CN111337957B - Autonomous Integrity Monitoring Method and System for Spaceborne Navigation Receiver - Google Patents
Autonomous Integrity Monitoring Method and System for Spaceborne Navigation Receiver Download PDFInfo
- Publication number
- CN111337957B CN111337957B CN202010276727.8A CN202010276727A CN111337957B CN 111337957 B CN111337957 B CN 111337957B CN 202010276727 A CN202010276727 A CN 202010276727A CN 111337957 B CN111337957 B CN 111337957B
- Authority
- CN
- China
- Prior art keywords
- satellite
- target
- navigation
- instantaneous
- information
- 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.)
- Active
Links
Images
Classifications
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01S—RADIO DIRECTION-FINDING; RADIO NAVIGATION; DETERMINING DISTANCE OR VELOCITY BY USE OF RADIO WAVES; LOCATING OR PRESENCE-DETECTING BY USE OF THE REFLECTION OR RERADIATION OF RADIO WAVES; ANALOGOUS ARRANGEMENTS USING OTHER WAVES
- G01S19/00—Satellite radio beacon positioning systems; Determining position, velocity or attitude using signals transmitted by such systems
- G01S19/01—Satellite radio beacon positioning systems transmitting time-stamped messages, e.g. GPS [Global Positioning System], GLONASS [Global Orbiting Navigation Satellite System] or GALILEO
- G01S19/13—Receivers
- G01S19/23—Testing, monitoring, correcting or calibrating of receiver elements
Landscapes
- Engineering & Computer Science (AREA)
- Radar, Positioning & Navigation (AREA)
- Remote Sensing (AREA)
- Computer Networks & Wireless Communication (AREA)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Position Fixing By Use Of Radio Waves (AREA)
Abstract
The invention provides an autonomous integrity monitoring method and system of a satellite-borne navigation receiver, which are applied to the satellite-borne navigation receiver and comprise the following steps: acquiring navigation information of a target satellite; calculating the instantaneous orbit root of the current epoch of the target satellite based on the navigation information; acquiring instantaneous orbit roots of a plurality of historical epochs of a target satellite; fitting a curve of the change of the orbit root of the target satellite along with time based on the instantaneous orbit roots of the plurality of historical epochs to obtain a target time curve; and carrying out autonomous integrity monitoring based on the instantaneous track number and the target time curve of the current epoch to obtain a monitoring result. The method and the device solve the technical problem that the fault cannot be effectively found and the abnormal resolving cannot be realized under the condition that the signals of less than 5 visible navigation satellites are received in the prior art.
Description
Technical Field
The invention relates to the technical field of satellite navigation, in particular to an autonomous integrity monitoring method and system of a satellite-borne navigation receiver.
Background
Integrity monitoring of a satellite-borne navigation receiver refers to the ability of the system to alert a user in time to terminate the signal when the navigation system fails or the error exceeds a specified navigation task, which reflects the accuracy of the current computed position result and alerts that it is abnormal when the position error exceeds a given threshold.
Autonomous integrity monitoring is one of important means for ensuring the integrity of a satellite navigation system and is also a key technology of a navigation receiver. The Integrity of a positioning result of a user is monitored according to a redundancy observation value of a user Receiver in the conventional Receiver Autonomous Integrity Monitoring (RAIM), and the primary purpose of the Integrity Monitoring is to monitor a failed satellite navigation signal in a navigation process so as to guarantee navigation positioning accuracy. In modern navigation receivers, autonomous integrity monitoring is also used to discover solution result anomalies to inform the user of corresponding responses. In order to enable receiver autonomous integrity monitoring, redundant observations are necessary. Generally, more than 5 visible satellites are needed for integrity detection; more than 6 satellites are needed to be possible to identify a faulty satellite. An enhanced version of RAIM is RAIM-FDE (FDE: Fault Detection exception), a Fault Detection and troubleshooting technique, which must use a minimum of 6 visible satellites because it is necessary to troubleshoot the faulty satellites. The RAIM algorithm is very important for applications with strict requirements on security. There are different implementation algorithms for RAIM: a pseudo-range residual error decision method, a pseudo-range comparison method, a check vector method and a maximum solution separation method.
