Aviation Investigation Report A98H0003
1.14.10 Assessment of Fire Damage (STI1-95)
- 18.104.22.168 - Aircraft Skin Panels
- 22.214.171.124 - Forward Airframe Structure Condition
- 126.96.36.199 - Air Distribution System Cockpit and Cabin
- 188.8.131.52 - Air Recirculation System
- 184.108.40.206 - Left and Right Forward Passenger Cabin Individual Air
- 220.127.116.11 - Forward Galleys
- 18.104.22.168 - Forward Lavatories
- 22.214.171.124 - Cockpit and Passenger Cabin Material
- 126.96.36.199 - Cabin Ceiling Panels
- 188.8.131.52 - Cockpit Ceiling Liner and Dome Light
- 184.108.40.206 - Avionics Circuit Breaker Panel
- 220.127.116.11 - Overhead Circuit Breaker Panel
- 18.104.22.168 - Heat and Fire Damage Pattern Aircraft Wires and Cables
The construction of the upper crown of the aircraft consisted primarily of 2024 aluminum alloy stressed skin exterior panels, rivetted to 7075 aluminum alloy airframe structure. The aircraft skin was painted on the outside with white exterior paint and painted on the inside with green FR primer.
There was no indication that the fire burned through the aircraft skin at any place along the fuselage, nor was there any indication of discolouration of the white exterior paint. There were varying degrees of soot accumulation on the interior surface of some of the aircraft skin panels in the area of the fire.
Forward of STA 475, insulation blankets were placed directly against the aircraft skin between the frame structure. A layer of over-frame insulation blankets was then added to cover the entire area. Between STA 475 and STA 755, no between-frame insulation blankets were installed, only the over-frame insulation blankets.
In some areas of fire damage, the lack of soot accumulation on the aircraft skin panels suggests that the insulation protected them from the fire, particularly where double layers of insulation blankets were installed. In those areas, it appeared that the between-frame insulation blankets remained in place until the aircraft struck the water.
There were two areas of predominately heavy soot accumulation on the aircraft skin, located between STA 401 and STA 420 from X= 25 to X= 25, and between STA 475 and STA 555 from X= 35 to X= 50. Between these two areas, from STA 420 to STA 475 located between X= 40 and X= 30, there was an area that had a mixture of light to moderate soot accumulation. This area is within the area above the forward passenger entry door tracks and operating cables. Another area of moderate soot accumulation was located from STA 374 to STA 401 between X= 45 to X= 25. Outside of these areas, the amount of soot on the recovered skin portions ranged from light to none.
Most of the airframe structural components were painted with green FR primer prior to their assembly, and were covered with insulation blankets during the aircraft construction (see Figure 4). For the inner surfaces of the frames and intercostals to become exposed to either soot or heat damage, the insulation blankets must first be compromised.
Material heat testing showed that a 10-minute exposure to temperatures below 204°C (400°F) did not discolour the FR primer; therefore, recovered structure with no discolouration of the FR primer was useful in establishing the boundary between heat-damaged and non-heat-damaged areas.
The airframe structure in the cockpit attic area with the most heat damage was forward of STA 366 between intercostal planes 15 left and 15 right (see Figure 24 and Figure 25). The intercostals appear to have created a barrier that, for the most part, impeded the fire from spreading in an outboard/downward direction. On the left side, heat damage extended outboard of intercostal plane 15, on the STA 366 frame, for approximately 10 to 15 cm (4 to 6 inches). On the right side, heat damage extended outboard of intercostal plane 15 on frame planes 10, 11, and 12 for approximately 30 cm (12 inches). This area is directly behind, and above, the upper avionics CB panel.
Between STA 366 and STA 401, the most significant heat damage pattern was along the crown of the aircraft between the plane 15 left and right intercostals. The severity of the heat damage increased toward STA 401. This area is above the cockpit centre overhead air diffuser, and immediately aft of the cockpit door.
The most forward airframe structure with heat damage resulting in bare metal was at STA 353 at X= 20, and the most rearward was on the STA 535 frame between X= 31 and X= 48. Between these two station locations, there were 17 additional frames or intercostals with similar patterns of heat damage.
