algorithm and prediction algorithm, the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system are carried out filtering, and export the rail control computing machine to;
The speed integral unit carries out the speed integration to the attitude of satellite angle signal of inertia device output, obtains the one or three shaft angle degree information, and exports the second federal filter unit, first to and unite and decide appearance unit and second and unite and decide the appearance unit;
First unites and decides the appearance unit, utilizes the satellite starlight elevation angle and the hypercomplex number attitude angle of star sensor output, and the error of the one or three shaft angle degree information of speed integral unit output is proofreaied and correct, and obtains the two or three shaft angle degree information and exports the second federal wave filter to;
Second unites and decides the appearance unit, utilizes the attitude of satellite angle information of earth sensor output, and the error of the one or three shaft angle degree information of speed integral unit output is proofreaied and correct, and obtains the three or three shaft angle degree information and exports the second federal wave filter to;
The second federal filter unit; Utilize the main filtering equations of federal filtering algorithm that first gain, second gain and the 3rd gain are adjusted; Utilize first gain that the one or three shaft angle degree information is carried out filtering, utilize second gain that the two or three shaft angle degree information is carried out filtering, utilize the 3rd gain that the three or three shaft angle degree information is carried out filtering; Filtered three information are carried out unrulyvalue rejecting; Information after rejecting is carried out the optimal estimation computing, obtain the threeaxis attitude angular velocity of satellite threeaxis attitude angle and satellite, and export the rail control computing machine to.
6. system according to claim 5 is characterized in that, said attitude track determination module also comprises:
First predicting unit; Utilize preset prediction algorithm; Uniting the two or the three shaft angle degree information of deciding the output of appearance unit to first predicts; The two or three shaft angle degree information to the first of the prediction of output is united and is decided the appearance unit, so that first unites and decide the appearance unit and revise calculating the two or the three shaft angle degree information that obtains;
Second predicting unit; Utilize preset prediction algorithm; Uniting the three or the three shaft angle degree information of deciding the output of appearance unit to second predicts; The three or three shaft angle degree information to the second of the prediction of output is united and is decided the appearance unit, so that second unites and decide the appearance unit and revise calculating the three or the three shaft angle degree information that obtains.
7. system according to claim 6 is characterized in that, the said second federal filter unit comprises:
Three subtracters; The one or three shaft angle degree information, the two or three shaft angle degree information or the three or three shaft angle degree information that each subtracter will receive; Carry out subtraction with the satellite threeaxis attitude angle of optimal estimation subelement output, difference to the first gain subelement, the second gain subelement or the 3rd gain subelement that obtains calculated in output;
The first gain subelement according to first gain, carries out filtering to the one or three shaft angle degree information with from the difference of the satellite threeaxis attitude angle of optimal estimation subelement, exports filtered synchronous subelement of the one or three shaft angle degree information to the very first time;
The second gain subelement according to second gain, carries out filtering to the two or three shaft angle degree information with from the difference of the satellite threeaxis attitude angle of optimal estimation subelement, exports filtered synchronous subelement of the two or three shaft angle degree information to the very first time;
The 3rd gain subelement according to the 3rd gain, carries out filtering to the three or three shaft angle degree information with from the difference of the satellite threeaxis attitude angle of optimal estimation subelement, exports filtered synchronous subelement of the three or three shaft angle degree information to the very first time;
Synchronous subelement of the very first time; To receive filtered the one or three shaft angle degree information, the two or three shaft angle degree information and the three or three shaft angle degree information and export the preliminary treatment subelement to; And according to the main filtering equations of federal wave filter; Calculate first error, second sum of errors the 3rd error; Utilize first error that first gain of the first gain subelement is set; Utilize second error that second gain of the second gain subelement is set, utilize the 3rd error that the 3rd gain of the 3rd gain subelement is set;
The preservice subelement according to preset threshold value, carries out unrulyvalue rejecting to receiving filtered the one or three shaft angle degree information, the two or three shaft angle degree information and the three or three shaft angle degree information, exports the information after rejecting to the optimal estimation subelement;
The optimal estimation subelement according to the optimal estimation algorithm, carries out computing to the information after rejecting, and obtains the threeaxis attitude angular velocity of satellite threeaxis attitude angle and satellite, and exports the rail control computing machine to.
