Star-based enhanced satellite orbit correction calculation method and system based on Hille difference

Abstract

The invention discloses a satellite-based enhanced satellite orbit correction calculation method and system based on Hill difference, and relates to the technical field of satellite navigation wide-area difference, and the method comprises the steps: determining a monitoring station which observes a target satellite; smoothing the pseudo-range observed quantity by using the carrier phase observed quantity of the GNSS satellite j observed by each monitoring station; based on the smoothed pseudo-range observed quantity, ionospheric delay is eliminated; calculating the ephemeris position and the satellite clock deviation of a GNSS satellite j by using the orbit parameter and the clock parameter broadcasted by the navigation message, and calculating the inter-satellite distance by combining the position of the monitoring station i; based on the calculated distance between the stations and satellites, the influence of satellite clock errors on satellite orbit errors is eliminated, and a state model of the satellite orbit errors is established based on a Hille difference equation; and resolving the state model to obtain the satellite orbit correction, and the method can improve the precision of the satellite orbit differential correction.

Description

Technical Field

[0001] The present application relates to the field of satellite navigation wide-area differential technology, and in particular to a method and system for calculating satellite-based augmented satellite orbit correction numbers based on Hill differential. Background Art

[0002] The orbital error of the Global Navigation Satellite System (GNSS) satellite is mainly caused by the distribution and coordinate error of the monitoring stations in the ground control part, the accuracy of the observations obtained by the monitoring stations tracking the satellites, the accuracy of the perturbation dynamic model of the satellites, etc. The Satellite Based Augmentation System (SBAS) estimates the satellite orbit error and generates the corresponding satellite orbit correction number to reduce the impact of the satellite orbit error on the user's positioning accuracy.

[0003] GNSS satellites rotate around the earth periodically at a constant angular velocity in space. The disturbance force they experience during movement is also periodic, and the disturbance force model used to solve the satellite orbit is also fixed, which causes the change of satellite orbit error to have a certain periodicity.

[0004] At present, the public satellite-based augmentation satellite orbit correction number calculation methods all use the least squares method or the Kalman filter method based on the velocity model. The existing satellite-based augmentation satellite orbit correction number calculation methods currently fail to establish a motion model based on the characteristics of the satellite orbit error, and simply use the least squares method or the Kalman filter method based on the velocity model for calculation. The calculated satellite orbit correction number cannot effectively reflect the true characteristics of the satellite orbit error. Summary of the invention

[0005] The embodiments of the present application provide a method and system for calculating satellite-based augmented satellite orbit correction numbers based on Hill differential, which is used to establish a dynamic change model of satellite orbit errors using the Hill differential equation. The calculated satellite orbit correction numbers can more realistically reflect the satellite orbit errors and improve the accuracy of satellite orbit differential correction numbers.

[0006] The embodiment of the present application provides a method for calculating satellite orbit correction numbers for satellite-based augmentation based on Hill difference, including:

[0007] Identify the monitoring station that observed the target satellite;

[0008] The pseudo-range observations are smoothed using the carrier phase observations of GNSS satellite j observed by each monitoring station;

[0009] Eliminate ionospheric delay based on smoothed pseudorange observations;

[0010] The orbit parameters and clock parameters broadcast by the navigation message are used to calculate the ephemeris position and satellite clock bias of GNSS satellite j, and the ephemeris distance is calculated in combination with the position of monitoring station i;

[0011] Based on the calculated ephemeris distance, the influence of satellite clock error on satellite orbit error is eliminated, and the state model of satellite orbit error is established based on the Hill difference equation;

[0012] The state model is solved to obtain satellite orbit correction numbers.

[0013] Optionally, the monitoring stations that determine the target satellite are observed include:

[0014] Among the monitoring stations that observe GNSS satellite j, the monitoring stations with elevation angles greater than the altitude cutoff angle are selected, and the total number of monitoring stations is M.