The most forward area with feathered-edge damage was the STA 374 frame between X= 17 and X= 9, and the most rearward was at the STA 515 frame between X= 15 and X= 22. In the area between these two frames, 17 additional frames or intercostals showed feathered-edge damage.
Mechanical fractures with broomstraw-like appearance were found at 16 separate locations on frames and intercostals between frame STA 374 and frame STA 466. Broomstraw was also noted on the R1 passenger door forward track at approximately STA 427 between X= 29 and X= 31.
The area of the fire damage above the ceiling in the front of the aircraft contained a network of primarily aluminum air ducts that were part of the aircraft's air distribution system (see Figure 3, Figure 8, Figure 24, Figure 26, and Figure 27). The ducts from this area were reconstructed from small pieces that were straightened, identified, fracture matched and sewn together with locking wire to replicate their original shape. Most of the aluminum ducts in the fire-damaged area were painted with FR primer prior to installation. In service, they were wrapped with thermal acoustic insulation blankets. The rebuilt air ducts provided information about the boundaries of the fire and the intensity of the heat.
Heat damage to the ducts that provided air to the cockpit ranged from no damage to heat damage resulting in bare metal. There was bare-metal heat damage starting near the top of Galley 1 and running forward to the cockpit manifold. A portion of a duct located above the cockpit door at approximately STA 396, X= 19 and Z= 72 had several resolidified aluminium deposits on the outer surface. The precise alloy of this aluminum could not be determined.
There was bare-metal heat damage on the ducts behind the cockpit air diffusers from STA 350 aft to STA 402, primarily between the plane 15 left and right intercostals. The exception was the window diffuser distribution duct, which had bare-metal heat damage along STA 392 to X= 32. The heat damage to the recovered portions of the diffusers varied, with higher heat on the upper surfaces and lower heat on the lower surfaces adjacent to the cockpit ceiling liner.
Damage to the riser duct assembly ranged from no heat damage to bare-metal heat damage. The vertical portion of this assembly had no heat damage. The first area of bare-metal damage to the riser duct assembly started at intercostal plane 15 right, where the lower surface had a region of bare-metal heat damage from X= 30 inboard to the joint at X= 20. This was in the vicinity of the Galley 2 vent end cap. There was bare-metal damage from approximately STA 395 to STA 442 between X= 25 and X= 10. A section of molded duct was installed in that area from STA 420 to STA 442 between X= 5 and X= 25. No portions of this molded duct were identified in the wreckage.
Between approximately STA 480 and STA 545, the main conditioned air ducts for zones 2, 3, and 4 transitioned from running near the top of the forward cabin drop-ceiling, upward to run near the crown of the aircraft. In this area there was bare-metal heat damage at places along the top of the ducts. The primer on the underside of these ducts was not damaged by heat. The recovered portions of two individual air ducts from approximately STA 555 to STA 595 at X= 70 and X= 70 had areas of light to moderate soot accumulation.
Recirculation air was supplied by four fans located above the passenger compartment ceiling at STAs 685, 725, 1009, and 1109 at X= 28 (see Figure 24). Each fan drew air through a filter and plenum assembly located at the corresponding station from X= 40 to X= 65. The plenum assemblies, which had been painted with FR primer on their aluminum parts, were reconstructed.
The plenums at STA 685 and STA 725 had no discolouration of the FR primer. They did have localized areas with heavy soot accumulation. The recovered portions of the fibreglass filter elements had dark grey colouration on one side and light grey colouration on the opposite side. The hoses connecting these plenum assemblies to the fan housings had localized light soot accumulation on the outer surfaces but had no soot accumulation on the interior surfaces. The recovered portions of the plenums at STA 1009 and STA 1109 had no heat damage or soot accumulation.
The recirculation duct was uninsulated between STA 569 and the recirculation fan. A check valve was installed in the duct to prevent reverse airflow when the fan was not operating. The reconstructed recirculation duct portions had been installed between the STA 685 fan and the cabin conditioned air duct, just forward of the muffler, at approximately STA 555. The duct had bare-metal heat damage on the uninsulated sections aft of STA 569.
The individual air supply to the centre section of the forward cabin was provided by two ducts, which were connected to a recirculation air duct at STA 575, X= 29 and STA 672, X= 24. Both individual air supply ducts showed bare-metal heat damage.
