Remote sensing load geographical positioning line of sight starting point determination method and system

Abstract

The application provides a remote sensing load geographical positioning line-of-sight starting point determination method and system, comprising the following steps: step 1, determining the position of the load line-of-sight starting point in the satellite body system; step 2, determining the accurate time of the remote sensing load detection moment; step 3, obtaining the three-axis attitude of the satellite orbit coordinate system near the load detection moment; step 4, interpolating to obtain the satellite attitude at the load detection moment; step 5, obtaining the satellite position determined by the satellite GPS system near the load detection moment; step 6, calculating the satellite position in the ephemeris celestial coordinate system; step 7, calculating the accurate position and speed of the satellite center of mass in the ephemeris celestial coordinate system at the load detection moment; step 8, calculating the accurate position of the satellite center of mass in the earth-fixed coordinate system at the load detection moment; and step 9, calculating the accurate position of the line-of-sight starting point in the earth-fixed coordinate system at the load detection moment. The method has strong robustness and can be universally applied to the accurate determination of the line-of-sight starting point of the remote sensing load geographical positioning.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of geographic positioning, in particular to a remote sensing load geographic positioning line-of-sight starting point determination method and system. BACKGROUND

[0002] Regardless of the type of remote sensing satellite, the positioning of its load data is very important. When the remote sensing load is geographically positioned, the detection or observation line-of-sight needs to be determined, including the line-of-sight direction and the line-of-sight starting point.

[0003] High-precision line-of-sight determination is an important guarantee for the geographic positioning of remote sensing loads. As the application of remote sensing loads to geographic positioning accuracy requirements improve, the accuracy requirements for determining the line-of-sight starting point also improve. The existing line-of-sight starting point generally uses the satellite centroid obtained by interpolation as the line-of-sight starting point, which may contain errors of meters to hundreds of meters.

[0004] Document 1 (Mash Nishihama et.al. MODIS Level 1A Earth Location: Algorithm Theoretical Basis Document Version 3.0, 1997) gives the line-of-sight starting point by interpolating the satellite centroid position. Documents 2 (Ye Jundu, Research on Load Line-of-Sight Determination Method Based on Star Measurement, Master's Thesis of University of Defense Science and Technology, 2014) and 3 (Fan Bin, Research on Load Line-of-Sight Determination Technology of Space Target Monitoring System, Master's Thesis of University of Defense Science and Technology, 2015) both focus on the correction of the line-of-sight direction and do not consider the error of the line-of-sight starting point obtained by interpolating the satellite centroid position.

[0005] Patent document CN112925708A (application number: CN202110217010.0) discloses a stationary orbit microwave satellite load and platform cooperative scanning imaging simulation method and system, which includes: step S1: simulating the imaging mode and optical axis pointing characteristics of the microwave load through a two-dimensional scanning motion mechanism; step S2: based on the two-dimensional scanning motion mechanism, a microwave load and satellite platform cooperative scanning imaging simulation method is proposed to simulate the influence of the satellite platform and the microwave load image positioning and registration on the ground; step S3: based on the influence of the microwave load and satellite platform cooperative scanning imaging on the microwave load image positioning and registration, the positioning errors before and after the microwave load line-of-sight navigation function is turned on are compared to verify the microwave load image positioning and registration technology.

[0006] In order to accurately calculate the line-of-sight starting point of the remote sensing load geographic positioning, the present application first determines the satellite orbit according to the GPS measured data, then calculates the accurate position of the satellite centroid in the Earth-fixed coordinate system at the load detection time, and determines the line-of-sight starting point in the Earth-fixed coordinate system in combination with the relative position of the load line-of-sight starting point in the satellite body system. SUMMARY

[0007] In view of the defects in the prior art, the purpose of the present application is to provide a remote sensing load geographical positioning line-of-sight starting point determination method and system.

