Boeing Co 767-238, VH-EAQ

History of the flight

Approximately eight minutes into a regular passenger transport
flight from Melbourne to Sydney, while the Boeing 767 aircraft was
climbing through flight level 160, the crew and passengers heard a
loud bang and felt severe vibration throughout the airframe. Engine
indication and crew alerting system (EICAS) messages on the flight
deck indicated the left (number-one) engine had no N1 turbine
rotation and an elevated exhaust gas temperature. After
discontinuing the climb and advising air traffic services (ATS),
the flight crew actioned the 'engine fire, severe damage and
separation' checklist and advised the cabin crew and passengers of
the engine failure and the intention to return to Melbourne.
Several aircraft crewmembers that were passengers aboard the flight
advised the flight crew (via the cabin services manager) that the
left engine had lost a fan blade and that it had perforated the
engine cowling. The flight crew made a PAN radio call to ATS and
requested emergency services be placed on local stand-by. After
configuring the aircraft for a single-engine approach and landing,
some adjustment of the airspeed was required to minimise the level
of vibration from the failed engine. The aircraft landed safely on
Melbourne airport runway 27, eighteen minutes after the engine had
failed and twenty-six minutes after departure.

After exiting the runway, the aircraft was stopped and airport
rescue and fire-fighting services carried out a safety inspection
before allowing the aircraft to taxi to the terminal buildings
using thrust from its remaining serviceable engine. Following
passenger disembarkation, the flight crew conducted an operational
debriefing with the cabin crew.

Injuries to persons

Damage to the aircraft

Damage to the aircraft was limited to the left engine assembly
and nacelle. While multiple punctures of the engine nose cowling
indicated the liberation of debris from the confines of the intake
area, none of this debris had struck the wing, fuselage or
tailplane of the aircraft.

Failure of the Pratt & Whitney JT9D-7R4 engine (serial
number P-716610) fitted to the aircraft was attributed directly to
the fracture and release of the outer half of a single low-pressure
compressor (fan) blade (part number 5001341-22, serial number
ND9278).

Liberation of the blade segment caused appreciable damage to the
remaining fan blades and extensive damage to the intake linings.
Ancillary damage to the engine included distortion of the fan
casing, loss of the fan speed (N1) sensor and the overload failure
of several nose-cowl bolts. Although the primary impact of the
released blade with the fan casing had resulted in the segment
being contained, the subsequent forward movement of the blade
allowed it to impact the nose-cowling with sufficient energy to
puncture the cowl wall and escape the engine intake. The initial
impact with the cowl occurred at the two-o'clock position (looking
forward), with the blade segment passing through the cowl with a
tangential trajectory, exiting at around the three-o'clock
position. From the impact point and angle, it was evident that the
blade segment had been ejected downward and beneath the aircraft.
Other debris liberated through the nose cowl or fairings included
the N1 sensor and one of the nose cowl lip bolts. Both components
were located adjacent to the initial blade impact point and thus
were likely to have been subject to a very large reactive force as
the blade segment struck the fan case. Figures one to four
illustrate the trajectory followed by the released blade segment
and the fan case components that perforated the engine cowling.

Aircraft information

Engine information

The subject engine (serial number P-716610) had been installed
on VH-EAQ since October 2001 and had operated for 319 hours and
through 200 cycles while fitted to the aircraft. Pratt &
Whitney first purchased the engine for leasing in 1998 and, since
that time, it had been installed on several different aircraft from
different airlines. At the time of failure, the engine had operated
for a total of 26,138 hours and through approximately 8,900
cycles.

The failed fan blade (part number 5001341-22, serial number
ND9278) was fitted to the engine in August 1998. Before this, the
blade had been held as a stock component since its repair and
refurbishment in 1991. Work done on the blade at that time included
two elevated-temperature straightening operations, where the blade
was heated to 650 degrees Celsius and the aerofoil shape re-formed.
The manufacturer's records indicated a subsequent blade service
life of 7,187 hours and 2,083 cycles. The total time and cycles
accumulated by the blade since manufacture was unknown.

