McDonnell Douglas 369E crash near Burns, OR, United States

HISTORY OF FLIGHTOn August 12, 2013, about 1400 Pacific daylight time, a McDonnell Douglas Helicopter Inc. (MDHI) 369E, N62PJ, sustained substantial damage during a forced landing following a loss of engine power near Burns, Oregon. The helicopter was registered to and operated by PJ Helicopters, Red Bluff, California, under the provisions of Title 14 Code of Federal Regulations Part 91. The commercial pilot and his passenger were not injured. Visual meteorological conditions prevailed, and a company visual flight rules flight plan was filed for the local public aircraft flight. The flight originated from a staging area about 1245. The pilot reported that during cruise flight at an altitude about 400 feet above ground level over densely wooded mountainous terrain, he heard a "loud bang" originate from the engine followed by a left yaw and a "severe medium frequency vibration." The pilot performed a 180-degree left turn autorotation to a nearby forest service logging road. During the landing, the main rotor blades struck trees adjacent to the road, and the left skid slid into a ditch, which resulted in the helicopter rotating 90-degrees and subsequently the separation of the left skid. Examination of the helicopter by the pilot revealed that the tailboom was separated, and the left landing skid was separated. The helicopter was recovered to a secure location for further examination. AIRCRAFT INFORMATIONThe helicopter was equipped with a Rolls-Royce (RR) M250-C20B turbo-shaft engine, which features a 6 stage axial and 1 stage centrifugal compressor section that directs the diffused air via an external 180 degree compressor discharge tube system to the combustor. The hot gases from the combustor are then directed against a two-stage gas producer turbine and subsequently a two-stage power turbine before being exhausted. The RR M250-C20B produces 420 shaft horsepower. The engine serial number (S/N) was CAE-836896. The engine logbook revealed that the engine was overhauled on October 13, 2010, at a time since new (TSN) of 9,202.8 hours, cycles since new (CSN) of 2,840, and cycles since overhaul (CSO) of 529. At the time of the overhaul, new post Service Bulletin (SB) Commercial Engine Bulletin (CEB) 1365 1st, 2nd, 3rd, and 4th stage turbine wheels were installed. At the time of the accident, the engine had 10,449.6 hours TSN and 1,246.8 hours since major overhaul. AIRPORT INFORMATIONThe helicopter was equipped with a Rolls-Royce (RR) M250-C20B turbo-shaft engine, which features a 6 stage axial and 1 stage centrifugal compressor section that directs the diffused air via an external 180 degree compressor discharge tube system to the combustor. The hot gases from the combustor are then directed against a two-stage gas producer turbine and subsequently a two-stage power turbine before being exhausted. The RR M250-C20B produces 420 shaft horsepower. The engine serial number (S/N) was CAE-836896. The engine logbook revealed that the engine was overhauled on October 13, 2010, at a time since new (TSN) of 9,202.8 hours, cycles since new (CSN) of 2,840, and cycles since overhaul (CSO) of 529. At the time of the overhaul, new post Service Bulletin (SB) Commercial Engine Bulletin (CEB) 1365 1st, 2nd, 3rd, and 4th stage turbine wheels were installed. At the time of the accident, the engine had 10,449.6 hours TSN and 1,246.8 hours since major overhaul. WRECKAGE AND IMPACT INFORMATIONThe Rolls Royce M250-C20B turbo-shaft engine was removed from the airframe, and subsequently sent to the Rolls-Royce manufacturing facilities in Indianapolis, Indiana, for examination. The investigation team met on September 18, 2013, for the engine teardown, and to establish further testing and/or examinations that were necessary to complete the investigation. Prior to engine disassembly, the N1 (Gas Producer) rotor was free to rotate, and its drive train continuity between the compressor and the starter generator pad was confirmed. The N2 (Power Turbine) rotor and associated drive train could not be rotated. The gearbox module, S/N CAG 34078, was intact, and appeared to be undamaged. An examination of the two chip detectors revealed one small shiny flake in the lower chip detector while the upper was clean. The compressor module, S/N CAC 38234F and the diffuser scroll case were intact, and appeared to be undamaged. The 1st stage compressor blades were intact, complete, and undamaged; the compressor rotor could be rotated by hand. The compressor assembly was not disassembled further. The gas producer turbine module S/N was CAT 36932. The gas producer turbine support was intact and undamaged. The No. 8 bearing support was undamaged, oil wetted, and oil darkened. The 1st stage turbine nozzle was intact and undamaged. All the airfoils were coated with a flaky black sooty deposit that was easily removed with finger pressure. The 1st stage turbine wheel was intact and undamaged. The curvic coupling was intact and undamaged. The 2nd stage turbine nozzle was intact and undamaged. The 2nd stage turbine wheel and its curvic coupling was intact and undamaged. The turbine-to-compressor coupling was intact and undamaged. The N1 (Gas Producer) spool bearing set consists of the No. 1, No. 2, No. 2-½ , No. 7 and No. 8 bearings. Bearings No. 7 and 8 were intact, undamaged, oil wetted, and could be turned by hand. The left-hand compressor discharge tube (CDT) was undamaged. The right-hand compressor discharge tube was pierced in two locations, and dented in multiple locations at the elbow location, which is in the plane of the 4th stage turbine rotor. The larger pierced hole on the CDT elbow was approximately 2-½ inches long with the width varying from 1/2 inch to 1 inch. The smaller hole had a rectangular shape approximately 1/2 inch x 3/8 inch. The edge material condition of the inside wall of both piercings of the CDT tube was 'petaled' inwards; consistent with a high-speed particle entering the tube, and the edge material condition of the outside wall was 'petaled' outwards, consistent with a high-speed particle material exiting the tube. The outer combustor case was intact and undamaged; however the combustor liner had a 2 inch long radial fracture of the fish-mouth seal at the 8 o'clock location. The combustor liner was also heat distorted at the 8 o'clock location. The heat distortion started as a point at one of the aft most dilution holes, and continued in an increasing pattern until meeting the fish-mouth. The 1st stage nozzle deflector shield heat shield was slightly distorted. The heat shield is a plate that is welded onto the deflector shield, and has a wavy pattern corresponding to the mounting dimples on the deflector shield. The clearance between the edges of the heat shield and the deflector shield is approximately 1/8 inch, however at the 8 o'clock location the edges were in direct contact with the deflector shield material. The exhaust collector was pierced in three locations. One of the exhaust collector center support casting arms was fractured in two locations and found loose. The three pierced holes were at the 9:30, 11:00, and 2:30 o'clock locations on the plane of the 4th stage turbine wheel. The hole at 9:30 o'clock was approximately 3.5 inches long in the circumferential direction, and maximum width of the hole was 1 inch in the axial direction. The material around the tear was petaled outwards, and there was no loss of material. The hole at 11 o'clock was approximately 1 inch long in the circumferential direction with a maximum width of 1 inch in the axial direction; the edges of the hole were fractured, and the entire patch of material was missing. The fracture at 2:30 o'clock was 4 inches long in the circumferential direction, and the opening width varied between ¾ and 1 inch. The missing piece of the housing material was approximately 2 square inches in total area. The power turbine support was intact and undamaged. The 3rd stage turbine nozzle was intact an