WPR13IA144

The failure and separation of a section of the No. 2 engine's combustion chamber's small exit duct (SED), which created an airflow disruption that led to an engine surge and subsequent fire. Contributing to the accident was the insufficient weld penetration that remained totally contained within the SED outer dome and did not penetrate through to the inner duct as required by the manufacturing specifications.

The flight was climbing through flight level 200 when the flight crew received a No. 2 engine audible fire warning, which was accompanied by a bang sound. The crew subsequently shut the No. 2 engine down and discharged both of the engine's fire extinguishing bottles. The cabin crew reported no visible fire, but the flight crew continued to receive a fire indication. The captain then declared an emergency, began a descent to the departure airport, and landed about 9 minutes after the event. After the airplane was stopped on the runway and the No. 1 (left) engine was shut down, an emergency evacuation was conducted using the main cabin door. An initial assessment by company personnel indicated that the fire damage to the airplane was limited to what was visible on the aircraft's exterior. A visual postincident examination of the engine revealed that sooting was present on external case surfaces, which were concentrated on the aft half of the No. 2 engine. A disassembly of the engine confirmed that about half of the SED outer dome within the combustion section had separated and was subsequently ingested into the gas path. To determine the reason for the separation, three locations on the outer dome were sectioned and examined. All three locations exhibited insufficient weld penetration so that the welds remained totally contained within the SED outer dome and did not penetrate through to the inner duct as required by the manufacturing specifications.