Forward passenger cabin individual air was supplied by a fan and plenum assembly identical to a recirculation air fan and plenum assembly. The fan was located at STA 990 at X= 24. The plenum assembly was between X= 40 and X= 65. The recovered portions of the plenum had no heat damage or soot accumulation. The individual air ducts were uninsulated and ran forward from STA 990 to a "Y" (split) at STA 750 at Z= 91, X= 21. One branch of the "Y" ran across the crown to the left side of the cabin at X= 76, and the other ran to the right side at X= 76. The recovered portions of these ducts had moderate to heavy soot accumulations on the exterior, with no heat damage. The aftermost section of individual air duct that was identified was from STA 934 to STA 955 at X= 22 and had light soot accumulation on the exterior.
Galleys 1, 2, and 3 were installed forward of the first-class cabin. The outer surfaces of the top of these galleys were exposed to the forward cabin drop-ceiling area through a cut-out in the ceiling panels. Identified portions of these galleys were reconstructed to examine their exposure to the fire environment.
Galley 1, which was installed on the left side of the cabin between the cockpit aft wall and the L1 door, had heavy soot accumulation and heat damage on the top outer surface (see Figure 27). Other components of Galley 1 had varying degrees of soot accumulation, particularly near the top of the unit where it had been exposed to the fire environment. The electrical equipment in the forward upper compartment of Galley 1 did not show any fire or soot damage. Wires installed inside the galley were not affected by the fire. A four-wire harness, which was routed through an access hole in the top of the galley, had a length of white spiral wrap that had localized soot accumulation. The wires in the harness had localized light-brown discolouration.
Galley 2 was installed on the right side of the cabin between the cockpit aft wall and the R1 door. No pieces of the outer surface of the top of this galley were identified. Pieces of Galley 2 from below the forward cabin drop-ceiling were identified and did not show heat damage or soot accumulation.
Galley 3 was installed in the forward-centre position in the aircraft, immediately aft of the L1 and R1 doors between STA 470 and STA 508. There were localized areas of light soot on pieces of Galley 3 on or near the top of the unit. The portions of the wires that were installed on the upper surface of the galley top showed some soot accumulation, whereas those wires installed within the galley were free of soot. Wires associated with the Galley 3 disconnect assembly had soot accumulation ranging from trace to heavy, and showed areas of heat damage.
The forward galley vent duct assembly comprised a single, uninsulated aluminum duct with branch connections to the three forward galleys (see Figure 3 and Figure 8). The upstream end of the vent duct was located above the forward cabin drop-ceiling above Galley 3. The duct ran horizontally forward above the conditioned air ducts to Galley 2, and continued laterally across the fuselage to a point above the top of Galley 1. The vent duct then ran vertically down the left side of the fuselage, outboard of Galley 1, to an area below the cabin floor. The duct continued aft under the cabin floor along the left side of the aircraft ending at the cabin air outflow valve located just forward of the left wing root. The forward galley vent system utilized a pneumatic jet pump operated by bleed air from the Pneumatic System 1. The jet pump was installed at approximately STA 872 to provide a single source of constant vacuum for the forward galleys, both in flight and on the ground. This installation provided a constant flow of between 200 and 400 CFM at the jet pump that was exhausted overboard through the outflow valve at STA 920.
From the upstream end of the forward galley vent duct, a branch duct extended vertically down to connect with the air intake grill near the top of Galley 3. A similar vertical branch extended from the vent duct toward the top of Galley 2; however, this branch was not connected to Galley 2 and the branch was closed off with a silicone elastomeric end cap (see Figure 4 and Figure 6). The vent connection to Galley 2 was not made because Galley 2 was not electrically powered, nor was an oven installed. A ceiling-mounted air intake plenum was installed in the corridor outside the cockpit door, near the inboard edge of Galley 1. The air intake plenum was connected to the forward galley vent duct by a 8-cm (3-inch) diameter hose that was routed across the Galley 1 ceiling. This hose connected to the vertical portion of the forward galley vent duct between the left fuselage and the outboard face of Galley 1. A second hose, measuring 5 cm (2 inches) in diameter, extended from the outboard face of Galley 1 and connected to the vertical portion of the galley vent duct at STA 398, X= 48, Z= 60, just below the 8-cm (3-inch) hose. This second hose drew air (odours) from Galley 1 into the vent duct.
