CN110793528A - Low-orbit satellite-based anchoring-based Beidou navigation constellation autonomous orbit determination method - Google Patents
Low-orbit satellite-based anchoring-based Beidou navigation constellation autonomous orbit determination method Download PDFInfo
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Abstract
The invention relates to a low-orbit satellite-based anchoring-based Beidou navigation constellation autonomous orbit determination method which comprises the following steps: bidirectional pseudo-range observation is carried out among the planets through the Beidou inter-satellite link, and bidirectional observed quantity information among the Beidou satellites, and prior position and speed information of the local satellite under the geocentric inertial coordinate system are sent to adjacent Beidou navigation satellites through the inter-satellite link; receiving Beidou downlink observation quantity information and low-orbit satellite autonomous orbit determination data sent by a low-orbit satellite; error correction is carried out on the pseudo range between the Beidou double satellites and the pseudo range between the Beidou satellite and the low orbit satellite, and errors of the receiving and transmitting equipment and link errors are eliminated; unifying the corrected pseudo range between the Beidou satellite and the low earth orbit satellite and the corrected pseudo range between the Beidou satellite and the low earth orbit satellite to the same moment, decoupling the inter-satellite distance and the relative clock error in the two-way observation information between the Beidou satellites, taking the pseudo range between the low earth orbit satellite, the inter-Beidou satellite distance and the relative clock error as observed quantities, and calculating the position, the speed and the clock error of the Beidou satellite by adopting an extended Kalman filtering method.
Description
Technical Field
The invention relates to the technical field of autonomous orbit determination of navigation constellations, in particular to a Beidou navigation constellation autonomous orbit determination method based on low-orbit satellite-based anchoring.
Background
At present, the autonomous orbit determination of the Beidou navigation constellation adopts two methods: 1. based on the autonomous orbit determination of the ground anchoring station, the ground operation and control station or the ground rover station anchors the Beidou constellation, the Beidou constellation is used as a pseudolite with known coordinates to complete Ka frequency band bidirectional precise measurement and time synchronization with other satellites of a navigation constellation, and the prior coordinates of the pseudolite are utilized to establish the connection between the Beidou constellation and the earth; 2. the navigation constellation is completely and autonomously orbit-fixed, and only the high-precision inter-satellite measurement link between the middle-high orbit and the middle-middle orbit in the navigation constellation is used as observation to finish the autonomous orbit-fixing of the constellation. The first method cannot completely cut off the association between the Beidou constellation and the ground, and has safety risk when ground anchoring facilities are unavailable; the second method has navigation constellation orbit and time divergence phenomenon in long-time autonomous orbit determination due to lack of earth reference information.
Under the war condition, the autonomous orbit determination capability of the navigation constellation is an important means for ensuring that the satellite navigation system does not depend on the ground and improving the safety and destroy resistance of the navigation constellation system. However, under the condition of lacking the Beidou ground system anchoring, the Beidou navigation constellation autonomous orbit determination for a long time faces the phenomena of service precision divergence and decline. The URE index of the Beidou autonomous orbit determination for 60 days is 3 meters. If the autonomous orbit determination time is further prolonged, the URE is increased rapidly, so that the Beidou satellite navigation system is difficult to provide high-precision positioning and time service.
Document 1: the patent: an autonomous orbit determination method for a navigation satellite constellation, patent number: CN104048664A belongs to the field of satellite autonomous orbit determination, and relates to a method for autonomous orbit determination of a navigation satellite constellation. Which comprises the following steps: (1): starting autonomous orbit determination, and initializing a system; (2): obtaining inter-satellite ranging between satellites; (3): calculating an observation matrix of the inter-satellite distance measurement; (4): forecasting the track by utilizing a track dynamics model; (5): respectively carrying out measurement updating and state estimation by using a Kalman filtering algorithm; (6): judging whether the autonomous orbit determination is finished or not, and if not, operating the steps (2) to (6) again; otherwise, the autonomous orbit determination program is finished exiting. The method has the following advantages: 1. the problem of rank deficiency when the navigation constellation only utilizes inter-satellite measuring distance to carry out autonomous orbit determination can be solved; 2. the long-term autonomous orbit determination precision is higher.
Document 2: the patent: navigation satellite autonomous navigation system and method based on X-ray pulsar, patent number: CN 101038169A. The navigation satellite autonomous navigation system based on the X-ray pulsar comprises: the system comprises an X-ray detector, a satellite-borne atomic clock group, a solar system planet parameter database, an X-ray pulsar model, a characteristic parameter database, a satellite-borne computer, a Strapdown Inertial Navigation System (SINS) and an autonomous navigation algorithm module library; the autonomous navigation method utilizes X-ray photons radiated by pulsar as external information input, extracts pulse arrival time TOA and angular position information, performs data processing through an autonomous navigation filter, acquires navigation parameters such as position, speed, time and attitude of a navigation satellite in real time, autonomously generates navigation messages and control instructions, and realizes autonomous orbit determination of a navigation constellation. The method has the advantages of providing long-time high-precision autonomous navigation for the navigation satellite, improving the fault-tolerant capability of autonomous navigation information processing of the navigation satellite, and being suitable for high-precision autonomous navigation of low earth orbit, deep space and interplanetary flying spacecrafts, dense atmosphere-free celestial body landers and surface tourists thereof.
