ДО ВИЗНАЧЕННЯ УРІВНОВАЖУВАЛЬНОЇ НАВАНТАГИ НА ШПАНГОУТ ОДНОПАЛУБНОГО ФЮЗЕЛЯЖУ

With the optimal design of the fuselage, a very important issue is the choice of the optimal position of the load-bearing floor in the cross-section of the fuselage.Depending on the relative position of the load-bearing floor, the reduced thickness of the floor, the scheme of fastening the floor to the frames and the ratio of the reduced thicknesses of the fuselage skin and the floor, the position of the center of stiffness of the fuselage cross-section changes, the torsional stiffness of the fuselage. This leads to a change in torque, a redistribution of shear flows, a redistribution of flattening loads on the frame from the bending of the fuselage.In this work, two schemes of fastening the floor to the frame are considered - a rigid, torque connection and a hinged one. In this case, the frame takes up additional load from the floor. The fuselage is considered as a thin-walled rod, loaded with horizontal and vertical shear forces, torque and flattening forces from the fuselage bending.For reliability, the calculation of the position of the center of stiffness in a double-closed cross-section was carried out by two methods: a fictitious force and a fictitious moment. The influence of various parameters on the location of the center of rigidity was investigated. The influence of the vertical position of the floor, the ratio of the reduced thicknesses of the floor and the fuselage skin and the cross-sectional area of the beams of the floor attachment to the fuselage on the position of the center of stiffness was evaluated. Diagrams of these dependencies were constructed based on the results of calculations. The dependence of the torsional stiffness on the position of the floor and the ratio of the reduced thicknesses of the floor and the fuselage skin was investigated. Based on the calculation results, a diagram of these dependencies was built. Various constructive solutions were considered for fastening the floor to the fuselage skin: with their direct connection and with the floor support only on the beam. The floor loading from flattening loads caused by the bending of the fuselage was studied. The diagram of the loading of the frame and the floor from flattening loads is shown.According to the diagrams, you can choose the optimal vertical position of the floor, the reduced floor thickness and the cross-sectional area of the beam

Design and Analysis of Composite Window Frame for a Regional Aircraft

This work presents the design and analysis of a thermoplastic composite window frame for integration into a regional aircraft. The main parameters which are investigated include buckling, damage and failure loads of a composite window frame subjected to shear loads repesentative of fuselage skin stress distribution due to flight loads. The attachment of such thermoplastic window frame to a thermoset fuselage skin was investigated including both adhesively bonded interface as well as riveting. Even though the bonded frame did meet the design criteria, its failure was very sudden, and the riveted assembly showed a considerably higher strength and structural integrity. The numerical simulation resulted in failure loads which matched very closely to experimental results.

Reliability Evaluation of Fatigue Crack Growth Rate of Heat-Treated TIG-Welded Al 6013-t4 by Two-Parameter Weibull

The limited data of fatigue crack growth (FCG) may cause an inaccuracy assessment of the fatigue crack growth rate (FCGR). For particular parts in aircraft such as fuselage skin, a high-reliability degree due to FCG must be determined accurately for the design and safety requirements. Generally, the 6xxx series of aluminum alloy is used as the material for the fuselage skin in the aircraft. In this study, reliability evaluation of FCGR of heat-treated TIG-welded Al 6013-t4 was investigated by two-parameter Weibull. The FCG tests were conducted by following the ASTM E647 under three different artificial aging time conditions of 6, 18, and 24 hours. The C and m constant values were obtained by drawing the regression line from FCG data following Paris’s equation and analyzed employing three methods; the least square fitting method (LSFM), a mean value method (MVM), and a probabilistic distribution method (PDM). The result showed that the PDM and MVM showed a better-fitted line to assess the C and m values than LSFM. From the reliability viewpoints, the two-parameter Weibull was proposed to be applied as the PDM. Furthermore, the MCM was successful in evaluating the probabilistic assessment of the FCGR with the 85% confidence interval.

Case Studies of Aircraft Fuselage Cracking

Fatigue plays a significant role in the crack growth of the fuselage skin structures. In addition, the fuselage may suffer also from the corrosion damage, and the wear defects. The proper maintenance and scheduled test intervals can avoid the sudden skin failure. Therefore, the inspection interval has to be shortened. Nevertheless, the young machines may be also suffering from the unexpected skin rupture. The cracks are emanating from the rivets and the holes under cyclic loading. The stress concentration around the notch has an effective role under the effect of cyclic loading. The cracks propagate toward the high stressed area such as the notches or other crack locations. The propagation into a critical crack size is rather fast and causes a sudden aircraft fuselage cracking. Hence, the number of cycles to failure will be decreased dramatically. During the last decades, the fracture toughness, design, and the new alloying element have been enhanced. The previous fuselage failures show that the inspections against the cracking are recommended even after a few thousand of cycles. To prevent the crack extending, the crack arresting is recommended to use around the fuselage.

Effect of repair methods on mechanical behaviors of repair area in composite laminate fuselage skin with different damage size

Composite laminates are widely used in the large civil aircrafts because of their excellent mechanical properties. The maintenance and repair of composite laminates become essential. In this paper, a new adhesive-rivet hybrid repair of composite laminate fuselage skin is presented. For the circular hole damage with the diameter of 90mm and 50mm, the finite element simulation models of adhesive repair and adhesive-rivet hybrid repair were built respectively. Uniform pressure load was applied on these finite element models. The mechanical properties of laminate motherboard, patch and adhesive film for these four models were analyzed. The effects of adhesive repair, adhesive-rivet hybrid repair on mechanical behaviors of repair areas of composite laminate fuselage skins with different damage size were studied. By analyzing the mechanical behaviors of these two repair methods, a suitable repair method can be obtained.