Two segments of the forward galley vent duct were identified; both were located on the vertical section of the vent duct that was routed between the left fuselage and the outboard face of Galley 1. The first segment was located at the cabin floor level (Z= 18); the second segment was located below the point where the 8- and 5-cm (3- and 2-inch) diameter hoses connected to the vent duct (Z= 11 to Z= 42). Both of the above-floor duct segments sustained high heat damage. Identified portions of the galley vent system from below the cabin floor, between STA 396 and STA 457, exhibited localized areas of moderate heat damage. Identified portions of the vent duct, located below the floor and aft of STA 457, exhibited no heat damage.
The Lavatory A module was installed on the left side of the aircraft forward cabin between approximately STA 465 and STA 495, immediately aft of the L1 door. A portion of the top of Lavatory A was recovered with a wire harness still attached (see Figure 27). Localized soot accumulations were noted on both the attic and cabin facing surfaces of the portions recovered. The attached wire harness was also sooted. Few additional pieces of this lavatory module were recovered and identified. Of those that were identified, there were no signs of heat damage.
The Lavatory B module was installed on the right side of the aircraft forward cabin between approximately STA 460 and STA 496, immediately aft of the R1 door. Only three pieces of this lavatory module were identified. There was soot accumulation on the cabin-facing side of these pieces, which were from the upper portion of the module. There was soot accumulation on the Lavatory B cabin placard.
There were no signs of heat damage or arcing on any of the wires associated with the forward lavatories. Light soot accumulation was noted on some of the wiring. The smoke detector control panel for these two lavatories, which was located above the forward cabin drop-ceiling over Galley 1, showed soot accumulation and heat damage.
There is no indication that either smoke alarm activated before the recorders stopped. The available information indicates that the fire did not start within one of the forward lavatories.
Materials from within the passenger cabin were examined for fire-related damage. One business-class seat cushion assembly, complete with seat cover, had several locations where melt-like features and discolouration were evident, consistent with the drop-down of hot materials. None of the melt-like features appeared to penetrate the underlying cushion. Sections from some fabric curtains used in the cabin had damage consistent with being exposed to heat, including discolouration, and stiffness or roughness or both. There were also areas where melted and resolidified material was adhering to the curtain fabric, including an MPET-covered insulation blanket, a blue-green material consistent with being from an electrical module block, and a material consistent with 7075 aluminum alloy. These curtains were most likely located in the aisle between lavatories A and B, and Galley 3.
Complete passenger comfort blankets and portions of these blankets were recovered. When not in use, these blankets are stored in overhead bins in the passenger cabin. Some blanket material had minor heat-related damage.
Areas of forward galley flooring and forward cabin carpet near the forward lavatories had localized areas of heat damage consistent with the drop-down of hot materials. A tiny portion of aluminum alloy that appeared to have melted and resolidified was found in a piece of forward galley flooring that had been installed along a wall. The exact type of alloy could not be identified.
One portion of cabin carpet had an area of numerous small holes through the carpet pile and backing, which were consistent with having been caused by drop-down of hot materials. This area was from approximately STA 472 to STA 505, between X= 26 and X= 46, with the highest concentration from approximately STA 482 to STA 493. This corresponds to the aisle area between Galley 3 and Lavatory B.
In the cockpit, there were numerous locations where localized heat damage had taken place. Microscopic fibre analysis confirmed that the spotted areas had been generated by fire drop-down damage. Attempts to recover traces of the drop-down material to identify it were unsuccessful, as it appears they were dislodged at the time of impact. The source for a large extent of the fire drop-down damage was most likely the cockpit ceiling liner melting and dropping down onto the carpet (see Section 22.214.171.124).
Several deposits were found on the right observer seat. Small amounts of resolidified 2024 aluminum alloy were deposited on the lap belt, and on the right side of the seat. When the resolidified metal was removed from the lap belt for analysis, a white deposit remained. Other white deposits were observed elsewhere on the same lap belt. Trace analysis of these deposits disclosed that they were primarily aluminum oxide, and microscopic fibre analysis revealed fused heat damage material at each location. The white deposits were determined to be the remnants of other locations where resolidified aluminum had also been deposited. It was not possible to determine the alloy of the aluminum from the remnants. A small amount of resolidified 6061 aluminium alloy was found on the rear of the right observer seat base.
