Best airfoil for high stall angle?

[JackBauer]

Hey guys,

I'm looking for the airfoil that has the highest lift/drag ratio at the highest possible angle of attack. It would be preferabbly symmetric, even better if it's a NACA 4 digit (or something that would be easy to model).

Any ideas? Thanks in advance

 

[GregLocock]

So what is your tradeoff between L/D and alphamax?

That is, which is better, one with an L/D of 19 at 20 degrees, or one with an L/D of 18 at 21 degrees?



Cheers

Greg Locock

Please see FAQ731-376 for tips on how to make the best use of Eng-Tips.

 

[JackBauer]

Hi Greg,

I guess my question is poorly stated, and does not reveal the the true intent of my investigation.

X = (lift)cos(90-a) - (drag)cos(90-a)

(a = maximum angel of attack)

The angle of attack will be swept through (nearly sinusoidally) from -a -> a, so it would be very nice not to have it act erracticly inbetween.

I've been thumbing through some of the NACA 4 series... it seems like the 00xx is a good range ot look through but I thought I'd consult the expertise of the members here before I take too many steps in the wrong direction

Thanks again.

 

[rb1957]

i think your equation isn't quite right ... if X is meant to be the vertical component of the wing force then it'd be Lift*cos(a)-Drag*cos(90-a) ? ...
but then you know what you're doing, just looks funny that two forces at 90deg to one another would have the same trig function (rather than sin and cos) ...

reading your post again, i think you're doomed to failure. first the lift force is going to dominate (being much bigger than drag), second, lift is linear with aoa (but then maybe that's not behaving erractically), third, drag is not usually expressed as a function of aoa (but maybe you could transform the Cd plot to achieve this)

good luck

 

[JackBauer]

Ah, I am truely dyslexic on the keyboard.

X = (lift)sin(a) - (drag)cos(a) is what I meant to say

I choose ended up going with the NACA 0012, but just out of curiosity... if it is a symmetric airfoil, why is it that programs such as designFOIL and xFOIL generate different lift coefficients for +/- angles of attack?

For example, for
AoA: -10 degrees
Cl: -1.193 <----
Cd: 0.0114

AoA: 10 degrees
Cl: 1.239 <----
Cd: 0.0114

If the airfoil is symmetric would it really matter if the anlge of attack is positive or negative? Or is it just bad programming? Th

 

[rb1957]

i don't know the programs ... maybe 10deg is getting close to stall and maybe non-linear "things" are happening ... try smaller angles

 

[WhiteRabbit]

Theory of Wing Sections shows a linear AOA vs CL curve. However, around +10 AOA it appears that it is beginning to have flow separation. I do not know if your programs are accounting viscosity or not. This book was derived using actual wind tunnel testing.

It may also be using a zero lift unit conversion to bring it back to a unitless coefficient. It may be adjusting the incidence angle to compensate.

Does the Cl show zero at 0 AOA?

As for the lower Cl vs Cd, it appears that you are looking for a wide drag bucket.

Check out Theory of Wing Sections, Ira G Abbott and Albert E. Von Doenhoff Library of Congress # 60-1601.

There are coordinates for the NACA series airfoils so you can manually input the points and spline the curves. This way you won't be restricting yourself to 1000 series symmetrical airfoils.

By the way...the highest Lift/Drag ratio is the holy grail of aerodynamics. The airfoil selection all depends on your flight envelope and the Reynolds Number you plan to use it.

 

[Sparweb]

By calling out a NACA 0012, you aren't "technically" done specifying the airfoil. The standard assumption is that the maximum thickness will fall at 30% chord. There is a way to modify the thickness function to move the point of maximum thickness forward or back. This may not have much effect on the lift curve, but it will on the drag "buckets", and hence, L/D at high angles of attack. Moving the thickness back makes it start looking like the 6-series airfoils, without the more complex designation format.

Just when you thought it was complicated enough...

You'd better read Abbott & von Doenhoff's Theory of Wing Sections. It's not expensive, and easily obtained because I think it's still in print (Dover Press).


Steven Fahey, CET

 

[alphaguru]

Also be aware that you may not need a 19 L/D for your mission. If it is only for stalling, takeoff, landing ect. you can select a airfoil that is suitable for the bulk of your mission (cruise possibly?) and utilize high lift devices to achieve the characteristics that you need for the other portions of your mission. Like slats to increase the stall angle and decrease the stall speed. These can then be undeployed. This will proballky allow you to reduce drag. Drag = fuel = weight = cost

 

[matt00000]

Jack If you are after lift at very high AoA go for a delta planform they can be symetrical.

You will get vortex lift at high AoA and they are
typically good up to 25degress AoA