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

HomeBuiltAirplanes.com

Help Support HomeBuiltAirplanes.com:

Vigilant1

Well-Known Member
Lifetime Supporter
Joined
Jan 24, 2011
Messages
5,879
Location
US
It will be interesting to see how much skin bulging will occur at various foam thicknesses, web layouts, skin thickness, and negative pressures.

And, it will be a good opportunity to do the "ball peen" handling toughness tests of skins.

I don't suppose anyone wants extra weight, or to spend extra time, or to build in a medium they don't enjoy. So, lots of compromises and sradeoffs.

Mark
 
Last edited:

wsimpso1

Super Moderator
Staff member
Log Member
Joined
Oct 18, 2003
Messages
7,795
Location
Saline Michigan
Can I ask your opinion on extending control surfaces chord for lighter weight of whole wing ? Example taken from BoKu carbon max and seeing how Cora ul is build with solid D nose and fabric rear part of the wing.

30% of D nose, 45% of fabric, 25% of controls, versus 65% of solid cored wing and 35% of control surface covered with fabric. For 10m2 / 107 sq ft we are wasting in ideal case 15-20 lb / 7-10 kg of weight if we change fabric to adequate solid skin. On latter one - half of it. Or nil, because transition from solid skin to fabric needs some filler, or you are cowering D cell with fabric too as done for most wooden wing with plywood D cell..
Hmm. See my other recent post. The big question is what is most important to you in your airplane? If lightness is more important than drag, the classic wooden wing with a plywood skin on all or part of the wing and then fabric is a terrific way to go. If drag is more important than weight, fiberglass is the terrific way to go. If drag and weight are both really important but cost is secondary, graphite fiber is the terrific way to go. You guys are talking about taking a little weight out of the heaviest wing, and messing up its drag too - why do that? If you want low weight and are willing to tolerate a little more drag, wood and fabric is pretty terrific, and you do not have to pioneer a new scheme that can kill you.

Maybe you guys are messing around in a corner of the design space where you can actually make a light laminar flow wing this way, but the usual result is that this is worse for drag and worse for weight than other options. You would have to prove it to me though...

I am back to challenging the hollowed core true believers to design and analyze and optimize some wing for some flight envelope in fiberglass with a solid core and then with a ventilated core and show us what you give up. Oh, and remember, your FOS for composites is 2.0, not the 1.5 you use in metal and wood. Have fun.

Billski
 

Vigilant1

Well-Known Member
Lifetime Supporter
Joined
Jan 24, 2011
Messages
5,879
Location
US
And a little voice inside says "if we just epoxy some 4oz fiberglass on the inside of those cutouts...stiffness no longer a problem. Invent a taping tool to push into the cavity...." (running for cover)
 

wsimpso1

Super Moderator
Staff member
Log Member
Joined
Oct 18, 2003
Messages
7,795
Location
Saline Michigan
Yes, wing center mass at 25% chord would be ideal for flutter. Rotor blades are balanced at 25%. Most wings are not, but why not have more foam density at front?
Bill tells part of why here, and I can not resist telling the rest... Sorry Bill.

The wing's shear center needs to be at or forward of the aero center for it to not try to diverge in pitch. In short stiff wings this is not too important, but in long high aspect ratio wings and in the compressibility regime (helo rotors qualify), this really is important for beating flutter. This is achieved in helo rotors by making the front of the blade solid and the rest a thin shell or of a very light material, and while also putting the CG about at the quarter chord point too. Now the foil does not try to diverge and will pretty much stay where you put it in pitch, even with it being oscillated one cycle per rotation.

In short stiff wings (like most of our homebuilts) and with low airspeeds, the shear center can almost be anywhere and it will usually be OK, unless the wing is exceptionally soft. Look at the large number of airfoils with the main spar back 35 to 40% c and same skin overall or one increment thicker forward than aft. Shear center is way aft in them, and they still work fine. They are trying to diverge in pitch, but they diverge small fractions of a degree, it does not matter much.

