Broomfield, CO, USA
N772UA
BOEING 777-222
United Airlines flight 328 was climbing through 12,500 ft mean sea level about 5 minutes after departure from Denver International Airport (DEN), Denver, Colorado, when the right engine, a Pratt & Whitney PW4077, sustained a full-length fan blade separation, or fan blade out (FBO) event. This resulted in the subsequent separation of the engine inlet lip skin, fan cowl support beam, and components of the inlet, fan cowls, and thrust reversers (TRs), as well as an engine fire. The flight crew declared an emergency and landed the airplane without incident at the departure airport about 24 minutes after takeoff. There were no injuries to the passengers or crew, and no ground injuries due to debris; however, a vehicle and a residence sustained damage when impacted by the inlet lip skin and fan cowl support beam, respectively. Fan Blade Impact Damage Examination of the engine revealed that the separated fan blade and other fan debris impacted the fan case, which successfully contained the fan blade fragments. Damage to the nacelle inner and outer barrels was observed, and a postaccident evaluation indicated that the displacement wave of the impact resulted in a deflection of the fan case and contact with the nacelle doors and hinges, which subsequently resulted in the failure of the inlet aft bulkhead and the fan cowl support beam. The failure of the bulkhead, along with the damage to the inner and outer barrels, allowed these structures, as well as the inlet lip skin, to separate from the engine. Following the separation of the inlet, air loads resulted in the separation of the fan cowls and the fan cowl support beam. Simulation studies indicated that the carbon fiber reinforced plastic (CFRP) honeycomb structure of the event engine inlet and inlet aft bulkhead was unable to dissipate and redistribute the energy of the loads imposed by the FBO event in the same manner as the aluminum structure inlet that was used during certification tests. Separation of the inlet and fan cowls due to an FBO event is not allowed under certification standards, and following this event, Boeing developed modifications to the inlet to ensure that inlets and fan cowls remain in place during an FBO event that may damage the aft bulkhead, inner barrel, or outer barrel and modifications to add strength and ductility to the inlet by incorporating additional metallic structure. Boeing also developed procedures for inspection and repair for moisture ingression damage to the fan cowls, which can degrade the strength of the cowls. These modifications were subsequently mandated by Federal Aviation Administration (FAA) Airworthiness Directives (AD) 2022-06-10 and 2022-06-11, effective April 15, 2022. Additional modifications are expected to the fan cowl. This event was the fourth in-service FBO event due to fatigue cracking recorded for PW4000-powered 777 airplanes and resulted in the most nacelle damage of the four events. In the first event in 2010, approximately 50 percent of the blade airfoil was released. Full-span separations occurred in 2018, 2021, and during this event. Engine Fire Propagation Seconds after the FBO event, the flight crew received a right engine fire warning. The crew completed the engine fire checklist, which included activating the fire switch and discharging both engine fire extinguishing bottles; however, the fire was not arrested and continued to propagate through the engine for the remainder of the flight due to damage the engine sustained during the fan blade out event. Although the cockpit fire warning light extinguished shortly before landing, this was likely the result of thermal damage to the engine fire detection system. The engine fire propagated as the result of several cascading failures following the FBO event. 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, which created rotating torsion loads through the engine core structure; and the continued fan shaft imbalance during the engine run-down, which created rotating bending loads through the core structure. The loading associated with the high dynamic activity of the attached main gearbox (MGB) 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. Pratt & Whitney is evaluating actions to improve the strength of the “K” flange and expects hardware to be available in 2025. The fire spread to the 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 melted the aluminum latch beams at the lower end of each TR and through the TR inner wall and translating sleeves. One of the last components to separate from the airplane was a section of the outboard TR translating sleeve, which was located about 30 miles southeast of the debris associated with the initial FBO event. The burn-through of the TR lower bifurcation area likely occurred within about six to nine minutes of the initial FBO event, though certification standards required that materials in this area withstand fire for a minimum of 15 minutes. Examination of the engine’s fire suppression system revealed that the engine driven hydraulic pump supply shutoff valve failed to close as designed upon the crew’s activation of the engine fire handle due to silicone lubricant contamination of