On May 14, 2024, a United Airlines (UAL) Boeing 787-9 airplane, serial number 36409, registration number N27957, powered by two GE Aerospace GEnx™-1B76A/P2 turbofan engines, experienced an uncontained high pressure turbine engine failure and inflight shutdown (IFSD) of the No. 1 (left) engine shortly after reaching their assigned flight altitude of FL310 following departure from the Singapore Changi Airport, Singapore. The airplane was operated as a Title 14 Code of Federal Regulations (CFR) Part 121 air carrier flight as a regularly scheduled passenger flight from the Singapore Changi Airport, Singapore, to the San Francisco International Airport, San Francisco, California (SFO). The flight crew reported hearing a pop sound and felt the airplane shudder, followed by the annunciation of a No. 1 engine EICAS (Engine Indicating and Crew Alerting System) message for left engine Fire Warning. The flight crew pulled the No. 1 engine fire handle, secured the engine, and performed an inflight air turnback to Singapore Changi Airport where an uneventful single engine landing was performed. There were no reported injuries. Upon landing, ground inspection of the No.1 engine revealed damage to the airplane, missing material from the left side thrust reverser cowling and a penetration of the engine’s high pressure turbine (HPT) case. The damage to the airplane and engine was documented by representatives from UAL and GE Aerospace (GE). The initial on-scene engine examination found two exit holes in the HPT case, centered at approximately the 9:00 o’clock position; the HPT stage 1 blades and stage 2 stators could be seen through the exit holes, and it appeared the HPT stage 1 aft blade retainer was missing material. There was no evidence of an under-cowl fire but localized heat distress was observed on electrical wiring harnesses, fire loop grommets/isolators and some of the HPT active clearance control (HPTACC)/low pressure turbine active clearance control (LPTACC) rubber bellows. Examination of the left side thrust reverser cowl found a large section of material, approximately 3 feet by 6 feet in size, missing from the translating sleeve and metal debris was found lodged into the inner diameter of the thrust reverser. The left side thrust reverser inner fixed structure contained an exit hole, approximately 13 inches by 15 inches in size, centered at about the 8:30 o’clock position. Since this event took place in international waters outside the territory of a State, the National Transportation Safety Board (NTSB) led the investigation as the State of Registry in accordance with ICAO Annex 13 guidelines. Additional parties to the investigation include the Federal Aviation Administration (FAA), Boeing, United Airlines, and GE Aerospace. The engine was removed from the airplane, the fan module removed from the engine, and the engine propulsor prepared for shipment to the United States for disassembly and examination. The left side thrust reverser cowl, various other components and recovered metal debris were also shipped to GE for further examination. The examination team, comprised of party members from GE, Boeing, FAA, UAL, and the NTSB, convened at the GE’s, Evendale, Ohio facility to perform the engine disassembly and examination. Disassembly of the engine found significant hard body damage and missing material throughout the HPT module. The HPT case contained two exit holes, and the case wall was distorted outward. There were no signs of fuel nozzle streaking, hot spots or burn throughs in the combustion chamber. Multiple HPT stage 1 nozzle segments were found disengaged. All HPT stage 1 and stage 2 blades were fractured transversely across the airfoils at various span lengths. Downstream damage to the low pressure turbine (LPT) module included one fractured LPT stage 1 blade but the LPT stage 7 blades appeared undamaged. The HPT stage 1 aft blade retainer was found lodged in place on the forward side of the HPT rotating interstage seal (RIS) and fractured, with the outboard portion of the retainer arm missing. The HPT stage 1 aft blade retainer and the HPT RIS were transported to GE’s material systems laboratory, separated, and the retainer further examined. The HPT stage 1 aft blade retainer presented a 360-degree circumferential fracture surface adjacent to the rabbet undercut fillet and the outboard retainer arm was missing (Figure 1). Figure 1. Aft Blade Retainer Ring Cross Section and Fracture Location Fractography revealed the crack initiated in the upper collet radius at a single point of origin which corresponded to a surface connected anomaly that measured 0.005 inches by 0.015 inches by 0.0005 inches (Figure 2). The surface connected anomaly was determined to be a cluster of particles enriched with Titanium, Aluminum, Zirconium, Magnesium, and Oxygen, which is consistent with a cluster of Oxide/Carbonitride/Nitride/Carbide (OCNC) particles. The