The existing method needs at least 5 visible navigation satellite signals to effectively find faults and solve the abnormity, and for a satellite-borne navigation receiver, particularly a medium and high orbit navigation receiver, 5 navigation satellites cannot be ensured due to factors such as geometrical relation, earth shielding and insufficient sensitivity. Therefore, the prior art has the technical problem that the fault cannot be effectively found and the abnormal calculation cannot be effectively carried out under the condition that the signals of less than 5 visible navigation satellites are received.
Disclosure of Invention
In view of the above, the present invention provides a method and a system for monitoring autonomous integrity of a satellite-borne navigation receiver, so as to alleviate the technical problem that a fault cannot be effectively found and an anomaly cannot be resolved when signals of less than 5 visible navigation satellites are received in the prior art.
In a first aspect, an embodiment of the present invention provides an autonomous integrity monitoring method for a satellite navigation receiver, which is applied to the satellite navigation receiver, and includes: acquiring navigation information of a target satellite; the target satellite is a satellite carried by the satellite-borne navigation receiver, and the navigation information includes: position information and velocity information; calculating an instantaneous orbital element of a current epoch of the target satellite based on the navigation information; acquiring instantaneous orbit roots of a plurality of historical epochs of the target satellite; fitting a curve of the orbit root of the target satellite changing along with time based on the instantaneous orbit roots of the plurality of historical epochs to obtain a target time curve; performing autonomous integrity monitoring based on the instantaneous orbit number of the current epoch and the target time curve to obtain a monitoring result; the monitoring result comprises any one of the following items: the navigation information is not abnormal, and the navigation information is abnormal.
Further, the instantaneous track root of the current epoch includes at least one of: semi-major axis of the track, eccentricity of the track, inclination angle of the track surface, right ascension of the ascending intersection point, angle distance of the near point and angle of the average near point.
Further, acquiring navigation information of the target satellite comprises: acquiring a navigation signal of a target satellite; the navigation signal is a signal from a navigation satellite; resolving position information and velocity information of the target satellite based on the navigation signal; the position information is three-dimensional position information of the target satellite, and the speed information is three-dimensional speed information of the target satellite; and determining the position information and the speed information as navigation information of the target satellite.
Further, fitting a curve of the orbital element of the target satellite changing with time based on the instantaneous orbital elements of the plurality of historical epochs to obtain a target time curve, including: performing precise orbit determination on the target satellite based on the instantaneous orbit roots of the plurality of historical epochs to obtain initial orbit information of the target satellite; and performing orbit forecasting on the target satellite based on the initial orbit information and the orbit dynamics model to obtain a target time curve of the change of the orbit root of the target satellite along with time.
Further, fitting a curve of the orbital element of the target satellite changing with time based on the instantaneous orbital elements of the plurality of historical epochs to obtain a target time curve, including: fourier transform is carried out on the orbital elements of the target satellite to obtain a target Fourier series; the target Fourier series comprises a plurality of fitting parameters; determining values of the plurality of fitting parameters based on instantaneous orbit roots of the plurality of historical epochs; substituting the values of the fitting parameters into the target Fourier series to obtain a target time curve.
Further, based on the instantaneous orbit number of the current epoch and the target time curve, performing autonomous integrity monitoring to obtain a monitoring result, including: acquiring the fitted instantaneous orbit number of the current epoch on the target time curve; calculating the difference between the instantaneous orbit root of the current epoch and the fitted instantaneous orbit root; judging whether the difference value is larger than a preset threshold or not; if so, obtaining a monitoring result that the navigation information is abnormal; if not, the monitoring result is that the navigation information is not abnormal.
In a second aspect, an embodiment of the present invention further provides an autonomous integrity monitoring system for a satellite-borne navigation receiver, including: the satellite-borne navigation receiver is used for acquiring navigation information of a carried target satellite; the navigation information includes: position information and velocity information; calculating an instantaneous orbital element of a current epoch of the target satellite based on the navigation information; the ground operation and control center is used for acquiring instantaneous orbit roots of a plurality of historical epochs of the target satellite; fitting a time-varying curve of the orbit root of the target satellite based on the instantaneous orbit roots of the plurality of historical epochs to obtain a target time curve; the satellite-borne navigation receiver is further used for carrying out autonomous integrity monitoring based on the instantaneous orbit number of the current epoch and the target time curve to obtain a monitoring result; the monitoring result comprises any one of the following items: the navigation information is not abnormal, and the navigation information is abnormal.