Two cockpit checklist booklets were recovered with the aircraft wreckage. Both were heat damaged. In these checklists, each double-sided page is contained in a plastic sleeve, and the pages are bound together along one edge. Each sleeve can be rotated about its bound edge and turned over and under the booklet so that one page is visible on one side of the booklet and the next page in the sequence is visible on the other side of the booklet.
One of the booklets had more heat-related damage than the other. Some of the edges of the plastic sleeves had been partially melted and fused together, fixing the booklet in the open position to pages 10 and 11. Page 10 was the Smoke/Fumes of Unknown Origin checklist, and Page 11 was the Smoke/Fumes Removal checklist. Page 11 was more significantly heat damaged than Page 10. The heat pattern appeared to be from the outside surface inward, suggesting the booklet was in a horizontal position with Page 11 upward at the time it was heated. On the second checklist booklet, a small burn mark was found at the top edge of Page 1 (Index) and extended through to Page 4 (ENG 2 A-ICE DUCT). Two mating heat-related damage marks were found on Page 2 (INTENTIONALLY LEFT BLANK) and on Page 3 (ENGINE FIRE).
Ceiling panels were used to separate the passenger space from the attic area throughout the cabin portion of the aircraft. The panels were suspended from the structure of the aircraft fuselage using suspension rods, beams, and attachment hardware. All the panels were manufactured as phenolic/glass skins, bonded to meta-aramid fibre paper honeycomb-like core. Some of the panels had white bondable PVF adhered to one side only and some showed it on both sides. A decorative PVF laminate was adhered to the face of the panels that were visible from the passenger compartment.
Three types of ceiling panels were used to construct the passenger cabin ceiling and another four types were used to construct the forward cabin drop-ceiling. Of these four, the CD 207 type panel was used to construct portions of the overhead bins and the two sliding ceiling panel assemblies at the forward doors. Similar types of panel construction were used to fabricate the close-out panels, forward cabin drop-ceiling, header panels and bridge/gap assembly panels.
Portions of the various panels were recovered and examined for soot and heat damage. As ceiling panels had been fractured into many pieces by impact forces, their installed locations in the aircraft could not be positively identified. Of the recovered pieces, many showed signs of heat damage or soot accumulation or both. The heat damage varied from discolouration to severe charring of the panel core. Most of the heat damage was on the attic side of the panels but some displayed heat damage on the cabin side as well. This could be considered an indication that the fire had penetrated the ceiling in these areas.
Some of the panels could be identified as to type by their construction. One piece of CD 207 panel was determined to be either a portion from one of the sliding forward door panel assemblies in the forward cabin drop-ceiling area or a panel portion from one of the overhead stowage bins in the first-class cabin. The damage to this piece was consistent with exposure to a temperature of 593°C (1 100°F) for 10 minutes. Four recovered pieces were identified as portions of bridge/gap cover assemblies in which the aisle and emergency lights were installed. These portions had areas of dark-brown discolouration in the shape of a half moon, which coincided with the installation location of the aisle and emergency lights. Another four of the recovered pieces were determined to be portions of the overhead stowage bins. Three of these portions showed no indication of heat damage. The fourth portion, identified as part of an overhead stowage bin ramp air duct, showed a soot pattern similar to that deposited on an adjacent panel.
Cockpit ceiling liners, constructed of a light grey thermoformable low-heat-release sheet material, are installed as the interior finish surface of the flight compartment. The material has a low forming temperature; that is, it melts within a relatively low temperature range. It begins to soften and sag at 246°C to 274°C (475°F to 525°F). The five sections that comprise the cockpit ceiling liner are attached to the aircraft structure with screws and nutplates.
