Engine Failure, Boeing Co 717-200, VH-VQA, Near Melbourne, Victoria

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 1,149^oC, 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 1,149^oC for 1 minute and 46 seconds
before decreasing (Figure 6). It is likely that the maximum TGT
reached during the failure was higher than 1,149^oC,
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

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).
Bundesstelle für Flugunfalluntersuchung
(BFU).
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'.
Fatigue
that occurs at relatively small numbers of cycles. Brooks, C.
(1993). Metalurgical Failure Analysis. USA: McGraw-Hill,
Inc.
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.

SAFETY ACTIONS

Engine manufacturer

Following this and several other similar failures worldwide, the
engine manufacturer re-designed the Stage-1 High Pressure Turbine
(HPT 1) blades and prioritised the removal of the remaining
affected blades from the world fleet. Significant design changes
were made to the HPT 1, Life Improvement Package (LIP) blade's
internal cooling passage fillet radius to reduce the stress
concentrations in that area. The vapour aluminised coating is also
no longer applied to the blade's internal passages.

The manufacturer instigated a blade replacement program with the
highest cycle usage engines in the world fleet, totalling 94 units,
to be returned to the factory for blade replacement first, with the
remaining engines being completed in highest cycle order.

The probability of a double in flight shut-down (DIFSD) event
occurring, where an aircraft was fitted with two BR700-715 engines
with time in service approximating the time at which the HPT 1
blade problems were occurring, was analysed. That analysis
identified that the possibility of a DIFSD existed. To mitigate the
immediate risk, the manufacturer required that operators with
aircraft with both engines affected, remove the higher cycle count
engine and replace it with an unaffected engine.

The manufacturer issued BR 700 Propulsion System, Service
Bulletin (SB), SB-BR700-72-900361⁶.
That SB introduced an on-wing, ultrasonic inspection of the HPT 1
LIP blades for in-service engines.

Operator

Once the mode of failure for the engine was known, the operator
independently checked their fleet to determine if any of their
aircraft had both engines with time in service approximating the
time at which the failures were occurring in the world fleet. One
engine change was carried out following that check.

6. BR 700 Propulsion System, Service
Bulletin (SB), ENGINE - HIGH PRESSURE (HP) TURBINE BLADES -
ULTRASONIC INSPECTION OF THE HP TURBINE STAGE 1 BLADES,
NON-MODIFICATION SB-BR700-72-900361; dated Jun 03/05.

ANALYSIS

The circumstances of the engine failure were such that there was
no prior indication to enable the crew to take any action that may
have minimised the extent of the engine damage.

The blade failed as a result of a low cycle fatigue cracking
mechanism associated with cracking of the blade's vapour aluminised
coating.

At 1435 Eastern Standard Time on 10 August 2004, a Boeing
Company B717-200 aircraft, registered VH-VQA, was climbing to
cruise altitude on a scheduled passenger service from Melbourne,
Victoria to Hobart, Tasmania. As the aircraft passed through flight
level 110, the crew heard a loud bang, with a corresponding
increase in indicated left engine vibrations and the left engine
began to spool down. The crew then shut the engine down in
accordance with the operator's procedures and returned for a
landing.

Post incident examination of the BR700-715 engine found metal
fragments and metallisation in the exhaust area.

The engine was forwarded to the engine manufacturer for a
detailed investigation that was supervised by a representative of
the German Federal Bureau of Aircraft Accident Investigation. That
investigation found that the engine failure was due the release of
a single blade from Stage-1 of the high pressure turbine (HPT),
following the development of low-cycle fatigue cracking in its
internal cooling passages. The manufacturer indicated that there
had been four similar BR700-715 engine failures, with another
engine failure under investigation.

Computer stress modelling, carried out by the manufacturer on
the HPT blades, found stress levels in the blade's internal cooling
passages, in the area of the occurrence blade's crack propagation,
that were potentially in excess of the manufacturer's original
design intent. The thickness of the vapour aluminised surface
coating in the internal cooling passages was also variable. In
certain operational conditions the coating could crack, with the
subsequent growth of the crack into the parent material.

As a result of this and the other engine failures the operator
and the engine manufacturer have completed a number of safety
actions to prevent re-occurrence.