AAIB Bulletin No: 3/99 Ref: EW/C97/9/6

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Cessna T303 Crusader, G-PUSI AAIB Bulletin No: 3/99 Ref: EW/C97/9/6 Category: 1.2 Aircraft Type and Registration: Cessna T303 Crusader, G-PUSI No & Type of Engines: 2 Continental TSIO-520-AE piston engines Year of Manufacture: 1983 Date & Time (UTC): 28 September 1997 at 1515 hrs Location: 15 miles east of Oxford (Kidlington) Airport Type of Flight: Private Persons on Board: Crew - 1 - Passengers - None Injuries: Crew - None - Passengers - N/A Nature of Damage: Left engine destroyed and minor airframe damage Commander's Licence: Private Pilot's Licence Commander's Age: 58 years Commander's Flying Experience: 2,900 hours (of which 1,600 were on type) Last 90 days - 60 hours Last 28 days - 25 hours Information Source: AAIB Field Investigation History of the flight Whilst in cruising flight at 2000 feet, the pilot heard and felt a 'bang' from the left engine. The aircraft lurched to the left and the pilot saw heavy smoke coming from the top and bottom of the left engine nacelle. He initiated the shutdown procedure for the left engine but then observed that the left propeller was stationary and found that the propeller control lever would not move to the full feather position. However the pilot then managed to carry out a successful single engine approach and landing at Oxford Airport. Strip examination of the left engine The engine, which had accumulated 167 hours and 40 minutes over 233 flights (sectors) since overhaul, was shipped by an aircraft maintenance organisation at Oxford Airport to the agency that had carried out the last overhaul of the engine. Following strip examination by the overhaul agency, during which the crankshaft was found to have separated into three sections, the AAIB was informed of the accident. AAIB examination A detailed metallurgical examination was conducted on the failed components. This revealed that the crankshaft had initially failed by a fatigue mechanism resulting from service induced cyclic stresses at the rear end of the No 6 crankpin (Figure 1 and Photograph 1). The failure was at an inclined angle to the horizontal axis of the crankshaft, which had allowed the crankshaft to continue to run. However after some further aircraft flights, with the crankshaft failed at the rear of No 6 crankpin, a second failure had occurred at the rear end of No 5 crankpin. Immediately following this second crankshaft failure, the engine crankcase had ruptured resulting in the propeller and the front of the engine, including the No 6 cylinder, separating from the remainder of the engine. Despite this massive failure, all engine parts were retained within the cowlings. Detailed visual and macroscopic examination of the surfaces of both the primary and secondary crankshaft fractures indicated the following: (a) The initiation region on both fracture faces of the primary fracture at the rear end of the No 6 crankpin had been mechanically damaged after separation. This mechanical damage had destroyed all evidence within the crack initiation region, precluding determination of the cause of the fatigue crack initiation. From the nature of the mechanical damage to these primary fracture surfaces it was evident that the engine had completed a small number of cycles/flights, probably of the order of five, after the crankshaft had fractured at this position. (b) From a fatigue progression band count on the primary fracture surface of the No 6 crankpin (Photograph 2) it was determined that approximately 150 +/- 20 'events' (ie engine cycles/flights) had occurred between the time that a fatigue crack 6 mm long had been present in the initiation region and the time that full fracture had occurred. However, it was not considered probable that the size of the crack, when the engine was last overhauled, would have permitted detection using the magnetic flow fluorescent particle crack detection technique which had been used. (c) Less than twenty 'events' (engine cycles) were associated with the secondary fatigue failure at the No 5 crankpin, indicating that this cracking had started before complete separation had occurred at the No 6 crankpin. A microsection was taken from the crankshaft material and examined. This revealed that the steel had a fine tempered martensitic microstructure, which appeared metallurgically satisfactory, and the shaft had been surface hardened by a nitriding process. The pistons, piston pins, connecting rods, counter balance weights and the main, big end and small end bearings were examined in the premises of the overhaul organisation to determine whether any reciprocating or rotary components could have been malaligned and thus adversely affected the fatigue resistance of the crankshaft. No evidence of malalignment was found, all unusual wear observed having occurred rapidly, probably during the final flight. The rear end faces of the big ends of the connecting rods had been forced against the adjacent faces of the cranks at the primary fracture position of the No 6 crankpin, and also at the No 3 crankpin. Associated very rapid friction-induced wear had taken place resulting in local temperatures which had approached 300°C. Manufacturer's examination The stripped engine, together with the failed sections of the crankshaft, were therefore forwarded to the engine manufacturer for further examination. The manufacturer concluded that the main fracture had occurred outside the No 6 crankpin fillet radius which, from their experience and knowledge, was not normally an area subjected to high stress. The manufacturer concluded that the failure had initiated due to hard contact between the crank face and the side of the No 6 connecting rod which had induced localised high stresses. Their assessment of the No 6 crankpin fracture indicated that the crack had been present for a large number of flights, probably in excess of 100. A more precise count was not possible due to the damage of the origin area. However the