Description

On 20 February 2021, a PW4077-engined Boeing 777-200 (N772UA) being operated by United Airlines on a scheduled domestic passenger flight from Denver to Hawaii as UAL328 experienced a right engine fan blade separation and subsequent engine fire as it climbed through approximately 12,500 feet about five minutes after takeoff. A MAYDAY was declared, the climb stopped and the corresponding engine fire response was completed as the flight began the return to Denver. Despite completion of the engine fire emergency response with both extinguisher shots used, the fire warning did not cease until the aircraft was on an extended downwind leg for runway 25 seven minutes prior to touchdown. Following an overweight landing, the aircraft was stopped on the runway until the engine had been cooled down and it was then towed clear and the passengers disembarked to buses and taken to the terminal. There was no aircraft damage beyond that to the failed engine and its pylon and no injuries to the 229 passengers or the 10 aircrew operating the flight. Debris fell onto and damaged a vehicle and a house but no injuries resulted.

A still image from a passenger in-flight video showing the failed engine nacelle damage and the continuing fire 11 minutes after the failure. [Reproduced from the Official Report]

Investigation

An Investigation was carried out by the US National Transportation Safety Board (NTSB). Relevant data was obtained from the SSFDR and the SSCVR. Data on the QAR was identical to that on the SSFDR and it was not downloaded. Recorded ATC communications during the flight were also available.

The 60 year-old Captain had a total of 28,062 hours flying experience which included 538 hours on type during and after gaining his type rating just over a year ago. The 53 year-old First Officer had a total of 18,612 hours flying experience which included 4,190 hours on type.

What happened

The flight took off from runway 25 at Denver with the Captain as PF and as it climbed through approximately 12,500 feet QNH at 280 KCAS, the crew decided to increase thrust slightly in order to minimise the time which might be spent in forecast moderate turbulence between 13,000 feet and FL230. Between five and seven seconds after thrust levers had been advanced, a loud ‘bang’ was heard and an uncommanded shutdown of the right engine was followed two seconds later by an EICAS fire warning for the same engine. 

The climb was stopped and a MAYDAY call made advising the intention to return to Denver and a left turn to do so was approved. The applicable checklists, including the one for an engine fire, were completed. However, the engine fire warning did not then cease and continued until shortly before landing on runway 26 at Denver 24 minutes later. 

The aircraft was stopped on the runway and the airport RFFS met the aircraft. They observed smoke and flames emanating from the remains of the right engine and responded by immediately applying water and foam to it which extinguished the fire and cooled the engine. After approximately 40 minutes, with all occupants still on board, it was assessed safe to tow the aircraft clear of the runway. The passengers were then disembarked to buses using airstairs and taken to the passenger terminal. 

The engine failure occurred as the aircraft was overflying the Broomfield area and it was found that multiple pieces of the engine inlet, fan cowls and thrust reversers which had separated from the aircraft had subsequently been found scattered over a 40 acre area which included a public park and residential properties. This ejected debris included direct hits on a vehicle hit by the engine inlet lip skin and a house hit by the fan cowl support beam. However, no reports of any consequential injuries were received.

Why it happened

The PW4077 engine is a dual-spool, axial flow, high-bypass turbofan engine that features in order from the front a single stage 2.8 metre diameter fan, a 6 stage low pressure compressor (LPC), an 11 stage high pressure compressor (HPC) an annular combustor, a 2 stage high pressure turbine (HPT) which drives the HPC and a 7 stage low pressure turbine (LPT) which drives the fan and the LPC. These engine components are contained by the engine inlet attached to the forward part of the engine, the fan cowls which are attached around the central part of the engine and the thrust reversers which are attached around the aft part of the engine. The engine is attached to the wing by a pylon through which fuel is supplied and the fire extinguisher discharge nozzles are located inside of the engine nacelle and cowling.

