Status:	Schlussbericht
Datum:	18 MAI 2001
Zeit:	11:20
Flugzeugtyp:	Airbus A300B4-605R
Fluggesellschaft:	Monarch Airlines
Kennzeichen:	G-MONS
Werknummer:	556
Baujahr:	1990-02-14 (11 years 3 months)
Triebwerk:	2 General Electric CF6-80C2A5
Besatzung:	Todesopfer: 0 / Insassen: 15
Fluggäste:	Todesopfer: 0 / Insassen: 262
Gesamt:	Todesopfer: 0 / Insassen: 277
Sachschaden:	schwer beschädigt
Konsequenzen:	Repaired
Unfallort:	nahe Casablanca (Marokko)
Flugphase:	Während des Fluges (ENR)
Betriebsart:	Internationaler Linienflug
Flug von:	London-Gatwick Airport (LGW/EGKK), Großbritannien
Flug nach:	Banjul-Yundum International Airport (BJL/GBYD), Gambia

Unfallbericht:
The aircraft was en route from London (Gatwick) Airport to Banjul when in stabilised cruise conditions at FL290, abeam Casablanca, there was a sudden onset of very noticeable vibration accompanied by some thrust loss from the No. 2 engine (a GE CF6-80C2A5 engine) . This was confirmed by the No. 2 engine N2 vibration indication rising rapidly to 5.8 units. Manual throttle was selected and maximum continuous thrust was applied on the No. 1 engine and the No. 2 engine throttle was retarded to idle. The N2 vibration decreased to just below 5.0 units then increased to 5.6 units before settling. A 'PAN' call was made and a precautionary diversion to Faro was initiated. The No. 2 engine throttle was kept at idle and the engine parameters were carefully monitored but these remained normal, with the exception of the N2 vibration. An uneventful overweight landing was completed at Faro with the fire service in attendance. The aircraft was taxied off the runway and the No. 2 engine shut down. The fire service confirmed that there was no sign of fire or obvious damage, however it was subsequently noticed that the No. 2 engine cowl was holed. The origin of the engine failure was established to be separation of an HPT Stage 2 rotor blade, which had broken off just above the blade platform. The fracture had propagated from the edge of a 0.25" deep notch, which had been worn in the blade leading edge. Similar notches were found on all of the Stage 2 blades. Several HPT Stage 2 nozzle segments around the 12 o'clock position were found to have large cracks in the aerofoil to outer platform fillet. The cracking had weakened the cantilevered nozzle segments and allowed them to deform rearwards under aerodynamic loading and contact the leading edges of the Stage 2 blades.

CONCLUSIONS: "The failure of the No 2 engine was caused by thermally induced fatigue cracking of the HPT Stage 2 NGV aerofoils around the aerofoil outer fillet. This allowed the NGV segment at the 12 o'clock position to 'sag' backwards and 'machine' notches in the HPT Stage 2 blades, ultimately causing the separation of a single blade due to fatigue and tensile overload. All of the HPT Stage 2 NGVs installed on the engine were post-G29/G30 modification standard, the standard which are prone to cracking of the aerofoil outer fillet. An additional factor was the fact that many of the NGV sections, but in particular the NGV segment at the 12 o'clock location, had been multiple-repaired, which further increased the probability of outer fillet cracking.
According to Service bulletin CF6-80C2 S/B 72-0952 Revision 4, the inspection of the HPT Stage 2 NGVs could be performed at the operator's convenience. The HPT on the G-MONS No 2 engine was boroscope inspected 237 cycles prior to the incident, but the Stage 2 NGVs were not inspected at that time. It is possible that, had the NGVs been inspected on this occasion, the aerofoil outer fillet cracking might have been identified and the incident could have been prevented. The engine manufacturer's recommendation that operators inspect the non-G29/G30 nozzles was not assigned a higher level of significance as, prior to the G-MONS event, there was no significant history of engine failures due to HPT Stage 2 NGV aerofoil outer fillet cracking.
The HPT case withstood the initial event of the release of a single HPT Stage 2 aerofoil and therefore met the engine certification requirements in this respect. However, LPT case penetrations in the planes of the LPT Stages 1 and 4, and LPT case rub-through in the plane of the LPT Stage 5, subsequently occurred due to the cascade of damage caused by engine debris travelling rearwards through the engine. This and previous events give cause for concern that, whilst the CF6-80C2 LPT case may satisfy the certification design requirement to contain a single blade-off, it cannot, in some cases, contain the debris subsequently generated within the LPT case as a consequence of an initial failure further upstream in the engine. This is particularly true for the LPT Stage 5 case rubthrough condition, which was an unforseen failure mode and for which the LPT case was not designed. In-service experience of actual failures suggests that further improvements in LPT containment capability are desirable, and that these need to address the LPT case rub-through failure mode."

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