Pilot Debrief

AIRBUS A320 232 near Long Beach, CA — 2014-09-18

Final reportENG14IA028
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Date
2014-09-18
Location
Long Beach, CA, USA
Airport
LGB
Aircraft
AIRBUS A320 232
Registration
N656JB
Category
Airplane
Highest injury
None
Fatalities
0
Phase of flight
Initial climb

Probable cause

The probable cause of the engine failure and subsequent undercowl engine fire was due to the fatigue fracture of a high pressure turbine stage 2 disk blade retaining lug that released two blades which impacted the low pressure turbine case causing a fuel line to fracture spraying fuel on the hot engine cases where it ignited. During a machining operation of the disk lug, a tool mark was introduced that set up the area for fatigue cracks to initiate.

NTSB narrative

On September 18, 2014, at approximately 0930 Pacific standard time (PST), an Airbus A320-232, registration number N656JB, flight number 1416, powered by two International Aero Engines (IAE) V2527-A5 turbofan engines, experienced a No. 2 (right) engine failure and subsequent undercowl fire during initial climb after departing Long Beach Airport (LGB), Long Beach, California The flightcrew shutdown the No. 2 engine, discharged both fire bottles, and performed an air turnback to Long Beach. The airplane made a successful and uneventful single-engine landing at LBG The incident flight was a 14 Code of Federal Regulations (CFR) Part 121 domestic passenger flight from LGB to Austin-Bergstrom International Airport (AUS) Austin, Texas. Day visual meteorological conditions prevailed at the time, and an instrument flight rules flight plan was filed. Examination of the outside of the engine revealed a fractured fuel pressure line to the station 2.5 low pressure compressor bleed valve slave actuator and evidence of thermal distress such as consumed, partially-consumed or oxidized insulation blankets, loop clamps cushions, wiring harness sheathing, and sooting of various components and cases. No case breaches or penetrations were noted although the LPT case did exhibited a localized outward bulge. Disassembly of the engine revealed that a single fir tree blade retaining lug from the high pressure turbine stage 2 disk had fractured and 2 HPT stage 2 blades had released. Metallurgical examination of the fractured HPT stage 2 disk lug revealed evidence of fatigue from multiple origins that propagated from the pressure side (PS) of the middle (No. 2) fillet towards the suction side (SS) almost through the entire width of the lug before finally fracturing due to progressive tensile overload. Closer examination of the fractured lug revealed a concave 'divot'/groove in the PS No. 2 fillet, immediately adjacent to the fracture surface that ran the entire length of the fillet. It was concluded that the groove appeared to be a tool mark resulting from the machining (broach) operation during the original manufacturing of the disk. Inspections of other HPT stage 2 disks manufactured using the same broaching tool at the fractured disk found the same grooving. Based on this event, the disk broaching procedures were reviewed and best practices were implemented to address these manufacturing deficiencies.

Analysis

Primary failure mode
Mechanical failure
First missed decision gate
Engine failure and fire warning could have prompted earlier decision to abort takeoff.

NTSB coding

  • Loss of engine power (partial) · Initial climb
  • Loss of Thrust on Takeoff

Evidence available

  • Photos
  • 3 docket documents
View NTSB final reportView NTSB docket

Docket documents3

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