Does someone know the difference between the 'NACA 0014' airfoil and the 'S 8035' airfoil?

Thanks,
Hans

Hans,

Here's a pic with the 2 airfoils overlayed. The 8035 has the high point a slight bit farther forward and the 0014 has a slightly "blunter" entry and is slightly thicker towards the rear. A little error in final sanding and they're close to the same!

HTH
Milton Dickey

Milton,

Thanks for your aswer. However I realise now that I was not very clear in explaining what I really did want to know.
For many aerobatic applications the NACA 00XX serie is used, for example the Giles ARF (From Flair) that I'm flying now. In some articles the work of mr. Selig is mentioned and the S 8035 is mentioned as a improved aerobatic airfoil.
I'm in the process of designing a slightly enlarged Giles based upon the current ARF that I wil use as a kind of template. Now for the selection of the airfoil I could use the same NACA 00XX airfoil or the Selig.
I'd like to know what I can expect of the aerodynamics (flight performance) from this airfoil in relation to the NACA 00XX

In general I've allways these kind of questions; what kind of performance can I expect for a certain airfoil. I can find hundres of airfoils (coordinates) on the net, but in what area's they excel is very difficult to find.

Hans

I found the coordinates of the S8035 at the UIUC website and its designation as aerobatic but, there are no low speed windtunnel tests to reveal its stall characteristics. Nor was I able to find any low speed windtunnel tests to reveal the NACA0014 stall characteristics for models.

An airfoil for aerobatic application should have low drag at high lift coefficients and high drag at low lift coefficients to limit acceleration so as to present a more uniform appearance of relative speed during maneuvers. It also should have an abrupt stall to facilitate snap rolls and spin entry. Hysterisis in stall characteristics requires higher airspeed for spin and stall recovery. Stall characteristics should have a high priority in aerobatic airfoil selection.

Hans,

Dr. Selig's airfoils are typically geared towards low speed, low reynolds number flight such as that found in RC aircraft. If you aren't familiar with reynolds number, it's effectively the vicosity of air, and with the size wings we typically run and the speed at which they fly, the wings are prone to stall earlier than normal due to something called laminar seperation bubbles. NACA series airfoils were designed for larger chord wings at higher speeds in which these seperation bubble do not occur. Hope this helps a bit.

Wallis

Wallis,

I'm sorry to say that it is difficult for me to effectively understand what you are saying and apply the conclusion to the choise of the airfoil.
Did you mean to say that a small wing will stall aerlier than a large wing. And that the Selig airfoils will counteract this effect for smaller wings as in RC. And therefore the choise of the S 8035 is preferred above the NACA 0014 if I want a 'later' stall.

Thanks anyhow,
Hans

Hans,

sorry for being unclear, but effectively yes, different chord sizes of the same airfoil will perform differently. The selig airfoils are typically designed to perform in RC applications. In fact, they are widely used in RC sailplanes, but would definitely work better for your aerobatic plane over the NACA OO14. Hope this is a little clearer

Wallis

The significance of reynolds number is that at the same reynolds number, the flow pattern over similar shapes will be similar. At a given viscosity and density, reynolds number is proportional to velocity times flow length (wing chord). As the reynolds number decreases the boundry layer thickens. the boundry layer is the layer over the surface where the flow is shearing (going from zero to the local flow velocity). At full scale reynolds numbers the rate at which the boundry layer thickens with decreasing reynolds number is much lower than at model reynolds numbers.

Because the boundry layer thickens at lower reynolds numbers the airfoil drag goes up as though the airfoil was thicker. Because there is less energy in the flow at lower speeds the flow transitions from laminar to turbulent sooner and from turbulent to detached sooner. At model speeds, laminar flow airfoils don't have enough laminar flow to be significantly lower drag than nonlaminar flow airfoils. Since seperation occurs sooner airfoils at low speeds can't achieve as high lift coefficients as the same airfoils at higher speeds.

Fortunately for aerobatic models low drag is not a high priority.

Thanks to all.
Only I guess I have to make a new set of templates for the foam cores. Did allready made the naca and found out later about the selig.

Hans

I wouldn't bother making a new set of templates. These aerofoils are so similar that I'd be surprised if you could make them accurately enough to tell them apart. Then once in the air I'm sure there will be no difference at all.

Leonard

Let's explore the effects of changing the position of the peak thickness of the airfoil. The standard NACA 00xx airfoils have the peak thickness at about the 30% point. What happens if this point is moved forward or aft. My theory is the further forward this point is (up to a certain point) the higher the drag but the smoother and more stable the wing will be during changing AoA manuevers (i.e. 3D stuff).

Ollie, any thoughts?

The X-Foil program models airfoils to produce data similar to that obtained in a windtunnel. Mark Drela has arranged with MIT to make this program available to us modelers. At model reynolds numbers it is a lot more convenient than building a model of the airfoil and testing it in the windtunnel. X-Foil for Windows is available as a free download at:
http://www.charlesriverrc.org/

AFSalmon:

You can test your maximum thickness location hypothesis against your choice of specific airfoils with X-Foil. I'll let you do the work. My intuition is that for symmetrical airfoils more than about 12 % thick and with blunt leading edges, the location of maximum thickness less than 30% of the chord results in a very abrupt stall. Such airfoils can have a straight line coefficient of lift vs. angle of attack characteristic. Airfoils like the R140 with a small leading edge radius and a maximum thickness aft of 30% of the chord may not have a straightline relationship between coefficient of lift and angle of attack.

Thanks Ollie, fascinating stuff, can't wait to try out that X-foil software. If the lift vs. AoA approaches a straight line by moving the peak thickness forward then that would indicate to me a smoother more predictable transition between AoA - up to the point of stall. As you mention then we are talking about a more abrubt stall point. However I'm curious if during the 3d flying mode if this point is not encountered as a model has already transitioned to flying off the prop rather than the wing. Hmmmmm, now I'm hooked.

Mike

One problem is judging the AOA from moment to moment during 3D maneuvers. Because AOA is relative to motion through the air and we are acostomed to judging motion relative to the ground, we lack the proper reference without instrumentation.

Another problem is that 3D maneuvers involve the vector sum of lift, drag, thrust and weight. When that sum isn't zero the model will accelerate in the direction of the resultant vector. The results are only obvious in the special cases where lift is zero and thrust is equal and opposite to weight plus drag (hovering) or, where lift equals weight and thrust equals drag (horizontal flight at a constant airspeed).

AFSalmon, It's definitely fasinating stuff. Just be careful when using X-foil. It really is great for the linear region of the Cl vs AOA curve, i.e, the non stalled region, but tends to over predict the max Cl, or stall AOA. In my experience, X-foil max Cl is about 10% over what will actually occur. Happy X-foiling!

Wallis Collie

Most 3Dmodels I have seen have aspect ratios of about 4 or 5 and props in the nose. The propwash swirls in a spiral and presents the root end of the wing with turbulent flow at nonuniform angles. The outer end of the wing encounters serious spanwise flow due to the tip vortex. Both of these conditions invalidate the airfoil characteristics measured in wind tunnels or predicted by X-Foil. That doesn't leave much of the wing where the measured or predicted airfoil characteristics are valid.

If you want to be able to use the airfoil data try a pusher configuration with a higher aspect ratio. Then, with a big engine and a light airframe, balance becomes the challenge.