Values changining after a satellite is initialized.

Question

Raghav Rathi 2024-8-5

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评论： Raghav Rathi 2024-8-15

Hi,

Below is the code which I am using in order to create a satellite scenario. I have defined all the required parameters manually that are required to initialize a satellte, but I am not able to understand why are they getting changed after the satellite is initialized.

startTime = datetime('27-Jun-2024 13:23:02'); 
stopTime = startTime + seconds(60*100);
sampleTime = 10;
sc = satelliteScenario(startTime,stopTime,sampleTime);
mu = 3.986004418*1e14;
n = 15.06398680175205;
semiMajorAxis = nthroot(mu,3)/(((2*n*pi)/86400)^(2/3)) ;
trueAnomaly = 341.2089;    
Eccentricity = 1.249e-04;
Inclination = 53.0534;
RightAscensionOfAscendingNode = 242.3386;
ArgumentOfPeriapsis = 94.5842;
sat = satellite(sc,semiMajorAxis,Eccentricity,Inclination,RightAscensionOfAscendingNode,ArgumentOfPeriapsis,trueAnomaly);
elements = orbitalElements(sat)
elements = struct with fields:
                       MeanMotion: 0.0628
                     Eccentricity: 1.2490e-04
                      Inclination: 53.1730
    RightAscensionOfAscendingNode: 242.7010
              ArgumentOfPeriapsis: 94.5031
                      MeanAnomaly: 341.2135
                           Period: 5.7355e+03
                            Epoch: 27-Jun-2024 13:23:02
                            BStar: 0

I would really appreciate if someone can help me figure out why are the values of RAAN and ArgumentOfPeriapsis getting changed after the satellite is initialized.

Thank you

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Umar 2024-8-6

编辑：Walter Roberson 2024-8-6

Hi @ Raghav Rathi,

Also, I have noticed carefully now that TLE information you provided is not directly being used to initialize the satellite, and instead, you are manually defining parameters such as semiMajorAxis, Eccentricity, Inclination, RightAscensionOfAscendingNode, ArgumentOfPeriapsis, and trueAnomaly. So, when you initialize a satellite using the Aerospace Toolbox satellite function with manual parameter inputs, it's possible that these values are being overridden by the TLE-derived elements. So, my approach at this point would be parsing the TLE data and extract the relevant parameters needed to initialize the satellite. This involves converting the TLE data into orbital elements such as semi-major axis, eccentricity, inclination, right ascension of ascending node, argument of periapsis, and mean anomaly. Once you have extracted these values from the TLE, you can then use them to initialize the satellite in your scenario. By doing so, you can ensure that the satellite's parameters are accurately defined based on the TLE information. By incorporating this approach into your code, you can achieve consistency between the manually defined parameters and those derived from the TLE data. This will help in resolving any discrepancies or changes observed after satellite initialization.

Secondly, be aware of potential issues related to data reliability. Data derived from TLEs older than 30 days can become unreliable, and it's recommended to calculate orbital positions using algorithms such as Simplified General Perturbations-4 (SGP4) and Simplified Deep-Space Perturbations-4 (SDP4). To resolve this problem, consider using TLE data directly as input for satellite initialization in MATLAB's Aerospace Toolbox. By obtaining reliable TLE files from sources such as the Space Track website, you can ensure that the orbital elements, including RAAN and Argument of Periapsis, are accurately defined based on current information.

I would also recommend reviewing materials by clicking the following links below.

https://www.mathworks.com/help/aerotbx/ug/matlabshared.satellitescenario.satellite.orbitalelements.html?searchHighlight=orbitalElements&s_tid=srchtitle_support_results_1_orbitalElements

https://www.mathworks.com/help/aerotbx/ug/satelliteScenario-key-concepts.html#mw_355a5ca3-9480-4991-8f53-fa1b11befe86

Hope this helps.

Raghav Rathi 2024-8-6

@Umar, Thank you for the explaination.

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Answer 1

Garmit Pant 2024-8-6

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在 MATLAB Online 中打开

Hello Raghav,

I ran your code snippet on MATLAB R2023b and was able to reproduce the same outputs. The discrepancies between the input angles and the output of the “orbitalElements” function are due to the way the different input methods to the “satellite” function are handled.

