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

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Riggerrob

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Wel, only if you have tube fuselage fittings .... Most use some sort of channel, I beam, box spar, etc, with the bending strength tailored along the spar to the expected bending moment. And they use bolts to make the connections.
Do those bolts carry all the flight loads?
Some of my sketches include large diameter bolt holes where spars pass through fuselage side walls and bolts transfer all flight loads to fuselage walls/bulkheads.
The next question is: what diameter bolts do you need?
 

wsimpso1

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While I'm deviating a bit - still within the large monolithic block discussion. Yup EPS. Coffee cup foam - as long as it IS fully sealed and away from fuel/UV/rodents...

If Walt Mooney was willing to run with it in a certified aircraft, it deserves a second look. (and I was wrong in the end they were talking foaming first then prepregs). As long as it meets engineering requirements - With the ongoing discussions on composites of extra layers for ground handling - getting a lighter monolithic block + quicker build still seems a fairly good option.

From Rohr 2-175

The wing, which was the airplane's primary structure, would be built in clamshell molds. First, foam "sugar" would be placed in the mold. Live steam would then be injected to expand the polystyrene beads to the mold contour. The mold would be opened, the core temporarily withdrawn, and "B" stage fiberglass prepreg would be laid up on the inside surfaces of the mold. The core would then be reinserted and the mold would be closed, compressing the core slightly and ensuring even pressure on the laminate. Heat would then be applied in the usual manner to cure the assembly. The use of prepreg would avoid the weight gain problem of a wet layup and would prevent the laminate from becoming resin-starved. The process as a whole is simple, repeatable and cheap.
OK, I shall bite. How sturdy does the clamshell mold have to be to do this? I seem to remember the EPS block molds at Olsonite (1975, so it is kind of far back in my memory) were made of beefy aluminum plate and lots of plate ribs, etc. Did not look like anything an amateur airplane builder would make two of to make one airplane.
 

wsimpso1

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Okay, another way to remove "lazy" foam (assuming it exists) from a monolithic XPS core. In this case the foam is removed in chordwise chunks leaving XPS ribs in the "normal" chordwise orientation. Like a solid XPS core, the composite skin would be 100% supported by foam. (This could use pictures, and is probably harder to describe than to actually do):
1) Cut solid core to desired airfoil shape
2) Using a hotwire, split the core from LE to TE
3) Mark the future location of the spar on the inside of the split face of the top and bottom core half. Also mark the future rib locations (say, every 6 " on center)
4) Use a foam cutting hoop to remove the foam between the ribs. You'll use a guide/template to assure the hoop goes the right depth into the foam leaving an adequate foam supporting layer under the skin.
a) Foam cutting hoop: A modified soldering >gun<. Described here by Pops and others. Make a squared-off hoop out of suitable rigid wire. One dimension of the hoop is the same as the open space between the ribs, so 5"in this case if our ribs will be 1" thick and they are 6" on center. The "depth" of the hoop is the deepest cavity you'll need to make, leaving the desired foam thickness under the foam. The foam cutting gun/hoop has a flat foot/sled so it can ride on supports while in use and have the hoop hang down.
b) Make a tapered template of particleboard/XPS etc for the foam cutting gun/hoop to ride on. It is placed on the inside face of the "split" core surface to guide the foam cutting gun/hoop along the correct line for removal of foam in each rib bay. The main purpose of the template is to guide the cutting depth of the hot loop tool. The small end of the tapered template is at the spar side (so the cutting hoop goes in deep at this thick part of the wing) and as the cuttting hoop goes toward the trailing edge the template raises it up to keep it the desired distance from the skin (1"? 3/4"?). So,the cutting hoop will exit the foam some distance from the back of the core, to allow for the desired foam thickness under the skin.
c) Cut out all the rib bays aft of the spar. Then, use another appropriate template to (with the proper taper) to remove the foam from the bays forward of the spar (if desired).
5) While the top and bottom core halves are apart, make any desired small spanwise cutouts for
wires, pushrods, cables, etc.
6) Re-glue the top and bottom core halves (all ribs and the place where the wire left the skin at the LE and TE). Assure everything is still aligned and no twisting has occurred due to internal foam stresses (consider putting everything back inside the original female top and bottom halves of the foam blank and apply weight/pressure while the glue cures).
7) Next, use a hotwire to cut out the locations for the spar caps and to split the cores top to bottom at the spar web. Build the spar webs and cap on the foam core.

