Abnormal engine indications - Boeing 777 aircraft, 9V-SRE, Tindal, NT, 18 November 2001

Investigation number

200105494

Occurrence date

18/11/2001

Location

Tindal

Report release date

18/03/2003

Report status

Final

Investigation type

Occurrence Investigation

Investigation status

Completed

Aviation occurrence type

Abnormal engine indications

Occurrence category

Incident

Highest injury level

None

Analysis Technical Analysis Report Safety Action Significant Factors Summary

The soft body damage identified as bending on the IPC stage-2
blades by the boroscope inspection on 8 April 2001, was consistent
with that incurred by the engine ingesting a soft bodied object
such as a bird or ice. These defects were determined to be within
serviceable limits, and as such would not have failed unless
additional abnormal forces were applied to them.

There was no evidence of any system faults or additional
internal deterioration of the engine after the increase in TGT was
detected on 15 October 2001. It is possible that the increase was
the result of the number-28 vane moving to the closed position
after the lever failed. The nudging of the number-28 vane towards
its normal closed position by the number-29 vane during engine shut
down may have been enough to allow the failed lever to assume its
normal position, thereby escaping easy detection.

The ATSB's technical analysis report determined that the failure
of the lever was probably a product of a progressive fatigue
cracking mechanism. The observed fracture features suggested that
the crack initiation occurred from the connecting-pin, hole.

The bias in the wear mark to one side of the failed lever
combined with the contact points evident on the number-28 and 29
vanes indicated that during engine operation, the number-28 vane
had remained in a predominantly closed position.

As found with previous VSV lever failures, the out of sequence
position of the vane created turbulence in the airflow. This would
have been felt on the IPC stage-2 blades inducing abnormal loads.
Two of the bent IPC stage-2 blades were unable to tolerate that
excitation and as a result their blade tips failed. The released
sections of blade then ricocheted within that stage before being
projected through the engine, impacting blades in other stages down
stream.

As the sections of blade and accumulated debris passed through
the engine, a piece of material impacted the leading edge of a HPC
stage-1 blade, chipping and cracking it. The crack then progressed
to a point where the blade failed and detached.

Although it was possible for the IPC stage-2 blade tips to fail
as a result of bird or ice ingestion during the incident flight,
the lack of supporting evidence, and the failure mechanism of the
HPC blade did not support such an event.

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Boeing 777-212ER, 9V-SRE

1. FACTUAL INFORMATION

1.1 Introduction

During a flight from Brisbane to Singapore, the crew of the
Boeing 777-212ER aircraft noticed the onset of abnormal vibration
levels and several 'thumps' from the right engine. While continuing
to monitor the engine, the vibration and thumps recurred and the
engine was subsequently shut down after the oil and exhaust gas
temperatures rapidly increased. Following a diversion to Darwin,
the crew conducted an uneventful single engine landing.

On initial inspection, maintenance personnel found a single
fractured first-stage variable stator vane (VSV) control lever
(figure 1). Later internal boroscopic inspection of the engine
found significant levels of mechanical damage within the
intermediate and high-pressure compressor stages.

The fractured lever and a selection of other levers from the
first-stage VSV assembly were removed from the engine for
examination by the ATSB.

Fig 1

Fig. 1 External view of the
stage-one variable stator vane actuator ring and the single
fractured lever (arrowed).

1.2 Visual examination and fractography (failed lever)

Initial inspection in the ATSB laboratory found the VSV lever
had fractured transversely through the end of the arm section
(figure 2), at a location coincident with the riveted connection to
the actuator pin. The fracture path followed a uniform arc,
extending from one side of the arm to the opposite and intersecting
the pin connection at the centre (figure 3).

Fig 2

Fig. 2 Underside of the VSV
lever removed from the engine.

Fig 3

Fig. 3 Underside of
the VSV lever showing the curved fracture path and the wear mark
produced by in-service movement after failure.

A prominent track mark had developed on the underside of the arm
where the relative movement between the separated arm and the pin
flange had produced appreciable wear. The effects of wear extended
to the fracture surfaces themselves, which were heavily eroded and
all fracture surface detail obliterated (figures 4 & 5). Apart
from the fracture, the arm had sustained little other mechanical
damage and showed no evidence of deformation or distortion
associated with the failure.

