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Posted for "Mike Hutchins" <mhutchins@attglobal.net>:
This is a very interesting discussion about the impact of gross weight
increases on an existing airframe. I would like to add a few points to the
discussion as well as share some old info from the factory's test program on
the LIV landing gear and wing.
I have attached two articles from the factory's newsletter. The pdf file,
LIV Gear, is an article discussing the gear strength of the LIV and the fact
that it was tested to 8.7Gs. What is not specified was the simulated weight
of the aircraft during the test. Was it the original design weight of 2900#
(probably), intermediate weight of 3200#, or the test weight of 2000# as
mentioned in the caption to the pictures in the second attachment, LIV
Landing Gear?
Also relevant to any discussion related to G induced failure of the wing is
the concept of subtracting the wing weight from the gross weight of the
aircraft. After all, the weight of the wing must first be overcome by the
aerodynamic forces acting on the wing before the wing can apply any force to
the fuselage and thus load the wing root. I'll grant you that there must
therefore be some level of stress induced in the spar by the loads necessary
to overcome the wing weight, however these are relatively small. In the
original design, 310# was subtracted from the 2900# gross to arrive at a
fuselage weight of 2590# (one would expect that fuel weight should also be
subtracted from these figures as well, however, in Martin Hollmann's design
analysis presented in his book, "Modern Aircraft Design" fuel weight was not
excluded. I'm not sure why, but my guess is that it was a conservative
assumption for a worst case scenario, i.e.. gross weight but no fuel, thus
the maximum load condition the wing spar would ever see in flight.
Originally, the L-IV was designed to carry 82 gallons of fuel).
Martin reports that the wing was loaded until failure at 9250# (presumably
this was an equivalent load, since only half a wing was tested. There was
no mention of sawing through the spar cap. Since composite structures do not
have a yield point the way metallic structures do, e.g. the ultimate tensile
strength of 2024-T3 at room temperature is 70,000 psi whereas the yield
strength is 50,000 psi, they must be tested to failure.). Thus the wing
failed at an equivalent G load of 8.3 Gs, again using the original fuselage
weight of 2590#. Attached is another picture, "Wing Test", showing Lance and
the L-IV wing loaded with lead shot to simulate the aerodynamic loads on the
wing. The weight is carefully placed according to numerous calculations
applied at each station along the wingspan and chord to accurately reflect
the aerodynamic contribution of the full wing surface. In addition, the wing
is mounted in the test jig to closely simulate its mounting in the aircraft.
Here is a link to Van's aircraft
(http://www.vansaircraft.com/public/rv-10int3.htm) with a picture showing
the placement grid and the final weight distribution for their new RV-10.
So, what does it all mean?
1. It takes a lot to wreck the gear on the L-IV
2. 3145# should be well within the wing's strength envelope, assuming normal
category operation and a safety factor of 2.0
3. 3637# is probably within the wing's strength envelope, although it would
require a less conservative approach, i.e.. including fuel weight in with
wing weight (3145 + 6 x 82 = 3637), and using normal category load limits.
4. 4000# would require all the above assumptions as well as diminishing the
safety margin to 1.6.
I don't have the engineering background necessary to really serve up an
answer beyond the above speculation, but it seems rational to consider 3637#
a safe max gross weight with 82 gallons of fuel and 3805# with 110 gallons
of fuel in the wings. You can always slow down a bit to increase your safety
margin.
I'm out on a limb and out of my element. Saw off the limb if I have ventured
too far from reality.
Best Regards,
Mike Hutchins
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