Sequence of events
At about 1200 Eastern Standard Time on 30 May 2005, a Boeing Co
747-300 (747), registered JA8184, was being pushed back from its
gate at Sydney International Airport for a scheduled passenger
flight to Osaka, Japan. During pushback, the ground staff heard a
loud cracking noise. The pushback was stopped and an inspection by
the ground crew identified a structural failure in the left wing
landing gear forward trunnion fork (trunnion), as shown in Figure 1
and 2. After an on-site inspection by the Australian Transport
Safety Bureau (ATSB), the aircraft was moved to a hangar for
maintenance and the fractured component was removed from the
aircraft and sent to the ATSB for a detailed examination.
Figure 1 : Left wing landing gear
Figure 2: Looking up and outboard into wing landing gear
well
Examination of fractured trunnion
The trunnion had sustained a complete through-section fracture,
located approximately mid-way between the ball-end and the fork-end
(Figure 3).
Figure 3 : Fracture location on trunnion
A general inspection of the fractured trunnion revealed a
discoloured (orange/brown) region on the fracture surface (Figure
4) in the upper outboard region. The corroded nature of that
region, compared with the adjacent bright fracture surfaces,
indicated the presence of a pre-existing defect, and that the
trunnion had been cracked for a period of time prior to the final
failure during the pushback.
Figure 4 : Fracture surface
The fractured component was examined in a metallurgical
laboratory under the supervision of the ATSB. Chemical analysis of
a sample taken from the trunnion near the fracture showed that the
material met the specification for AISI/SAE 4340M alloy steel.
Metallographic examination confirmed a fine-grained lightly
tempered martensitic microstructure, typical of the 4340M alloy in
the hardened and tempered condition. The inner and outer surfaces
were also observed to have been finished with a metallic type
plating and painted with a surface primer and top coat.
Hardness measurements taken indicated that the material had an
ultimate tensile strength of approximately 275,600 psi (1900
MPa).
Wall thickness measurements taken around the fractured trunnion
circumference showed that the minimum local wall thickness of 0.137
inches (3.48 mm) corresponded with the corroded and discoloured
area of cracking.
A detailed technical examination of the corroded region
identified two transverse fatigue cracks originating at the inner
surface of the trunnion bore (noted as C1 and C2 in Figure 5). The
cracks had initiated approximately 11mm apart and had joined to
form a single crack. This crack continued to grow until the final
fracture occurred during the pushback.
Figure 5 : Fatigue crack development
Crack C2 presented well defined fatigue progression marks,
including several distinct regions of fatigue and corrosion (Figure
6). The region bounded by the red dotted line was a distinct region
of heavily corroded fatigue cracking and was about 4 mm long and
1.6 mm deep.
Figure 6 : Fatigue progression and corrosion
marks
Close examination of the origins in the inner wall of the
trunnion found that crack C1 had initiated from multiple closely
spaced origins at the root of a machining groove, giving the
appearance of a longer crack following the machining groove (or
mark). Crack C2 had also originated at the root of a machining
groove from multiple closely spaced origins, but over a much
smaller distance before aligning with the principal stress
plan1, resulting in the apparent
difference in the planes of the cracks as shown in Figure 7. The
plane of the initial cracks in both C1 and C2 were approximately
parallel and were aligned with the machining grooves in the inner
surface.
Figure 7 : Plane of crack origins
The inner and outer surfaces did not have a consistent surface
roughness. Well defined machining marks were observed in the large
diameter bore and to a lesser extent on the small bore. However,
the outer surface and the taper region in the bore were relatively
smooth without defined machining marks. The well defined machining
marks on the large and small diameter bore blended into the
smoother surface of the taper region (that is, there was no abrupt
change in surface roughness).
Examination of the surface microstructure in the region of the
cracks revealed that the smooth regions (outer surface and taper
section) had a thin layer of deformed material typical of a cold
working process such as shot peening2. The area of surface
deformation in the taper region ran out just before the radius (a
few millimetres from the cracks). Figure 8 shows the differences in
the surface roughness at a microscopic level (scale is 25µm, or
0.025 mm).
Figure 8 : Smooth surface (upper); surface with distinct
machining marks (lower)
The sections taken from adjacent to the primary cracks for
micrographic examination contained multiple independent fatigue
cracks of various sizes. One example is shown in Figure 9. Each of
these cracks originated in the root of the machining groove and
were associated with shallow intergranular penetrations, which also
existed in the roots of the machining grooves (Figure 8).
