N417QX

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. Impeller blade tips exhibited light rub with corresponding rub marks on the impeller exducer. No damage was noted on diffuser or axial compressor stages.

The accessory gear box (AGB) was not disassembled due to the absence of evidence suggesting internal damage.

The combustion section's fuel nozzles were intact and in good condition with light sooting consistent with engines of similar hours/cycles. The outer combustion liner had dark uniform sooting on the liner walls and material burn through in two locations. Small metal fragments were found resting in the bottom of the outer liner. The inner combustion liner was intact but exhibited sooting of liner walls. There were no indications of thermal damage or burn through. Fuel nozzle floating collars had thermal erosion which is considered a normal condition on the PW150A engine.

The high pressure turbine (HPT) vane assembly includes the SED inner duct and outer dome that are welded together and function to direct combustion gas flow. Pieces of the outer dome had separated from the vane assembly and were subsequently ingested into the gas path. A section of the SED outer dome was submitted to the PWC materials lab for weld analysis. The exposed inner duct of the SED exhibited localized thermal distress including burn through and distortion at approximately the 11 O' Clock position. Sooting, metal splatter and discoloration were noted on all HPT vanes with some metallic fragments adhering to the leading edge of the vane airfoil. Two metal fragments were found resting at the bottom of the HPT blade shroud against the aft side of the vanes.

The HPT disk assembly exhibited impact damage and metal splatter along the blade leading edges 360 degrees around. All blades were heavily sooted. Material loss was noted on leading edge blade tips resulting in exposure of internal blade cooling passages. Uniform tip rubs were present on HPT blades around the disk as evidenced by shiny metal and material smearing.

The low pressure turbine (LPT) stator was intact and in good con...