Pensacola, FL, USA
N766SW
BOEING 737-700
The Boeing 737-700 airplane was climbing through flight level 310 when fan blade No. 23 in the left CFM56-7B engine fractured at its root, with the dovetail (part of the blade root) remaining within a slot of the fan disk. The separated fan blade impacted the engine fan case and fractured into multiple fragments. The blade fragments traveled forward into the inlet and caused substantial damage that compromised the structural integrity of the inlet, causing most of the inlet structure to depart from the airplane. A large portion of the inlet contacted and punctured the left side of the fuselage, creating a hole of sufficient size to cause the cabin to depressurize. The flight crew conducted an emergency descent and landed safely at Pensacola International Airport, Pensacola, Florida, about 21 minutes after the fan-blade-out (FBO) event occurred. The fan blade fractured due to a low-cycle fatigue crack that initiated in the blade root dovetail under the blade coating near the outboard edge. Metallurgical examination of the fan blade found that its material composition and microstructure were consistent with the specified titanium alloy and that no surface anomalies or material defects were observed in the fracture origin area. The fracture surface had fatigue cracks that initiated close to the dovetail leading edge convex side area, which is where the greatest stresses from operational loads, and thus the greatest potential for cracking, were predicted to occur. The fan blades were not certified as life-limited parts. The accident fan blade (as well as the six other cracked fan blades in the accident engine) failed with 38,152 cycles since new. Similarly, the fan blade associated with an April 2018 FBO accident (case number DCA18MA142) failed with 32,636 cycles since new. Further, 19 other cracked fan blades on CFM56-7B engines had been identified as of January 2020, and those fan blades had accumulated an average of about 33,000 cycles since new when the cracks were detected. After this accident, CFM reevaluated the fan blade dovetail stresses and determined that the fatigue cracks initiated in an area of high stress on the dovetail and that the dovetail was experiencing peak stresses that were higher than originally predicted. CFM found that the higher operational stresses resulted from coating spalling, higher friction levels when operated without lubrication or a shim, variations in coating thickness, higher dovetail edge loading than predicted, and a loss or relaxation of compressive residual stress (the stress that is present in solid material in the absence of external forces). Before the application of the dovetail coating during manufacturing and before the reapplication of the coating that is stripped during each overhaul, the entire blade, including the dovetail, is shot-peened to provide a compressive residual stress surface layer for the material, which increases the fatigue strength of the material and relieves surface tensile stresses that can lead to cracking. A loss of residual stress could be the result of a fan blade's exposure to high temperatures during the application of the dovetail coating as part of the overhaul of a blade set, but no evidence indicated that the accident fan blade dovetail was subjected to an overheat situation during the coating repair process. However, higher-than-expected dovetail operational stresses could also lead to the loss/relaxation of residual stress and premature fatigue crack initiation, which occurred during this event. Residual stress measurements were taken from multiple areas of the dovetail surface on fan blade No. 23 and eight other blades from the accident engine, including three blades with no identified cracks. All nine blades had abnormal residual stress profiles compared with the reference profile data. One method that CFM recommended to maintain the fan blade loads within the predicted range and reduce the overall stresses on the blade root in the contact areas is repetitive relubrication of the fan blade dovetails. As part of the relubrication procedure, the fan blades were visually inspected for crack indications. The investigation of this accident found fan blade cracks that had initiated and propagated underneath the dovetail coating. Because such cracks might not be detected during a visual inspection, CFM implemented, in March 2017, an on-wing ultrasonic inspection method to detect cracks with the coating still on the dovetail. A review of the fan blade overhaul process found that a fluorescent penetrant inspection (FPI) was performed (as specified in the CFM engine shop manual) during the fan blade set's last overhaul