Aviation Accident Summaries

Aviation Accident Summary WPR13GA374

Burns, OR, USA

Aircraft #1

N62PJ

MCDONNELL DOUGLAS HELICOPTER 369E

Analysis

The commercial pilot was conducting a local public flight. He reported that, during cruise flight about 400 ft above ground level over densely wooded mountainous terrain, he heard a "loud bang" emanate from the engine, followed by a left yaw and a "severe medium frequency vibration." The pilot subsequently 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 subsequent separation of the left skid. Postaccident examination of the airframe and engine revealed that one 4th-stage turbine wheel blade had fractured at the blade root and that several additional turbine blades exhibited damage. Metallurgical examination of the fracture surface area revealed that it exhibited signatures consistent with low-cycle fatigue cracking that had progressed to high-cycle fatigue cracking, which led to the failure of the blade due to overload and resulted in the subsequent loss of engine power. Fluorescent penetrant inspection of the 4th-stage turbine wheel did not reveal any other cracks. In addition, the microstructure, hardness, and composition of the 4th-stage turbine wheel conformed to the engineering drawing requirements. A review of the maintenance logbooks revealed that the 4th-stage turbine wheel was part of a post-service bulletin (SB) commercial engine bulletin (CEB)-1365 or "enhanced" power turbine assembly design. According to the engine manufacturer, the pattern and appearance of the fracture on the accident wheel was similar to fractures on the wheels involved in eight other failures of the post-SB CEB-1365 power turbine assembly design. Representatives from the engine manufacturer reported that the low-cycle fatigue crack was likely initiated by a combination of two factors: (1) a high, positive thermal gradient in the airfoil trailing edge material near the hub during the engine starting process; and (2) the subsequent high, negative thermal gradient and high-speed stress into the blade during transient operation, such as autorotation.

