Jonker, A.S.Boshoff, Rudolph2016-10-252016-10-252015http://hdl.handle.net/10394/19141MIng (Mechanical Engineering), North-West University, Potchefstroom Campus, 2016During flight, aircraft are subjected to dynamic flight as well as a dynamic environment, resulting in variable stresses in the wings and other structural components. These variable stresses cause cumulative material damage known as fatigue. This document contains the experimental process of developing a Constant Life Diagram (CLD) which can be used to predict the fatigue life of glass fibre samples subjected to axial fatigue stress. The properties of composite materials such as those used to produce the body and wings of the JS1 sailplane (aramid, carbon and glass fibre) are highly dependent on the manufacturing method, layup sequence and orientation, the weaving pattern of the raw fibre, fibre orientation, resin to fibre ratio as well as various other factors. When one or more of these aspects are altered, the material, mechanical and fatigue properties change. A combination of all of these aspects constitutes a new unique material as such. Because of the complex nature of composite materials as well as the fact that composite fatigue is a relatively new field of study it is not yet possible to approximate the fatigue life of composite materials mathematically with correction factors as is done for metallic materials. Instead a unique CLD needs to be constructed based on the experimental results for each unique variant1 of the material. The engineering problem at hand is that sailplane designers lack the predictive tool to design against fatigue. To deal with this problem, a number of tests were done on the specific materials used at Jonker Sailplanes Co. to construct the wings of the JS1 sailplane: glass fibre 92125 bonded with MGS287 epoxy, four (4) layers laid up in a (0°/0°/0°/0°) sequence. Four (4) load ratios were tested, spanning the compression-compression, compression-tension and tension-tension loading spectra. Load ratios 𝑅=0.1 in tension-tension and its compressive mirror 𝑅=10 were used along with 𝑅=−1 wherein the magnitude of the compressive and tensile stresses is equal. The last load ratio which was used to construct the CLD is the ratio 𝑅=−0.55. The material does not react the same under tensile and compressive loading, with the material being more resilient to tensile loading than to compression. The load ratio at 𝑅=−0.55 indicates the situation wherein the tensile and compressive stress applied to the material are an equal percentage of the material’s ultimate strength. During testing the applied stress was gradually decreased (for each load ratio separately) resulting in a logarithmic relation between the applied stress and fatigue life. The data can in turn be used to predict what the fatigue life will be at a given stress or vice versa what stress can be applied if a designer wants the component to have a specified fatigue life. Data was processed using confidence intervals so that the final CLD makes a conservative prediction. An estimation of the optimistic prediction is also available. The CLD, in its nature, enables predictive value even for the load ratios that were not tested, which makes it a very practical tool to predict the fatigue life of the specific composite material it was constructed for. Because aircraft are subjected to variable, repeated, alternating, and fluctuating stresses during flight, it is important to be able to predict the fatigue life to enable better design parameters instead of blindly increasing the safety factor. The CLD constructed in this study provides such a predictive tool.enThesis