[0008] The remote sensing load geographical positioning line-of-sight starting point determination method provided by the present application comprises:

[0009] Step 1: determining the position of the line-of-sight starting point of the load in the satellite body system according to the installation position of the load in the satellite body system;

[0010] Step 2: determining the accurate time of the remote sensing load detection moment according to the information of the satellite on-board internal timing system;

[0011] Step 3: obtaining the three-axis attitude Euler angle of the satellite in the satellite orbit coordinate system determined by the satellite attitude and orbit control system near the accurate time of the load detection moment;

[0012] Step 4: interpolating the three-axis attitude Euler angle of the satellite near the accurate time of the load detection moment to obtain the three-axis attitude Euler angle of the satellite at the load detection moment;

[0013] Step 5: obtaining the satellite position determined by the satellite GPS system near the accurate time of the load detection moment;

[0014] Step 6: calculating the satellite position in the ephemeris celestial coordinate system near the accurate time of the load detection moment;

[0015] Step 7: calculating the accurate position and speed of the satellite center of mass in the ephemeris celestial coordinate system at the load detection moment;

[0016] Step 8: calculating the accurate position of the satellite center of mass in the earth-fixed coordinate system at the load detection moment;

[0017] Step 9: calculating the accurate position of the geographical positioning line-of-sight starting point in the earth-fixed coordinate system at the load detection moment.

[0018] Preferably, the satellite positions determined by the satellite GPS system in step 5 are wherein t1, t2, …, t k is the time corresponding to the position; k is the total number of position samples;

[0019] The accurate time t of the remote sensing load detection moment m satisfies t1-t y ≤ t m ≤ t k ≤ t y ≤ t y is an orbit extrapolation time threshold;

[0020] The satellite position in the ephemeris celestial coordinate system in step 6 The calculation formula of the precise position of the satellite center of mass in the Earth-fixed coordinate system at time t

[0021]

[0022] Wherein, i = 1, 2, …, k, T ecr2eci (t i ) is the conversion matrix of the Earth-fixed coordinate system to the ephemeris celestial coordinate system at time t i ; T ecr2eci (t i ) is a matrix determined by the precession, nutation, earth rotation and polar motion at time t i .

[0023] The position of the satellite in the Earth-fixed coordinate system is converted to the ephemeris celestial coordinate system respectively to obtain the satellite position in the ephemeris celestial coordinate system near the precise time of the load detection time

[0024] Preferably, the step 7 comprises:

[0025] Step 7.1: According to the satellite position in the ephemeris celestial coordinate system The least square method is used to determine the precise orbit of the satellite in the ephemeris celestial coordinate system combined with the orbit dynamics recursive model.

[0026] Step 7.2: According to the precise orbit, the position m and the velocity of the satellite center of mass at time t

[0027] Preferably, the calculation formula of the precise position of the satellite center of mass in the Earth-fixed coordinate system at time t

[0028]

[0029] Wherein, T eci2ecr (t m ) is the conversion matrix of the ephemeris celestial coordinate system to the Earth-fixed coordinate system at time t m , and T eci2ecr (t m ) is equal to the transpose of the conversion matrix T m (t ecr2eci ) of the Earth-fixed coordinate system to the ephemeris celestial coordinate system at time t m .

[0030] Preferably, the position of the load line-of-sight starting point in the satellite body system in step 1 is The satellite three-axis attitude Euler angles at time t m obtained in step 4 are respectively the yaw ψ, the roll angle φ, and the pitch angle θ. ​​

[0031] Then the geolocation of the line-of-sight starting point at the load detection moment in step 9 is accurate in the earth-fixed coordinate system The following method is used for calculation:

[0032]

[0033] Wherein, T sc2orb (t m ) is the conversion matrix from the satellite body system to the satellite orbit system, T sc2orb (t m ) satisfies:

[0034]

[0035] T orb2eci (t m ) is the conversion matrix from the satellite orbit system to the ephemeris celestial coordinate system, T orb2eci (t m ) satisfies:

[0036]

[0037] Wherein, is the direction of the substar point, is the direction of the negative angular momentum vector, is the direction of the circumferential velocity vector, T eci2ecr (t m ) is the conversion matrix from the ephemeris celestial coordinate system to the earth-fixed coordinate system at the moment t m .