Factual Information

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 oclock location. The combustor liner was also heat distorted at the 8 oclock 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 oclock 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 oclock locations on the plane of the 4th stage turbine wheel. The hole at 9:30 oclock 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 oclock 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 oclock 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 and undamaged. The 3rd stage turbine wheel was intact. Five of the curvic coupling teeth were damaged on their contact faces, however they were still intact. The outer shroud was intact; however a 90 degree segment of the outer shroud labyrinth seal was heavily scored, consistent with contact against the blade track. The 4th stage turbine nozzle was generally intact; however a 190 degree segment of the 4th stage turbine blade track was fractured in the plane of the 4th stage turbine wheel, consistent in location with the fracture of the exhaust collector housing. The 4th stage blade track was heavily rotationally scored. All the nozzle vanes were present and undamaged. The 3rd stage turbine blade track was rotationally scored, consistent with contact against the 3rd stage turbine wheel labyrinth seals. The 4th stage turbine wheel was missing one blade. The missing blade was fractured near the inner rim. A 4-blade span of the shroud was also missing; it was located outboard of the liberated and three adjacent lagging blades. The three lagging adjacent blades were fractured near the tips, were missing the entire rim segment, and were twisted. The leading and lagging adjacent blades were fully attached to the inner platform rim. All the other blades were present and intact, however all were dented at the trailing edges approximately ½ inch inwards from the shroud. Nine of the curvic coupling teeth, which mate to the 3rd stage turbine wheel, were smeared on their contact faces; however, all were still intact. The N2 (power turbine) bearing set consists of the No. 3, No. 4, No. 5 and No. 6 bearings. The No. 5 bearing snap ring was found dislodged from its groove. The forward right hand portion of the lower fire shield was deformed and torn in an area of approximately 4 inches by 2 inches. This damage was adjacent to the damage to the right hand compressor discharge tube ADDITIONAL INFORMATIONThere have been four previous NTSB-investigated failures of 4th stage turbine wheels in RR M250-C20B engines installed in MDHI 369 helicopters, which have had the post-SB CEB-1365 power turbine assembly incorporated. There has been one 4th stage turbine failure in a Bell OH-58 operated by the US Army; however it was reported that this helicopter was used extensively for auto-rotation training, and no further details were available. Service Bulletins Rolls-Royce SB CEB-1365 The enhanced power turbine section was developed by Rolls-Royce as a product improvement, designed to increase both power and fuel efficiency. SB CEB-1365 hardware was a major re-design of the 3rd and 4th stage turbine assembly, with the main differences between the pre- and post-SB CEB-1365 being different airfoil size, shape, tilt, lean, flow, and quantity of airfoils per stage for both turbine nozzles and wheels. The enhanced power turbine design was released for new production engines built after August 1999. It was then released as a customer option to fielded engines viaRolls-Royce SB CEB-1365 in November 1999. The modification applied to all M250-C20 series engines, with the exception of turbo-prop variants, and was to be complied with as a customer option. Release of enhanced power turbine to M250–B17F/2 turbo-prop variants occurred in August, 2008, while release of all other turbo-prop applications was in November, 2009. The previous "non-enhanced" power turbine part numbers were discontinued from production in August 2009, and discontinued from Service/Spares orders in March 2013. Thus, the SB CEB-1365 enhanced power turbine is the only current production and service released hardware. E.1.3.2 Rolls-Royce Alert SB CEB -A-1400 CEB-A-1400, entitled Steady State Operation Avoidance Range Limit was originally released in December 2006, and specified turbine (N2) revolutions per minute (rpm) of the speed ranges that should be avoided. The current revision, CEB-A-1400-revision 3, dated January 19, 2009, advises to avoid steady state engine operation in the N2 speed avoidance range of 75 - 88 percent for operation above 85 shaft horsepower (SHP). Steady-state operation below 85 SHP or transient operation thru the speed avoidance range is allowable. The revised SB requires an entry in the maintenance records documenting steady-state operation in the speed avoidance range when operating above 85 SHP. On November 29, 2010, Rolls-Royce issued CEB-A-1407, which was only applicable to M250-C20B engines installed on MDHI 369 models equipped with post-CEB-1365 hardware. It required a one-time inspection of the 3rd and 4th stage turbine wheels for possible airfoil cracks to be completed within 1,750 hours, and required the removal of the applicable turbine wheels for fluorescent-penetrant and visual inspections. The SB mandated the replacement of any wheel that contained any cracks of an airfoil trailing edge where the platform fillet engages. Additionally, it warned operators to avoid prolonged engine operation in the 75-88 percent N2 speed range, failure of which could result in possible turbine wheel airfoil fracture. The compliance code of this SB was 6: To be complied with at or within next 1,750 hours gas producer turbine overhaul. Revision 1 of Rolls-Royce CEB-A-1407 on February 7, 2011, pertained to M250-C20B engines installed with an enhanced power turbine 4th stage power turbine wheel P/N 23055944 (applicable to the accident engine and helicopter), revealed that the analysis of recent failures of enhanced power turbine wheels determined the most possible cause to be operation of the engine in the N2 speed avoidance range of 75 to 88 percent. The SB called for a one-time inspection of the 3rd and 4th stage turbine wheels for cracks in the airfoils. The compliance code of this SB was 6: To be complied with at the next 1,750 hours. Revision 2 of Rolls-Royce CEB-A-1407, issued on August 2012 stated that 3rd or 4th stage turbine wheels that were installed after September 2011 were not subject to the one-time inspection due to the issuance of an MDHI Helicopter bald rotor blade tracking procedure revision and prior speed avoidance actions regarding tail rotor balancing. The compliance code of this SB was 6: To be complied with at the next 1,750 hours. In addition, this alert was referenced in FAA Airworthiness Directive (AD) 202-14-06. Revision 3 of Rolls-Royce CEB-A-1407, issued on May 19, 2014, after a total of four - 4th stage failures, increased the compliance code from 6 to 2 To be complied with at each 1,750 hour gas producer turbine wheel replacement, a continuing repetitive rather than a one-time inspection. It also retracted Revision 2 by requiring all turbine wheels, even those installed after September, 2011, to be subject to the inspection. This Alert CEB is referenced in FAA AD 2015-02-22. In November 2013, Rolls Royce issued a change in their maintenance manual (MM) requiring the removal and scrap of all post SB CEB 1365 3rd and 4th stage turbine wheels if any temperature limits were exceeded. MDHI Operation Manual Changes On March 21, 2007, MDHI issued Revision 16 of the MDHI 369E flight manual, which incorporated the Rolls-Royce CEB-A-1400 operational limitations to the power turbine by introducing a rotational speed avoid range and a turbine transient temperature limit to operators of the MDHI 369 helicopter. The turbine speed avoid range of 75 – 88 percent was introduced, and the transition time while passing through the range was now limited to 1 minute. Transient temperature limits during start and shutdown were now limited to 10 seconds above 810 degrees C and a maximum of 927 degrees C. On September 30, 2010, MDHI issued Revision 17 of the 369E flight manual, which incorporated two additional limitations. Firstly, steady state operations between 75 and 88 percent were now required to be recorded. Secondly, a caution was added, warning operators to maintain the main rotor speed above 420 rpm during auto-rotation practice maneuvers, to avoid entering the speed avoid range. Airworthiness Directives AD 2012-14-06, released on July 10, 2012, required a one-time visual inspection and fluorescent penetrant inspection (FPI) on post-SB CEB 1365 3rd and 4th stage turbine wheels for cracks in the trailing edges of the turbine blades. It was issued to mandate the incorporation of C

Probable Cause and Findings

The total loss of engine power during cruise flight due to the separation of one of the 4th-stage turbine wheel blades due to low-cycle fatigue cracking that had progressed to high-cycle fatigue cracking, which ultimately led to the overload failure of the blade.

 

Source: NTSB Aviation Accident Database

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