DCA25LA061

The airplane suffered a fracture in the forward pressure bulkhead canted web causing a rapid depressurization during flight. The crew declared an emergency, descended and continued to the destination airport where they landed safely. The accident airplane was manufactured in 1975 and had accumulated 36,034 flight cycles. Examination of the fractured canted bulkhead web revealed multiple fatigue cracks that initiated at the forward and aft surfaces of the web adjacent to a vertical line of rivets at the airplane centerline. The largest fatigue cracks propagated from the aft face through 90%-95% of the web thickness prior to intersecting smaller fatigue cracks which propagated from the forward face. The fatigue cracking was present near 21 consecutive fastener holes. Several areas exhibited black/brown and greenish deposits on the mating fracture faces in areas that contained the larger fatigue cracks indicating the larger fatigue cracks were present for some time before the accident flight. Trace evidence of chlorine was present near the edge of the fracture. Chlorides are known contaminants that promote corrosion in aluminum alloys. The investigation did not determine the source of chlorine contamination. Detailed examination of the fatigue crack initiation sites revealed the presence of isolated areas of intergranular and branched cracking. Intergranular cracking, chlorine contamination, and a constant or prolonged state of applied or residual tensile stress typically are associated with stress corrosion cracking in aluminum alloys. There were no material discrepancies noted on the fractured web. Areas between fatigue cracks and at the upper and lower ends were due to overstress separation. A fatigue crack striation study indicated that the largest cracks likely initiated from the aft surface of the web about 4,098 cycles prior to the accident flight. Lockheed published Alert Service Bulletin (SB) A382-53-67 on March 10, 2016, applicable to all model 382, 382B, 382E, 382F, and 382G airplanes based on information from a previous investigation DCA16LA096. The bulletin recommended a detailed visual inspection of the forward side of the FS 93 bulkhead web for cracks and any loose or missing fasteners; however, an inspection of the aft surface was not required. The bulletin specified the initial inspection should be performed within 90 days on any aircraft that had accrued more than 15,000 flight cycles with repetitive inspections recommended every C-check not to exceed 3 years or 3,000 flight cycles. According to maintenance records, the most recent SB inspection on the airplane accident was completed on November 7, 2022, with no discrepancies noted. In addition, there was no record of any previous repairs to the canted bulkhead web. Following this accident the manufacturer reported that a new Service Bulletin is in development to add an ultrasonic inspection of the FS 93 canted bulkhead web for cracking. They estimated that the bulletin would becompleted and FAA approved by the end of calendar year 2026. In addition, the manufacturer anticipated that a maintenance action to remove and replace the FS 93 bulkhead web after a specified number of cycles would be added to the maintenance program once further engineering analysis was completed.

Safety Action Lockheed reported that a new SB is in development to add an ultrasonic inspection of the FS 93 canted bulkhead web for cracking. They estimated that the bulletin would be completed and FAA approved by the end of the calendar year 2026. Lockheed also anticipated that a maintenance action to remove and replace the FS 93 bulkhead web after a specified number of cycles would be added to the maintenance program following completion of their engineering analysis. The accident airplane was purchased by Lynden Air Cargo in 1997 and had accrued 70,704:36 hours and 36,034 cycles at the time of the accident. After a similar accident in 2016, Lockheed developed an alert SB recommending a detailed visual inspection of the forward side of the FS 93 canted bulkhead web for cracks and any loose or missing fasteners. The first SB inspection of the accident airplane’s bulkhead web was accomplished on March 23, 2016, after the airplane had accrued 63,827:54 hours and 31,705 cycles with no discrepancies. Subsequent SB inspections of the bulkhead web were accomplished on September 11, 2016 (64,217:31 hours / 32,021 cycles), July 30, 2019 (66,552:24 hours / 33,465 cycles), and November 7, 2022 (68,898:23 hours/ 34,728 cycles) with no discrepancies reported. In addition, there was no record of any previous repairs to the canted bulkhead web. On December 18, 2024, about 0400 Alaska standard time (AKST), Lynden Air Cargo, LLC flight LYC141, a Lockheed L-382C (Hercules) airplane, N403LC, experienced a rapid depressurization while climbing through FL220 (22,000 ft) after departure from the St. Mary’s airport (KSM), St. Mary’s, Alaska. There were no injuries to the four occupants onboard the airplane. A postaccident inspection revealed that the airplane sustained substantial damage. The cargo flight was operating under the provisions of Title 14 Code of Federal Regulations Part 121 from KSM to the Ted Stevens Anchorage International Airport (ANC), Anchorage, Alaska. The flight crew consisted of the captain who was the pilot monitoring, the first officer (FO) who was the pilot flying and the flight engineer. In addition, a load master was present on the flight. The flight crew reported that after departure from KSM and while climbing through about FL220 they heard a loud bang followed by a loss of pressurization. They donned their oxygen masks and initiated a descent to 13,000 ft mean sea level (msl). With no visible signs of damage, the flight crew suspected a bleed air problem. The quick reference handbook (QRH) was reviewed and all bleed air systems were turned off. The flight crew elected to continue to their destination. About 30 minutes before arrival to ANC, the number 3 engine’s torque began to fluctuate. Concerned that the torque fluctuations were related to the bleed air systems, they decided on a precautionary shutdown of the number three engine in accordance with the QRH and company procedures. An emergency was declared with air traffic control andthe flight continued to ANC and landed without further incident. Post flight inspection of the fuselage station (FS) 93 forward pressure bulkhead behind the radome (figure 1) found a through-crack, about 14 inches long, in the canted web area adjacent to the fasteners along the centerline of the airplane at buttock line (BL) 0. The fractured area was excised by maintenance personnel and shipped to the NTSB Materials Laboratory for further examination. Figure 1. Location of the through-crack found in the canted