Côte-d'Or Region, OF, FR
HB-JCA
AIRBUS A220
The No. 1 engine low pressure compressor (LPC) stage 1 integrally bladed rotor (IBR) failure was caused by a stage 1 IBR high cycle fatigue crack that originated at the runout of an airfoil leading edge root radius. Multiple analytical methods, including two-dimensional (2D) computational fluid dynamics (CFD), acoustic testing, and instrumented flight testing identified a coupled LPC stage 3 and stage 1 IBR instability caused by an acoustic coincidence with the 2.5 bleed valve duct cavity. At high engine N1 speeds, the stage 3 IBR blade tips generate vortices/turbulent airflow and given the right conditions, the turbulent airflow can create an acoustic tone as it passes over the 2.5 bleed valve duct cavity, located immediately aft of the LPC stage 3 IBR (Figure 1). The acoustic tone drove a LPC stage 3 IBR blade 1st bending mode excitation that was then mechanically transferred through the LPC module and excited a LPC stage 1 IBR stiffwise bending mode that was present at the same approximate frequency. The resultant stresses on the LPC stage 1 IBR blades exceeded material limits and subsequently led to crack formation and eventual progression to overload failure. Figure 1- Engine LPC Diagram Identifying the IBR 3 / 2.5 Bleed Valve Acoustic Interaction Factors that contributed to the LPC stage 3 and stage 1 IBR acoustic coincidence and blade excitation within the engine operating range were: installation of electronic engine control (EEC) Software V2.11.7.2 and the low time rub in period on the LPC IBR blade tip clearances. EEC Software V2.11.7.2 changed the LPC inlet guide vane schedule to rotate the IGV’s in the closed direction at specific high power engine conditions to improve engine stall/surge margin. The revised vane schedule inadvertently created conditions that were favorable for generation of the 2.5 bleed valve duct cavity acoustic tone and IBR mode excitation. New engines have tighter clearances between the LPC IBR blade tips and the outer air seals. The reduced clearance created unsteady loading at the blade tip region and resulted in stronger acoustic coupling/flutter response. After a rub in period, the clearance increases, and the occurrence of flutter onset is reduced. This incident was the second of four Airbus A220-300 PW1500G LPC stage 1 IBR separations that occurred between July 25, 2019 and February 12, 2020. The findings from the first LPC stage 1 IBR failure investigation are available in the NTSB investigation number ENG19IA029 docket.
HISTORY OF FLIGHT On September 16, 2019, about 1925 coordinated universal time (UTC), a Swiss International Air Lines (SWISS) Airbus Canada A220-300, registration HB-JCA, equipped with two Pratt & Whitney (P&W) PW1524G geared turbofan engines, experienced a No. 1 (left) engine failure during climb, just prior to reaching cruise altitude, FL350, over the Côte-d’Or region of northeastern France. The flight crew initiated quick reference handbook procedures, declared an emergency, and returned to Geneva Airport (GVA), Geneva, Switzerland where they made an uneventful single-engine landing. A post-flight examination of the engine revealed a hole in the low pressure compressor (LPC) case and a separated LPC stage 1 integrally bladed rotor (IBR). A majority of the LPC stage 1 IBR was recovered from the bypass duct/thrust reverser structure. The thrust reverser exhibited outer barrel impact damage, but there was no evidence of radial uncontainment through the nacelle. There were no passenger or crew injuries reported. The regularly scheduled passenger flight was operating from GVA to London Heathrow Airport (LHR), London, United Kingdom. In accordance with ICAO Annex 13, the National Transportation Safety Board (NTSB) accepted the delegation of this investigation from the French Bureau d’Enquêtes et d’Analyses pour la Sécurité de l’Aviation Civile (BEA), who appointed an Accredited Representative to assist with the investigation. The Transportation Safety Board of Canada (TSB) appointed an Accredited Representative as the State of Design and Manufacture of the airplane. The Swiss Transportation Safety Investigation Board (STSB-AV) appointed an Accredited Representative as the State of the Operator. DAMAGE TO THE AIRPLANE There was no damage to the airplane structure. TEST AND RESEARCH Engine Examination and Disassembly The incident engine, serial number P736195, was shipped from GVA to the P&W Columbus Engine Center in Columbus, Georgia, USA for examination and disassembly. The LPC forward case was missing material from the 3 