[0015] Optionally, using the carrier phase observations of the GNSS satellite j observed by each monitoring station to smooth the pseudorange observations includes:

[0016] The carrier observation is converted to satisfy:

[0017]

[0018] in, f1 is the frequency of carrier L1, f5 is the frequency of carrier L5; and are the carrier phase observations at the L1 and L5 frequency points at time t respectively;

[0019] use To smooth the noise in the pseudorange observations:

[0020]

[0021] Among them, L k Indicates L1 or L5 frequency point, is the pseudo-range observation value of the corresponding frequency point, is the pseudorange observation after smoothing of the corresponding frequency point, and τ is the smoothing time.

[0022] Optionally, based on the smoothed pseudorange observations, eliminating ionospheric delays includes:

[0023] The pseudo-range observations after smoothing at L1 and L5 frequencies are used to eliminate the ionospheric delay. for:

[0024]

[0025] Optionally, the orbit parameters and clock parameters broadcast by the navigation message are used to calculate the ephemeris position and satellite clock bias of GNSS satellite j, and the inter-station distance is calculated in combination with the position of monitoring station i to satisfy:

[0026]

[0027] in, is the ephemeris position of GNSS satellite j, is the satellite clock bias, is the location of monitoring station i;

[0028] The distance between the stations and satellite clock bias from Eliminate it to obtain the pseudorange residual

[0029]

[0030] Among them, ΔR j (t) = [Δx j (t)Δy j (t)Δz j (t)] T is the orbit error of GNSS satellite j in the X, Y, and Z directions in the Earth-centered Earth-fixed coordinate system; is the unit direction vector from monitoring station i to GNSS satellite j; ΔB j (t) is the clock error of GNSS satellite j; is the residual error, is the variance.

[0031] Optionally, based on the calculated station-satellite distance, eliminating the effect of satellite clock error on satellite orbit error includes:

[0032] The satellite orbit error is eliminated by the inter-station single difference method, and the pseudorange residual is Each value of Subtract:

[0033]

[0034] in, is the residual error, and the variance is

[0035] Convert equation (7) into the corresponding matrix form:

[0036]

[0037] in, V o jThe covariance matrix of

[0038] The state model of satellite orbit error based on the Hill difference equation includes establishing the Hill difference equation describing the satellite orbit error:

[0039] x″(t)-2ωy ′ (t)-3ω 2 x(t)=0 (9)

[0040] y″(t)+2ωx ′ (t)=0 (10)

[0041] z″(t)+ω 2 z(t)=0 (11)

[0042] Among them, x(t), y(t) and z(t) are the satellite orbit errors in the three directions of XYZ in the satellite inertial coordinate system at time t; ′ x(t) and x″(t) are the first and second derivatives of x(t), respectively; y ′ y(t) and y″(t) are the first and second derivatives of y(t), respectively; z ′ z(t) and z″(t) are the first and second derivatives of z(t), respectively; ω is the orbital angular velocity;

[0043] The matrix form of the solution to the Hill difference equation is:

[0044] X(t)=F(t)X0 (12)

[0045] Among them, X(t)=[x(t)y(t)z(t)x ′ (t)y ′ (t)z ′ (t)] T ; X0 = X(0), is the initial value of satellite orbit error;

[0046]

[0047] Then the state equation of the satellite orbit error is:

[0048] X(t n )=F(t n -t n-1 )X(t n-1 ) (13)

[0049] Among them, t n For time series.

[0050] Optionally, solving the state model to obtain satellite orbit corrections includes:

[0051] The observation equation and state equation used to solve the satellite orbit correction number are expressed as:

[0052]

[0053] X r (t n )=F(t n -t n-1 )X r (t n-1 )+W r (t n-1 ) (15)

[0054] in, t n The satellite orbit error at time t n The satellite orbit error change rate at time λ and are the latitude and longitude of the satellite sub-satellite trajectory, β = arccos[cos(i) / cos(λ)], i is the satellite orbit inclination; W r (t n ) is noise, and its covariance matrix is:

[0055] Q o (t n )=diag([0.001 0.001 0.001 0.001 0.001 0.001])

[0056] Use Kalman filtering to calculate the satellite orbit correction number in the satellite inertial coordinate system

[0057]

[0058] P o (t n |t n-1 )=F(t n -t n-1 ) o (t n-1 )F T (t n -t n-1 )+Q o (t n-1 ) (17)

[0059]

[0060] P o (t n )=[IKo (t n )H o (t n )]P o (t n |t n-1 ) (20)

[0061] in, I is the identity matrix;

[0062] The satellite orbit correction number in the Earth Centered Earth Fixed (ECEF) coordinate system can be obtained through the matrix C

[0063] .

