Cracked No. 2 engine combustion chamber outer case
Cracked No. 2 engine combustion chamber outer case
Keywords Aircraft, Aerospace corrosion,Engines, Safety
A Boeing 747-136 departed from Heathrow for New York. Shortly after take off, at an estimated height of 100 feet and IAS of 172 kt, the flight crew was alerted to a No 2 engine problem by the illumination of an amber caution light on the No 2 Exhaust Gas Temperature (EGT) gauge. The light signifies EGT exceeding 915°C(engine shutdown required if EGT exceeds 940°C). The flight engineers immediately reduced the thrust lever setting almost simultaneously,noting that other No 2 engine parameters had suffered a major decrease and informed the other crew members that the engine had failed. The crew shut down the engine, completing the Engine Fire Checklist immediate actions by around 500 feet, and continued the climb to FL 100. After jettisoning fuel over the English Channel, an uneventful three engine landing was made back at Heathrow.
Initial examination revealed a 69 inch long circumferential crack around the No 2 engine combustion chamber outer casing (CCOC). The engine manufacturer of the JT9D-7A engine,Pratt & Whitney, considered that the resultant disruption of the gas flow was likely to have caused the engine to surge.
Figure 3a JT-9D Combustion chamber outer case engine schematic
Figure 3b JT-9D Combustion chamber outer case L-flange schematic
The engine is a conventional twin spool modular turbofan (Figure 3a). The CCOC forms part of the K-Module; it is a circular case, approximately 41 inches in diameter and 10.5 inches long, that surrounds the annular combustion chamber and forms the main structure of the engine carcass in this area (Figure 3b). It is constructed of Inconel 728, a nickel-chrome-iron alloy. An internal flange (the L flange)formed at the front of the CCOC is bolted to the diffuser case and an external flange (the M flange) at the aft end is bolted to the high pressure (HP) turbine case (Figures 4a and 4b). Under normal operating conditions the CCOC reaches a maximum temperature of approximately 550°C and has a maximum internal pressure of 300 psig.
The CCOC has two basic configurations, the No 2 engine having one of the earlier fabricated versions,with a dual radius internal fillet machined between the L flange and the case wall. For this case the required material thicknesses are 0.19-0.20 inch for the flange and 0.118-0.138 inch for the wall over most of its length, increasing to 0.290-0.350 inch at the forward end. The minimum permissible forward and aft radii for the L Flange fillet was 0.025 inch and 0.057 inch respectively. Each engine is fitted with a dual loop nacelle overtemperature system. Each loop operates a flight deck temperature gauge and a nacelle fire warning system. A flight deck test of the system is scheduled prior to each engine start.
Figure 4a Allowable dimensions for fillet radii of combustion chamber outer case L-flange
Figure 4b Combustion chamber outer case
Powerplant examination
No clear signs of abnormal heating or other damage to the components mounted in the area of the CCOC crack were apparent. Both loops of the No 2 nacelle overtemperature systems failed electrical continuity checks carried out after engine removal from the aircraft. Individual components checked satisfactorily and the system passed the checks after re-assembly. Checks of the No. 2 engine EGT indicating system and of the No. 2 engine oil system, including magnetic chip indicators (MCDs), filter debris and oil sample analysis, revealed no abnormalities.
The LP spool of the engine was tight to turn and fine metal fragments were evident in the exhaust system. Bulk strip examination of relevant engine components at the engine overhauler under AAIB supervision revealed significant internal damage, including heavy rotational rubbing between the blade tips and the outer airseal of the HP turbine and appreciable rotational rubbing between the LP and HP shafts. Both rivets attaching a locking plate for the HP turbine retaining nut had been sheared off and the nut had unscrewed approximately 10o. The severed rivets had lodged between the LP and HP shafts and were partially responsible for the damage to them but the shafts had also made direct contact with each other while rotating. Appreciable abnormal damage was present to the HP turbine first stage nozzle guide vanes (NGV) consistent with overtemperature effects,with major leading edge burn-through of six vanes and deposition of re-solidified material on the surface of most of the vanes. Assessment by an overhaul agency found no evidence of cooling passage airflow anomalies and concluded that the melted material had been present for only a brief period before the engine had ceased to operate.