Billski
 

wsimpso1

Super Moderator
Staff member
Log Member
Joined
Oct 18, 2003
Messages
7,795
Location
Saline Michigan
And a little voice inside says "if we just epoxy some 4oz fiberglass on the inside of those cutouts...stiffness no longer a problem. Invent a taping tool to push into the cavity...." (running for cover)
It is done. Bike frame guys make tubes or whole frames inside two piece female molds. The suppliers make soft plastic tubes in a variety of diameters and a sealing tool to close one end. The other end goes inside the tube you are fabricating, and is led out through the vacuum bag, then exposed to atmospheric pressure. When the two piece mold is closed, wrapped in a vacuum bag, and the air pumped out, the plastic tube inflates to press the prepreg material against the inside of the mold. Then you run a temperature ramp to first allow the pre-pregs to debulk against the mold, then to cure the resin.

You are only limited by your imagination. And your ability to analyze the thing and see what is needed where to make it work. Oh, and to see if it has any advantage over other schemes...

Billski
 

Vigilant1

Well-Known Member
Lifetime Supporter
Joined
Jan 24, 2011
Messages
5,879
Location
US
It is done. . . .

You are only limited by your imagination. And your ability to analyze the thing and see what is needed where to make it work. Oh, and to see if it has any advantage over other schemes...
No, I'm just joking. But I did think about ways to get a little reinforcement onto the other side of that foam (and it wouldn't take much). But I will speak no more of it, now.

Mark
 

ragflyer

Well-Known Member
Joined
Apr 17, 2007
Messages
310
In Rutan solid core wings the foam is only there as buckling support for the skins and spar flanges. The characteristic for the foam to do its job is rigidity not strength. Unlike a classic sandwich where the foam/core takes the shear load the Rutan wing has BID wrapped all around, and this takes the shear loading leaving the core as only support.

In a Rutan wing the core is particularly effective in supporting the skins and as such enables both the facings and web to take virtually its full ultimate tension and compressive load without prematurely buckling or wrinkling. I have back analyzed to verify this.

Now as you hollow out the foam the buckling support becomes compromised. At some point when the thickness of foam between the skin and hole is small enough the skin/ facings will wrinkle (before its ultimate load) either inward or outward into or above the holes in the foam. This is a well understand phenomenon in composite structures called wrinkling. The same thing happens when you choose a very low density core. The facings will pull out or dig it to the foam causing wrinkles or ripples. In honeycomb there is somewhat related phenomenon called dimpling where the facings dimples into the hole.

So the question is how big can your holes in the foam get before wrinkling occurs. The only way to know is testing. There is no analytical solution to this. FEA can give you an answer but is of little value if you cannot validate with tests. Botton line test.

PS. by the way wrinkling in classic sandwich structures has a pretty simple formula (Hoff formula) but does not apply to Rutan type wings (with or without holes).
 

Vigilant1

Well-Known Member
Lifetime Supporter
Joined
Jan 24, 2011
Messages
5,879
Location
US
If we know the in-plane compression load on the wing skin, can we determine (with high confidence) how much skin deflection is allowable before buckling occurs?

Bad for us: Highest exterior pressure loads on the skin occur in the same flight regime where in-plane skin compression loads are greatest.

Good for us: the zone of highest external skin load pressures is fairly localized (near the leading edge)
 

ragflyer

Well-Known Member
Joined
Apr 17, 2007
Messages
310
I do not believe membrane stresses ( for reasonable sizes holes) are a significant issue here. Remember foam is not a solid material, well it is a foam- it is porus and filled with air. The only difference between 'solid' foam and your hollowed foam is the size of the holes.
 

BBerson

Light Plane Philosopher
HBA Supporter
Joined
Dec 16, 2007
Messages
14,611
Location
Port Townsend WA
The original Rand KR-1 used a 1" thick 2lb per cubic foot rigid polyurethane and a single 4oz. per sq. yard Dynel outer skin. Later, fiberglass was used. And 2" thick foam ribs spaced at 20".
 
Last edited:

ragflyer

Well-Known Member
Joined
Apr 17, 2007
Messages
310
One relatively conservative approach would be to restrict the holes to only aft of the spar and design to take all the torsion load in the D nose. The skin loads aft of the spar then are very low and should not be critical. Of course a simple test will tell you for sure.
 

ragflyer

Well-Known Member
Joined
Apr 17, 2007
Messages
310
Of course take the idea above to an extreme then all you have is a Foam/glass D nose and aft of the spar you hang a few ribs and cover with fabric.

