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Orbit propagation consistency #120
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Hi Fede, Thanks for the data. I see what you're talking about. I'll have to squirrel this away till I have some time to devote to it. In the meantime, the easiest way to get 42 to agree with GMAT is to use GMAT outputs as 42's inputs. I've done this in the past for three-body orbits, and most of the code is in place for central orbits as well. Looking at the code, it appears that there's one or two missing connections. Search 42init.c for INP_SPLINE, and 42ephem.c around line 500, and you'll see what I mean. I should be able to get this done within a week. Regards, |
Hi Eric, thank you for checking the data and for your answer. Thanks |
Hi Fede, I just committed changes so that you can ingest an orbit and interpolate it using a cubic spline. Look for Demo/EllipSpline.txt and Demo/Orb_Ellipse.txt as an example. |
Hi Eric, Since I would really like to use 42 also to simulate orbital maneuvers, do you have a rough estimate of the time that you'll need to fix the propagation core in 42? I was considering to interface 42 with another propagator to simulate the orbital propagation step by step by adding also the propulsion force. But it seems a useless effort for many reasons. Thanks |
Hi Fede, "Fixing the propagation core in 42" is a months-if-ever timeframe. This is not my day job, and the only sure way to achieve 100% agreement with GMAT is to reverse-engineer it. (Old saying: "He with one watch knows what time it is; he with two watches is never sure.") Let me make a suggestion, though, that might be made to fit your needs. The Orb_ input file defines the Reference Orbit. The spacecraft can wander from its reference orbit under the influence of whatever forces aren't bookkept by the reference orbit. So whatever GMAT can model, put those influences in the reference orbit, and apply the leftover forces (thrusters, solar sails, etc) to the S/C, with "Orb Prop" set to ENCKE. This plays to the strengths of both GMAT and 42, and, if it fits your requirements, you're good to go today. Regards, |
Hello @ericstoneking and @f-e-d-e @ericstoneking, First, I would like to thank you, you already helped me a lot with issues I encountered while using your code. After reading your discussion, I want to confirm my understanding that orbit propagation is not accurate in 42. @ericstoneking, I am using your code as part of my work, so it is very important for me to have the correct results. Regards |
One thing to check: are you exerting your Ideal Actuator force in B[0] or N? Looking at 42actuators.c, it's expected to be expressed in B[0]. Beyond that, you might be on your own. I threw together a quick example on my installation, and I get plausible results. But I'm not pushing it as hard as you are, so your mileage may vary. |
Hello Eric!
I'm currently struggling to understand how to correctly propagate the orbit dynamics in 42 and I'd like your help in solving some doubts.
I'll report here the results for a 6U satellite (no sensor or actuator modeled, only a single body) in LEO using only the spherical harmonics gravitational field and no other perturbations. I've compared the results obtained with 42 with the ones given by GMAT in the same testing conditions (LEO orbit, simulation over 90 days using a 10s integration step, EGM96 gravity field model, max degree 2, RK4 numerical integrator).
I've done different tests in 42 considering different configurations for the input files and for the integration options, but none of them provides reasonable results in terms of orbital parameters profile. As I was mentioning, 42 simulation is pretty straightforward as I'm considering a simple satellite equipped with the PASSIVE_FSW. For each simulation test in 42, I retrieve the keplerian parameters from the output files PosN and VelN and I compare their profile with GMAT.
In the attached folder the following set of test results is available:
Here the zip folder containing the results:
42_orbit_propagation.zip
I've done a lot of other tests considering different inputs (by changing the Dynamics Flags options, activating both the secular drift and the gravity perturbation forces) but none of them seem to provide reasonable outputs. Am I missing something? How should I propagate the orbital dynamics of a LEO satellite to take into account correctly the gravity perturbation forces?
Thanks
-Fede
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