in August 2007 to detect cracks. As a result of this accident, CFM implemented, in November 2016, an eddy current inspection (ECI) technique for the fan blade dovetail as part of the overhaul process (in addition to the FPI). An ECI has a higher sensitivity than an FPI and can detect cracks at or near the surface (unlike an FPI, which can only detect surface cracks). The shims for fan blade Nos. 22 through 24 had a newer design configuration that was introduced after the blades and their associated hardware were first installed on the accident disk. Wear patterns on the shims from blade Nos. 23 and 24 showed that a significant area of coating was missing from the blade dovetails at the time that the shims were installed. The location of the missing coating was in the area where overhaul of the blades was required within 50 cycles. The National Transportation Safety Board could not determine, with the available evidence, how long the dovetail coating damage existed and whether the cracks were large enough to be detected by an FPI during an overhaul in response to the coating damage (when the dovetail coating damage first became significant enough during visual inspections conducted at relubrication to trigger the overhaul). The damage to the accident inlet and fan case showed that there were significant differences between the accident FBO event and the engine FBO containment certification tests. For example, during the accident FBO event, the fan blade fragments that went forward of the fan case and into the inlet had a greater total mass and a different trajectory (a larger exit angle) and traveled beyond the containment shield. Also, the inlet damage caused by these fan blade fragments was significantly greater than the amount of damage that was defined at the time of inlet certification. Given the results of CFM's engine FBO containment certification tests and Boeing's subsequent structural analyses of the effects of an FBO event on the airframe, the post-FBO events that occurred during this accident could not have been predicted.
On August 27, 2016, about 0922 central daylight time, Southwest Airlines (SWA) flight 3472, a Boeing 737-7H4, N766SW, experienced a left engine failure while climbing through flight level 310 en route to the flight's assigned cruise altitude. Portions of the left engine inlet departed the airplane, and fragments from the inlet impacted the left side of the fuselage, creating a hole. The airplane depressurized, and the flight crew declared an emergency and diverted to Pensacola International Airport (PNS), Pensacola, Florida, where the airplane made an uneventful single-engine landing. The 2 pilots, 3 flight attendants, and 99 passengers were not injured. The airplane was substantially damaged. The regularly scheduled passenger flight was operating under the provisions of Title 14 Code of Federal Regulations (CFR) Part 121 from Louis Armstrong New Orleans International Airport, New Orleans, Louisiana, to Orlando International Airport, Orlando, Florida. The flight was the first of the day for the airplane and the flight crewmembers. According to the cockpit voice recorder (CVR), the left engine started at 0903:21, and the right engine started 1 minute later. Air traffic control (ATC) cleared the flight for takeoff at 0907:08. The takeoff and climb were uneventful until 0921:45, when the CVR recorded a "bang" sound followed by the sound of a decrease in engine rpm. The flight data recorder (FDR) showed that the airplane's altitude was 31,259 ft. FDR data also showed that, by 0921:49, the left engine fan speed had decreased from 99% to 39% rpm. At 0921:52, the captain asked the first officer to declare an emergency. Three seconds later, the first officer contacted ATC to declare an emergency and told ATC that the airplane was descending. At 0922:06, the flight crew completed the SWA Engine Fire or Engine Severe Damage or Separation checklist. At 0922:17, the CVR recorded a sound similar to the cabin altitude warning horn; FDR data showed that the cabin altitude warning parameter had transitioned from "No Warn" to "Warn" 1 second later. (The parameter remained at "Warn" until the FDR data ended at 0929:37 due to a loss of electrical power to the FDR.) At 0922:18, the left engine fan speed had further decreased to 33% rpm. At 0922:43, ATC told the flight crew that PNS was about 70 miles ahead of the airplane's position. At 0922:51, the CVR recorded the sound of the flight crewmembers using their oxygen masks. Two seconds later, ATC cleared the airplane to flight level 180. At 0926:47, the captain communicated with the cabin crewmembers about the engine failure and instructed them to secure the cabin. At 