wrote:


hi richard: could you possibly give me some numerical illustrations?
I am not sure where your numbers are coming from (17.2?, 295?), but
this is probably just my ignorance. obviously, going from 0 to 1 sqft
of wing is a useful weight/lift tradeoff... :-) is your point that the
current wing size (of any plane) is already the optimal weight/lift
tradeoff? strange that better materials over the last 50 years would
not have changed the optimum.

regards,

/iaw

I'll give it a try, but you owe me lunch now!


No, not trying to say that all wings are already optimal.
Especally for all missions.

Just that one must look carefully at
proposed performance gains / weight increase.

Write this down: "It's ALL about WEIGHT"

Material gains over the last 100(!) years have been amazing.
For instance, can you imagine a wooden 747
Think I'm kidding?
Check out this 600 passenger design proposal by Bel Geddess,
presented at the New Your Worlds Fair in 1940.
And it's a SPAN LOADER, too!

http://www.home.earthlink.net/~tp-1/ged.pdf




the mystery numbers?
Nothing mysterious here.
It's just arithmetic.

refer to page 25 of Aerodynamics for Naval Aviators

Anybody have a link to this?
It is an excellent starting point for basic aerodynamics.


Going back to...
L = Cl S q

This is the basic lift equasion (well, one of many forms)
Rearrange that equasion to solve for L, CL, S, or even q.

Cl is the coeffecient of lift that we are looking at.

S is Wing Surface Area is Sq Feet

q (rho) is dynamic pressure in pounds per square foot.
q = 1/2 density * velocity squared
Density in Slugs per cubic ft.
V in feet per second).

using q = (sigma V^2)/295 (V in knots, TAS)

295 is a conversion factor that converts FPS to knots

Greek leter Sigma is used for density ratio.
That's ambient pressure / standard day pressure
At sea level standard day, sigma is 1, and drops out.


L = Cl (Sigma V^2 /295) S

So if sigma is 1 then L = CL V^2 S / 295

then solving for V

V = 17.2 Sqrt(L / Cl Sigma S)

17.2 is ~ the square root of 295

Simplified for stall speed at sea level...

Vstall = 17.2 (WEIGHT / CLmax S) Since Lift = Weight

Stall speed implies Maximum Coeffeceint of Lift for a given airfoil.

We need as much lift as weight for straight and level flight,
So L = W = L (are the same)


Well, it's late and I'm bushed.
I've read that over several times, and now it doesn't make sense
to me either.

We'll do lunch another day, ok?

Richard