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10

That's a great software-based experiment! What is this about? It's about drag and Newton's 2nd law of motion! $$F = \frac{dp}{dt} = ma$$ but in the context of orbital mechanics. We can re-arrange Newton's law as $a = F_D/m$ where $F_D$ is the drag force, and the drag equation is $$F_D = \frac{1}{2} \rho v^2 C_D A$$ where $\rho$ is the density at that ...


9

From Force=Mass*Acceleration for a given starting drag force (from the constant drag area, altitude and velocity) increasing the mass will reduce the acceleration (deceleration in this case) slowing the satellite and causing the perihelion to lower. So being heavier does not directly make the satellite fall slower, but does change how quickly atmospheric ...


9

GMAT has been released under the Apache license since R2013a, so I'd say it's about as wide open as you can get.


5

To answer your questions, Do the terms (J2, etc.) change with time? Yes, they do. The Earth is still rebounding from the end of the last glaciation, and the Earth's rotation rate is decreasing due to transfer of angular momentum to the Moon's orbit. The end result is that $J2$ is decreasing by about 3 parts per million per century. You don't need ...


4

I am not familiar with GMAT, but there is another route to solve this challenge using an extensively validated open-source solution. You can use the Orekit Astrodynamics library to build one Moon-centered and one Earth-centered orbit. Orekit has the ability to compute what is known as "Intersatellite visibility" (in STK as Access Times), which is ...


4

You can find a ton of information by looking in the "General Mission Analysis Tool (GMAT) Mathematical Specifications" (located at <GMAT installation directory>\<version>\docs\GMATMathSpec.pdf, or find the latest version here) You can look at section "4.2.4 Solar Radiation Pressure", here is the R2018a version: 4.2.4 Solar Radiation Pressure ...


4

From what I understand, you need to ensure that both spacecraft are phased by 180 degrees. There are different orbital elements used to assess the phasing of spacecraft. The main one is the true anomaly, but it's ill-defined for circular orbits (in which case I would recommend using the "true longitude," defined as AoP + RAAN + TA, it is unrelated ...


4

The notion of "the most realistic" is quite debatable. In fact, if you want the most realistic, you should be including the a precise 3D model of your spacecraft which includes the albedo of each infinitesimal surface of the spacecraft, take into account the shadowing of some parts of the spacecraft with respect to others (which will affect the reflectivity ...


4

I'll expand on my comment: I'm not familliar with the software, but if you start the satellite at the same everything except time (12 hours difference) then the satellite will be starting in a different gravitational potentials for the two cases. I am guessing that if you ran it 100 times spacing your starting Epochs evenly over 24 hours, you'd see their "...


2

To answer your first question: No, these coefficients do not change over time, at least not in the time scale you are working. These terms depend on the internal structure of the Earth, whose changes are slow and take thousands of years. The higher order terms may vary from one model to another, but their influence in the movement is not significant. ...


2

From the provided plots, O1 is the orbit which returns to the initial AoP after the five year period. Hence, that is the best answer to the question "which of the two orbits is the most stable after a five year period?" Moreover, the overall amplitude of the change in AoP (the difference between the minimum value and the maximum value) throughout the five ...


2

Here is a NASA repository of space weather data, including f10.7. https://omniweb.gsfc.nasa.gov/form/dx1.html I would bet GMAT gets its data there or a similar site. EDIT Apologies everyone for the lacking answer, I'm new to StackExchange. That original link actually only has historical data, so it probably isn't the best source. I found this old GMAT ...


2

there's effectively none. Hypergolics are not allowed in cubesats, and delta-v req. LEO->LLO is 4 and 8 km/s (high-thrust low-thrust) https://en.wikipedia.org/wiki/Delta-v_budget So, spec it for moonlaunch rideshare, if possible. You might want to have a look at Cornell university electrolysis solution: https://en.wikipedia.org/wiki/Cislunar_Explorers ...


2

In the InternalODEModel of your script, you have J2 enabled. Set the degree to zero and the problem should be fixed. Create ForceModel InternalODEModel; GMAT InternalODEModel.CentralBody = Earth; GMAT InternalODEModel.PrimaryBodies = {Earth}; GMAT InternalODEModel.Drag = None; GMAT InternalODEModel.SRP = Off; GMAT InternalODEModel.RelativisticCorrection = ...


