I need a good CL equation 6

[2288]

Can some one tell me where to find a good CL equation???

 

[abuhanifah]

You must be specific, CL for what,is it for only an airfoil, wing or wing body (fuselage and wing).

Generally, L=(1/2)* air density* (Speed)^2*wing area*CL
work out the CL = 2*L / air density*Speed^2*wing area

 

[2288]

thank you for responding but i need to find one (cl equation) that does not need (L) i have heard there is one so if you can help me any more i would thank you... sorry i did not say that in my first one

 

[pja]

You are probably thinking about Cl_alpha which
is just the slope of the Cl vs alpha plot..classical
2d linear potential flow theory give this as 2*pi for a 2d airfoil..for a 3d wing you have to modify it somewhat
to take into account Aspect ratio and sweep. R.T Jones
gave a formula which went like 2*pi/(p*AR+2) where p is the ratio of the wing semi-perimeter to wing span
and AR is the aspect ratio. This slightly overestimates
the lift curve slope for AR>2 .

 

[sanab123]

Dear 2288,
it will b better if u specify what all variables u have the values with u. And what d u want to calculate ?? That will help in giving u the xact solution.

Sanab.

 

[2288]

thank you sanab... i am trying to find lift for the wings i build... i have caculators and comp. programs that give me CL but i am looking for an equation so i can do it my self one of my caculators does not ask for lift so i tryed to make an equation based on what it used (weight alt. speed and wing area... i made one and it worked up until a point... so i am now looking for one that has already been proven...

-2288

 

[knobhead]

Remember in level flight that weight=lift...? Excessive accuaracy is a sign of poor breeding. -Socrates.

 

[2288]

thank you,

i can not belive i did not see that i under stand now thank you very much knobhead... i would still like to get as much as i can on this subject...

-2288

 

[MechAero]

Hello Guys:

I have further suggestions to our friend.
If the wing is already built and you know the profiles you used (that's the shape of the ribs of the wing), then with patient you could try to calculate the aerodynamic properties of the wing (Lift Coefficient slope, Roll Speeed Coefficient slope, etc.) using Multhopp's method.
Have you tried it before?

Regards,

Cristian

 

[2288]

Thank you-

i have not heard of Multhopp's Method before but i will try to find more about it online or in my books... if you have any tips about where to find more i would appreciate them very much...

-2288

 

[MechAero]

Well, I remember it is described in an old NACA paper.
Try to find it in the NACA Technical Report Server.

Let me know if you can't get it.

Regards,

Cristian

 

[2288]

i found a lot on that web site thank you... i will need do do some more searching on it thow because Multhopp's method pulls up a lot of pages... but i am sure i will find some-thing thank you

-2288

 

[2288]

does any one have a formula for the multhopp method?

-2288

 

[macoovacany]

If you want a quick method that gives reasonable accuracy then I would probably recommend doing as PJA has suggested. (Although PJA prefers R.T Jones, Iam going to go back to Prandtl).

L = (1/2) * rho * V^2 * S *CL

CL = dCL/dalpha * (alpha - alpha0)

alpha0 == Zero lift line for an aerofoil section (wings with varying sections or with twist will make this different. Look at the numbers and guess. It'll be closer than you think.)

dCl / dalpha == 2*pi *(AR/(AR+2))

accurate to within 5%.

More accuracy requires more sophistcation. Unfortunately, it isn't linear.

daccuracy / dcomplexity > 0
d^2accuracy / dcomplexity^2 > 0 ( allitle bit more accurate means a lot more complex.)

Alternate formula is =

dCl / dalpha == a2d / (1+(a2d/ (pi*AR)))

a2d is lift curve slope for 2d section. (If you let it equal the theoretical 2*pi value then
(AR/(AR+2)) appears.)

Not for swept wings.

Regards,
Timbo

 

[2288]

you have written "d" in front of some words, Cl, alpha, accurcy, so on and so forth... what are they?

 

[macoovacany]

dy / dx means rate of change of y with respect to x.

(Ignore "daccuracy / dcomplexity > 0
d^2accuracy / dcomplexity^2 > 0 ( allitle bit more accurate means a lot more complex.)" It was my attempt at humour.)

Take out a sheet of paper and draw a set of axes on it. Label the horizontal axis x and the vertical y. Now draw an arbitary curve. The closer the curve is to being parrallel with the vertical axis the larger the value of dy/dx.

Note that the value of dy/dx can change along the curve.

The best way to think about dy/dx is the amount y (the vertical value) will change by changing x a little bit.

A steep line means there is a large change of y with respect to x. A shallow line means that there a small change of y with respect to x.

I suppose now might be a good time to ask about your background.


Timbo

 

[2288]

thank you for that...

you wanted to know my background so... i am not a true aeronautical i am going to become one (i hope you can tell that i am not out of college yet) i build things by myself right now and i needed some equations for lift CL and so on... because i want to know them in my head and not depend on a computer (if you recall i said i have several computer programs)... so i found this site and asked you guys, and as you can see i do have a lot of questions but i know a lot and wish to expand what i know...

 

[2288]

and this is not home work

 

[macoovacany]

Alpha is in radians, not degrees. 1 radian ~ 57.3 degrees.

Timbo

 

[dan100]

CL is just wing loading divided by Q (dynamic pressure)in coherent measures. Once you learn this it all becomes so much easier to understand....
Dan100