Abstract
The failure mechanism of a foam core sandwich panel joint specimen was evaluated as future possible airplane fuselage structure. Specimens were manufactured from UT500/#135, graphite/epoxy fabric prepreg, and PEI (polyether imide) foam core and subjected to static tension in servo hydraulic loading frame. Delamination length, onset load and point were identified through the test. The energy release rate at the delamination end was calculated with the crack closure method in the fracture mechanics approach and the results were compared with the interlaminar fracture toughness value of UT500/#135 for ENF (End Notch Flexure) test. The calculated value of G II was much higher than the experimented fracture toughness values. The modified calculation with resin impregnation at the tapered core edge (filler) much reduced this difference. This fact also indicated that the resin filler improved the strength of foam core sandwich panel join.
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