Further, the system also comprises a satellite-ground communication device used for carrying out data transmission between the satellite-borne navigation receiver and the ground operation and control center.
In a third aspect, an embodiment of the present invention further provides a satellite navigation receiver, including: the device comprises a navigation signal receiving unit, an instantaneous ephemeris calculation unit, a historical data storage unit, an ephemeris fitting calculation unit and an integrity comparison unit, wherein the navigation signal receiving unit is used for acquiring navigation information of a target satellite; the real-time target satellite is a satellite carried by the satellite-borne navigation receiver, and the navigation information comprises: position information and velocity information; the instantaneous ephemeris calculation unit is used for calculating the instantaneous orbit root of the current epoch of the target satellite based on the navigation information; the historical data storage unit is used for acquiring the instantaneous orbit roots of a plurality of historical epochs of the target satellite; the ephemeris fitting calculation unit is used for fitting a time-varying curve of the orbit root of the target satellite based on the instantaneous orbit roots of the plurality of historical epochs to obtain a target time curve; the integrity comparison unit is used for carrying out autonomous integrity monitoring based on the instantaneous track number of the current epoch and the target time curve to obtain a monitoring result; the monitoring result comprises any one of the following items: the navigation information is not abnormal, and the navigation information is abnormal.
Further, the navigation signal receiving unit is further configured to: acquiring a navigation signal of a target satellite; the navigation signal is a signal from a navigation satellite; resolving position information and velocity information of the target satellite based on the navigation signal; the position information is three-dimensional position information of the target satellite, and the speed information is three-dimensional speed information of the target satellite; and determining the position information and the speed information as navigation information of the target satellite.
The embodiment of the invention provides an autonomous integrity monitoring method and system for a satellite-borne navigation receiver, which jointly transform a position information and speed information resolving result into an instantaneous orbit number of a target satellite, and discover resolving abnormality by using a change rule of the orbit number of the target satellite, so that the autonomous integrity monitoring of the satellite-borne navigation receiver can be carried out without the minimum requirement of 5 navigation satellite signals of a conventional method, and the capability of the satellite-borne navigation receiver for discovering the abnormal navigation result is improved.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, and it is obvious that the drawings in the following description are some embodiments of the present invention, and other drawings can be obtained by those skilled in the art without creative efforts.
Fig. 1 is a flowchart of an autonomous integrity monitoring method for a satellite-borne navigation receiver according to an embodiment of the present invention;
FIG. 2 is a schematic diagram of a change of a semi-major axis of an orbit of an actual satellite with time according to an embodiment of the present invention;
FIG. 3 is a schematic diagram illustrating the inclination of the orbital plane of an actual satellite according to the time variation provided by the embodiment of the invention;
FIG. 4 is a schematic error diagram of integrity monitoring using a target time curve of a semi-major axis of a track according to an embodiment of the present invention;
FIG. 5 is a schematic error diagram illustrating integrity monitoring using a target time curve of orbital plane inclination angles according to an embodiment of the present invention;
FIG. 6 is a graph illustrating how the root semi-major axis changes with time according to an embodiment of the present invention;
FIG. 7 is a schematic error diagram illustrating integrity monitoring using a target time curve of a root track semimajor axis according to an embodiment of the present invention;
FIG. 8 is a schematic view of another exemplary embodiment of a track surface inclination angle variation with time;
FIG. 9 is a schematic error diagram of integrity monitoring using a target time curve of a track surface inclination angle according to an embodiment of the present invention;
FIG. 10 is a schematic diagram of an autonomous integrity monitoring system for a satellite navigation receiver according to an embodiment of the present invention;
fig. 11 is a schematic diagram of a satellite navigation receiver according to an embodiment of the present invention.
Detailed Description
The technical solutions of the present invention will be described clearly and completely with reference to the accompanying drawings, and it should be understood that the described embodiments are some, but not all embodiments of the present invention. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
The first embodiment is as follows:
fig. 1 is a flowchart of an autonomous integrity monitoring method for a satellite navigation receiver, which is applied to the satellite navigation receiver according to an embodiment of the present invention. As shown in fig. 1, the method specifically includes the following steps:
step S102, acquiring navigation information of a target satellite; the target satellite is a satellite carried by a satellite-borne navigation receiver, and the navigation information comprises: position information and velocity information.