An annotated external view of the engine. [Reproduced from the Official Report] 

An initial examination of the engine determined that origin of its failure and the subsequent sustained engine fire had been the full length separation of a single fan blade which then resulted in the separation of the engine inlet lip skin, the fan cowl support beam and various components from the inlet, fan cowls and thrust reversers. A more detailed examination then found that the separated fan blade and other fan debris had impacted the fan case, which successfully contained the fan blade fragments. However, the displacement wave of this impact had resulted in a deflection of the fan case and contact with the nacelle doors and hinges and this had resulted in the failure of the inlet aft bulkhead and the fan cowl support beam. 

The position of the primary fan blade failure and the adjacent secondary partial failure. [Reproduced from the Official Report]

The failure of the bulkhead, along with the damage to the inner and outer barrels, had allowed these structures and the inlet lip skin to separate from the engine following which air loads resulted in the separation of the fan cowls and the fan cowl support beam. Simulation studies indicated that the Carbon Fibre Reinforced Plastic (CFRP) honeycomb structure of the engine inlet and the inlet aft bulkhead had failed to dissipate and redistribute the energy of the loads arising from the blade shedding event in the same way that the aluminium structure which was used during certification tests had done. It was noted that separation of the inlet and fan cowls due to fan blade shedding event is not allowed under the applicable engine certification standards. It was also noted that the event under investigation was the fourth in-service fan blade failure event due to fatigue cracking recorded for PW4000-powered Boeing 777 aircraft and had led to more nacelle damage than in any of the previous three such events, although apart from the 2010 event which had only involved part of a blade, the previous events in 2018 and 2021 had both involved full length blade separations.

The continuation of the engine fire despite the prompt use of both fire extinguisher shots was a matter requiring an additional line of investigation. It was also considered that the fact that the fire warning had ceased shortly before landing was not because the fire had ceased but due to thermal damage to the engine fire detection system. 

Propagation of the fire through the engine was assessed to have been the consequence of a series of “cascading failures” as follows:

• The engine core was subjected to high dynamic loads due to the energy of the initial blade release. 
• The fan blade rubbing against the case created rotating torsion loads through the engine core structure.
• The continued fan shaft imbalance during the engine run-down created rotating bending loads through the core structure. 
• The loading associated with the high dynamic activity of the attached main gearbox (MGB) which ultimately resulted in the failure of the “K” flange bolts that attached the MGB to the engine. 
• The remaining “K” flange bolts then fractured, resulting in the total separation of the “K” flange, which allowed hot, compressed gases to escape the engine core and provided an ignition source in the engine nacelle.
• As the “K” flange was part of the MGB support structure, the failure of the flange also allowed the MGB to rotate and the MGB-mounted servo fuel heater to contact the engine core-mounted fuel oil cooler. 
• As a result of this contact, a high-pressure fuel cavity within the servo fuel heater was fractured open, releasing high-pressure fuel into the nacelle, where it was ignited by the hot, compressed gases that escaped through the “K” flange separation.
• The fire spread to the Thrust Reverser (TR) lower bifurcation area, burned away the support structure for the nacelle drain access door and exited the lower aft TR area. 
• The undercowl fire then melted the aluminium latch beams at the lower end of each TR and through the TR inner wall and translating sleeves. The outboard TR translating sleeve was one of the last components to separate from the aircraft which accounted for its subsequent discovery 30 miles from the main debris field. 

It was assessed that the burn-through of the TR lower bifurcation area had probably occurred between six and nine minutes after the initial fan blade failure whereas the applicable engine certification standard had required that materials in this area withstand fire for a minimum of 15 minutes.

Also, an examination of the engine’s fire suppression system found that the engine driven hydraulic pump supply shutoff valve had failed to close as designed when the flight crew pulled the engine fire handle because of silicone lubricant contamination of electrical contact components in the valve’s DC motor. This failure had allowed a limited amount of hydraulic fluid to leak into the engine compartment and feed the undercowl fire. 