0929:22, the flight crewmembers commented about high vibration levels and indicated that they were going to keep the airplane's speed up. The FDR recording ended at 0929:37; the last recorded data (1 second earlier) showed that the airplane was at an altitude of about 17,400 ft and that the left engine fan speed was 17% rpm. At 0931:38, the flight crew told the cabin crew that there would not be an evacuation. At 0932:10, the CVR recorded the flight crewmembers removing their oxygen masks. At 0935:17, the flight crew made another comment about high vibration levels. At 0937:39, ATC told the flight crew to expect the instrument landing system (ILS) runway 17 approach. Fifteen seconds later, the flight crew selected flaps 1 and 5 to assess the airplane's handling. At 0938:19 and 0938:30, the flight crew selected autobrake level 3 and flaps 15, respectively. At 0939:10, ATC cleared the airplane for the ILS runway 17 approach. Twelve seconds later, the flight crew completed the descent and approach checklists. At 0940:01, the flight crew reported that the airport was in sight; 12 seconds later, the crew lowered the landing gear. At 0941:40, the flight crew again commented about the vibration. Eleven seconds later, the CVR recorded the sound of the 500-ft automated callout. At 0942:30, the CVR recorded the sound of the airplane touching down on the runway, about 21 minutes after the left engine failure occurred. This report will discuss that the left engine failure occurred when a fan blade fractured and exited the engine fan case, which is referred to as a fan-blade-out (FBO) event. An FBO event consists of four phases (the impact phase, the engine surge phase, the engine rundown phase, and the windmilling phase) during which the airplane structure is subjected to various loads. AIRCRAFT INFORMATION The airplane was equipped with two CFM International CFM56-7B24 turbofan engines, with one engine mounted under each wing. CFM International was established in 1974 as a partnership between General Electric Aviation (GE), a US manufacturer, and Safran Aircraft Engines (Safran), a French manufacturer formerly known as Snecma. (GE manufactured the CF6 engine, and Snecma manufactured the M56 engine; those engine designations were combined to form the new company and engine names.) The CFM56-7B24 engine is one of several CFM56-7B models (other models are the -7B18, -7B20, -7B22, -7B26, and -7B27) installed in Boeing 737 next-generation-series airplanes (the 737-600, -700, -800, and -900). The CFM56-7B references throughout this report apply to all of the engine models. The inlet, which is part of the Boeing 737-700 airframe, is an aerodynamic fairing that guides air into and around an engine. The inlet is attached to the front of each engine's fan case. The inlet is part of the nacelle, which houses the engine. (Other nacelle components include the fan cowl and the thrust reverser.) The CFM56-7B engine and the Boeing 737-700 inlet are further described in the sections below. Engine The CFM56-7B is a high-bypass, dual-rotor, axial-flow turbofan engine. A single-stage high-pressure turbine (HPT) drives the nine-stage high-pressure compressor (HPC). A four-stage low-pressure turbine (LPT) drives the booster assembly, which comprises the engine fan and three-stage low-pressure compressor. The engine rotates clockwise (aft looking forward). The engine consists of three major assemblies: the fan, engine core, and LPT. GE is responsible for manufacturing the HPC, combustion chamber, and HPT (collectively referred to as the engine core). Safran is responsible for manufacturing the LPT and the booster assembly. Both companies assemble the engines; those assembled by GE are identified by an even engine serial number (ESN) prefix (for example, 874), and those assembled by Safran are identified by an odd ESN prefix (for example, 875). The ESN of the accident engine, 874112, showed that GE assembled the engine. The fan and booster assembly comprises the front and aft spinner cones, fan disk, fan blades, booster rotor, booster vanes, and associated hardware. The fan disk, which is secured to the booster, has 24 fan blade slots. In accordance with the instructions in Boeing's 737-600/700/800/900 Aircraft Maintenance Manual, the fan blades are numbered sequentially (1 through 24) in the counterclockwise direction (forward looking aft). Each CFM56-7B fan blade had a chord (width) of about 11 inches at its widest point. The nominal weight of each fan blade is about 10.83 pounds. The fan blades are made of a titanium alloy (known as Ti-6-4), and the dovetail part of the fan blade, which slides into the fan disk, has a copper-nickel-indium coating to protect the