[0038] According to the remote sensing load geolocation line-of-sight starting point determination system provided by the application, comprising:

[0039] Module M1: according to the installation position of the load in the satellite body system, the position of the load line-of-sight starting point in the satellite body system is determined;

[0040] Module M2: according to the load on-board internal timing system information, the accurate time of the remote sensing load detection moment is determined;

[0041] Module M3: the three-axis attitude Euler angle of the satellite orbit coordinate system determined by the satellite attitude and orbit control system near the accurate time of the load detection moment is obtained;

[0042] Module M4: according to the three-axis attitude Euler angle near the accurate time of the load detection moment, the satellite three-axis attitude Euler angle at the load detection moment is obtained by interpolation;

[0043] Module M5: the satellite position determined by the satellite GPS system near the accurate time of the load detection moment is obtained;

[0044] Module M6: calculating satellite position in the ephemeris celestial coordinate system near the precise time of the payload detection time;

[0045] Module M7: calculating the precise position and velocity of the satellite center of mass in the ephemeris celestial coordinate system at the payload detection time;

[0046] Module M8: calculating the precise position of the satellite center of mass in the earth-fixed coordinate system at the payload detection time;

[0047] Module M9: calculating the precise position of the geolocation line-of-sight starting point in the earth-fixed coordinate system at the payload detection time.

[0048] Preferably, the satellite positions determined by the satellite GPS system are in turn wherein t1, t2, …, tk are the corresponding times of the positions; k is the total number of position samples; k wherein t1, t2, …, tk are the corresponding times of the positions; k is the total number of position samples;

[0049] The precise time t of the remote sensing payload detection time m satisfies t1-t y ≤ t m ≤ t k + t y , t y is an orbit extrapolation time threshold;

[0050] The satellite position in the ephemeris celestial coordinate system is calculated by the following formula:

[0051]

[0052] wherein i = 1, 2, …, k, T ecr2eci (t i ) is the conversion matrix from the earth-fixed coordinate system to the ephemeris celestial coordinate system at t i ; T ecr2eci (t i ) is a matrix determined by the precession, nutation, earth rotation, and polar motion at t i .

[0053] The positions in the earth-fixed coordinate system are respectively converted to the ephemeris celestial coordinate system to obtain the satellite positions in the ephemeris celestial coordinate system near the precise time of the payload detection time .

[0054] Preferably, the module M7 includes:

[0055] Module M7.1: calculating the satellite position in the ephemeris celestial coordinate system according to the satellite position in the ephemeris celestial coordinate system

[0056] Module M7.2: Calculate t according to precise orbit calculation m Position of satellite center of mass at time t and velocity

[0057] Preferred, the precise position of the satellite center of mass in the Earth-fixed coordinate system at time t is calculated by the formula:

[0058]

[0059] where T eci2ecr (t m ) is the transformation matrix from the ephemeris celestial coordinate system to the Earth-fixed coordinate system at time t m , and T eci2ecr (t m ) is the transpose of the transformation matrix T m (t ecr2eci ) from the Earth-fixed coordinate system to the ephemeris celestial coordinate system at time t m .

[0060] Preferred, the position of the line-of-sight starting point of the load in the satellite body coordinate system at time t is calculated by the formula:

[0061] The Euler angles of the satellite three-axis attitude at time t m are yaw ψ, roll angle φ, and pitch angle θ, respectively.

[0062]

[0063] where T sc2orb (t m ) is the transformation matrix from the satellite body coordinate system to the satellite orbit coordinate system, and T sc2orb (t m ) satisfies:

[0064]

[0065] T orb2eci (t m ) is the transformation matrix from the satellite orbit coordinate system to the ephemeris celestial coordinate system, and T orb2eci (t m ) satisfies:

[0066]

[0067] where is the direction of the substar point, is the direction of the negative angular momentum vector, is the direction of the circumferential velocity vector, and T​eci2ecr (t m ) for t m The transformation matrix from the celestial coordinate system to the Earth-fixed coordinate system at a given time epoch.

[0068] Compared with the prior art, the present invention has the following beneficial effects:

[0069] This invention provides a method for determining the starting point of the line of sight for remote sensing payload geolocation. First, the satellite orbit is determined based on GPS measured data. Then, the precise position of the satellite's center of mass in the Earth-fixed coordinate system at the time of payload detection is calculated. Combined with the relative position of the payload's line of sight starting point in the satellite's own coordinate system, the starting point of the line of sight in the Earth-fixed coordinate system is determined. This method is reasonable, computationally simple, and easy to implement. It can accurately obtain the starting point of the line of sight for remote sensing payload geolocation, and can still stably output an accurate starting point even when there are a few outliers (exceeding the measurement error) in the GPS measured data. Attached Figure Description

[0070] Other features, objects, and advantages of the present invention will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0071] Figure 1 This is a flowchart of the present invention;

