Re: spar caps


Allan Farr
 

Im no expert but I read a book by one of the UKs leading fibreglass experts
and from that I understand that overloading fibreglass causes it to be
irreversibly weakened (similar to metal fatigue?). Presumably for a/c that are
regularly overloaded the structural strength safety margin is gradually being
eroded.

Allan F

Q2 not flying

-----Original Message-----
From: Q-LIST@... [mailto:Q-LIST@...] On Behalf Of
johntenhave
Sent: Thursday, February 14, 2008 4:43 PM
To: Q-LIST@...
Subject: [Q-LIST] Re: spar caps



Charlie,

I am delighted that I am not the only one whose alarm bells are
beginning to sound like tinnitus.

That said, Larry has illustrated my previous assertions beautifully.

The problem is that some poor souls still think that this
discontinuous outpouring is justification for the most dangerous
behaviour.

You make the best of points when you highlight the actual strength of
the materials which result from the wet layup process. The safety
margins are eroded before the part is cured.

Gentlemen, please ignore Larry's advice. It is just plain (and plane)
wrong.

I am presently revisiting the calcs to illustrate how close to the
edge the 1000 lb loading is to the limit and the dangers of overload.
I will restate the critical point that irrespective of the canard
strength, it is the weakest link that will fail and the result will be
catastrophic.

The headstone manufacture will not care one jot if it was a wing
failure or a canard failure that proved one's undoing...

John

--- In HYPERLINK "mailto:Q-LIST%40yahoogroups.com"Q-LIST@...,
oneskydog@..-. wrote:

Larry ,

When you stray from the plans the responsibility is awesome as the
"new"
designer you hold lives in your hand. When you fix a damaged part
original
materials must be used to restore the laminate to the design
specifications to
restore just the right amount of strength and stiffness. Unless you
have a note
from Burt do not redesign these planes you are in way over your
head. Your
mumbo jumbo with terms like using stiffness when it should be
strength, using
laminate strength claims made in a catalog as "A"basis allowable's
to do your
back of the envelope section calculations make me nervous. I have
also seen
the quality of your repairs on the flipped Tri-Q and frankly was
alarmed. I
cannot sit silent and let you preach design mods to the builders.

I teach advanced composite manufacturing and processing as well as
repair
but am not qualified to design composite parts. In one class the
students build
a carbon fiber "I" beam from unidirectional pre-impregnated carbon
fiber
tape. They also have to calculate the strength and first ply failure
and predict
the maximum load the beam will take. We then take the beams to the
test lab
and using a really cool 4 point bend fixture we crush the living
**** out of
them. The students who are engineers are always surprised at the
spread
between the predicted and the actual values

For your 30 G claim at 1000 lbs I offer this information. Our "I" beam
design has a +/- 45 degree shear web 0.080 inches thick, caps are
0.130 inches
thick with 19 unidirectional plies for tension/compression-. The
overall beam is
3" high and 3" wide so we are in the range of a Q canard section
somewhere
between the root and tip.

The breaking loads vary from around 5,000 lbs to 9,000 lbs. Quite a
variance
for the same laminate built on hard tooling vacuum bagged and
autoclave
cured. More sobering is that it is not even close to 30 G's if your
working load
is a thousand pounds. You cannot duplicate the quality of these
parts in your
garage or hanger and we haven't even talked about adhesively bonded
joints.

I do not want to be to harsh and you might be a hot commercial
pilot with
what a gazillion hours and a ME degree but are you a practicing
composites
structural designer? Playing composite airplane designer gets
people killed and
maimed. Follow the dammed plans for your structural parts, exercise
your
creativity on tertiary structure and finish.

The above is my humble opinion and I do not wish to offend qualified
people.

Regards

One Sky Dog

Charlie Johnson Materials/Process Engineer
Air Force Research Laboratory / Advanced Composites Laboratory



In a message dated 2/13/2008 9:44:12 A.M. Mountain Standard Time,
larry2@... writes:

So, how stiff (strong) will the result be? We do not have real
numbers, but e glass is about 1/4 the stiffness and strength of
carbon fiber. AS&S lists a 5 inch wide CF as being 54,000 PSI. This
means that a square inch of this material would withstand a
stretching force of 54K lbs. Given the 5 inch width, it would take
about 17 (16.67) layers in the lamination to achieve the stiffness.
On the Q2, there are equal layers of UNI on both the top and bottom
surface spread out over the skin surface. The thickness of the
airfoil at any point adds to the overall stiffness because the
resistance layers of the UNI are forced apart. Think bending moment.
The effective strength of the construction is the UNI layer strength
at a given point times 1/2 the core (foam) thickness. From this info,
anyone can start creating a strength ratio for either of the Q
airfoils. However, as has been pointed out, the center section has an
added hat section for increased strength. The hat section acts more
like a box spar. The UNI "spar caps" are just that. They use the foam
to create the effect of a thick, light-weight spar.





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