[0065] An embodiment of the present application also proposes a satellite-based augmentation satellite orbit correction number solution system based on Hill differential, including a processor and a memory, wherein a computer program is stored in the memory, and when the computer program is executed by the processor, the steps of the satellite-based augmentation satellite orbit correction number solution method based on Hill differential as described above are implemented.

[0066] The embodiment of the present application uses the Hill differential equation to establish a dynamic change model of satellite orbit error. The calculated satellite orbit correction number can more realistically reflect the satellite orbit error and improve the accuracy of the satellite orbit differential correction number.

[0067] The above description is only an overview of the technical solution of the present application. In order to more clearly understand the technical means of the present application, it can be implemented in accordance with the contents of the specification. In order to make the above and other purposes, features and advantages of the present application more obvious and easy to understand, the specific implementation methods of the present application are listed below. BRIEF DESCRIPTION OF THE DRAWINGS

[0068] Various other advantages and benefits will become apparent to those of ordinary skill in the art by reading the detailed description of the preferred embodiments below. The accompanying drawings are only for the purpose of illustrating the preferred embodiments and are not to be considered as limiting the present application. Also, the same reference symbols are used throughout the accompanying drawings to represent the same components. In the accompanying drawings:

[0069] Figure 1 The basic process of the method for calculating satellite orbit correction numbers based on Hill difference of the satellite-based augmentation is shown in the embodiment of the present application. DETAILED DESCRIPTION

[0070] The exemplary embodiments of the present disclosure will be described in more detail below with reference to the accompanying drawings. Although the exemplary embodiments of the present disclosure are shown in the accompanying drawings, it should be understood that the present disclosure can be implemented in various forms and should not be limited by the embodiments set forth herein. On the contrary, these embodiments are provided to enable a more thorough understanding of the present disclosure and to fully convey the scope of the present disclosure to those skilled in the art.

[0071] The present application embodiment provides a method for calculating satellite orbit correction numbers based on Hill difference. Figure 1 As shown, the following steps are included:

[0072] In step S101, a monitoring station that has observed a target satellite is determined. In some examples, determining the monitoring station that has observed a target satellite includes:

[0073] Among the monitoring stations that observe GNSS satellite j, the monitoring stations with elevation angles greater than the altitude cutoff angle are selected, and the total number of monitoring stations is M.

[0074] In step S102, the pseudorange observations are smoothed using the carrier phase observations of the GNSS satellite j observed by each monitoring station (i=1, 2, ..., M).

[0075] In some specific examples, smoothing the pseudorange observations using the carrier phase observations of the GNSS satellite j observed by each monitoring station includes:

[0076] The carrier observation is converted to satisfy:

[0077]

[0078] in, f1 is the frequency of carrier L1, f5 is the frequency of carrier L5; and are the carrier phase observations at the L1 and L5 frequency points at time t respectively;

[0079] because The integer ambiguity of the two moments before and after is basically the same. The embodiment of the present application uses To smooth the noise in the pseudorange observations:

[0080]

[0081] Among them, L k Indicates L1 or L5 frequency point, is the pseudo-range observation value of the corresponding frequency point, is the pseudorange observation after smoothing of the corresponding frequency point, and τ is the smoothing time.

[0082] In step S103, the ionospheric delay is eliminated based on the smoothed pseudorange observations.

[0083] In some specific examples, eliminating ionospheric delay based on the smoothed pseudorange observations includes:

[0084] The pseudo-range observations after smoothing at L1 and L5 frequencies are used to eliminate the ionospheric delay. for:

[0085]

[0086] In step S104, the orbit parameters and clock parameters broadcast by the navigation message are used to calculate the ephemeris position and satellite clock deviation of the GNSS satellite j, and the inter-station distance is calculated in combination with the position of the monitoring station i.