[0072] Figure 2 The satellite positions output by the GPS system;

[0073] Figure 3 This involves comparing the distance between the satellite's center of mass in the Earth-fixed coordinate system, obtained through precise orbit calculations, and the corresponding geocentric distance between the satellite's position output by the GPS system. Detailed Implementation

[0074] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0075] Example:

[0076] Geographic positioning of remote sensing payloads requires determining the direction and starting point of the detection or observation line of sight. Errors in the starting point of the line of sight directly affect positioning accuracy. The main source of errors in the starting point of the line of sight is the satellite position error at the moment of detection (or exposure) by the remote sensing payload, while the secondary influencing factor is the installation position of the payload relative to the satellite's center of mass.

[0077] like Figure 1, in order to improve the accuracy of the remote sensing load geographic positioning line of sight starting point, first according to the load in the satellite body under the installation position, determine the load line of sight starting point in the satellite body under the position.

[0078] According to the load on the satellite internal timing system information, determine the remote sensing load detection (or exposure) time accurate time.

[0079] Obtain the satellite orbit coordinate system under the satellite attitude and orbit control system determined by the satellite orbit coordinate system near the load detection time accurate time three-axis attitude Euler angle (yaw angle, pitch angle, roll angle). And according to the three-axis attitude Euler angle near the load detection time accurate time, interpolation to get the satellite three-axis attitude Euler angle of load detection time.

[0080] Satellite positioning through satellite GPS system is determined, but the satellite position of GPS system output is generally equal time interval data (such as 1 second interval), and there is no synchronization relationship between the remote sensing load detection (or exposure) time accurate time.

[0081] The satellite position determined by satellite GPS system is Where t1, t2, …, t k is the corresponding time of position, k is the total number of position samples, and the remote sensing load detection (or exposure) time accurate time t m y ≤t m ≤t k +t y , t y is the orbit extrapolation time threshold (generally can be set to 20 min, that is, the data within 20 min between the load detection time and the available GPS measured data can be effectively derived to the satellite accurate position).

[0082] In order to calculate the satellite center position, first convert the position (i = 1, 2, …, k) to the satellite position

[0083]

[0084] Where, T ecr2eci (t i ) is the conversion matrix of the earth fixed coordinate system to the ephemeris celestial coordinate system at t i , T ecr2eci (t i ) is the matrix determined by the precession, nutation, earth rotation and polar motion at t i . Convert the position in the earth fixed coordinate system to the ephemeris celestial coordinate system respectively, get the satellite position in the ephemeris celestial coordinate system near the load detection time accurate time​

[0085] Based on the measured satellite positions in the epochal celestial coordinate system The precise orbit of a satellite in the epochal celestial coordinate system can be determined using the least squares method combined with a recursive orbital dynamics model. Theoretically, the orbit can be determined from just two positions; more samples can improve the accuracy of orbit determination. Even if a small number of outliers exceeding the measurement error exist in the measured values, the t value can still be accurately obtained. m Position of the satellite's center of mass at any given time speed

[0086] Then, the precise position of the satellite's center of mass in the Earth-fixed coordinate system at the moment of payload detection is calculated.

[0087]

[0088] Among them, T eci2ecr (t m ) for t m The transformation matrix from the celestial coordinate system to the Earth-fixed coordinate system at time epoch, T eci2ecr (t m ) equals t m The transformation matrix T from the Earth-fixed coordinate system to the epochal celestial coordinate system. ecr2eci (t m The transpose of ).

[0089] Finally, the exact location of the starting point of the geographic positioning line of sight at the moment of payload detection is calculated in the Earth's fixed coordinate system.

[0090] If the position of the load line-of-sight starting point within the satellite's own system is t m The satellite's three-axis attitude Euler angles at a given time are yaw ψ, roll φ, and pitch θ. Therefore, the precise location of the payload's geographic positioning line of sight at that time within the Earth-fixed coordinate system is... Calculated using the following method:

[0091]

[0092] Among them, T sc2orb (t m () represents the transformation matrix from the satellite's intrinsic system to its orbital system. Taking the 3-1-2 transformation sequence as an example, T sc2orb (t m )satisfy

[0093]

[0094] T orb2eci (t m T is the transformation matrix from the satellite orbital system to the epochal celestial coordinate system.orb2eci (t m )satisfy:

[0095]

[0096] in, The direction of the sub-star point. The direction of the negative angular momentum vector. The direction of the circumferential velocity vector.

