Airfoil simulation

BoredEngineer · 2016-04-05T12:34:46

In continuation to over 2 years old topic: http://www.gamedev.net/topic/661025-simulation-of-drag-in-flightsims/ I've got back to working on airfoil simulation. The focus of the project now is not a flight simulation specifically but variety of different machines, some of them, happen to utilize airfoils. I've got something working over the last week and tried to teach one of the existing machines to fly. First "successful" take-off: Second attempt with better balanced wings and control surfaces: Since then I've improved controls (control surfaces where stalling) and added some rudimentary GUI. If you want to give it a try, compiled version is here: https://www.dropbox.com/s/55dazkm8q4s0axn/MMT_Content.zip?dl=0 The chassis of the vehicle has rather high drag and not really build as an airplane. So doing some fancy maneuvers is rather difficult, for example, I can't do a loop on it as it looses velocity super fast as soon as you start to pitch up. I have suspicion that total drag is too high (could be because of the chassis itself) and with zero throttle terminal velocity is around 65 km/h. Anyway, for proper test I'll add something like a Junkers J1 as I already have a low poly model of it. --------------------- End of introduction ----------------------- There are couple of issues I can see with current model. Calculation of airfoil drag is done using this formula: Cd = Cd0 + Cl^2 / ( pi * Ar * e) where Ar is aspect ratio (span/chord) and "e" is efficiency factor. Cl was calculated for the symmetrical airfoil and so I'm ignoring pitching moment, curve of Cl looks like this: https://www.dropbox.com/s/hdw58m4hel64cfh/NACA_0012_Cl_smooth.PNG?dl=0 For a proper airplane I'm planning on cutting wings into multiple pieces, which means that Cd approximation won't work anymore as using aspect ratio for segment of the wing (rather than full wing) will drastically lower drag coefficient. I have a separate curve for Cd: https://www.dropbox.com/s/yu1ww6p8m6cp7ne/NACA_0012_Cd_smooth.PNG?dl=0 which I can apply directly, but it doesn't take into account the shape of the wing as a whole. So I'm baffled of how to properly describe end sections of the wings which might not be straight but have more of an ellipsoid form. For the full wing this is suppose to be taken into account by efficiency factor. Another point that I want to improve is to model control surfaces as a part of the wing and not as a separate airfoil as it is done right now (very inaccurate). The question is, can I simply sum up Cl and Cd coefficients of wing section and aileron section, taking into account their respective angles of attack, or I have to get new curves with different states of control surfaces and blend between them? The other important part I'm missing is ground effect, at least this seams to be more significant than proper simulation of the engine in respect to gliding behavior of the aircraft.

Math and Physics Programming

Started by BoredEngineer March 21, 2016 03:30 PM

9 comments, last by BoredEngineer 8 years ago

BoredEngineer

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Author

April 05, 2016 12:34 PM

Got some good progress:

Executable is here: https://www.dropbox.com/s/55dazkm8q4s0axn/MMT_Content.zip?dl=0

Implemented algorithm posted earlier. What I do now is split wing into segments, so some segments can have control surfaces. Contribution from each segment is scaled in proportion to wing area that segment covers. Cl/Cd/Cm are not parametrized but stored as curves for a single Reynold's number of 500k. It looks like this for NACA0012:

https://www.dropbox.com/s/gf0fppzi3l8uqtp/NACA0012_Cl_With_Flaps.JPG?dl=0

https://www.dropbox.com/s/ib04mky85acra3t/NACA0012_Cd_With_Flaps.JPG?dl=0

https://www.dropbox.com/s/mtfv0fexsvr2ejp/NACA0012_Cm_With_Flaps.JPG?dl=0

Red curve is flaps at 0 degrees deflection, green is at +15 and blue is at -15.

I've got curves from JavaFoil for a range of -20 to +20 degrees of angle of attack and then smoothed them out with curves for high angle of attack according to this article:

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

Another thing that was added is taking into account beta angle of attack as this:

q = 1/2 * p * V2 * cos2 ?

Overall, using approximations from Thin Line Theory and combining control surfaces with wing surface got more interesting results. Airplane reacts less linearly to steering. Both Fokker Dr.I and "Flying Aerosled" are using new model. Fokker took about an hour to balance, had to tweak position of center of mass and angle of horizontal stabilizer. The damn thing really likes to flip over the wheel during take-off. Moving center of gravity by 10 centimeters to the front, guarantees flip, moving it 10 centimeters to the back makes it unstable in flight. Wing geometry is limited to rectangular wings without tapper or sweep, I'll add function to account them in area and aspect ratio calculations later.

Overall I'm happy with the result. But what I can't do so far is to get airplane into unrecoverable spin. I can stall wings or tail, but as fuselage has less airdrag along the frontal section (propeller drag is not modeled) it just turns itself back into position where wings can catch air. What should I do about this? Is propeller drag plays important role in this process?

The other thing I didn't managed to get working is delta wing. Perhaps NACA0012 is not suitable for them or I need a higher quality and more precise pitching moment curve.

I haven't found formulation for pitching moment in approximation of thin line theory and using this:

Torque = 1/2 * p * V2 * WingArea * WingChord * Cm
not sure if this is precise enough to balance delta wing pitch.

Edit: Fixed links to coefficient curves

Airfoil simulation

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Airfoil simulation

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