ERA09LA317
ERA09LA317

HISTORY OF FLIGHT

On June 2, 2009, about 1110 eastern daylight time, a Hughes 369D, N8356F, registered to Aviation Advantage LLC, operated by Aerial Solutions, Inc., was landed hard during an autorotative landing to a clearing near Greenville, Virginia, following loss of engine power. Visual meteorological conditions prevailed at the time and no flight plan was filed for the local 14 Code of Federal Regulations (CFR) Part 133 external load flight from Greenville, Virginia. The helicopter sustained substantial damage and the commercial pilot, the sole occupant was seriously injured. The flight originated about 55 minutes earlier from Greenville, Virginia.

The purpose of the flight was aerial tree trimming near power lines, with approximately 1 hour duration. The pilot stated that a ground crew member gave him a 45 minute fuel check, and he then spent 5 minutes performing a “bottom pass.” He pulled up out of the right of way and while returning to a landing zone (LZ), he performed a power check. He also did an instrument scan while on final approach to the LZ; all instruments were in the green. He lowered the saw wheel to a marked spot at the LZ, and maneuvered the helicopter backwards and down to lower the saw engine to the ground. During set down stabilization, he heard a loud sound from the engine and the helicopter yawed to the left. The chief pilot of the operator estimated that the helicopter was between 90 and 100 feet above ground level (agl) when the engine failure occurred. The pilot further stated that he reduced collective, attempted to release the saw, and nosed the helicopter over to increase airspeed to achieve flare, but the helicopter impacted the ground. He released his seatbelt and shoulder harness and exited the helicopter.

DAMAGE TO AIRCRAFT

Damage to the helicopter consisted of a deformed tailboom, and a crease in the left side of the tailboom. The left skid was separated, and compression wrinkles were noted above and aft of the aft right skid tube. The pilot’s seat exhibited downward deformation, and extensive structural damage was noted in the area of the pilot’s seat. Bending damage to two of the main rotor blades was noted. One main rotor blade was fracture near the blade root, and one tail rotor blade was bent.

PERSONNEL INFORMATION

The pilot, age 41, holds a commercial pilot certificate with a rotorcraft helicopter rating, and was issued a second class medical certificate with no limitations on January 20, 2009.

The NTSB Pilot/Operator Aircraft Accident/Incident Report submitted by the pilot and operator indicated his total flight time in all aircraft was 8,676 hours, of which 7,780 were in the accident make and model. During the last 90 days, 30 days, and 24 hours he reported accruing 105, 32, and 7 hours respectively, all of which were in the accident make and model helicopter.

The pilot’s last flight review or equivalent was performed in the accident make and model helicopter on April 28, 2009.

AIRCRAFT INFORMATION

The helicopter was manufactured in 1976 by Hughes Helicopters as model 369D, and was designated serial number 1260063D. It was equipped with an engine torque exceedance instrument, and powered by a Rolls-Royce Corporation (formerly Allison Engine Company) 250-C20B engine rated at 420 shaft horsepower (SHP) with 375 SHP usable.

Review of the maintenance records revealed the engine installed at the time of the accident, serial number (S/N) CAE 823103, had been installed in various airframes since October 1999. The engine was installed into the accident helicopter on July 16, 2008, at helicopter total time of approximately 9,597 hours, and remained installed until removed for postaccident examination. The helicopter total time at the time of the accident was approximately 10,518 hours.

According to personnel from the engine manufacturer, a new Enhanced fourth stage power turbine wheel (P/N) 23055944 and S/N X555507 was manufactured, then sold in a kit on August 24, 2006.

The maintenance records associated with the turbine assembly revealed during an overhaul by a Canadian repair station in October 2006; the new Enhanced design fourth stage turbine wheel part number (P/N) 23055944 and S/N X555507 was installed. At the time of overhaul, the turbine assembly total time and cycles since new were approximately 7,077 hours, and 8,124, respectively. The turbine assembly was removed for a 1,750 hour inspection on May 30, 2008, and sent to a U.S. FAA certified repair station for the inspection; the fourth stage turbine at that time had accumulated 1,768.9 hours since new.