Document 3: the patent: an earth-Lagrange combined constellation autonomous orbit determination method based on inter-satellite ranging (patent number: CN107421550A) discloses an earth-Lagrange combined constellation autonomous orbit determination method based on inter-satellite ranging, which comprises the following steps: the method comprises the following steps: establishing a state equation of an earth-Lagrange combined constellation autonomous orbit determination system; step two: establishing a measurement equation of an earth-Lagrange combined constellation autonomous orbit determination system; step three: determining a filtering method for realizing orbit parameter estimation; step four: the earth-Lagrange combined constellation autonomous orbit determination method based on the selected filtering algorithm is specifically realized. According to the method, by introducing Lagrange satellites, the problem of 'deficit rank' existing when autonomous orbit determination is carried out only by using inter-satellite ranging information is effectively solved, and the complexity of system equipment is reduced; the statistical characteristic of the system noise is estimated on line in real time through the self-adaptive nonlinear filtering algorithm, the requirement on noise prior information is low, the stability of the autonomous orbit determination filtering algorithm is improved, and the autonomous orbit determination precision is improved.
Document 4: the article: navigation constellation entire network autonomous orbit determination based on inter-satellite ranging and ground emission sources. The method is characterized in that a small number of ground emission sources work randomly to provide ground reference, the inter-satellite distance measurement and the ground emission source information are fused to carry out constellation whole network orbit determination, the orbit determination precision is further improved, and finally, a simulation experiment is utilized to verify the reasonable effectiveness of the method.
Documents 1 and 3 describe a method for solving the problem of reference loss in autonomous orbit determination of a navigation constellation by using special orbital characteristics of a Lagrangian point satellite, document 2 describes a method for providing an external reference for the navigation constellation through pulsar observation, and document 4 describes a method for establishing a relationship between the navigation constellation and the earth by using a ground anchor station. Among them, the methods described in documents 1 and 3 require support from a satellite at a lagrangian point in the earth, and the engineering cost is very high; the autonomous navigation precision of the method introduced in the document 2 is poor, and the method introduced in the document 4 belongs to a navigation constellation semi-autonomous orbit determination method and is not supported by disengaging from the ground.
Disclosure of Invention
The technical problem solved by the invention is as follows: the method overcomes the defects of the prior art, provides the Beidou navigation constellation autonomous orbit determination method based on low-orbit satellite-based anchoring, and supports autonomous orbit determination of navigation satellites by using low-orbit satellites as satellite-based anchoring stations.
The technical scheme of the invention is as follows: a low-orbit satellite-based anchoring-based Beidou navigation constellation autonomous orbit determination method comprises the following steps:
(1) carrying out bidirectional pseudo-range observation among the planets through a Beidou inter-satellite link, and sending bidirectional observed quantity information among the Beidou satellites, and prior position and speed information of the local satellite in a geocentric inertial coordinate system to an adjacent Beidou navigation satellite through an inter-satellite link; the two-way observation quantity information between the Beidou satellites is pseudo-range between the Beidou satellites;
(2) receiving Beidou downlink observation quantity information and low-orbit satellite autonomous orbit determination data sent by a low-orbit satellite; the low-orbit satellite autonomous orbit determination data comprises the position and the speed of the low-orbit satellite in the geocentric inertial coordinate system; the Beidou downlink observation quantity information is pseudo-range between a Beidou satellite and a low-orbit satellite, which is obtained when the low-orbit satellite receives a Beidou downlink navigation signal;
(3) error correction is carried out on the pseudo range between the Beidou double satellites and the pseudo range between the Beidou satellite and the low-orbit satellite, and errors of the receiving and transmitting equipment and link errors are eliminated;
(4) unifying the corrected pseudo range between the Beidou satellite and the low earth orbit satellite and the corrected pseudo range between the Beidou satellite and the low earth orbit satellite to the same moment through epoch reduction, decoupling the inter-satellite distance and the relative clock error in the two-way observation information between the Beidou satellites, taking the reduced pseudo range of the low earth orbit satellite, the inter-Beidou satellite distance and the inter-Beidou satellite relative clock error as observed quantities, taking the position, the speed and the clock error of the Beidou navigation satellite under a geocentric inertial coordinate system as system state quantities, establishing a system function through an orbit dynamics model, finishing the recursion of the system state quantities under the geocentric inertial system by adopting a numerical integration method, taking the recursion as the prior estimation of the system state quantities, and solving the partial differential of the system state quantities through the system function to generate a system matrix; and generating an observation matrix through the space geometric distribution of the Beidou navigation satellite and the low-orbit satellite, and continuously iterating and updating by adopting an extended Kalman filtering method to obtain the position, the speed and the clock error of the Beidou satellite.
The step (3) comprises the following substeps:
(3.1) carrying out equipment time delay error correction on the pseudo range between the Beidou double satellites and the pseudo range between the Beidou satellite and the low orbit satellite, and eliminating errors introduced by time delay of the receiving and sending equipment contained in the pseudo range;
(3.2) calculating to obtain an antenna phase center correction value according to the attitude information of the Beidou satellite and the antenna phase center offset parameters of the inter-satellite link transceiving antenna and the downlink signal transmitting antenna, and performing antenna phase center correction on the pseudo range between the Beidou double satellites and the pseudo range between the Beidou satellite and the low-orbit satellite to obtain the pseudo range between the Beidou double-satellite centroids and the pseudo range between the Beidou satellite and the low-orbit satellite antenna centroids; (3.3) judging whether the inter-Beidou satellite observation is over-ionosphere observation or not, discarding a pseudo range passing through the ionosphere, and taking a pseudo range between the remaining Beidou dual-satellite antenna phase centers as a corrected pseudo range between the Beidou dual satellites;
and (3.4) performing double-frequency ionosphere correction on the pseudo range between the Beidou satellite and the low-orbit satellite antenna phase center to obtain the pseudo range between the Beidou satellite with the ionosphere error eliminated and the low-orbit satellite antenna phase center, and taking the pseudo range as the corrected pseudo range between the Beidou satellite and the low-orbit satellite.
