inexpensive but good alternatives to graphlite carbon pultrusions?

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wsimpso1

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For our gliders, we make unitary spars that ensure a bulletproof shear connection between the upper and lower caps. Then we bond and tape them into the upper wing skin. On wing closure, we do a blind closure of the lower skin along the leading edge, drag spar, and between the lower skin and the lower spar cap. The leading edge joggle bondline is pretty forgiving, and you can inspect the drag spar bondline, at least the aft edge of it, from the flap cove. You can also inspect the main spar to lower skin bondline at several discrete locations through various access hatches, but it is relatively forgiving of mistakes.

For our bonding paste we use an undiluted epoxy resin system (not the thinned stuff common for laminations), add flox to thicken it and give tensile strength, and add cabocil to make it thixotropic (so it doesn't) sag or run when applied to a vertical surface. However, the cabocil makes the paste hard and brittle, so we generally don't use more than about 1 part cab to 3 parts flox by volume.

I came to similar conclusions. I use flange tranfer techniques for fitting forward spar, aft spar and ribs. My leading edge is closed with a tape on the outside.

Concerns over fumed silica? Wear a mask, and be careful... Otherwise, you can use Gougeon Proset adhesive, which handles like Vasoline, does not sag, shear softens during assembly, has yellow and blue colors so you know it is adequately mixed when you have a uniform green, and sands ok where it squeezes out. Oh, and it post cures like the rest of the Proset line... In exchange, it does cost more than whipping up your own adhesive.

Billski
 

autoreply

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So what's the alternative to blind closures?
Build a proper composite wing. No hard points, spars or other nonsense. Either have a monocoque wing skin that carries all shear, torsion and bending, or use a coreless warren-truss like web that carries the shear and bury the caps in the skins.


It's been done and it's better in almost every conceivable way. Often practical issues (two removable wings) or the difficulties of proving new structural concepts get in the way. Reality I guess.

Looking at more conventional composite wings (spar+skins), they're very easy to inspect. Shove a PVC pipe and a webcam+leds in and inspect the glue joint. Done multiple 3000 hr inspections that way, easy.
 

pictsidhe

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How stiff does the skin need to be anyway? I can calculate aero load deflection, but I'm unsure how much is too much. Buckling in bending is a can of worms to calculate, so I'd probably use an analog computer: test load a section. For hangar rash, test again. I'm thinking 1/2" socket wrench drop from 3-4'. I have no trouble destruction testing stuff first, it's been a while since I Laminated anything so practice would be good.
 

pictsidhe

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Thickness/ local span is something I use a lot. I can do that quickly in my head, unlike bending stresses and deflections.
 

wsimpso1

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Seems like something half between core or no core would be an option.
Such as a sandwich of three laminations with the center lamination more like a thick light fabric. Something cheap, about 1/16" thick with good shear and adhesion to the face skins.

Either you are building with a core or building without a core. It is a binary thing. Now if you want it as stiff in bending as 0.020 aluminum, 0.020 graphite in a symmetric laminate will probably do it.

But if you are going to go to the trouble of having a mold and doing any core at all, why not use enough core to allow you to remove most of the ribs and all of that adhesive attahing everything together?

As to the assertion that you can not go this thin, well, I understand that sailplanes frequently have 2 plies of 4-5 oz graphite on each side of the cores. Given that they survive handling for assembly, disassembly, and road transport, they might indeed be pretty sturdy and still meet your lightness requirements.

Billski
 

BBerson

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I think calling it core or not is just for convenience.
What do you call the center layer of three ply plywood? If all three plies of the plywood are the same wood, then it is a three ply lamination with different orientation of the grain. If the middle ply is a lighter wood then you have "lite ply" plywood and might call the center layer "core" (used by R/C models).
That was my suggestion, find a lighter weave for the center ply of a three ply lamination.
Or, as discussed, it could be just two ply with the inner ply a lighter grid or stringer arrangement to expand the thickness.

I really like the idea of stiffened fiberglass. The cost is about three times less, I think.
But also, the translucent ability to see the ribs is neat. Yes I like ribs and have no desire to eliminate all the ribs. That would require molds. A stiffened skin could use fewer ribs.
 

autoreply

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Either you are building with a core or building without a core. It is a binary thing.
Not necessarily.

The big advantage of composites is that they can be made highly directional. A skin that's fairly floppy chordwise but very stiff spanwise by using solid stiffeners (the earlier discussed cloth over spanwise pultrusions) can be light enough to work.

I think calling it core or not is just for convenience.
What do you call the center layer of three ply plywood? If all three plies of the plywood are the same wood, then it is a three ply lamination with different orientation of the grain. If the middle ply is a lighter wood then you have "lite ply" plywood and might call the center layer "core" (used by R/C models).
That was my suggestion, find a lighter weave for the center ply of a three ply lamination.
Or, as discussed, it could be just two ply with the inner ply a lighter grid or stringer arrangement to expand the thickness.

