| High pressure compressor blades are the core component of turbofan engine. Its fatigue reliability has a great influence on engine performance and flight safety. Fatigue fracture is the main failure mode of blade during the engine service period. If the blades occurred fatigue fracture in the engine working progress, the engine performance would be affect severely and even the high speed falling debris posed a serious threat to the safety of compressor rotor and cartridge receiver. Therefore it has great significance to research the problem about fatigue crack initiation and propagation of high pressure compressor blades for improving engine reliability. The issue of blades fatigue life prediction goes through the whole life period in individual stage of designing, building, manufacturing, operation and maintenance. Hence it has an important scientific significance and engineering application value to build an accurate fatigue life prediction model of high pressure compressor blades in condition of considering the complex engine working process. In this study, combined with fatigue test and numerical simulation,fatigue life of crack initiation and propagation was predicted for compressor blades. The work can be summarized as follow:In the study vibration fatigue tests including constant loading test, variable loading test and crack propagation test for Ti–6Al–4V titanium alloy of blade imitation specimen were conducted in order to obtain the damage accumulate and crack propagation parameters of blade material and analyze the effect of stress difference and cycle ratio on blade fatigue life under the variable loading. According to the results of crack propagation tests, crack initiation and propagation rules for blade simulated specimen were studied and microscopic feature of fatigue fracture was observed by super depth of field microscope and scanning electron microscope(SEM). Furthermore, the influence of stress raito, overloads effect and combined high and low loading on crack propagation rate were analyzed.Secondly, based on the small crack theory of Haddad, a collective model of fatigue life was established, the model combined Chaboche’s damage accumulation model with Walker’s crack propagation model to predict fatigue life of crack initiation and growth. Moreover this model consider the influence of different stress ratio, variable loading and over loading in fact working condition with complex loading for compressor blades.Thirdly, damage identification of blades is based on vibration analysis. However the damage degree is the main foundation for predicting residue life. The geometric model and finite element model of blade were established to analyze the vibration performance of static and dynamic frequency for undamaged complete blades as well as the natural and forced vibration performance for damaged blades with crack. Then the problems of frequency veering and mode shape switching caused by crack parameters including crack length and location. Based on the analysis results of vibration performance for crack blades and RBF neural network theory, parameter identification of crack blades was studied.Finally, according to the typical working conditions in the actual engine operation, aircraft engine running state was simplified to five typical work cycle and the aircraft engine loading spectrum in 1000 hours was drawn. The stress processes of high pressure compressor blades were computed on condition of considering centrifugal loading, stable aerodynamic loading and vibrational loading. Based on the collective model of predicting the fatigue of crack initiation and propagation, fatigue life of complete blades and damaged blades with crack were predicted. |
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