electrical contact components in the valve’s DC motor. The failure of the valve to close allowed a limited amount of hydraulic fluid to leak into the engine compartment and feed the undercowl fire. FAA AD 2022-06-10 and 2022-06-11 required installation of debris shields on the TR inner wall lower bifurcation area, as well as repeated functional checks of the engine driven hydraulic pump supply shutoff valves to ensure proper operation in response to fire switch activation. Fan Blade Fatigue Failure and Inspection Process The separated fan blade was fractured transversely across the chord of the airfoil near the fan hub fairing as the result of a fatigue crack, which originated at the surface of an internal radius in a hollow cavity within the blade. The event blade had accumulated 2,979 cycles since overhaul; at the time of the event, overhaul inspection was required every 6,500 cycles. As part of the overhaul, blades were inspected for both external and internal cracks using a proprietary thermal acoustic imaging (TAI) process. The most recent TAI inspection of the event fan blade occurred about five years before the event, in 2016. Inspection imagery 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. 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 occurred, and the blade was approved for continued service. Following an FBO event in 2018 involving another PW4077 engine, the data from the 2016 inspection of the blade involved in this event were reviewed again; once more, the indications were not identified as anomalous and the blade continued in service. Two of the low-level indications identified during the 2016 TAI inspection were likely associated with the fatigue crack that grew to result in the blade failure. The accident blade had accumulated 15,262 cycles since new, which was less than one quarter of the expected life for a nominal blade, and only 2,979 cycles since its last overhaul, less than half the prescribed inspection interval at the time. Metallurgical examination identified two conditions which contributed to the reduced fatigue life of the accident blade: a surface carbon contamination; and a geometric discontinuity that occurred during manufacturing. In assessing fatigue life of this blade relative to the nominal expectation, the reduced fatigue capability from the surface carbon contamination accounted for approximately 2/3 of the difference, and the increased stress from the geometric discontinuity accounted for approximately 1/3 of the difference. Following this event, Pratt & Whitney performed an immediate TAI inspection of the entire fleet before the next flight and issued a service bulletin introducing ultrasonic testing (UT) blade inspections to occur both immediately and at regular intervals. Additionally, the frequency of required TAI inspections was increased from every 6,500 cycles to every 1,000 cycles. The increased inspection interval and the immediate TAI inspection were made mandatory on April 15, 2022, when the FAA issued AD 2022-06-09. Additionally, the new UT inspection that was developed by Pratt & Whitney for the flowpath and midspan areas has shown a capability to detect small cracks that are below the threshold of detectability for the TAI inspection. The blades are now inspected by UT every 275 cycles. Examination of the crack in this event and previous fan blades failure events have shown the growth rates of the fatigue crack, from detectable size to full-wall penetration, are relatively stable and predictable in each case, since the sources for premature fatigue initiation are surface related and do not have a significant impact on growth through the thickness of the blade. The increased TAI inspection interval and the new UT inspections should provide multiple opportunities to detect cracks in the high-stress areas.
HISTORY OF FLIGHTOn February 20, 2021, about 1309 mountain standard time, a Boeing 777-222, N772UA, operated by United Airlines (UAL) as flight 328, experienced a right engine fan blade separation and subsequent engine fire shortly after takeoff from Denver International Airport (DEN), Denver, Colorado. The two pilots, eight crew members, and 229 passengers onboard were not injured. The airplane was operated as a Title 14 Code of Federal Regulations Part 121 scheduled passenger flight. The airplane departed DEN about 1304 enroute to Daniel K. Inouye International Airport (HNL), Honolulu, Hawaii. The captain was the pilot flying, and the first officer was the pilot monitoring. The pilots reported that preflight weather forecasts indicated moderate turbulence from about 13,000 ft mean sea level (msl) to 23,000 ft msl, and as the airplane climbed through about 12,500 ft msl at an airspeed about 280 knots (kts), they increased engine power in order to minimize the time spent climbing through the altitudes where turbulence was forecast. About 5 to 7 seconds after advancing the throttles, the cockpit voice recorder (CVR) captured a loud “bang,” and the flight data recorder (FDR) showed an uncommanded shutdown of the No. 2 (right) engine. Shortly thereafter, an engine fire warning activated on the engine indicating and crew alerting system (EICAS). The flight crew declared an