General Question Orbital Elements for ISS

[PaulG]

HI. I want to enter the orbital elements for the ISS based on real life positioning. I got the elements from http://calsky.com.

My question is on exactly how I should enter these into Orbiter

Kepler Elements
Semi-major Axis: 6730.773 km
Eccentricity: 0.0008685
Inclination: 51.6408°
Argument of Perigee: 246.9642°
Right Ascension of Ascending Node: 261.4942°
Mean Anomaly at Epoch: 257.7333°
Epoch (UTC): 16. Apr 2009 19:43:12 (JD=2454938.32166763)

I assume that since the Inc is 51°, this is all based on
the equatorial frame so I should set that first.
Since it is the Equ fram, RAofAN is the LAN right? If so, is
Lpe then RAoAN + Argument of Perigee and
eps RAoAN + Mean Anomaly at Epoch?

I assume then, for Epoch, I should enter in the
JD - 2400000.5 right?

Thanks,
Paul

 

[ripley1]

I think you can use scenario editor for this. When you run a scenario,click
(F4-custom-scenario editor-edit) Hope this help

 

[PaulG]

I know. My question is:

How do the Kepler Elements I copied above equate to the parameters in the editor? For instance, Lpe is not one of the Kepler Elements, so do I just add LAN and the Arugment of Perigee? Are those values correct in the equatorial frame?

 

[DaveS]

Get the TLEs you want and plug them in with this modified Scenario Editor: [ame="http://www.orbithangar.com/searchid.php?ID=2617"]Scenario Editor TLE[/ame]

 

[PaulG]

Well, its close. The 2 Line states:

ISS 73.0 44.5 27.5 -2.0 d 402
1 25544U 98067A 09108.52667130 .00010800 00000-0 84397-4 0 6977
2 25544 051.6408 252.7625 0008582 253.7918 187.6845 15.72234264596569

I notice that the Ecc in orbiter is 0.000647941 instead of 0.0008582

Other values may be off. So why this discrepancy? How would I take the values above and MANUALLY put these into Orbiter? What calculations need to be done?

Thanks.

 

[Quick_Nick]

Here's some info on Orbiter's elements:
http://www.orbiterwiki.org/wiki/Orbital_elements#Orbiter_Elements

Just calculate, put it together, and put the ELEMENTS line in a scn file under a ship.

(See a little higher on the page for TLE info)

 

[PaulG]

I don't understand how this relates, as the only calculation is "a".

Why is the Ecc in Orbiter different than the Ecc in the 2 line element file?

 

[Quick_Nick]

Sorry, I wasn't really answering your last post, just the idea of the thread. Just kind of an alternative to other possible ways of implementing elements.

 

[tblaxland]

Well, its close. The 2 Line states:

ISS 73.0 44.5 27.5 -2.0 d 402
1 25544U 98067A 09108.52667130 .00010800 00000-0 84397-4 0 6977
2 25544 051.6408 252.7625 0008582 253.7918 187.6845 15.72234264596569

I notice that the Ecc in orbiter is 0.000647941 instead of 0.0008582

Other values may be off. So why this discrepancy? How would I take the values above and MANUALLY put these into Orbiter? What calculations need to be done.

You get this when you use the TLE Scenario Editor DaveS linked to? I suspect it is because of the way the SGP4 algorithm works. The elements in a NORAD two-line element set are not the actual osculating orbital elements for the given epoch. They are a mean orbital element set and the SGP4 algorithm defines the method for translating that into osculating elements for a desired epoch. The TLE Scenario Editor does that work for you.

Back to your OP: In the scenario editor, you will want to set the "Orbit Reference" to "Earth" and "Frame" to "ref. equator". Once you have done that, then Longitude of Ascending Node is the same as Right Ascension of Ascending Node (only true for Earth-centric orbits). Longitude of Perigee = Argument of Perigee + Longitude of Ascending Node. Mean Longitude at Epoch = Longitude of Perigee + Mean Anomaly at Epoch.