The Australian Transport Safety Bureau recommends that the UK Civil Aviation Authority review the repair and overhaul processes for the failed torque links attachment lugs and also for the wheel failure identified in occurrence 199903327, to ensure that they conform to the appropriate airworthiness requirements.
Summary
On 9 October 1999, a Fokker F28-100 aircraft, on a direct service from Brisbane with 84 persons on board, experienced severe vibration through the airframe during landing at Norfolk Island. The crew stopped the aircraft on the runway and, after a preliminary examination, taxied the aircraft to the terminal where the passengers disembarked normally. There were no injuries.
Investigation revealed that the left main landing gear upper torque link attachment lugs had broken. The upper torque link attachment point on the landing gear main fitting was an integrally forged double lug with a stiffening web between the two lugs. The maintenance documentation showed that the main landing gear had completed 16,579 cycles since new and 658 cycles since last overhaul.
The Australian Transport Safety Bureau (ATSB) conducted specialist fracture analysis of all the broken landing gear components. The specialist report concluded that the failure of the torque link attachment lugs was associated with the extension of pre-existing cracking in the lug-stiffening web while torque was transmitting through the torque links. The initial cracking in the web was caused by stress corrosion. The propagation of the fatigue crack was consistent with a loading regime that involved the sideways flexing of the wheel rim. This will occur when a turning moment (torque) is applied to the main landing gear while the wheels are rotating, such as during ground turning or crosswind landings.
The evidence showed that the region of the pivot pin bore and locating pinhole had been reworked during overhaul. At that time material had been removed by localised surface grinding to remove corrosion. The pivot bore surface was then shot peened and repainted with a chromate based paint primer. However, the paint primer exhibited poor adhesion, and the shot peening coverage was haphazard. Consequently, these measures had been ineffective in preventing stress corrosion.
The final failure of the torque link attachment lugs occurred during the initial stage of the landing and occurred while the landing gear was being subjected to significant torque loads. It is likely that the torque loads were associated with crosswind conditions. The crew report for a previous landing incident with this aircraft at Norfolk Island, indicated that crosswind components of 15 knots or higher are regularly experienced during operations at Norfolk Island.
The operator's maintenance facility reported that part of the left main gear shimmy damper was found to have been wrongly re-assembled during last overhaul. The fracture analysis evidence indicates this would have had minimal if any influence on the start or development of the fatigue cracks that led to the failure of the torque link attachment lugs.
A previous incident occurred on 4 July 1999, involving the left main landing gear of this aircraft, also while landing at Norfolk Island. In that incident, ATSB occurrence number 199903327, the outboard main landing gear wheel broke away from the wheel hub during the landing roll. The ATSB specialist fracture analysis report (see below) found the wheel failure had started and progressed in similar circumstances to those for the torque link attachment lugs.
During short final to landing at a heliport, as the helicopter approached a hover at about 6 ft AGL, the engine spooled down to idle RPM. The pilot continued with the landing and landed the helicopter without further incident. Following the landing the engine remained at idle RPM and did not respond to pilot inputs.
Investigation revealed internal damage to the engine compressor section. This damage was identified to be caused by the breaking off and subsequent ingestion of a compressor third stage stator vane blade. An analysis of the vane blade root indicated a pre-existing fatigue crack at the forward edge of the blade. The most recent inspection of the compressor had occurred 218.9 hours prior to the occurrence. The compressor case had accumulated 513.7 hours since overhaul. Erosion of the blade vane was not a factor. The investigation was unable to determine the origin of the fatigue crack.
The crew of a Dash 8 reported on the mandatory broadcast zone frequency that they were inbound to Dubbo, at a position 40 NM south-east, and on descent from FL180. The only response to this transmission was from the pilot of a Piper Cherokee who advised that he was approximately 15 NM east of Dubbo at 4,000 ft and inbound. Approaching 4,000 ft the crew of the Dash 8 requested the position of the Cherokee. The pilot advised that he was now 8 NM from Dubbo and descending to 2,400 ft. They then asked the pilot if he was south of "the highway".
The pilot of the Cherokee confirmed that he was south of "the highway". The crew of the Dash 8 indicated that they would remain north of the highway and join a 5 NM final approach to runway 23, and requested that he remain south.
When the crew subsequently reported that they were 5.5 NM from Dubbo and about to turn final for a straight-in approach to Runway 23, they observed a Piper Cherokee pass from their right to left at an estimated distance of 400 m and 200 ft below.
Weather conditions at the time were reported to be CAVOK.
"The highway" to which the crew was referring was the Mitchell Highway that runs south-east from Dubbo to Wellington, almost directly beneath their track. The pilot of the Cherokee, who held a private licence, was undertaking a solo navigation exercise as part of the training for upgrading to a commercial licence. He reported that he was just north of his planned track from Gulgong to Dubbo. When asked by the crew of the Dash 8 if he was south of the highway he assumed that they were referring to the Dubbo to Dunedoo road, aligned east-north-east from Dubbo and that he could see to his north. He was not familiar with the Dubbo area and was not aware of the existence of the Mitchell Highway although this road was annotated as a highway on his Visual Navigation Chart. Additionally, he thought that the Dash 8 would pass behind him and join for a 5 NM final north of this road.
The use of a line feature to assure separation between aircraft is an accepted and generally sound technique. However, the use of the generic term "the highway" by the crew of the Dash 8 introduced an ambiguity that neither the crew of the Dash 8 nor the pilot of the Cherokee was aware of at the time. Specifying the road by name or description should have removed this ambiguity.
Occurrence summary
Investigation number
199904771
Occurrence date
11/10/1999
Location
10 km NE Dubbo, Aero.
State
New South Wales
Report release date
28/02/2000
Report status
Final
Investigation type
Occurrence Investigation
Investigation status
Completed
Mode of transport
Aviation
Aviation occurrence category
Near collision
Occurrence class
Incident
Highest injury level
None
Aircraft details
Manufacturer
De Havilland Canada/De Havilland Aircraft of Canada
The aircraft overran the end of runway 21L after landing. The runway was wet from very heavy rain. The crew had flown the approach with flaps 25 selected, and had intended to select idle reverse thrust after touchdown.
The aircraft encountered very heavy rain on late final approach and deviated above the glideslope. Just before touchdown, the captain instructed the first officer (the flying pilot) to go around. As thrust was being increased, the main wheels contacted the runway. The captain then retarded the thrust levers and the crew commenced manual braking about 7 seconds later. Reverse thrust was not selected.
Touchdown occurred about 1,000 m along the runway and the aircraft entered the overrun at 88 kts. Partial dynamic hydroplaning had occurred during the landing roll. This, along with the absence of reverse thrust, reducing the stopping forces available to slow the aircraft.
The aircraft nose landing gear and one main landing gear separated during the overrun sequence.
The captain ordered a precautionary disembarkation of the aircraft about 20 minutes after it came to rest.
The Bureau is examining the feasibility of a study into the phenomena of low-level windshear to be undertaken by a suitable research institution and involving the Bureau of Meteorology, Airservices Australia, Civil Aviation Safety Authority, Australian Transport Safety Bureau and industry.
Analysis
At the time of the occurrence the environmental wind was strong, and the investigation concluded that it was likely that the downdrafts and associated surface outflows from the entrained convective activity were distorted in the direction of the prevailing airstream, and that this accounted for the gusting conditions that were present at the time of the occurrence.
