Strange I started to post and RC Universe said,

"You can't post two consecutive messages within such a short amount of time."

and I said, "Sure I can, I can type very fast so it's no problem!"

It ingnored me totally, no manners.

Anyway here is a really fine Neutral Point Calculator.

http://www.geistware.com/rcmodeling/cg_super_calc.htm

It seems that the formula used to calculate NP doesn’t take into account that the stab operates in disturbed air, so the NP will be further forward than the calculation suggests.

For example, if the tail has an area 20% of the wing, and the tail arm is 2.5 times wing chord, the calculation would put the NP at 67% MAC giving a static margin of 41% if you choose the CG location at 25% of MAC.

I prefer the empirical calculation NP = 25 + 40*Vbar (as a % of MAC) where Vbar is the tail Volume ratio.

Adam,
I worked with the author on this and I want to know do you have a concern with the equation used or do you have an equation that you feel is better?

Hi Geistware,
Your formula for NP = Tail arm * (Stab Area / (Wing Area + Stab Area)) requires that the wing and stab have the same aspect ratio and also the stab should not be operating in the disturbed air from the wing, which is not so usual.
The empirical formula that uses the tail volume ratio Vbar gives a result much closer to the reality.

Depending on the TASK , this ratio can be changed --a lot -
That is the reality of the arrangement -
Example
the ratios for a stable glider is different from that of a extreme aerobat setup.
Also - if the aerobat is required to do "departure " maneuvers - that is tumble - a smaller forward mounted stab works better .
For the record - this is the arrangement used on the CAP type 232 .
The best setup is defined by the task -However
The ratios noted are fine for stable craft.

I guess you are referring to balance point (CG) location, which the above calculator doesn't calculate, since you have to enter your own value as % of MAC.
The calculator is supposed to calculate the Neutral Point (NP) location, which seems to be too far aft, since the resulting Static Margin for a typical CG location is about 40%, which is far from the typical 5% to 15% as suggested in Geistware's website.

-I was looking at how one selected the ideal neutral point -
I did not look at the formula .
On the type stuff we fly - one can look at the arrangement and pretty well guess (fairly accurately) what the character of the model will be .
This is from that humbling teacher -- experience.

I would like to thank you all for the conversation.
I am going to try to get the author involved in this conversation to see if more light can be shed on the formulas. I know nothing about aerodynamics but I am glad that my site generated so much coversation!

??[sm=confused.gif]?? There is no such thing as selecting a neutral point. The neutral point is a fixed characteristic of an airplane of fixed shape. Or were you talking about designing an airplane to have a specific neutral point location?

As for the calculator, it does not seem to be very accurate. I have an excell spreadsheet I could forward that calculates these things quite well, the only problem being it requires one value to be looked up from a chart. THe chart could be scanned and published, alongside the calculator, I guess.

Of course the chart is from a published stability book, so putting it on the internet may not be the way to go.

I think were asking too much from a simple formula for getting a "ballpark" of where NP is located. I did a number of these calculations recently using the results to determine a range for the possible C.G locations. Basically, I consider this range to be the C.G. for a first flight and approximately the limit on how far back the final C.G. could be.

I used the results for figuring 4 things:
1) How long to make the nose of the plane
2) How much flexibility I need for mounting engine and battery
3) Determining the location of some parts (servos, receiver, fuel tank, landing gear etc), I had to move them forward
4) Deciding if I needed to loss weight in the rear of the plane, I did.


If you consider that the calculation only uses areas, you have to conclude it is only a rough estimate. To do this right you'd need to include Cl (lift coefficient) and Cm (pitching moment) of the wing, Cl of the tail, combined with cruise attitude, incidences of the wing and tail, and downwash at the tail. Probably the AeroEngineers can add to this list. This still only determines the NP for straight and level flight. As Dick notes, you may need to adjust your C.G. thinking for the type of flying you do.

Carl

An excellent discussion on this so far. If the forumulas can be fine tuned to provide a more accurate tool we'll all benifit.

