Which size you have used ? 200ml or some bigger ones ? I am thinking of 500-700ml, and more on flat shape.
Not-so-solid massive core wings: Lightening the core foam
- Thread starter Vigilant1
- Start date
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Usually that time is the pot life, and they specify the pot size. Epoxy makes heat as it cures, and that heat raises temperatures more in the cup than it does spread out on the part. You would have to do an experiment to see how much more time you get on an open part and on a vacuum bagged part - cores and the extra plies all have some insulative properties. I have made parts with Gougeon 125/229 that took 90 minutes to be under full vacuum, and they came out fine. Oh, climate controlled shop 70-75F (21-24C) and modest humidity.
You have this massive solid core wing. Now what? How do you attach it to the airplane?
Zip ties. Or some duct-tape.
If serious - you will have one or two spars. Skin loads goes to them. And you have many options how to go next..
This topic is on how to get rigid skin on bird which is in 150-200mph max range. Which is fine with fabric..
You won't need to remind TiPi of this. Here's the very nice build log for the SD-1 he's buiding, including a picture of the work the mice did to enlarge some holes in a stack of his stored XPS ribs.
Maybe a step before final wing closure:
1) Insert mouse baits and mothballs into rib bays. Replace baits (on back of bellcrank inspection covers) during annual inspection.
Well if you want a nice white skull mount you put the skull in a drawer of maggots, they clean it up nice so maybe some kind of critter can be of aid.
Hephaestus
Well-Known Member
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.
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.
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.
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There are a few ways to do it, but it can be a little different/more complicated with composite spars.
1) We can build a composite built-up box spar of some type. Slide them into a slot in the fusleage. This post has a picture of the wing/spar of the Mead Adventure, another plane with a solid foam wing (though there aren't many out there).
2) We can build the spar caps and web directly on the solid wing core, and add an extra length of foam (I'd probably use PVC foam) out of the root end of the solid wing to extend the spar into the fuselage. There, the two wing spars overlap and attach to each other and the fuselage. As with everything else, the foam isn't passing any significant loads, it's just holding the composite skins in position which helps prevent them from buckling under compressive loads.
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It does sound like a possibly efficient process for rapidly producing wings in an industrial environment. Two-part pressure molding, the use of steam "popped" polystyrene beads within those molds, pre-preg-- none of that sounds like the kind of thing designed for a one-off homebuilt. But, maybe the FAA's interpretation of the 51% rule will eventually get relaxed enough that a kit company could send out the EPS-filled wings already skinned and ready for ailerons and position lights.
I'd suspect that Rohr used EPS rather than XPS primarily because it was a very fast way to fill that 3D shape. If the wing skin really needs the compressive modulus typically associated with a 60 PSI foam, then doing it in EPS will be heavier than XPS, not lighter (by about 1lb per cubic foot). Still, especially in the days before cheap CNC cutting, it would have saved them a lot of labor hours to just throw the EPS "sugar" into the mold and turn on the steam rather than shape a block of XPS.
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If I am thinking right that expanded stuff is the coffee cups we used to dissolve into gasoline as kids to make kind of a napalm because , you know ... it was fun to play with.
https://www.google.com/search?client=opera&q=napalm&sourceid=opera&ie=UTF-8&oe=UTF-8
https://www.google.com/search?client=opera&q=napalm&sourceid=opera&ie=UTF-8&oe=UTF-8
External local pressures on a wing skin
Some refinements on my earlier estimates of the pressures on a wing skin. This will (someday) be used to set parameters for testing various composite wing skin and foam core configurations. Goal now: Get an estimate of how many PSI (max) I'll need to be pulling up on the skin of the test piece/model.
I'm sure there are better ways to do this, and am open to any suggestions to improve this "methodology."