The flight data recorder fitted to EBS was not equipped to record the aircraft's groundspeed, and the investigation was unable to determine the actual external winds that affected it during the approach and landing. However, from the meteorological data that was available, it was probable that the roll rate encountered by EBS as it commenced the landing flare resulted from an encounter with low-level windshear. It is likely that this was produced by a downdraft from one of the convective storm cells passing through the terminal area at the time.
Although the pilot in command responded in a timely manner with appropriate control input, under the dynamic conditions that were encountered, it is unlikely there was sufficient available aileron/spoiler authority to counteract the high rate of roll that had suddenly been experienced. This resulted in the number 1 engine pod momentarily striking the ground as the aircraft touched down.
Low-level windshear may occur as a result of thunderstorms, land/sea breezes, low-level jet streams, mountain waves and frontal systems. There have been accidents and incidents associated with low-level wind shear in Australia. Pilots should be aware that it is a phenomena that may occur at any location, is difficult to predict, and can present a hazard to aircraft on approach and departure.
Summary
History of the Flight
On arrival at Perth, the crew of a Boeing 747-238B, VH-EBS, were cleared to conduct an instrument approach for landing on runway 03. The pilot in command was the handling pilot for the sector, and the crew subsequently reported that although they were in visual meteorological conditions during the approach, turbulence was encountered. Information "Whisky" was being broadcast on the automatic terminal information service (ATIS), and provided information to the crew that the wind speed and direction at the aerodrome was 330 degrees magnetic at 20 knots. The ATIS included information that the wind speed and direction at 200 ft above ground level was 330 degrees magnetic at 30 knots, and advised crews to expect moderate turbulence below 4,000 ft.
During the approach the co-pilot requested a wind check from the aerodrome controller. The controller advised the crew that the runway 03 threshold wind was 300 degrees at 12 kts, giving a crosswind of 12 kts. The controller requested the crew to advise the spot wind at 1,000 ft, and the co-pilot reported that the 1,000 ft spot wind was 280 degrees at 35 kts.
On short final, at approximately 500 ft above ground level, the pilot in command discontinued the approach when the aircraft experienced turbulence rendering the approach unstable. The co-pilot notified air traffic control (ATC) that EBS was conducting a missed approach, and the controller issued an instruction to the crew to climb to 1,500 ft. The controller then issued further instructions for EBS to climb to 3,000 ft and instructed the crew to take up an easterly heading to intercept the 9 mile arc, from the Perth distance measuring equipment beacon, to position the aircraft for another approach onto runway 03.
As EBS proceeded towards the south to intercept the 9 mile arc for the second approach to runway 03, the controller reassessed the prevailing wind conditions. The wind had been steadily backing to a more southerly direction, and the controller considered that the wind had begun to favour operations on runway 24. The controller notified the crew of EBS that runway 24 was available for landing, and the crew advised that they would accept an approach for that runway. The controller then issued radar vectors to the crew to position EBS onto the approach for runway 24. As EBS was on final approach the controller advised the crew that the threshold wind for runway 24 was 290 degrees at 23 kts, and that the wind at 200 ft was 290 degrees at 35 kts.
The crew reported that the approach to runway 24 was conducted normally and with the autopilot engaged. However, turbulence had prevailed throughout the approach. Flaps 30 was the landing flap setting, and as the aircraft flared for touchdown it suddenly experienced an unexpected roll to the right and the pilot in command applied a control wheel input to the left to counter the roll. The aircraft then suddenly experienced a severe roll to the left. Although the pilot in command applied an immediate control wheel input to the right to arrest the roll, the aircraft touched down in a left wing down attitude, and the number 1 engine pod briefly struck the runway surface.
The crew reported that the touchdown was smooth, and appeared to be on the centreline of runway 24. They also reported being unaware that the number 1 engine pod had struck the ground during the touchdown. As the aircraft taxied in to the international apron a flight attendant advised the crew that a passenger had reported seeing brown fluid leaking from the number 1 engine. After the aircraft had parked the number 1 engine was inspected for damage. The casing of the high speed external gearbox fitted to the engine was fractured adjacent to the gearbox mount position, and the number 1 engine thrust reverser was damaged. The pilot in command then notified the controller that EBS had sustained a podstrike during the landing on runway 24.
The subsequent inspection of runway 24 revealed a scrape mark on the runway approximately 490 metres from the threshold of runway 24. The scrape mark was approximately 30 metres in length and was located approximately 18 metres left of the runway centreline just outside the outer edge of the runway touchdown zone markings. Examination of the manufacturer's data for the B747-200 series showed engine number 1 to be 21.2 metres outboard from the aircraft centreline. This was consistent with position of the scrape mark on runway 24.
Flight Data
Air traffic control radar plots and the flight path derived from EBS's flight data recorder (FDR) were examined during the investigation. They revealed that the pilot in command discontinued the first approach onto runway 03 at 04:07 co-ordinated universal time when EBS was at approximately 500 ft above ground level. Following the discontinued approach onto runway 03, EBS was vectored to the southeast of the airport, then back towards the northeast when ATC reconfigured the terminal airspace for operations onto runway 24. EBS commenced the approach onto runway 24 at 04:23, and the approach concluded at 04:28 when the aircraft landed.
The FDR roll angle plot revealed that as EBS was 35 ft above ground level it commenced an uncommanded roll to the right. The pilot in command immediately applied 29.5 degrees of left control wheel to counteract the roll. However, the roll continued to increase, and EBS was in an 8.0 degrees right wing down attitude as it reached 2 ft above ground level. The roll then suddenly reversed, and within 2 seconds EBS was in an 8.4 degrees left wing down attitude. Although the pilot in command immediately responded with 40.7 degrees of right control wheel to counteract the roll, the aircraft touched down still in an 8.4 degrees left wing down attitude.
Groundspeed was not a recorded parameter on the Lockheed LAS209F FDR that was fitted to EBS, and the investigation was therefore unable to determine the actual wind conditions that it encountered throughout the approach. However, variations in the FDR computed airspeed plot throughout the occurrence sequence were consistent with the reported turbulent conditions.
Aircraft Data
Roll control of the Boeing 747 aircraft is provided by inboard and outboard ailerons and spoilers. The manufacturer advised that a control wheel deflection of 40.7 degrees to roll the aircraft to the right would result in outboard aileron deflections of left outboard +13.8 degrees and right outboard -22.2 degrees. The maximum outboard aileron deflection is +15 and -25 degrees, with +ve signifying trailing edge down and -ve signifying trailing edge up. With the same control wheel deflection of 40.7 degrees to roll the aircraft to the right, the resultant inboard aileron deflections would be left inboard +18.2 degrees and right inboard -17.9 degrees. Normally the maximum inboard aileron deflection is +/- 20 degrees.
The spoilers consist of 12 panels on both wings starting with no 1 on the left outboard wing and extending to no 12 on the right outboard wing. For a control wheel deflection of 40.7 degrees to roll the aircraft to the right, spoilers 1-7 would be deflected 0 degrees (faired with wing), spoiler 8 would be deflected 9.7 degrees, and spoilers 9-12 would be deflected 15.3 degrees. Aileron and spoiler deflection would be reversed for a control wheel deflection of 40.7 degrees to roll the aircraft to the left.