bulkhead web is shown within the red circle. (Source: Lynden Air Cargo LLC) Lockheed Inspection – N403LC Accident Airplane The canted bulkhead web was removed and sent to Lockheed for inspection after an ultrasonic inspection was performed in April 2025. The bulkhead had a cut-out area along BL 0 where the repair was performed after the accident flight. There was an additional area that was cut out on the right side of the bulkhead for unknown reasons. Lockheed performed a surface eddy current (SEC) inspection of the aft surface of the bulkhead web. Crack indications were found adjacent to the right side of additional buttock line (BL) 0 fastener holes. There was an indication that extended from the upper edge of the BL 0 cut-out area and an indication to the right of the first BL 0 fastener hole above the cut-out. There were indications to the right of the two BL 0 fastener holes below the cut-out area. Additional crack indications were found to the right of 4 fastener holes at right buttock line (RBL) 7, 5 fastener holes at RBL 15, 2 fastener holes at RBL 27, 2 fastener holes at left buttock line (LBL) 15, and 5 fastener holes at LBL 27. One of the indications at LBL 27 had progressed through the thickness of the web. Through-thickness crack indications were also found adjacent to the lower left fastener hole and the two lower right fastener hole locations in the bulkhead web. Previous Accident On February 16, 2016, about 0745 Alaska standard time, a Lynden Air Cargo, LLC., Lockheed 382G airplane, N401LC, experienced a rapid depressurization during cruise flight. The flight crew declared an emergency and diverted to the nearest airport where they landed safely. The NTSB investigated the accident, DCA16LA096, and determined the probable cause was the fatigue failure of the forward pressure bulkhead canted web that originated from stress corrosion cracking areas. Contributing to the failure were inspection procedures and intervals inadequate to identify the crack. The investigation found multiple fatigue cracks that initiated from the aft and forward faces of the web adjacent to 21 consecutive rivet holes at the centerline of the airplane at BL 0. The fatigue fractures from the aft face propagated through 85%-90% of the web thickness while the fatigue fractures from the forward face propagated through the remaining thickness. Several of the individual fatigue fractures had linked up to form a continuous crack about 10.5 inches long thatshowed evidence of black-like deposits on and adjacent to the fractures. A fatigue crack striation study performed for the investigation indicated that the cracks likely initiated about 5,000 cycles prior to the accident flight. Postaccident examination of the submitted portion of the fractured bulkhead web revealed a 14.4 inch vertically oriented through-crack located right-adjacent to a row of fastener holes where a C-channel stiffener was fastened to the aft side of the bulkhead. The fastener hole locations were labeled H1 through H23 beginning with the upper most hole on the submitted web section. The overall profile of the crack was relatively flat and straight relative to the line of fastener holes but exhibited localized steps in between many of the holes. The crack path deviated toward the right at the upper and lower ends near hole 4 and between holes 22 and 23, respectively. Additional vertically oriented cracks were observed above and below the primary crack right-adjacent to holes 1 – 2 and 23. Figure 2. Photographs of the forward surface (left) and aft surface (right) of a portion of the canted bulkhead web with the fastener holes numbered H1 – H23. The forward surface had a green coating with a satin finish. Circular witness marks were observed around each fastener hole, consistent with the presence of a rivet or fastener head. The aft surface had a lighter green/brownish coating with a matte finish and a lighter green band which extended along the rivet line, consistent with the footprint of a stiffener element which had been attached before the fasteners were removed. The fracture was exposed and the left and right fracture surfaces were examined under magnification (4X – 80X stereomicroscope) and various lighting conditions. The visual exam revealed features resembling ratchet marks emanating from the aft and forward surfaces between holes 4 - 22, which appeared more prominent along the aft surface. The fracture surface was generally oriented perpendicular to the adjacent bulkhead web surfaces on many offset planes in this region and exhibited a relatively shiny appearance. Several curved thumbnail shaped regions were observed to radiate from the aft and forward sides, consistent with crack arrest marks. These features were all consistent with progressive fatigue cracking from multiple initiation sites. Closer examination of the fatigue crack origin areas near the aft and forward edges revealed intergranular fracture features which extended to a depth up to 0.003 inch. The upper and lower ends of the fracture between holes 3 – 4 and holes 22 – 23, respectively, where the fracture path deviated away from the fastener row, exhibited a relatively rough morphology, matte appearance, and was slanted through the bulkhead web thickness, consistent with overstress separation. Closer examination of the aft and forward web surfaces adjacent to the fracture surface and near fastener hole locations revealed smaller cracks oriented parallelto the primary fracture surface. When viewed in cross-section, some of these surface initiated cracks extended through the clad layer into the base metal and branched into multiple crack pathways. The presence of chlorine was detected at multiple locations near the aft edge of the fracture surface. These observations coupled with the intergranular fracture features were consistent with a stress corrosion cracking mechanism. The manufacturer's drawing specified the web material as 7075-T6 aluminum alloy with a 7072 aluminum alloy cladding layer on both surfaces. Testing was performed to verify the material composition, heat treatment condition, and microstructure along with the thickness of the web, clad layer, and paint layers. All were within specifications. To estimate the fatigue crack growth rate from the aft surface, a striation count study was carried out over the fracture surface near hole H20. After cleaning, a series of scanning electron microscope (SEM) images were obtained along a relatively linear path from a fracture origin area at the aft surface to the end of the crack near the forward surface. Striation spacing was measured and averaged over the incremental distances represented by these images. From these average growth rates, a cumulative striation count was estimated to be from 3,331 to 4,098 fatigue loading cycles.