emergency with air traffic control (ATC) and completed multiple checklists, including the engine fire checklist. As part of the engine fire checklist, the crew discharged both right engine fire extinguishing bottles; however, the engine fire warning continued to display on the EICAS until shortly before landing. The crew landed the airplane on runway 26 at DEN at 1328 and the airplane was met by aircraft rescue and firefighting (ARFF), which applied water and foaming agent to the right engine for about 40 minutes. The airplane was then towed off the runway, where the passengers disembarked via air stairs and were bussed to the terminal. Figure 1 below, a still image captured from in-flight video recorded by a passenger, shows the damage to the engine nacelle as well as the under-cowl fire about 11 minutes after the fan blade separation. Figure 1. Still image from passenger in-flight video showing engine nacelle damage and under-cowl fire about 11 minutes after fan blade separation (Courtesy Boeing via YouTube). At the time of the event, the airplane was over Broomfield, Colorado; multiple pieces of the engine inlet, fan cowls, and thrust reversers separated from the airplane and were found scattered over an area of about 40 acres, including a public park and residential areas. There were no ground injuries reported. PERSONNEL INFORMATIONCaptain The captain, age 60, held an airline transport pilot certificate with a rating for airplane multiengine land and multiple type ratings, including the B-777. His most recent first-class Federal Aviation Administration (FAA) medical certificate was issued on February 23, 2021. Operator records indicated that the captain had 28,062 total hours of flight experience, including 414 hours on the B-777 in the previous 12 months. His most recent proficiency check was completed on February 5, 2021. First Officer The first officer, age 54, held an airline transport pilot certificate with a rating for airplane multiengine land and multiple type ratings, including the B-777. Operator records indicated that the first officer had 18,612 total hours of flight experience, including 355 hours on the B-777 in the previous 12 months. His most recent proficiency check was completed on November 27, 2020. AIRCRAFT INFORMATIONOverview The Boeing 777-222 is a long range, twin-engine, transport category airplane. The primary wing and fuselage structure is of all metal construction, primarily aluminum alloys. The control surfaces and engine cowlings are of composite construction, which comprises graphite epoxy carbon fiber reinforced plastic (CFRP), fiberglass, or honeycomb sandwich. The incident airplane was manufactured in September 1995. Engines The airplane was powered by two Pratt & Whitney (P&W) PW4077 turbofan engines. The right engine was manufactured in 1995 and installed on the accident airplane in August 2016. At the time of the event, the engine had accumulated 12,384 hours and 2,979 cycles since overhaul and 81,768 hours and 15,262 cycles since new. The PW4077 is a dual-spool, axial flow, high-bypass turbofan engine that features a single-stage, 112-inch diameter fan (low pressure compressor [LPC] 1st stage), a 6-stage LPC, 11-stage high pressure compressor (HPC), annular combustor, a 2-stage high pressure turbine (HPT) that drives the HPC, and a 7-stage low pressure turbine (LPT) that drives the fan and LPC. Each engine is attached to a pylon on its respective wing. The engine inlet is attached to the forward end of the engine, the fan cowls are attached around the center portion of the engine, and the thrust reversers are attached around the aft portion of the engine. (see figure 2.) Engine flanges are identified alphabetically from the front of the engine aft, with the A-flange located where the inlet attaches to the fan case and the T-flange at the aft end of the exhaust case. (see figure 3.) Figure 2. Engine installation drawing for 777-200 (Source: Boeing). Figure 3. PW4000 112-inch significant engine flanges (Source: Pratt & Whitney) Inlet The engine inlet is a cantilevered structure attached to the forward flange of the engine fan case through the inlet attach ring with 52 bolts. The inlet consists of two concentric cylindrical structures joined together by forward and aft bulkheads (see figure 4). The hollow aluminum lip skin is attached to the forward bulkhead and provides an aerodynamic surface for the leading edge of the inlet and a passage for engine anti-ice air. The inlet aft bulkhead consists of the aluminum inlet attach ring and aluminum outer ring chord with a CFRP honeycomb sandwich composite web. The inlet forward bulkhead consists of the aluminum inner and outer ring chords with a stiffened aluminum web. The inlet outer barrel comprises three CFRP honeycomb sandwich panels. A section of the outer barrel in the lower right quadrant is comprised of a titanium skin, where the anti-ice exhaust duct is located. The inlet inner barrel is comprised of two CFRP honeycomb sandwich panels. The inner face sheet of the inner barrel is perforated for noise suppression and the outer face sheet is solid. Figure 4. Inlet cross-section drawing for 777-200 (Source: Boeing). Fan Cowl The fan cowl provides an