[0087] In some specific examples, the ephemeris position of GNSS satellite j is calculated using the orbit parameters and clock parameters broadcast by the navigation message. and satellite clock bias Combined with the location of monitoring station i Calculate the distance between stations and satellites

[0088]

[0089] in, is the ephemeris position of GNSS satellite j, is the satellite clock bias, is the location of monitoring station i;

[0090] The distance between the stations and satellite clock bias from Eliminate it to obtain the pseudorange residual

[0091]

[0092] Among them, ΔR j (t) = [Δx j (t)Δy j (t)Δz j (t)] T is the orbit error of GNSS satellite j in the X, Y, and Z directions in the Earth-centered Earth-fixed coordinate system (the error between the satellite ephemeris position and the true position); is the unit direction vector from monitoring station i to GNSS satellite j; ΔB j (t) is the clock error of GNSS satellite j (the residual error after correction using the satellite clock bias in the navigation message); is the residual error, and the variance is

[0093] In step S105, based on the calculated station-satellite distance, the influence of the satellite clock error on the satellite orbit error is eliminated, and a state model of the satellite orbit error is established based on the Hill difference equation.

[0094] In some specific examples, based on the calculated station-satellite distance, eliminating the influence of satellite clock error on satellite orbit error includes:

[0095] The single difference method is used to eliminate the satellite orbit error caused by the satellite clock error. Each value of Subtract:

[0096]

[0097] in, is the residual error, and the variance is

[0098] Convert equation (7) into the corresponding matrix form:

[0099]

[0100] in, V o j The covariance matrix of

[0101] The state model of satellite orbit error based on the Hill difference equation includes establishing the Hill difference equation describing the satellite orbit error:

[0102] x″(t)-2ωy ′ (t)-3ω 2 x(t)=0 (9)

[0103] y″(t)+2ωx ′ (t)=0 (10)

[0104] z″(t)+ω 2 z(t)=0 (11)

[0105] Among them, x(t), y(t) and z(t) are the satellite orbit errors in the three directions of XYZ in the satellite inertial coordinate system at time t; ′ x(t) and x″(t) are the first and second derivatives of x(t), respectively; y ′ y(t) and y″(t) are the first and second derivatives of y(t), respectively; z ′ z(t) and z″(t) are the first and second derivatives of z(t), respectively; ω is the orbital angular velocity.

[0106] The matrix form of the solution to the Hill difference equation is:

[0107] X(t)=F(t)X0 (12)

[0108] Among them, X(t)=[x(t)y(t)z(t)x ′ (t)y ′ (t)z ′ (t)] T ; X0 = X(0), is the initial value of satellite orbit error;

[0109]

[0110] Then the state equation of the satellite orbit error is:

[0111] X(t n )=F(t n -t n-1 )X(t n-1 ) (13)

[0112] Among them, t n For time series.

[0113] In step S106, the state model is solved to obtain satellite orbit correction numbers.

[0114] In some specific examples, solving the state model to obtain satellite orbit corrections includes:

[0115] The observation equation and state equation used to solve the satellite orbit correction number are expressed as:

[0116]

[0117] X r (t n )=F(t n -t n-1 )X r (t n-1 )+W r (t n-1 ) (15)

[0118] in, t n The satellite orbit error at time t n The satellite orbit error change rate at time λ and are the latitude and longitude of the satellite sub-satellite trajectory, β = arccos[cos(i) / cos(λ)], i is the satellite orbit inclination; W r (t n) is noise, and its covariance matrix is:

[0119] Q o (t n )=diag([0.001 0.001 0.001 0.001 0.001 0.001])

[0120] Use Kalman filtering to calculate the satellite orbit correction number in the satellite inertial coordinate system

[0121]

[0122] P o (t n |t n-1 )=F(t n -t n-1 ) o (t n-1 )F T (t n -t n-1 )+Q o (t n-1 ) (17)

[0123]

[0124] P o (t n )=[IK o (t n )H o (t n )]P o (t n |t n-1 ) (20)

[0125] in, I is the identity matrix;

[0126] The satellite orbit correction number in the Earth Centered Earth Fixed (ECEF) coordinate system can be obtained through the matrix C

[0127]

[0128] The embodiment of the present application uses the Hill difference equation to establish a dynamic change model of satellite orbit error. The calculated satellite orbit correction number can more realistically reflect the satellite orbit error and improve the user positioning accuracy.