Specifically, firstly, a navigation signal of a target satellite is obtained, wherein the navigation signal is a signal from a navigation satellite; then, based on the navigation signal, the position information and the speed information of the target satellite are calculated, wherein the position information is the three-dimensional position information of the target satellite, and the speed information is the three-dimensional speed information of the target satellite; and finally, determining the position information and the speed information as the navigation information of the target satellite.
In the embodiment of the invention, the navigation information of the target satellite can be solved only by acquiring the navigation signals of at least 4 visible navigation satellites.
And step S104, calculating the instantaneous orbit root of the current epoch of the target satellite based on the navigation information.
And step S106, acquiring the instantaneous orbit roots of a plurality of historical epochs of the target satellite.
And S108, fitting a curve of the orbit number of the target satellite changing along with time based on the instantaneous orbit numbers of the plurality of historical epochs to obtain a target time curve.
Step S110, performing autonomous integrity monitoring based on the instantaneous track number and the target time curve of the current epoch to obtain a monitoring result; the monitoring result comprises any one of the following items: the navigation information is not abnormal, and the navigation information is abnormal.
The invention provides an autonomous integrity monitoring method of a satellite-borne navigation receiver, which jointly converts a position information and speed information resolving result into an instantaneous orbit number of a target satellite, and discovers resolving abnormality by using a change rule of the orbit number of the target satellite, so that the minimum requirement of 5 navigation satellite signals of a conventional method is not needed, and the autonomous integrity monitoring of the satellite-borne navigation receiver can be carried out only by completing the resolving of the target position information and the speed information by 4 navigation satellite signals at minimum. The method and the device solve the technical problems that faults cannot be effectively found and the resolving abnormality cannot be effectively found under the condition that signals of less than 5 visible navigation satellites are received in the prior art, and improve the capability of a satellite-borne navigation receiver for finding the navigation resolving result abnormality.
In an embodiment of the invention, the instantaneous track root of the current epoch includes at least one of: semi-major axis a of the tracksEccentricity of track esInclination angle i of the orbital plane, right ascension omega of the intersection point, and angular distance omega of the near placesMean anomaly angle M.
In particular, the Semi-major Axis of the orbit (Semi-major Axis) asEccentricity of track esThe two parameters determine the shape and size of the keplerian ellipse; the inclination angle i of the orbital plane is an included angle between the orbital plane of the satellite and the equatorial plane of the earth, the right ascension channel omega of the ascension intersection point is an geocentric included angle between the ascension intersection point and the spring minute point of the equatorial plane of the earth, and the relative orientation between the orbital plane of the satellite and the earth sphere is uniquely determined by two parameters i and omega; angular distance omega from near to earthsIs the center-of-earth angle, omega, between the point of intersection and the point of approach on the orbital planesExpresses the orientation of the keplerian ellipse in the orbital plane; the mean anomaly M of the satellite is a function of the time-uniform variation related to the average angular velocity of motion, enabling the determination of the instantaneous position of the satellite in orbit.
In the embodiment of the invention, a curve of the orbital element number of the target satellite changing along with time can be fitted through different implementation modes to obtain a target time curve. For example, the target time curve can be obtained by using an orbit prediction method or a fourier series method.
Optionally, obtaining the target time curve by using a method of orbit prediction includes:
performing precise orbit determination on the target satellite based on the instantaneous orbit roots of a plurality of historical epochs to obtain initial orbit information of the target satellite;
and performing orbit prediction on the target satellite based on the initial orbit information and the orbit dynamics model to obtain a target time curve of the change of the orbit number of the target satellite along with time.
Specifically, the core problem of numerical method orbit prediction is to solve an initial value problem of an ordinary differential equation, where an orbit state quantity X is set to { a, e, i, ω, Ω, M }, and a corresponding equation differential form is:
wherein, FeThe gravity of the earth center can be obtained based on the position information of the target satellite; f (X, t) is perturbation force; t is t0，X0Indicating the track state quantity at an initial time t0When the corresponding value is X0(ii) a Solving equation (1) by adopting a numerical method, the value of X on the discrete node can be obtained, namely corresponding to { t; (ii) a 1, 2, N } of Xl：X1，X2，...，XN. For example, for an actual satellite with an orbital altitude of about 700km, the actual satellite is precisely orbited by using the instantaneous orbital element of the historical epoch of the previous 24 hours to obtain initial orbital information of the actual satellite, then the initial orbital information is substituted into the formula (1) to perform orbital prediction on the actual satellite, ephemeris of the actual satellite for the next 12 hours is determined, and finally a target time curve of the actual satellite is obtained. Fig. 2 is a graph illustrating changes in the orbital semimajor axis of an actual satellite according to an embodiment of the present invention with time, and fig. 3 is a graph illustrating changes in the orbital plane inclination of an actual satellite according to an embodiment of the present invention with time.