Figure 9 : Secondary fatigue crack indicated by
arrow
Component manufacture
The landing gear trunnion was manufactured to the aircraft
manufacturer's specifications by an approved external supplier.
Both the supplier and the aircraft manufacturer informed the ATSB
that the trunnion was manufactured at some time prior to 1975;
however the original manufacture documentation (including the
manufacture plan and conformance records) was destroyed in 1994.
The trunnion specifications were supplied by the aircraft
manufacturer. Those documents included construction drawings and
process specifications.
In the trunnion specifications, the aircraft manufacturer
specified the use of 4340M steel, heat treated to an ultimate
tensile strength of 275,000 to 300,000 psi. Therefore, the material
used in the manufacture of the failed component met the steel alloy
and strength requirements of the design.
The minimum allowable wall thickness specified3 for the trunnion at
the location of the failure was 0.180 inches (4.57 mm). Therefore,
the minimum wall thickness measured at the crack of 0.137 inches
(3.48 mm) was 0.043 inches (1.09 mm) thinner than the design
allowed.
Component maintenance
The maintenance documents supplied by the aircraft operator
indicated that the trunnion had been fitted to five aircraft and
had amassed a total of 25,095 landing cycles during its service
life. The records also showed that the trunnion had been overhauled
by the operator's component repair workshop on four occasions
(Table 1). The landing gear assembly had an overhaul interval of 8
years or 12,000 cycles, whichever occurred first. Therefore, the
landing gear was not due for overhaul for another 4 years, or 9,509
cycles.
Table 1 : Overhaul history
Overhaul |
Date | Total cycles at overhaul |
|
1 |
October 1979 |
3,517 |
|
2 |
November 1987 |
14,069 |
|
3 |
March 1992 |
16,508 |
|
4 |
October 2001 |
22,604 |
The overhaul records showed that on each occasion, the trunnion
had undergone repair work. The documents indicated that the repairs
were limited to the lugs and were within the repair limits
permissible by the aircraft manufacturer. There was no record of
any repair work carried out in the bore of the trunnion or on the
outer surface in the region of the bore diameter transition.
Comparison of the overhaul instructions maintained by the
operator with those supplied by the aircraft manufacturer confirmed
that the operator's workshop maintained the correct instructions
for the overhaul. The overhaul instructions for the component did
not require a dimensional check for wall thickness or a specific
check for surface finish (roughness) in the bore.
As one of the first processes in the overhaul, the protective
finishes (including metallic plating) were removed from the surface
of the trunnion. These finishes are removed using a chemical
process, some of which, including water for cleaning, can have a
corrosive effect on the trunnion material.
The overhaul procedure for the landing gear components required
that the component undergo a magnetic particle inspection (MPI) to
detect any defects, including cracks, that may have developed
during service. The manufacturer did not provide specific
instructions on how to perform the MPI on this particular
component, but provided a general process that the operator was to
use as the basis for a component specific process. Neither the
overhaul procedure nor the MPI process directed the MPI
technician's attention to the radius in the bore diameter
transition as a possible location for cracking. The MPI process
used was capable of highlighting cracks of less than 0.5mm.
The overhaul records provided by the operator indicated that an
MPI of the component was carried out at each overhaul. The MPI
procedure defined by the operator was in accordance with the
aircraft manufacturer's recommended procedure. It used the
equipment and materials recommended by the manufacturer and
specified sufficient examinations to highlight any cracks in the
component. The overhaul records indicated that the item had been
found satisfactory on each occasion, suggesting that no cracks had
been detected.
The component maintenance manual for the repair of high-strength
steel landing gear parts directed the operator to obtain advice
from the manufacturer if cracks were detected during the MPI. The
manufacturer did not have a record of any request for advice
relating to the failed component.
The crack was in a location that was not readily viewable during
a normal visual ground check. There was no requirement to perform a
detailed inspection for cracks in the body of the trunnion between
overhauls.
- The principal stress plane is a
plane that the stress in the part acts perpendicular to. In this
case, the principal stress plane was not aligned with the machining
marks. - Shot peening is a process where
small 'shot' beads are fired against the surface of a component
producing a residual compressive surface stress and a thin layer of
deformed material. This process has been demonstrated to increase
the fatigue life of high-strength steel components. - In the manufacture drawings for the
component.