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Panel Analysis

Panel Analysis

Panel Analysis

Hi folks,

I am analysis a repair problem. Please refer attached picture. The initial plate (305mmx206mmx1.6mm) has supported some installations. To facilitate installations, the holes were blanked in the panel.

Now, the new design is void of those installations. Hence these holes need to be covered. To cover those, designer is simply fastening one more panel (same dimensions) of thickness 1.8mm to the existing panel. The material of the both panels is 2024-T3. The panel is subjected to a pressure of 2dp (Not a primary structure/skin panel). When I carry out panel bending analysis using Roark, the plate is failing. Here I am assuming the lower panel with holes (original plate) is ineffective in taking the loads.

I don't have the reference of original analysis. Now I need to clear the repair. I would like to ask your opinion regarding this. Do I need to suggest increase in thickness or material? I feel it is a geometric non linearity problem, Is there any way to clear the present scheme by hand calculations so that change in material/thickness is avoided. Your reply is appreciated. Thanks in advance.


RE: Panel Analysis

It depends on what case you have selected from Roark for your reverse analysis of the panel.
There are thick plates under pressure but some cases ignore the "membrane" effect.
There are thin membrane analysis methods but those in Roark aren't preferred.
Chapter A17 in Bruhn is much more suitable to pressure loads on aircraft skins.

Also, where does the "2dp" load come from? Do you mean you are using a safety factor of 2 x the limit load, to find the ultimate load in the panel? This might be a very high load for any aluminum sheet to handle. Tyipcal aircraft skin pressures are 8 to 9 psi, and ultimate load cases tak you to 12 or 13 psi. Pressure loads on skins tend to cause about 15 psi structural stress in the skins. If this is not a skin panel, then what is applying the pressure?


RE: Panel Analysis

so you've skinned over a area with a thicker skin than the OEM put there and think you're failing. Then how could the OEM pass ?

turn around, if the OEM passes how can your's fail ?

If you analyzed as a flat plate under pressure, using Roark plate equations, then yes you are going to fail. The point is that this isn't a plate, but a membrane (and I don't think Roark can help)

another day in paradise, or is paradise one day closer ?

RE: Panel Analysis

Always good to reverse-analysis your way through the OEM structure to figure out the original loads.
My take on Istambul's question:
In his sketch, there are holes in his panel (assume lightening holes). Hence, the OEM configuration cannot retain pressure.
The panel may be skinned over for a different reason, but now that one side is isolated from the other, a pressure differential can develop.
Maybe a cabin decompression case from FAR 25.365... The OP didn't mention it but this is one way this could happen.


RE: Panel Analysis

What are the constraints to your geometry? Which locations was the first (pre) plate attached to the rest of the aircraft? And how is it fixed? Fixed only in translational? OR fixed in all degrees of freedoms (translational+rotational)?

Once your boundary conditions (for each edge)(constraints and exact pressure load) are clear, we might be able to figure the exact equation of your assembly.
Since you used mm for lengths, can you tell me the exact pressure in MPa (MegaPascals)?

Aerospace Engineer, M.Sc. / Aircraft Stress Engineer

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