And then, carrying out autonomous integrity monitoring based on the instantaneous track number of the current epoch and a target time curve to obtain a monitoring result. Specifically, the method comprises the following steps:
step 1101, acquiring the number of fitted instantaneous orbits on a target time curve relative to a current epoch;
step S1102, calculating the difference between the instantaneous orbit number of the current epoch and the fitted instantaneous orbit number;
step S1103, judging whether the difference value is greater than a preset threshold;
step S1104, if yes, the monitoring result is obtained that the navigation information is abnormal;
step S1105, if not, the monitoring result is that the navigation information is not abnormal.
In the embodiment of the invention, the position information and speed information resolving results are jointly converted into the instantaneous orbit number of the target satellite, and the resolving abnormality is found by utilizing the change rule of the orbit number of the target satellite, namely, if the judgment difference value is greater than the preset threshold, the resolved navigation information received by satellite-borne navigation is determined to be abnormal, and the autonomous integrity monitoring fails.
For example, according to the method provided by the embodiment of the present invention, the instantaneous orbit number of the current epoch of the actual satellite is calculated, and then the error (i.e. the difference) between the instantaneous orbit number of the current epoch (e.g. the orbit semi-major axis and the orbit plane inclination angle) and the target time curve is calculated, and when the error between any orbit number and the target time curve exceeds a preset threshold (e.g. 3 times of standard deviation), the point is determined as an abnormal point, and the autonomous integrity monitoring fails; otherwise autonomous integrity monitoring passes. Fig. 4 is a schematic diagram of an error of integrity monitoring using a target time curve of a semi-major axis of a track according to an embodiment of the present invention, and fig. 5 is a schematic diagram of an error of integrity monitoring using a target time curve of a track surface inclination angle according to an embodiment of the present invention. The open dots in fig. 4 and 5 represent outlier points where the error exceeds a preset threshold.
Optionally, obtaining the target time curve by using a fourier series method includes:
fourier transform is carried out on the orbital elements of the target satellite to obtain target Fourier series; the target Fourier series comprises a plurality of fitting parameters;
determining values of a plurality of fitting parameters based on instantaneous orbit roots of a plurality of historical epochs;
and substituting values of the plurality of fitting parameters into the target Fourier series to obtain a target time curve.
For example, for a semi-major axis of the track, the curve after fourier transformation can be expressed as:
thereby obtaining the root semi-long shaft,
wherein, a, b, w and
is a fitting parameter and is an unknown number. Then, based on the instantaneous track number values of multiple historical epochs, the sum of a, b, w and a is easily calculated by a general optimization method
For example, for an actual satellite with an orbital altitude of about 700km, fitting calculation is performed using the values of the instantaneous orbital roots of the historical epochs from the previous 3 hours, and the calculated fitting parameters are a-2677.332636, b-1.676301, w-0.002079,
fig. 6 is a graph illustrating the variation of the root semi-major axis with time according to an embodiment of the present invention. Then, according to the target time curve in fig. 6 and the instantaneous orbit root of the current epoch of the actual satellite, autonomous integrity monitoring is performed, for example, a preset threshold is determined to be 3 times of a standard deviation, and an error schematic diagram of integrity monitoring performed by using the target time curve of the root orbit semi-major axis in fig. 7 is obtained. As shown in fig. 7, the open circles in the graph represent outliers where the error exceeds a preset threshold.