The low-orbit satellite completes the on-satellite autonomous orbit determination by receiving and processing ubiquitous exogenous navigation signals so as to acquire the autonomous orbit determination data of the low-orbit satellite.
The ubiquitous exogenous navigation signals comprise GPS satellite navigation signals, GALILEO satellite navigation signals, GLONASS satellite navigation signals, QZSS navigation system L6 frequency point enhancement signals, GALILEO navigation system L6 frequency point enhancement signals, OmniSTAR navigation system L frequency band enhancement signals, Starfire navigation system L frequency band enhancement signals or global precision L frequency band enhancement signals.
The low-orbit satellite refers to a man-made satellite with the flight height below 2000 kilometers.
The orbital dynamics model is represented as:
a(r,t)＝ag(r,t)+as(r,t)+am(r,t)+asrp(r,t)
wherein a (r, t) represents the acceleration caused by the orbit perturbation received by the Beidou navigation satellite, r represents the position of the Beidou satellite, t represents the time, ag(r, t) represents the acceleration caused by the earth gravity field received by the Beidou navigation satellite, as(r, t) represents the acceleration of the Beidou navigation satellite caused by the sun, am(r, t) represents the acceleration of the Beidou navigation satellite due to the moon, asrpAnd (r, t) represents the acceleration caused by sunlight pressure received by the Beidou navigation satellite.
The big dipperTime-receiving inter-satellite pseudo range rho of A satellite launching-Beidou B satelliteABCorresponding apparatus delay correction
Time-receiving inter-satellite pseudo range rho of Beidou satellite B to Beidou satellite ABACorresponding apparatus delay correction
And the pseudo range rho of the low orbit satellite when the Beidou A satellite transmits and the low orbit a satellite receivesAaCorresponding apparatus delay correction
(Beidou is calculated as follows:
wherein the content of the first and second substances,
is a star transmit device delay;delaying for B star receiving equipment;
a, satellite receiving equipment delay;b, satellite sending equipment delay;
low earth orbit satellite a receives the device delay. These devices transmit and receiveThe time delay can be obtained by calibration before delivery.
The antenna phase center correction amount
And
the following are calculated respectively:
wherein T is a transfer matrix from a satellite body coordinate system to a geocentric inertial coordinate system, satellite attitude information is provided by a satellite platform attitude control system in real time,and
antenna phase center offset parameters of an inter-satellite transmitting antenna, an inter-satellite receiving antenna, a downlink transmitting antenna and a low-orbit satellite receiving antenna of the Beidou navigation satellite under each satellite body coordinate system are respectively set; r isA、rBAnd raPosition vectors of Beidou navigation satellites A and B and a low-orbit satellite a in a geocentric inertial coordinate system are respectively.
The specific method for judging whether the double-satellite pseudo range observation among the Beidou satellites is the over-ionosphere observation in the step (3.3) is as follows:
judging whether an included angle ∠β formed by the connecting lines of the Beidou navigation satellites A and B and the connecting line of the Beidou navigation satellite B and the earth mass center is larger than an included angle ∠α formed by the connecting line of the Beidou navigation satellite B and the earth mass center, if so, the inter-satellite double-satellite pseudo range observation is non-ionosphere observation, otherwise, the inter-satellite double-satellite pseudo range observation is ionosphere observation;
wherein R isionoRepresents the radius of the ionosphere; r isA、rBAnd raRespectively, the position vectors of the beidou navigation satellite A, B in the earth-centered inertial coordinate system.
The pseudo range P after ionospheric influence eliminationiono-freeComprises the following steps:
wherein: p1And P2Respectively representing pseudo-range observed values corresponding to the dual-frequency navigation signals, f1、f2Respectively, a first carrier frequency and a second carrier frequency of the dual-frequency navigation signal.
Compared with the prior art, the invention has the beneficial effects that:
(1) the invention utilizes the global coverage, quick movement and low-cost low-orbit mobile constellation as a satellite-based anchoring station to provide an external spatial reference for the Beidou navigation constellation in an autonomous orbit determination mode, thereby inhibiting the phenomenon that the orbit precision diverges along with time when the constellation autonomously determines the orbit and ensuring the long-term high-precision autonomous orbit determination.
(2) According to the autonomous orbit determination method, the observation matrix is generated through the space geometric distribution of the Beidou navigation satellite and the low orbit satellite, and compared with the autonomous orbit determination through the independent use of the Beidou navigation satellite inter-satellite observation, the autonomous orbit determination method can enhance the geometric strength of observation and the change speed of the geometric relation, so that the accuracy of autonomous orbit determination is further improved.
(3) The low-orbit navigation enhanced constellation is the key development core content of the next generation PNT system in China, and can utilize the low-orbit GNSS monitoring function and the link function between the low-orbit satellites and the middle-low-orbit satellites of the constellation to complete the generation and the transmission of low-orbit observation information/orbit determination results.
