Hi All
I have been checking out the new Profili software and am reviewing some aerofoils for a new design.
The main data is understandable except for the graph Cm~alpha.
Can someone please explain the significance of this data to me. I am comparing three different aerofoils with the same RN.

The Cm as I understand it, is the "turning moment" around the C/4 point that the airfoil develops at given angles of attack.

If you have a NACA0008 section at 0* then the moment (tendancy for the wing to rotate downward) will be zero. If you increase the angle to 2* then there will be lift generated but the airfoil will try to twist downward.

The effect of this on the overall scheme of things is, as I understand it, quite small for most airfoils in the model world.

My own experience is with ff glider using high camber (5-7%) airfoils. Part of the set-up in those particular models is the down moment of the airfoil which is comparatively large and is one of the factors in allowing the CG to be well back (like 65%-85% C is not uncommon). BUT that is a very specialised use and not one to aim for with other aircraft.

Hi Probligo
Thanks for your reply. The Cm on all three aerofoils is negative at zero alpha reducing to zero at max angle. I do not know the significance of this. Is it stable or not.

At this point I think I know as much as you.

HELP!!!

The amount of Cm indicates how large a tail volume you need. That's not how fat the rear of the fuselage is but rather how long the tail is compared to how large the stab is.

For example a flying wing airfoil like the MH110 will show a positive curve indicating that it's stable without any other tail. Low pitching moment values will tell you that you can get away with less than normal. Symetrical wings for example hardly need any tail. Especially if the balance is anywhere near the 25% chord point. Of course the amount of tail will also determine the response to control inputs so that's why you see lots of models with very large stabs and elevators even with a symetrical airfoil.

BTW. If you play with a little test wing all by itself you'll see the outcome of the negative moments. Build a couple of 3 x 12 x 1/16 sheet balsa wings. One curved with ribs to a conventional airfoil arc of about 1/8 inch and the other with a forward high point of the same 1/8 inch but at the 15% mark and with a 1/16 reflex curve at the rear 1 inch mark. Try gliding them and they will both flip up (usually) or down (sedom but it depends on how you launch them) and watch them flip like a dervish. Now start adding nose weight to get the balance to near the 25% mark. At some point the arced wing will tuck violently and nothing you do will make it fly stably. Removing some weight helps but it goes from less violent tuck under to the upwards back flipping in the blink of an eye.

The reflexed wing will respond to nose weight by starting to act almost normal with a stalling flight transitioning to a normal flight as you move the balance forwars and finally to a controlled dive.

All because the reflex makes the section have a positive Cm rather than a negative one.

I highly recommend Profili, it is a great program for the small cost.

I agree (not that it is in any way necessary :-) that for flying wings a section Cm is a really big thing. For airplanes with tails it is a small factor by comparison to what the tail or elevator inputs and so can be ignored in the analysis for the most part.

The reason that the CGs can be ran so far aft on a lot of planes is that the Neutral Point is also way aft because of the tail volume thing. I have seen free flights with the CG aft of the trailing edge of the wing. The longer the moment arm and bigger the tail the futher aft the Neutral point is. The Neutral Point doesn't change with varying camber but the tail deflection or incidence angle of the stabilizer to trim the airplane will change with camber.

Take for example a glider that is flying with a given setup. It has a certain stability determined by the CG location. Assume it has a tall T tail to get rid of downwash effects. Lowering flaps will cause an increase in lift and a negative moment (nosedown). The stability of the airplane has not changed but the tail will have to be changed to a more negaive setting to balance the moments.

Hi All
Thanks for all of your input. Bruce I think I now have the gist of it.

Can someone confirm the equation utilization that this Cm value plugs into? Is it analagous to lift equation:
L= 1/2 * rho * V^2 * S * Cl ?

In other words, is it M = 1/2 * rho * V^2 * S * Cm ?
If you use the same units for viscosity, velocity & area in which L results in lbs, does M result in a (foot * lbs) moment value?