For example, for orbital plane inclination angle i, the curve after fourier transform can be represented as:
wherein, c, d, e, w1，w2，
And
are fitting parameters and are all unknowns. Then, based on the value of i in the instantaneous track numbers of a plurality of historical epochs, c, d, e, w are easily calculated by a general optimization method1，w2，
And
for example, a fitting calculation is performed using the value of i in the instantaneous orbit root of a 14-hour historical epoch for an actual satellite having an orbit height of about 700km, and the calculated fitting parameters are d-1.721010, e-0.000095, c-0.000031, and w1＝0.002078，w2＝0.000140，
Fig. 8 is a schematic view of another track surface inclination angle variation with time according to the embodiment of the present invention. Then, according to the target time curve in fig. 8 and the instantaneous orbit number of the current epoch of the actual satellite, autonomous integrity monitoring is performed, for example, a preset threshold is determined to be 3 times of a standard deviation, and an error diagram of integrity monitoring performed by using the target time curve of the orbital plane inclination angle in fig. 9 is obtained. As shown in fig. 9, the open circles in the graph represent outliers where the error exceeds a preset threshold.
Optionally, the method provided in the embodiment of the present invention further includes: and if the monitoring result is that the navigation information is abnormal, sending an alarm signal.
Example two:
sometimes, the satellite-borne navigation receiver is not enough in calculation force and is not suitable for directly performing fitting calculation on a satellite, and the fitting calculation can be placed on the ground. In view of this, the embodiment of the present invention further provides an autonomous integrity monitoring system for a satellite navigation receiver, which can place fitting calculation in a ground operation and control center.
FIG. 10 is a schematic diagram of an autonomous integrity monitoring system of a satellite navigation receiver according to an embodiment of the present invention. As shown in fig. 10, the system includes: a satellite navigation receiver 100 and a ground operation and control center 200.
Specifically, the satellite navigation receiver 100 is configured to obtain navigation information of a target satellite to be carried; the navigation information includes: position information and velocity information; an instantaneous orbital element of a current epoch of the target satellite is calculated based on the navigation information.
In the embodiment of the invention, the navigation information of the target satellite can be solved only by acquiring the navigation signals of at least 4 visible navigation satellites.
The ground operation and control center 200 is used for acquiring the instantaneous orbit number of a plurality of historical epochs of the target satellite; and fitting a time-varying curve of the orbit root of the target satellite based on the instantaneous orbit roots of the historical epochs to obtain a target time curve.
The satellite-borne navigation receiver 100 is further configured to perform autonomous integrity monitoring based on the instantaneous orbit number of the current epoch and the target time curve to obtain a monitoring result; the monitoring result comprises any one of the following items: the navigation information is not abnormal, and the navigation information is abnormal.
Optionally, as shown in fig. 10, the system further includes a satellite-ground communication device 300 for data transmission between the satellite-borne navigation receiver and the ground operation and control center.
According to the autonomous integrity monitoring system of the satellite-borne navigation receiver, the process of fitting the target time curve is issued to the ground operation and control center, the operation pressure of a satellite on the navigation receiver is relieved, and the stability of the autonomous integrity monitoring process is improved.
Example three:
fig. 11 is a schematic diagram of an on-board navigation receiver according to an embodiment of the present invention, and as shown in fig. 11, the on-board navigation receiver includes: the system comprises a navigation signal receiving unit 10, an instant ephemeris calculating unit 20, a historical data storing unit 30, an ephemeris fitting calculating unit 40 and an integrity comparing unit 50.
Specifically, a navigation signal receiving unit 10 for acquiring navigation information of a target satellite; the real-time target satellite is a satellite carried by a satellite-borne navigation receiver, and the navigation information comprises: position information and velocity information.
And an instantaneous ephemeris calculation unit 20, configured to calculate an instantaneous orbital element of the current epoch of the target satellite based on the navigation information. Alternatively, the instantaneous ephemeris calculation unit 20 may calculate the instantaneous orbit root of the current epoch of the satellite-borne navigation receiver itself from the acquired position information and velocity information, and the input of the instantaneous ephemeris calculation unit 20 is the navigation information output by the navigation signal receiving unit 10 and the output is the instantaneous orbit root of the current epoch of the target satellite.
In an embodiment of the invention, the instantaneous track root of the current epoch includes at least one of: semi-major axis a of the tracksEccentricity of track esInclination angle i of the orbital plane, right ascension omega of the intersection point, and angular distance omega of the near placesMean anomaly angle M.
And the historical data storage unit 30 is used for acquiring the instantaneous orbit roots of a plurality of historical epochs of the target satellite. Optionally, the historical data storage unit 30 is further configured to store the instantaneous orbit number of the historical epoch, which is input as the instantaneous orbit number output by the instantaneous ephemeris calculation unit 20 and output as the instantaneous orbit number of the stored historical epoch.