(4) The low-orbit satellite is used as the satellite-based anchoring station of the Beidou navigation constellation, so that the long-term high-precision autonomous orbit determination of the Beidou navigation satellite is realized under the condition of being separated from the ground support.
Drawings
FIG. 1 is a low-orbit satellite-based anchored Beidou navigation constellation autonomous orbit determination method according to an embodiment of the present invention;
FIG. 2 is a processing flow chart of the autonomous orbit determination method of the Beidou navigation constellation based on low-orbit satellite-based anchoring in the embodiment of the invention;
fig. 3 is a schematic diagram of an inspection of an over-ionization layer of the inter-Beidou satellite measurement link according to the embodiment of the invention.
Detailed Description
The invention is further illustrated by the following examples.
The invention provides a low-orbit satellite-based anchoring-based Beidou navigation constellation autonomous orbit determination method. The main idea is as follows: the global coverage, quick movement and low cost low-orbit mobile constellation is used as a star-based anchoring station to provide an external space reference for the Beidou navigation constellation in an autonomous orbit determination mode, so that the phenomenon that the orbit precision diverges along with time when the constellation autonomously determines the orbit is inhibited, and long-term high-precision autonomous orbit determination is ensured. The low-orbit satellite is an artificial satellite with the flight height of less than 2000 kilometers. The satellite autonomous orbit determination is completed by receiving and processing ubiquitous exogenous navigation signals so as to obtain self coordinates and time information; meanwhile, the low-orbit satellite monitors and receives a navigation signal broadcast by a high-orbit satellite in the Beidou constellation, and then the Beidou observation quantity, the self coordinate and the time information are framed and sent to the Beidou navigation satellite through a low-high orbit communication link. The Beidou navigation constellation completes inter-satellite bidirectional measurement through inter-satellite links, each Beidou navigation satellite uses adjacent inter-satellite bidirectional observation quantities to conduct distributed constellation autonomous orbit determination, and when measurement information of low-orbit satellites is effective, low-orbit satellite observation quantities are introduced into an observation model of autonomous orbit determination so as to eliminate and weaken the integral rotation problem caused by reference loss during autonomous orbit determination of the Beidou constellation by means of space references of the low-orbit satellites. The autonomous orbit determination method of the Beidou navigation constellation based on low-orbit satellite-based anchoring is shown in FIG. 1.
The autonomous orbit determination method for the navigation constellation comprises a Beidou navigation satellite and a low orbit satellite onboard processing part, and the data processing flow is shown in figure 2.
The Beidou navigation satellite carries out distributed autonomous orbit determination, each Beidou navigation satellite respectively processes the Beidou inter-satellite bidirectional observation quantity information of the adjacent navigation satellite and the Beidou downlink observation quantity information sent by the low orbit satellite related to the Beidou navigation satellite, and the satellite position and time estimation is finished by using a satellite-borne extended Kalman filter; the low-orbit satellite observes the Beidou navigation satellite and autonomously determines the orbit through an external source signal.
Beidou navigation satellite onboard processing
(1) Carrying out bidirectional pseudo-range observation among the planets through a Beidou inter-satellite link, and sending bidirectional observed quantity information among the Beidou satellites, and prior position and speed information of the local satellite in a geocentric inertial coordinate system to an adjacent Beidou navigation satellite through an inter-satellite link; the Beidou inter-satellite bidirectional observed quantity information comprises pseudo ranges between Beidou double satellites;
(2) receiving Beidou downlink observation quantity information and low-orbit satellite autonomous orbit determination data sent by a low-orbit satellite; the low-orbit satellite autonomous orbit determination data comprises the position and the speed of the low-orbit satellite in the geocentric inertial coordinate system; the Beidou downlink observation quantity information is pseudo-range between a Beidou satellite and a low-orbit satellite, which is obtained when the low-orbit satellite receives a Beidou downlink navigation signal;
(3) error correction is carried out on the pseudo range between the Beidou double satellites and the pseudo range between the Beidou satellite and the low-orbit satellite, and errors of the receiving and transmitting equipment and link errors are eliminated;
the step (3) comprises the following substeps:
(3.1) carrying out equipment time delay error correction on the pseudo range between the Beidou double satellites and the pseudo range between the Beidou satellite and the low orbit satellite, and eliminating errors introduced by time delay of the receiving and sending equipment contained in the pseudo range;
the original distance observed quantity contains errors introduced by time delay of the inter-satellite transceiver, and the errors can be calibrated on the ground in advance and can be directly added to the observed quantity during on-orbit calculation. The time-receiving inter-satellite pseudo range rho of the Beidou A satellite transmitter and the Beidou B satellite receiverABCorresponding apparatus delay correction
Time-receiving inter-satellite pseudo range rho of Beidou satellite B to Beidou satellite ABACorresponding apparatus delay correction
And the pseudo range rho of the low orbit satellite when the Beidou A satellite transmits and the low orbit a satellite receivesAaCorresponding apparatus delay correction
(Beidou is calculated as follows:
wherein the content of the first and second substances,
is a star transmit device delay;
delaying for B star receiving equipment;
a, satellite receiving equipment delay;
b, satellite sending equipment delay;
low earth orbit satellite a receives the device delay. The transceiving time delay of the devices can be obtained through calibration before leaving a factory.