HISTORY OF FLIGHT On March 5, 2013, about 1120 Pacific standard time, a Bombardier Inc. DHC-8-402 airplane, N417QX, sustained minor damage following an in-flight engine fire during initial climb near Danville, California. The airplane, which was registered to Horizon Air Industries of Seatac Washington, was being operated in accordance with 14 Code of Federal Regulations Part 121. Visual meteorological conditions prevailed at the time of the accident, and an instrument flight rules (IFR) flight plan had been filed and activated. The flight departed the Norman Y. Mineta San Jose International Airport (SJC), about 1108, with its destination being the Boise Air Terminal (BOI), Boise, Idaho. In a written statement provided to the National Transportation Safety Board (NTSB) investigator-in-charge (IIC), the pilot-in-command (PIC) reported while climbing en route through 18,000 feet mean sea level (msl), the flight crew heard a muffled "boom" in the flight deck, immediately followed by the crew noticing that the number 2 engine's internal turbine temperature (ITT) was exceeding limits. The PIC stated that some yaw was observed as well, which indicated a loss of power on the number 2 engine. The flight crew received and engine fire warning for the number 2 engine, which was then followed by the crew performing the emergency Engine Fire Checklist. At this time an emergency was declared with Air Traffic Control (ATC). The PIC revealed that in accordance with the emergency checklist, the crew discharged the first [fire] bottle; however, the fire warning on the flight deck continued. A second bottle was discharged, which was consistent with the emergency checklist, but the fire warning remained illuminated. The flight crew then requested ATC clearance directly back to SJC. The PIC revealed that while the Oakland International Airport (OAK), Oakland, California was slightly closer in terms of miles, the aircraft's altitude would have required circling, so the crew elected to proceed directly towards SJC. At this time the PIC reported that the flight crew advised ATC of the situation, and asked them to have the fire trucks ready on landing. The PIC stated that the weather at the time was consistent with visual meteorological conditions, with inflight visibility better than 25 miles. The pilot revealed that after touchdown at SJC, the aircraft was brought to a complete stop on the runway. The flight crew asked emergency personnel if the fire could be seen on the exterior of the engine; no smoke or fire was reported. The flight crew then elected to make an evacuation through the main cabin door. The passengers were taken to the side of the runway and the crew secured the airplane. Buses were used to transport the passengers and crew to the airport terminal. ENGINE EXAMINATION General According to an NTSB powerplants aerospace engineer who participated in the postaccident examination of the engine, the engine was removed from the airplane and shipped to the Pratt and Whitney Canada (PWC) facility in St. Hubert, Quebec, Canada where it was inspected and disassembled. During disassembly, a section of the small exit duct (SED) located within the combustion section was found to have separated and been ingested into the gaspath. All internal hardware aft of the combustion section exhibited varying degrees of thermal damage and sooting consistent with an airflow disruption. The remaining intact section of the SED was examined at the PWC materials lab where improper weld penetration between the SED outer dome and the combustion housing inner duct was identified based on drawing specifications. The advanced pneumatic detection (APD) fire loop exhibited thermal distress including deformation of internal diaphragm switches. The deformation resulted in a fire warning to the flight crew that remained latched even after temperatures levels had decreased below warning limits. In response to APD system related failures, Kidde Aerospace made design improvements to the system and Bombardier released service bulletins (SBs) 84-26-08A, 84-26-09A, and 84-26-12A for removal and replacement with the updated model. Airworthiness directives (ADs) were subsequently released by Transport Canada and the United States Federal Aviation Administration (FAA) mandating replacement. A quality investigation into the improper weld penetration during production of the SED assembly identified an inadvertent adjustment to the electron beam (EB) current level during the manufacturing process as the probable cause. A lot of five SED assemblies were identified as being potentially affected by the quality escape. Two of the assemblies, including the SED in the incident aircraft engine experienced an in-flight failure and the remaining three engines containing suspect assemblies were removed for repair. Engine Information History According to Horizon Air's maintenance records, engine serial number (ESN) PCE-FA0184 had accumulated the following hours and cycles: 18,764 hours time since new (TSN), 19,351 cycles since new (CSN), 4,244 hours time since repair (TSR), and 4,391 cycles since repair (CSR). Engine Description The PWC model PW150A engine is a three spool free-turbine turboprop engine incorporating a three-stage axial compressor followed by a centrifugal compressor stage, each driven by independent axial turbines, a reverse flow annular combustor, and a two-stage power turbine that drives an offset reduction gearbox. The engine fuel flow is controlled by a Full Authority Digital Electronic Control. The engine has a maximum rating of 5,071 shaft horsepower (SHP) and a continuous rating of 4,580 SHP. Engine Disassembly Examination and Documentation An initial visual examination of the engine was conducted at SJC. Sooting was observed on external case surfaces, concentrated on the aft half of the engine. The engine's flexible lines exhibited thermal distress including melting and charring of line sheathing. A borescope examination of the engine revealed that a piece of the SED located in the combustion section had separated. The 1st stage turbine nozzle vanes exhibited hard particle damage along the leading edge. The engine was shipped to PWC in St. Hubert, Quebec, Canada for further inspection and teardown. As Received Upon receipt of the engine at the PWC facility in Quebec, it was reported that the engine was intact without indications of case breaches or deformation. The external engine components were sooted from the aft turbine support case flange forward to the accessory gearbox (AGB). Soot concentration was heaviest on the aft portion of the engine and was progressively lighter moving forward. The indicated turbine temperature (ITT) harness exhibited thermal distress including charring and blistering 360° around the engine. Other flexible lines such as the flexible fuel manifold harness exhibited blistering and had a grayish color consistent with thermal exposure, but maintained flexibility. The oil pressure supply line to the No. 7 bearing was thermally damaged exposing the internal braided cord. The engine magnetic chip detector plugs were removed and inspected for indications of abnormal metal wear accumulation. The reduction gearbox (RGB) scavenge plug and A/C generator plug were free of debris. The turbo machine (T/M) magnetic chip detector plug had metal "fuzz", consistent with normal engine operation according to PWC. The RGB and main oil filters were removed and were free of debris or blockages. Power turbine gear train continuity was verified by manually spinning the propeller shaft and observing concurrent rotation of the 2nd stage power turbine. Disassembly and Inspection Findings The reduction gearbox (RGB) was not disassembled due to the absence of evidence suggesting internal damage. The RGB propeller shaft spun freely without binding. The compressor assembly/impeller assembly was intact and in good condition. Impell