Morgans wrote:

"Stealth Pilot" wrote in message
.
TLAR only gets it correct is the eye is exceptionally practised.
(tlar - that looks about right)
Stealth Pilot
Australia

I do love TLAR, but where does one find figures needed for things like
downforce required by the tail, gust factor loadings, lft distributions for
varios airfoios and configurations, ect?
--
Jim in NC


It's all in the numbers, Jim.

Your teachers always told that math would come in handy someday.

Well, take tail loads?

We start with the wing airfoil performance curves.

One curve plots section lift at any given angle of attack.

One curve plots section drag "

Back in the slide rule days the third curve represents "center of
pressure" location (again, per angle of attack but expressed in percent
chord. i.e.: where the summed center of the pressure field is in
relation to the section chord.

Makes an easy model to visualize what's happening.

Now days, the third curve is the Coefficient of Moment, or the
rotational
force the airfoil generates at that angle of attack.
No as touchy feely, but I gotta admit that the coefficient method is
easier to calculate with.

Next, there is the CG question.

For a pitch stable airplane, the center of lift will be behind the
center of gravity. If you visualize this, the nose falls.

A down load on the tail lifts the nose.
How much down load?
Just enough to bring the nose back level.

Knowing where the CG and CL are physically located we also know the
distance between them (the Arm).

For straight and level flight, we know the lift (equal to weight).

So we can do a little arithmetic and find the pitch moment for our
hypothetical airframe.

(We'll skip the airfoil CM for now, ok?
The CG/CP moment is by far the greater issue of the two.
But in reality ALL moments get included.)

So, take that pitching moment and divide it by the Tail Arm (distance
from CG to elevator?) to find out what the load on the tail will be
(pounds).

It's really fairly simple arithmetic so far.
The biggest surprise is how small the actual control loads are.

Some 10 feet back to the elevator makes a very long Arm.

Ten pounds back here can have more impact on CG location than 100
pounds in the back seat. In fact, it better, or the back seat might
be the way wrong place!

Bounds checking shows that many airfoils have a higher CM at higher
AOA.

I think that implies that at low speeds, tail loads are actually higher?
Why?
More force is needed on the tail to hold the nose up that high.
:^)

At higher speeds the CM is generally lower because the AOA is lower.

Ok, too many blank looks again...

Visualize it this way?

At high AOA the center of pressure is generally forward some.
The air is attached to the front third of the wing (or less?),
so the lift force is transferred to the wing in that area only.
(drag too)

As the AOA comes down (and speed is higher to make same lift) the
center of pressure is "blown" aft. (?)

It just makes an easy mental image to help remember how it all\
fits together.

So a steep CM curve (old style) or a larger range of CM values indicate
an airfoil with an active center of pressure. (also ?)


As for the other specific things you mentioned?

Lift distribution is more of a plan form thing but there are other
considerations as well.

One aspect is the planform shape.
Rectangle (Hershey Bar), Elliptical? Delta

Another aspect is wing twist.
Twisting the tips down makes less lift at the tips.

Take a rectangular wing and wash the tips down a bit and you get can a
nice elliptical lift distribution from a Hershey bar.

Do the same thing to an elliptical plan form and it might not even fly.
(washed the lift right off the back of the wing!)

I'm still working on gust loads...
They are described as so many feet per second of gust
but I have trouble wrapping that up neatly.

My best guess for a reasonable approximation is this...

Do a vector diagram with the airplane's forward speed (in fps)
on the X axis, and the gust vector pointing doen (vertical at xxx fps)
and note the angle of the resultant.

Go back to the airfoil performance data and recalculate how much lift
will be generated at that speed if the AOA suddenly increased by that
angle.

Divide that lift number by the flying weight to get the G load that
would be imposed.

There is a lot more to it, of course.
More than I know for sure.
But it's a start.


Richard