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Just invented a new airfoil

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Aerowerx

Well-Known Member
Joined
Dec 1, 2011
Messages
6,092
Location
Marion, Ohio
As some of you may know, I have been playing around with tailless aircraft design. Although it can be made to work with any airfoil, the most common ones use have a bit of reflex (turned-up at the trailing edge).

Recently I was looking at the Northrop N1M and N9M, and wondering if they could be replicated, at least used as an inspirational starting point. One interesting thing is that they both use symmetrical airfoils.

If necessity is the mother of invention, then curiosity must be the father. I was looking at the Riblett airfoils and wondered if they would work better than the NACA foils that Northrop used. In his book, Riblett lists his "A" cusped airfoils and his "-" noncusped airfoils. I immediately noticed that the cusp looks like what is used on a reflexed tailless aircraft airfoil, except is is on both the top and bottom surface.

As an experiment, I took the top surface of a cusped foil and stuck it on the bottom of a noncusped. I used a symmetrical foil with no camber for this test, as I am thinking of using it at the center section. I also scaled it to 20% thickness. The basis for this was the GA35, which has maximum thickness at 35% of chord.

This is what the back half of the airfoil looks like:
Capture.jpg

And some of the polars:
Capture1.JPGCapture2.JPGCapture3.JPG

The polars for the noncusped GA35-020 is included for comparison. Notice something interesting. The CL vs alpha and CL/CD vs alpha curves pass through the 0,0 (or close to it) point as it would for a symetrical airfoil. But the new airfoil has a positive pitching moment and, in fact, is essentially flat between about alpha=4 and 12 degrees.

[Edit]Forgot to mention, the polars are at Re=4000000.

I have designated this new airfoil as a GA35x020. It would be interesting to see how it would work for a plank-type tailless aircraft.

If anyone else would like to try it, here are the coordinates:
GA35x020
1.0000 0.00000000
0.9500 0.00400000
0.9000 0.01136000
0.8500 0.02054667
0.8000 0.03080000
0.7500 0.04158667
0.7000 0.05245333
0.6500 0.06294667
0.6000 0.07270667
0.5500 0.08144000
0.5000 0.08886667
0.4500 0.09465333
0.4000 0.09848000
0.3500 0.10000000
0.3000 0.09894667
0.2500 0.09540000
0.2000 0.08924000
0.1500 0.08014667
0.1000 0.06740000
0.0750 0.05902667
0.0500 0.04864000
0.0250 0.03480000
0.0125 0.02504000
0.0075 0.01949333
0.0050 0.01605333
0.0025 0.01166667
0.0000 0.00000000
0.0025 -0.01143333
0.0050 -0.01604000
0.0075 -0.01930667
0.0125 -0.02458667
0.0250 -0.03438667
0.0500 -0.04824000
0.0750 -0.05842667
0.1000 -0.06662667
0.1500 -0.07922667
0.2000 -0.08825333
0.2500 -0.09454667
0.3000 -0.09845333
0.3500 -0.09994667
0.4000 -0.09913333
0.4500 -0.09620000
0.5000 -0.09144000
0.5500 -0.08516000
0.6000 -0.07760000
0.6500 -0.06897333
0.7000 -0.05957333
0.7500 -0.04974667
0.8000 -0.03988000
0.8500 -0.03002667
0.9000 -0.02016000
0.9500 -0.01029333
1.0000 -0.00042667
 
Last edited:
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