Data for Boeing 747-200 series aircraft fitted with Rolls Royce RB211-524 engines showed that the ground clearance of the number one engine pod was 188 cm at an operating empty weight of 164,610 kgs. This clearance was reduced to 158 cm when the aircraft was at its maximum taxi weight of 352,894 kgs. The plan view of the Boeing 747-200 series aircraft showed the number 1 engine to be 21.2 metres outboard of the aircraft centreline, and 15.2 metres outboard of the outboard wheel of the wing landing gear assembly. Under static conditions and with 0 degrees nose pitch, a body roll of 7.05 degrees at the operating empty weight would cause the number one engine pod to contact the ground. A body roll of 5.93 degrees at the maximum taxi weight would also result in ground contact of the number one engine pod.
The weight of EBS at the time of the occurrence was approximately 230,000 kgs.
Meteorological Information
At the time of the occurrence, Perth was under the influence of an unstable air flow as a result of a complex low-pressure system situated to the south of Western Australia. A series of fast moving cold fronts were embedded in the strong to gale force westerly airstream, and the unstable atmosphere resulted in widespread rain showers, squalls and occasional thunderstorm activity.
The trend type forecasts (TTF's) for Perth from 00:33 leading up to the time of the occurrence indicated that gusty wind conditions could be expected in the terminal area. Additionally, the TTF's from 01:31 indicated that thunderstorms were also likely to be present in the area. The aerodrome forecast current for Perth at the time of the occurrence also indicated the likely presence of gusty conditions and rain showers. At 01:00 an airport warning was issued for Perth containing information that a series of squall lines were expected to cause wind gusts to 45 knots during the day, and that thunderstorms were predicted. At 01:13 information concerning en route weather phenomenon with the potential to affect the safety of aircraft operations (SIGMET) was issued. The SIGMET, valid from 02:00 until 08:00, forecast the presence of severe turbulence below 4,000 ft for the Perth region, and was passed to the operator by the Bureau of Meteorology.
The Perth aerodrome ATIS was changed to information "Whiskey" at 03:03. It provided information that the wind speed and direction was 330 degrees magnetic at 20 kts, and warned pilots to expect moderate turbulence below 4,000 ft. A windshear alert was also provided, with the wind speed and direction at 200 ft being 330 degrees magnetic at 30 kts. The windshear alert was included on the ATIS because there was a 10 kts difference between the wind speed on the ground and the wind speed at 200 ft. The ATIS was amended to information "X-Ray" at 04:27, providing information that the wind speed and direction were 290 degrees magnetic at 25 kts. The revised ATIS continued to provide a warning of moderate turbulence below 4,000 ft and also a windshear alert.
Wind shear is defined as a sudden change in wind direction and/or speed with height or horizontal distance. In most cases wind shear does not present a hazard to aircraft and the majority of pilots will be familiar with changes in wind direction and speed as they ascend or descend. However, at low altitudes (below 1000 ft) during critical stages of landing and takeoff, wind shear can present a significant hazard to aircraft because there is a limited ability to undertake a recovery manoeuvre if the aircraft configuration changes. Low altitude windshear events are small scale and short lived and only affect the approach / departure flight path for a short period of time. The ability to identify such events based on traditional airport observations is limited, although systems which can detect wind shear and provide alerts in a timely manner are available.
The Bureau of Meteorology
anemometer at Perth airport was the source of wind data transmitted on the ATIS. The anemometer sampled the wind at 1-second intervals, and a display of the anemometer wind data was located in the control tower. Controllers were able to select the display for instantaneous, 2-minute average wind speed and direction, or 10-minute peak wind speed. Data from the Bureau of Meteorology anemometer was recorded and archived. The control tower also had displays of threshold wind data obtained from anemometers located adjacent to the threshold of each runway. A display also provided wind data from an anemometer located on the control tower cabin. The controllers could select the threshold anemometer and tower cab displays to provide instantaneous, 2-minute average wind speed and direction, or 10-minute peak wind speed. The controllers reported that their usual practice was to leave the tower cab displays selected to the instantaneous setting, with selection to the 2-minute average wind speed and direction or the 10-minute peak wind speed settings being made to determine the development of any significant trends. Data from the threshold and tower cab anemometers was not recorded and archived.
Radar imagery taken at 20 minute intervals during the occurrence period showed a significant line of enhanced rain echoes passing through Perth at 03:20 in a generally easterly direction at approximately 35 to 40 kts. Scattered convective showers were present behind the line of precipitation, with rain echoes being randomly spaced and largely unorganised. The radar imagery also revealed showers in the vicinity of Perth aerodrome at the time of the occurrence.
The 1-minute data recorded by the Bureau of Meterorology anemometer at Perth aerodrome showed that a fast moving front passed across Perth aerodrome at 04:12, shortly after the pilot in command discontinued EBS's approach onto runway 03. The front was associated with a change in wind direction and a significant increase in windspeed for a period of approximately 3 minutes. At 04:26 the 1-minute anemometer data revealed a marked increase in wind speed which persisted until approximately 04:30. The wind speeds increased to 23 - 25 kts during this period, with maximum wind speeds being recorded at 27 - 35 kt. At 04:28, the time of the occurrence, the maximum wind speed was approximately 29 kts. However, during the period 04:26 to 04:30 there was little variation in the recorded wind direction, and it remained relatively constant from the west. The duration of the increase in wind speed was considered characteristic of an outflow from a convective rainshower.
As a result of this incident, and following contact between the ATSB and the aircraft manufacturer, several temporary revision changes have been made to the aircraft maintenance manuals for the Beechcraft 1900,1900C and 1900D. These revisions detail changes to the fire bottle activation testing procedures, and introduce a check to ensure that sufficient voltage is available at the fire bottle squib to operate the bottle. The manufacturer has also introduced a more secure method of attaching the landing light wiring in this area on all of the new production aircraft. This method involves utilising a length of spirally wrapped electrical insulation tubing around the wiring leading to the landing lights and stand-by fuel pump. An extra cable tie and tubing stand-off is also utilised to further guarantee wiring separation from the fuel lines in the area.
Following discussions with the ATSB, the Civil Aviation Safety Authority issued airworthiness directive, AD/BEECH/1900/30, effective on the 20 September 1999. This AD details the requirement to inspect the affected left and right wing zones on 1900 aircraft for evidence of electrical wiring chafing and rub or burn marks on the aluminium fuel lines.
The Australian Transport Safety Bureau, (formerly the Bureau of Air Safety Investigation) issued the following interim recommendations on the 21 September 1999. The responses to these recommendations, without alteration to the text, are attached to this report.
The Australian Transport Safety Bureau classifies the responses according to the guidelines in the Bureau's Policy and Procedures manual. These response classifications are as follows:
CLOSED - ACCEPTED
ATSB accepts the response without qualification.
CLOSED - PARTIALLY ACCEPTED
ATSB accepts the response in part but considers other parts of the response to be unsatisfactory. However, ATSB believes that further correspondence is not warranted at this time.
CLOSED - NOT ACCEPTED
ATSB considers the response to be unsatisfactory but that further correspondence is not warranted at this time.
OPEN
The response does not meet some or all of the criteria for acceptability for a recommendation that ATSB considers to be significant for safety. ATSB will initiate further correspondence.
