Quantcast

Not-so-solid massive core wings: Lightening the core foam

HomeBuiltAirplanes.com

Help Support HomeBuiltAirplanes.com:

wsimpso1

Super Moderator
Staff member
Log Member
Joined
Oct 18, 2003
Messages
7,639
Location
Saline Michigan
But at which point it is applied ? Or shell i apply it to stiffness centroid ? ~ at main spar .
Yes, you apply the torque to the stiffness centroid. But finding it can get fussy in a composite structure.

The fussy part? In a structure all made of the same material, yc = Sum(Ay)/Sum(A). In structures with varying E's, yc = Sum(EAy)/Sum(EA)... In UNI Glass spar caps, E is about 4.4M, while the +/-45 skin is about 1.8M. Want to work with Graphite caps, the disparity doubles. The easy way to look at it is any unidrectional elements look like they are a multiple of their actual areas. Let's bite it off one piece at a time:

Most of our wings have things cut out of the trailing edge - flaps and ailerons - so the wing skins go from about 0.00C to about 0.75c, with the lower skin wrapped back up to the bottom of the top skin as a drag spar and to close the aft edge. Under those circumstances, I have done piecewise calcs of my Riblett 37A315 and found the skin centroid to be 0.38c aft of the leading edge and 0.02c above the chord line (as stated in post 254 above).

Spars typically have caps of unidirectional material and wraps and webs of +/- 45. Symmetric rectangular spars have their centroid about their middle. I prefer symmetric caps and channel shapes, so mine have the neutral axis in the middle of spar, but the fore-aft line will be shifted to the web side. Good first estimate for a symmetric spar will be to just use the visual center of the spar. Some folks find asymmetric spar caps to be appealing, in which case a downtown job on all this will first require finding Sum(EAx)/Sum(EA) in both axes for the spar.

Most of us working in composite wings are working in Laminar Flow foils. Max thickness runs in the .35c - 0.4c range. The Riblett 37A315 has max thickness at about the centroid of the skin, and conveniently, also the place where you could fabricate your lightest spar. Hmm. I suspect that for running basic models to see what is possible, you could just call it 0.375c aft and 0.02c up. If ,however, you are running a more traditional foil with the spar at 0.25c, you will have to do the Sum(EAx)/Sum(EA) to find total centroid. This also requires some spar sizing to get in the ballpark. I have discussed this elsewhere on hba.com. Rule of thumb: size caps to carry bending moment; Size web to carry shear; use your favorite multiplier greater than 1.6 on the web and cap areas; Make sure everything is wrapped in two plies. This will probably do for the current exercise, but for final design a more thorough check of failure criteria, iteration for low total weight, and manufacturing ease are all in order.

Some folks find using an equivalent G to be convenient. J for the wing section with a spar is then Sum(GJ)/Gskin where Gskin is the shear modulus (Q66 in composite parlance) of the skin material and the other G's vary with the material and fiber orientation. After that, yes, tau = Tr/J, and r is how far the point being looked at is from the centroid of the composite system. When you first get started, composites can make you nuts, but after a while it becomes normal for you, and your friends just think you are nuts... Oh, and Excel is your friend in all of these calcs. If you want I will check your math. Send functioning Excel spreadsheet files. We can communicate on PM.

Billski
 
Last edited:

ragflyer

Well-Known Member
Joined
Apr 17, 2007
Messages
308
The 1/70 rule is roughly close (more like 1/60) to what you get for plywood in shear buckling. For aluminum a/b tending to infinity

1. all edges simply supported for shear it is 1/45 and 1/30 for compression.

2. for fixed edges shear is 1/60 and compression 1/40

The general formula is s = k*E*(t/b)^2. k depends on edge conditions and a/b ratio.
hence t/b is proportional the square root of E all other things being equal.
 

ragflyer

Well-Known Member
Joined
Apr 17, 2007
Messages
308
for composites due to their stiffness carrying in all directions it gets a bit more complicated and is not a simple b/t ratio and is a function of D11, D22, D12, and D66... too lazy to type the formula unless someone is really interested...
 

stanislavz

Well-Known Member
HBA Supporter
Joined
Aug 21, 2016
Messages
909
Location
Lt
for composites due to their stiffness carrying in all directions it gets a bit more complicated and is not a simple b/t ratio and is a function of D11, D22, D12, and D66... too lazy to type the formula unless someone is really interested...
I am course plus one or two example to verify fea..
 

wsimpso1

Super Moderator
Staff member
Log Member
Joined
Oct 18, 2003
Messages
7,639
Location
Saline Michigan
I recognize elastic stability and composite notation for load and strain vectors, [Q], [G], [ABBD] matrices. OK, so the rules of thumb are based in E, type of loading, and proportions... Cool.
 

stanislavz

Well-Known Member
HBA Supporter
Joined
Aug 21, 2016
Messages
909
Location
Lt
I recognize elastic stability and composite notation for load and strain vectors, [Q], [G], [ABBD] matrices. OK, so the rules of thumb are based in E, type of loading, and proportions... Cool.
Did you get my mail ?
 

Vigilant1

Well-Known Member
Lifetime Supporter
Joined
Jan 24, 2011
Messages
5,754
Location
US
General comment: Thanks to all who are taking a lot of time and effort to make this a very valuable thread (from its very humble beginnings).
There's lots of great information here in one place, with enough context help make it useful. A bit "stream of consciousness," but that's the nature of a forum, with backhandhforth. I don't know of a for-the-layman source that spells these issues out (I need to check out Stress Without Tears). Strojnik has some, Hiscocks has some, Raymer can help with the aero portions, but structural issues are left to others, etc.
 

stanislavz

Well-Known Member
HBA Supporter
Joined
Aug 21, 2016
Messages
909
Location
Lt
This thread have really nice discussion going. But it will have to be divided.. Mine posts are for foam-less thinking. But - each discussion must have opposition.

Btw, got some samples of cf net

ArtCarbgitter.jpg

Cheap bid, really interested how it show itself. At 170gsm, it will form thicker layout than 450gsm. And no idea how to test it in fea or analitycally.
 

stanislavz

Well-Known Member
HBA Supporter
Joined
Aug 21, 2016
Messages
909
Location
Lt
Some more info for allowable deflection to maintain laminar flow. Found in one old book.

Deflection 0.001 of chord is ok if it is spread at 0.1 of chordvise. Ie - on 1 meter of wing, 1mm bump/hole is ok if it at the middle of plate sized 0.1 m of chord.
 

autoreply

Well-Known Member
Joined
Jul 7, 2009
Messages
10,762
Location
Rotterdam, Netherlands
Some more info for allowable deflection to maintain laminar flow. Found in one old book.

Deflection 0.001 of chord is ok if it is spread at 0.1 of chordvise. Ie - on 1 meter of wing, 1mm bump/hole is ok if it at the middle of plate sized 0.1 m of chord.
That seems a bit higher than I'm willing to believe.

Typical waviness is 0.002" over a 2 inch span, or about 50 microns ;-)
 

stanislavz

Well-Known Member
HBA Supporter
Joined
Aug 21, 2016
Messages
909
Location
Lt
That seems a bit higher than I'm willing to believe.

Typical waviness is 0.002" over a 2 inch span, or about 50 microns ;-)
Do not know. It was from russian book, live data, tested in a wind tunnel. Important factor was no abrup chamges in shape.
 
Top