Orbital elements of manmade stuff

Max-Q

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Where can I find the orbits for manmade objects in heliocentric orbit? I refer to things like Apollo 12's S-IVB, LM-4 Snoopy, that Tesla, etc.
After several hours spent fruitlessly googling, I am pretty much at a dead end. Also, how to convert the data to what would work in Orbiter.
Specifically in the case of LM-4, it was initially assigned an asteroid designation, but when it was found to be manmade, the asteroid designation was dropped. Surely the data is still out there!
 

BrianJ

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Hi,
all the items you mention are listed on JPL Horizons:
They used to have a convenient drop-down list of all the spacecraft they had data for - but they changed their website, so now you have to search for something like "Tesla" or "LM" or "S-IVB".
Whether they have data for the time period you require, I don't know.
If you're not familiar with their ephemeris table format and options etc., give a shout here.
(If using state vectors you need to swap Y<->Z and convert from km to m).
Cheers,
Brian
 

Max-Q

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Taking the Apollo 12 S-IVB: I am not sure how to get this data into orbiter.

Code:
  To extrapolate to future times, the following heliocentric IAU76 J2000
  ecliptic osculating elements can be manually input into Horizons as starting
  conditions for a ballistic numerical integration using the telnet or e-mail
  interfaces:

  EPOCH = 2440561.50 ! A.D. 1969-Dec-06 00:00:00.0 (TDB)
   EC= 3.788077679293159E-02 QR= 9.895270525906216E-01 IN= 5.742003975916762E-02
   OM= 3.492340571440683E+01 W = 4.664715199038439E+01 Tp=  2440569.022123539820

... where

    EPOCH    Epoch Julian Day Number, Barycentric Dynamical Time
      EC     Eccentricity, e
      QR     Periapsis distance, q (au)
      IN     Inclination w.r.t xy-plane, i (degrees)
      OM     Longitude of Ascending Node, OMEGA, (degrees)
      W      Argument of Perifocus, w (degrees)
      Tp     Time of periapsis (Julian Day Number)

  Uncertainties and errors in the extrapolated position prediction will increase
  as the time from the EPOCH increases. Predictions for this man-made object
  years past the EPOCH are problematic due to unmodeled-but-cumulative solar 
  pressure and out-gassing accelerations, yet no further tracking data to
  characterize the forces.

  Supplemental orbital elements (FYI, at same epoch):
   N = 9.449436351342855E-01 MA= 3.528920172383751E+02 TA= 3.523284678961845E+02
   A = 1.028486936673184E+00 AD= 1.067446820755747E+00 PR= 3.809751043498384E+02
 
... where

      N      Mean motion, n (degrees/day)
      MA     Mean anomaly, M (degrees)
      TA     True anomaly, nu (degrees)
      A      Semi-major axis, a (au)
      AD     Apoapsis distance (au)
      PR     Sidereal orbit period (day)

This seems to be the data I want, and mostly it seems to plug into scn editor... everything except Mean longitude at epoch (scn editor). I guess Time of periapsis (data above) can be converted to get a value in degrees? I guess I should also ask, what is the easiest way to integrate this trajectory out to today? I understand that if I set the epoch correctly, the object will be in the right place today... except for things like earth close approaches, etc.
 
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BrianJ

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Hi,
I guess you could use the Scenario Editor, or you could use ELEMENTS in your scenario, instead of RPOS and RVEL state vectors.
See your Doc/OrbiterConfig.pdf for definition of ELEMENTS parameters, and your Scenarios/DeltaGlider/BrightonBeach.scn as an example.

The Horizons data gives you Semi-major Axis (need to convert to meters for Orbiter), Eccentricity, Inclination and Longitude of Ascending Node.
You'll need to calculate (just a bit of addition!):
Longitude of Periapsis = Longitude Ascending Node + Argument of Perifocus
Mean Longitude = Longitude Ascending Node + Argument of Perifocus + Mean Anomaly
And your Utils/Date.exe will convert the Epoch Julian date to MJD for you.

Orbiter should propagate the Keplerian elements to the date you set in your scenario Environment section.
However, it won't take into account any orbit perturbations from gravity sources or radiation pressure etc.
So, over a span of ~50 years, there may well be considerable innacuracy, especially if there were any close approaches to planets etc.

You could use Orbiter itself to propagate the orbit by setting the scenario date close to the Epoch date and running the simulation. At a time acceleration of x10000, 50 years would take about 2 days of running the simulation. (I'd be using lower time acceleration for any close-approaches, too).

There might be a better way to propagate the orbit over long time periods using GMAT or Horizons itself - but that's outside my limited knowledge!

Cheers,
Brian
 
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