IR19990172
The Australian Transport Safety Bureau (formerly the Bureau of Air Safety Investigation) recommends that the Civil Aviation Safety Authority alert all operators to initiate an immediate wiring and fuel-line inspection of the Beech 1900 fleet in Wing Zones 531 and 631 as a matter of the highest priority.
On the 27 October 1999, the following response to IR199900172 was received from the Civil Aviation Safety Authority:
CASA has reviewed Air Safety Interim Recommendation IR 9990172. Your staff briefed the relevant CASA specialist staff on the circumstances surrounding the inflight fire in VH-NTL on 16 September 1999. The serious nature of the incident prompted this Authority to issue AD/BEECH 1900/30, Electrical Loom Inspection, on 17 September 1999, to be effective on 20 September 1999 and requiring an inspection of the area before further flight.
A report has now been received for all aircraft on the Australian register, showing that no similar problems exist in Beech 1900C/D aircraft operated in Australia. None-the -less, the conclusions in your report are generally supported. CASA will seek advice from the manufacturer regarding the appropriateness of the electrical circuit protection and the instructions for maintenance of this wire.
Response classification - CLOSED-ACCEPTED
IR19990173
The Australian Transport Safety Bureau (formerly the Bureau of Air Safety Investigation) recommends that the Federal Aviation Administration alert all operators to initiate an immediate wiring and fuel-line inspection of the Beech 1900 fleet in Wing Zones 531 and 631 as a matter of the highest priority.
On the 7 October 1999, the following response to IR199900173 was received from the Federal Aviation Administration:
The Wichita Aircraft Certification Office (ACO) received the following Safety Recommendation on October 1, 1999:
Safety Recommendation 99.371; "The Bureau of Air Safety Investigation recommends that the Federal Aviation Administration alert all operators to initiate an immediate wiring and fuel-line inspection of the Beech 1900 fleet in Wing Zones 531 and 631 as a matter of the highest priority."
The original 1900s and 1900Cs (serial numbers beginning with "UA" and "UB" respectively) use a fuel bladder versus a total wet wing in the later 1900Cs (serial numbers beginning with 'UC") and 1900D series (serial numbers that begin with "UE"). The specific area of concern for this Safety Recommendation is applicable only to the 'UC" serial numbered 1900Cs and 1900D aircraft models because the components involved in this incident are located elsewhere in the Model 1900s and original 1900Cs ('UA"s and "UB" s). However, there are some wiring and fuel systems components in this general area in these earlier model aircraft ("UA" and "UB" serial numbers) so the review included them as well.
In the FAA's investigation, which included looking at the incident pictures, reviewing new production 1900D aircraft and reviewing the 1900, 1900C and 1900D Maintenance Manuals, the following items were noted:
The incident pictures revealed "tie-wrap" impressions on the plumbing in the area where the arcing is believed to have occurred. This indicates that at one time, the tie-wraps that are used to construct the stand-offs for the electrical wiring were in place.
The 1900 and 1900C Maintenance Manual in Section 5-20-02, page 5, (First 200-hour-interval detailed inspection) for item 13.b. states "LEADING EDGE AND NACELLE PLUMBING AND WIRING Zone inspection areas: 511, 521, 522, 531, 541, 611, 621, 622, 631 and 641. Wing panel inspection areas: 54, 55, 56, 57, 58, 59, 60, 61 (UA-1 and after, UB-1 and after); 23, 24, 25, 26, 27, 28 and 29 (UC-1 and after). b. Check the wiring for chafing and security of attachment." In addition, on the same page of the same section for item 14.a states, "PLUMBING Zone inspection areas: 531, 532, 631 and 632. Wing panel inspection areas: 4, 17, 18 (UA- 1 and after, UB- 1 and after); 8, 9, 11, 12, 15, 18, 21, 23, 25, 29 (UC- 1 and after). a. Visually check for leaks, chafing or damage and proper attachment." This inspection is one of six that are to be repeated every 1200 hours per section 5-20-00, page 3 of the same manual.
The 1900D Maintenance Manual in Section 5-20-02, page 204, (First 200-hour-interval detailed inspection) for item 9.b. states "LEADING EDGE AND NACELLE PLUMBING AND WIRING Zone inspection areas: 521, 621, 522 and 622. Panel inspection areas: 511, 611, 531AB and 631AB. b. Check the wiring for chafing and security of attachment." In addition, on page 205 of the same section for item 14.a states, "PLUMBING Zone inspection areas: 500, 600, 730 and 740. Panel inspection areas. 531AT and 631A.T. a. Visually check for leaks, chafing or damage and attachment." This inspection is one of six that are to be repeated every 1200 hours per Section 5-20-00, page 204 of the same manual.
A FAA representative ran a search on the FAA Service Difficulty Database for "chaffing". What was found were SDR items 162360 and 332583 (there were actually several more, but these were the most relevant). Both of these items appeared to be different (one being in the right outboard nacelle, the other being in the wheel well). The FAA representative did not find any other occurrences of items that resembled this particular scenario.
A FAA representative also visually inspected the new Model 1900D aircraft that were coming off of the assembly line. The two production aircraft that were inspected had the electrical wire stand-offs in place, and the FAA representative concluded that these stand-offs provided adequate clearance to the fuel lines.
The FAA representative that was investigating this incident was not aware of any instances where the plastic tie wraps that are used for stand-offs have failed (without being cut by something).
The FAA concludes that the Maintenance Manuals already provide for wiring and fuel-line inspection of the Beech 1900, 1900C and 1900D fleet in Wing Zones 531 and 63 1. These wiring and fuel-line inspections are required by the same Maintenance Manuals to be repeated every 1,200 hours. The FAA believes that these inspections are adequate and that no additional Airworthiness Directive action is required. Therefore, we recommend this Safety Recommendation be closed.
Response classification - CLOSED-NOT ACCEPTED.
IR19990174
The Australian Transport Safety Bureau (formerly the Bureau of Air Safety Investigation) recommends that Raytheon Aircraft alert all operators to initiate an immediate wiring and fuel-line inspection of the Beech 1900 fleet in Wing Zones 531 and 631 as a matter of the highest priority.
On the 18 November 1999, the following response to IR199900174 was received from the Raytheon:
The attached Safety Communique' No. 164 and Temporary Revision No. 26-1 to the Beech 1900D Airliner Maintenance Manual, P/N 129-590000-15, are for your information.
October 1999
ALL BEECH MODEL 1900 SERIES OPERATORS, CHIEF PILOTS, DIRECTORS OF OPERATIONS, DIRECTORS OF MAINTENANCE, AND ALL RAYTHEON AIRCRAFT AUTHORIZED SERVICE CENTERS, AND INTERNATIONAL DISTRIBUTORS AND DEALERS
MODELS: BEECH 1900, SERIALS UA-2 AND UA-3; 1900C, SERIALS UB-1 THROUGH UB-74, AND UC-1 THROUGH UC-174; 1900C (C-12J), SERIALS UD-1 THROUGH UD-6; AND 1900D), SERIALS UE-1 THROUGH UE-384.
SUBJECT: FIRE CAUSED BY ELECTRIC WIRE CHAFING FUEL LINE AND FIRE EXTINGUISHER TEST
A report has been received of a fire that occurred in the right main wheel well and adjacent outboard wing leading edge area of a Beech 1900D airliner. The event occurred when the aircraft was taxiing to the terminal following a night landing. The fire was quickly extinguished by ground personnel. No injuries were incurred by the flight crew or the passengers.