Records from the U.S. repair station indicate in part that the fourth stage turbine wheel remained installed at the completion of the 1,750 hour inspection. The turbine assembly was approved for return to service, installed on July 14, 2008, and the engine was installed in the helicopter on July 16, 2008. The turbine assembly remained installed until removed for the postaccident investigation. The engine had been operated approximately 2,690 hours since the new Enhanced design fourth stage turbine wheel was installed, and approximately 921 hours since the 1,750 hour inspection was last performed.

The engine was last inspected in accordance with a 100-Hour inspection on May 12, 2009. The engine total time at that time was recorded to be approximately 22,889 hours.

METEOROLOGICAL INFORMATION

A surface observation weather report taken at Shenandoah Valley Regional Airport (SHD), Staunton/Waynesboro/Harrisonburg, Virginia, at 1120, or approximately 10 minutes before the accident indicates the wind was from 190 degrees at 7 knots, the visibility was 10 miles, and clear skies existed. The temperature and dew point were 29 and 21 degrees Celsius respectively, and the altimeter setting was 30.08 inches of Mercury.

WRECKAGE AND IMPACT INFORMATION

Initial inspection of the engine following recovery of the helicopter was performed by a representative of the engine manufacturer with Federal Aviation Administration (FAA) oversight. The inspection revealed neither the N1 nor N2 drive trains could be manually rotated. No visible damage was noted to the compressor module. Both exhaust stacks exhibited small areas of peening on the inside surface. The exhaust collector was in position, but exhibited breaching in the path of the No. 4 power turbine wheel. The opening extended counter-clockwise from the 2 o’clock to the 6 o’clock positions. The outer rim and knife seals of the No. 4 power turbine wheel as viewed from the opening of the exhaust collector exhibited separation of a section of the No. 4 wheel rim spanning an area of approximately 3 airfoils. The middle airfoil was fractured near the hub. The engine was removed from the helicopter and shipped to the manufacturer’s facility for further examination.

Disassembly inspection of the engine at the manufacturer’s facility with FAA oversight confirmed one airfoil of the fourth stage turbine wheel was separated near the hub. Components of the engine consisting of the first, second, third, and fourth stage turbine wheels, second, third, and fourth stage turbine nozzle assemblies, power turbine outer shaft and outer shaft nut, power turbine inner race shaft and nut, No. 6 bearing, exhaust collector support, and power turbine support were submitted for detailed examination by the Safety Board’s Materials Laboratory located in Washington, DC. The power turbine governor and fuel control unit were retained for bench testing at the manufacturer’s facility.

The Safety Board’s Materials Laboratory examined the fourth stage turbine wheel with a representative of the engine manufacturer (materials engineer) present. The examination revealed impact damage to several airfoils, one airfoil was fractured at the hub root filet, and portion of the outer shroud was fractured from the wheel in the area of the fractured airfoil. Neither the fractured airfoil nor the shroud pieces were recovered. The fracture surface of the airfoil contained fused white powder deposits consistent with fire retardant. Additionally, overstress fracture was noted to the shroud piece or pieces fractured from the shroud and from the tips of three additional airfoils. Visual and fluorescent liquid penetrant inspection of the trailing edge of each intact airfoil from the hub diameter to about 0.5 inches towards each blade tip revealed no cracks.

Further examination of the airfoil of the fourth stage turbine wheel that was fractured near the hub root filet revealed fracture features consistent with fatigue initiation at the airfoil’s trailing edge near the pressure side. The fatigue terminus was located about 0.50 inch from the initiation region, which is about 0.04 inch from the outside diameter of the hub. No surface defects were noted along the root filet radius adjacent to the fracture initiation region; the filet radius adjacent to the fracture origin measured approximately 0.034 inch. Fracture features in the region between the fatigue terminus location and the leading edge of the airfoil were consistent with overstress fracture. The fractographic features in the fatigue origin region are consistent with high cycle fatigue; no anomalies or injurious defects were noted at the origin. The chemical composition of the 4th stage power turbine wheel casting and the hardness of the hub near the fractured airfoil were within manufacturer’s specification. No fabrication anomalies such as porosity or casting defects were noted in the area of the fractured airfoil. A concentration of subsurface carbides along the fillet radius adjacent to the fractured turbine blade was noted. According to a representative of the engine manufacturer, such microstructural attributes are typical of cast precipitation-hardenable nickel-based alloy castings and are believed to be due to either mold chill effects or a metallurgical interaction with the mold coating.