It is more trouble to do things this way than to cut out spanwise channels as described in the OP.
Pros:
- Like the OP method (spanwise cavities) and a regular solid core wing, the outer laminate skin is bonded >everywhere< to XPS foam for support against buckling.
- Faster construction than cutting, aligning, and attaching "a thousand ribs" to a spar. Less adhesive weight, and all the foam junctions will have a filleted edge (from the cutting hoop) rather than a 90 degree junction with adhesive bond.
- ?? Better support against skin buckling/foam failure than the spanwise ribs-and-voids in the OP??

Con:
- Takes longer to remove the foam in this fashion than to remove spanwise chunks with a simple hotwire. Still, once the taper of the template is figured out and the foam cutting hoop is made, I'd think the cutting would go quickly.
- Unlike the method in the OP, there's no easy way to leave the to-be-removed foam in place to support the skin temporarily during vacuum bagging. This might be a fatal flaw.
I like the spanwise scheme better. Sounds way more able to be manufactured cleanly in a home shop. Hotwire the perimeter and a small circle out of the center of each core to be removed later. Yes, it is done all the time on Long-EZ and derivatives for rudder cable and tip wiring, for control surface definition, etc. Laminate the spar and skins , then hotwire the core sections to be removed. Oh, and once you have them in your hands, you may ask yourself why you bothered...

Billski
 

wsimpso1

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Do those bolts carry all the flight loads?
Some of my sketches include large diameter bolt holes where spars pass through fuselage side walls and bolts transfer all flight loads to fuselage walls/bulkheads.
The next question is: what diameter bolts do you need?
Yes, they do. The trick is both having bolts with enough strength and making the hard points they pass through to have enough strength too, and then to do all of that at somewhere close to minimum weight for the job. There are calculations to allow folks to do such stuff. A summary is presented in Shigley's Mechanical Engineering Design. In my Third Edition it is Chapter 6. I have friends who came to the US in the 1980's to go to graduate school, and they discovered that the Chinese ME Design text was a shamlessly copied bootleg translation of Shigley. I know some talented German engineers who have the book on their shelf from college. Really useful book ...

Big enough ones to do the job. That is not a smart aleck comment - you would be surprised how many folks just do not understand that an engineering degree prepares one not to guess, but to make good estimates of what loads are present and then enable you to deal with them.

You might be surprised how small some can be. Go to any maintenance hanger with the covers off of a strut braced Cessna wing root. IIRC, you will see AN4's & 5's in double shear, one each on 152's and 172's. I think the huge knuckle joints in the root of B-17's use two AN8's per side, but they are in about a dozen times shear, maybe more. All that stuff is calculated by people trained in engineering and usually checked by someone else also equally trained. My bird has two AN8's in shear at BL 50 and another AN8 at BL18 or so. They have a generous FOS, as all fasteners should.

Billski
 
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wsimpso1

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How thick is the "kerf" created by a hot-wire?
How does it compare to the thickness of a composite wing skin?