aerodynamic closure around the engine fan cases and the doors open to allow maintenance access to the engine. The CFRP honeycomb sandwich construction cowls are semi-cylindrical doors fastened to four hinges at the upper ends; two on the cowl support beam, one floating hinge, and one hinge on the engine. The fan cowl support beam is a CFRP honeycomb sandwich panel attached at the forward end to the inlet attach ring and to the fan case at the aft end through aluminum fittings. The fan cowls interface with the inlet at the forward edge through a v-blade on the fan cowls that seats in a v-groove on the inlet aft bulkhead. The fan cowls interface with the thrust reversers at the aft edge through a sliding contact seal. Thrust Reversers The thrust reversers (TRs) provide an aerodynamic enclosure around the core of the engine, direct the fan exhaust, and actuate to provide reverse thrust during landing. The two semi-cylindrical TR halves comprise three main components; the translating sleeve, the fan duct cowl, and the aft cowl. The CFRP honeycomb sandwich inner wall of the fan duct cowl and the titanium aft cowl enclose the engine core and comprise the fire zone in the TR. The TRs are hinged at the upper end to the pylon and open to provide maintenance access. The main structural skeleton of the TR consists of the aluminum hinge beam at the upper end, the aluminum torque box at the forward end, the aluminum latch beam at the lower end, and the aluminum aft support ring and titanium aft cowl at the aft end. The CFRP honeycomb sandwich inner wall is connected to the TR at the upper and lower bifurcations. The CFRP honeycomb sandwich translating sleeve forms the outer surface of the TR and the outer wall of the fan duct cowl in the closed position. The translating sleeve slides aft along a mechanism attached to the torque box when actuated for reverse thrust. Rubber fire seals are installed in each TR half to help contain an undercowl fire within the interior of the fan duct inner wall and aft cowl. The fabric-reinforced silicone rubber seals are installed along the upper and lower bifurcation walls and down the upper aft edge of the aft cowl. Kapton-faced thermal insulation blankets are installed on the upper and lower bifurcations and on the inside of the inner wall in the fire zone to protect the composite structure from radiant engine heat and fire. Engine Fire Protection and Extinguishing Systems The B-777 engine fire protection comprised two systems: an engine fire and overheat detection system, and an engine fire extinguishing system. The engine fire and overheat detection system comprised two detector loops in each engine nacelle. Normally, both loops must detect a fire or overheat condition to cause an engine fire warning or overheat caution message to display on the EICAS. If a fault was detected in one loop, the system automatically switched to single-loop operation. If there were faults in both detector loops, no fire or overheat detection would be provided. The EICAS advisory message DET FIRE ENG (L or R) would be displayed if the engine fire detection system failed. An engine fire warning would be accompanied by several indications, including an aural fire bell, the illumination of master WARNING lights, an EICAS warning message (FIRE ENG [L or R]), the illumination of the affected engine fire switch, unlocking of the engine fire switch, and the illumination of the engine FUEL CONTROL (L or R) switch fire warning light. Each engine was equipped with two fire extinguisher bottles, which were located inside the engine nacelle and cowling and activated by engine fire switches in the flight deck. When the switch is pulled out and rotated in either direction, a single extinguisher bottle is discharged into the associated nacelle cavity. When the switch is rotated in the other direction, the remaining extinguisher bottle is discharged into the same engine. Activation of the fire switch is also designed to isolate the engine by closing the fuel spar valve, de-energizing the engine fuel metering unit cutoff solenoid, closing and depressurizing the engine driven hydraulic pump supply shutoff valve, closing the pressure regulator and shutoff valve, removing power from the thrust reverser isolation valve, and tripping the generator and backup generator fields. AIRPORT INFORMATIONOverview The Boeing 777-222 is a long range, twin-engine, transport category airplane. The primary wing and fuselage structure is of all metal construction, primarily aluminum alloys. The control surfaces and engine cowlings are of composite construction, which comprises graphite epoxy carbon fiber reinforced plastic (CFRP), fiberglass, or honeycomb sandwich. The incident airplane was manufactured in September 1995. Engines The airplane was powered by two Pratt & Whitney (P&W) PW4077 turbofan engines. The right engine was manufactured in 1995 and installed on the accident airplane in August 2016. At the time of the event, the engine had accumulated 12,384 hours and 2,979 cycles since overhaul and 81,768 hours and 15,262 cycles since new. The PW4077 is a dual-spool, axial flow, high-bypass turbofan engine that features a single-stage, 112-inch diameter fan (low pressure