The overhead liner is installed immediately aft of the overhead CB panel (see Figure 9 and Figure 10). The liner includes the cockpit dome light assembly and has openings for the air conditioning system diffusers and diffuser controls. Some of the dome light components displayed light soot deposits; the identified portions of the dome light assembly showed a few signs of heat damage. Only a small number of pieces of the centre overhead liner were identified. Heat damage on the fuselage-facing surface of the pieces was indicated by a dark brown to black discolouration. The cockpit-facing surface of the pieces had localized heat damage indicated by hints of taupe discoloration. The heat also caused thinning, necking, surface melting, and wrinkling of the material. The surfaces showed localized light to moderate soot, wrinkles, bubbles, and edge melting.
The left and centre-left liners enclose the area bound by the left edge of the overhead switch panel housing, aft to the ceiling panel and then bordered the dome light and left diffuser outlet aft to the cockpit coat closet. The lower edge of the liners follow the top of the clearview and aft windows back to the cockpit coat closet. The left liner has cut-outs to accommodate two air conditioning slide controls, individual air supply, captain's map light and speaker control box, observer's map light and on/off dimmer control, audio and microphone jack panels and three inspection panels. The cut-out for the captain's map light and speaker control box is located in the forward lower corner of the left liner. The left liner also incorporates an escape rope compartment door, located adjacent to the captain's air conditioning slide control. The centre-left liner has cut-outs for the spare lamps compartment door and an individual air supply.
The majority of the left liner was identified and reconstructed. The fuselage-facing surface of the liner had heavy soot deposits over most of the surface, along with heat damage. The heat damage varied from dark brown to black discolouration with localized small wrinkles and surface melting. The cockpit-facing surface had moderate to heavy soot deposits from the forward to the aft end of the liner, with localized areas exhibiting surface wrinkles caused by heat. The microphone/headset hook, located approximately one third of the way from the front of the liner, displayed moderate soot accumulation on the surface. A piece of the escape rope compartment door had no heat damage. The only piece of the centre-left liner that was identified was a small portion of the spare lamps compartment door that had remained attached to its hinge assembly. The cockpit-facing surface had melting and bubbling with dark brown and black discolouration. The surface that faced into the Spare Lamps compartment had localized light soot deposits.
The right and right aft liners enclose the area between the right edge of the overhead switch panel housing, the upper avionics CB panel, lower avionics CB panel, and the cockpit video monitor. The lower edge of the liners follow the top of the clearview and aft windows back to the video monitor. The right liner has cut-outs to accommodate the air conditioning slide control, individual air supply, first officer's map light and speaker control box and, audio and microphone jack panel. The cut-out for the first officer's map light and speaker control box is located in the forward lower corner of the right liner. The right liner also incorporates an escape rope compartment door, located adjacent to the first officer's air conditioning slide control.
The majority of the right and right aft liners were identified and reconstructed. The fuselage-facing surface of the right liner had light accumulations of soot, along with several areas of heat damage. The heat damage varied from light to dark-brown discolouration near the microphone hook, to black discolouration with blisters and surface melting near the overhead CB panel. The cockpit-facing surface had localized areas of light soot accumulation. The heat discolouration varied from dark grey, to dark grey with light taupe in localized areas. There was slight blistering of the surface. The right aft liner displayed light soot only, with no heat damage.
The avionics CB panel consisted of upper and lower panels, located above the work table at the right observer's station (see Figure 12). The main body of each panel was constructed of aluminum. The upper avionics CB panel had five rows of CBs, identified alphabetically from A to E. Individual CBs were numbered sequentially, starting from the left, and were identified according to the row in which they were installed (e.g., CB D1). The lower avionics CB panel had one row of CBs, identified as Row F. The front (inboard) face of both the upper and lower avionics CB panels was painted grey, and the back (outboard) face was not painted.
About 75 per cent of the upper and lower CB panels were identified, reconstructed, and placed in the reconstruction mock-up (see Figure 27 and Figure 28). The most forward sections of the upper and lower CB panels were not recovered. The upper avionics CB panel had been exposed to heat coming from both the front and back sides. The lower CB panel did not show any heat damage. The heat damage pattern on the front face of the upper CB panel was shown by the change in colour of the paint. Parts of the panel displayed damage consistent with temperature reference exemplar coupons exposed to temperatures from 430°C to 620°C (800°F to 1 150°F) for 10 minutes. Discolouration and heat damage was present on the back side of this panel on some electrical components and on some bare-metal surfaces. Although there were similarities at some locations, the damage pattern on the back side was less than on the front side, particularly on those pieces near the forward end of the panel where the front side of the panel showed considerably more heat damage than the back side. Individual CBs showed soot on the white indicator ring; soot could only have been deposited if the CB had tripped and subsequently been exposed to combustion by-products. Details concerning the CBs on the avionics CB panel and wiring in the vicinity are included in separate sections of this report.