(3.2) calculating to obtain an antenna phase center correction value according to the attitude information of the Beidou satellite and the antenna phase center offset parameters of the inter-satellite link transceiving antenna and the downlink signal transmitting antenna, and performing antenna phase center correction on the pseudo range between the Beidou double satellites and the pseudo range between the Beidou satellite and the low-orbit satellite to obtain the pseudo range between the Beidou double-satellite centroids and the pseudo range between the Beidou satellite and the low-orbit satellite antenna centroids;
autonomous constellation orbit determination of the Beidou navigation satellite system needs to determine a satellite centroid coordinate, but the distance between Beidou satellites and low orbit measurement is obtained relative to a phase center of a transmitting-receiving antenna, so that phase center correction processing is needed.
The antenna phase center correction amount
And
the following are calculated respectively:
wherein T is a transfer matrix from a satellite body coordinate system to a geocentric inertial coordinate system and is obtained by attitude information of a Beidou satellite platform, the satellite attitude information is provided by a satellite platform attitude control system in real time,
and
antenna phase center offset parameters of an inter-satellite transmitting antenna, an inter-satellite receiving antenna, a downlink transmitting antenna and a low-orbit satellite receiving antenna of the Beidou navigation satellite under each satellite body coordinate system are respectively obtained through calibration before delivery; r isA、rBAnd raThe position vectors of the Beidou navigation satellites A and B and the low-orbit satellite a in the geocentric inertial coordinate system can be calculated by using the prior orbit position.
(3.3) judging whether the inter-Beidou satellite observation is over-ionosphere observation or not, discarding a pseudo range passing through the ionosphere, and taking a pseudo range between the remaining Beidou dual-satellite antenna phase centers as a corrected pseudo range between the Beidou dual satellites;
microwave ranging signals are subject to additional delays as they pass through the ionosphere. For observation among Beidou satellites, over ionosphere inspection needs to be carried out, and for observation through the ionosphere, the observation is discarded without use.
The specific method for judging whether the double-satellite pseudo-range observation among the Beidou satellites is the over-ionosphere observation is shown in fig. 3:
∠β is an included angle formed by a connecting line of Beidou navigation satellites A and B and a connecting line of the Beidou navigation satellite B and the earth mass center, ∠α is an included angle formed by an earth ionosphere tangent line of the Beidou navigation satellite B and a connecting line of the Beidou navigation satellite B and the earth mass center, if ∠β is larger than ∠α, the inter-satellite double-satellite pseudo range observation is non-ionosphere observation, otherwise, the inter-satellite double-satellite pseudo range observation is ionosphere observation;
wherein R isionoRepresents the radius of the ionosphere, typically the earth radius +1000 km; r isA、rBAnd raRespectively, the position vectors of the beidou navigation satellite A, B in the earth-centered inertial coordinate system.
And (3.4) performing double-frequency ionosphere correction on the pseudo range between the Beidou satellite and the low-orbit satellite antenna phase center to obtain the pseudo range between the Beidou satellite with the ionosphere error eliminated and the low-orbit satellite antenna phase center, and taking the pseudo range as the corrected pseudo range between the Beidou satellite and the low-orbit satellite.
For low earth orbit satellite observations, because they are in ionosphere coverage, dual-frequency ionosphere corrections are made for all observations. Pseudorange P after ionospheric effects eliminationiono-freeComprises the following steps:
wherein: p1And P2Respectively representing pseudo-range observed values corresponding to the dual-frequency navigation signals, f1、f2Respectively, a first carrier frequency and a second carrier frequency of the dual-frequency navigation signal.
(4) Unifying the corrected pseudo range between the Beidou satellite and the low earth orbit satellite and the corrected pseudo range between the Beidou satellite and the low earth orbit satellite to the same moment through epoch reduction, decoupling the inter-satellite distance and the relative clock error in the two-way observation information between the Beidou satellites, taking the reduced pseudo range of the low earth orbit satellite, the inter-Beidou satellite distance and the inter-Beidou satellite relative clock error as observed quantities, taking the position, the speed and the clock error of the Beidou navigation satellite under a geocentric inertial coordinate system as system state quantities, establishing a system function through an orbit dynamics model, finishing the recursion of the system state quantities under the geocentric inertial system by adopting a numerical integration method, taking the recursion as the prior estimation of the system state quantities, and solving the partial differential of the system state quantities through the system function to generate a system matrix; and generating an observation matrix through the space geometric distribution of the Beidou navigation satellite and the low-orbit satellite, and continuously iterating and updating by adopting an extended Kalman filtering method to obtain the position, the speed and the clock error of the Beidou satellite.
And (4.1) measuring the epoch reduction.
The inter-satellite bidirectional pseudo range observation after various error corrections is as follows:
ρAB(t1,t3)＝|rA(t1)-rB(t3)|+c[δtB(t3)-δtA(t1)]+εAB
ρBA(t2,t4)＝|rB(t2)-rA(t4)|+c[δtA(t4)-δtB(t2)]+εAB
the low-orbit pseudo range observation after various error corrections is as follows:
ρAa(t5,t6)＝|rA(t5)-ra(t6)|+c[δta(t6)-δtA(t5)]+εAa
wherein, δ tAAnd δ tBIs the clock difference of A star and B star, epsilonAB，εABAnd εAaTo observe the noise. The observed quantity needs to finish time synchronization through reduction, and the position and clock error of the navigation satellite in the observed quantity are unified to the same moment.
