Maximum thickness of wing chord

I'm interested in learning about what determines where the maximum thickness of a symetrical stunt wing should be. Different designs vary between say 20 and 40% of the wing chord. Are there norms? What determines the position? What happens as the max. point is moved back? Simple answers please!

Simple answers? You have got to be kidding. Just wait.

OK, I'll try to make this simple..

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The center of pressure moves aft!

((Sorry, I have no idea.. I'll be interested to find out.))

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"I was under medication when I made the decision to burn the tapes." -Richard Nixon

Lovely graphic though - what do you take to induce that!!!!!

LAST EDITED ON Dec-21-03 AT 03:44 PM (CST)
 
Who was it(Charles Mackey?) that built an experimental stunter called the Bulldog several years back with a similar airfoil to this. I think the Bulldog 's airfoil was 4" thick at the high point.

I remember Bob Palmer mentioning that his Thunderbird airfoil had the high point pretty well back. He was thinking "laminar flow", like the P-51, as I remember. But laminar flow would not apply for our low Reynold's Numbers.

On the other hand, my Fancher DOCTOR has the high point very far forward, and in fact it is supposed to balance at about 15% chord.

I will be monitoring (but remember, I have a short attention span in aerodynamic discussions).

>On the other hand, my Fancher DOCTOR has the high point very
>far forward, and in fact it is supposed to balance at about
>15% chord.

To first approximation, the high point position doesn't really factor into the CG location.

>Who was it(Charles Mackey?) that built an experimental
>stunter called the Bulldog several years back with a similar
>airfoil to this. I think the Bulldog 's airfoil was 4" thick
>at the high point.

A small write up and photograph of Charles Mackey's X-33 Bulldog is on pages 96 & 97 of the Mar/Apr, 1997 Stunt News. The write up says the airfoil is .3333333, etc. It looks somewhat like a Zilch Xpendable fuselage with fat wings. I wrote Mr. Mackey about getting plans. He replied that he usually didn't draw plans, he built, flew and then went on to something new and different.

Hey! Looks like a Polliwog Fractal.

Flaps or no Flaps? Most people here are talking flaps. In that case, read Al Rabe's articles on the wing sections he developed for the Sea Fury in the 70's. His main thought was to fair in the curve of the airfoil into the deflected flap to try and get smoother airflow and less drag in hard maneuvers. Has worked very well. Doing that the highpoint of the unflapped part of the wing tends to end up around 33%. Was a real key insight into making a realistic semi scale model generate enough lift to fly a decent stunt pattern.

For unflapped wings it is mainly the shape of the streamline used. It is hard to beat the NACA 0018. You can stretch to to put the high point further forward or back as needed for the spars. A nice, smooth foam wing helps. It is quite possible to mess up the airflow by poorly placed spars, sheeting, D tubes, etc. near the high point.

>
>For unflapped wings it is mainly the shape of the streamline
>used. It is hard to beat the NACA 0018. You can stretch to
>to put the high point further forward or back as needed for
>the spars.

If memory serves me right, I think NACA 0018 is 18% thickness at 29.7% chord. I suppose what started me thinking is we know this works well (but why?). There are lots of NACA modifications which vary the max. thickness between 20% and 40%. I guess I hoped it would be straightforward to discover what happens as this parameter is changed.

as far as I've been able to learn over the years, there is no "simple" answer, except maybe, that is what the air "likes". part of the reason the NACA 0018 works well is that it is very "smooth" to the air- no changes the profile that cause abrupt changes in pressure or velocity. So when the flow does separate towards the back, it does so smoothly and consistently. Moving the highpoint back, at least on an unflapped airfoil, generally doesn't do anything good, as Bob points out below. I've made some pretty awful airfoils just playing around with the shape behind 50%. A highpoint at the 15-30% chord position seems to work the best, as long as the profile is a smooth one.

