FACTUAL INFORMATION
At 1435 Eastern Standard Time on 10 August 2004, a Boeing
Company 717-200 aircraft, registered VH-VQA, was climbing to cruise
altitude on a scheduled passenger service from Melbourne, Vic. to
Hobart, Tas. with six crew and 52 passengers on board. As the
aircraft passed through flight level (FL) 110, the crew heard a
loud bang, with a corresponding increase in indicated left engine
vibrations. The left engine began to spool down and the turbine gas
temperature (TGT) indications began to increase significantly.
The crew initially brought the left engine power lever back to
idle. However, the TGT continued to increase, indicating a maximum
of 1,149oC, before they shut the engine down and
discharged a fire bottle into the cowling area in accordance with
the operator's procedures. They then notified Melbourne air traffic
control (ATC) of the engine failure and returned to Melbourne.
The operator examined the left engine and found metal fragments
in the exhaust area and some metallisation1 of the exhaust duct.
At the time of the failure, the BR700-715 engine, serial number
13148, had completed 10, 321 hours and 8,888 cycles since new, and
6,474 hours and 5,417 cycles since repair.
Engine investigation
The operator removed the engine and forwarded it to the engine
manufacturer in Germany for detailed investigation, under the
supervision of a representative of the German Federal Bureau of
Aircraft Accident Investigation (BFU2), on behalf of the Australian Transport
Safety Bureau (ATSB).
The manufacturer conducted a visual inspection of the engine's
exterior, noting a bulge around most of the circumference of the
high-pressure turbine (HPT) casing (Figure 1), in line with the
Stage-1 HPT (HPT 1). A borescope examination of the engine interior
showed that one HPT 1 blade was almost completely missing, with the
remaining HPT 1 blades separated just above the blade platforms
(Figure 2). There was also significant damage to the subsequent HPT
and low-pressure turbine stages. Examination of the engine's
compressor assembly revealed no significant damage. All of the high
energy debris from the failure had been fully contained3.
A detailed examination of the engine revealed that the reason
for the engine failure was the release of a single HPT 1 blade. The
blade failed following the development of low-cycle fatigue4 (LCF) cracking in its internal cooling
passages. All other engine damage was considered to be a
consequence of the initial HPT 1 blade failure.
Figure 1: Bulged HPT casing
Figure 2: Damage to HPT 1 and HPT 2 rotors
Blade design considerations
The failed HPT 1 blade (Figure 3) was a life improvement
package5 (LIP) blade. The blade was a
shrouded-tip aerofoil design, with multi-passage internal cooling
(Figure 4). There was a vapour aluminised coating on the blade's
external aerodynamic surfaces and internal cooling passages.
The manufacturer indicated that there have been four similar
failures of LIP HPT blades in the BR700-715 engine type, with
another engine failure still under investigation. One failure
occurred prior to this event in November 2003. The remainder
occurred after this incident.
Figure 3: The failed HPT 1 blade (position
21)
Following those failures, the manufacturer conducted additional
computer stress modelling on the LIP blades. That modelling found
that there were stress levels in the larger trombone radius
feature, within the blade's cooling passages (Figure 4) that were
potentially in excess of the manufacturer's original design intent.
The manufacturer also found that the thickness of the vapour
aluminised coating inside the blade's internal cooling passages was
variable and difficult to predict. In certain operational
conditions, dependent upon high strains in areas of stress
concentration and local temperature, the coating could crack with
the possibility of subsequent growth into the coated (parent)
material. The area from which the failure occurred was confirmed to
be the most susceptible to this behaviour (Figure 5).
Figure 4: Intact HPT 1 blade (left); blade internal
cooling passage showing trombone feature (right)
Figure 5: Computer generated stress diagram from the
manufacturer indicating the point of potentially excessive stress
and crack origin
Flight data recorder information
The ATSB's examination of the aircraft's flight data recorder
(FDR) for the occurrence flight found that the left engine had
surged as the aircraft passed through 10,240 ft. The engine
pressure ratio (EPR) and engine rotational speed indications
decreased abruptly, while the turbine gas temperature (TGT) for the
engine began to increase. HPT vibration values for the engine
increased from a level of 0.5 units before the failure to a maximum
of 6.3 units over a three-second period. The manufacturer's
high-limit for vibrations was 4.0 units.
The FDR readout indicated that the TGT for the engine continued
to increase following the engine failure and remained at an
indicated maximum of 1,149oC for 1 minute and 46 seconds
before decreasing (Figure 6). It is likely that the maximum TGT
reached during the failure was higher than 1,149oC,
however the aircraft systems do not record above that
temperature.
The FDR report indicated that there were no anomalies observed
in the performance of the left engine prior to the failure.
Figure 6: FDR data plot of key engine parameters at the
time of the failure
1 Metal pulverised by the
turbine becomes molten and flows rearward attaching to the
subsequent turbine and exhaust assemblies (US Department of the Air
Force (1987). Safety Investigative Techniques (AF Pamphlet
127-1, Volume II. Washington DC: Author).
2 Bundesstelle für
Flugunfalluntersuchung (BFU).
3 FAA AC 33-5, paragraph
5.c. definitions state '…Contained means that no fragments are
released through the engine structure, but fragments may be ejected
out of the engine air inlet or exhaust'.
4 Fatigue that occurs at
relatively small numbers of cycles. Brooks, C. (1993).
Metalurgical Failure Analysis. USA: McGraw-Hill,
Inc.
5 The Life improvement
Package 3 (LIP3) was a suite of HP Turbine modifications that
included the HPT blade P/N BRH20351. The manufacturer introduced
the package by SB-BR700-72-100801.