The fire originated in the equipment bay of the right-wing leading edge, just aft of the landing light (MS. 124.20, F.S. 280.50). The fire spread into the wheel well area before it could be extinguished by ground personnel.
The fire was detected by the crew when the master caution annunciator illuminated, followed by the right AC bus "fail' and the right fuel 'pressure low' annunciators. The crew then observed smoke and flames coming from the right nacelle area at which time the appropriate emergency procedures were initiated. It was later determined that the right engine fire extinguisher system did not function when activated.
Heat and fire related damage was confined to the right main landing gear wheel well area, and some slight damage to the wing equipment bay.
The cause of the fire has been determined to be electrical arcing from an unsecured landing light chafed power wire contacting the transfer system fuel line located behind the landing light. Chafing damage of the wire insulation resulted in wire strands being exposed, thus allowing for electrical arcing to the fuel transfer line causing a fuel leak that then ignited.
Raytheon Aircraft Company is issuing this Safety Communique in order to urge all operators of affected 1900 series airplanes to inspect wiring in the left and right wing equipment bays for signs of distress or damage. Any damaged wiring is to be replaced or repaired. All wiring is to be routed and secured in such a manner as to prevent contact or chafing on any fuel lines, pneumatic lines, equipment and/or structure per best shop practice to maintain no less than 1/4 inch positive separation. The equipment bays may be accessed by removal of wing access panels No. 631 AT, 631 AB, 531 AT, and 531 AB.
The cause of the right engine fire extinguisher not functioning may be due to either a lack of electrical continuity through fire extinguisher "Push To Extinguish" switch or the "Firewall Fuel Valve" control "T" handle. Investigation is ongoing.
Within the next week, Raytheon Aircraft Company will be issuing temporary revisions to CHAPTER 5 -TIME LUTS/MAINTENANCE CHECKS and CHAPTER 26 - FIRE PROTECTION of the appropriate Maintenance Manuals which will establish a required periodic testing of the left and right fire extinguisher circuits. Although this requirement will be added to the appropriate detail inspection, Raytheon Aircraft Company recommends a check be conducted of the fire extinguisher circuit at the next scheduled inspection after receipt of these temporary revisions.
This inspection should be conducted as soon as possible, but no later than the next detail inspection on all effected aircraft over 1000 flight hours total time in service.
Response classification - CLOSED-ACCEPTED.
Significant Factors
The right-wing landing light wiring cable tie stand-offs were not installed.
The right-wing landing light wiring was in contact with the surface of one or more fuel lines in the right-wing equipment bay.
The wiring had electrically arced on the surface of one or both of the fuel lines resulting in holes being made in the fuel line walls with resultant fuel leaks from each line.
The fuel had ignited resulting in fire damage to the adjacent aircraft structure.
The fuel leaks were unable to be stopped by the flight crew.
Analysis
The investigation was unable to determine with any certainty at what time during the flight the fire began. Indications are that the fire was probably not an in-flight fire. The damage to the surrounding structure of the wheel well appeared to indicate plastic deformation and some melting of the aluminium structure, and a probable maximum fire temperature of around 700 o C. The crew had reported good illumination from the landing lights for landing. This indicated that all lights were still operating, and that at that time the mechanical indicating fuse was intact. The pilot's actions in not selecting the standby boost pump to on, following the R FUEL PRESS LOW indication, may have inadvertently been a mitigating factor in the fire. The pump could have supplied the fire with extra fuel under pressure at a critical time.
Several years prior to the incident the right wing de-ice boots had been removed and the wing repaired as a result of hail damage. This repair entailed some disconnection and disturbance of the pneumatic lines and electrical wiring running through the forward area of the wing bay. It is possible that the missing landing light electrical wiring cable tie stand-off had been removed and not replaced at this time. The fact that the cable tie impressions on the pneumatic line fire sleeving were covered with soot from the fire, also suggests that the tie was not in place. The area immediately surrounding the landing light cable tie stand-off was also less severely heat affected than other areas in the zone where the remains of cable ties still existed. For example, in an adjacent area there was still considerable evidence of the cable tie that secured and positioned the standby pump electrical wiring. Had the landing light cable tie been in place at the time of the incident, some remains of it should still have been evident.
As there was no visible evidence of chafing between the landing light wiring and the fuel lines, it is possible that the electric arcing may have been the result of the plastic deformation and subsequent breakdown of the wire's ETFE insulation. This could have occurred while the wiring was in contact with the fuel lines, following the generation of excessive heat in the landing light wiring due to the excessive 'contact bounce' in the K11 relay.
The fuel ignition source may have initiated from one of several sources. For example: the electrical arcing between the electrical wiring and the fuel tubing, arcing of a fuel drenched electrical aircraft component, the main gear up position indicator switch, and/or possible static electricity generated by the fuel escaping from the damaged fuel lines. The fuel 'washing' marks against the upper panel suggests that the fire was not burning in that area. Further, the area immediately surrounding the motive flow line leak only exhibited evidence of heat damage and sooting.
The suitability of the ampere rating of the enclosed link, current limiting fuse was also discussed with the aircraft's manufacturer, due to the fact that the fuse delay had allowed the wiring to arc through the fuel lines. Following a review of the wiring system and the current limiter's rating, the aircraft's manufacturer decided that it was appropriate for the task.
The fire may have started following the holing of the engine supply line in the rear of the wing zone, and spread from there. There was evidence of a well-established fire in this area, with some of the ethylene-tetrafluoroethylene copolymer (ETFE) wiring insulation completely burnt away. It is also possible that the holed fuel lines allowed the fuel to run down onto the landing light gear up position indicator switch. This switch initiating the fire when it was momentarily powered during the landing gear extension cycle.
It is probable that the electrical arcing and fire occurred either immediately prior to, or just following landing. Had this fire been burning in flight it is likely that a more serious outcome would have resulted. The inability of the flight crew to isolate the fuel leaks, together with the extreme heat of an inflight fire, could have resulted in the wing spar losing structural integrity, and a possible in-flight loss of the right wing.
Summary
After landing, and while taxying to the terminal, the co-pilot of the Beechcraft 1900D aircraft turned the landing lights off. He then contacted Air Traffic Control, cancelling SARWATCH. During this radio transmission, the MASTER WARNING and right AC bus (R AC BUS) warning captions illuminated, closely followed by illumination of the right fuel low pressure (R FUEL PRESS LOW) warning.
The crew immediately carried out the company check list actions for the right AC bus failure, but decided not to implement the actions for the right fuel low pressure warning as the aircraft was close to the terminal. The checklist actions for the right low fuel pressure indication required the standby boost pump to be switched on. The co-pilot then detected an acrid smell in the cockpit and alerted the pilot to flames he had observed coming from the underside of the right engine nacelle. The pilot in command immediately brought the aircraft to a stop, shutting down both engines.
Although there was no engine fire warning indication, the crew operated both engine fire handles, making several unsuccessful attempts to discharge the right engine fire bottle. The co-pilot then evacuated the passengers through the forward cabin door, directing them to the flood lit terminal apron area. The pilot in command alerted the RAAF fire personnel by radio of the fire, before turning off the aircraft power and vacating the aircraft.
