Lnewqban

So, the discussion boils down to the position of the neutral point along the chord for different airfoils.

I don't know if for laminar airfoils with the max thickness at around 50% of the chord the NP will still be located at 25% of the chord measured from the LE.
It is very possible because the neutral point is at the centroid of the chord-wise lift distribution.

http://repository.lib.ncsu.edu/ir/bi...5063/1/etd.pdf

However, it has been verified that NP is located at 25% for any geometry of airfoils with the thickest point around the 25%:

http://scherrer.pagesperso-orange.fr...tral_point.pdf

Typical symmetrical and cambered airfoils have no influence on the location of the CG of the plane.
That is the reason that allows airplanes with non-symmetrical airfoils fly inverted with no need for readjusting the CG location.

RAPPTOR

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I tried program. Said 8 inches at root. Plans say 7 at root??

BMatthews

8 inches to what? If it's 8 to the NP and the CG on the plans call for 7 inches then it's 1 inch ahead of the NP. That'll be a positive value of some amount for the stability margin.

But at the same time 7 and 8 inches back from the leading edge is a LOT of room. This indicates either a lot of wing sweep or a very big wing chord. If it's a big wing chord then 1 inch isn't very much. You may be getting all concerned about a very slight difference.

In any case as always use the CG shown on the plan. Presumably the designer played with the final location and found that the model flies well with it as shown on the plan.

buzzard bait

I have been using a formula for about 25 years from Gordon Whitehead's book, Scale Models for Everyday Flying. I've also used one from Bill McComb's book. They are algebraically the same. I assume they are also the same as the online calculators. They use wing area, stab area, chord, and tail arm. Airfoil is not a factor.

The formula method has proven to be absolutely reliable for me, for sport and scale models having flat bottom to full symmetrical airfoils. I have used it for every plane I have owned for the last 25 years.

I have found that some kit manufacturers are quite conservative about balance point. I have developed enough confidence with it that I will even make maiden flights with the balance point aft of the one on the plans, if that is what the formula tells me. It works every time. In those cases it saves me the trouble of achieving a forward balance point that is unnecessary.

The very first time I used it, however, it was to check a plane that was wildly unstable in pitch even though I had the balance point exactly according to plan. I added nose weight and got it flying right, and then went home and tried out the formula. I discovered that the plans were wrong and the calculation came out to almost exactly what I had gotten by trial and error.

After that, I use a formula every time. No more errors.

Jim

BMatthews

Yes and no. Again much depends on the angle of attack and the shape of the airfoil.

In the full size world the whole range of "laminar flow" airfoils to produce less drag is based on shapes with a more rearward point of maximum thickness. The idea being to produce a longer pressure ramp on the pressure distribution. And ideally to hold the air longer as it passes over the upper surface.

At model sizes we actually gain for a change on the whole Reynolds number thing. As the Reynolds number falls with smaller wing chords and lower flying speeds the tendency is to maintain laminar flow over rougher and poorer shapes.

But we suffer in other ways. For example the lower the Rn value the lower the maximum angle of attack at which the stall occurs. For really small or really slow models a typical maximum angle of attack is only 6 degrees. Mid size and speed models get up to around 8 degrees and giant scale models likely get to around 10 degrees before stalling.

So back to rearward maximum thickness point airfoils on models.....

The problem with this is that when the laminar flow does break down at higher angles of attack it breaks down fast and badly. And due to the nature of the upper surface shape on the high point rearward airfoils we'll get this breakdown at a lower angle of attack than we see with airfoils with a more forward maximum thickness shape. So in general the full size laminar shapes don't work well at our size and speed range. Instead the more curved shape of airfoils with the max thickness and high point of the camber more forward provide a shape that is more prone to guiding the air smoothly over the shape. So the air holds on longer which delays the stall to a higher angle of attack than you'd get with a max thickness to the rear.

And in fact there are two "rules of thumb" from way back when we didn't have the computer power to design tailor made airfoils so easily. Those being that for a given airfoil that was about average a slightly more rounded leading edge shape would slightly delay the stall to a higher angle of attack. Similarly moving the point of maximum thickness forward a little also tended to delay the stall to a higher angle. But either if taken to extremes resulted in degrading of the performance.

But for a while this was taken to extremes in the fun fly models of the 80's. Those models used massively thick airfoils with very bluntly rounded leading edges combined with super light wing loadings to produce models that were extremely stall resistant. .40 powered 44" span models were capable of doing 6 foot diameter loops when the wide ailerons were mixed with coupled flaps with the elevators.