![]() The overall program objective was to investigate the effects of holes and notches on residual strength. This evaluation was a part of a more general study on the damage tolerance of six honeycomb sandwich composite curved panels, each containing a different damage scenario. The purpose of the study was to evaluate the AE technique as a tool for detecting notch tip damage initiation and evaluating damage severity in such structures. In addition, the presence of multiple cracks reduced the residuaĪcoustic emission (AE) was monitored in notched full-scale honeycomb sandwich composite curved fuselage panels during loading. However, the number of cycles to grow a fatigue crack to a predetermined length was reduced by 37% and 27% for the longitudinal lap joint and circumferential butt joint panels, respectively. In general, results show that multiple cracking did not have an effect on the overall global strain response. ![]() Results include comparisons of strain distributions, fatigue crack growth characteristics, and the damage growth process during residual strength test for the two joint configurations. For fatigue crack growth predictions, the corresponding mixed mode stress-intensity factors were calculated using the Modified Crack Closure Integral (MCCI) method. Comparisons with strain-gage data verified the finite element models. The strain distributions and fracture parameters governing crack formation and growth were determined. Geometric nonlinear finite element analyses were conducted to support the tests. Third, the crack growth and residual strength were measured during quasi-static loading to failure. Second, fatigue crack formation and growth were monitored and recorded in real time using the Remote Controlled Crack Monitoring (RCCM) system under constant amplitude loading up to a prescribed amount of fatigue crack growth. First, strains were measured under quasi-static load conditions to ensure proper load introduction into the panels. The panels were tested in the Full-Scale Aircraft Structural Test Evaluation and Research (FASTER) facility. For each joint configuration, one panel contained only a lead crack and the other contained a lead crack with multiple cracks located along the outer critical fastener row of the joints. Four panels were tested, two panels with a longitudinal lap splice and two with a circumferential butt joint. An experimental and analytical study was conducted to determine the effects of multiple cracks on the fatigue crack growth and residual strength of curved fuselage panels.
0 Comments
Leave a Reply. |
AuthorWrite something about yourself. No need to be fancy, just an overview. ArchivesCategories |