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I want to chcek the SGP4 orbit Propogator. But MY CODING IS NOT WORKING. Can someone check and help me with the coding. I want to use orbital elements not the tle file.

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Toshi
Toshi 2024-2-4
评论： Toshi 2024-3-4
Ran in:
Here is my code;
% Orbital elements of starlink
a = 6925000; % Semi-major axis in km
e = 0.0001406; % Eccentricity
i = 53.0538; % Inclination in degrees
RAAN = 66.9644; % Right Ascension of Ascending Node in degrees
arg_perigee = 87.41; % Argument of Perigee in degrees
M = 272.6995; % Mean Anomaly in degrees
epoch = 23285; % Epoch time in days
[positions,velocities]= orbitPropagate(datetime(2020,2,2,12,0,2),a,e,i,RAAN, ...
arg_perigee,M,epoch);
Unrecognized function or variable 'orbitPropagate'.
  2 个评论
Sam Chak
Sam Chak 2024-2-25
Hey @Toshi, I wanted to inquire about the coding issue we discussed earlier. It's a fascinating problem, especially for satellite enthusiasts like us. I'm genuinely curious to know if any recent developments or breakthroughs have occurred that might have led to its resolution.
Toshi
Toshi 2024-3-4
Hi @Sam Chak,
I am still learning. But I think there are many new features available in matlab for satcom.There is PseudoRange Function available, high precision orbit propogator available.

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回答（1 个）

Aiswarya
Aiswarya 2024-2-12
Hi Toshi,
The issue with your code as mentioned in the error message is because you are using an incorrect function name. The built-in function to do the same is 'propagateOrbit'. This gives you option to use orbital elements or tle file as well. You can modify your function call as follows:
[positions,velocities]= propagateOrbit(datetime(2020,2,2,12,0,2),a,e,i,RAAN,arg_perigee,M);
You may refer to the following documentation for more information on the 'propagateOrbit' function:
https://www.mathworks.com/help/aerotbx/ug/propagateorbit.html?searchHighlight=orbitPropagate&s_tid=srchtitle_support_results_6_orbitPropagate
  1 个评论
Toshi
Toshi 2024-3-4
Thank you for the help.
I have one more question if you can help : how can i get the orbital elements at stoptime
startTime = datetime(2024,3,1,22,07,37);
stopTime = startTime + days(3);
sampleTime = 60;
sc = satelliteScenario(startTime,stopTime,sampleTime);
tleFile = "o3b.tle" ;
sat1 = satellite(sc,tleFile);
sat1 = satellite(sc,tleFile, ...
"Name","sat1", ...
"OrbitPropagator","sgp4");
elements1 = orbitalElements(sat1);
[positionSGP4,velocitySGP4] = states(sat1);
display(elements1);
display(velocitySGP4);
display(positionSGP4);
satelliteScenarioViewer(sc);

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