[0129] An embodiment of the present application also proposes a satellite-based augmentation satellite orbit correction number solution system based on Hill differential, including a processor and a memory, wherein a computer program is stored in the memory, and when the computer program is executed by the processor, the steps of the satellite-based augmentation satellite orbit correction number solution method based on Hill differential as described above are implemented.

[0130] It should be noted that in the various embodiments of the present application, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also includes other elements not explicitly listed, or also includes elements inherent to such process, method, article or device. In the absence of further restrictions, an element defined by the sentence "includes a ..." does not exclude the presence of other identical elements in the process, method, article or device including the element.

[0131] The serial numbers of the above-mentioned embodiments of the present application are for description only and do not represent the advantages or disadvantages of the embodiments.

[0132] Through the description of the above implementation methods, those skilled in the art can clearly understand that the above-mentioned embodiment methods can be implemented by means of software plus a necessary general hardware platform, and of course by hardware, but in many cases the former is a better implementation method. Based on such an understanding, the technical solution of the present application, or the part that contributes to the prior art, can be embodied in the form of a software product, which is stored in a storage medium (such as ROM / RAM, a magnetic disk, or an optical disk), and includes a number of instructions for a terminal (which can be a mobile phone, a computer, a server, an air conditioner, or a network device, etc.) to execute the methods described in each embodiment of the present application.

[0133] The embodiments of the present application are described above in conjunction with the accompanying drawings, but the present application is not limited to the above-mentioned specific implementation methods. The above-mentioned specific implementation methods are merely illustrative and not restrictive. Under the guidance of the present application, ordinary technicians in this field can also make many forms without departing from the purpose of the present application and the scope of protection of the claims, all of which are within the protection of the present application.

Claims

1. A method for calculating satellite orbit correction numbers for satellite-based augmentation based on Hill difference, characterized in that: include: Identify the monitoring station that observed the target satellite; The pseudo-range observations are smoothed using the carrier phase observations of GNSS satellite j observed by each monitoring station; Eliminate ionospheric delay based on smoothed pseudorange observations; The orbit parameters and clock parameters broadcast by the navigation message are used to calculate the ephemeris position and satellite clock bias of GNSS satellite j, and the inter-station distance is calculated in combination with the position of monitoring station i; Based on the calculated station-satellite distance, the influence of satellite clock error on satellite orbit error is eliminated, and the state model of satellite orbit error is established based on the Hill difference equation; The state model is solved to obtain satellite orbit correction numbers.

2. The method for calculating satellite orbit correction numbers based on Hill difference for satellite-based augmentation according to claim 1, characterized in that: The monitoring stations that have observed the target satellite include: Among the monitoring stations that observe GNSS satellite j, the monitoring stations with elevation angles greater than the altitude cutoff angle are selected, and the total number of monitoring stations is M.

3. The method for calculating satellite orbit correction numbers based on Hill difference for satellite-based augmentation according to claim 1, characterized in that: The pseudo-range observations are smoothed using the carrier phase observations of GNSS satellite j observed by each monitoring station, including: The carrier observation is converted to satisfy: in, f1 is the frequency of carrier L1, f5 is the frequency of carrier L5; and are the carrier phase observations at the L1 and L5 frequency points at time t respectively; use To smooth the noise in the pseudorange observations: Among them, L k Indicates L1 or L5 frequency point, is the pseudo-range observation value of the corresponding frequency point, is the pseudorange observation after smoothing of the corresponding frequency point, and τ is the smoothing time.

