WPR13GA374

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.

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.

The 4th stage turbine wheel and other related parts were submitted to the Rolls-Royce metallurgical laboratory and the NTSB Materials Laboratory for further examination. A Senior Materials Engineer examined the 4th stage turbine wheel along with the additional related parts. The engineer reported that the 4th stage turbine wheel was originally cast as a single article that contained 32 airfoils; however one of the airfoils was fractured just outboard of the rim of the wheel. The fractured airfoil was designated as airfoil #1, and the rest of the airfoils were incrementally numbered proceeding in a clockwise direction when viewing the wheel from the leading edge side. The airfoils smoothly transitioned into a continuous outer shroud that ran around the circumference of the wheel. The outer shroud was fractured adjacent to airfoils #29 and #2, and was missing across the intermediate airfoils #30–#32. The outer shroud fractures had rough appearances that were consistent with overstress fractures. Airfoil #1 was sectioned from the wheel, and first examined using an optical microscope. The fracture, as described below, initiated along the airfoil trailing edge and progressed by fatigue toward the leading edge. The length of the fatigue crack, measured from the trailing edge, was approximately 0.54 inch. Beyond the progression of the fracture, the fracture surface exhibited a comparatively rough appearance that was consistent with the remaining ligament failing in overstress. The fracture initiated approximately 0.025 inch above the rim where a fillet radius blended into the airfoil trailing edge. As the crack initially progressed toward the airfoil leading edge, it also progressed inward toward the rim. At a distance of 0.028 inch from the trailing edge, the crack began to progress outward away from the rim as it continued its progression toward the leading edge. The fracture surface of airfoil 1 was examined using a scanning electron microscope (SEM), and exhibited curved fatigue crack propagation features near the trailing edge. The curvature and orientation of the features were used to trace the crack path back to an initiation site along the airfoil trailing edge. The crack initiated along a straight segment of the airfoil trailing edge approximately 0.0045 inch from a curved segment on the pressure side of the airfoil. The initiation site exhibited features consistent with oxidation and mechanical rubbing/smearing. Examination of the fracture surface in the SEM revealed a boundary at which the fatigue fracture morphology changed as the crack progressed toward the airfoil leading edge. From the initiation site up to the boundary, the fracture surface exhibited the curved fatigue crack progression features described above. Beyond the boundary the curved crack progression, features diminished in appearance, and the fracture surface began to exhibit feathery features. The transition occurred as far as 0.038 inch from the trailing edge of the airfoil. The feathery features persisted as the fatigue crack progressed toward the leading edge until it reached the overstress region described above. The Rockwell hardness of the wheel was measured in accordance with ASTM E18 on the curvic teeth on the trailing edge side of the wheel, and measured 35.5 HRC, 34.0 HRC, 33.6 HRC, 34.6 HRC, and 36.4 HRC. The hardness values were in accordance with the material requirement. After the conclusion of the group exam, the wheel was sent to Rolls-Royce to inspect the other airfoil trailing edges for the presence of cracks. The wheel was cleaned with grit blast, and the remaining airfoil trailing edges were inspected using a fluorescent penetrant. No additional cracks were detected. The microstructure of the fractured airfoil was also examined by preparing a metallurgical specimen. The fractured airfoil was cross sectioned, ground, and polished in accordance with ASTM E3. The sample was then etched using a mixture of 33% acetic acid, 33% nitric acid, 33% water, and 1% hydrofluoric acid. The microstructure had an appearance of a typical cast microstructure, and there were no apparent microstructural anomalies at or near the trailing edge. For further details, see the Materials Laboratory Factual Report in the public docket for this accident. Representatives from Rolls-Royce engineering reported that after extensive analysis, the initiation of the low cycle fatigue (LCF) crack was most likely caused by a combination of two factors: The first being a high positive thermal gradient in the airfoil trailing edge material near the hub during the engine starting process. The second factor was during transient operation, such as auto-rotation, which produces a combination of high negative thermal gradient and high speed stress into the blade. There 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