Two of the operator's maintenance engineers, awaiting the aircraft's arrival, had noticed the flames emanating from the wheel well area as the aircraft approached. They had immediately picked up two dry chemical powder fire extinguishers and approached the aircraft. Following the feathering of the right propeller, they discharged the contents of both fire extinguishers into the right main landing gear wheel well area, extinguishing the fire. The military fire tender arrived soon after to assist.
Investigation
The investigation found that there had been a fuel-fed fire in the area to the rear of the right main landing gear wheel well, and in the right wing equipment bay area positioned immediately outboard of the right engine nacelle. The fire had severely damaged the airframe structure, wiring and components in both areas. The greatest damage was evident in the rear of the wheel well. The aluminium inner fender panel assembly, positioned at the rear of the right wheel well, had partially melted during the fire, leaving a trail of molten aluminium that extended back along the taxiway.
Fuel to the fire had been supplied from two damaged aluminium alloy fuel tank lines in the right wing equipment bay. One line was positioned toward the front of the enclosed equipment bay area, and the other toward the rear. Examination of the surfaces of both fuel lines indicated that they had been in contact with powered electrical wiring. This contact had resulted in electrical arcing, with holes being burnt completely through the walls of both fuel lines. The wiring supplied power to the right, wing mounted, 450 watt landing light.
The damaged fuel line, positioned in the forward area of the bay, was the fuel transfer motive flow line that supplied the operating pressure to the forward fuel transfer jet pump (See Fig 1). The jet pump transferred the fuel from the main wing tanks to the wing mounted collector tank, ensuring a constant fuel supply for the engine driven pump. This line was pressurised with fuel from the engine driven fuel pump, or by the electric standby boost pump when that pump was turned on.
The damaged fuel line in the rear of the bay was part of the main fuel supply from the wing mounted collector tank to the engine driven pump. The line was positioned between the fuel filter shut-off valve, just forward of the wing main spar and the fuel filter assembly.
The standby electric pump, located in the bottom of the collector tank, served as a backup to the engine driven pump in the event of a failure of that pump, and could be manually selected on by the flight crew. The pump also activated automatically during a normal engine start sequence.
Low fuel pressure from either the engine driven pump or the standby pump was indicated by the illumination of the left or right fuel pressure low (L or R FUEL PRESS LOW) warning annunciator.
The fuel leak in the engine driven pump supply line was stopped by maintenance personnel soon after the incident, by manually closing the fuel shut-off valve positioned on the front of the fuel collector tank. There was no mechanical method of isolating the leak from the motive flow line. The fuel flow from this line was temporarily stemmed by the fitting of a rubberised electrical wiring loom clamp over the hole in the line.
The electrically activated right engine firewall shut-off valve was found to be in the open position.
Electrical
The landing light system operating voltage was 28 volts DC. The 450 watt landing light receives its power from the K11 relay, positioned in the rear of the right engine nacelle area. The relay contactor switching wiring was protected by a 10 amp circuit breaker positioned in the aircraft underfloor area. The electrical wiring leading to and from the K11 relay to power the landing light, was protected from a prolonged overcurrent situation by a 35 amp mechanical indicating, enclosed link, current limiting fuse. This device was utilised to allow a transient high current draw that would occur during the landing light initial illumination. Examination of the fuse revealed that the mechanical indicating pin had triggered. This indicated that the internal fusible link had melted.
The landing lights were normally switched on during descent at the transition altitude of 10,000 ft. This was done as a part of the operator's transition checklist actions. In this instance, the crew advised that the lights were turned as the aircraft descended through 11,300 ft. No abnormal operation of the lights was noticed at any time, with good illumination of the runway for landing.
The electrical wiring was examined and found to be of the correct specification as detailed by the manufacturer. The wiring had a copper core with the insulation surrounding the wire manufactured from white extruded ethylene-tetrafluoroethylene copolymer (ETFE). The surface of the wiring and aluminium fuel line tubing was microscopically examined, with no evidence found of any rubbing on or around the arcing points.
The aircraft manufacturer indicated that the normal method for the positioning and securing of the electric wiring, in areas such as the right-wing equipment bay, was by utilising plastic cable ties (See Fig 3). The ties would be routed through lengths of plastic tubing that acted as cable stand-offs. These were to used to securely position the electrical wiring, and ensure that it did not come into contact with the adjacent fuel lines.
An inspection of other Beechcraft 1900 aircraft in the operator's fleet revealed cable tie and tubing stand-offs securing the landing light wires. These were attached around the fire sleeving on the wing de-ice boot pneumatic lines in the forward area of the bay.
The inspection of the area around the motive flow fuel line on the accident aircraft, revealed that one of the two landing light wires was in contact with the surface of the fuel line. No plastic stand-offs were fitted to space the landing light wiring away from the fuel line. There was however, evidence of a soot-covered imprint of a plastic cable tie on the surface of the fire sleeving (See Fig 4). This fire sleeving surrounded the adjacent wing de-ice boot pneumatic line.
The remains of another plastic stand-off, on the nearby positioned standby fuel pump power wiring, was still evident (See Fig 5).
The aircraft had been subjected to severe hail damage in 1995. During the repairs following this damage, the right-wing leading edge de-ice boots were removed and the pneumatic de-ice lines were disturbed.
The landing light wiring, positioned above the damaged main fuel supply line, was manufactured longer than required. This excess wiring had then been doubled back on itself and tied, with a cable tie, along the main wiring loom into this area. This wiring had also been in contact with the fuel line.
When examined the contacts on the K11 landing light relay exhibited signs of arcing due to excessive 'contact bounce'. 'Contact bounce' is an oscillation of the relay contacts. This condition can be exaggerated as a result of the effects of heat and in-service wear on the springs and latches that control and damp the relay contact movement. With excessive 'contact bounce' present, it is possible for the wiring served by the relay to heat up due to a continually higher than normal current flowing through the wiring. Following removal of the landing light wiring insulation by the ATSB, there was evidence seen of excessive heat on the wiring surface. Plastic deformation of the wiring's ETFE insulation was noted at the point of arcing on the fuel line positioned at the front of the bay.
Fire
The fire damage was most severe in the rear of the wheel well area, this was evidenced by the melted aluminium alloy panel, the distorted wheel well surrounding structure and some destroyed ETFE wiring insulation.
The fuel, from the leaking equipment bay lines, had flowed along the face of the wing spar towards the wheel well area as the amount of fuel increased. The fuel was then able to flow onto the top of the right main gear up position indicator switch, through a hole in the inner fender assembly panel immediately above. The wheel well area was open to the airflow and to the propeller wash at the rear lower end of the inner fender panel.
Air was also drawn from behind the fender panel by the inverter cooling fan positioned in the rear of the engine nacelle area. This cooling air flowed from the wheel well through the rear of the equipment bay, and was ducted along under the right side of the engine nacelle cowling. The duct exhibited signs of heavy sooting and some of the inverter cooling fan plastic blade tips had melted.
The area at the rear of the right-wing equipment bay, through which the inverter cooling air was drawn, was the most heat-affected area of the bay. Some of the ETFE wiring insulation in this area was completely burnt away. The forward end of the bay, in the area of the other fuel leak, exhibited some wiring damage and medium to heavy sooting. In this area there was also evidence of 'washing' of fuel against the upper access panel (See Fig 6). The wiring and components that were immediately adjacent to the fuel leak in the area were not as heat affected as in other parts of the zone.