Fig 4

Fig. 4 Arm section fracture
surface showing degree of wear and loss of detail.

Fig 5

Fig. 5 Pin section fracture
- adjoins the surface shown in figure 4.

On close inspection, the fracture path appeared to intersect the
bore of the rivet hole, with slight upward 'dishing' of the arm
section beneath the rivet head (figure 6). A clearance or gap was
not evident between the pin shaft and the bore of the arm hole
through which the pin shaft was riveted.

Fig 6

Fig. 6 Gap between the pin
flange and the control arm produced by upward 'dishing' of the arm
beneath the rivet head.

The examination did not show any evidence of binding or
excessive friction between the actuator pin and the mating bushing,
nor did any other component show significant indications of
miss-installation or anomalous operation.

After departing Brisbane en-route to Singapore, the crew of the
Boeing 777-212ER aircraft heard and felt two thumps through the
airframe and noticed a severe vibration indication of the right
engine. The vibration subsided before re-occurring moments later
with an increase in the engine's exhaust gas temperature also
indicated. The crew conducted an in-flight engine shutdown and
requested Air Traffic Control clearance to divert to Darwin where
an uneventful single-engine landing was carried out.

An inspection conducted by ground engineers at Darwin found a
stage-1, variable stator vane (VSV) control lever broken on the
right engine. A boroscope inspection of the engine interior was
then carried out with a number of compressor blades found damaged.
The engine was removed from the aircraft and sent to an overhaul
facility for disassembly and evaluation.

Engine description

The Rolls-Royce Trent 800, was a triple spool turbofan engine.
Its construction consisted of a single-stage low pressure fan
connected to a five-stage low pressure turbine. An eight-stage
intermediate pressure compressor (IPC) connected to a single stage
intermediate pressure turbine and a six-stage high pressure
compressor (HPC) connected to a single-stage high pressure turbine
(see fig 1).

To maintain maximum efficiency during all power settings, the
airflow through the engine needed to be controlled to prevent
stalling or surging. This was achieved by a single stage of
variable inlet guide vanes installed between the fan and the first
stage of the IPC and two stages of variable stator vanes, VSV stage
1 and VSV stage 2 (VSV-1 and VSV-2) installed between IPC stages,
one/two and two/three respectively. Each Variable Stator Vane was
attached to a lever that transferred the linear input from the
controlling actuators and unison rings to a rotational movement of
the vane (see fig 2). These levers consisted of an arm and
connecting pin (see fig 3).

During engine start, these vanes would have been in their most
closed position with internal engine bleed valves open. As the
power was increased, the bleed valves would close and the vanes
move toward their full open position allowing optimum airflow
through the engine.

Post incident engine inspection

Prior to disassembly, the engine's exterior was inspected with
the broken VSV-1 lever identified as being in the number-28
position. No evidence was observed of bird impact or other external
defects. Rigging and clearance checks of the VSV control system
were carried out with no anomalies found. The remaining VSV-1 and
VSV-2 levers were then removed and crack tested using a dye
penetrant inspection. No evidence of cracking was found on any of
those levers.

After separating the engine-to-modular level, the IPC and HPC
modules were completely disassembled for a detailed inspection.

IPC inspection

Removal of the IPC casing revealed six stage-2 blades displaying
soft body impact damage¹ resulting in bending of the blades. On two of those
blades the corners had also detached. Three other blades displayed
hard body impact damage² with cuts and nicks (small cuts) on their surfaces. Two
blades with minor nicks were found in stage 3, with only one blade
in stage 5 showing nick damage. All of the stage-8 blades displayed
hard body impact damage on their trailing edges, a few blades also
having nicks on their leading edges. There was no evidence of
damage to the disc material adjacent to the blade roots on any of
the eight stages.

The IPC case lining was examined with only minor damage evident.
The number-28 VSV 1 vane, had a wear mark on the leading edge lower
corner with a noticeable worn stepped area on its horizontal
surface above its base. The adjacent number-29 vane had a contact
mark at a point mid span on the vane and one on its base. The
remaining VSV-1 and all the VSV-2 vanes were found to be undamaged.
Dark deposits were evident around the base of each VSV. These
deposits formed a black ringed area around all except for the
number-28 vane where the mark was crescent shaped.