Goes back to the earlier comments by Billski - what are you trying to achieve?
 

stanislavz

Well-Known Member
HBA Supporter
Joined
Aug 21, 2016
Messages
929
Location
Lt
Maybe you guys are messing around in a corner of the design space where you can actually make a light laminar flow wing this way, but the usual result is that this is worse for drag and worse for weight than other options. You would have to prove it to me though...
Thank you. So my turn. Full cored or xps in any form is no-no for me, due to mainly unavailable proper xps in EU. Do some samples, with bone dry xps (few hours in oven at 40-45 degree of celsius), and some sanding i was able to get proper bond on 2/3 of tests. Other aspect - too heavy for any aircraft slower than 150 mph.

So, I am a fan of chord-vise core lightning. Aka milion ribs as Vigilant mentioned. As i dig more on it - it is quite widely used. First one for me was on Cri-cri, later - on some glider, here is one example for glider : Udo Rumpf Wing Project

Examples which i was able to find shows rib spacing in 45-100 mm range for laminar flow applications and skinned with 0.5mm / 0.02" aluminum or 120-150 mm if skinned with plywood. Or 225-300 mm in rib spacing for-non laminar flow airfoils.

And on weight number. 100kg/m3 pvc 6mm foam with ribs spaced 45 mm - gives your 13 kg /m3 0.8 pound per cubic foot. Which is on pair with lowest density eps. 90 mm spacing gives you half of it etc..

But - doing it in aluminum skin requires some fancy structural adhesives, and jigs for gluing it. Both this could be solved by going to CF in place of aluminium and even making ribs from CF itself. Only if one will be able to mold and glue ribs to skin in one operation. If not - extra adhesive to glue ribs to pre formed skin will ruin your mass balance. 100kg/m3 pvc 6mm foam is same as one ply of 300 gsm cf.

More on comparision. Bare minimum wing skin from 200gsm cf / 4 mm 75kg/m3 pvc foam / 200gsm cf gives you 1.1 kg per 1m2. Or 2.2kg per both wing skins - top and bottom. And if we move foam from skin to ribs, for 13.5% airfoils i am getting this data : airfoil area is ~ 10% of chord. So for 1 m2 chord wing: 1.6 kg for skins, 0.6kg for foam gives your 100mm spaced ribs from 6mm 100kg/m3 foam or 75mm spaced for 75kg/m3 etc..

Live examples have 2x300gsm fiberglass skin and ribs spaced each 225-300mm. Totally ok on safe side.

And for this "extra new" styled wing - you will have to test only front D cell in your pressure chamber. + torque.
 

Speedboat100

Banned
Joined
Nov 8, 2018
Messages
1,900
Location
Europe
Bill tells part of why here, and I can not resist telling the rest... Sorry Bill.

The wing's shear center needs to be at or forward of the aero center for it to not try to diverge in pitch. In short stiff wings this is not too important, but in long high aspect ratio wings and in the compressibility regime (helo rotors qualify), this really is important for beating flutter. This is achieved in helo rotors by making the front of the blade solid and the rest a thin shell or of a very light material, and while also putting the CG about at the quarter chord point too. Now the foil does not try to diverge and will pretty much stay where you put it in pitch, even with it being oscillated one cycle per rotation.

In short stiff wings (like most of our homebuilts) and with low airspeeds, the shear center can almost be anywhere and it will usually be OK, unless the wing is exceptionally soft. Look at the large number of airfoils with the main spar back 35 to 40% c and same skin overall or one increment thicker forward than aft. Shear center is way aft in them, and they still work fine. They are trying to diverge in pitch, but they diverge small fractions of a degree, it does not matter much.

Billski

What would happen if the planes wing mass center was at 18-20% position of the chord ?
 

wsimpso1

Super Moderator
Staff member
Log Member
Joined
Oct 18, 2003
Messages
7,795
Location
Saline Michigan
That's quite a find, thanks.
So, in the modeled case, the Reynolds number is 1,000,000 and the AoA is 15 deg. The airspeed field ("Mach") is listed as zero, I don't quite know why that is. Maybe it doesn't matter because Cp is a ratio between freestream pressure ("0") and the dynamic pressure ("1"). So, you can use the depiction to find the actual pressures simply by providing your own value of q for the airspeed and air density of interest (?)

