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.

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'.

[Sam Chak]

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]

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.

Answer 1

[Aiswarya]

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

[Toshi]

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);