I ask these questions because I am considering using "scrap" foam as female molds while vacuum-infusing wings. By "scrap" I mean the outside of the block of foam that you just carved/hot-wired the wing panel out of. As soon as the airplane is complete, those odd-shaped foam blocks go to the garbage.
The primary tool will still be a large, flat and stiff table.
I doubt if these female molds will be smooth enough for final surface. Instead, my concept is to use "scrap" female foam as cradles to stabilize components ... against distortion created by vacuum-bags.
Yes, this might limit builders to curing only the top skin at one step. Then flip everything over and cure the bottom skin during the second step.
Do you live near a city with an airport? Chances are good there is an EAA chapter nearby that you can meet folks and see what they are building. Probably ought to join in the fun. There you can see bunches of this stuff. I have three chapters near me. There was a fire damage Long-EZ in the back of the EAA chapter hangar for over a year and a Cozy in process in another hangar here. Next chapter east of us has an Open-EZ and a Cozy both in progress, another Open-EZ that has recently started flying. Then lots of Vans, and others.

I built molds from hotwired foam. Described it on here multiple times. Advanced search tool works great.

Billski
 

Hephaestus

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OK, I shall bite. How sturdy does the clamshell mold have to be to do this? I seem to remember the EPS block molds at Olsonite (1975, so it is kind of far back in my memory) were made of beefy aluminum plate and lots of plate ribs, etc. Did not look like anything an amateur airplane builder would make two of to make one airplane.
No idea, haven't worked out any numbers - or even sat down to try to wrap my head around the process.

Never tried EPS forming don't know the first thing about it truly.

Partly why I was asking about it - figure the diversity here may lead to some new understanding.

Typically we know that molds for 1-off can be typically made lighter and more temporary than those for serial production... Don't know if that looks like eggcrate, xps foam block, masonite, concrete or what...
 

BBerson

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I think his question was how to attach a strut to a foam wing. The only example I can think of is the Sripol design.
 

stanislavz

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The only example I can think of is the Sripol design.
For skin less wings - yes. For skin loaded wings - no, you will like to add some ribs at root, to collect all loads from skin and give it to pins mounted on fuselage.
 

wsimpso1

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I think his question was how to attach a strut to a foam wing. The only example I can think of is the Sripol design.
Not sure which post you were talking about Bill. Struts can be attached several ways. Obvious ones are to locally beef up the spar for a hardpoint that will carry bolts or laminate a couple lift tabs on the spar.
 

wsimpso1

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I could not stand the uncertainty any longer, cracked open Roark's to get us some numbers on these sorts of things.

First, I established a local dynamic pressure of 1.02 psi pulling on the skins, and used a uniform pressure model. I postulated a 48" chord, 7.2" thick, and started with uniform thickness skin panel 12" by 48".

Massive cores 3.6" from midline, foam sees 1.02 psi tension, and skin deflection maxes out at about 0.00082". Low stress in the foam, and tiny deflections.

My standard Triax / 3/8" thick6 pcf divinycel foam / Biax, it sees 174 psi in the fiberglass skin, and 0.022" deflection at mid panel. Way understressed, and deflection will cooperate with laminar flow, but not with a lot of margin.

Then I put the same Triax skin over a constant thickness and blue foam core, and bumped the core thickness until foam stresses got below 35 psi. 1.65" of blue foam, 34 psi In the foam, 0.009" deflection. That will support laminar flow...

Blue foam core dropped to 0.625" thick with stress of 32 psi and deflection of 0.002.

Then I went to a 4" width, and only needed 1/2" foam thickness to get stresses down.

Yeah, the model does not include the ramped pressure and who knows whose airplane the particulars fit, but you have to leave foam under the fiberglass skin to make it work for flight. Go to impact loads, and I do expect it will be kind of fragile.

Have fun!

Billski
 
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stanislavz

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I could not stand the uncertainty any longer, cracked open Roarrks to get us some numbers on these sorts of things.

First, I established a local dynamic pressure of 1.02 psi pulling on the skins, and used a uniform pressure model. I postulated a 48" chord, 7.2" thick, and started with uniform thickness skin panel 12" by 48".

Massive cores 3.6" from midline, foam sees 1.02 psi tension, and skin deflection maxes out at about 0.00052". Low stress in the foam, and tiny deflections.