When the number-28 and 29 vanes were positioned so that the wear
marks on both vanes aligned, the number-29 vane was found to be in
its normal full open position while the number-28 vane was noted as
sitting beyond its normal closed position. The dark crescent area
around the base of the number-28 vane also coincided with the angle
of the vane's root. When the number-29 vane was moved to the closed
position it was seen to nudge the number-28 vane up towards its
normal closed position.

¹ Having been
impacted by an object made from a softer material than the blade
itself.

² Having been
impacted by an object made from the same or harder material than
the blade itself.

Figure 2, 3

Failed VSV-1 lever

The Australian Transport Safety Bureau (ATSB), conducted a
metallurgical examination of the failed number-28, VSV-1 lever (see
fig 3). The examination found that:

'The lever had fractured transversely through the end of the arm
section, at a location coincident with the riveted connection to
the actuator pin. The fracture path followed a uniform arc,
extending from one side of the arm to the opposite and intersecting
the pin connection at the centre. A prominent track mark had
developed on the underside of the arm where the relative movement
between the separated arm and the pin flange had produced
appreciable wear. On close inspection, the fracture path appeared
to intersect the bore of the rivet hole with a slight upward
'dishing' of the arm section beneath the rivet head'.

The examination also determined that:

'During riveting, the expansion of the rivet shaft could induce
tensile stresses within the bore of the rivet hole if the diameter
was insufficient to allow for the expansion. Tensile stresses of
this nature would be expected to predispose the lever arm to the
initiation and propagation of fatigue cracking'.

The examination of a further four VSV-1 levers was conducted,
with welding and partial fusion between the lever and connecting
pin evident, and varying degrees of cracking also evident on all
four levers. For the full technical report refer to attachment
A.

A further investigation by the engine manufacturer, identified
the presence of a double-sided chamfer to the lever holes on a
small number of levers. This removal of material during the lever
manufacture may have led to the overheating and partial welding of
the lever material during the rivet forming.

HPC inspection

In the HPC, all of the stage-1 blades displayed severe hard body
impact damage with one blade found to have failed, detaching above
the blade root. Stages 2 to 6 showed hard body impact damage to
varying degrees on all the blades.

Close examination of the failed HPC stage-1 blade, found a
chipped area in the leading edge, with the fracture surface
revealing a number of crack progression marks indicating that the
failure was progressive over a number of cycles and not
instantaneous. The exact number of cycles required to fail the
blade could not be determined (see fig 4).

Engine history

The engine commenced service in December 1998 and had completed
a total of 8923 hours and 2373 cycles at the time of this incident.
Its service history showed that on 8 April 2001, a routine
boroscope inspection detected damage to a number of IPC stage 2
blades in the form of bending and curling to their tips. This
damage was assessed to be within the manufacturer's allowable
limits so the engine remained in service. On 17 October 2001, a
substantial shift in the turbine gas temperature (TGT) was detected
giving a warning that the engine's efficiency had deteriorated
significantly. A boroscope inspection was carried out on the engine
with only the previously recorded IPC stage-2 bent blade damage
found. No other damage was observed on the engine. Checks of the
air system, engine bleed air and monitoring systems were carried
out, however the reason for the TGT shift could not be
determined.

Previous VSV lever failures

Although the manufacturer had not experienced previous failures
of VSV-1 levers, failures of VSV-2 levers had been recorded on two
separate occasion. On those occasions the connecting pin's had
fretted through the body of the lever due to inadequate riveting
during the manufacturing process. The result of the levers failing
was the closure of their associated variable stator vanes, which
created a disruption to the airflow behind them. The vibration
subsequently experienced by the blades passing the area resulted in
fragments of disc material breaking off and migrating through the
engine, damaging blades further down stream. These failures were
indicated by a progressive increase in the engine's TGT over
periods of 2 to 4 weeks.

The manufacturer issued a service bulletin, RB211-72-D516 to all
operators recommending an inspection of, 'the six VSV-1 and VSV-2
levers either side of the actuating mechanism control rod
connection, for significant relative movement between the lever and
connecting pin'.

Aircraft Details

Manufacturer

The Boeing Company

Model

777

Registration

9V-SRE

Serial number

226

Operation type

Air Transport High Capacity

Departure point

Brisbane, Qld

Destination

Singapore

Damage

Nil