View of upper surface of right wing, showing fuel 'washing' on upper access panel
A typical hydrocarbon fuelled ground fire would burn at a temperature in the range of 870 0 C to 1093 oC. An inflight fire, with the added oxygen, would burn in excess of 1093 oC, often up to 1371 oC and higher. The aluminium alloy in aircraft becomes plastic at 454 o C and melts at about 677 oC, while the extruded ETFE insulation on the electrical wiring, melts at approximately 300 o C. The cable ties are made from either Teflon or nylon and have a similar melting point of 280 o C to 300 oC
Engine isolation and fire extinguishing system The aircraft was equipped with an engine isolation and fire extinguishing system that was activated by the operation of a fire emergency tee handle. The operation of the fire handle cuts fuel to the selected engine and arms the engine fire extinguisher bottle. The pilot can then discharge the fire extinguisher by depressing an instrument panel mounted switch.
During the shutdown of the aircraft prior to passenger evacuation, the flight crew had activated the engine fire handles and attempted to discharge the right engine fire bottle several times. The fire bottle, however, had not discharged and the fire wall fuel shut-off valve had not closed. An inspection of these emergency systems revealed that the components had not operated because of fire damage that had occurred to the electrical wiring to these components. The right side firewall shut-off valve circuit breaker was found to have tripped during the incident. Regardless, the operation of the system would not have had any effect in this instance, due to the fire being in an area outside the engine fire zone.
Following an inspection of the manufacturer's maintenance procedures for the aircraft type, it was discovered that there was no procedure for determining the serviceability of the fire bottle activation system that ensured there was sufficient voltage at the fire bottle electrical connection to activate the bottle. This has been brought to the attention of the aircraft manufacturer.
As a result of its investigation into this incident, Airservices Australia suggested the following safety actions:
"Sydney Tower experiment with the use of blocking strips for aircraft crossing runways to see if a satisfactory method of usage can be found which is beneficial to controllers."
"Team Leaders discuss with their teams the relevance of surface movement controllers taxiing aircraft, which require runway crossings, by taxiways which provide the ADC with the optimum view of the aircraft and present the best opportunities for expeditious runway crossings."
Australian Transport Safety Bureau action
As a result of this and other occurrences the Australian Transport Safety Bureau, formerly the Bureau of Air Safety Investigation, is currently investigating a safety deficiency. The deficiency relates to the use of conditional clearances for runway entry and runway crossings by vehicles and aircraft and procedures used by air traffic controllers to alert themselves that vehicles or aircraft are on an active runway.
Any recommendation issued as a result of this deficiency analysis will be published in the Bureau's Quarterly Safety Deficiency Report.
Significant Factors
The ADC did not adequately scan the runway prior to issuing a takeoff clearance to the crew of the Saab 340 or immediately before the takeoff was commenced.
The ADC forgot about a conditional clearance issued to the crew of the B767 to cross the runway at taxiway Lima.
Analysis
In establishing that the take-off path of an aircraft was unobstructed, controllers were required to make two separate visual observations of the take-off path; one before issuing the take-off clearance and another before take-off was commenced. In addition, controllers needed to have a high level of situational awareness about the movement of other aircraft on the airfield, particularly aircraft subject to conditional clearances to cross or enter an active runway given to aircraft under the control of the SMC. In this particular incident, the controller relied on short term memory to maintain awareness of the clearance issued to the B767.
Memory prompts can assist controllers to maintain situational awareness. Items such as pens, blank flight progress strips or the like were commonly used by controllers to act as memory prompts, but their use was inconsistent in application. Use of some form of memory prompt in this particular incident may have helped to maintain the controller's situational awareness and enhanced the effectiveness of the controller's visual scanning of the take-off path.
Summary
The aerodrome controller (ADC) had given a conditional clearance for a Saab 340 to line up on runway 16R behind a landing Boeing 737 (B737). The ADC then gave the surface movement controller (SMC) a conditional clearance for a Boeing 767 (B767) to cross runway 16R at taxiway Lima when clear of the landing B737.
The SMC issued the clearance for the B767 to cross the runway as the B737 vacated the runway at taxiway A4. The ADC observed the B737 vacate the runway at taxiway A4 and cleared the Saab 340 for take-off. The pilot of the Saab 340 rejected the take-off clearance and advised the ADC that there was a B767 crossing the runway.
The Manual of Air Traffic Services (MATS) 6-2-3 paragraph 31 stated:
"Before clearing an aircraft for take-off, and immediately before the take-off is commenced, the tower controller shall make a visual check from the control tower to determine, as far as practicable, that the take-off path is not obstructed." The ADC made a visual check of the runway, however, he only scanned between the runway 16R threshold and the point where the B737 was vacating the runway at taxiway A4. The ADC did not continue the visual check through to the upwind end of the runway. The B767 was crossing the runway at taxiway Lima, which was between taxiway A4 and the upwind end of the runway. The ADC had forgotten about the conditional clearance given to the SMC for the B767 to cross the runway.
There was no standard practice in Australia for the use of "blocking strips" or "memory prompts" by controllers to alert them of the presence of aircraft not under their direct control crossing or entering an active runway. In this particular incident, the ADC did not use, nor was he required to use, a memory prompt to remind him of the conditional clearance given to the SMC for the B767.
As a result of this occurrence, the BoM's analysis and report of the meteorological aspects of the occurrence included the following recommendations:
As BoM's observation stations at Parafield and Edinburgh will be closed before the winter of 2000, BoM should consider installation of a "Skycam" on a city building to better appreciate the extent of fog and low cloud when the conditions that led to this occurrence are present. In addition, BoM should urgently consider a research project on guidance material for prediction of fog events at Adelaide and Edinburgh airports.
As BoM's observing site is poorly located at Adelaide Airport in reference to fog and low stratus cloud to the north of the airport, BoM should consider relocation of the site to eliminate this impediment.
BoM should conduct a workshop for local forecasters on fog events and local guidance before the winter of 2000.
On 16 December 1999, BoM reported:
BoM was considering "Skycam" installations on a national basis, expecting that trials would be conducted in the eastern states due to higher traffic levels.
The relocation of the BoM's observation site at Adelaide Airport was included in the airport upgrade. In addition, a radio was installed in the BoM's airport office enabling forecasters to monitor the Automatic Terminal Information Service (ATIS) from Adelaide, Parafield and Edinburgh.
Research and workshops were an ongoing requirement in BoM, but depend upon staff availability.
On 19 July 2000, BoM reported:
BoM assessed the range of Skycam units available and obtained the agreement of the aviation industry to fund the installation of one unit at a major airport (not Adelaide). The industry has undertaken to assess the value of the information obtained from that unit before making any decision on possible funding of further units.
The BoM offices at RAAF Edinburgh and Parafield closed in December 1999. Since those closures, an automatic weather observation station was installed and has been operating at RAAF Edinburgh. The tower controllers at Parafield have also been providing some observations.
The relocation of the BoM office at Adelaide Airport was scheduled for December 2000, but could be delayed until early 2001. The proposed new site for the office was closer to the runways than the present office and was expected to provide better views of fog areas and low cloud than were available from the present site.
It has not been possible to hold workshops due to staffing limitations.
Summary
As the aircraft approached runway 23 for landing, the crew observed a bank of fog drifting toward the aerodrome from the north-east. By the time the aircraft arrived at the aerodrome, the runway threshold was obscured by the fog. As a result, the crew elected to conduct a missed approach.