Blade inspection

Various non-destructive inspections had been carried out on the
blade since overhaul, including eddy current inspections after the
thermal straightening operations and periodic visual inspections of
the blade while in operational service. Prior to installation on
VH-EAQ, the engine underwent a foreign object damage inspection
(conducted every 200 cycles) and an eddy current inspection of the
leading edge (conducted by the operator every 350 hours). No
further inspections had been performed or were required at the time
of failure. The requirements and frequency of these on-wing
inspections were specified in the aircraft manufacturer's
maintenance manual (B767-72-31-02/601) and in Pratt & Whitney
service bulletin SB 72-255. At the time of the failure, these
documents contained no requirement to carry out a periodic eddy
current inspection of the blade trailing edges while the engine was
in-service. SB 72-255 stated that 'Eddy current inspection may be
used as an option at the operator's discretion'.

After the 1991 refurbishment work, the manufacturer's records
indicated that the failed blade was inspected to the engine manual
requirements using a single-pass eddy current technique. The eddy
current procedure was specified as having the capability to detect
crack-like defects as shallow as 0.25mm (0.010") along the blade
edges. No defects were detected as a result of this procedure and
the blade was subsequently accepted for service.

Cabin aspects

The cabin services manager (CSM) reported the initial engine
failure event as "like hitting a brick wall; obviously not
turbulence". The CSM described a noisy, high level vibration
throughout the cabin, causing some unsteadiness to the crew
standing in the cabin service areas. After the vibration had
abated, the crew commenced securing the cabin and awaited
instruction from the flight deck. Several aircraft crewmembers
travelling as passengers reported damage to the left engine nacelle
to members of the cabin crew. The CSM passed those observations on
to the flight crew. The CSM reported no adverse passenger reactions
during the event or during the subsequent return to Melbourne.

Flight recorder

The aircraft was fitted with an L3 Communications (LORAL) model
FA2100 solid-state flight data recorder (SSFDR). An excerpt of the
data from the recorder containing information from the previous
flight and the incident flight was obtained by the ATSB. That data
was analysed by the ATSB and used to prepare a summary of events
and actions during the incident flight.

The FDR information indicated that the left engine failed at
00:19:59UTC (11:19:59 Eastern Summer Time) and was characterised by
a sudden increase in the engine broadband vibration and a decrease
in the engine pressure ratio (EPR). At that time, the aircraft was
climbing through an altitude of 16,134 feet and maintaining 311
knots airspeed. Both left and right engines were operating at an N1
speed of approximately 94 percent. Vibration levels peaked around
two seconds following the initial event and the engine exhaust gas
temperature (EGT) peaked at 633 degrees C, six seconds after.

Within the next fourteen seconds, the flight crew had retarded
the left engine thrust lever, disengaged the auto-throttle and move
the left engine fuel cut-off lever to the OFF position. The left
engine fire switch was pulled at 00:21:47, however neither fire
bottle was discharged. All actions taken were as documented in the
'Engine fire, severe damage or separation' section of the B767-238
quick reference handbook.

Comparison of the engine broadband vibration levels found no
specific differences between the incident flight (before the
failure) and the previous flight. Examination of the graphically
presented information showed that at approximately twenty seconds
before the major vibration transient associated with the fan blade
release, a smaller transient occurred in the base vibration levels
(figure 5). Short-term escalations in engine vibration levels are
anomalous and often indicative of transient events such as
compressor aerofoil stalls and surges or foreign object
ingestion.

Tests and research

The ATSB examined the released blade segment, assisted by
authorised representatives from Pratt & Whitney.

Liberation of the fan blade segment occurred as a direct result
of fatigue cracking developing within the trailing edge of the
blade aerofoil section. A single transverse high-cycle fatigue
crack had developed from a 0.6mm deep pre-existing defect at the
blade trailing edge, approximately 290 millimetres above the root
face. Multiple surface arrest marks indicated to the growth of the
cracking over multiple flight cycles. Final tensile overload of the
remaining cross-section released the outer blade section after the
fatigue crack had grown to a length of approximately 85
millimetres.

The characteristics of the defect at the fatigue origin
identified it as a crack-like feature formed under localised
tensile loads. Heat tinting of the defect surfaces indicated the
exposure of the region to the elevated temperatures associated with
the blade overhaul. The implication from this was that the defect
was either present before the overhaul or was produced by the
overhaul operations. The defect location was within an area of
repair blending at the blade trailing edge. While the blending had
reduced the chord-wise width of the blade to one millimetre below
the specified minimum limit, it was not considered to have
significantly contributed to the development of fatigue cracking
from the trailing edge defect. Non-destructive testing procedures
carried out following the blade re-work had failed to detect the
trailing edge defect before the blade was re-introduced into
service within engine P-716610.

A copy of Technical Analysis report number 9/02 detailing the
examination of the failed blade is available from the bureau on
request.