A little trick I read of years ago for calculating the neutral point for model rockets was to cut out a side view from cardstock and then balance the profile cutout over a pivot to determine the NP for the rocket. For stability the CG then needed to be so far in front of that. It was suggested in those magazine writeups that this trick could be used for model airplanes as well. Could anyone comment on this?

JohnG, how critical are the numbers from the chart? Would it be pratical to use a small list of "constants" that are picked to work over a range of designs? This would introduce a small (hopefully) error but if the NP could be figured that way to within a few % error it would be good enough for our purposes.

The two figures below show the different results for NP location depending on which formula one uses.
The first figure shows the NP using Geistware's formula, the other one is based on the empirical formula using the tail volume ratio Vbar.
I think the latter is much closer to the reality.

Adam,

I don't think you should call the result of the D = 25 + 40*Vbar equation
the Neutral Point.

It is supposed to be an estimate of a target CG with something like 15% static margin. I haven't checked the formulas, but if you move forward 15% of the MAC on the top drawing you're pretty close to the suggested CG from the Tail Volume Coefficent Equation.

Carl

Siefring,

If you use the NP shown on second drawing as your CG location, then you'll have the CG at 51% of MAC.
And, if you use Geistware's calculator, you'll get a Static Margin of 43.5% for a CG location at 30% of MAC.
Just type in the numbers from a typical plane and you'll see.

following is the complete formula used by Martin Simons to calculate the neutral-point of an aircraft:

NP = h0 + ns Vs ( as/aw) (1 - (de/da ))

where:
NP - neutral-point
h0 - aerodynamic center of the wing, typically 0.25
ns - stabilizer efficiency typcially 0.6 (0.9 for a T-tail)
Vs - stabiliser volume coefficient
as - lift curve slope of stabiliser
aw - lift curve slope of wing
de/da - change in stabiliser downwash angle versus change in wing angle-of-attack, typically 0.5 to 0.33

Mine is somewhat more simplified...

John -I design to get the neutral point I want .
However - we will be flying a new design - using the collective pitch prop very soon and my guess is the we want to be as close as possible to having the model such that it will be shorter coupled and with minimum tail group size - so that when reverse pitch is applied -it will stop and back up for at least a short interval .
I don't think there are any books which cover this - so I will have to cut and try .
I think the reverse appied in a vertical dive or climb will do this best.

Adam,

I see what you mean, but I'm not surprised that the Geistware NP estimate is off.

I made a mistake interpreting "your" equations. I use the formula

CGTarget(As %of MAC) ~ 10+40*TailVol

for a CG estimate with 15% Margin. The 25+40*Vbar version adds the 15% margin back in, so it is an estimate of the NP.


The Neutral Point calculation in the Geistware Site assumes the lift coefficient for each slice of the tail and wing are same. So total lift is proportional to area and is acting at 25% of the MAC for both the tail and wing. The NP can then be calculated by solving for the point where there is no pitching moment. It seems to do the calculation correctly. How accurately this describes reality is debatable. If you add the tail efficiency kluge that is used in the Simons formula, you will get numbers more to your liking. Try an example and reduce the tail area to 0.6*TailArea.

You can then go a step further and reduce the tail volume again by the downwash term (1 - (de/da ))~(1-.33)=.66 then the two estimates are very close indeed. The Geistware version has a (WingArea+TailAera*.6*.66) in the denominator vs. WingArea for the Simons formula. All this assuming as/aw=1.

Carl

I've been working on a calculator that uses the Tail Volume coefficient to find the NP location.
The calculator gives both the wing area and the stab area, so you don't need to calculate it, just measure the plane and type in the figures, no math needed at all.
You also get the CG location by entering the Static Margin you want (typical between 5% and 15%).
You may check it here https://rcplanes.online/cg_calc.htm.

OK Everyone, Version 6.0 of the calculator are up and does balance calculations as well as CG. Please review it here.

http://www.geistware.com/rcmodeling/cg_super_calc.htm

Thanks Geistware. The link and your post has been added to the Calculators and Tools thread for posterity.