blade from wear and provide a better fit with the fan disk (compared with no coating). Before the application of the fan blade coating, the entire blade, including the dovetail, is shot-peened to increase the fatigue strength of the material and reduce surface tensile stresses that can lead to cracking. (Shot-peening is a process that adds a compressive residual stress surface layer to material, and residual stress is the stress that is present in solid material in the absence of external forces.) A spacer is installed under each fan blade root primarily to limit fan blade radial (outward) movement. The spacer also ensures that axial (longitudinal) loads would be transmitted to the fan blade axial retention feature if an FBO event or bird ingestion were to occur. A platform is installed on both sides of each fan blade to provide a smooth aerodynamic flow path between the blades. A shim is installed over each fan blade dovetail to prevent fretting (wear) of the fan disk pressure faces and improve lubrication durability, which reduces the amount of stress on the fan blade dovetail and the fan disk pressure faces. The shims for fan blade Nos. 22 through 24 had a newer design configuration than the other shims installed in the fan assembly, which had the previous design configuration. The newer design configuration for the shims was introduced in February 2008 to improve reliability. Both shim configurations are authorized for CFM56-7B fan blades, and both configurations are interchangeable. The fan disk and spacers are manufactured from a titanium alloy (Ti-6-4). Both shim configurations are manufactured from a nickel-chromium-iron alloy (alloy 718). The platforms are manufactured from an aluminum alloy. Figure 1 shows the CFM56-7B engine fan assembly. Source: CFM Figure 1. Fan assembly. The fan frame assembly is the main forward support for the installation of the engine to the airframe and includes the fan frame, the fan case, and the fan outlet guide vanes. The fan case, which is made of an aluminum alloy, was designed to provide fan blade radial containment if an FBO event were to occur and transmit FBO loads to the fan frame and the inlet. Although the fan case provides the primary FBO radial containment protection, the inlet, which is attached to the fan case A1 flange, provides additional FBO protection. Figure 2 shows a cross-section of the engine and airframe components. Source: CFM Figure 2. Cross-section of engine and airframe components. Inlet The inlet is attached to the engine at the fan case A1 flange with 24 bolted assemblies (fasteners). The inlet directs smooth uninterrupted airflow into the engine fan and core sections and is an external aerodynamic pressure surface that directs the airflow over other nacelle components. Boeing created the design specifications for the inlet. Rohr Industries, which later became UTC Aerospace Systems (UTAS), was the inlet manufacturer. (In November 2018, UTAS and Rockwell Collins merged and became Collins Aerospace.) The inlet consists of the inlet lip, inner and outer barrels, forward and aft bulkheads, and attach ring. The inlet lip is an aerodynamic surface to reduce airplane drag. The inner and outer aft ends of the inlet lip are attached to the forward bulkhead. The inner and outer barrels are concentric structures connected to the forward and aft bulkheads. The inner barrel consists of an acoustic honeycomb core; an aluminum containment shield (also referred to as a containment doubler), which is bonded over an area comprising about the aft 11 inches of the inner barrel; two bolted splice plates at the three and nine o'clock positions, which is where the two halves of the inner barrel join together; perforate skin; and back skin. The outer barrel comprises three panels with two support frames. The attach ring is located at the aft end of the inlet and is secured to the engine fan case A1 flange. The interface between the inlet attach ring and the engine fan case A1 flange is considered to be critical to the inlet's FBO capability because engine loads are reacted at that interface. The aft bulkhead and the inner barrel are also secured to the attach ring. Figure 3 shows a cross-section of the inlet. Source: Boeing. Figure 3. Inlet cross-section. Note: The upper fasteners attach the aft bulkhead to the outer barrel. The lower fasteners attach the aft bulkhead to the attach ring, which is attached to the inner barrel and the engine fan case A1 flange. The fan case and the A1 flange are shown in figure 2. The inlet lip, outer barrel, and attach ring are made of an aluminum alloy. The forward and aft bulkheads are made of a titanium alloy. The inner barrel is primarily made of an