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Case study: Failure Analysis of LPTR blade
The document summarizes a case study of the failure of a low-pressure turbine rotor (LPTR) blade in a jet engine. Visual examination found a fatigue crack had initiated at the leading edge of the blade and propagated about 1 mm intergranularly and then about 11 mm transgranularly. Testing showed the microstructure at the leading edge had experienced dissolution of gamma prime precipitates, reducing the hardness and creep resistance. The failure mechanism involved initial stress rupture cracking at the leading edge followed by fatigue crack propagation. High operating temperatures caused the failure by dissolving precipitates and degrading the blade's strength properties.
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Case study: Failure Analysis of LPTR blade
- 1. CASE STUDY Failure ofa Low-Pressure Turbine Rotor (LPTR) Blade Dr. Muhammad Ali Siddiqui
- 2. Content Background withliterature review about super alloys, properties, alloy design and manufacturing process Testing Procedure and Results Visual Examination of General Physical Features Scanning Electron Microscopy and Fractography Metallography and Hardness Discussion Conclusion 2
- 3. Background and LiteratureReview 3
- 4. • Fracture: low-pressureturbine blade causing extensive damage to the Jet engine. 4
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- 14. 14 Ni Al Ni Al Thermodynamic favourordered structure
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- 16. 16 70% of intermetallicmake it suitable for applications
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- 22. 22 Investment (lost wax)casting Materials 2019, 12(11), 1781; https://doi.org/10.3390/ma12111781
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- 25. 26 ceramic topcoat is depositedby electron beam physical vapor deposition (EBPVD) or air plasma- spraying (APS) Sandwiched between the topcoat and the metallic bond coat is the thermally grown oxide (TGO) Surface treated: THERMAL-BARRIER COATINGS (TBCs) Examples: Aluminide coatings
- 26. 27 First stage (thestage directly following the combustor) of a modern gas turbine faces temperatures around 1,370 °C Modern military jet engines, like the Snecma M88, can see turbine temperatures of 1,590 °C. Cross sectional view
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- 30. LPTR bladefailed Ni-base superalloys causing extensive damage to the engine. 31 Background
- 31. Visual Examination 32 Observation: •Fractured =Airfoil section at a distance of about 25 mm from the blade root platform •The fracture surface was flat and perpendicular to the blade axis. Fig : Photograph of the failed LPTR blade
- 32. Low magnification Examination 33 Fig: Entire fracture surface Observation: Discolored due to oxidation and exposure to high temperatures well-defined crescent-shaped area with smooth fracture features Fatigue features Fatigue crack found at leading edge. Remaining fracture surface had a crystalline appearance
- 33. 34 Intergranular = 1mm from the leading edge Transgranular = 11 mm SEM and Fractography Fig: Intergranular fracture at the leading edge Fig: Beach Marks Fig: striations
- 34. 35 Metallography and Hardness Fig:Microstructure at leading edge Fig: Microstructure at midcord Observation: precipitates dissolution at the leading edge. At the leading edge 350 HV At the mid-chord section 400 HV
- 35. Discussion 1. Failure Mechanism: •Intergranular mode (IM) to about 1 mm from the leading edge • Followed by a transgranular mode (TM) of about 11 mm – IM Principal mode in high-temperature creep/stress rupture. – Cracking occur principally along the grain boundaries normal to the major stress axis of the blade. – TM The incremental crack growth is evidenced by the presence of closely spaced striations. 36
- 36. • Therefore, itappears that the crack initiated at the leading edge of the blade by stress rupture and propagated fast to a distance of about 1 mm. • In the second stage, this crack acted as a notch for stress concentration and led to the propagation of the crack by fatigue. • The fatigue crack then propagated progressively up to about 11 mm from the leading edge before giving way to overload fracture. 37 Cont…
- 37. 2. Cause ofFailure: • Dissolution of gamma prime (γ-prime) precipitates at the leading edge clearly indicates that the blade was exposed to high temperatures the hardness the creep resistance of the blade drastically. 38
- 38. Conclusion • The crackingof the blade took place by stress rupture. • Once the initial crack had formed due to stress rupture, the crack further propagated under the cyclic loading experienced by the blade during service. • The factors responsible for such cracking are high operating temperatures and stresses that causes the dissolution of the precipitates. 39
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