Example aircraft of interest:
Max Weight: 1400 lbs
Wing: 120 sq ft area, 30' span, 4' chord. Airfoil: NACA 4412 (for simplicity: no flaps)
Service loads: +4.4G, - 1.76 G (200% safety factor = +8.8G, - 3.52 G)
Maneuvering Speed (VA) : 110 MPH
Cruise speed (Vc): 130 MPH
Never Exceed (Vne): 160 MPH
Vdive: 180 MPH
The pressure around a particular wing in flight depends on two things: the dynamic pressure and the AoA. The two places in the allowable flight envelope where we are most likely to find maximum pressures are:
Case 1) A pull to Clmax at our maneuvering airspeed (110 MPH). What AoA will that be? TOWS shows the Clmax of the NACA 4412 at our Reynolds number is about 12 degrees AoA. Per NACA Report 563, the maximum negative pressure coefficient around the airfoil at 12deg AoA is -3.7 . The dynamic pressure (q) at 110 MPH is .21 PSI. So, the minimum pressure anywhere on the airfoil skin under these conditions will be - 3.7 x 0.21PSI = - 0.78 PSI
Case 2) A maximum allowable G (+4.4) pullout at Vdive. Lift needed: 1400lbs (MTOW) x 4.4G x 2 (FOS) = 12,320 lbs. To generate that lift at 180 MPH (Vdive) and with our 120 sq ft wing requires a Cl of 1.24. TOWS tells us that the venerable NACA 4412 will make that at an AoA of 8 degrees. Per NACA Report 563, the maximum negative pressure coefficient around the airfoil at 8 deg AoA is -1.8. The dynamic pressure at 180 MPH is .57 PSI. So, the minimum pressure anywhere on the airfoil skin under these conditions will be - 1.8 x 0.57 PSI = - 1.03 PSI
It looks like the Vdive pullout scenario is the most strenuous case, and our physical model will need to be able to generate about 1 lb of pull on every inch of the wing's skin, at least in a small localized area. The skin area subject to this 1 PSI pull is very small (near the top of the LE), and most of the rest of the skin is either under modest positive pressure or less than 1/2 PSI negative pressure.
One gotcha: I applied the 200% safety factor to Case 2 above. I didn't apply any safety factor to Case 1, since it wasn't clear to me how to best meet the intent of that safety factor. At Va, the wing can't generate twice as many Gs, so I can't realistically find an AoA to accomplish that function (as I did in Case 2). Maybe, since the issue we designing for is the strength of the foam/skin system, my Case 2 pullout scenario should have been at 4.4 G (not 8.8 G). Just do both scenarios as "max stress within allowable G and airspeed" events, then found the max PSI, then applied the FOS of 200% to the max PSI for both scenarios. I'll try that next.
ETA: A Re-do of the dive scenario (Case 2) with a 4.4G pullout. At 180 knots, this requires a 0.62 Cl. TOWS tells us that requires a 2 deg AoA. The lowest coefficient of pressure at that AoA is -0.77, so the lowest pressure is -0.77 x 0.57 (dynamic pressure) = - 0.44 psi
So, the most challenging case is Case 1 (Va max G pull) producing a - 0.78 psi pressure. Apply a 200% FOS and we need to test our physical model to approx -1.6 PSI ultimate. Still, under normal ops, it looks like the max expected pressure anywhere will be about -0.8 psi, and if the model survives that (repeatedly) and returns to the original shape without damage, that will be a good sign.
Sorry for the meandering post.
Some refinements on my earlier estimates of the pressures on a wing skin. This will (someday) be used to set parameters for testing various composite wing skin and foam core configurations. Goal now: Get an estimate of how many PSI (max) I'll need to be pulling up on the skin of the test piece/model.
I'm sure there are better ways to do this, and am open to any suggestions to improve this "methodology."