The position reduction method is as follows, by means of a correction termThe positions of A star and B star in the observation can be reduced to a uniform t0Time of day, wherein rA(t1)，rB(t3)，vA(t1)，vB(t3) The forecasted track position and velocity can be used.
The clock error reduction method is as follows, by means of the correction term
The clock error of A star and B star in observation can be reduced to uniform t0At a time, wherein
The prior clock error parameters stored on the satellite can be continuously refined and estimated in the autonomous orbit determination process.
The new observed quantity generated after the measurement time is reduced is as follows:
ρAB(t0,t0)＝|rA(t0)-rB(t0)|+c[δtB(t0)-δtA(t0)]+εAB
ρBA(t0,t0)＝|rB(t0)-rA(t0)|+c[δtA(t0)-δtB(t0)]+εAB
ρAa(t0,t6)＝|rA(t0)-ra(t6)|+c[δta(t6)-δtA(t0)]+εAa
(4.2), clock error and distance decoupling.
The distance and the relative clock error between the satellites are decoupled through the two-way processing of the Beidou satellite observation, and then the distance can be used as the observation of Kalman filtering.
For low-orbit observed quantity, because only one-way measurement is needed, clock offset and distance cannot be decoupled through two-way processing, and distance and time decoupling needs to be completed by estimating time information of a navigation satellite in a Kalman filter.
(4.3) Kalman Filter
The invention generates an observation matrix through the space geometric distribution of the Beidou navigation satellite and the low-orbit satellite. The orbit dynamics model needs to consider the factors such as the earth high-order gravity field, the gravity of the lunar and lunar trisomy, the solar radiation pressure and the like, and the mechanics model is expressed as follows:
the orbital dynamics model is represented as:
a(r,t)＝ag(r,t)+as(r,t)+am(r,t)+asrp(r,t)
wherein a (r, t) represents the acceleration caused by the orbit perturbation received by the Beidou navigation satellite, r represents the position of the Beidou satellite, t represents the time, ag(r, t) represents the acceleration caused by the earth gravity field received by the Beidou navigation satellite, as(r, t) represents the acceleration of the Beidou navigation satellite caused by the sun, am(r, t) represents the acceleration of the Beidou navigation satellite due to the moon, asrpAnd (r, t) represents the acceleration caused by sunlight pressure received by the Beidou navigation satellite.
The observation model is mainly determined by the space geometric distribution of the Beidou navigation satellite and the low earth orbit satellite. The inter-satellite observation of the Beidou navigation satellite does not contain time information, the time and distance information of low-orbit observation are not decoupled, and the time difference between the Beidou navigation satellite and the low-orbit satellite needs to be additionally estimated. The observation matrix is specifically represented as follows:
wherein the content of the first and second substances,represents the unit position vector between the Beidou navigation satellites,
and the unit position vector between the Beidou navigation satellite and the low-orbit satellite is represented. n and m represent the number of inter-satellite observations and low-orbit observations, respectively. The observation model establishes the correlation between the system state and the inter-satellite observation and the low-orbit observation.
Two, low orbit satellite on-board processing
1. And (5) observing a Beidou navigation signal and an external source signal.
And observing the Beidou navigation signal by the low-orbit satellite to obtain Beidou downlink observation quantity information, namely pseudo-range between the Beidou satellite and the low-orbit satellite. And the low-orbit satellite observes the exogenous navigation signal to generate navigation enhancement information such as exogenous navigation observed quantity information, exogenous navigation signal pseudo-range observed quantity, carrier phase observed quantity, navigation satellite precise orbit clock error and the like. The ubiquitous exogenous navigation signals comprise GPS satellite navigation signals, GALILEO satellite navigation signals, GLONASS satellite navigation signals, QZSS navigation system L6 frequency point enhancement signals, GALILEO navigation system L6 frequency point enhancement signals, OmniSTAR navigation system L frequency band enhancement signals, Starfire navigation system L frequency band enhancement signals or global precision L frequency band enhancement signals. As listed in the following table:
2. and the low-orbit satellite autonomously orbits.
If the number of the exogenous navigation signal observation data is more than 4 (5 are needed for double-system observation, and 1 is added to the number of the observation needed when the number of the systems is increased), the low-orbit satellite geometric autonomous orbit determination can be carried out by the SPP (standard single-point positioning) technology. If the exogenous navigation enhanced signal is effective, performing geometric precise autonomous orbit determination by using PPP (precise point positioning) in an effective time period to obtain the position and speed information of the low-orbit satellite in the geocentric inertial coordinate system.
3. And uploading the orbit determination result and the Beidou observation group frame.
And performing data framing on the autonomous orbit determination result of the low-orbit satellite and the Beidou downlink observation quantity information, and uploading the Beidou navigation satellite through the low-high orbit communication link.
Parts of the specification which are not described in detail are within the common general knowledge of a person skilled in the art.
