Abstract

Compression After Impact (CAI) test is normally used for the assessment of damage tolerance of composite laminates. The applicable specification for the execution of these tests is AITM 1-0010. However, for thicknessess under 4mm were buckling occurs, the applicable specification is the AITM 1-0077, which introduces the use of anti-buckling tooling covering the entire surface of the specimens. The tooling, which is used in the aeronautical industry  for the testing of thin specimens presents two problems which are the difficulty of assembly and of obtaining realiable results, and the visibility of the specimen during the test.

This work comprises the design, manufacturing and validation of a novel anti-buckling device for compression after impact thin specimens. For that purpose, a study of the specification toolings along with a review of the already developed anti-buckling tooling literature has been carried. As part of the validation process three series of impacted specimens of 3mm thickness.

This activity is developed as part of the project CERES, which is framed under Clean Sky 2, and has the main objective of validating the manufactured anti-buckling tooling. This tooling has been developed as an innovative solution for its application in compression after impact thin specimens. 

Abstract

This work focuses on analyzing the uniaxial response of carbon fiber reinforced polymer matrix laminates with the configurations [±45]_2S under tension and [±45]_4S under compression. These present non-linear stress-strain evolutions that allow them to withstand large deformations before losing their load-carrying capacity. Both responses are characterized by a first linear stage, followed by a plateau in which the strain grows without increasing the stress level, and finally by a re-stiffening phase. But there are quantitative differences that lead to different failure patterns, which are explained with the help of the state of stress in the plane of maximum shear stress. For a better understanding of the process, tensile load-unload-reload tests are performed to verify if the energy recovered in each cycle could be related to the loss of the apparent stiffness. During the first two stages of the mechanical response, the laminate suffers a progressive damage with a reduction in stiffness related to the dissipated energy, but this pattern is not repeated in the last stage of strain hardening. Based on the experience of other authors, it is assumed that the re-stiffening stage follows a different pattern due to possible microstructural changes in the matrix during the plateau. These are promoted by a narrowing process under tensile loads and a local widening of the test zone under compression. The dimensional changes perpendicular to the load direction are observed thanks to the strain fields obtained by Digital Image Correlation.

Abstract

Primary structures fabricated with carbon/epoxy pre-impregnates has become very popular, especially in aeronautics and automotion. Despite the excellent properties of resistance and stiffness, composite materials present a high sensitivity to impact damage which can cause a sudden drop of mechanical properties affecting the service behavior and reducing the useful life. Here is where damage tolerance concept is introduced.In this work, firstly impact tests have been made in composite lamintes reinforced with carbon fiber, and then, the residual strength for the different stacking sequences have been evaluated using CAI (Compression After Impact) tests.

Abstract

This paper describes the Fatigue & Damage Tolerance Phase for the S19.1 Full Component Test of the A350 XWB Aircraft.   The aim of this test was to demonstrate that the fatigue phenomena does not occurs, as well as no crack growing on composite materials. A350XWB S19.1 Full Scale test is the first test developed on a S19.1 aircraft section, therefore, the process to carry out the test constitute an important step for the development of experimental models which allow for the better understanding of this kind of aircraft component. The process carried out to obtain relevant data for the CFRP F&DT test is described in this paper: Checking point of the structure selection. Specimen instrumentation. Spectra generation. Whiffletrees selection for structure load application. Load optimization for interfaces correlation.   The FCT results shown that the fatigue and damage tolerance Certification requirements are fulfilled. 

Abstract

With thin structures used in the aircrafts, aeronautical manufacturers are worried about the response of these thin laminates towards impact and post impact loads. Thin laminates studied to show reduced compression after impact strengths with thin plies and hence continues to be a primary concern to the aeronautical industries. In this work, we aim to use ply grade hybridization (mixing different ply grades) and unsymmetrical laminate designs as an effort to improve the low velocity impact response and compression after impact strength of thin laminates. Using different materials, uni-directional and non-crimp fabrics, through experimental, numerical and optimization methods, we demonstrate that hybridization along with unsymmetrical designs can improve the CAI strength by upto 40% over the thin ply baseline laminates.