Example aircraft of interest:
Max Weight: 1400 lbs
Wing: 120 sq ft area, 30' span, 4' chord. Airfoil: NACA 4412 (for simplicity: no flaps)
Service loads: +4.4G, - 1.76 G (200% safety factor = +8.8G, - 3.52 G)
Maneuvering Speed (VA) : 110 MPH
Cruise speed (Vc): 130 MPH
Never Exceed (Vne): 160 MPH
Vdive: 180 MPH
The pressure around a particular wing in flight depends on two things: the dynamic pressure and the AoA. The two places in the allowable flight envelope where we are most likely to find maximum pressures are:
Case 1) A pull to Clmax at our maneuvering airspeed (110 MPH). What AoA will that be? TOWS shows the Clmax of the NACA 4412 at our Reynolds number is about 12 degrees AoA. Per NACA Report 563, the maximum negative pressure coefficient around the airfoil at 12deg AoA is -3.7 . The dynamic pressure (q) at 110 MPH is .21 PSI. So, the minimum pressure anywhere on the airfoil skin under these conditions will be - 3.7 x 0.21PSI = - 0.78 PSI
Case 2) A maximum allowable G (+4.4) pullout at Vdive. Lift needed: 1400lbs (MTOW) x 4.4G x 2 (FOS) = 12,320 lbs. To generate that lift at 180 MPH (Vdive) and with our 120 sq ft wing requires a Cl of 1.24. TOWS tells us that the venerable NACA 4412 will make that at an AoA of 8 degrees. Per NACA Report 563, the maximum negative pressure coefficient around the airfoil at 8 deg AoA is -1.8. The dynamic pressure at 180 MPH is .57 PSI. So, the minimum pressure anywhere on the airfoil skin under these conditions will be - 1.8 x 0.57 PSI = - 1.03 PSI
It looks like the Vdive pullout scenario is the most strenuous case, and our physical model will need to be able to generate about 1 lb of pull on every inch of the wing's skin, at least in a small localized area. The skin area subject to this 1 PSI pull is very small (near the top of the LE), and most of the rest of the skin is either under modest positive pressure or less than 1/2 PSI negative pressure.
One gotcha: I applied the 200% safety factor to Case 2 above. I didn't apply any safety factor to Case 1, since it wasn't clear to me how to best meet the intent of that safety factor. At Va, the wing can't generate twice as many Gs, so I can't realistically find an AoA to accomplish that function (as I did in Case 2). Maybe, since the issue we designing for is the strength of the foam/skin system, my Case 2 pullout scenario should have been at 4.4 G (not 8.8 G). Just do both scenarios as "max stress within allowable G and airspeed" events, then found the max PSI, then applied the FOS of 200% to the max PSI for both scenarios. I'll try that next.
ETA: A Re-do of the dive scenario (Case 2) with a 4.4G pullout. At 180 knots, this requires a 0.62 Cl. TOWS tells us that requires a 2 deg AoA. The lowest coefficient of pressure at that AoA is -0.77, so the lowest pressure is -0.77 x 0.57 (dynamic pressure) = - 0.44 psi
So, the most challenging case is Case 1 (Va max G pull) producing a - 0.78 psi pressure. Apply a 200% FOS and we need to test our physical model to approx -1.6 PSI ultimate. Still, under normal ops, it looks like the max expected pressure anywhere will be about -0.8 psi, and if the model survives that (repeatedly) and returns to the original shape without damage, that will be a good sign.
Sorry for the meandering post.
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Hephaestus
Well-Known Member
Still makes me wonder, Single use molds are a little easier in this day and age. Sheetmetal and disposable CNC'd inexpensive construction materials, If you can salvage sections (ie prefinished skin) from one mold to the next - it's not the worst method we've ever discussed here.
Also allows adding say - a composite box in there to create specific voids in specific areas. That then gets foamed into place within the wing structure.
The original prototype aircraft were hand laid fiberglass over monolithic xps blocks...
Please do not forhet - we are on home building area...
Due to high activity in this thread could we try to reinvent bd4/bd5 wing ? No more than 10 replies. Lets go from milion ribs to 100 000 wing panels.
For example : my hands are fully ok to reach and laminate panel with 600mm 24" length/depth. And we will have to place inside at least 2-4 pe strips with some cf for local reinforcements against buckling. As in BoKu carbon max.
6/7 panels per wing - a week of time, you will have left and right molds anyway..
Which kind of spar we would like to use ? I beam bulid up spar ? Or big fat tube ?