Claims (10)
1. A low-orbit satellite-based anchoring-based Beidou navigation constellation autonomous orbit determination method is characterized in that a Beidou navigation satellite executes the following steps:
(1) carrying out bidirectional pseudo-range observation among the planets through a Beidou inter-satellite link, and sending bidirectional observed quantity information among the Beidou satellites, and prior position and speed information of the local satellite in a geocentric inertial coordinate system to an adjacent Beidou navigation satellite through an inter-satellite link; the two-way observation quantity information between the Beidou satellites is pseudo-range between the Beidou satellites;
(2) receiving Beidou downlink observation quantity information and low-orbit satellite autonomous orbit determination data sent by a low-orbit satellite; the low-orbit satellite autonomous orbit determination data comprises the position and the speed of the low-orbit satellite in the geocentric inertial coordinate system; the Beidou downlink observation quantity information is pseudo-range between a Beidou satellite and a low-orbit satellite, which is obtained when the low-orbit satellite receives a Beidou downlink navigation signal;
(3) error correction is carried out on the pseudo range between the Beidou double satellites and the pseudo range between the Beidou satellite and the low-orbit satellite, and errors of the receiving and transmitting equipment and link errors are eliminated;
(4) unifying the corrected pseudo range between the Beidou satellite and the low earth orbit satellite and the corrected pseudo range between the Beidou satellite and the low earth orbit satellite to the same moment through epoch reduction, decoupling the inter-satellite distance and the relative clock error in the two-way observation information between the Beidou satellites, taking the reduced pseudo range of the low earth orbit satellite, the inter-Beidou satellite distance and the inter-Beidou satellite relative clock error as observed quantities, taking the position, the speed and the clock error of the Beidou navigation satellite under a geocentric inertial coordinate system as system state quantities, establishing a system function through an orbit dynamics model, finishing the recursion of the system state quantities under the geocentric inertial system by adopting a numerical integration method, taking the recursion as the prior estimation of the system state quantities, and solving the partial differential of the system state quantities through the system function to generate a system matrix; and generating an observation matrix through the space geometric distribution of the Beidou navigation satellite and the low-orbit satellite, and continuously iterating and updating by adopting an extended Kalman filtering method to obtain the position, the speed and the clock error of the Beidou satellite.
2. The low-orbit satellite-based anchored Beidou navigation constellation autonomous orbit determination method according to claim 1, characterized in that: the step (3) comprises the following substeps:
(3.1) carrying out equipment time delay error correction on the pseudo range between the Beidou double satellites and the pseudo range between the Beidou satellite and the low orbit satellite, and eliminating errors introduced by time delay of the receiving and sending equipment contained in the pseudo range;
(3.2) calculating to obtain an antenna phase center correction value according to the attitude information of the Beidou satellite and the antenna phase center offset parameters of the inter-satellite link transceiving antenna and the downlink signal transmitting antenna, and performing antenna phase center correction on the pseudo range between the Beidou double satellites and the pseudo range between the Beidou satellite and the low-orbit satellite to obtain the pseudo range between the Beidou double-satellite centroids and the pseudo range between the Beidou satellite and the low-orbit satellite antenna centroids; (3.3) judging whether the inter-Beidou satellite observation is over-ionosphere observation or not, discarding a pseudo range passing through the ionosphere, and taking a pseudo range between the remaining Beidou dual-satellite antenna phase centers as a corrected pseudo range between the Beidou dual satellites;
and (3.4) performing double-frequency ionosphere correction on the pseudo range between the Beidou satellite and the low-orbit satellite antenna phase center to obtain the pseudo range between the Beidou satellite with the ionosphere error eliminated and the low-orbit satellite antenna phase center, and taking the pseudo range as the corrected pseudo range between the Beidou satellite and the low-orbit satellite.
3. The low-orbit satellite-based anchored Beidou navigation constellation autonomous orbit determination method according to claim 1, characterized in that: the low-orbit satellite completes the on-satellite autonomous orbit determination by receiving and processing ubiquitous exogenous navigation signals so as to acquire the autonomous orbit determination data of the low-orbit satellite.
4. The low-orbit satellite-based anchored Beidou navigation constellation autonomous orbit determination method according to claim 1, characterized in that: the ubiquitous exogenous navigation signals comprise GPS satellite navigation signals, GALILEO satellite navigation signals, GLONASS satellite navigation signals, QZSS navigation system L6 frequency point enhancement signals, GALILEO navigation system L6 frequency point enhancement signals, OmniSTAR navigation system L frequency band enhancement signals, Starfire navigation system L frequency band enhancement signals or global precision L frequency band enhancement signals.
5. The autonomous orbit determination method of Beidou navigation satellite system based on low-orbit satellite-based anchoring according to claim 1, characterized in that the low-orbit satellite is an artificial satellite with a flight altitude below 2000 km.
6. The low-orbit satellite-based anchored Beidou navigation constellation autonomous orbit determination method according to claim 1, characterized in that the orbit dynamics model is expressed as:
a(r，t)＝ag(r，t)+as(r，t)+am(r，t)+asrp(r，t)
wherein a (r, t) represents the acceleration caused by the orbit perturbation received by the Beidou navigation satellite, r represents the position of the Beidou satellite, t represents the time, ag(r, t) represents the acceleration caused by the earth gravity field received by the Beidou navigation satellite, as(r, t) represents the acceleration of the Beidou navigation satellite caused by the sun, am(r, t) represents the acceleration of the Beidou navigation satellite due to the moon, asrpAnd (r, t) represents the acceleration caused by sunlight pressure received by the Beidou navigation satellite.