aluminum alloy. The 24 bolted assemblies that attach the inlet to the fan case include a crushable spacer that is designed to absorb energy and compress during an FBO event. Aerodynamic and inertia loads on the outer barrel are transmitted to the inner barrel through the forward and aft bulkheads. Loads on the aft bulkhead are transmitted to the interface between the inlet attach ring and the fan case A1 flange. Engine and Airframe Certification The CFM56-7B engine was jointly certificated by the Federal Aviation Administration (FAA) and its counterpart in France, the Direction Générale de L'Aviation Civile (DGAC). The FAA issued the type certificate for the engine model on December 17, 1996, and the certification basis was 14 CFR Part 33, Airworthiness Standards: Aircraft Engines (amendment levels 33-1 through 33-15). The DGAC certificated the engine in December 1996 under Certificat de Type Moteur M21, which was superseded by European Aviation Safety Agency (EASA) type certificate EASA.E.004 in 2006. (In 2004, EASA assumed responsibility for the certification of CFM engines.) There were no significant differences between the FAA and DGAC certification requirements, and the engine met all requirements of both agencies. Because the engine was dual certificated, the certification basis also included Joint Aviation Requirements JAR-E Change 8 (dated May 4, 1990). To meet Part 33 requirements and obtain data for Boeing (as the airframe manufacturer) to use to meet certification requirements under 14 CFR Part 25, Airworthiness Standards: Transport Category Airplanes, CFM performed eight development FBO rig tests and two engine FBO containment certification tests. The purposes of the FBO tests were to (1) understand the fan blade fragmentation and kinematics (fragment energy levels and trajectories) after blade separation, (2) determine the fan case containment capability (radial containment), (3) define the loads and displacements from the initial impact and the resulting engine imbalance, (4) calculate the speed and weight of any ejected fragments (forward containment), and (5) demonstrate the proposed production hardware configuration. Each rig test had a specific set of objectives and used various fan blade and fan case configurations. The first four FBO rig tests were designed to define the fan blade fragmentation and kinematics for fan case radial containment capability, fan blade axial retention, and fan blade interaction. The fourth FBO rig test included a full set of production-representative fan blades, a production-representative fan case, and a combination of actual and production-representative engine accessories. Also included in this test was a Boeing production-representative inlet, which had the same size, shape, and stiffness of the intended production inlet at that time. The production representative inlet included crushable spacers but did not include a containment shield. After a fan blade release, fan blade airfoil and tip fragments generally travel in a forward spiral/helical pattern around the fan case and inlet. The forward movement is due to several factors, including the slope/conical nature of the fan case and inlet surfaces, the angle of attack of the blades, and aerodynamic loading in the forward direction. According to the results of FBO rig test 4, which was conducted in August 1995, the separated fan blade fractured into five pieces with the blade tip panel traveling forward of the fan case and into the inlet at an estimated 10° to 15° helix angle (where the fragment crossed the A1 flange). The blade tip panel then spiraled around the inlet and penetrated the inlet inner and outer barrels. The inner barrel penetration was about 26° forward of the fan blade rotational plane, which was beyond the ± 15° fragment spread angle (impact area) referenced in FAA Advisory Circular 20-128, "Design Considerations for Minimizing Hazards Caused by Uncontained Turbine Engine and Auxiliary Power Unit Rotor Failure," as a practical design consideration to minimize the hazards of uncontained fragments. The test also revealed a small penetration hole in th
A low-cycle fatigue crack in the dovetail of fan blade No. 23, which resulted in the fan blade separating in flight and impacting the fan case. This impact caused the fan blade to fracture into fragments that traveled farther than expected into the inlet, which compromised the structural integrity of the inlet and led to the in-flight separation of inlet components. A portion of the inlet struck the fuselage and created a hole, causing the cabin to depressurize.
Source: NTSB Aviation Accident Database
Aviation Accidents App
In-Depth Access to Aviation Accident Reports