[0097] T eci2ecr (t m ) for t m The transformation matrix from the celestial coordinate system to the Earth-fixed coordinate system at a given time epoch.

[0098] The effectiveness of the method of this invention is verified by the following on-orbit remote sensing data from a certain satellite. The satellite position output by the GPS system around 3:30 UTC on July 10, 2021 is shown below. Figure 2 As shown. The position sequence is transformed to the epochal celestial coordinate system. The position of the satellite's barycenter in the epochal celestial coordinate system, obtained through precise orbit calculation, is then transformed to the Earth-fixed coordinate system. The geocentric data corresponding to the calculated precise satellite position and the satellite position output by the GPS system is compared as follows: Figure 3 As shown. By Figure 3 It can be seen that there are a few outliers in the satellite position output by the GPS system around 3:40, but the calculated accurate satellite position is not affected by these outliers and stably outputs the high-precision satellite centroid position. Therefore, the accurate line-of-sight starting point can be calculated according to the procedure of this invention.

[0099] Those skilled in the art will understand that, in addition to implementing the system, apparatus, and their modules provided by this invention in purely computer-readable program code, the same program can be implemented in the form of logic gates, switches, application-specific integrated circuits, programmable logic controllers, and embedded microcontrollers by logically programming the method steps. Therefore, the system, apparatus, and their modules provided by this invention can be considered a hardware component, and the modules included therein for implementing various programs can also be considered structures within the hardware component; alternatively, modules for implementing various functions can be considered both software programs implementing the method and structures within the hardware component.

[0100] Specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the specific embodiments described above, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Unless otherwise specified, the embodiments and features described in this application can be arbitrarily combined with each other.

Claims

1. A method for determining the origin of a line of sight for geolocating a remote sensing payload, characterized in that, The method comprises the following steps: Step 1: determining the position of the line-of-sight starting point of the load in the satellite body system according to the installation position of the load in the satellite body system; Step 2: determining the accurate time of the load detection moment according to the information of the internal timing system on the satellite; Step 3: obtaining the three-axis attitude Euler angle of the satellite in the satellite orbit coordinate system determined by the satellite attitude and orbit control system near the accurate time of the load detection moment; Step 4: interpolating the three-axis attitude Euler angle of the satellite at the load detection moment according to the three-axis attitude Euler angle near the accurate time of the load detection moment; Step 5: obtaining the satellite position determined by the satellite GPS system near the accurate time of the load detection moment; Step 6: calculating the satellite position in the ephemeris celestial coordinate system near the accurate time of the load detection moment; Step 7: calculating the accurate position and speed of the satellite center of mass in the ephemeris celestial coordinate system at the load detection moment; Step 8: calculating the accurate position of the satellite center of mass in the earth-fixed coordinate system at the load detection moment; Step 9: calculating the accurate position of the geographical positioning line-of-sight starting point in the earth-fixed coordinate system at the load detection moment; The position of the load line-of-sight origin in the satellite body system in step 1 is , the satellite three-axis attitude Euler angles at the time obtained in step 4 are yaw , roll angle , and pitch angle ; The precise position of the line-of-sight starting point in the earth-fixed coordinate system at the time of the load detection in step 9 is determined This is calculated by the following method: wherein is the transformation matrix from the satellite body frame to the satellite orbital frame, satisfies: is the transformation matrix from the satellite orbit system to the epoch celestial coordinate system, satisfies: wherein is the direction of the sub-satellite point, ; is the direction of the negative angular momentum vector, ; is the direction of the circumferential velocity vector, ; is the direction of the conversion matrix of the ECI to the ECF at the instant of the epoch.

2. The method according to claim 1, wherein, The satellite positions determined by the satellite GPS system in step 5 are in turn wherein is the time corresponding to the position; is the total number of position samples; Remote sensing payload detection time accurate time Satisfy , Is the orbit extrapolation time threshold; Satellite position in the ECI at the epoch of step 6 The formula for this is: wherein , is the conversion matrix of the Earth-fixed coordinate system to the ephemeris celestial coordinate system at the time instant t; is the matrix determined by the precession, nutation, Earth rotation and polar motion at the time instant t; Converting the position in the Earth-fixed coordinate system into the ECI coordinate system respectively, to obtain the satellite position in the ECI coordinate system near the precise time of the payload detection moment .

