This paper presents the impact of selecting the fatigue material model on the numerically determined fatigue life of a compressor blade. In the work, the first stage compressor blade of the PZL-10W turbine engine was used. The research object contained a geometric notch with a known location and shape. In numerical studies, 8 fatigue estimation methods were used in the ε-N analysis (based on the Manson-Coffin-Basquin model). At the same time, three methods for estimating material constants associated with the cyclic hardening were employed. On the basis of the selected models, 24 sets of fatigue parameters were obtained, which were used in numerical studies. The numerical tests were carried out under resonant conditions with amplitudes of 1.5 and 1.8 mm. The numerical tests were confirmed by the experimental fatigue tests. As a result of the above-mentioned tests, the impact of selecting the material fatigue model and hardening model on the obtained results was determined and they were referred to the initiation of the crack with the length a = 0.2 mm (achieved during experimental studies). The obtained results will constitute the basis for further fatigue tests.

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This work presents the results of the numerical analyses pertaining to the influence of resonance vibration amplitude on the fatigue life of a compressor blade with a defect made by a collision with a hard object (FOD). The research object was the first stage compressor blade of the PZL-10W engine. The numerical simulation of the notch formation was performed for the tested blade. The material fatigue models (for e-N analysis), three cyclic hardening models, and two mean stress correction models were used in the numerical analyses. As a result of the numerical analysis, the information on the distribution of principal stress was obtained. The values of the principal stresses were used for numerical e-N fatigue analysis using the aforementioned models of fatigue, hardening, and mean stress correction. Obtained results were compared to previously published experimental research, where a notch was created at the leading edge in 8 blades. The blades damaged under laboratory conditions were subjected to experimental fatigue tests during which the effect of resonance amplitude on the number of damage cycles was determined. As a result of the comparison work carried out, the impact of the vibration amplitude on the durability of the element with plastic deformation was determined.

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The work presents the results of numerical fatigue analysis of a turbine engine compressor blade, taking into account the values of initial stresses resulting from surface treatment-shot-peening. The values of the residual stresses were estimated experimentally using X-ray diffraction. The paper specifies the values of the residual stresses on both sides of the blade and their reduction due to cutting through the blade-relaxation. The obtained values of the residual stresses were used as initial stresses in the numerical fatigue analysis of the damaged compressor blade, which was subjected to resonant vibrations of known amplitude. Numerical fatigue ε-N life analysis was based on several fatigue material models: Manson’s, Mitchell’s, Baumel-Seeger’s, Muralidharan-Manson’s, Ong’s, Roessle-Fatemi’s, and Median’s, and also on the three models of cyclic hardening: Manson’s, Xianxin’s, and Fatemi’s. Because of this approach, it was possible to determine the relationship between the selection of the fatigue material ε-N model and the cyclic hardening model on the results of the numerical fatigue analysis. Additionally, the calculated results were compared with the results of experimental research, which allowed for a substantive evaluation of the obtained results. These results are of great scientific and practical importance. The problem of determining the fatigue life of blades with defects operating under resonance vibrations is one of the original tasks in the field of fracture mechanics and experimental mechanics. The results obtained are of great importance in the aviation industry and can be used during engine maintenance and inspections to assess the suitability of blades with defects in terms of the needs of further work. This aspect of engineering maintenance is of great importance from the aircraft safety point of view.