The results of the examination of the components other than the fourth stage turbine wheel revealed secondary damage due to airfoil fracture of the fourth stage power turbine wheel. A detailed report with accompanying pictures is contained in the public docket for this case.

TESTS AND RESEARCH

The power turbine governor and fuel control unit installed at the time of the accident, and also the power turbine governor and fuel control unit removed from the engine the day before the accident were bench tested at the manufacturer’s facility with FAA oversight. The bench testing of all components revealed no discrepancies that would have resulted in an engine malfunction. A detailed report with accompanying pictures is contained in the public docket for this case.

Revision 3 of Rolls-Royce Commercial Engine Bulletin CEB-A-1400, dated January 19, 2009, (applicable to the accident engine), indicates to avoid steady state engine operation in the N2 speed avoidance range of 75 to 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 bulletin requires an entry in the maintenance records documenting steady state operation in the speed avoidance range when operating above 85 shp.

According to the company director of maintenance (DOM), following release of CEB-A-1400 in December 2006, company flightcrew members and mechanics were briefed about the details of the bulletin. The pilots were advised that this range could be encountered during a slow acceleration from ground idle to flight rpm, and the mechanics were advised that the range could be encountered when balancing the main rotor hub without the main rotor blades, or during vibration surveys of the engine. The pilot and mechanics were advised to report any steady state operation in the speed avoidance range. To his knowledge, there was no reported steady state operation of the engine in the N2 speed avoidance range. Additionally, the DOM stated that he later created a graphic depiction of the speed avoidance range which he used during training classes and general meetings to, “…underline the importance of the service bulletin.”

Review of the maintenance records from January 7, 2006, to June 1, 2009, revealed no entry indicating any steady state engine operation in the N2 speed avoidance range.

The Safety Board has investigated one other accident in which one airfoil of an Enhanced fourth stage turbine wheel fractured near the hub due to a fatigue crack. A resulting loss of engine power occurred and the helicopter sustained substantial damage during an autorotative landing. The accident involved a McDonnell Douglas 369E helicopter, N556CP, operated by a public use agency, and occurred on October 29, 2007. More information regarding this accident, National Transportation Safety Board case number CHI08GA028, is available online at http://www.ntsb.gov/aviationquery/index.aspx

Personnel from Rolls-Royce Corporation reported that since 2005, excluding both NTSB investigations, three other fatigue failures of an airfoil of an Enhanced fourth stage turbine wheel were reported. The investigation of the three fractures of the airfoil revealed all initiated at the trailing edge near the hub. Including the two NTSB investigations and the three additional cases, four of the five involve a variant of the McDonnell Douglas 369 helicopter, while the final case involves a variant of a Bell helicopter.

Rolls-Royce Corporation released revision 1 of Rolls-Royce Commercial Engine Bulletin CEB-A-1407, dated February 7, 2011, pertaining to 250-C20B engines installed with an Enhanced Power Turbine fourth stage power turbine wheel P/N 23055944 (applicable to the accident engine and helicopter), revealed analysis of recent failure in part 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 bulletin called for a one-time inspection of the third and fourth stage turbine wheels for cracks in the airfoils.

According to a letter from Rolls-Royce to NTSB dated August 12, 2010, their engineering investigative findings related to the airfoil failures of the fourth stage turbine wheel due to fatigue revealed the possibility of a higher stress state in the airfoil at the trailing edge root than originally modeled. The stress, which occurs during engine start results from thermal differentials in the airfoil geometry that produces a residual stress at the trailing edge root which can lead to a fatigue crack. Once a crack occurs, it can propagate in high cycle fatigue followed by overload failure if the engine is operated steady state in the speed avoidance range specified in CEB-A-1400. The letter from Rolls-Royce also indicates that they are currently refining their analysis in order to design a more robust fourth stage turbine wheel that does not exhibit the fatigue cracking at the trailing edge root.

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