If this model represented our 48" chord wing, we can assume the area inside the wing (black in the depiction) is at freestream static pressure. At 150 mph (dynamic pressure "q" = 0.4 psi (16.6 kiloPascals) at sea level) , then the maximum negative pressure we see in the depiction is the dark blue "Cp = -5.0" region on the nose. At this airspeed, q x (-5.0) = - 1.95 psi (or -83 kP). That is, on each square inch of the outside wing surface touching the dark blue, the wing skin (and the foam under it) is being pulled outward with a force of about 2 pounds (or, each cm2 is being pulled outward with 0.14 kgf).
On the underside of the wing, where the Cp is moderately positive (e.g. +0.4), we'd have some pushing in on the wing skin. Where the Cp is +0.4, every sq inch would see 0.4 x 0.4psi = 0.16 pound of "inward" pressure against the skin.

If the depiction is for our wing at a lower airspeed or ambient pressure, then the forces would be lower, proportionate with q.

I think a simple physical model could be constructed. It would be fiddly and look like an exploded harpsicord, but it would be cheap and able to (roughly) test various skin/foam core configurations. More later . . .
Now here is where I think you need to be modeling in some way.

In solid foam with a suitably sturdy skin laminate, we already have thousands of airplanes proving this works just fine. In a wood spar and rib fabric covered airplane, we already have over 100 years experience showing that a plywood skinned D-tube and pretty close spaced ribs work fine, and fully skinned with close spaced ribs work fine too on faster airplanes. Similarly, sheet metal works fine. And finally, sandwich cored composite skins with vey few ribs also works well, if the skin sandwich core is thick enough.

These skin and rib structures all have the same basic restrictions. Between ribs, the skinning is carrying the airloads to a periphery of stiffer structure. Stresses (in homogenous skin materials) go with shorter panel size squared divided by panel thickness squared. Panel deflections (again in homogenous materials) go with panel dimensions to the fourth power divided by E and shorter panel thickness cubed. Both stress and deflection increase as the longer side increases, but infinitely long is only a little worse than a long side twice as long as the short side length. This all sets rib spacing and panel sizes on skin and rib structures. Stresses move in one set of proportions, while deflections move in another. Modeling is great on this because you can play with proportions and see how it does.

If you are determined to play with it empirically, might I suggest making full scale models of your cutouts in the same foam and composite skins, and pressurize them internally with a water column or bicycle pump so you can ramp pressure up and down accurately, then record deflections and first failures. You will likely hear creaking and popping or see your dial indicator jumping to announce failure. I also suggest a lighted bore scope and an inspection port to examine for internal cracking as you go. Once you establish max pressure for acceptable stresses and deflections for one set of skin thickness and cutout shape/thickness, then you can probably extrapolate to other proportions and try them. Going about this intelligently may allow you to keep your number of test pieces modestly small.

Once you start thinking about this, you can probably come up with a method to do load cycling. For instance you can mount your whole test piece with water column on a rotisserie, that gives several revolutions a minute, and get a thousand load cycles in a day or two. I imagine you could even put a microphone on the fiberglass and a Raspbery Pi programmed to count cycles and listen for pops and creaks to stop test at first failure.

Remember that if you establish a dimension set and a max pressure for first fiber failure in the facing or first crack in the foam, you are limited to one half that pressure in that configuration in the airplane. Remember too that airfoil shape change may stop you well short of strength issues. In many applications, skin deflection limits for laminar flow design the skin panel size and sandwich core thickness. Perhaps true here too. Oh, the length of the air cels should be at least twice the width to approximated the long cells you will get in the airplane.

Good luck with the sample prep and please write it up for us, or maybe even for KitPlanes.

Billski
 
Last edited:

wsimpso1

Super Moderator
Staff member
Log Member
Joined
Oct 18, 2003
Messages
7,795
Location
Saline Michigan
Why not build a carbon fiber/Kevlar shell that attaches to the spars (crude black lines), then foam around it and vacuum bag a carbon fiber/kevlar shell over it?


View attachment 104614
This is a sandwich skin composite wing, but made, shall we say, the hard way. The easy way is to make simple female molds and vacuum bag each top and bottom skin in one session, then assemble the wing with those skins, main spar, and a few ribs. Drag spar would be integral with the lower skin, assuming trailing edge devices. This way would require at least as much tooling, an internal structure sturdy enough to stand the vacuum bagging, and most likely a lot more fill and fair work to finish the outside. Much better to just do it the easy way.

Billski
 
Top