An investigation was made to determine the causes of surface contact fatigue failure of a case hardened driver pinion located in the intermediate shaft of a reducer gearbox used in a sugar and alcohol mill. The examination of the component revealed the presence of a cemented layer substantially thicker than that generally specified for pinions devised for this application. This, associated with the massive presence of brittle threadlike carbon-rich cementite phase (Fe3C) in prior austenite grain boundaries of the pinion teeth, favored surface crack nucleation and propagation during cyclic loading, leading to spallation of the contact surface with the counterpart gear, which impaired the system's operation. Poor carburization practice was discovered as the root cause of the mechanical failure, thus demanding the implementation of a new manufacturing route to avoid problems in similar load-bearing rotating components.

Abstract: Reinforced thermoplastic structural detail parts and assemblies are being developed to be included in current aeronautic programs. Thermoplastic composite technology intends to achieve improved properties and low cost processes. Welding of detail parts permits to obtain assemblies with weight reduction and cost saving. Currently, joining composite materials is a matter of intense research because traditional joining technologies are not directly transferable to composite structures. Fusion bonding and the use of thermoplastic as hot melt adhesives offer an alternative to mechanical fastening and thermosetting adhesive bonding. Fusion bonding technology, which originated from the thermoplastic polymer industry, has gained a new interest with the introduction of thermoplastic matrix composites, which are currently regarded as candidate for primary aircraft structures. This paper reviewed the state of the art of the welding technologies devised to aerospace industry, including the fields that Universidade Estadual Paulista Júlio de Mesquita Filho and Universidade de São Paulo are deeply involved.

A weight-based analysis was made of the translaminar Charpy impact toughness performance of conventional and advanced composite materials for aircraft fabrication. The materials were carbon-epoxy (C-Ep) and hybrid fiber-metal TiGr (Titanium-Graphite) laminates. 5 mm-thick three-point bend specimens were tested over a temperature range of -70 to 180 ºC to reproduce typical in-service conditions of supersonic jetliners. The energies required for the processes of damage initiation (Ei), damage propagation (Ep), and whole fracture (Et = Ei + Ep), were evaluated at two loading rates, namely, 2.25 and 5.52 m/s in an instrumented Charpy impact testing machine. C-Ep laminates with unidirectional fiber tapes arranged in cross-ply architecture consistently showed the best performance in terms of damage initiation toughness, whereas the hybrid fiber-metal laminate TiGr excelled in terms of propagation toughness. On the other hand, the overall performance of bi-directional fabric C-Ep laminates was very disappointing. The impact behavior of composite laminates was substantiated by a qualitative analysis of topographic aspects of fracture surfaces.

In Part II of this work, quasi-static tensile properties of four aeronautical grade carbon-epoxy composite laminates, in both the as-received and pre-fatigued states, have been determined and compared. Quasi-static mechanical properties assessed were tensile strength and stiffness, tenacity (toughness) at the maximum load and for a 50% load drop-off. In general, as-molded unidirectional cross-ply carbon fiber (tape) reinforcements impregnated with either standard or rubber-toughened epoxy resin exhibited the maximum performance. The materials also displayed a significant tenacification (toughening) after exposed to cyclic loading, resulting from the increased stress (the so-called wear-in phenomenon) and/or strain at the maximum load capacity of the specimens. With no exceptions, two-dimensional woven textile (fabric) pre-forms fractured catastrophically under identical cyclic loading conditions imposed to the fiber tape architecture, thus preventing their residual properties from being determined.

In Part I of this study, quasi-static and impact bending properties of four aeronautical grade carbon-epoxy laminates have been determined and compared. Materials tested were unidirectional cross-ply (tape) and bidirectional woven textile (fabric) carbon fiber lay-up architectures, impregnated with standard and rubber-toughened resins, respectively, giving rise to 1.5 mm-thick laminates. Quasi-static mechanical properties assessed in transversal mode loading were modulus of elasticity, flexural strength and tenacity at the maximum load, whereas the net absorbed energy was determined under translaminar impact conditions. Two-dimensional woven carbon fiber reinforcements embedded in a rubber-toughened matrix presented the best mechanical performance under static loading. Under dynamic loading conditions, woven fiber fabric pre-forms were favorably sensitive to increasing impact energies regardless the nature of the employed epoxy resin. However, it was concluded that great care should be taken with this material within the low energy impact regimen.