Examination of the CCOC revealed that the crack ran around the L Flange fillet extending over 195oof the circumference, from 050o to 245o (orientations throughout are relative to the top of the case, measured clockwise as viewed from the rear). It completely penetrated the section, separating the L flange from the case wall over the 69 inch length of the crack, and in places a gap of around 0.25 inch had opened up. Fluorescent dye penetrant inspection (FPI) of the unfractured part of the L Flange fillet by the engine overhauler, after removal of the CCOC from the module and cleaning of the surface, indicated the presence of a number of unopened cracks. However, subsequent examination of a representative section from one of these areas by the Structural Materials Centre (SMC) of DERA at Farnborough found no sign of additional cracking. It was noted that a removable surface deposit probably could have given false FPI indications.
Further assessment of the fracture was undertaken by SMC, with a representative of the engine manufacturer present, followed by examination of the case in the USA by the engine manufacturer. L Flange fillet forward and aft radii were 0.0287 and 0.0788 inch respectively, both of which exceeded the minimum requirements, and the case wall and flange thicknesses were also within limits. The material was found by Dispersive X-Ray analysis to be consistent with Inconel 718, the hardness was within requirements, and the microstructure of a sample section was consistent with the required material in its properly heat treated state.
Optical examination of the separated fracture showed that the surfaces were generally clean and bright,typical of a fresh final fracture, but also exhibited a discoloured area over an approximately 5.5 inch circumferential length that had apparently been present for a period with the engine operating. This initial fracture passed through the forward radius of the fillet and was oriented approximately radially. It exhibited two distinct morphologies; a relatively smooth, blue-gold surface over an approximately 0.115 inch depth from the radius surface, and a red-brown, more woody textured surface continuing for an additional 0.095 inch. The fracture terminated in a non-oxidised tensile shear lip through the flange chamfer.
Detailed examination of the initial fracture by Scanning Electron Microscopy (SEM) revealed striation marks in the blue-gold surface region, indicative of fatigue crack progression. The red-brown region did not show striations and was likely to have been a region where the fracture had progressed at a much higher rate. Surface detail near the origin of the fatigue crack had been obliterated by severe rubbing damage and oxidation, but the manufacturer assessed that the fracture had originated at multiple sites along the fillet radius at approximately the 135oposition. It was assessed that the fracture had been due to low cycle fatigue(LCF), i.e. step progression in response to stress cycles associated with a significant change in engine operating conditions and probably directly related to flight cycles rather than high cycle fatigue associated with vibratory stresses. The total number of cycles involved in producing the intial crack,determined by integration of striation spacing versus distance from the origin,was estimated by the manufacturer as approximately 5,200 (predicted as accurate within ±10 per cent). The SMC estimate was much higher.
Background and discussion
When the incident occurred, the CCOC had considerably exceeded the FPI interval recommended by the Pratt & Whitney Overhaul Manual; the CCOC service at the time of the incident represented 176 per cent by hours and 145 per cent by cycles of the recommended interval since the last FPI inspection. The fabricated CCOC of the type that ruptured in this incident was the first standard used on the JT9D engine. The engine manufacturer believed that the effects of extensive cracking of the CCOC would be for the engine carcass and shafts to bend and for the rotating assemblies to rub heavily and considered that there would be no significant effects on the engine mounts or the pylon and no concern for an engine non-containment. This incident and other similar cases suggested that rapid extension of the crack was likely to occur during take-off when the engine was at high power and components were relatively highly stressed.
Three other cases of gross fracturing around the L Flange fillet have been reported. Two preceded this incident and one followed it. The post-incident actions included the engine manufacturer providing a management plan for the part and/or a retirement plan for the cases.
The recommendations by the AAIB included the following. The effects of the CCOC cracking in the incident reported here and in the other known cases had apparently been largely confined to the engine and had not had major repercussions on other parts of the aircraft. However, there was a likelihood that such a failure would result in sudden complete loss of power from one engine at a critical stage of flight, as in this incident. Additionally, even though the effects in the cases so far had been relatively benign, the possibility that such an extensive rupture of a major structural component of the engine hot section, with consequent leakage of high temperature gas and disruption of the high speed rotating assemblies, could hazard the aircraft could not be dismissed. In at least one of the other cases the crack direction had turned from circumferential to axial and this suggested the potential for an even more hazardous rupture.
A recommendation is therefore made that the CAA, in conjunction with the FAA, review the history and engineering analysis of the fabricated type of combustion chamber outer case used on the JT9D engine and mandate measures aimed at preventing recurrence of instances of extensive cracking of the case.
Reference:AAIB Bulletin 12/98.
