EDIT2: Just to be sure, expected value of pitchingMoment = Cm * wingArea * q * wingChord is in radians, right?

pitchingMoment is just engineer-speak for torque, so is in Newton metres.

Cool, thank you!

I thought of adding pitching moment when I get Cl/Cd/Cm for assymetrical aerofoil. So just checked them in JavaFoil and to my surprise, default symetrical aerofoil actually has some Cm values, they are small but still. I guess wikipedia article just makes some generalization.

I'll add a pitching moment at the next iteration. Right now physical simulation is very unstable, maybe because I apply each Lift+Drag contribution from aerofoils separately.

pitchingMoment is just engineer-speak for torque, so is in Newton metres.

I thought of adding pitching moment when I get Cl/Cd/Cm for assymetrical aerofoil. So just checked them in JavaFoil and to my surprise, default symetrical aerofoil actually has some Cm values, they are small but still. I guess wikipedia article just makes some generalization.

I think that's because all the published curves are given for moments around the quarter-chord point, not the actual aerodynamic centre.

I think that's because all the published curves are given for moments around the quarter-chord point, not the actual aerodynamic centre.

Makes sense now.

I was looking into what kind of calculations can be done in JavaFoil, mainly getting Cl/Cd/Cm coefficients. I know that you don't use lookup table anymore and I might take the same approach, for a simple reason that Mach and Reynolds numbers are used as constants during calculation and I can't figure out how to take them out of calculated values. So I've looked into interpolating a set of Cl/Cd/Cm, calculated for different Reynolds numbers at least. Unfortunately linear interpolation won't work so I might have to go with using RBF for this. Main argument against manually building my own Cl/Cd/Cm functions is that I want to run simulation for different than earth atmosphere, which would mean that I couldn't really compare what I'm getting as end result to flight in earth atmosphere.

EDIT: To figure out more details about how I can get coefficients without JavaFoil or taking them from somewhere, I ordered "Aircraft Design: A Conceptual Approach" by Daniel P. Raymer. As a bonus side I hope to figure out the rest of the puzzle - side forces and etc.

Got coefficients exported from JavaFoil for symmetrical NACA 0012 aerofoil at Re=100000.

Cl looks fine, with just a minor weird slope between 7.5 and 10 degrees:

https://www.dropbox.com/s/7uyczcnqgadnxn9/NACA_0012_Cl.PNG?dl=0

But Cd and Cm looks weird, some weird spikes here and there:

https://www.dropbox.com/s/xkcp5f4idmue4fq/NACA_0012_Cd.PNG?dl=0

https://www.dropbox.com/s/2vvenin03vchp4l/NACA_0012_Cm.PNG?dl=0

I guess there is something wrong with precision of calculations or it just can't handle +-180 degrees AoA well.

Got coefficients exported from JavaFoil

...

I guess there is something wrong with precision of calculations or it just can't handle +-180 degrees AoA well.

JavaFoil will only be close when the wing isn't stalled. It should be able to predict the stall itself OK, and a few degrees beyond as Cl drops, but beyond that it's just not relevant.

For Cl you should find that Cl peaks and then drops as the flow transitions from laminar to turbulent. Then Cl will rise again, following a curve that is the same for all normal aerofoils, with a second peak of about 1 at 45 degrees. It then drops down to zero at 90 degrees.

So it should look more like in this article:

http://www.aerospaceweb.org/question/airfoils/q0150b.shtml

But isn't it weird that for such aerofoil (NACA 0015) with a thin trailing edge, at angles 90-180, which means trailing edge is going against flow, mirrored magnitude of Cl is similar to 0-90? I mean, the same aerofoil installed 180 flipped shoudn't work in the same way, right?

I don't think it's that surprising. Probably all "normal" symmetrical aerofoils, when viewed across the whole range like this, will look fairly similar (and all fairly similar to a flat plate). However, the choice of aerofoil for a plane will depend on really small details - e.g. just a tiny difference in drag around the angle of attacks you expect in normal flight will be important for fuel economy/glide slopes, and how it behaves in a stall will be important for safety. The symmetric tear-drop shape is there to minimise drag at low angles of attack, not to maximise lift close to the stall. Also, there will be additional constraints - e.g. needing to be a certain thickness for structural reasons.

You are right. There is indeed a difference but it's very small and hard to see on that graph.

I've crafted coefficients manually for now:

https://www.dropbox.com/s/hdw58m4hel64cfh/NACA_0012_Cl_smooth.PNG?dl=0

https://www.dropbox.com/s/yu1ww6p8m6cp7ne/NACA_0012_Cd_smooth.PNG?dl=0

https://www.dropbox.com/s/1kk7b1c36zolpsk/NACA_0012_Cm_smooth.PNG?dl=0

Go the book and it looks like it has empirical formulas for all the coefficients. So it might be possible to plug Mach and Reynold numbers as variables as well. That would be nit.

I don't need them to be super realistic but more like reasonably - not what JavaFoil provided for 45-180 degrees :)