Boeing 747-436, G-BNLK, Sydney, NSW, 10 August 2002

Investigation number

200203671

Occurrence date

10/08/2002

Location

6 km N Sydney, Aero.

Report release date

18/12/2003

Report status

Final

Investigation type

Occurrence Investigation

Investigation status

Completed

Aviation occurrence type

Fire

Occurrence category

Serious Incident

Highest injury level

None

Technical Analysis Report Safety Action Significant Factors Factual Information

Technical Analysis Report: Boeing Commercial Aircraft
Group, 747-436, G-BNLK

1. FACTUAL INFORMATION

1.1. Examination brief

The disassembled components of an electric air-cooling fan were
received by the ATSB Technical Analysis unit for examination and
analysis of the damaged fan impeller. The fan unit had been fitted
to a Boeing 747 aircraft (registration G-BNLK) to provide forced
air circulation for a forward galley chiller unit. During the early
stages of a flight on 10 August 2002, a small fire developed in the
forward cargo compartment adjacent to this unit. Physical and
recorded evidence suggested the fire had initiated from electrical
arcing that was a result of a wiring short-circuit near the fan
terminal housing.

1.2. Samples received

Data plates affixed to the fan housing identified the unit as a
three-phase unit (part number 73259E, serial number 3676),
manufactured by Sunstrand (San Diego, California) in 1994. The fan
was an axial flow design, with a single bell-shaped impeller
manufactured from a moulded resin material. The motor and fan
outlet guide vane assembly shared an integral housing which also
carried the (damaged) electrical terminal housing. The motor was a
brushless (induction) design, with the armature supported on sealed
rolling element bearings. Surrounding the impeller was an aluminium
shroud, which formed the fan intake and also provided for the
mounting and support of the unit and its associated ducting.

1.3. Visual examination

1.3.1. Impeller

The impeller unit had been effectively 'cobbed', with all eleven
blades fractured at or immediately adjacent to the impeller hub.
The uneven, irregular nature of the fractures suggested the failure
occurred as a cascading fragmentation event, with multiple sections
of blade breaking away and striking others, causing further
break-up. A study of all fracture surfaces failed to identify any
evidence of pre-existing defects or cracking that may have
precipitated the initial blade failure, nor was any indication
found of unusual hard-object impact damage that may have suggested
foreign object ingestion. There was some evidence however that
suggested early damage to the impeller blade forward corners - many
blades showed breakage of a curved lip of material from the corners
in a manner that suggested possible contact with the fan
shroud.

The inside surfaces of the impeller body carried a heavy coating
of a powdery brown dust, which was also evident inside the armature
core and around the end of the motor housing. This material was
loose and easily wiped away by hand and a sample was taken for
later qualitative chemical analysis.

1.3.2. Fan shroud

Produced as a machined die-casting from an aluminium alloy, the
fan shroud encased the full impeller length and showed no evidence
of having failed to fully contain the fractured impeller blades.
Around the blade tip path, the shroud internal surfaces showed
several circumferential wear bands that indicated significant tip
interference. These were most clearly defined at the forward and
rear limits of the blade path. Random indentation and scratching
damage was noted around the shroud 'throat' region - this was
consistent with the effect of multiple fragment impacts produced by
the failing impeller blades. No specific evidence of hard-object
impact damage was found across the blade path. Dark, waxy stains
produced by an unidentified liquid draining through the fan shroud
were noted at the low-point of the assembly and a series of
tide-marks were formed on the front face of the shroud as the
liquid had accumulated and later drained away.

1.3.3. Fan motor and housing

Being a brushless three-phase induction motor, the unit was
comparatively simple in construction, with a star-wound stator
containing a compact cage-type armature. Internal inspection found
evidence of rub and erosion of the iron stator former over the
innermost twenty millimetres of the stator length, however the
armature did not reflect this and showed no evidence of operational
damage. Several areas where material had been removed by bevelling
the core corners were evidence that the armature had been
re-balanced at some time following original manufacture. Both
armature shaft bearings rotated smoothly by hand and showed no
notable indications of distress or abnormal operation. The rear
armature bearing housing contained particulate debris that appeared
to be the fine brown dust mixed with lubricant lost from the
bearing unit. The contact points of the leaf spring and the housing
bottom showed light fretting damage, with some evidence of rotation
also noted.

