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.

1. CASE STUDY
Failure of a Low-Pressure
Turbine Rotor (LPTR) Blade
Dr. Muhammad Ali Siddiqui

2. Content
 Background with literature 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
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3. Background and Literature Review
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4. • Fracture: low-pressure turbine blade  causing extensive
damage to the Jet engine.
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Source: https://www.wikiwand.com/en/Turbine_blade

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Ni

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Ni
Al
Ni Al
Thermodynamic favour ordered structure

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70% of intermetallic make it suitable for
applications

17. Other Alloying
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Source: https://www.semanticscholar.org/paper/High-temperature-materials-for-aerospace-Ni-based-Perrut-
Caron/9fa926f08f061dc68b29ebe7b4bfdfb7a60c56bf
https://doi.org/10.1016/j.crhy.2018.10.002

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Investment (lost wax) casting
Materials 2019, 12(11), 1781; https://doi.org/10.3390/ma12111781

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ceramic topcoat is
deposited by 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

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First stage (the stage 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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29. Case Study
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30.  LPTR blade failed
 Ni-base superalloys
 causing extensive damage to the engine.
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Background

31. Visual Examination
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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
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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

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Intergranular = 1 mm from
the leading edge
Transgranular = 11 mm
SEM and Fractography
Fig: Intergranular fracture at the
leading edge
Fig: Beach Marks
Fig: striations

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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.
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36. • Therefore, it appears 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.
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Cont…

37. 2. Cause of Failure:
• 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.
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38. Conclusion
• The cracking of 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.
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39. Thank You
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