3. The method of claim 1, wherein, The step 7 comprises: Step 7.1: Satellite position in ECI The precise orbit of the satellite in ECI is determined using least squares combined with an orbit dynamics propagator. Step 7.2: Precise orbit computation according to Position of the satellite's center of mass at the time instant and velocity .

4. The method of claim 1, wherein, The accurate position of the satellite center of mass in the earth-fixed coordinate system in step 8 The calculation formula is: wherein, is the conversion matrix from the ECI coordinate system to the ECF coordinate system at the time instant, is equal to the conversion matrix from the ECF coordinate system to the ECI coordinate system at the time instant is the transpose of.

5. A remote sensing payload geolocating line-of-sight origin determination system, characterized by, The method comprises the following steps: Module M1: determining the position of the line-of-sight starting point of the load in the satellite body system according to the installation position of the load in the satellite body system; Module M2: determining the accurate time of the load detection moment according to the information of the internal timing system on the satellite; Module M3: obtaining the three-axis attitude Euler angle of the satellite in the satellite orbit coordinate system determined by the satellite attitude and orbit control system near the accurate time of the load detection moment; Module M4: interpolating the three-axis attitude Euler angle of the satellite at the load detection moment according to the three-axis attitude Euler angle near the accurate time of the load detection moment; Module M5: obtaining the satellite position determined by the satellite GPS system near the accurate time of the load detection moment; Module M6: calculating the satellite position in the ephemeris celestial coordinate system near the accurate time of the load detection moment; Module M7: calculating the accurate position and speed of the satellite center of mass in the ephemeris celestial coordinate system at the load detection moment; Module M8: calculating the accurate position of the satellite center of mass in the earth-fixed coordinate system at the load detection moment; Module M9: calculating the accurate position of the geographical positioning line-of-sight starting point in the earth-fixed coordinate system at the load detection moment; The position of the line of sight of the load at the satellite body is , the satellite three-axis attitude Euler angles at the time of are respectively yaw , roll angle , and pitch angle ; The precise position of the line-of-sight starting point of the geographic positioning at the load detection moment in the earth-fixed coordinate system The calculation formula is: wherein is the transformation matrix from the satellite body frame to the satellite orbital frame, satisfies: is the transformation matrix from the satellite orbit system to the epoch celestial coordinate system, satisfies: wherein is the direction of the sub-satellite point, ; is the direction of the negative angular momentum vector, ; is the direction of the circumferential velocity vector, ; is the direction of the transformation matrix from the ECI to the ECF at the instant of the epoch.

6. The remote sensing payload geo-location line-of-sight origin determination system of claim 5, wherein, The satellite positions determined by the satellite GPS system are in turn wherein is the time corresponding to the position; is the total number of position samples; Remote sensing payload detection time accurate time Satisfy , Is the orbit extrapolation time threshold; Satellite position in the epochal celestial coordinate system The formula for calculating the satellite position in the epochal celestial coordinate system is: wherein , is the conversion matrix of the Earth-fixed coordinate system to the ephemeris celestial coordinate system at the time instant t; is the matrix determined by the precession, nutation, Earth rotation and polar motion at the time instant t; Converting the position in the Earth-fixed coordinate system into the ECI coordinate system respectively, to obtain the satellite position in the ECI coordinate system near the precise time of the payload detection moment .

7. The remote sensing payload geo-location line-of-sight origin determination system of claim 5, wherein, The module M7 comprises: Module M7.1 : satellite position in the ephemeris celestial reference frame determining the precise orbit of the satellite in the ephemeris celestial reference frame using a least squares method combined with an orbit dynamics recursive model; Module M7.2: Precise orbit computation according to Position of the satellite's center of mass at the time instant and velocity .

8. The remote sensing payload geo-location line-of-sight origin determination system of claim 5, wherein, Precise position of satellite center of mass in earth fixed coordinate system The formula for calculating the precise position of the satellite center of mass in the earth fixed coordinate system is: wherein, is the conversion matrix from the Earth-fixed coordinate system to the epoch celestial coordinate system at the time instant, is equal to the conversion matrix from the epoch celestial coordinate system to the Earth-fixed coordinate system at the time instant the transpose of.

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