No abstract

For the last few decades, composite materials have been more popular than other conventional metal materials in the aircraft industry. Having better mechanical properties (strength, fatigue life, impact strength, corrosion resistance, etc.) and being lighter than conventional engineering materials, composites have become very important in defense industry as well. In spite of the fact that some of the composite materials such as aramid-based composites have been effectively used in body protection, they have not been so successful in heavy armored vehicles which are generally equipped with different types of add-on armor blocks for protecting against threats. These add-on armors are mostly composed of armor steels and ceramics. This study specifically aims to investigate high-velocity impact behavior of unidirectionally oriented carbon fiber reinforced/epoxy layer sandwiched with armor steel plates that are exposed to kinetic energy projectile. Carbon fibers are normally very brittle to transverse loading direction, contrarily, to its axial tension or compression direction. This is the reason why it is claimed that this high compression strength property of carbon fibers could be used for manufacturing a layer in order to replace ceramics in add-on multilayer composite armor. In order to prove this hypothesis, an experimental analysis has been carried out by performing impact tests on these manufactured add-on armor test samples. Testing was carried out in accordance with the STANAG 4569 level-4 standard. The results indicated that the multilayer carbon fiber reinforced epoxy composite–armor steel hybrid panels can provide level-4 protection with a lower areal density compared to Rolled Homogenous Armor.

The aim of this study was to explore easy technologies for manufacturing of integral composite structures to evaluate the manufacturing concepts. This has been done by determining the buckling and post-buckling behavior of hat-stiffened composite panels under compression loading. A shadow Moiré technique was used to monitor the out-of-plane displacement of the skin panel. The experimental results of hat-stiffened panels for initial buckling and post-buckled response of the panels were compared with numerical results obtained from linear and non-linear finite element methods. It was found to be in reasonably good agreement with each other. The panels showed good post-buckling strength and total failure began with the local buckling of the hat stiffeners.

The scarf-joint technique is one of the latest techniques used for repairing composite aircraft structures. But this technique is mostly used at depot level repairing activities since it requires autoclave and other equipments. This article focuses on scarf joint comprised of vacuum and autoclave precured and co-cured composite patches bonded to autoclave and vacuum precured parent laminates. Autoclave and vacuum cured parent laminates and scarf joints were prepared and exposed to the same temperature and moisture environment for comparison. All specimens were loaded in tension at three temperatures. Interlaminar shear strength (ILSS) tests were also carried out for the parent materials. As noted, the tensile strength and ILSS decrease when the material has been exposed to moisture and tested at elevated temperature. But, no significant difference was reported for either tensile strength or ILSS between autoclave and vacuum cured materials. The room temperature repair efficiencies are reported for single scarf repairs comprised of vacuum co-cured and precured patches. These vacuum cured repair efficiencies were found to be similar to the efficiency of the autoclave precured patch repair. This result supports the feasibility of scarf joint repairs with precured or co-cured patches under vacuum curing conditions in field level facilities. Therefore, repairs with vacuum precured or vacuum co-cured patches requiring less equipment seems to be a serious potential alternative to the composite patch repair requiring autoclave conditions which might be only available at depot level maintenance centers.

For the last three decades, composites have become very preferable materials to be used in the automotive industry, structural parts of aircraft and military systems and spacecraft, due to their high strength and modulus. Composite materials are sometimes exposed to invisible or visible damage due to impact loading during their service life. In this study, the effect of impactor geometry with four different contact surfaces on woven carbon fibre-reinforced composite plates having three different thicknesses are investigated. In the first stage, composite plates were manufactured with the ply orientations of [45/-45/0/90/45/-45]2s, [45/-45/0/90/45/-45]3s, [45/-45/0/90/45/-45]4s based on conventional usage. In the second stage, carbon fibre-reinforced composite test panels were exposed to low velocity impact tests to obtain force-time, energy-time and force-displacement curves. Finally, semi and full penetration of composite panels and damage magnitude were determined. It was found that the impactor geometries with lower contact surfaces such as conical and ogive types were much more penetrative on composite plates than the other geometries, but they caused larger damage area in the vicinity of the impact point.