7. The low-orbit satellite-based anchored Beidou navigation satellite system according to claim 1The seat autonomous orbit determination method is characterized in that the pseudo range rho between the Beidou A satellite transmission and the Beidou B satellite time receivingABCorresponding apparatus delay correction
Time-receiving inter-satellite pseudo range rho of Beidou satellite B to Beidou satellite ABACorresponding apparatus delay correctionAnd the pseudo range rho of the low orbit satellite when the Beidou A satellite transmits and the low orbit a satellite receivesAaCorresponding apparatus delay correction(Beidou is calculated as follows:
wherein the content of the first and second substances,
is a star transmit device delay;delaying for B star receiving equipment;
satellite reception device delay;satellite transmission equipment delay;
low earth orbit satellite a receives the device delay. The transceiving time delay of the devices can be obtained through calibration before leaving a factory.
8. The low-orbit satellite-based anchored Beidou navigation constellation autonomous orbit determination method of claim 1, characterized in that the correction amount of the antenna phase center is
Andthe following are calculated respectively:
wherein T is a transfer matrix from a satellite body coordinate system to a geocentric inertial coordinate system, satellite attitude information is provided by a satellite platform attitude control system in real time,
and
antenna phase center offset parameters of an inter-satellite transmitting antenna, an inter-satellite receiving antenna, a downlink transmitting antenna and a low-orbit satellite receiving antenna of the Beidou navigation satellite under each satellite body coordinate system are respectively set; rA, rB and raAre respectively Beidou navigation satellite A, Beidou navigation satellite B and low orbit satellite a in the center of the earthA position vector of a linear coordinate system.
9. The autonomous orbit determination method of Beidou navigation constellation based on low-orbit satellite based anchoring according to claim 1, characterized in that the specific method for judging whether the inter-Beidou satellite double-satellite pseudorange observation is the over-ionosphere observation in the step (3.3) is as follows:
judging whether an included angle ∠β formed by the connecting lines of the Beidou navigation satellites A and B and the connecting line of the Beidou navigation satellite B and the earth mass center is larger than an included angle ∠α formed by the connecting line of the Beidou navigation satellite B and the earth mass center, if so, the inter-satellite double-satellite pseudo range observation is non-ionosphere observation, otherwise, the inter-satellite double-satellite pseudo range observation is ionosphere observation;
wherein R isionoRepresents the radius of the ionosphere; r isA、rBAnd raRespectively, the position vectors of the beidou navigation satellite A, B in the earth-centered inertial coordinate system.
10. The low-orbit satellite-based anchored Beidou navigation constellation autonomous orbit determination method according to claim 1, characterized in that said pseudo range P after ionosphere influence eliminationiono-freeComprises the following steps:
wherein: p1And P2Respectively representing pseudo-range observed values corresponding to the dual-frequency navigation signals, f1、f2Respectively, a first carrier frequency and a second carrier frequency of the dual-frequency navigation signal.
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Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20070159382A1 (en) * | 2005-11-16 | 2007-07-12 | In-One Joo | GPS signal repeater and GPS receiver of stationary orbit satellite, and method for positioning stationary orbit satellite using the same |
| US20100103041A1 (en) * | 2008-10-24 | 2010-04-29 | Seiko Epson Corporation | Method of determining reliability of long-term predicted orbit data, method of providing long-term predicted orbit data, and information providing apparatus |
| CN104048664A (en) * | 2014-07-01 | 2014-09-17 | 南京航空航天大学 | Autonomous orbit determination method of navigation satellite constellation |
| CN105468882A (en) * | 2014-07-28 | 2016-04-06 | 航天恒星科技有限公司 | Satellite automatic orbit determination method and system |
| CN106338753A (en) * | 2016-09-22 | 2017-01-18 | 北京航空航天大学 | Geosynchronous orbit constellation orbit determination method based on ground station/satellite link/GNSS combined measurement |
| CN107421550A (en) * | 2017-07-25 | 2017-12-01 | 北京航空航天大学 | A kind of earth Lagrange joint constellation autonomous orbit determination methods based on H_2O maser |
-
2019
- 2019-09-27 CN CN201910927460.1A patent/CN110793528B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20070159382A1 (en) * | 2005-11-16 | 2007-07-12 | In-One Joo | GPS signal repeater and GPS receiver of stationary orbit satellite, and method for positioning stationary orbit satellite using the same |
| US20100103041A1 (en) * | 2008-10-24 | 2010-04-29 | Seiko Epson Corporation | Method of determining reliability of long-term predicted orbit data, method of providing long-term predicted orbit data, and information providing apparatus |
| CN104048664A (en) * | 2014-07-01 | 2014-09-17 | 南京航空航天大学 | Autonomous orbit determination method of navigation satellite constellation |
| CN105468882A (en) * | 2014-07-28 | 2016-04-06 | 航天恒星科技有限公司 | Satellite automatic orbit determination method and system |
| CN106338753A (en) * | 2016-09-22 | 2017-01-18 | 北京航空航天大学 | Geosynchronous orbit constellation orbit determination method based on ground station/satellite link/GNSS combined measurement |
| CN107421550A (en) * | 2017-07-25 | 2017-12-01 | 北京航空航天大学 | A kind of earth Lagrange joint constellation autonomous orbit determination methods based on H_2O maser |
Non-Patent Citations (2)
| Title |
|---|
| QIAO JING ET AL.: "BDS Autonomous Orbit Determination using Inter-satellite Tracking and Onboard Accelerometers", 《 PROCEEDINGS OF THE INSTITUTE OF NAVIGATION PACIFIC PNT MEETING》 * |
| 朱俊等: "基于星间测距和地面发射源的导航星座整网自主定轨", 《国防科技大学学报》 * |
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