And the ephemeris fitting calculation unit 40 is configured to fit a time-varying curve of the orbit root of the target satellite based on the instantaneous orbit roots of the plurality of historical epochs to obtain a target time curve. Specifically, the ephemeris fitting calculation unit 40 has an input of the instantaneous orbit number of the history epoch output from the history data storage unit 30, and outputs a target time curve of each orbit number fitted.
The integrity comparison unit 50 is used for carrying out autonomous integrity monitoring based on the instantaneous track number of the current epoch and the target time curve to obtain a monitoring result; the monitoring result comprises any one of the following items: the navigation information is not abnormal, and the navigation information is abnormal.
Specifically, the integrity comparison unit 50 makes a difference between the instantaneous track number of the current epoch and the target time curve of the track number, and when the error exceeds a preset threshold, the autonomous integrity monitoring fails; otherwise autonomous integrity monitoring passes. The input of the system is the instantaneous orbit number of the current epoch output by the instantaneous ephemeris calculation unit 20 and the target time curve output by the ephemeris fitting calculation unit 40, and the output is the monitoring result.
The invention provides a satellite-borne navigation receiver, which jointly converts a position information and speed information resolving result into an instantaneous orbit number of a target satellite, and discovers resolving abnormality by utilizing a change rule of the orbit number of the target satellite, so that the minimum requirement of 5 navigation satellite signals of a conventional method is not needed, and the autonomous integrity monitoring of the satellite-borne navigation receiver can be carried out only by completing the resolving of the target position information and the speed information by 4 navigation satellite signals at minimum. The method and the device solve the technical problems that faults cannot be effectively found and abnormal resolving cannot be achieved when signals of less than 5 visible navigation satellites are received in the prior art, and improve the capability of a satellite-borne navigation receiver for finding the abnormal navigation results.
Optionally, the navigation signal receiving unit 10 is further configured to:
acquiring a navigation signal of a target satellite; the navigation signal is a signal from a navigation satellite; resolving position information and speed information of the target satellite based on the navigation signal; the position information is three-dimensional position information of a target satellite, and the speed information is three-dimensional speed information of the target satellite; the position information and the velocity information are determined as navigation information of the target satellite.
In the embodiment of the invention, the navigation information of the target satellite can be solved only by acquiring the navigation signals of at least 4 visible navigation satellites.
Optionally, the integrity comparison unit 50 is further configured to: acquiring the fitted instantaneous orbit number related to the current epoch on the target time curve; calculating the difference between the instantaneous orbit root of the current epoch and the fitted instantaneous orbit root; judging whether the difference value is greater than a preset threshold or not; if so, obtaining a monitoring result that the navigation information is abnormal; if not, the navigation information is not abnormal according to the obtained monitoring result.
Optionally, the satellite-borne navigation receiver provided in the embodiment of the present invention further includes: and the warning unit is used for sending a warning signal when the monitoring result is that the navigation information is abnormal.
Finally, it should be noted that: the above embodiments are only used to illustrate the technical solution of the present invention, and not to limit the same; while the invention has been described in detail and with reference to the foregoing embodiments, it will be understood by those skilled in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some or all of the technical features may be equivalently replaced; and the modifications or the substitutions do not make the essence of the corresponding technical solutions depart from the scope of the technical solutions of the embodiments of the present invention.