to 11 o’clock positions. The LPC mid case had a 360 degree circumferential crack approximately ¾ inch forward of the LPC mid-to-aft case flange and was missing material from the 7 to 10 o’clock positions. The remaining LPC mid case material from the 1 to 7 o’clock positions was bent radially outward from the engine centerline. LPC stages 2 and 3 located aft of the separated LPC stage 1 IBR exhibited secondary impact damage. A borescope inspection (BSI) of the engine core revealed impact damage and material loss through all eight high pressure compressor (HPC) stages and both the high pressure turbine (HPT) stage 1 blades and stage 2 vanes were missing material. Metallurgy The LPC stage 1 IBR fragment recovered from the incident engine bypass duct at GVA was shipped to the P&W Materials and Processes Engineering Laboratory in East Hartford, Connecticut for analysis. The primary fracture surface revealed a high cycle fatigue crack that originated at the runout of an airfoil leading edge root radius. Scanning electron microscopy (SEM) analysis identified the fracture origin approximately 0.0033 inch (0.084 mm) beneath the material surface. There was no evidence of material or processing anomalies near the crack origin and the fracture surface showed no evidence of cyclic markers or arrest lines. Testing and Analysis to Identify Failure Mode Computer modeling, component testing, and instrumented flight tests were conducted to identify the root cause of the failures. The testing identified that at specific engine operating conditions an acoustic tone was generated by the 2.5 bleed valve located immediately aft of LPC stage 3 rotor that excited coupled LPC stage 3 and stage 1 IBR rotor modes. The sound waves from the acoustic tone excited a LPC stage 3 IBR 1st bending mode that then mechanically transferred through the LPC module and excited a coupled LPC stage 1 IBR stiffwise bending mode. The acoustic coincidence and blade flutter response created LPC stage 1 IBR blade stress levels that exceeded material limits and resulted in leading edge blade root crack formation and subsequent IBR fracture. Factors that contributed to the acoustic coincidence included an electronic engine control (EEC) software revision that altered the LPC variable inlet guide vane (IGV) schedule and tighter LPC IBR blade tip clearance due to low time on the engine. There were three PW1524G-3 and one PW1521G-3 LPC stage 1 IBR separations between July 25, 2019 and February 12, 2020, that occurred on multiple operators. The incident detailed in this investigation was the second of the four PW1500G series LPC stage 1 IBR failures. The engine parameters at the time of the four LPC stage 1 IBR failures and the resulting engine damage were consistent in each of the events. Corrective Actions Corrective actions were released by P&W, Federal Aviation Administration (FAA), Airbus Canada, and Transport Canada to reduce the likelihood of additional events. A recurrent BSI of the LPC stage 1 IBR was mandated, and the inspections identified two additional LPC stage 1 IBR crack findings in the fleet. An engine low rotor speed (N1) restriction above FL290 was implemented to reduce the likelihood of mode excitation. EEC software update V2.11.9 was released to revert the LPC vane schedule back to the original vane schedule that was programmed prior to the LPC stage 1 IBR fractures. Another EEC software update, V2.11.10, was later released to automate the 95.2% N1 speed limit at altitudes above FL290. P&W has also begun installation of a redesigned 2.5 bleed valve in the fourth quarter of 2021 to provide adequate frequency separation from the coupled LPC response that is the driver of the LPC stage 1 IBR separations.
A No. 1 (left) engine low pressure compressor (LPC) stage 1 integrally bladed rotor (IBR) separation due to a high cycle fatigue crack (HCF) that originated at the runout of an airfoil leading edge root radius. The HCF crack developed because of a mechanically coupled LPC stage 3 and stage 1 IBR mode excitation and blade flutter response. The excitation was driven by an acoustic tone generated by turbulent airflow passing over the 2.5 bleed valve duct cavity while the engine was operating at high speeds in specific flight conditions. A primary contributor to the failure mode was an electronic engine control (EEC) software update that changed the LPC vane schedule and increased the likelihood of LPC stage 1 IBR blade flutter onset within the engine operating range.
Source: NTSB Aviation Accident Database
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