Due to high activity in this thread could we try to reinvent bd4/bd5 wing ? No more than 10 replies. Lets go from milion ribs to 100 000 wing panels.
For example : my hands are fully ok to reach and laminate panel with 600mm 24" length/depth. And we will have to place inside at least 2-4 pe strips with some cf for local reinforcements against buckling. As in BoKu carbon max.
6/7 panels per wing - a week of time, you will have left and right molds anyway..
Which kind of spar we would like to use ? I beam bulid up spar ? Or big fat tube ?
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.
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.
It depends on many setting, on eps it is about 0.7-1mm, 3-4 * wire gauge. I think it is used by rc guys for balsa sheeting.
Well, my hot wire cuts never seem to recede below the template protecting the foam, my guess is the kerf is about 0.050", which is more than the thickness of my skin laminates. I always use the template to protect the piece of foam I need the dimensions close on. When I keep the cutoffs to make saddles, I do not worry over this issue, the part core is to size under the laminate, and the supports are what they are.
As to the process, see these:
Billski's Fiberglass Bird
Flaps are starting post 9 and ailerons start on post 24. I had done the horizontal tail long before by the same methods, and the vertical tail done the same way is coming up.
I have never laminated between female cutoffs and the male part core except where it is tooling. In a part, way too much resin gets left in the part that way. I laminate one side at a time as illustrated.
Oh, and I am positive that I only figured out something that had been figured out before, as did you.
Billski.
Testing plan
I think it might be enlightening to run a few low-tech tests relevant to the lightened core theme. I can order supplies now and get to it sometime after the holidays. Thoughts and suggestions are welcome.
Based on conversations here and elsewhere, I see three aspects worth examining:
Phase 1) Resistance to hangar rash. How well does the outer surface survive the indignities expected in aviation service? I'll start a separate thread on this, since it isn't strictly limited to "lightened solid foam core" idea. It'll look something like a hillbilly version of the ASTM D7766/D7766M impact testing of sandwich panels (a modified "Procedure C": edge-supported sample with a dropped, smooth semi-hemispherical 5/8" impactor ). Having some comparative information on how well various laminate/substrate combinations hold up will be useful in determining what to test in the subsequent phases. I don't think there will be a "pass/fail" to this process, a surface that isn't tough enough to take flight-school abuse might be fine foe a pampered owner-flown Sunday flier.
Phase 2) Plane deflections due to membrane loads: This would test some combinations of skin laminate schedules atop varying thicknesses of foam--maybe different types of foam. Probably 4" wide test coupons about 12" long, firmly supported at both ends (clamped to the laminate skin, not through the foam). For a wing with a hollowed-core foam core, the largest pressures due to flight loads would "push" from the inside to the outer surface, so the tests of primary interest would be with the planks supported at their ends and the evenly distributed weight pressing down on the bare foam. Still, wings are also subject to loads the other way, so I'd flip them over and apply weights to the laminate side. I'd measure deflection with increasing loads.
The information gathered from these simple tests may be useful to folks who need base data to plug into more sophisticated software tools. It might be useful as a starting point for making further choices in simple design cases (rib spacing vs surface control for laminar flow, etc). And, I think it may provide information useful for refining configurations worth testing in the next phase.
Phase 3) Deflections of lightened core wing sections. Here, I'd mock up some sample sections of a wing (maybe the nose and a representative section aft of the spar). Forces simulating aero loads would be placed on the sections and deflections measured. Since the loads on the top surface won't be the same as loads on the bottom surface, I think the test section will need to be stabilized with something like a spar (or I'll be chasing the thing around my garage). After doing the Phase 2 tests, we should have a better idea of the relationship between foam thickness, foam density/stiffness, and skin deflection, and this will help decide what types of configurations would be worth testing as (partial) wing sections. Ellipse-shaped voids, or more aggressive rectangular or triangular voids, etc. It is possible that if some limited FG or CF reinforcement of the voids looks practical and useful, and I could test that.