The fan motor electrical supply was routed behind the outlet
guide vanes and terminated in an external housing fitted with an
eight-pole connector. A localised area of heavy electrical arcing
was clearly evident adjacent to the connector and had resulted in
the melting and loss of around ten to fifteen millimetres of
material in a v-shaped notch from both the front and side faces of
the connector housing (figure 15). The metal loss had exposed some
of the internal wiring and had produced an appreciable level of
heating around the contact area, as evident from the tinting of the
surrounding surfaces.

1.4. Dust analysis

The sample of brown dust recovered from the underside of the
impeller body was qualitatively analysed using energy-dispersive
x-ray spectroscopy techniques under the scanning electron
microscope. The results of this testwork showed the material to be
comprised primarily of an iron-oxide compound, with traces of
chromium, aluminium and silicon.

2. ANALYSIS

2.1. Impeller failure

All of the failed impeller blades exhibited brittle overload
fractures. No evidence of material flaws, cracks or other
pre-existing damage was shown by the blade stubs, however this was
not conclusive, given the opportunity was not available to examine
all of the blade remnants. Resin materials such as that used to
manufacture the impeller have an inherent lack of ductility and as
such, are prone to cracking and fracture under impact or elevated
stress conditions. While no direct evidence was found to suggest
the ingestion of foreign object/s, the damage sustained was not
unlike that expected from such an event, and thus this possibility
cannot be discounted.

2.2. Motor damage

The abrasion exhibited by the inner sections of the stator
assembly was believed to be evidence of armature contact,
presumably caused by the failure and collapse of the end bearing
unit. The dust under the fan impeller was believed to be an
accumulation of wear products from this contact - both the stator
former and the armature core were ferrous alloys, while the
remainder of the motor and fan assembly was predominantly aluminium
based materials. Given that the installed armature showed no
indication of stator contact and the end bearing was sound, it was
concluded that the stator damage was a product of a previous
failure that had been repaired by replacement of the armature and
the re-use of the remaining components, including the impeller.

The collapse and failure of an armature shaft bearing and the
shaft misalignment that results would be expected to alter the
impeller - shroud clearances, with a risk of contact between the
shroud and the impeller blade tips if the misalignment became
severe enough. Impeller blade tip contact, if it did not produce
immediate blade breakage, may produce latent cracking damage that
could lead to later blade failures if not detected. While there was
no evidence found to suggest that pre-existing damage of this
nature existed, the possibility remained that this damage had been
sustained given the evidence of the rear bearing failure and the
contact marks inside the fan shroud.

2.3. Vibration

Rotary equipment such as the cooling fan relies upon accurate
dynamic balancing to minimise the vibration induced during
operation. Events such as the impeller failure will disrupt the
balance of the assembly and can lead to significantly increased
vibration levels. Wiring or piping that is installed against or in
contact with the vibrating equipment may sustain fretting or
erosion damage if the external protection or insulation has not
safeguarded against this event.

3. CONCLUSIONS

3.1. Findings

The cooling fan impeller had sustained gross breakage of all
blades at or adjacent to the body of the impeller.
All fractures were brittle in nature and showed no evidence of
pre-existing defects.
No evidence of foreign object damage was found.
The fracture profiles of some blades suggested preferential tip
breakage before the complete blade failure.
The fan motor showed evidence of a previous rear bearing
failure that had produced contact and wear between the armature and
the stator. The armature had subsequently been replaced, however
the stator and other motor components remained in service.
The previous bearing failure may have allowed operating contact
between the impeller blade tips and the fan shroud. This contact
may have produced blade cracking which predisposed the impeller to
failure in the manner observed.
After failure, the fan impeller would have presented a
significant unbalanced load to the fan.

History of the flight

Shortly after take-off from runway 34L at Sydney, the flight crew of the Boeing 747-400 aircraft received a forward cargo compartment fire warning on the Engine Indicating and Crew Alerting System (EICAS). On receiving the warning message the crew actioned the appropriate checklist, activated the fire suppression system and transmitted a MAYDAY. At the same time, flight attendants noticed a fine mist and the smell of smoke in the passenger cabin. The crew then returned the aircraft to Sydney, where an uneventful overweight landing was conducted.

Prior to landing, the EICAS fire warning message ceased. This indicated that the aircraft fire suppression system may have successfully extinguished any fire, however the cabin fumes were still evident. After landing, the flight crew stopped the aircraft on the runway where emergency services came to their assistance. After confirming with the flight crew that the fire warning message was no longer present, the emergency services assessed the aircraft from the ground, then allowed the passengers and cabin crew to disembark to a safe distance via mobile stairs positioned at the aircraft's front left door. Once the passengers and cabin crew were clear of the aircraft, the emergency services opened the forward cargo door.