It was noted that the heat damage on parts of the upper avionics CB panel was consistent with damage seen on temperature reference coupons that were exposed to temperatures from 427°C to 620°C (800°F to 1 150°F) for 10 minutes.
The overhead CB panel is located just aft of the overhead switch panel in the cockpit ceiling, above and behind the captain's and first officer's seats (see Figure 12). The panel has seven rows of CBs identified alphabetically from A to G. Like the avionics CB panel, individual CBs are numbered sequentially, starting from the left, and are identified according to the row in which they were installed (e.g., CB A1). An integrally illuminated, polycarbonate lightplate base assembly was installed on each row of CBs.
Most of the CB panel was identified, reconstructed, and placed in the reconstruction mock-up (see Figure 27 and Figure 29). On the cockpit-facing surface, approximately two thirds of the polycarbonate lightplate base on Row A had been melted and folded back over onto itself, forming a fused mass of material near the top right corner of the panel. Part of the right end of the lightplate base on Row B was also melted and fused into this mass. The top right corner of the panel where the paint was missing had heat damage consistent with temperature reference coupons that were exposed to temperatures from 427°C to 621°C (800°F to 1 150°F) for 10 minutes, and the area where the paint was discoloured had heat damage consistent with temperature reference coupons that were exposed to temperatures from 343°C to 398°C (650°F to 750°F) for 10 minutes. The rear surface of the panel also showed signs of localized high-heat damage at the top right corner.
Individual CBs showed soot on the white indicator ring. Soot could have been deposited if the CB had tripped and subsequently been exposed to the smoke environment.
All of the wire segments identified as being from the heat-damaged section of the aircraft were compared to exemplar wires before being incorporated into the reconstruction mock-up. The exemplar wires were created by heating them at specific temperatures for specific times in a controlled heat environment. When the exemplar wires were heated, it was observed that the ETFE wires were more susceptible to heat than the polyimide wires. This was consistent with the damage observed on the wire segments that were recovered. The examination of the wire segments from the heat-damaged area helped define the heat pattern and boundaries of the fire. There were wires from the area of the fire that had no discernible damage. There was a range of damage on other wires, from light soot, to complete melting and destruction of the wire insulation.
The wires that were routed between the upper avionics CB panel and the avionics disconnect panel had areas of localized soot accumulation, and some minor heat damage. The bus feed wires that were routed between the upper avionics CB panel and the upper main CB panel had some areas of localized light soot accumulation, and some minor heat damage, near the upper avionics CB panel. The ETFE wiring that was routed above this area, closer to the aircraft structure, showed more pronounced heat damage.
The IFEN-related PTFE 8 AWG jumper wires from the forward end of the lower avionics CB panel were not heat damaged. Polyimide wires in the same area near the lower avionics CB panel showed soot accumulation, but no heat damage. Based on the appearance of the surrounding area, it was considered likely that the IFEN-related 8 AWG and 12 AWG wires that were routed from the aft end of the lower avionics CB panel upward toward the bottom of the avionics disconnect panel had little heat damage. The portions of these wires that were located near the ceiling structure area, including where the 12 AWG IFEN wires entered the conduit, were heat damaged.
 In the Swissair Product 99 interior configuration applicable to the occurrence aircraft (version M1130), Galley 2 was not electrically powered, and no oven was installed.
 Thermoformable plastics or thermoplasts are moldable when heated; they harden when cooled but soften again during subsequent heating.
 The paint disappears from the surface of the temperature reference coupons within this temperature range. It is probable that exposure to higher temperatures for shorter durations could also generate the same damage.
 No colour discrimination could be made on this material when heated to any temperature within this range of temperatures. It is probable that exposure to higher temperatures for shorter durations could also generate the same discolouration.
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