It was established that a fatigue crack which had originated at the surface of an internal radius in a hollow cavity within the blade had been the cause of the transverse fracture of the separated fan blade near the fan hub fairing. This blade had accumulated 2,979 cycles since overhaul in 2016 which at the time of the event was required every 6,500 cycles. At such overhauls, blades were subject to inspection using a proprietary Thermal Acoustic Imaging (TAI) process which in the 2016 examination had “revealed multiple low-level indications, two of which were in the fatigue crack origin area that were reviewed further and interpreted as being generated by camera sensor noise or loose contamination within the cavity”. The Investigation took the view that “given the observed indications and the inspection criteria in place at the time, the blade should have received a second TAI inspection, or the images should have undergone a team review”. However there was no record that either of these inspections had occurred before the blade was approved for continued service. It was found that a similar failure to adequately assess the fitness of fan blades for continued service had featured in the of the 2018 PW4077 fan blade failure. 

It was also noted that the 15,262 cycles since new accumulated by the accident blade were less than a quarter of the expected life for a nominal blade and the 2,979 cycles accumulated since its most recent overhaul were less than half the prescribed inspection interval at the time. Metallurgical examination identified two conditions which had contributed to the reduced fatigue life of the accident blade as “surface carbon contamination and a geometric discontinuity that occurred during manufacturing”. The relative effect of these two circumstances on the actual fatigue life of the blade relative to the nominal expectation was considered to be 67% due to surface carbon contamination and 33% due to increased stress due to defective manufacture.

The Probable Cause of the accident was formally documented as “the fatigue failure of the right engine fan blade”

Contributory Factors to the fan blade failure were identified as:

1. the inadequate inspection of the blades which failed to identify low-level indications of cracking;
2. the insufficient frequency of the manufacturer’s inspection intervals, which permitted the low-level crack indications to propagate undetected and ultimately resulted in the fatigue failure.

A Contributory Factor to the severity of the engine damage following the fan blade failure was the design and testing of the engine inlet, which failed to ensure that the inlet could adequately dissipate the energy of, and therefore limit further damage from, an in-flight fan blade failure. 

A Contributory Factor to the severity of the engine fire was the failure of the “K” flange following the fan blade failure which allowed hot ignition gases to enter the nacelle and cause damage to several components that fed flammable fluids to the nacelle. This allowed the fire to propagate past the undercowl area and into the thrust reversers, where it could not be extinguished.

Safety Action taken as a result of the findings of the Investigation whilst it was in progress were noted as having included the following:

• The FAA initially issued Emergency AD 2021-05-51 which effectively grounded all Boeing 777-200 and 777-300 aircraft powered by PW4000 2.8 metre fan engines pending a one-time TAI inspection of the first stage LPC blades. 
• Pratt & Whitney issued an Alert Service Bulletin ASB PW4G-112-A72-361 requiring Ultrasonic and Thermal Acoustic Image (TAI) Inspection of 1st stage LPC Blade Assemblies to Find Airfoil Cracks which included both immediate and repetitive Ultrasonic inspections for three specific high-risk areas on 2.8 metre hollow fan blades and also introduced a required TAI inspection every 1,000 cycles for all 1st stage LPC blades on these engines. The inspections included in this ASB were subsequently made mandatory by the issue by the FAA of AD 2022-06-09.   
• Boeing developed an interim design solution incorporating engine nacelle modifications and subsequently issued multiple alert SBs for fan cowl inspections, modifications to the inlet cowls and thrust reversers on 777-200 and 777-300 aircraft fitted with PW4000 series engines and inspection/repair of fan cowls for moisture ingression. These SBs were subsequently mandated by the issue of ADs 2022-06-10 and 2022-06-11.  

The Final Report was published on 7 September 2023. No Safety Recommendations were made. 

Editor's Note: Some inconsistencies in the Final Report which was the primary source of the above summary were resolved by reference to the event ‘Docket’ which if required is accessible via the link at the end of the Final Report. The same source provided a small amount of relevant additional information incorporated in the summary.

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