When the “satellite” object is initialized by passing Keplerian elements defined in the Geocentric Celestial Reference Frame (GCRF), the function internally computes the ‘position’ and ‘velocity’ of the satellite in the True Equator Mean Equinox (TEME) frame. These are then further used to compute the Keplerian elements’ values in the TEME frame. Thus, the values returned by the “orbitalElements” function are different from the inputs since they represent the transformed values of the input in the TEME frame.

When the “satellite” object is initialized using a file-based input argument, the values remain consistent. You can use the following code snippet to write your TLE information to a TXT file and then use it to initialize a “satellite” object:

% Define the TLE data
line1 = 'Satellite 1';
line2 = '1 47641C 21012X   24179.14770833  .00007445  00000+0  49891-3 0  1797';
line3 = '2 47641  53.0534 242.3386 0001249  94.5842 341.2135 15.06386781     5';
% Open a file for writing
filename = 'satellite.tle';
fileID = fopen(filename, 'w');
% Check if the file opened successfully
if fileID == -1 
    error('Failed to open file for writing.'); 
end
% Write the TLE lines to the file
fprintf(fileID, '%s\n', line1);
fprintf(fileID, '%s\n', line2);
fprintf(fileID, '%s\n', line3);
% Close the file
fclose(fileID);
% Display a message indicating success
disp(['TLE file written to ', filename]);
TLE file written to satellite.tle

% Set the scenario
startTime = datetime('27-Jun-2024 13:23:02'); 
stopTime = startTime + seconds(60*100);
sampleTime = 10;
sc = satelliteScenario(startTime,stopTime,sampleTime);
% Add the satellite
sat = satellite(sc,'satellite.tle');
elements = orbitalElements(sat)
elements = struct with fields:
                       MeanMotion: 0.0628
                     Eccentricity: 1.2490e-04
                      Inclination: 53.0534
    RightAscensionOfAscendingNode: 242.3386
              ArgumentOfPeriapsis: 94.5842
                      MeanAnomaly: 341.2135
                           Period: 5.7356e+03
                            Epoch: 27-Jun-2024 03:32:41
                            BStar: 4.9891e-04

For further understanding, kindly refer to the following MathWorks Documentation:

Refer to the ‘Input Arguments’ section to understand different methods of initialising a “satellite” object: https://www.mathworks.com/help/releases/R2023b/aerotbx/ug/satellitescenario.satellite.html#mw_0e009042-cc9a-46c0-9514-d5019692ca22
Refer to the ’Two Line Element (TLE) Files’ section to understand the TLE standard: https://www.mathworks.com/help/releases/R2023b/satcom/gs/satelliteScenario-key-concepts.html#mw_e08739b4-c5da-4983-898b-56e18cc71f87

I hope you find the above explanation and suggestions useful!

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Umar 2024-8-6

Hi @ Garmit Pant,

The point that you made was very important to notice, “ When the “satellite” object is initialized by passing Keplerian elements defined in the Geocentric Celestial Reference Frame (GCRF), the function internally computes the ‘position’ and ‘velocity’ of the satellite in the True Equator Mean Equinox (TEME) frame. These are then further used to compute the Keplerian elements’ values in the TEME frame. Thus, the values returned by the “orbitalElements” function are different from the inputs since they represent the transformed values of the input in the TEME frame.”

This is because GCRF serves as a reference for defining the initial orbital parameters, and these are then transformed into the TEME frame for further analysis and calculations. I would like to elaborate further that the process involves converting the Keplerian elements from the GCRF to the TEME frame, allowing for a more precise representation of the satellite's orbital characteristics. This transformation is crucial for accurately determining the satellite's position and velocity relative to the Earth's true equator and mean equinox. By utilizing this approach, it becomes possible to establish a comprehensive understanding of the satellite's orbital dynamics within a specific reference frame. Also, this process plays a pivotal role in satellite operations and contributes to the overall accuracy and efficiency of space missions. Very good analysis.