During the missed approach, the crew noticed that the threshold area of runway 05 was clear, so they requested an immediate visual approach to runway 05 before the fog drifted further to the south-west. Due to other instrument flight rules traffic, Air Traffic Control (ATC) could not issue an immediate clearance for the approach. By the time that clearance was available, the remainder of the runway was obscured by fog. A B737 aircraft had been able to land on runway 05 following a VOR/DME approach, so the A320 crew attempted to conduct a similar approach. However, that attempt resulted in a second missed approach. The aircraft tracked to the north-east of the aerodrome and the crew informed ATC that they would conduct an instrument landing system (ILS) approach to runway 23, and then land using the aircraft's autoland system. With 1,500 kg of fuel remaining, the aircraft landed without incident in the fog. Visibility was 250 to 350 m.
The aircraft was certificated for autoland approaches, but the ground equipment was not. The ILS transmitter was a Category 1 unit with a minimum visibility of 1,200 m required for landing. The crew decided to conduct an autopilot-coupled approach with automatic landing, as fog was also present at the Royal Australian Air Force base at Edinburgh, rendering that aerodrome unsuitable as an alternate. The crew considered that Whyalla, the nearest suitable aerodrome, was likely to have similar weather conditions to Adelaide.
Fog had not been forecast for Adelaide when the crew submitted their flight plan. Consequently, the aircraft did not carry fuel for holding at Adelaide or for diversion to an alternate.
However, fog had been forecast for both Edinburgh and Parafield. The Bureau of Meteorology (BoM) reported that this was not unusual, as records showed that in the past 20 years, fogs formed at both Adelaide and Edinburgh on about 50% of occasions, with Edinburgh proving to be the greater risk. On the day of the occurrence, moisture levels were higher to the north of Adelaide, with fog forming at Edinburgh at 0700 Central Standard Time. What was unusual about this event was that the advection of fog from the north took place at a greater speed than the surface wind and that the onset time of fog at Adelaide Airport was 40 minutes later than any recorded onset time at that location in the past 30 years.
BoM records showed that Adelaide Airport averaged 4.9 fog events per annum. The highest annual total for events was nine, recorded in both 1956 and 1983. At the time of the incident on 20 August, there had been 11 fog events recorded at Adelaide Airport during 1999.
The crew of a SAAB, conducting a scheduled passenger service from Griffith to Sydney, broadcast that they were taxiing to depart from runway 06. The pilot of a Cherokee advised that he was on downwind for runway 36. The crew acknowledged this transmission and then established the position and intentions of the pilot of a Dromader who was on an extended downwind leg for a low-level approach to runway 06. The crew of the SAAB then advised that they were entering and backtracking on runway 06. Approximately 90 seconds later, when the crew advised they were rolling on runway 06, the pilot of the Cherokee responded that he was on late finals to runway 36. The crew continued their take-off and overflew the landing Cherokee by a reported 400 feet.
Each aircraft was in radio communication on the Griffith common traffic advisory frequency of 126.7 MHz.
The crew of the SAAB later reported that they had not heard the pilot of the Cherokee respond to their taxiing broadcast. Their attention had been directed toward the pilot of the Dromader who had adjusted his approach to assist their departure. They had not seen the Cherokee and consequently it was not until the pilot reported on late finals to runway 36 that they realised there was a traffic conflict. The crew reported that at this time the SAAB had accelerated to a speed such that rejecting the take-off was potentially more hazardous than continuing.
The higher terrain to the south of the aerodrome may have made the Cherokee more difficult to detect against the background. Additionally, a line of trees to the south of runway 06 obscured the final approach path to runway 36 from the view of pilots at the 06 threshold.
The reason why the crew of the SAAB did not recall hearing the response of the Cherokee pilot to their taxiing report was not determined. However, it is likely that the decision to expedite their departure ahead of the arriving Dromader created a self-imposed high workload that led to a loss of awareness of the Cherokee.
As a result of this incident and chafing found on the galley connectors of another B737 aircraft, the operator issued Engineering Release E/R B73-24-00-22 on 30 September 1999. This required the operator to inspect all B737 aircraft for chafed or burnt wiring on the galley power connectors and to reinstall insulation tape if required. The inspection showed that four other aircraft needed to have the galley connector wiring insulation retaped with none exhibiting any chafed or burnt wiring damage.
Abstract
Performance monitoring of the auxiliary power unit (APU) was being carried out during ground maintenance of the Boeing 737-300. Maintenance staff reported that the electrical system voltage had surged twice and that the aircraft interior lights had dimmed before voltage and frequency indications returned to normal. No circuit breakers had tripped. When maintenance personnel in the aircraft cockpit noticed a burning smell, they shut down the APU to investigate the fault.
The forward No. 2 galley ovens were found to be inoperative. During further troubleshooting, ground power was applied twice, which resulted in smoke emanating from a ceiling access panel above the forward No. 2 galley assembly.
The fault was traced to the electrical connector supplying 115VAC three-phase power to the forward No. 2 galley ovens. The connector and associated wiring behind an overhead panel were burnt. The seven-pin connector had a five-wire bundle comprising phases A, B, C, neutral D and static ground. The three-phase wires had desoldered from the pins on the connector. Arcing damage was found on the endbell and cable clamp halves of the connector.
The insulation blanket and galley panels next to the connector were blackened and heat-damaged. Charring of the insulation material had occurred over an area of about 75 mm x 100 mm. The film covering the blanket had burnt for a further 150 mm to 250 mm.
Figure 1: VH-CZB damaged insulation blanket
The galley manufacturer and operator established that arcing and heating of the connector and wiring was due to inadequate soldering and insertion of the phase wires into the connector-pin buckets. Only the static ground-wire connection was found to meet the manufacturer's soldering process standard.
The connector at the other end of the damaged cable was also examined. The solder terminations of phases A, B, C and neutral D were also found to be substandard, with solder smeared on the outside of the contact pin wall and only a small amount of solder bonding the cable strands to the contact pin.
The phase and neutral wire-identification numbers stamped on the insulation were of a different size and orientation from those supplied by the galley manufacturer (as shown on the static ground wire insulation). The wire-identification number stamps matched those held by the operator.
The arcing damage found on the endbell and cable-clamp halves of the connector had no matching wire-bundle insulation damage, showing that the damage had occurred before this incident. The resulting damage to the cable probably explains why the phase A, B, C and neutral D wires had been replaced and re-identified.
Forward No. 2 galley was the original fitment on this aircraft, delivered in September 1986. The operator reported that disconnection of the subject connector would occur only during scheduled galley removal at each D-check. The sole D-check on this aircraft occurred in July 1994. The operator reviewed maintenance and pilot discrepancies raised against electrical power and galley equipment systems since aircraft delivery, but did not find any failures requiring loom replacement. Similarly, a review of D-check work documents showed no galley electrical system discrepancies.
The insulation blanket and film-like covering from above the forward No. 2 galley were heat damaged as a result of the electrical connector fault. The operator reported that the blanket-covering material (Orco Film AN-26) did not appear to be flame retardant to the level of FAR 25.853. The operator believed that the accumulation of organic material on the surface of the film over 13 years, especially corrosion inhibiting compound mist, would have changed the film's flammability performance. As a result of the accident involving Swissair flight 111, the U.S. Federal Aviation Administration is researching the flammability of thermal acoustical insulation materials.
