Hello Orekit team,

I am new to Orekit and I am trying to write a code that would propagate backwards the orbit of a satellite which is supposed to be performing an EOR and then comparing the results to the ones obtained from a simple backwards propagated orbit using a keplerian propagator to validate the concept.

The code I wrote for the backwards propagation of the maneuver is the following:

```
public static DragForce dragForce(double crossSection, double dragCoef){
Frame frame = FramesFactory.getITRF(IERSConventions.IERS_2010,true);
IsotropicDrag drag = new IsotropicDrag(crossSection,dragCoef);
OneAxisEllipsoid earthShape = new OneAxisEllipsoid(Constants.WGS84_EARTH_EQUATORIAL_RADIUS,Constants.WGS84_EARTH_FLATTENING,frame);
SimpleExponentialAtmosphere earthAtmosphere = new SimpleExponentialAtmosphere(earthShape, 0.0949,20000.0,7500);
return new DragForce(earthAtmosphere,drag);
}
public static void main(String[] args) throws IOException {
File orekitData = new File("/home/razvan/Downloads/orekit-data");
DataProvidersManager manager = DataContext.getDefault().getDataProvidersManager();
manager.addProvider(new DirectoryCrawler(orekitData));
Frame inertialFrame = FramesFactory.getEME2000();
TimeScale utc = TimeScalesFactory.getUTC();
AbsoluteDate initialDate = new AbsoluteDate(2023, 01, 31, 17, 46, 49.924704, utc);
double mu = 3.986004415e+14;
double a = 504848; //semi-major axis in metres
double e = 0.00101990;
double i = FastMath.toRadians(98.6875);
double omega = FastMath.toRadians(131.2378);
double raan = FastMath.toRadians(39.9845);
double lM = FastMath.toRadians(337.9425);
double mass = 6000;
double crossSection = 10;
double dragCoef = 2.2;
Orbit initialOrbit = new KeplerianOrbit(a, e, i, omega, raan, lM, PositionAngle.MEAN, inertialFrame, initialDate, mu);
double dP = 1;
double minStep = 0.001;
double maxStep = 1000;
double initStep = 20;
double[][] tolerance = NumericalPropagator.tolerances(dP,initialOrbit, OrbitType.KEPLERIAN);
AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(minStep, maxStep, tolerance[0], tolerance[1]);
integrator.setInitialStepSize(initStep);
NumericalPropagator propagator = new NumericalPropagator(integrator);
propagator.setAttitudeProvider(new LofOffset(inertialFrame, LOFType.TNW));
double duration = 2*24*3600;
ManeuverTriggers triggers = new DateBasedManeuverTriggers(firingDate, duration);
double thrust = 250*FastMath.pow(10,-3);
double isp = 2000;
propagator.addForceModel(new NewtonianAttraction(mu));
propagator.addForceModel(dragForce(crossSection,dragCoef));
propagator.addForceModel(new
ConfigurableLowThrustManeuver(ThrustDirectionAndAttitudeProvider.buildFromFixedDirectionInSatelliteFrame(Vector3D.PLUS_I),triggers,thrust,isp));
File csvFile = new File("bck_propagation_data.csv");
FileWriter fileWriter = new FileWriter(csvFile);
PrintWriter printWriter = new PrintWriter(fileWriter, true);
String header = "time,pos_X,pos_Y,pos_Z,inclination";
printWriter.println(header);
propagator.getMultiplexer().add(20, (state) -> {
LocalOrbitalFrame localFrame = new LocalOrbitalFrame(inertialFrame,LOFType.TNW, state.getOrbit(), "TNWframe");
Transform inertialToLocal = inertialFrame.getTransformTo(localFrame,state.getDate());
printWriter.printf(Locale.US, "%s, %.3f, %.3f, %.3f, %.3f %n",
state.getDate(), inertialToLocal.transformVector(state.getPVCoordinates().getPosition()).getX(), inertialToLocal.transformVector(state.getPVCoordinates().getPosition()).getY(),
inertialToLocal.transformPVCoordinates(state.getPVCoordinates()).getPosition().getZ(), FastMath.toDegrees(state.getI()));
}
);
propagator.propagate(initialDate,firingDate );
System.exit(0);
```

This is a code that I wrote reading the forum and the tutorials of Orekit and adapting the answers to my needs. However, I have the following questions:

- What is the cause for which I receive the following error when running this code and how could I solve the issue:

```
Exception in thread "main" org.orekit.errors.OrekitException: NaN appears during integration near time -20
at org.orekit.errors.OrekitException.unwrap(OrekitException.java:154)
at org.orekit.propagation.integration.AbstractIntegratedPropagator.integrateDynamics(AbstractIntegratedPropagator.java:511)
at org.orekit.propagation.integration.AbstractIntegratedPropagator.propagate(AbstractIntegratedPropagator.java:425)
at org.example.Main.main(Main.java:128)
Caused by: org.hipparchus.exception.MathIllegalStateException: NaN appears during integration near time -20
at org.hipparchus.ode.nonstiff.EmbeddedRungeKuttaIntegrator.integrate(EmbeddedRungeKuttaIntegrator.java:269)
at org.orekit.propagation.integration.AbstractIntegratedPropagator.integrateDynamics(AbstractIntegratedPropagator.java:477)
... 2 more
```

- To me it seems a little bit strange that when converting to RTN the value that I obtain when calling .getY() is around the value of the semimajor axis. Is that normal or do I make a mistake when computing RTN? The results I am talking about were obtained when performing the backward propagation without the aerodynamic drag force model added. To me it was strange as well that the Z coordinate was always 0. That I thought to be normal considering that I configured the low thrust maneuver to be in along-track direction, but am I right?

Thank you very much,

Razvan