In aeronautics, sandwich structures are widely used for secondary structures like flaps or landing gear doors. In the case of landing gear doors, the junction is made by a local reinforcement called an insert. This insert is made by a resin molded in the Nomex™ sandwich core. Such structures are still designed mainly using test results and the lack of an efficient numerical model remains a problem. The purpose of this study is on the one hand to perform experiments in order to be able to identify the failure modes and on the other hand to propose an efficient numerical model. Pull-out tests with cycling were conducted and 3D displacement measured by optical methods. The potential failure modes are numerous (delamination, local fiber breaking, skin/core debonding, core crushing, core shear buckling, potting failure, etc.). Experiments demonstrated that, for the lower loads, the non-linearity and the hysteresis are mainly due to core shear buckling. From this observation, the nonlinear behavior of the core is identified by a 3 point-bending test. The shear-modulus damage law is then implemented on a non-linear finite element model and an acceptable correlation of the tests is achieved. As a consequence, some improvements of the technology will be proposed, manufactured and tested.

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Sandwich structures are widely used in the aerospace field, also for primary parts. However, due to the low strength of core properties, the failure behavior under high stress concentration such as joining position is hard to evaluate. This study is, at present, a key task to enable future exploitation of the joint for structural sandwich consisted of Carbon Fiber Reinforced Polymer (CFRP) face and a Honeycomb core made of phenolic impregnated NomexTM paper. Previous experimental investigations provided the failure mechanism of sandwich plates with hard points in the form of inserts, and special attention is focused on the problem of sandwich panels with inserts of the fully potted types. Numerical simulations are achieved in this work with good correlation based on experimental test.

The connection concept of placing only two bolts in offset misalignment against the bending load along the wing span was used for an aerobatic airplane designed in Thailand as a KIT plane to minimize the impact of drilling numerous holes. This concept can deviate the force direction on the holes. The two suitable drilling positions should have the lowest resultant force and highest strength of fiber reinforcement structure. To investigate the strength of the fastener hole when the force deviated from the original orientation, specimens, made from twill weave carbon fiberepoxy and laid up at ±45-degree orientation, were tested under bearing load according to ASTM D5961 standard. The experimental results revealed that the bearing strength of CFRP material decreases when the force deviation angle increases, so zero-angle deviation of the resultant force on the drilling hole is the most suitable orientation to absorb bolt bearing load. The most suitable pattern of two offset misalignment holes is the greatest horizontal distance at zero vertical distance when it was considered only the effect of the bearing strength and the deviation angle. Moreover, the failure pattern begins to deviate along the fiber orientation when the inclination angle increased.

A bending test method was designed to investigate the failure characteristics of drilling carbon-fiber reinforced composite (CFRP) plates. The CFRP plates were fabricated and drilled to join other parts for use in aircraft structures (riveting or bolting). Each sample had two equal-sized holes at 0, θ, or -θ degrees at a fixed distance of 36 mm in the longitudinal direction. The sample dimensions were identified following a study on tensile testing singleholed samples based on the ASTM D5961 standard. The tensile test focused on three cases of laminate: [0]14, [22.5]14, and [45]14. The maximum mean forces for [0]14, [22.5]14, and [45]14 laminate configurations were 6.32, 6.59, and 7.12 kN respectively. These results were implemented into a bending test which included two configurations of laminate: [0]14 and [45]14 and two equal-sized holes with a variable distance of 0 and 9 mm in the transverse direction of the plate. The bending test configuration is presented in this paper. The results revealed different load-bearing capabilities in each of the three types of specimen configuration. Samples with two holes lying at ±9 degree affected the overall strength in the [45]14 case, and as in [0]14 there was no difference between the eccentricity and the aligned holes. In [45]14, the eccentricity decreased the maximum allowable force by 1.57 percent in the normal eccentricity case and by 15.94 percent in the alternate eccentricity case compared to the aligned case.