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Summary

Summary

Sequence of events

On 17 November 2003, at 1630 UTC, as Boeing Company 747-438 (747), registered VH-OJI, rotated during takeoff from Changi International Airport, Singapore, the crew heard a loud bang and the aircraft yawed left. The aircraft's instrument indications were consistent with a failure of the number-2 (left inboard) engine. The crew shut down the engine and, after dumping fuel, returned to Changi.

An initial inspection found that there had been a failure within the engine's first-stage high-pressure compressor assembly (HPC 1). Following removal from the aircraft and return to Australia, the engine was disassembled at the operator's maintenance facility, where the failure was confirmed as the loss of one blade and the extensive thermal and mechanical damage of the remaining blades from the HPC 1 disc. The liberated blade stub was recovered from the engine confines.

Engine

The engine was a Rolls-Royce RB211-524G2-T-19 model, serial number 13211 (see figures 1 and 2). The designation 'T' in the engine model number indicated that the original engine had been upgraded with power-train components from the larger Rolls Royce Trent 700 series engine. That enabled greater fuel efficiency and lower exhaust gas temperatures. At the time of the failure, the engine had operated for a total of 50,847 hours and through 6,659 cycles. Module 41, the engine's high-pressure compressor and turbine system, was fitted at the last engine refurbishment on 18 August 2000 and had operated for 14,166 hours and through 1,456 cycles since installation.

Components received

The HPC 1 blades (including the liberated blade) were removed from the engine by the operator and forwarded to the ATSB for examination. For the purposes of the investigation, the blades were identified by numbering in a counter-clockwise direction (looking rearward), commencing from the blade adjacent to the position of the liberated blade. The blade numbered 21 was not received. All examinable blades carried the part number FK28595 H124, embossed on the underside of the root section.

Blade condition - retained blades

Preliminary examination of the HPC 1 blades that remained on the compressor disc during the failure showed that all sustained extensive damage to the tips and edges such that the blade profiles resembled that of an arrowhead. The melting and loss of the leading and trailing edge corners of the blades (figures 3 and 4) is characteristic of a titanium fire within the assembly. Evidence of metallisation was noted on the blade aerofoil and platform surfaces. A close visual inspection of the blade root section revealed no evidence of cracking around the dovetail corners, nor was there any evidence of isolated or general mechanical damage such as may be sustained during handling or installation.

A study of the condition of the dovetail bedding surfaces found that all exhibited varying levels of surface galling damage across the full width (figure 5), with the damage appearing most pronounced toward the rear (outflow) end of the blade root. Low power stereomicroscopy study of the galling showed a predominantly axial orientation to the damage (transverse to the axes of the dovetail faces).

Several areas showed heavy localised galling, producing a gouging effect with notable metal loss (figure 6). Evidence of frictional heating with associated surface tinting (blueing) surrounded the heavily galled area.

Review of the distribution of dovetail galling found that some blades showed light galling up to and encompassing the transition radius between the upper dovetail edge and the root body (figure 7).

Failed (liberated) blade

The remnant stub of the single blade liberated from the compressor disk, despite extensive damage, showed the 'arrowhead' form produced by the loss of the leading and trailing edge corners (figure 8). Although the blade root fracture features were damaged beyond recognition, the similarity in form to previous failures indicated a high-cycle fatigue cracking mechanism.

The blade root showed the apparent fracture of the complete length of the dovetail toe along the leading (concave) side of the blade and roughly one-half of the toe length on the trailing side (figure 9). Damage and abrasion sustained following the blade loss prevented any useful examination of the fracture surfaces. However, the fracture profiles did not exhibit any associated plastic deformation or bending of the blade root body as could be expected if an external load had been applied to the blade, through impact with a foreign object or other internally liberated component.

Cracking and subsequent failure of the blade's dovetail root corners allowed the blade to move radially outward from its slot, under the influence of centrifugal operating loads. As the blade contacted the compressor housing, the resulting friction initiated the titanium fire that melted the blade corners, before the blade completely released from the slot.

The aerofoil section of the blade stub showed uniform transverse bending in a direction opposing the rotation of the compressor disc (figure 10). The layer of metallisation and melted debris on the inside of the bend showed a degree of cracking and fissuring that suggests deposition prior to the bending of the blade.

Service information

The engine manufacturer was aware of seven similar HPC 1 blade release failures in RB211-524G/H-T series engines worldwide. The ATSB had previously investigated one of those occurrences (see ATSB report BO/200205895). The engine manufacturer was also aware of a failure involving a Trent 700 series engine.

On 6 August 2003, in response to the first five failures, the engine manufacturer published Service Bulletin RB211-72-E181, applicable to the occurrence engine, that introduced a revised dry film lubricant on the stage-1 high pressure compressor blade root. The reason for issue of the service bulletin was the inadequacy of the earlier lubricant in preventing incomplete blade bedding and uneven wear, leading to subsequent high-cycle fatigue cracking and potential failure of the blade root. Accomplishment of the Service Bulletin was required to be denoted by a change in the blade part number, from FK28595 to FW26617. Compliance with the Service Bulletin was listed as 'Recommended', with suggested accomplishment when the engine or engine module was disassembled for refurbishment or overhaul.

The engine manufacturer also indicated that scoring and sharp edge damage in the blade root area, leading to local stress concentrations, may have contributed to HPC 1 blade failures. To prevent that damage, the manufacturer revised the manufacturing process, changing the blade root machining process from broaching to milling.

 
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