Rag and tube construction and computer models?

"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


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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

Richard Lamb wrote...

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

SHH! If you tell *all* our secrets
*everybody'll* be doin' it! g

Nice post.

Dave 'mystique is 50%, the rest is algebra' Hyde

That's a keeper...

--
Dan D.



..
"Richard Lamb" wrote in message ...
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

Many thanks for the comments and interesting stories guys....

As usual, the answers range ALL over the spectrum so I am not sure much was
resolved...

But interesting things were told, good points were made, and somebody somewhere
probably learned something....

So, all in all I think it was worth it....

take care

Blll

Coupla extra points. (Not for Richard, for you other guys.)

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!

-------------------------------------------------------

Don't stop there. In fact, don't even start there... not if it's a
tail-dragger. Cuz if you got the little wheel in back and the fan up front,
your worse-case isn't going to be your in-flight tail loads but the kink you'll
put in the fuselage when you're having a bad hair day and try rotating too
soon... or leveling out your flare too late. One reason for the kinks is the
fact the moment for the tail wheel is usually more than for the elevator.

So get a handle on that one first, making sure the fuselage has enough strength
for an occasional bad landing. When you get to the flight loads, odds are
they'll be less than your worse-case landing/take-off loads. (All the better
to appreciate a trike gear, with the engine mount doing double-duty for the
landing gear loads.)

-------------------------------------------------------

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)

-----------------------------------------------------

Second Point: Listen to the man. Or build yourself some practice airfoil
sections, make up a wind tunnel and spend a lot of time watching smoke trails.
Because if you keep the NOSE of your airfoil clean back to at about 25% of the
chord, the remaining 75% of the upper camber can look like cottage cheese and
the silly thing will still fly jus' fine.

NACA figured this out in the 1920's which makes it something of a
head-scratcher to see the Famous Designers of today degrading the main working
portion of their wings with protruding rivet heads. Keep that portion of the
wing clean, you'll see a lower stall and higher cruise. (And if you don't,
I'll give you back the money you paid for this :-)

-R.S.Hoover

PS -- I don't mean to say NACA figured out the cottage cheese. I figured that
one out myself when I was designing my All-Dairy composite... the one with the
bricks of butter for the landing gear.

On Sat, 10 Apr 2004 01:07:18 -0400, "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?

in all the flying I've done it the tailwind I still have no idea what
the tail download figure is.
I think it peaks about 65lb but I'm not sure.

The tailwind has a hybrid aerofoil that has no published data.
(so all that you posted richard is accurate but quite useless )

the trim with 20degrees of flap is full back, with light fuel in
cruise it is full forward, suggesting to me that it may have tail
upload in cruise when light. dunno. this genuinely interests me since
I wanted to do a reverse engineering exercise to see just where the
strengths and weaknesses lie in the design.

I've been stuck on my first question for 2 years now.
"what is the magnitude of the tail download?"

still dont know.
Stealth Pilot

Stealth Pilot wrote:

On Sat, 10 Apr 2004 01:07:18 -0400, "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?

in all the flying I've done it the tailwind I still have no idea what
the tail download figure is.
I think it peaks about 65lb but I'm not sure.

The tailwind has a hybrid aerofoil that has no published data.
(so all that you posted richard is accurate but quite useless )

the trim with 20degrees of flap is full back, with light fuel in
cruise it is full forward, suggesting to me that it may have tail
upload in cruise when light. dunno. this genuinely interests me since
I wanted to do a reverse engineering exercise to see just where the
strengths and weaknesses lie in the design.

I've been stuck on my first question for 2 years now.
"what is the magnitude of the tail download?"

still dont know.
Stealth Pilot


I've always felt the same way about that one.
But before we can analyze the master's work...

Not knowing what the airfoil performance curves look like makes it a
guessier answer. We don't know exactly where the center of lift is
located.

But we can work from you trim description and the known numbers and
come up with a guess.

Wittman's airfoil _is_ a little strange.

From the plans, the CG Range is 15% to 28% of the chord.
Figuring the chord as 48 inches, that means 7.2 inches forward limit,
(aft of leading edge) to 13.44 for the aft limit (also aft of LE).

(That 15% forward limit seems awful far forward to me - but it woiks)

Anyway, the difference between those two points is 6.25 inches.
That is the maximum range of CG locations. We might assume that the
center of pressure is at or forward of the forward limit.

So, working from that...

1425 pounds gross times 6 inches = 8550 inlb moment (worst case).
Tail is about station 144, and the main spar is at station 24.
Sounds like 10 feet (120 inches) to the tail.

So we can divide 8550/120 and get 71 pounds tail load.
(real close to your 65 pound number!)

But different assumptions will give different results.
All depends on how valid our assumptions are at the beginning...

I'm real skeptical that the tail on the Tailwind is producing an
upload at high speed. The airplane handles too well for that to
be the case.

Remember that the leading edge of the stab is probably a little nose
down. Full forward trim _should_ still result in a net down load on
the tail.

Any better, Stealth guy?

Richard

Steve said that. However, he also said that it was done many years after
the plans were first offered for sale. I believe it was about the time he
changed it from the W-8 to the W-10. There were a few tube sizes that were
increased in size then, particularly at the top, front of the cabin to carry
the spar loads. Of course, it was because of the heavier Lycoming engines
being used rather than from failures.

"Cy Galley" wrote in message
news:1uJdc.117$xn4.5040@attbi_s51...
Steve did more than just "eyeball" engineering. He had some contacts at
the
University of Wisconsin that he sent his drawings and parts down to have
them analyzed.

On Sat, 10 Apr 2004 16:06:27 GMT, Richard Lamb
wrote:


hedge trimming

From the plans, the CG Range is 15% to 28% of the chord.
Figuring the chord as 48 inches, that means 7.2 inches forward limit,
(aft of leading edge) to 13.44 for the aft limit (also aft of LE).

(That 15% forward limit seems awful far forward to me - but it woiks)

the actual flight tested values for my aircraft are 10.4" forward
limit and 16.5" aft limit.

1300lb auw (well actually 590kg with an empty weight of 362kg.
empty cg moment arm is 214mm.

hedge trimming

But different assumptions will give different results.
All depends on how valid our assumptions are at the beginning...

I'm real skeptical that the tail on the Tailwind is producing an
upload at high speed. The airplane handles too well for that to
be the case.

Remember that the leading edge of the stab is probably a little nose
down. Full forward trim _should_ still result in a net down load on
the tail.

I tend to agree although I'm sometimes not sure. I'd love to see a
windtunnel test on the "new" aerofoil to see just what the pitching
moment was.
Stealth Pilot