fracture surface contained an area of intergranular fracture morphology with visible beach marks, consistent with hold-time fatigue growth and the remaining fracture surface was consistent with overload fracture (Figure 2). Figure 2. Fractography Images of Fracture Surface The investigation of this retainer supplements the ongoing investigation into two other cracked HPT stage 1 aft blade retainers; one was an opportunistic find during a scheduled maintenance activity and the other was an in flight failure that led to an uncontained failure event similar to the UAL event covered within this report. A fault tree was established to consider various aspects of the component design, material, and operational factors to identify possible causes and contributing factors. GE’s analysis of material test data suggests that surface connected OCNCs have a reduced crack incubation/nucleation life at high stress conditions, and that surface connected OCNCs do not behave according to traditional crack growth threshold behavior. Due to the difference between the original calculated life capability (e.g. crack initiation) and the observed operational cycles, GE performed a material sensitivity study to assess the sensitivity of the collet feature in the HPT stage 1 aft blade retainer to various parameters identified within the fault tree. The event HPT stage 1 aft blade retainer is manufactured from Rene 65, a wrought nickel-base superalloy. GE initiated material test plans to explore the Rene 65 material behavior with focus on low cycle fatigue (LCF) and sustained peak low cycle fatigue (SPLCF) capabilities, by using standard test bar specimens and coupons representative of the collet feature. The studies found that activating surface connected OCNCs can in fact reduce material fatigue life capability and that not every OCNC will initiate a crack, suggesting an unknown bimodal behavior. The calculated stress conditions (high peak, high mean stress, high stress-ratio (minimum stress to maximum stress), and high temperature) for the GEnx HPT stage 1 aft blade retainer collet feature were found to be uniquely high when compared to the stress conditions of other rotating hardware. This higher stress environment is the result of the assembly interference fits between the HPT stage 1 disk, the HPT stage 1 retainer, and the HPT RIS in combination with the operating loads (Figure 3). The collet feature was found to have a significant SPLCF debit (low crack growth life) when a hold-time fatigue condition was taken into account. Figure 3. Aft Blade Retainer Ring – Collet Feature Stress Map Multiple corrective actions have already been taken by GE as a result of this investigation and the findings from three other cracked HPT stage 1 aft blade retainers. The corrective actions focused on inspection/removal of suspect HPT stage 1 aft blade retainers, and continuous fleet monitoring for specific shifts in core engine vibration trends. The short-term containment actions include service bulletins (SB) for the removal/inspection/replacement of the HPT stage 1 aft blade retainer at the piece part level, SBs for the ultrasonic inspection of the HPT stage 1 aft blade retainers at the module level, and the use of a data analytic tool to monitor core vibration trends in all GEnx engines. At the time this report was nearing completion, there were at least 477 HPT stage 1 aft blade retainers removed from service for inspection and at least 55 HPT modules ultrasonically inspected, with zero crack findings to date. A data analytic tool was created to perform continuous fleet monitoring and evaluation of core engine vibration trends for all engines within the GEnx-1B and GEnx-2B engine fleets. The data analytic tool was based on the engine performance/vibration data from the three engines that contained the cracked HPT stage 1 aft blade retainers, and the analytical rules established such that the core engine vibration shifts of these engines would have resulted in an alert for their removal. In the event the data analytic tool identifies a specific engine with a core vibration alert, a more thorough assessment would be performed by GE’s engineering team on a case-by-case basis. In February 2025, the data analytic tool alerted on an operational engine, the engine was removed and a detailed ultrasonic inspection resulted in a positive crack indication. The long-term corrective action being pursued by GE will re-establish the manufacturing process for the previously approved HPT stage 1 aft blade retainer configuration made of a Rene 88 alloy, a nickel-based superalloy powder metal. GE intends to make the Rene 88 retainer the primary configuration. The Rene 65 retainer will remain certified for use and listed in the illustrated parts catalog (IPC) but it will no longer be available as a spare part, further diminishing the population of the Rene 65 retainers through attrition.