compressor [LPC] 1st stage), a 6-stage LPC, 11-stage high pressure compressor (HPC), annular combustor, a 2-stage high pressure turbine (HPT) that drives the HPC, and a 7-stage low pressure turbine (LPT) that drives the fan and LPC. Each engine is attached to a pylon on its respective wing. The engine inlet is attached to the forward end of the engine, the fan cowls are attached around the center portion of the engine, and the thrust reversers are attached around the aft portion of the engine. (see figure 2.) Engine flanges are identified alphabetically from the front of the engine aft, with the A-flange located where the inlet attaches to the fan case and the T-flange at the aft end of the exhaust case. (see figure 3.) Figure 2. Engine installation drawing for 777-200 (Source: Boeing). Figure 3. PW4000 112-inch significant engine flanges (Source: Pratt & Whitney) Inlet The engine inlet is a cantilevered structure attached to the forward flange of the engine fan case through the inlet attach ring with 52 bolts. The inlet consists of two concentric cylindrical structures joined together by forward and aft bulkheads (see figure 4). The hollow aluminum lip skin is attached to the forward bulkhead and provides an aerodynamic surface for the leading edge of the inlet and a passage for engine anti-ice air. The inlet aft bulkhead consists of the aluminum inlet attach ring and aluminum outer ring chord with a CFRP honeycomb sandwich composite web. The inlet forward bulkhead consists of the aluminum inner and outer ring chords with a stiffened aluminum web. The inlet outer barrel comprises three CFRP honeycomb sandwich panels. A section of the outer barrel in the lower right quadrant is comprised of a titanium skin, where the anti-ice exhaust duct is located. The inlet inner barrel is comprised of two CFRP honeycomb sandwich panels. The inner face sheet of the inner barrel is perforated for noise suppression and the outer face sheet is solid. Figure 4. Inlet cross-section drawing for 777-200 (Source: Boeing). Fan Cowl The fan cowl provides an aerodynamic closure around the engine fan cases and the doors open to allow maintenance access to the engine. The CFRP honeycomb sandwich construction cowls are semi-cylindrical doors fastened to four hinges at the upper ends; two on the cowl support beam, one floating hinge, and one hinge on the engine. The fan cowl support beam is a CFRP honeycomb sandwich panel attached at the forward end to the inlet attach ring and to the fan case at the aft end through aluminum fittings. The fan cowls interface with the inlet at the forward edge through a v-blade on the fan cowls that seats in a v-groove on the inlet aft bulkhead. The fan cowls interface with the thrust reversers at the aft edge through a sliding contact seal. Thrust Reversers The thrust reversers (TRs) provide an aerodynamic enclosure around the core of the engine, direct the fan exhaust, and actuate to provide reverse thrust during landing. The two semi-cylindrical TR halves comprise three main components; the translating sleeve, the fan duct cowl, and the aft cowl. The CFRP honeycomb sandwich inner wall of the fan duct cowl and the titanium aft cowl enclose the engine core and comprise the fire zone in the TR. The TRs are hinged at the upper end to the pylon and open to provide maintenance access. The main structural skeleton of the TR consists of the aluminum hinge beam at the upper end, the aluminum torque box at the forward end, the aluminum latch beam at the lower end, and the aluminum aft support ring and titanium aft cowl at the aft end. The CFRP honeycomb sandwich inner wall is connected to the TR at the upper and lower bifurcations. The CFRP honeycomb sandwich translating sleeve forms the outer surface of the TR and the outer wall of the fan duct cowl in the closed position. The translating sleeve slides aft along a mechanism attached to the torque box when actuated for reverse thrust. Rubber fire seals are installed in each TR half to help contain an undercowl fire within the interior of the fan duct inner wall and aft cowl. The fabric-reinforced silicone rubber seals are installed along the upper and lower bifurcation walls and down the upper aft edge of the aft cowl. Kapton-faced thermal insulation blankets are installed on the upper and lower bifurcations and on the inside of the inner wall in the fire zone to protect the composite structure from radi
The fatigue failure of the right engine fan blade. Contributing to the fan blade failure was the inadequate inspection of the blades, which failed to identify low-level indications of cracking, and 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. Contributing 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 out event. Contributing to the severity of the engine fire was the failure of the “K” flange following the fan blade out, which allowed hot ignition gases to enter the nacelle and imparted damage to several components that fed flammable fluids to the nacelle, which allowed the fire to propagate past the undercowl area and into the thrust reversers, where it could not be extinguished.
Source: NTSB Aviation Accident Database
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