Errors using propagateOrbit after function update

I was using the propagateOrbit function to generate time series of state vectors from an initial state vector, start time, and stop time. I have a datetime array from start to stop then make call:

propagateOrbit(datetimes, startState(1:3), startState(4:6), PropModel='sgp4');

This was working up until a few weeks ago. I'm using R2024b and I know that some changes were made to the function in R2025a, but I'm not sure if that would break things. This is the error I get:

Array indices must be positive integers or logical values.

Error in Aero.internal.spaceweather.validateFluxData

Error in Aero.internal.spaceweather.validateFluxData

Error in Aero.spacecraft.NumericalPropagatorOptions/localSetSpaceWeatherDataFile

Error in Aero.spacecraft.NumericalPropagatorOptions

Error in propagateOrbit

Error in generateTruthData (line 64)
    propagateOrbit(datetimes, startState(1:3), startState(4:6), PropModel='sgp4');
    ^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^^

This tells me that it's trying to use the numerical propagator (which I don't specify options for), but I'm specifying SGP4 as the propagation model. Could it be an issue with the state vectors? Is anyone else running into a similar issue? Thanks!

[Torsten]

Did you check "datetimes" and "startState" before the call to "propagateOrbit" ? Can you supply data for the inputs to reproduce the error ?

Answer 1

[Ayush]

Hi Nicholas,

I understand that you are working with “Aerospace Toolbox” and facing error while using the “propagateOrbit” function to generate a time series of state vectors from an initial state vector and a datetime array, specifying the SGP4 propagation model.

From the below attached MATLAB documentation, the input syntax for “propagateOrbit” is as follows:

[positions,velocities] = propagateOrbit(time,rEpoch,vEpoch)

I have replicated your code in my system (MATLAB R2024b):

Since, the values of “datetimes”, “startState” is not mentioned in your query, I have used these values from MATLAB documentation example.

This is the code which is working fine for me and giving appropriate results for “r” (position) and “v” (velocity):

rEpoch = [5927630.386747557; ...
    3087663.891097251; ...
    1174446.969646237]; 

vEpoch = [-5190.330300215130; ...
    8212.486957313564; ...
    4605.538019512981]; 

[r, v] = propagateOrbit(datetime(2022, 1, 3, 12, 0, 0), rEpoch,vEpoch, 'PropModel', 'sgp4');
disp([r,v])

There are few other possible workarounds to avoid the error:

1. Check your “startState” and “datetimes”: Ensure that “startState(1:3)” and “startState(4:6)” in your code are not parallel as mentioned in the attached documentation. And, it is preferred to have “datetimes” in UTC.
2. Use Two Line element (TLE) structure:

To access one example on TLE, you can type the following in the command window:

    openExample('aero/ReadDataFromTLEFileAndCalculatePositionAndVelocityExample')

Below is the code to use “propagateOrbit” using TLE file in above example:

    tleStruct = tleread('leoSatelliteConstellation.tle')
    [r,v] = propagateOrbit(datetime(2022, 1, 3, 12, 0, 0),tleStruct,"PropModel","sgp4");

Following documents will be helpful:

1. Input syntax and state values: propagateOrbit
2. About Two-line element structure: tleread

Hope this helps clarify your options.