I really like the idea of stiffened fiberglass. The cost is about three times less, I think.
But also, the translucent ability to see the ribs is neat. Yes I like ribs and have no desire to eliminate all the ribs. That would require molds. A stiffened skin could use fewer ribs.
It's not a binary process (ribs or no ribs). The stiffer the skin, the less ribs you need. No need for molds either.

The earlier discussed skin that's very stiff spanwise can be made flat, put into the proper airfoil shape after curing and then some ribs can be laminated in. Simple foam with some carbon laminated over them on the inside would work.

Why use carbon? The Poisson ratio can be almost zero. Very important if you make a flat panel, then wrap it. A material with a high poisson ratio will warp near sharp bends.
 

BBerson

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I don't know what poisson ratio is, other than my steel book says the poisson ratio for steel is .3.
Can you list the poisson ratio for carbon, fiberglass and aluminum sheet?
I suspect fiberglass is poor in poisson, and that's why we are trying to make fiberglass stiffer using extra thickness or micro stringers.

BJC- I think mat is way too heavy for core. I am thinking about .010- .020" core thickness to make stiffer fiberglass of equal weight of .020" aluminum, which is standard wing skin for a single seater (some use .016").

The Jupiter 1 sport plane (post 28), had two or three ply thick solid fiberglass skins. Perhaps too heavy but it worked. I do like the ease of assembly with translucent fiberglass.

When comparing sandwich, I think it is important to point out that a foam sandwich can't be allowed to buckle because it would crack. Usually the lightest skins are thin and do elastically buckle in service. So you need to decide if your skins are capable of buckling or not or if buckling is desired or not.

Edit: the article from August ’67 Sport Aviation.
 
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pictsidhe

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Poisson, fish.
Oh, that Poisson?
Named after a Frenchman called Mr Fish.
Take a rubber band, stretch it. Notice how it gets thinner? Poisson's ratio is transverse strain to axial strain.
 

lr27

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I thought it was how many times more fish my brother caught than I did (early trauma). ;-)
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As far as a sandwich panel buckling, it depends. If you mean buckling like I've seen on metal wings in normal flight, a thin core might not crack. It depends on just how sharp the bend is when buckled, how thick the core is, and the elastic properties of the core.
 

BBerson

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That's what I was saying. You could design a thin core that could withstand elastic buckling.
But I don't think a thin layer of 4 pound foam is strong enough to avoid disbond. Just my guess.
Maybe denser (stronger) and thin flexible foam would work. But my hunch is something other than foam would be desired. I was thinking about Dynel fabric (synthetic fiber)
Used by Ken Rand on KR-1 prototype. Density is low, I think, but I don't know how good the bond is.
 
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wsimpso1

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That's what I was saying. You could design a thin core that could withstand elastic buckling.
But I don't think a thin layer of 4 pound foam is strong enough to avoid disbond. Just my guess.
Maybe denser (stronger) and thin flexible foam would work. But my hunch is something other than foam would be desired. I was thinking about Dynel fabric (synthetic fiber)
Used by Ken Rand on KR-1 prototype. Density is low, I think, but I don't know how good the bond is.

Let's get beyond guessing. Most fiber-resin systems hit their elastic limits around 1-2% deformation, resins and foams limits are 5% and above. Since they essentially move together under bending and torsion and in-plane shear, neither the resin nor the foam are in any real danger under those loads if the facings stay together.

Buckling prevention is a function of compression/shear loading of the panel vs bending stiffness. It does not take much core thickness to get really large bending stiffnesses. The top para covers the issue - in bending, the facings and core bend together and if the facings are OK, the cores will be OK too.

Now set up out-of-plane shear (load perpendicular to face sheets), remember a couple things:

Each facing has to stand the shear;
The foam can not crush;
The interlaminar shear must not fail the foam nor the bond line.

Sizing the facing to stand the shear is pretty straightforward. What is the airload on the panel (determined by the perimeter of support from spars, ribs, etc, that pick up each panel) divided by the perimeter of the panel? That is your average load per unit length (lb/inch). Then what is the facing thickness? Divide your load per unit length by your thickness and you have the average shear stress in the edges of the panel. You can also work out the maximum loads and stresses. Until you get to really big panels, a panel that can stand being handled and assembled is pretty stout to this.

Foam crush strengths do not care about thickness. If the skins do not misbehave, the shear loads result in bending strains of the sandwich, which are superimposed upon the tensile or compressive strains due to wing bending and in-plane shear from torsion. If the facings fail, all bets are off, the laminate has failed...

Interlaminar shear is a look at how much load is passed between adjacent lamina through the depth of the laminate. This actually designs how far the shear web has to lap onto the spar caps. If we designed our wing panels with only facings separated by cores, we would really have to worry over it in the panels too. Most of the time, we skip the core in the edges of the panels, just glass-to-glass or carbon-to-carbon or do a high density core there. A lot of airplanes done this way.

Now if your wing panels are so big that deflections between closed out edges puts the foam in danger, you need to connect the skins either more strongly (higher density cores) or remove a strip of core in places (get more facing-to-facing bonds).

The books on analysis and design of composites cover all of this, and really are necessary to KNOW that your design is OK.

Billski
 

BBerson

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Do any of the composite books cover solid fiberglass wing skin? Like the Jupiter 1.
My interest is, and I think the topic of this thread is inexpensive alternatives ideas.
Based on my experience of owning a sandwich skin aircraft, which I consider to be overly heavy, as Boku also mentioned, I wouldn't build that way. Weight is the enemy! :)
My Grob is 230 pounds heavier than a very similar metal Vivat motorglider with same engine and same wing span.

I am just looking at ideas for using the carbon pultrusions, which apparently have incredible strength. The problem is I think the pultrusions can't be be used with traditional metal skin.
The pultrusions should be used with fiberglass skin, it seems. Carbon skin is expensive and therefore out of the question.
So my logic leads to simple fiberglass sheet skin. It could be homemade skin.
Even better would be factory made skin.
But this is all new, I haven't ever seen any simple sheet fiberglass skinned aircraft other than Dyke Delta. Which is non-structural skin according to Mr.Dyke, if I remember correctly.

So of course there will be guessing until tests are made. For me this is all just conversation for the future. Most likely I will stick with fabric skin. But who knows.
 
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Vigilant1

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Each facing has to stand the shear;
The foam can not crush;
The interlaminar shear must not fail the foam nor the bond line.
Billski, IIRC, in a recent post you mentioned that you'd used straight epoxy to seal the foam in prep for laminating, rather than using epoxy/micro slurry, and that you'd seen little/no weight increase. Do you suppose there's any reason to be concerned that straight epoxy might not give the same foam/laminate bond strength immediately and/or long term? I don't have any reason to doubt the ability of straight epoxy to do the job, but I also have very limited tech knowledge of these materials.
 

autoreply

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Do any of the composite books cover solid fiberglass wing skin? Like the Jupiter 1.
My interest is, and I think the topic of this thread is inexpensive alternatives ideas.
Based on my experience of owning a sandwich skin aircraft, which I consider to be overly heavy, as Boku also mentioned, I wouldn't build that way. Weight is the enemy! :)
My Grob is 230 pounds heavier than a very similar metal Vivat motorglider with same engine and same wing span.
But that has everything to do with Grob and nothing with sandwich skins. The Grob Astir for example, a later design than the Std Libelle has wings that are almost twice as heavy. Grob build extremely heavy. Done plenty of repairs on Grob ships, quite often the laminate is twice as thick as on otherwise similar ships from other manufacturers.

I am just looking at ideas for using the carbon pultrusions, which apparently have incredible strength. The problem is I think the pultrusions can't be be used with traditional metal skin.
The pultrusions should be used with fiberglass skin, it seems. Carbon skin is expensive and therefore out of the question.
So my logic leads to simple fiberglass sheet skin. It could be homemade skin.
Even better would be factory made skin.
But this is all new, I haven't ever seen any simple sheet fiberglass skinned aircraft other than Dyke Delta. Which is non-structural skin according to Mr.Dyke, if I remember correctly.

So of course there will be guessing until tests are made. For me this is all just conversation for the future. Most likely I will stick with fabric skin. But who knows.
The cost of carbon per square meter (or per lbs) is irrelevant. What's relevant is how much the carbon for an airframe costs. Your Grob is what, 150 sqft of wing area? You might be shocked how little the CFRP for such a wing will cost compared to the weight savings ;)
 

autoreply

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Billski, IIRC, in a recent post you mentioned that you'd used straight epoxy to seal the foam in prep for laminating, rather than using epoxy/micro slurry, and that you'd seen little/no weight increase. Do you suppose there's any reason to be concerned that straight epoxy might not give the same foam/laminate bond strength immediately and/or long term? I don't have any reason to doubt the ability of straight epoxy to do the job, but I also have very limited tech knowledge of these materials.
Straight epoxy will always be stronger than micro+epoxy. Don't have data on the weight difference.
 

Vigilant1

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My Grob is 230 pounds heavier than a very similar metal Vivat motorglider with same engine and same wing span.
Sure, but can that Vivat survive having a tree fall directly on the TE of the wing with no damage?. They'd still be straightening extrusions and ordering new panels. You've got a sturdy bird there!
 

BBerson

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The cost of carbon per square meter (or per lbs) is irrelevant. What's relevant is how much the carbon for an airframe costs. Your Grob is what, 150 sqft of wing area? You might be shocked how little the CFRP for such a wing will cost compared to the weight savings ;)

Irrelevant ?
Aluminum skin 2024 t3 .020" is $2.56 per square foot at Aircraft Spruce.
Carbon sheet .020" is $17.56 per square foot here: http://www.carbonfiberglass.com/Carbon-Fiber-Sheets-Composite-Panels-Fibre
 

BBerson

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Sure, but can that Vivat survive having a tree fall directly on the TE of the wing with no damage?. They'd still be straightening extrusions and ordering new panels. You've got a sturdy bird there!

True, but as I get older I can barely lift the wing. I was seriously considering ways to lighten the wings. Does not need to be tree proof. :gig:
 
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