A hot spot was detected on the left side of the forward cargo bay at body station STA900, where the side wall lining was found to be heat affected. Removal of the lining revealed burned insulation blanket material, discolouration of the aircraft skin and burned/broken electrical wires that powered the forward galley chiller boost fan situated in the area (see Fig 1). As the fire was no longer evident, ground engineers isolated the chiller boost fan electrical circuit and towed the aircraft clear of the runway.


FIGURE 1: Forward cargo bay with expanded view of chiller boost fan location

Aircraft structural damage

Non-destructive testing to check for cracking and conductivity of the aircraft skin adjacent to the affected area was carried out. No cracks were detected, however the conductivity test revealed three locations where the skin had been substantially affected by heat (see Fig 2). The most severely affected area required a temporary skin repair before the aircraft could be flown back to the operator's maintenance facility in the United Kingdom, where the heat-affected aircraft skin was replaced.


FIGURE 2: Heat affected areas

Sidewall lining and insulation blankets

The fibreglass sidewall lining between STA880 to STA900 was visibly heat damaged with discolouration observed on the side facing into the cargo compartment. Inspection of the reverse side revealed burned layers of fibreglass confined to a localised area approximately 30cm x 45cm (see Fig. 3). The insulation blankets that lined the aircraft skin were made of a fibreglass core with a metallised TedlarTM film on one side and a MylarTM film on the other and had been subjected to localised heat and fire (see Fig. 4).

Samples of the sidewall lining and insulation blanket were sent to the United States of America, Federal Aviation Administration (FAA) technical centre and the aircraft manufacturer for analysis and testing.


FIGURE 3: Sidewall lining	FIGURE 4: Insulation blanket

The examinations determined that both the sidewall lining and insulation blanket samples complied with the appropriate material specifications for aircraft use.

The flammability testing, conducted by the FAA, on samples of the insulation blanket included a vertical Bunsen burner test, which was mandated in Federal Aviation Regulation FAR 25.853 - Appendix F. The samples tested met the requirements, but due to their limited size, the result was not conclusive as to the integrity of the entire blanket.

The aircraft manufacturer's tests revealed contamination on the insulation blanket samples. This contamination consisted of environmental dust, fibres and corrosion inhibiting compound. These contaminants were consistent with general contamination found during evaluations of other in-service insulation blankets and were considered to be normal.

The aircraft manufacturer's 'flame propagation cotton swab tests' found areas on the blanket samples that were self extinguishing while other areas showed "flame propagation uncharacteristic of that expected for new insulation blankets". It was unknown whether contamination, in-service ageing, or heat exposure, or a combination of these, altered the blanket's flame propagation characteristics.

Boost fan system

A galley chiller boost fan system was installed in the aircraft to provide forced air circulation over the forward galley chiller units increasing their cooling efficiency. The system incorporated a vaneaxial-type three-phase fan, powered by the aircraft's number-3 alternating current electrical system. Control power was supplied by the aircraft's direct current electrical system, with operation being automatic on selection of the galley chillers to ON. Circuit protection was provided by a 20 ampere circuit breaker and a cargo fire cutoff relay.

Chiller boost fan

An inspection of the boost fan revealed a burn hole and sooting on its casing adjacent to the electrical terminal (see Fig. 5). The electrical wiring to the fan was found to have four of its seven wires broken, with all of the wires displaying sooting discolouration (see Fig's. 6 and 7). The soot marks corresponded to those on the fan casing and when positioned together, revealed that the wires had separated at a point adjacent to the corner of the electrical terminal. The failure of the wires produced electrical arcing, which melted the casing, resulting in the burn hole observed.

Further inspection found that all of the fan impeller blades had failed just above their roots (see Fig. 8). Neither the impeller nor the fan shroud showed signs of hard body impact damage.


FIGURE 5: Electrical terminal	FIGURE 6: Broken wires

FIGURE 7: Sooting evident	FIGURE 8: Fan impeller blades failed

Technical examination of the fan found that the impeller was made from a moulded resin material. There was no evidence of any pre-existing defects or cracking found on the blade fracture surfaces. However, a number of blades showed breakage of a curved lip of material from their forward corner. This condition was consistent with overload fatigue possibly due to the blade tips contacting the fan shroud. Such a condition may have occurred prior to the blades total failure. The inside surface of the impeller was coated with a brown powder, determined to be primarily iron oxide.

The aluminium alloy shroud contained several circumferential wear marks that were adjacent to the impeller blade path. Although there were random scratches, no evidence of gouging or penetration of the shroud skin was found.

