NACA profile calculations

I have got into plotting NACA profiles by computing. Humbly beginning
with the simplest, which I presume to be NACA 4-digit, I found the
Wikipedia page as a prime source of information.

Plotting a symmetrical profile is not that hard: at regular intervals
(=X along the chord), one calculates the thickness (=2*Y) and plots a
dot at X,+Y and another at X,-Y . The formula for the thickness is
netaly given in the wikipedia page, neat!

The first step of assymetrical is just as easy: draw the camber line by
calculating its offset from the chord at regular intervals. From the
camber line. intrados and extrados points are again a vector away; and
the magnitude of the vector is calculated in the same way. What beats me
however is that, if I interpret the wikipedia page correctly, this
vector should be considerd perpendicular to the camber line, where I
would expect it to be vertical, i.e. perpendicular to the chord.

Any comments here?
Is my understanding of the wikipedia page correct?
If so, what's a useable algorithm to plt intrados and extrados?

And also, less poignant but still:
How do I calculate from the basic parameters the behaviour of the
profile, i.e. the Cl curves? And perhaps other nice info like the
leading edge diameter?

Generally: any pointers to in-depth information that a non-engineer can
mentally digest?

TIA,

If so, what's a useable algorithm to plt intrados and extrados?

http://en.wikipedia.org/wiki/NACA_ai...t_NACA_airfoil

dy_c/dx you can get by differentiating between two arbitrarily chosen
points on the camber line. You'll probably need closer spacing of
these points near the leading edge.

How do I calculate from the basic parameters the behaviour of the profile, i.e. the Cl curves?

What are you looking for? Slope will be very close to 2 pi ;-) The
angle of attack of zero lift is -2 f/t (f/t is maximum camber relative
to chord). The cm is more difficult to estimate -- it's easiest to
just plug the airfoil into XFOIL and do the calculation (or do you
prefer
http://www.iag.uni-stuttgart.de/IAG/...lettheorie.pdf
?). The rest, like clmax, are viscous effects and can't simply be
deducted/estimated from the basic parameters.

And perhaps other nice info like the leading edge diameter?

For NACA 4-digit-series, see the Wikipedia page.

Generally: any pointers to in-depth information that a non-engineer can mentally digest?

We would need more information on what you are actually looking for.
It could be structural properties of the airfoil (which XFOIL can
handle as well ;-).

Oliver

jan olieslagers wrote:
/snip/
How do I calculate from the basic parameters the behaviour of the
profile, i.e. the Cl curves? And perhaps other nice info like the
leading edge diameter?

Generally: any pointers to in-depth information that a non-engineer can
mentally digest?

TIA,

I think you would be best served by a 2-D flow visualizer of which there
are several out there. I don't think hand computation of reasonable
values for lift n drag are reasonable objectives.

Brian W

Oliver Arend schreef:
If so, what's a useable algorithm to plt intrados and extrados?

http://en.wikipedia.org/wiki/NACA_ai...t_NACA_airfoil

dy_c/dx you can get by differentiating between two arbitrarily chosen
points on the camber line. You'll probably need closer spacing of
these points near the leading edge.
So weit war ich ja schon...*

How do I calculate from the basic parameters the behaviour of the profile, i.e. the Cl curves?

What are you looking for? Slope will be very close to 2 pi ;-)
**Bis hier bin ich auch noch dabei!

The angle of attack of zero lift is -2 f/t (f/t is maximum camber relative
to chord). The cm is more difficult to estimate -- it's easiest to
just plug the airfoil into XFOIL and do the calculation

XFOIL muss es sein was mir fehlte!***

(or do you prefer
http://www.iag.uni-stuttgart.de/IAG/...lettheorie.pdf
?).

Mm, mal schauen, nur nicht gerade jetzt.****

The rest, like clmax, are viscous effects and can't simply be
deducted/estimated from the basic parameters.