My standard Triax / 3/8" thick6 pcf divinycel foam / Biax, it sees 195 psi in the fiberglass skin, and 0.022" deflection at mid panel. Way understressed, and deflection will cooperate with laminar flow.

Then I put the same Triax skin over a constant thickness and blue foam core, and bumped the core thickness until foam stresses got below 35 psi. 2.1" of blue foam, 31psi In the foam, 0.004" deflection. That will support laminar flow, but does not remove much foam...

So I halved the panel width to 6". Now my cored skin at 1/4" gets max stress of 102 psi, and 0.003" deflection.

Blue foam core dropped to 1.1" thick with stress of 31 pounds and deflection of 0.001.

Then I went to a 4" width, and only needed 7/8" foam thickness to get stresses down.

So, no matter how small you make the panels with removed cores, the foam removed is not going to be huge. Yeah, the model does not include the ramped pressure and who knows whose airplane the particulars fit, but you have to leave quite bit of foam under the fiberglass skin to make it work for flight. Go to impact loads, and I do expect it will be pretty fragile.

Have fun!

Billski
I will sever my right hand for that - please add plate without foam at all.

And your examples are for flat panels ? I will test panels with radius of 31", and 16" wide, with ribs at 2", 4" and 6" for foam-less, and for 24" with foam.

1606928776857.png
 
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Sockmonkey

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Did I post this earlier in this thread? I'm too lazy to check.

Now you wouldn't have to do it in exactly this way, but you see what I'm getting at.
Epoxy the different foam types together, hot wire it into shape, and glass it.
 

wsimpso1

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I will sever my right hand for that - please add plate without foam at all.

And your examples are for flat panels ? I will test panels with radius of 31", and 16" wide, with ribs at 2", 4" and 6" for foam-less, and for 24" with foam.

View attachment 104848
Cutting limbs is hardly necessary... Are you saying you want to model just the fiberglass laminate with no foam support? This thread is about the foam and taking some of it out... You want to know about taking all of the foam out. If you want to do more on this, you will need a little training on plate theory and Roark's chapter 11, table 11.4 cases 8a and 8d appear most appropriate.

Yes, the modeling I performed was using flat panel modeling. For a chord of 48", 15% thickness, and 0.3 design Cl, that is an average of about a 62" curvature. Thickness of 0.4"makes a thickness of about 0.6% of radius - pretty close to flat. You are planning radii of 31" which may be a good match to the forward part of the wing, but that means the panel thickness is still about 1.3% of radius - still very small errors due to curvature at this level.
 
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wsimpso1

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Ran the numbers for on, two, and four plies of BIAX cloth and a 24" separation and 1.02 psi.

One ply of 18 oz BIAX gets 491 psi and 0.505" deflection in the center of the part - This is mighty thin and moving a lot;
Two plies gets 533 psi and 0.273" deflection - this is about the minimum in a sandwich panel, one ply on each side;
Four plies 629 psi and 0.162" deflection - Lots of cowlings and fairings with no cores at this level.

All hope of maintaining laminar flow is gone at these levels of deflections. This is why cored panels and massive cores are common.

Billski
 
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Riggerrob

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How about building wings like the deHavilland Technical School's TK.4 light racing plane. It was built by students back during the 1930s. Wings were made of huge balls wood blocks. Vertical holes removed some balsa to lighten the structure, then the entire wing was wrapped in plywood. Only one was built and it crashed a few months later.
Could this technique be updated with modern foam and composites?
I can picture vacuum-infusing the top skin with cylindrical cores still in. Lift the wing out of its cradle and pull out lightening holes.
Just not sure how to vacuum-infuse a smooth bottom skin without temporary cores.
 

wsimpso1

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Did I post this earlier in this thread? I'm too lazy to check.

Now you wouldn't have to do it in exactly this way, but you see what I'm getting at.
Epoxy the different foam types together, hot wire it into shape, and glass it.
You can try just about anything.... please tell us why you would run different densities in different parts of the foil. Oh, epoxy will absolutely stop the hot wire saw.
 
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