1. FACTUAL INFORMATION
1.1. Examination brief
During a climb to a new cruise level with maximum climb thrust
set, the flight crew of a scheduled passenger service Boeing
747-338 aircraft noted an increase in vibration from the number-3
engine, accompanied by warning indications from the engine
monitoring instrumentation. The engine was shut down and the flight
continued to its destination.
Subsequent engineering examination of the RB211-524D2 engine
(serial number 12682) revealed the failure of a single blade from
the low-pressure turbine second stage (LPT-2, figure 1). Sections
of the failed blade, including the root block, and portions of the
separated airfoil section were recovered from the engine and
forwarded to the Australian Transport Safety Bureau (ATSB) for
examination (figure 2).
Figure 1: RB211-524 engine profile showing location of
LPT-2 blades (coloured)
Figure 2: Blade fragments recovered from the failed
engine
1.2. Visual characterisation
The part number LK83851 stage-2 low-pressure turbine blade
(serial number TA6939B) had fractured through the base of the
airfoil section; the fracture path extending diagonally from just
above the dovetail transition at the rear face, to a position
approximately 33 mm above the blade base, measured along the
forward face (figure 3).
Figure 3: LPT-2 blade root section showing the fracture
profile and its point of origin (arrowed)
The entire fracture was confined to the region beneath the blade
platform and airfoil section transition. The fracture surface
morphology confirmed a progressive fatigue cracking mechanism
(figure 4), with optical and electron microscopy identifying a
single origin approximately 790 µm from the rear face of the root
block (figures 5 - 7).
Figure 4: Plan view of the blade fracture showing the
fatigue cracking morphology and point of initiation
Figures 5 - 7: Blade fracture at increasing
magnifications with origin identified
Other than mechanical damage attributable to the loss of the outer
blade section, the blade root block showed no evidence of external
tool marks, handling damage or other physical features that could
be held as contributory to fatigue initiation.\
The blade root block carried the embossed and engraved
identification:
Trailing edge face: C2179 1H1153 < 22>
Leading edge face: LK83851 TA6939B < 9> < 13>
< 29> < 201>
The engine manufacturer confirmed that the numbers represented
as <xx> above indicated the application of various repair
schemes during prior blade overhaul/s. On review, none of those
schemes contained actions involving activity at or in the vicinity
of the fatigue origin. The manufacturer also indicated that the
engraved (vibro-etched) markings '< 201>' had been placed
closer to the shank edge than allowed by the particular repair
scheme, and that markings too close to section corners have the
potential to adversely affect the blade integrity. In this instance
however, the cracking showed no association with the < 201>
marking, or any other similar feature.
1.3. Metallography & micro-analysis
A serial grind-polish-examine process was used to investigate
and characterise the features at and surrounding the fatigue
origin. The plane of grinding was parallel to the trailing edge
root block face and a total of four metallographic planes were
examined through the origin area. A fractured and intrusive dark
oxide-like material inclusion was observed during the second stage
of examination (figure 8).
Figure 8 and 9: Metallographic cross-section through
origin and EDS spectra of inclusion present at arrow
Energy dispersive x-ray analysis (EDS, figure 9) of the
inclusion identified a range of elements from the typical base
metal chemistry (Ni, Nb, Ti), as well as zirconium (Zr) which was
foreign to the base metal alloy, but later identified as a
component of the refractory moulding material used during the
investment casting of the turbine blades. Dimensionally, the
visible sections of the inclusion measured approximately 40 µm in
the major dimension, with the engine manufacturer indicating that
the maximum root defect dimension allowable by the relevant quality
acceptance standards (QAS) was in the order of 508 µm (0.2").
The general blade microstructure was typical of a cast
nickel-based high temperature alloy, with massive semi-script form
carbides within a solid-solution (?, gamma) matrix.
1.4. Chemical analysis
Spectrographic analysis techniques identified the base LPT blade
alloy as an IN713LC nickel-chromium-aluminium-molybdenum material.
The engine manufacturer confirmed this as the intended LPT blade
alloy.
1.5. Fatigue loading and operating conditions
The initiation of fatigue cracking within any component is an
indication that the magnitude and number of local stress cycles
sustained have exceeded the limits of the material at the crack
origin. Components such as the LPT-2 blades that do not have a
prescribed fatigue life limit are engineered such that the nominal
and expected transient design stresses and stress cycles are kept
below the threshold for fatigue initiation. Where fatigue cracking
does initiate in these items, it is, by implication, an indication
that either the component has been subject to general stresses
above the design limits, or that some anomaly has acted to locally
concentrate the stresses in the affected area. The presence of
notches, tool marks, physical defects, inclusions or
microstructural abnormalities can all lead to local stress
concentration effects, thus providing conditions suitable for
fatigue crack initiation.
In the analysis of previous LPT-2 blade failures, the engine
manufacturer had implicated uneven fuel distribution issues in the
excitation of blade vibration sufficient to produce high-cycle
fatigue cracking. In this instance however, there was no reported
evidence from the engine examination to suggest that fuel
distribution or combustion conditions had been problematic, or that
those conditions had contributed to the blade failure.
2. ANALYSIS
The LPT-2 blade fracture morphology indicated that failure had
resulted from a progressive high-cycle fatigue (HCF) mechanism,
initiated from a surface or near-surface inclusion at the rear
trailing corner of the blade root section. The inclusion chemistry
and morphology suggested its origin from the refractory mould used
to cast the blade during initial manufacture.
Given that inclusions and other entrained discontinuities are
inherently difficult to avoid in components produced by casting
processes, limits are conventionally placed on the size and
location of such features to maintain the structural integrity of
the items under static and dynamical loading conditions. In this
instance, although the dimensions of the visible portion of the
inclusion were below the manufacturer's allowance for
discontinuities of that nature, it was possible that the defect in
total may have been considerably larger, being broken up and partly
lost during the blade fracture and separation.
3. CONCLUSIONS
The following conclusions were drawn from the examination of the
failed LPT-2 blade from RB211-524D2 engine serial number 12682.
1. Blade failure resulted from the initiation and growth of a
fatigue crack from the rear trailing corner of the blade root
section.
2. Physical evidence suggested the initiation of cracking from a
casting inclusion formed during blade manufacture.