4. The method for calculating satellite orbit correction numbers based on Hill difference for satellite-based augmentation according to claim 3, characterized in that: Based on the smoothed pseudorange observations, eliminating ionospheric delay includes: The pseudo-range observations after smoothing at L1 and L5 frequencies are used to eliminate the ionospheric delay. for:

5. The method for calculating satellite orbit correction numbers based on Hill difference for satellite-based augmentation according to claim 4, characterized in that: The orbital parameters and clock parameters broadcast by the navigation message are used to calculate the ephemeris position and satellite clock bias of GNSS satellite j, and the inter-station distance is calculated in combination with the position of monitoring station i to satisfy: in, is the ephemeris position of GNSS satellite j, is the satellite clock bias, is the location of monitoring station i; The distance between the stations and satellite clock bias from Eliminate it to obtain the pseudorange residual Among them, ΔR j (t) = [Δx j (t)Δy j (t)Δz j (t)] T is the orbit error of GNSS satellite j in the X, Y, and Z directions in the Earth-centered Earth-fixed coordinate system; is the unit direction vector from monitoring station i to GNSS satellite j; ΔB j (t) is the clock error of GNSS satellite j; is the residual error, is the variance.

6. The method for calculating satellite orbit correction numbers based on Hill difference for satellite-based augmentation according to claim 5, characterized in that: Based on the calculated station-satellite distance, the effects of satellite clock error on satellite orbit error are eliminated by: The satellite orbit error is eliminated by the inter-station single difference method, and the pseudorange residual is Each value of Subtract: in, is the residual error, and the variance is Convert equation (7) into the corresponding matrix form: in, The covariance matrix of The state model of satellite orbit error based on the Hill difference equation includes establishing the Hill difference equation describing the satellite orbit error: x″(t)-2ωy ′ (t)-3ω 2 x(t)=0 (9) y″(t)+2ωx ′ (t)=0 (10) z″(t)+ω 2 z(t)=0 (11) Among them, x(t), y(t) and z(t) are the satellite orbit errors in the three directions of XYZ in the satellite inertial coordinate system at time t; ′ x(t) and x″(t) are the first and second derivatives of x(t), respectively; y ′ y(t) and y″(t) are the first and second derivatives of y(t), respectively; z ′ z(t) and z″(t) are the first and second derivatives of z(t), respectively; ω is the orbital angular velocity; The matrix form of the solution to the Hill difference equation is: X(t)=F(t)X0 (12) Among them, X(t)=[x(t)y(t)z(t)x ′ (t)y ′ (t)z ′ (t)] T ; X0 = X(0), is the initial value of satellite orbit error; Then the state equation of the satellite orbit error is: X(t n )=F(t n -t n-1 )X(t n-1 ) (13) Among them, t n For time series.

7. The method for calculating satellite orbit correction numbers based on Hill difference for satellite-based augmentation according to claim 6, characterized in that: Solving the state model to obtain satellite orbit corrections includes: The observation equation and state equation used to solve the satellite orbit correction number are expressed as: X r (t n )=F(t n -t n-1 )X r (t n-1 )+W r (t n-1 ) (15) in, t n The satellite orbit error at time, t n The satellite orbit error change rate at time λ and are the latitude and longitude of the satellite sub-satellite trajectory, β = arccos[cos(i) / cos(λ)], i is the satellite orbit inclination; W r (t n ) is noise, and its covariance matrix is: Q o (t n )=diag([0.001 0.001 0.001 0.001 0.001 0.001]) Use Kalman filtering to calculate the satellite orbit correction number in the satellite inertial coordinate system P o (t n |t n-1 )=F(t n -t n-1 )P o (t n-1 )F T (t n -t n-1 )+Q o (t n-1 ) (17) P o (t n )=[I-K o (t n )H o (t n )]P o (t n |t n-1 ) (20) in, I is the identity matrix; The satellite orbit correction number in the Earth Centered Earth Fixed (ECEF) coordinate system can be obtained through the matrix C 8. A satellite-based augmented satellite orbit correction number calculation system based on Hill difference, characterized in that: The method comprises a processor and a memory, wherein a computer program is stored in the memory, and when the computer program is executed by the processor, the steps of the method for calculating the satellite-based augmented satellite orbit correction number based on Hill difference as described in any one of claims 1 to 7 are implemented.

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