Claims (5)
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010276727.8A CN111337957B (en) | 2020-04-09 | 2020-04-09 | Autonomous Integrity Monitoring Method and System for Spaceborne Navigation Receiver |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010276727.8A CN111337957B (en) | 2020-04-09 | 2020-04-09 | Autonomous Integrity Monitoring Method and System for Spaceborne Navigation Receiver |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN111337957A CN111337957A (en) | 2020-06-26 |
| CN111337957B true CN111337957B (en) | 2021-01-01 |
Family
ID=71184596
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202010276727.8A Active CN111337957B (en) | 2020-04-09 | 2020-04-09 | Autonomous Integrity Monitoring Method and System for Spaceborne Navigation Receiver |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN111337957B (en) |
Families Citing this family (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112327333B (en) * | 2020-10-26 | 2024-03-29 | 陕西星邑空间技术有限公司 | Satellite position calculation method and device |
| CN114764145A (en) * | 2021-01-15 | 2022-07-19 | 千寻位置网络有限公司 | Method, device and equipment for determining abnormal reason of satellite monitoring data |
| CN116184454B (en) * | 2023-02-08 | 2024-03-12 | 国家卫星海洋应用中心 | Satellite orbit parameter determination method, device, equipment and readable storage medium |
| CN118169718A (en) * | 2024-03-14 | 2024-06-11 | 长光卫星技术股份有限公司 | A satellite precise tracking method using Beidou short message to send real-time orbit prediction |
Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109521448A (en) * | 2018-12-18 | 2019-03-26 | 清华大学 | Satellite-based navigation receiver positioning time service method and device based on orbital tracking prediction |
| CN110007317A (en) * | 2019-04-10 | 2019-07-12 | 南京航空航天大学 | An Advanced Receiver Autonomous Integrity Monitoring Method for Star Selection Optimization |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9726764B1 (en) * | 2009-12-07 | 2017-08-08 | Rockwell Collins, Inc. | System and mehtod for providing space-based precision position correlations for promoting improved availability, accuracy and integrity |
| CN102401903A (en) * | 2010-09-17 | 2012-04-04 | 郑州威科姆科技股份有限公司 | Autonomous integrity implementation method for Beidou second-generation receiver |
-
2020
- 2020-04-09 CN CN202010276727.8A patent/CN111337957B/en active Active
Patent Citations (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109521448A (en) * | 2018-12-18 | 2019-03-26 | 清华大学 | Satellite-based navigation receiver positioning time service method and device based on orbital tracking prediction |
| CN110007317A (en) * | 2019-04-10 | 2019-07-12 | 南京航空航天大学 | An Advanced Receiver Autonomous Integrity Monitoring Method for Star Selection Optimization |
Also Published As
| Publication number | Publication date |
|---|---|
| CN111337957A (en) | 2020-06-26 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN111337957B (en) | Autonomous Integrity Monitoring Method and System for Spaceborne Navigation Receiver | |
| CN107449443B (en) | Integrity monitoring of radar altimeter | |
| EP3006900B1 (en) | Systems and methods for attitude fault detection based on integrated gnss-inertial hybrid filter residuals | |
| EP2887097B1 (en) | Architectures for high Navigation Satellite System integrity multi-constellation solution separation | |
| RU2388661C2 (en) | Method to control aircraft engine | |
| JP4783394B2 (en) | Sensor failure adaptation method | |
| EP3588225A1 (en) | Methods and systems for detecting data anomalies | |
| US9694902B2 (en) | Device for making available navigation parameter values of a vehicle | |
| CA3079279A1 (en) | Reducing bias impact on gnss integrity | |
| US7592952B2 (en) | Method of locating a vehicle from satellite data | |
| CN101833101A (en) | Completeness or adequateness monitoring method and device based on local area augmentation system (LAAS) | |
| KR101503001B1 (en) | System and method for determining fault of gnss using estimated value of baseline length and ground facility antenna of multiple gnss | |
| EP4089497B1 (en) | Methods and systems for depicting avionics data anomalies | |
| EP3346282A1 (en) | Electronic monitoring device for monitoring at least one radionavigation signal during an approach phase to a landing runway, related monitoring method and computer program | |
| CA2337829C (en) | Gps satellite drift monitor | |
| CN106601032B (en) | A kind of multipath landform integrality detection method based on lower view sensor | |
| JP2002520623A5 (en) | ||
| US20070113629A1 (en) | Method for detecting errors in sensor values and error detection device | |
| EP3964795B1 (en) | Fault detection, exclusion, isolation, and re-configuration of navigation sensors using an abstraction layer | |
| US8108146B2 (en) | Method and device for determining a consolidated position of a traveling object, particularly an aircraft | |
| CN104965209A (en) | Method, device and system for calculating actual navigation performance | |
| CN108717197A (en) | The computer programmed product and receiver that navigation solution protection class determines method, is associated | |
| CN114397677B (en) | Faulty satellite detection method at receiver end based on non-parametric estimation | |
| US8514127B2 (en) | Method and system of calculation for the evaluation of the precision performance of a satellite navigation system | |
| EP4214532B1 (en) | Method for checking the integrity of calibration on a plurality of landmarks, associated computer program product and associated integrity checking device |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |