Composite wing spar web design - composite webs, foam webs, etc.

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BrianW

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I believe you are talking about slurrying the foam. It is customary on open wet layups of foam to fill the surface with a "slurry" - equal volumes of resin and microballoons - and squeegee off as much as you can before applying glass. This has been done to make the layer of resin in the broken cells of the foam lighter than if it were straight resin.

As to paste, the terminology is usually "dry micro" - aprox 4 parts micro per one part resin. This used to fill nail holes, dents, etc prior to laying glass.

Billski
I found Bill's answer both responsive and practical as to the proportions, and reasons for using a slurry - it adds lightness.
I imagine that either method provides adequate adhesion to an overlay then?

Revealing the depth of my innocent-bystander status - I also recall that some foams are dissolved in some resins - and "blue construction/insulation foam" was being advocated a while ago?
 

BrianW

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/.../AeroComp/.../ planes were CompMonster, CompAir 4, 6, 7, 10. All their wings were made with 1/2” foam board wrapped in cloth and bonded with Vinyl-ester resin. They made a spar about an inch thick. They had two wings, a tapered and a symmetrical. They ran up to 650hp turbines on those planes and never had a spar failure. Too bad the rest of the plane worked as well.
Another practical informative post. The terminal sentence is tantalizing. I guess it meant to signify the opposite? And what was it that did NOT work well? <g>
 

BrianW

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This guy has a good site showing how the shear web that connects the spar caps are built in a Long-ez. Chapter 19 is the wing
These structures are so quick and easy to build you could make a load-test item pretty quick
Magnificent photo-journal from aryjglantz pointed to by Kent. Thank you! (Look = blue foam - slurry, and the glass/resin mix whose name I cannot recall! )
 

cblink.007

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Better to get some basic ideas here on this forum, then consult engineers who are in the know. Not that those of us here offering input lack experience and credentials, but we all know about what is said on the internet. So, reaching out to the engineering community simply keeps us all honest!

I will be the first to admit that while I know aero, my structures know-how is seriously wanting, so I turn my questions to those who can answer them.

This being said, our 25% scale subscale test bird is all foam-glass-epoxy, but in order to make the structure to "act" like the full scale variant in particular dynamic environments, with the help of a structures-savvy friend, we had to make some changes to the internal layout. We know we won't be able to correlate all the data the instrumentation will give us, but it will give us the confidence to move forward if all goes well! This thing is plenty stiff!!!

Can't wait to finish and put her through her paces!!

20200503_145527.jpg
20191227_123728.jpg
 
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proppastie

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never had a spar failure
When I see that statement here on an experimental forum I have to wonder....if they did would you know it?.....FAA cares very little about experimental and Part 103 vs Certified Aircraft, where one report from a mechanic can ground a whole fleet.
 

tdfsks

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so .1%=1000 microstrain? 1/1000 = .001 makes sense,....thank you.....was wondering what 4000 micro strain meant.
Apologies I should have clarified this in my original post:

Strain = Extension / Length (non dimensional)
Strain (%) = Strain x 100

1 microstrain = 1 x 10-6 (usually denoted by a Greek mu and epsilon but cannot do Greek font here)

1000 microstrain = 1 x 10-3 = 0.001 = 0.1% (as pointed out above).

Now with some real numbers ......

As an example take a uni glass epoxy spar cap with the following properties:

Tensile Strength (FTU) = 180,000 psi
Tensile Modulus (E) = 7,000,000 psi
Compressive Strength = 130,000 psi

Now, E = stress / strain, so strain = stress / E

Tensile strain at failure = 180,000/7,000,000 = 0.025714 or 25714 microstrain

So the 4000 microstrain at limit load that I suggested in my original post has a margin of 25714 / 4000 = 6.4 against rupture which is far more than a typical metallic structure which would have a minimum margin of about 1.5 between limit and ultimate loads. This margin allows for some of the other effects I discussed in the original post.

The stress corresponding to 4000 microstrain is stress = E x Strain = 7,000,000 x 4000 x 10^6 = 28,000 psi or 193 MPa.

You will find that this is pretty typical of the stress levels in spars that are designed by engineers who know what they are doing. Compressive strains will be lower (around 3000 microstrain) so the compression cap would be larger to move the neutral axis toward that cap to ensure strains are lower on the compression side than the tensile side) in the critical case (positive bending). This is really no different to metal spars. You do not design a metal spar to reach yield stress at limit loads. Stresses are held lower to ensure good fatigue performance and to allow for some margin for corrosion / repairs etc.

Stresses could be used in lieu of strains but over the years, microstrain has become the most common way of expressing the allowable load levels, at least in the projects I have worked and papers I have read.

The above numbers are for a real material but for illustration only. Please don't use them to design anything. The properties for other styles of reinforcement, other resins and different layup processes will vary considerably.

One final point that should be made here: Many of you are sitting there wondering why this is so difficult. This is because with composites, the airframe manufacturer is also the material manufacturer. You are not buying an off the shelf sheet of metal manufacturerd to QQ-A-250/11 or MIL-S-18729 with known properties. You are buying constituent materials (resin, glass, carbon) who's individual properties, whilst important, are not used in design. You are then developing your own manufacturing process to turn these into a laminate and you are responsible for the resulting properties. There are many possible combinations of materials and resin content varies, cure conditions vary etc ... there are a multitude of variables. That is why there are no handbook values of laminate allowables and why each manufacturer has to figure out their own laminate design allowables that result from their unique manufacturing process.
 
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stansusman

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Sorry, I'm catching this late and someone has likely said this already, I would look at a long EZ structure it's simple ( low tech) re;iable scale-able but its mostly a great basic educational tool, plans are all over
 

proppastie

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This margin allows for some of the other effects I discussed in the original post.
Do you have experience with room temp cures, hand layup, with or without vacuum bagging, and might you comment on the strength difference vs typical autoclave cured aerospace industry parts.
 

Vigilant1

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In my non-expert opinion: considering the inherent variability in the finished materials when built by an amateur at home AND the difficulty in chatacterizing the in-part stresses to be expected, it seems to me that there"s a good case to be made for testing critical parts to destruction. That gives 4 big pieces of info:
- Is it strong enough?
- What are the indicators (if any) of impending failure?
- How/where does it fail?
- What is the true factor-of-safety (FOF) of this part as built? (Maybe it can be made lighter? Maybe it is too close to the service requirement)

Carbon is getting cheaper, spending a little bit of money and time to "know for sure" on critical parts is probably a good investment for an amateur builder of a new design. Yes, do all the estimates you can practically do, then build the piece, strap it into an appropriate test rig, load it in a realistic way, and carefully record the results.
 
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wsimpso1

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I found Bill's answer both responsive and practical as to the proportions, and reasons for using a slurry - it adds lightness.
I imagine that either method provides adequate adhesion to an overlay then?

Revealing the depth of my innocent-bystander status - I also recall that some foams are dissolved in some resins - and "blue construction/insulation foam" was being advocated a while ago?
Thanks for the compliment.

Yes, you do not need to lighten the fill of open foam cells with micro, but micro is also less expensive than epoxy resin, so I appeal to efficiency as well. Let us also point out the reason for squeegeeing off as much as possible. Both micro and air in the glass-resin lamination is bad for strength. In open wet layups, much of good shop practice is removing air bubbles from the laminate. Well, micro is also air bubbles, and it too does evil things to the strength and fatigue resistance of our structures. So, wet the whole foam surface with slurry, work it around until all of the little holes are filled, then squeegee off as much as you can.

The full topic of open wet lamination, hot wire cutting cores, reinforcements, and so on are covered nicely in:

Yes, some foams and resin systems are not compatible.

Styrene foams can both be safely hotwire cut and laminated using epoxy resins. For airfoil shaped parts, we almost exclusively use Dow flotation billet, as we know it contains no silicones or other "poisons", is light and produces accurate parts when hot wired. Go near styrene foam with vinylester resins or polyester resins or gasoline or many other solvents, and the foam collapses. Epoxy only on Blue foam.

No other commercial foamed plastic both hotwires safely and can stand structural resins. For instance, polyurethane foams make highly toxic gases when hotwired, so we shape it with templates and saws, knives, sanding sticks, and it will stand epoxy and vinylester resins.

Oh, and no polyester resin in airplane structures. Tooling? Yes, but not in airplane parts. This includes body filler too. Just don't. Polyester is weak and shrinks forever.

Composite airplanes should be epoxy and in some cases vinylester.

Clear?

Billski
 
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wsimpso1

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So if I'm using a unidirectional material and it is rated for X PSI in compression, I just calculate the compression stress in the cap and size it appropriately ? That is it ?
No, it is more complicated than that. A review of compression crippling in Bruhn will go a long way here. In composites, we usually avoid wide thin caps, instead making them narrower and fairly thick to get to area. Yes, bonding to a substantial web and skins will help some with preventing this crippling or buckling, but avoiding really thin really wide sections is good practive

How about the web, if I'm sandwiching a piece of foam with a composite material ? How do the properties of the foam play into it ? What is the difference between a web of thickness T, versus a web with 2 thicknesses of 1/2T, separated by a bit of foam ? Seems to me it is way stronger, but I know the foam is compressible in X, Y and Z and will sheer between the layers easily compared to if the web was just thickness T.
Design a web by the classic methods, and it is a skinny thing. You look at it and can just SEE it wrinkling up. As I said, put in a core, split the web and put some on one side, some on the other, and be happy. Also, in order to get the web connected to the caps, the web material has to wrap onto the wide surface of each cap. This does several things: The spar will stay in one piece; The web gets curved at the change from web to wrap around the cap, giving it more buckling resistance, and; The web supports the caps against their own buckling/crippling.

Now all of this does not add much strength in carrying shear, that still has to be done by designing it to simultaneously stand the shear load and stand the extension/compression of the caps. As I keep mentioning, the light way to strength involves beefed caps. If you just size the caps to carry compression and size the web to carry the shear, the shear web will fail way early. Search out the lightest combination that makes strength, and the caps are beefed up a significant amount while the web is beefed up some too. All of that with a FOS of 2.0 minimum and you can put fails out of reach.

Billski
 

Hot Wings

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Contrary to the urban myth that composites do not suffer from fatigue, they can fatigue and the use of these low allowables is important to ensuring a long and trouble free life (lot of papers in the sailplane community on this)
I'd be interested in learning more about this topic. I'd start a new thread if anyone else was also interested.
From my very limited reading it appears that composites have a knee in the SN curve similar to steel?
Is this due to the fracture of the resin, micro fractures of the reinforcement, or a combination of both?
Has there been any research into using self healing polymers in composites to improve fatigue life?
 

wsimpso1

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This is an older video...but the techniques check out!!
For you novices, some warnings. Epoxy typically carries some sensitizing agents, that can make you allergic to your airplane. Yes, some more recent epoxies are lower sensitizing, but they are not non-sensitizing, while others are even worse than the epoxies of old...

The guy leading you through the processes wore "invisible gloves", a barrier material, on his hands to allow what looked like bare hand work with the epoxy. Many folks have used it, others are just religious about nitrile gloves. Either way, protection against absorption of epoxy through the skin is a wise precaution.

Next, some small amount of the epoxy will evaporate, and usually some amount is also a sensitizer. Folks, either really good ventilation or a respirator with filters for organic vapor capture is a good idea, or you can again become sensitized (allergic to) your airplane.

There are airplane projects out there that were partially built,and then abandoned, sold, given away, etc, because the builder became allergic to it. Do not be that guy. Please protect yourself from breathing epoxy vapors or absorbing epoxy through your skin.

Billski
 

BrianW

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Thanks for the compliment.

Yes, you do [not?] need to lighten the fill of open foam cells with micro, but micro is also less expensive than epoxy resin, so I appeal to efficiency as well. Let us also point out the reason for squeegeeing off as much as possible. Both micro in [ s/b and?] the glass-resin lamination is bad for strength. In open wet layups, much of good shop practice is removing air bubbles from the laminate. Well, micro is also air bubbles, and it too does evil things to the strength and fatigue resistance of our structures. So, wet the whole foam surface with slurry, work it around until all of the little holes are filled, then squeegee off as much as you can.

The full topic of open wet lamination, hot wire cutting cores, reinforcements, and so on are covered nicely in:

Yes, some foams and resin systems are not compatible.

Styrene foams can both be safely hotwire cut and laminated using epoxy resins. For airfoil shaped parts, we almost exclusively use Dow flotation billet, as we know it contains no silicones or other "poisons", is light and produces accurate parts when hot wired. Go near styrene foam with vinylester resins or polyester resins or gasoline or many other solvents, and the foam collapses. Epoxy only on Blue foam.

No other commercial foamed plastic both hotwires safely and can stand structural resins. For instance, polyurethane foams make highly toxic gases when hotwired, so we shape it with templates and saws, knives, sanding sticks, and it will stand epoxy and vinylester resins.

Oh, and no polyester resin in airplane structures. Tooling? Yes, but not in airplane parts. This includes body filler too. Just don't. Polyester is weak and shrinks forever.

Composite airplanes should be epoxy and in some cases vinylester.

Clear?

Billski
Clear! [except two minor typos marked in square brackets thus] I ordered Moldless Composites: cost $18~ good: shipping $12 ~ not so good. Thanks
Brian
 
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