Further disassembly of the fan revealed wear on the electrical motor stator, indicating that it had been subjected to armature rubbing. The armature did not display similar wear patterns. Rubbing of this nature usually occurs as a result of bearing failure or excessive wear, leading to armature oscillations. For the full technical report see Attachment 'A'

Chiller boost fan service history

The chiller boost fan entered service in 1994, with the last overhaul being in June 2000, after removal from service because of electrical failure. The maintenance records for that overhaul stated: "Unit noisy due worn bearings, all other parameters ok. Reported defect not confirmed. Disassembled, cleaned and inspected, bearings renewed, unit reassembled and tested to spec". The fan was then fitted to the incident aircraft on 2 August 2000. No subsequent maintenance was recorded.

Chiller boost fan circuit breaker and electrical relay

The installed circuit breaker was a 20 ampere three-phase, push-pull high performance, trip free type, designed for aircraft installations. It's design allowed for increased amperage through the circuit for a specific time before tripping (breaking the circuit) and was used in large motor load applications where the inrush current would trip a standard circuit breaker. The length of time taken to cause the circuit breaker to trip varied according to the current it received. The aircraft manufacturer advised that "At 385 per cent or 400 per cent [load rating], this breaker will trip between 2.3 to 10 seconds". This prevented aircraft electrical power surges from "nuisance" tripping of the circuit breaker and rendering the boost fan inoperative.

A number of tests were conducted on the circuit breaker, including a 'load withstanding test'. This required the controlled increase in current through the circuit breaker, with time to trip recorded. This test was conducted at 105 per cent, 140 per cent and 200 per cent values, as per the manufacturers test procedures.

The installed relay was a 25 ampere, electromagnetic, three pole, single throw, normally open type. This was also subjected to a number of tests including 'Coil resistance', 'Coil hold and drop voltage' and 'Voltage drop and switching test across all three phases'.

These tests were performed under the supervision of the United Kingdom Air Accident Investigation Branch. Both components were found to comply with their operational specifications, with no adverse mechanical or operational functions found during the testing. As a result, both components were considered to be serviceable.

Quick Access Recorder

The aircraft's Quick Access Recorder (QAR) data was analysed by the Australian Transport Safety Bureau with the following information retrieved.

During climb the number-3 alternating current system showed a momentary increase in load from a nominal 31 per cent to 54 per cent, which equated to an increase in current draw of 57 amperes.

Four seconds later, the load was again recorded and had returned to the nominal 30 percentage range, where it remained for the rest of the flight.

Approximately 1 minute later the QAR recorded a forward cargo fire.

Approximately 3 minutes later, the first cargo fire bottle low quantity message appeared, indicating that extinguishant had been discharged successfully.

Other recorded data received from the aircraft's central maintenance computer (CMC) confirmed the arming of the fire bottles approximately 2 minutes after the fire warning and the discharging of the last two fire bottles after the aircraft landed.

Cargo fire detection/extinguishing system

The aircraft incorporated two dual loop smoke detectors in each cargo compartment. Air from throughout the compartment was drawn through the detectors and sampled. In normal operation, both loops must sense smoke for a fire warning to be activated. If the system detects a loop fault during self-test at aircraft power on, it would reconfigure to a single loop operation.

The cargo fire module located on the overhead instrument panel in the flight deck incorporated forward and aft compartment ARM buttons and a DISCH discharge button. On sensing smoke, the relevant ARM button, along with an EICAS message would be illuminated, alerting the crew of the fire. The crew must then push the ARM button in. This action disables electrical power to a number of circuits, including the galley chiller fan circuit. Extinguishing is then achieved by pressing the DISCH button (see Fig. 9).


FIGURE 9: Overhead instrument panel with expanded view of cargo fire panel

Four fire extinguisher bottles (A, B, C and D) service the cargo compartments, each having discharge lines to both the forward or aft compartments. On depression of the DISCH switch, bottles A and B discharge flooding the selected compartment with extinguishing agent. Bottles C and D are not discharged until 30 minutes later. If the aircraft reaches the ground before the 30 minutes are up, the bottles will discharge on touch down. The system was designed to give up to 180 minutes of discharge time.

Aircraft Details

Manufacturer

The Boeing Company

Model

747

Registration

G-BNLK

Serial number

24053

Operation type

Air Transport High Capacity

Departure point

Sydney, NSW

Departure time

1500 hours EST

Destination

Singapore

Damage

Minor

Technical Analysis Report: Boeing Commercial Aircraft
Group, 747-436, G-BNLK