8. system according to claim 6 is characterized in that, the said first federal filter unit comprises:
The information distribution subelement according to the error evaluation result of information conservation principle and senior filter output, is that first subfilter and second subfilter are distributed weights;
First subfilter; According to the state equation of the weights that distribute, satellite orbit kinetic model, state equation, three dimensional local information and the pseudorange of GNSS range finding; Carry out finding the solution of state equation, obtain first result, be used to predicted value from senior filter; First result is revised, export revised first result to senior filter;
Second subfilter; According to the weights that distribute, the state equation of satellite orbit kinetic model, the state equation of star sensor angular observation; Carry out finding the solution of state equation, obtain second result, be used to predicted value from senior filter; Second result is revised, export revised second result to senior filter;
Senior filter; According to three dimensional local information, revised first result and revised second result; Utilize preset prediction algorithm; Obtain predicted value and export first subfilter to and second subfilter, revised first result and revised second result are carried out error evaluation, output error assessment result to information distribution subelement; Utilize linear minimum variance algorithm computation to obtain the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system, and export the rail control computing machine to.
9. system according to claim 3 is characterized in that, said ground subsystem comprises:
Receiver module, the wireless communication link of subsystem exports the attitude of satellite information and the satelliteorbit information that receive to data acquisition module on foundation and the star;
Data acquisition module; Resolve attitude of satellite information and satelliteorbit information; Obtain the positional information that relative velocity, satellite relative inertness are between relative position information, star between star, velocity information, the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite of satellite relative inertness system, and export driver module to;
Driver module utilizes picture Core Generator, the data that receive and information, generates satellite in rail flight picture and output;
Time synchronized module, output time are stabbed to data acquisition module and driver module, the data that synchronous data collection module and driver module receive.
10. an attitude of satellite track is confirmed method, it is characterized in that, this method comprises:
Subsystem is gathered satellite navigation signals, attitude of satellite angle signal and the satellite starlight elevation angle of multiple Navsat constellation on A, the star;
Subsystem obtains pseudorange, pseudorange rates and carrier phase data on B, the star from satellite navigation signals, and pseudorange, pseudorange rates and carrier phase data are carried out error correction and calculus of differences, relative velocity between relative position information and star between the acquisition star;
Subsystem utilizes least square optimal estimation algorithm and spacetime to unify algorithm on the star; Advance to generate three dimensional local information based on pseudorange, pseudorange rates and carrier phase data; Utilize flight dynamics model and federal filtering algorithm; The three dimensional local information and the satellite starlight elevation angle are carried out federal filtering algorithm, generate the positional information of satellite relative inertness system and the velocity information of satellite relative inertness system;
Subsystem utilizes satellite navigation signals, attitude of satellite angle information and the satellite starlight elevation angle on the star; Unite and decide appearance, generate three shaft angle degree, utilize federal filtering algorithm; Three shaft angle degree are carried out optimal estimation, obtain the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite;
Subsystem utilizes the positional information of relative velocity between relative position information between star, star, satellite relative inertness system, velocity information, the threeaxis attitude angle of satellite and the threeaxis attitude angular velocity of satellite of satellite relative inertness system, the track of adjustment subsatellite and the attitude of satellite on C, the star.
11. method according to claim 10 is characterized in that, step B is said to carry out error correction and calculus of differences comprises to pseudorange, pseudorange rates and carrier phase data:
B1, the pseudorange of multiple Navsat constellation and carrier phase data are carried out that ephemeris error correction, ionosphere timedelay are proofreaied and correct, integer ambiguity is found the solution and jumped in week and detect, obtain the pseudorange and the carrier phase data of revised multiple Navsat constellation;
B2, utilize the model of relative orbit key element, the pseudorange and the carrier phase data of revised multiple Navsat constellation are carried out calculus of differences, obtain between star relative velocity between relative position information and star.
12. method according to claim 10 is characterized in that, said least square optimal estimation algorithm and the spacetime of utilizing of step B is unified algorithm, generates three dimensional local information according to pseudorange, pseudorange rates and carrier phase data and comprises:
B3, the pseudorange according to multiple Navsat constellation, pseudorange rates and carrier phase data utilize seven parameter Bursa conversion formulas to accomplish the unification of coordinate system;
The time unification of B4, the pseudorange with multiple Navsat constellation, pseudorange rates and carrier phase data is to Coordinated Universal Time(UTC) UTC;
B5, utilize integrated positioning equation and least square method successive iteration, calculate and obtain three dimensional local information.
13. method according to claim 10 is characterized in that, said satellite navigation signals, attitude of satellite angle information and the satellite starlight elevation angle of utilizing of step B united and decided appearance, generates three shaft angle degree and comprises:
B6, the attitude of satellite angle signal that inertia device is exported carry out the speed integration, obtain the one or three shaft angle degree information;
B7, the satellite starlight elevation angle and the hypercomplex number attitude angle of utilizing star sensor to export are proofreaied and correct the error of the one or three shaft angle degree information, obtain the two or three shaft angle degree information;
B8, the attitude of satellite angle information that utilizes earth sensor to export are proofreaied and correct the error of the one or three shaft angle degree information, obtain the three or three shaft angle degree information.
Priority Applications (1)
Application Number  Priority Date  Filing Date  Title 

CN201110315103.3A CN102494686B (en)  20111017  A kind of satellite attitude orbit determines system and method 
Applications Claiming Priority (1)
Application Number  Priority Date  Filing Date  Title 

CN201110315103.3A CN102494686B (en)  20111017  A kind of satellite attitude orbit determines system and method 
Publications (2)
Publication Number  Publication Date 

CN102494686A true CN102494686A (en)  20120613 
CN102494686B CN102494686B (en)  20161214 
Family
ID=
Cited By (27)
Publication number  Priority date  Publication date  Assignee  Title 

CN103116361A (en) *  20130221  20130522  北京控制工程研究所  Method for determining orbital transfer interval under control of satellite momentum wheel 
CN103217982A (en) *  20130221  20130724  北京控制工程研究所  Orbit control method based on wheelcontrolled mode 
CN103398726A (en) *  20130813  20131120  中国科学院长春光学精密机械与物理研究所  Judgment method of validity of least square attitude solution of star sensor 
CN103646127A (en) *  20131120  20140319  中国空间技术研究院  Satellite orbit gesture visual threedimensional displaying method 
CN105182987A (en) *  20150818  20151223  北京航天长征飞行器研究所  Pose correction method for powered phase of aircraft 
CN105246781A (en) *  20130529  20160113  雷斯昂公司  Satellite orbital determination (OD) using doppler and kepler orbital elements 
CN105510936A (en) *  20141126  20160420  航天恒星科技有限公司  Satelliteborne GNSS combined orbit determination method 
CN105698764A (en) *  20160130  20160622  武汉大学  Error modeling compensation method and system of optical remote sensing satellite image timevarying system 
CN105799954A (en) *  20141231  20160727  上海新跃仪表厂  Spacebased modular aircraft for conducting decentralized deployment on micronano load and orbital transfer guidance method of modular aircraft 
CN105866808A (en) *  20160621  20160817  上海航天控制技术研究所  Method for confirming influence of orbit determination errors of navigation receiver to satellite attitude precision 
CN106092099A (en) *  20160602  20161109  哈尔滨工业大学  Spacecraft is relative to positional increment orbit determination method 
CN107064981A (en) *  20170410  20170818  千寻位置网络有限公司  Differential positioning method and system based on GNSS, service terminal 
CN107101649A (en) *  20170525  20170829  北京航天自动控制研究所  A kind of inorbit error separating method of spacecraft Guidance instrumentation 
CN108009462A (en) *  20161031  20180508  中南大学  It is a kind of to be applied to rail inspection filtering method of the basic string rail of instrument to data 
CN108512590A (en) *  20180323  20180907  中国空间技术研究院  A kind of jointtrial system and method for satellite attitude and orbit control subsystem and GNSS subsystems 
CN109444931A (en) *  20181008  20190308  闽江学院  A kind of method and device of static state pseudorange OnePoint Location 
CN109625331A (en) *  20181226  20190416  上海微小卫星工程中心  Satellite controller and satellite control method 
CN109709588A (en) *  20181211  20190503  中国人民解放军63921部队  A kind of more star highprecision measuring rail systems of high rail satellite 
CN109856995A (en) *  20190304  20190607  北京空间飞行器总体设计部  A kind of whole star control subsystem analog platform towards test method verifying assessment 
CN109901600A (en) *  20190308  20190618  宁波天擎航天科技有限公司  A kind of spacecraft flight control method, system and device 
CN110516379A (en) *  20190829  20191129  珠海迈越信息技术有限公司  A method of satellite signal receiving apparatus is installed based on augmented reality 
CN110763231A (en) *  20191015  20200207  哈尔滨工程大学  Errorfree attitude updating method suitable for fiber optic gyroscope filtering signal 
CN111487657A (en) *  20200321  20200804  哈尔滨工程大学  Beidou realtime precise orbit determination method based on satellite perturbation 
CN111645882A (en) *  20200604  20200911  北京航天方舟空间技术有限公司  Satellite autonomous orbit determination method, device, equipment and computer storage medium 
CN111854764A (en) *  20200720  20201030  中国科学院微小卫星创新研究院  Spacecraft attitude determination method and system based on intersatellite measurement information 
CN113504728A (en) *  20210722  20211015  北京微纳星空科技有限公司  Method, device and equipment for generating task instruction and storage medium 
CN115098828A (en) *  20220826  20220923  北京控制工程研究所  Method and device for calculating loworbit satellite orbit in near circle 
Citations (8)
Publication number  Priority date  Publication date  Assignee  Title 

JPH02141400A (en) *  19881121  19900530  Nippon Telegr & Teleph Corp <Ntt>  Orientation control system for satellite 
JP2004210032A (en) *  20021227  20040729  Mitsubishi Electric Corp  Formation flying satellite 
CN101093397A (en) *  20060623  20071226  航天东方红卫星有限公司  System for controlling satellite attitude and track based on network on satellites 
CN101556155A (en) *  20090520  20091014  上海微小卫星工程中心  Small satellite attitude determination system and method thereof 
CN101554926A (en) *  20090520  20091014  上海微小卫星工程中心  Attitude control system for space vehicle and method thereof 
US20100066599A1 (en) *  20080915  20100318  Sony Ericsson Mobile Communications Ab  System and method of transferring location assistance information between electronic devices 
CN102114918A (en) *  20101231  20110706  北京航空航天大学  Attitude control feedback loop based on combined fixed attitude of multirate sensor 
CN102176037A (en) *  20101224  20110907  航天恒星科技有限公司  Cofrequency multisystem navigation signal receiving and processing method 
Patent Citations (8)
Publication number  Priority date  Publication date  Assignee  Title 

JPH02141400A (en) *  19881121  19900530  Nippon Telegr & Teleph Corp <Ntt>  Orientation control system for satellite 
JP2004210032A (en) *  20021227  20040729  Mitsubishi Electric Corp  Formation flying satellite 
CN101093397A (en) *  20060623  20071226  航天东方红卫星有限公司  System for controlling satellite attitude and track based on network on satellites 
US20100066599A1 (en) *  20080915  20100318  Sony Ericsson Mobile Communications Ab  System and method of transferring location assistance information between electronic devices 
CN101556155A (en) *  20090520  20091014  上海微小卫星工程中心  Small satellite attitude determination system and method thereof 
CN101554926A (en) *  20090520  20091014  上海微小卫星工程中心  Attitude control system for space vehicle and method thereof 
CN102176037A (en) *  20101224  20110907  航天恒星科技有限公司  Cofrequency multisystem navigation signal receiving and processing method 
CN102114918A (en) *  20101231  20110706  北京航空航天大学  Attitude control feedback loop based on combined fixed attitude of multirate sensor 
NonPatent Citations (8)
Title 

JAEIK PARK: "Hardwareintheloop simulations of GPSbased navigation and control for satellite formation flying", 《ADVANCES IN SPACE RESEARCH》 * 
吴显兵: "星载GPS低轨卫星几何法定轨及动力学平滑方法研究", 《中国优秀硕士学位论文全文数据库 基础科学辑》 * 
潘旺华: "基于多传感器信息融合的卫星姿态确定技术研究", 《中国优秀硕士学位论文全文数据库 工程科技II辑》 * 
王军武: "基于GPS的微小卫星定姿及定轨研究", 《中国优秀硕士学位论文全文数据库 工程科技II辑》 * 
王岩: "卫星智能自主控制系统的研究", 《中国优秀硕士学位论文全文数据库 工程科技II辑》 * 
祝芙英: "单颗导航卫星及探月飞行器的轨道确定研究", 《中过优秀硕士学位论文全文数据库 基础学科辑》 * 
邢艳军: "微小卫星轨道姿态一体化确定算法研究", 《航天控制》 * 
雷辉: "用非差分方法确定单颗导航卫星的轨道", 《天文学进展》 * 
Cited By (46)
Publication number  Priority date  Publication date  Assignee  Title 

CN103217982A (en) *  20130221  20130724  北京控制工程研究所  Orbit control method based on wheelcontrolled mode 
CN103116361B (en) *  20130221  20131120  北京控制工程研究所  Method for determining orbital transfer interval under control of satellite momentum wheel 
CN103217982B (en) *  20130221  20150819  北京控制工程研究所  A kind of method for controlling scrolling based on wheel control pattern 
CN103116361A (en) *  20130221  20130522  北京控制工程研究所  Method for determining orbital transfer interval under control of satellite momentum wheel 
CN105246781A (en) *  20130529  20160113  雷斯昂公司  Satellite orbital determination (OD) using doppler and kepler orbital elements 
CN103398726A (en) *  20130813  20131120  中国科学院长春光学精密机械与物理研究所  Judgment method of validity of least square attitude solution of star sensor 
CN103398726B (en) *  20130813  20160120  中国科学院长春光学精密机械与物理研究所  A kind of availability deciding method of least square attitude solution of star sensor 
CN103646127B (en) *  20131120  20160629  中国空间技术研究院  Satellite orbit gesture visual threedimensional display packing 
CN103646127A (en) *  20131120  20140319  中国空间技术研究院  Satellite orbit gesture visual threedimensional displaying method 
CN105510936A (en) *  20141126  20160420  航天恒星科技有限公司  Satelliteborne GNSS combined orbit determination method 
CN105799954A (en) *  20141231  20160727  上海新跃仪表厂  Spacebased modular aircraft for conducting decentralized deployment on micronano load and orbital transfer guidance method of modular aircraft 
CN105799954B (en) *  20141231  20180605  上海新跃仪表厂  Spacebased disperses the modular aircraft for disposing micronano load and its becomes rail method of guidance 
CN105182987A (en) *  20150818  20151223  北京航天长征飞行器研究所  Pose correction method for powered phase of aircraft 
CN105698764A (en) *  20160130  20160622  武汉大学  Error modeling compensation method and system of optical remote sensing satellite image timevarying system 
CN105698764B (en) *  20160130  20180123  武汉大学  A kind of Optical remote satellite image timevarying system error modeling compensation method and system 
CN106092099B (en) *  20160602  20181002  哈尔滨工业大学  spacecraft relative position increment orbit determination method 
CN106092099A (en) *  20160602  20161109  哈尔滨工业大学  Spacecraft is relative to positional increment orbit determination method 
CN105866808B (en) *  20160621  20180420  上海航天控制技术研究所  The definite method of influence of the navigation neceiver Orbit Error to attitude of satellite precision 
CN105866808A (en) *  20160621  20160817  上海航天控制技术研究所  Method for confirming influence of orbit determination errors of navigation receiver to satellite attitude precision 
CN108009462A (en) *  20161031  20180508  中南大学  It is a kind of to be applied to rail inspection filtering method of the basic string rail of instrument to data 
CN107064981A (en) *  20170410  20170818  千寻位置网络有限公司  Differential positioning method and system based on GNSS, service terminal 
CN107064981B (en) *  20170410  20190924  千寻位置网络有限公司  Differential positioning method and system based on GNSS, service terminal 
CN107101649B (en) *  20170525  20190823  北京航天自动控制研究所  A kind of inorbit error separating method of spacecraft Guidance instrumentation 
CN107101649A (en) *  20170525  20170829  北京航天自动控制研究所  A kind of inorbit error separating method of spacecraft Guidance instrumentation 
CN108512590A (en) *  20180323  20180907  中国空间技术研究院  A kind of jointtrial system and method for satellite attitude and orbit control subsystem and GNSS subsystems 
CN108512590B (en) *  20180323  20200814  中国空间技术研究院  Joint test system of satellite attitude and orbit control subsystem and GNSS subsystem 
CN109444931B (en) *  20181008  20201215  闽江学院  Static pseudo range point positioning method and device 
CN109444931A (en) *  20181008  20190308  闽江学院  A kind of method and device of static state pseudorange OnePoint Location 
CN109709588B (en) *  20181211  20201009  中国人民解放军63921部队  Highorbit satellite multisatellite highprecision orbit determination system 
CN109709588A (en) *  20181211  20190503  中国人民解放军63921部队  A kind of more star highprecision measuring rail systems of high rail satellite 
CN109625331A (en) *  20181226  20190416  上海微小卫星工程中心  Satellite controller and satellite control method 
CN109856995A (en) *  20190304  20190607  北京空间飞行器总体设计部  A kind of whole star control subsystem analog platform towards test method verifying assessment 
CN109901600A (en) *  20190308  20190618  宁波天擎航天科技有限公司  A kind of spacecraft flight control method, system and device 
CN110516379A (en) *  20190829  20191129  珠海迈越信息技术有限公司  A method of satellite signal receiving apparatus is installed based on augmented reality 
CN110516379B (en) *  20190829  20221209  珠海迈越信息技术有限公司  Method for installing satellite signal receiving device based on augmented reality technology 
CN110763231A (en) *  20191015  20200207  哈尔滨工程大学  Errorfree attitude updating method suitable for fiber optic gyroscope filtering signal 
CN110763231B (en) *  20191015  20221118  哈尔滨工程大学  Errorfree attitude updating method suitable for fiber optic gyroscope filtering signal 
CN111487657B (en) *  20200321  20220715  哈尔滨工程大学  Beidou realtime precise orbit determination method based on satellite perturbation 
CN111487657A (en) *  20200321  20200804  哈尔滨工程大学  Beidou realtime precise orbit determination method based on satellite perturbation 
CN111645882A (en) *  20200604  20200911  北京航天方舟空间技术有限公司  Satellite autonomous orbit determination method, device, equipment and computer storage medium 
CN111645882B (en) *  20200604  20220322  北京航天方舟空间技术有限公司  Satellite autonomous orbit determination method, device, equipment and computer storage medium 