Take a look at Igor Burger's info on flapped airfoils. He's got some nice graphs that show what can happen when the airflow separates at the hinge line. It can generate big jumps in drag with very little change in angle of attack or control movement. that goes a long way to explaining why some planes with not so good airfoils will fly just fine until you need a quick change of control or get in a tight spot. Suddenly the plane gets much draggier than you expect, loses speed, and starts to flop around. That can cause a lot of anguish if it happens in the third corner of the hour glass!!

For NACA00xx, xx is the % thickness (0018 is 18% as you said) and the high point is at 30% of total chord.

Although I was joking earlier, I'm pretty sure that moving the high point aft does move the center of lift aft.. What the significance of that is, I wouldn't pretend to know.

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"I have often regretted my speech, never my silence." -Publilius Syrus

Back in 1950, when I designed the WILDMAN 60, I asked a club member to recommend an airfoil. He gave me the numbers for NACA 0018 out of Abbott & Dohenoff, and it was an immediate success.

>chord. Are there norms? What determines the position? What
>happens as the max. point is moved back? Simple answers
>please!

Looks like you have answers except for effect.

My personal observation was that further back the high point of the airfoil, the harder it is to do squares and keep lift. Round maneuvers are much easier. There are lots of other factors that play into this, though, and my observation and experience is purely subjective.

Modern powerplants help as well...

This is not a simple topic.

I'm not a CL guy yet (have built one, not flown yet), but it seems to me you want the airfoil that produces the most lift for an acceptable amount of drag, at the Reynold's numbers the plane flies at. I don't know what Re that is as yet, but I'm guessing around 300k? For a low aspect ratio design, the lift becomes a little less important as if you go to a high Cl (i.e. a lot of lift for the speed you are going) the plane will bog down and have lots of drag. (If any of you guys fly RC and have flown an IFO in a tight turn you will know what I'm talking about.)

The "laminar" sections don't seem to work all that well at low Re (say maybe under 500k????), and if the Re is low enough, the flow may well be laminar anyway. I suspect that this Re may be lower than that in CL stunt. For high Re, laminar flow foils are probably good, if you need to minimize drag and go as fast as you can. These have the high point further back, but they also have particular shapes that have to be kept and the workmanship has to be very good. ALso, every time you splat a bug or ding the wing, the flow will be turbulent behind it.

It is my impression that as the high point is moved back, the nose radius becomes sharper and the foil is likely to stall more easily.

At low Re, the high point will probably be further forward. I think some of the low Re (like 100k or less) modern airfoils (i.e. Drela) I've seen have the high point at maybe 22% and are very thin, and almost flat surfaces for the last 75% or so. Kind of like a thinned down version of some full sized aerobatic airfoils. Stunt can probably go much thicker.

If you are willing to put in a few days of work, and are technically minded, the Xfoil program, which is free, can tell you more than you want to know about this. I'm not at all sure it's worth the trouble.

Someone mentioned "center of lift". Without flaps, on a symmetrical airfoil, the "center of lift" is going to stay put at approximately 25% of chord, no matter where the max thickness is. Center of lift is not a good term because if an airfoil has camber or flap deflection and is at a zero lift angle of attack (as in a dive) the "center of lift" can be at infinity, i.e. a few light years behind the airplane. (That's what happens when you divide by 0.) It's more correct to speak about aerodynamic center, and pitching moment, as both stay about the same in normal flight including maneuvers. AC (I think that's the correct term, anyway) is always near 25% if the airfoil is not stalling or something. If you multiply pitching moment by the square of the airspeed, the density, and some other factors I forget, you will get the torque the wing produces about the AC.

But I see that I haven't provided a simple anwer.

>I'm interested in learning about what determines where the
>maximum thickness of a symetrical stunt wing should be.
>Different designs vary between say 20 and 40% of the wing
>chord. Are there norms? What determines the position? What
>happens as the max. point is moved back? Simple answers
>please!
>
>Adrian