I'd like to do the tests using simple equipment and make them easily repeatable by others (so if someone wants to try a new laminate/foam/lightened core combination, it can be easily compared to this, added to the data set, etc). I'll probably also throw in adhesion tests with some readily available hardware store foams. Also, I think it would be useful to include some tests of already flying profiles (e.g. 21oz FG over solid core, 1mm aircraft plywood, etc) to make the data on lightened core sections more relevant by giving them context.
So, that's the idea, I'm open to suggestions. I have no composite experience, so will be learning to vacuum bag, etc. Also, like everybody, I've got other responsibilities and life can get in the way, so nobody should hold their breath or delay a project waiting for me to get to work. I'm a bit of a layabout and my interests are fickle. Fair warning . . .
I think it might be enlightening to run a few low-tech tests relevant to the lightened core theme. I can order supplies now and get to it sometime after the holidays. Thoughts and suggestions are welcome.
Based on conversations here and elsewhere, I see three aspects worth examining:
Phase 1) Resistance to hangar rash. How well does the outer surface survive the indignities expected in aviation service? I'll start a separate thread on this, since it isn't strictly limited to "lightened solid foam core" idea. It'll look something like a hillbilly version of the ASTM D7766/D7766M impact testing of sandwich panels (a modified "Procedure C": edge-supported sample with a dropped, smooth semi-hemispherical 5/8" impactor ). Having some comparative information on how well various laminate/substrate combinations hold up will be useful in determining what to test in the subsequent phases. I don't think there will be a "pass/fail" to this process, a surface that isn't tough enough to take flight-school abuse might be fine foe a pampered owner-flown Sunday flier.
Phase 2) Plane deflections due to membrane loads: This would test some combinations of skin laminate schedules atop varying thicknesses of foam--maybe different types of foam. Probably 4" wide test coupons about 12" long, firmly supported at both ends (clamped to the laminate skin, not through the foam). For a wing with a hollowed-core foam core, the largest pressures due to flight loads would "push" from the inside to the outer surface, so the tests of primary interest would be with the planks supported at their ends and the evenly distributed weight pressing down on the bare foam. Still, wings are also subject to loads the other way, so I'd flip them over and apply weights to the laminate side. I'd measure deflection with increasing loads.
The information gathered from these simple tests may be useful to folks who need base data to plug into more sophisticated software tools. It might be useful as a starting point for making further choices in simple design cases (rib spacing vs surface control for laminar flow, etc). And, I think it may provide information useful for refining configurations worth testing in the next phase.
Phase 3) Deflections of lightened core wing sections. Here, I'd mock up some sample sections of a wing (maybe the nose and a representative section aft of the spar). Forces simulating aero loads would be placed on the sections and deflections measured. Since the loads on the top surface won't be the same as loads on the bottom surface, I think the test section will need to be stabilized with something like a spar (or I'll be chasing the thing around my garage). After doing the Phase 2 tests, we should have a better idea of the relationship between foam thickness, foam density/stiffness, and skin deflection, and this will help decide what types of configurations would be worth testing as (partial) wing sections. Ellipse-shaped voids, or more aggressive rectangular or triangular voids, etc. It is possible that if some limited FG or CF reinforcement of the voids looks practical and useful, and I could test that.
I'd like to do the tests using simple equipment and make them easily repeatable by others (so if someone wants to try a new laminate/foam/lightened core combination, it can be easily compared to this, added to the data set, etc). I'll probably also throw in adhesion tests with some readily available hardware store foams. Also, I think it would be useful to include some tests of already flying profiles (e.g. 21oz FG over solid core, 1mm aircraft plywood, etc) to make the data on lightened core sections more relevant by giving them context.
So, that's the idea, I'm open to suggestions. I have no composite experience, so will be learning to vacuum bag, etc. Also, like everybody, I've got other responsibilities and life can get in the way, so nobody should hold their breath or delay a project waiting for me to get to work. I'm a bit of a layabout and my interests are fickle. Fair warning . . .
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Granted.
I was just thinking of building a new set of wings to bolt to a completed Bede fuselage.
If you are starting with a mere pile of of fuselage parts, you can fabricate wing attach fittings any shape you want.
