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