2

...I don't know how this can be quantified within the simulator... ... I'd like to be able to evaluate the phase error... If I correctly understand your question, in your case you need to create a variable in GMAT, in that variable you will write a difference between longitudes of satellites after finishing of phasing (at any time endeed). Something like ...


2

I had read many years ago in one book the rough empirical rule for orbital maneuvering and rendezvous: suppose that two spaceships follow the same traectory (circular) with distance between them, say, 100 km, and the second ship try do intercept the first one in the one orbital period. The second ship need to change orbit from circular to elliptical with the ...


2

For more complicated scenarios like these, you probably need to use GMAT's Python interface. I would recommend the following: Setup both spacecraft to be propagated together: I'm not sure this is available in the GUI, but in the script, change the Propagate statement to Propagate Synchronized (cf. the docs). Instead of propagating for the whole orbit, ...


2

The basic approach is to make a long list of times, compute positions and observing angles at each one, and check whether line of sight (LOS) is obscured by anything. Do it at, say, 5 minute intervals, and then for any interval during which the LOS changed state, repeat the procedure using 5 second intervals. This won't catch an outage shorter that happens ...


2

At the top of the GMAT window there are animation controls. The buttons are play, stop, faster, and slower. They are the red/blue buttons. They only light up if the appropriate plot window is active (e.g., ground track plot or orbit view plot). I highlighted them in the image below. As Chris mentioned, there is a new plotting interface available in GMAT ...


2

In the "Output" section of the "Resources" tab, right-click and add "OpenFrames Interface". After the mission is completed, you'll be able to play back the mission at a reasonable speed. Also, I would recommend you to update to R2020a: I think that everything is compatible with R2019a.


1

The way I've modeled that in the past is to using Python to generate a script with N spacecraft in it, each with their own initial states (chosen from the known distribution), and using a Propagate Synchronized statement. Each phase of the mission would have its own script: upper stage in one script, and use the final state of that propagation statement to ...


1

Based on your example, it looks like an indexing problem. You defined anArray to be two-dimensional, with 25 rows and 1 column. Then, in your For loop, you access the array with a one-dimensional index. When you do this, GMAT assumes that the first index, the row index, is always 1. To check this, look at the "Array" section of the reference guide. ...


1

Try to set TOI.Element1 to -1 (Initial Value) in the Vary 'Vary TOI' command. It looks like you use too small MaxStep (0.002) for the TOI.Element1 in both Vary commands. Try to change it to 0.2 or 0.1. GMAT needs more iterations with your settings for the convergence. Honestly, I did not see (and use) before more than one burn (impulsive or finite) in the ...


1

About of how does GMAT calculates RadApo and INC: GMAT calculates the RadApo property from the keplerian elements, which are obtained directly from the cartesian state ($\vec{r}$ and $\vec{v}$) so any change in the spacecraft state (a maneuver which applies a $\Delta v$) is reflected at once in RadApo. The same happens with INC, which is a keplerian element....


1

This is not always the case. Suppose the satellite with the higher mass has, say, the form of a sphere, and the lighter one that of a long massive spear with a sharp point gradually extending to the small diameter of the spear, which has a higher mass at the front than at the back. On entering the atmosphere, the spherical satellite will experience more ...


1

Are you required to use 2018a instead of 2020a? If not, I would recommend 2020a. The VF13 pluging allows for SQP optimization: this is needed because in a low-thrust optimization scenario, your problem is not of rank 1, so there isn't an obvious solutio. There are other SQP solvers, such as SNOPT, IPOPT, or OpEn. Of those, IPOPT is an interior point ...


1

I discovered this issue in R2020a few days ago. When comparing the results for the same orbit and simulation parameters in GMAT (MSISE00/JacchiaRoberts) and Orekit (NRLMSISE00), there is a difference of almost 100 km along track after 10 days for a LEO small satellite. I know this a atmosphere issue because without atmospheric model, the results from GMAT ...


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