Schade*****

Oliver, besten dank!
KA

* Nothing new there
** This also was more or less mastered before
*** This must be what I was missing
**** Will have a look, not right now though
***** Too bad.

brian whatcott schreef:
jan olieslagers wrote:
/snip/
How do I calculate from the basic parameters the behaviour of the
profile, i.e. the Cl curves? And perhaps other nice info like the
leading edge diameter?

Generally: any pointers to in-depth information that a non-engineer
can mentally digest?

TIA,

I think you would be best served by a 2-D flow visualizer of which there
are several out there. I don't think hand computation of reasonable
values for lift n drag are reasonable objectives.

Hand computing wasn't really my idea. I've some nice machines down here
called computers. In today's world of office productivity it is lightly
overlooked, still there it is: computers are cheap and plenty today -
and computation is what they were originally meant for. I do mean to
apply this mechanical computation to describe and plot airfoils.

As for the flow visualisers: I'll be grateful for a suggestion of one I
can run from the Linux command line.

Thanks again,
KA

On Dec 7, 10:37*am, jan olieslagers
wrote:


As for the flow visualisers: I'll be grateful for a suggestion of one I
can run from the Linux command line.

Thanks again,
KA

http://www.mh-aerotools.de/airfoils/javafoil.htm ?

jan olieslagers wrote:
And also, less poignant but still:
How do I calculate from the basic parameters the behaviour of the
profile, i.e. the Cl curves? And perhaps other nice info like the
leading edge diameter?

Generally: any pointers to in-depth information that a non-engineer can
mentally digest?

Have you checked out some of the links at CFD Online:

http://www.cfd-online.com/

In particular, they have links to various software around the web:

http://www.cfd-online.com/Links/

jan olieslagers wrote:

And also, less poignant but still:
How do I calculate from the basic parameters the behaviour of the
profile, i.e. the Cl curves? And perhaps other nice info like the
leading edge diameter?

Generally: any pointers to in-depth information that a non-engineer can
mentally digest?

TIA,

Jan,

I'd not think the CL and CD curves can be "calculated" from the airfoil shape
very easily. Those, and other behaviors, are usually derived from wind tunnel
testing.

Having said that, I'd not be too surprised to find someone has written a program
that approximates CL/CD from a database of some kind.

Anyway, question...

Are you trying to plot the airfoil shape from a table of ordinates?
Or are you trying to generate a shape from a mathematical algorythm?

jan olieslagers wrote:
/snip/
I think you would be best served by a 2-D flow visualizer of which
there are several out there. I don't think hand computation of
reasonable values for lift n drag are reasonable objectives.

/snip/
As for the flow visualisers: I'll be grateful for a suggestion of one I
can run from the Linux command line.

Thanks again,
KA

I tried an early version of Drela's Xfoil. It is available now in
several suitable flavors he
http://web.mit.edu/drela/Public/web/xfoil/

Can't remember whether these distributions include the data for the many
foils held by U.Illinois. Probably do?

Brian W

cavelamb schreef:
jan olieslagers wrote:

And also, less poignant but still:
How do I calculate from the basic parameters the behaviour of the
profile, i.e. the Cl curves? And perhaps other nice info like the
leading edge diameter?

Generally: any pointers to in-depth information that a non-engineer
can mentally digest?

TIA,

Jan,

I'd not think the CL and CD curves can be "calculated" from the airfoil
shape
very easily. Those, and other behaviors, are usually derived from wind
tunnel
testing.

Having said that, I'd not be too surprised to find someone has written a
program
that approximates CL/CD from a database of some kind.

Anyway, question...

Are you trying to plot the airfoil shape from a table of ordinates?
Or are you trying to generate a shape from a mathematical algorythm?

My idea was to first plot a lot of ordinates through an algoritm, then
plot the airfoil around these. But it seems there is so much good work
already done that I'd be reinventing the wheel. I'll have to study all
the links given here, that'll take some doing to begin with.
But I did have understood the airfoil's behaviour could be calculated
from the parameters that describe its form - wrong assumption, apparently.
Thank you!
KA