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28 Aug 2016 - 737-700, N766SW (29806/537), FF 12/4/2000, operated by Southwest was passing FL310 in the climb about 80nm west of Pensacola whilst enroute from New Orleans to Orlando on 28 Aug 2016 when it suffered an engine failure. This caused the inlet cowl to detach from the engine. The cowl damaged the left side of the fuselage causing a loss of cabin pressure.

The NTSB published their final report in December 2020 and determined the probable cause(s) of this accident to be: 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.

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The inlet cowl is aerodynamically shaped to send air into the fan, it is also contains acoustic absorbtion material to reduce noise. It has a fan inlet temperature probe (T12) and two ram air scoops for engine anti ice and EEC cooling.

N766SW Uncontained engine failure

Analysis

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 fanblade- 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-thanexpected 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.

The full report can be found here

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