CN111854764A (en) *  20200720  20201030  中国科学院微小卫星创新研究院  Spacecraft attitude determination method and system based on intersatellite measurement information 
CN113504728A (en) *  20210722  20211015  北京微纳星空科技有限公司  Method, device and equipment for generating task instruction and storage medium 
CN113504728B (en) *  20210722  20220405  北京微纳星空科技有限公司  Method, device and equipment for generating task instruction and storage medium 
CN115098828A (en) *  20220826  20220923  北京控制工程研究所  Method and device for calculating loworbit satellite orbit in near circle 
CN115098828B (en) *  20220826  20221104  北京控制工程研究所  Method and device for calculating loworbit satellite orbit in near circle 
Similar Documents
Publication  Publication Date  Title 

Hasan et al.  A review of navigation systems (integration and algorithms)  
Noureldin et al.  Fundamentals of inertial navigation, satellitebased positioning and their integration  
US6424914B1 (en)  Fullycoupled vehicle positioning method and system thereof  
CN103221839B (en)  GNSS signal processing with regional augmentation message  
CN101395443B (en)  Hybrid positioning method and device  
JP2022534387A (en)  Satellites for broadcasting highprecision data  
Shytermeja et al.  Proposed architecture for integrity monitoring of a GNSS/MEMS system with a fisheye camera in urban environment  
US8922426B1 (en)  System for geolocation  
Hassan et al.  A review of system integration and current integrity monitoring methods for positioning in intelligent transport systems  
Fernández et al.  ATENEA: Advanced techniques for deeply integrated GNSS/INS/LiDAR navigation  
Travis III  Path duplication using GPS carrier based relative position for automated ground vehicle convoys  
Afia et al.  A lowcost gnss/imu/visual monoslam/wss integration based on federated kalman filtering for navigation in urban environments  
Jiang et al.  An analysis of PPPGPSbased decentralized train multisensor navigation system  
Skaloud  Reliability of Direct Georeferencing Phase 1: An Overview of the Current Approaches and Possibilities., Checking and Improving of Digital Terrain Models/Reliability of Direct Georeferencing.  
CN102494686A (en)  Satellite attitude orbit determining system and method  
Bauer et al.  Nonlineofsight mitigation for reliable urban gnss vehicle localization using a particle filter  
Rizos  Introducing the global positioning system  
Martin et al.  Performance comparison of single and dual frequency closely coupled gps/ins relative positioning systems  
CN111801595A (en)  Multipath management for global navigation satellite systems  
Afia et al.  A GNSS/IMU/WSS/VSLAM hybridization using an extended kalman filter  
Liu et al.  GNSSbased localization for autonomous vehicles: Prospects and challenges  
Elsheikh  Integration of GNSS precise point positioning and inertial technologies for land vehicle navigation  
Skaloud  Reliability in direct georeferencing: An overview of the current approaches and possibilities  
CN102494686B (en)  A kind of satellite attitude orbit determines system and method  
Tabb  MultiAntenna GPS for Improved Carrier Phase Positioning in Autonomous Convoys 
Legal Events
Date  Code  Title  Description 

C06  Publication  
PB01  Publication  
C10  Entry into substantive examination  
SE01  Entry into force of request for substantive examination  
C14  Grant of patent or utility model  
GR01  Patent grant  
CP01  Change in the name or title of a patent holder 
Address after: 16th Floor of No.63 Satellite Building, Zhichun Road, Haidian District, Beijing, 100190 Patentee after: Beijing Guoke Huanyu Science and Technology Co., Ltd. Address before: 16th Floor of No.63 Satellite Building, Zhichun Road, Haidian District, Beijing, 100190 Patentee before: Beijing Guokehuanyu Space Technology Co., Ltd. 

CP01  Change in the name or title of a patent holder 