N5225C

HISTORY OF FLIGHT On July 22, 2014, about 1120 Pacific daylight time, an MDHI 369D, N5225C, collided with terrain and rolled downhill several times near Oso, Washington. Olympic Air was operating the helicopter under the provisions of Title 14 Code of Federal Regulations (CFR) Part 133. The commercial pilot sustained serious injuries, and the helicopter airframe sustained substantial damage during the accident sequence. The local external load flight departed at an undetermined time. Visual (VMC) meteorological conditions prevailed, and no flight plan had been filed.

A Federal Aviation Administration (FAA) inspector examined the wreckage on site. He noted that the main rotor blades fragmented into many pieces that were scattered along the hillside. He identified four of the five blades. During recovery, personnel located the fifth blade about 900 feet away, and with much less damage.

TESTS AND RESEARCH

Follow Up Examination July 28, 2014

The National Transportation Safety Board (NTSB) investigator-in-charge (IIC) and investigators from Boeing and Rolls-Royce examined the wreckage in Auburn, Washington, on July 28, 2014. A complete report is part of the public docket for this accident.

There was no evidence of preimpact damage to the airframe. There were no anomalies with the airframe or engine that would have precluded normal operation.

Four of the five main rotor blades exhibited substantial damage. They were bent, buckled, had spar damage, leading edge gouges and dents, and trailing edge separation. Some blades were warped and fragmented.

The fifth blade, serial number (SN) 091B, had substantially less damage than the other four blades. The blade subassembly was liberated from the root assembly; it fractured with bonding separation at the blade root, and the blade airfoil was relatively intact.

All five main rotor blades were manufactured by Helicopter Technology Company (HTC), LLC, under Parts Manufacturing Authority (PMA) issued by the FAA.

Follow Up Examination

The NTSB IIC and another NTSB investigator examined the wreckage on September 23, 2014. They retrieved the blade roots for all five main rotor blades, and submitted them to the NTSB Material's Laboratory for examination.

National Transportation Safety Board (NTSB) Material's Laboratory Blade Examination

There were several components that comprised the main rotor blade assembly. Skin doublers were bonded to the outside of the blade at the root end. Upper and lower root fittings were joined to the root end of the blade subassembly with a blue-colored film adhesive on a scrim cloth carrier and five attachment bolts. Three of the attachment bolts were arranged spanwise along the fitting, and two bolts (inboard of the aforementioned three spanwise bolts) at the root end were arranged chordwise relative to one another. For the purposes of this investigation, the bolt located at the outboard end of the root fitting was labeled as the "No. 1 attachment bolt," and the bolt adjacent to it was labeled as the "No. 2 attachment bolt."

According to the manufacturer of the main rotor blade assemblies, the blade and fittings were joined as follows:

1. A primer containing strontium was applied to each fitting and the doubler

2. The adhesive film was sandwiched between the fittings and the blade and the adhesive was cured

3. The excess adhesive was trimmed from the edge of the fitting

4. A sealant was applied around the edge of the fitting

5. Paint coatings were then applied

Examination of Separated Rotor Blade

Visual examination prior to disassembly revealed that rotor blade SN 091B was fractured primarily chordwise through the second bolt hole from the outboard end of the root fitting, about 3 inches from the inboard end of the blade subassembly. On the inboard end of the separated blade subassembly, blue-colored adhesive film remained bonded to the upper and lower surfaces in the area that mated to the inner surfaces of the upper and lower root fittings. There were cracks in the white paint along the interface between the retained root end of the blade subassembly and both upper and lower root fittings. The No. 1 attachment bolt was missing, and was not recovered. The No. 1 attachment bolt hole in the upper fitting had been enlarged to a size and shape similar to the bolt head of the other attachment bolts. The No. 1 attachment bolt hole in the lower fitting had also been enlarged into an elongated slot with a width similar to the minor thread diameter of the other attachment bolts; the slot was elongated in a primarily outboard direction.

The remaining four attachment bolts were found in their installed configuration, and removed from the rotor blade root assembly, upon which the upper and lower root fittings separated from the blade subassembly. Blue-colored adhesive remained bonded to the blade subassembly; no evidence of blue-colored adhesive was observed on the inner surfaces of the upper and lower root fittings, which was consistent with disbonding of both root fittings from the blade subassembly. Wear debris was observed on the bonding faces of the upper and lower root fittings around the No. 1 and No. 2 attachment bolt holes. The adhesive film on the liberated portion of the blade subassembly exhibited a rubbed appearance around the first and second bolt holes as well.

The fracture through the blade subassembly at the No. 2 attachment bolt hole was further examined. Flat, comparatively smooth fracture faces, and periodic crack arrest marks consistent with fatigue were observed on the upper and lower spar arms, the upper and lower blade skin, and the upper and lower doublers. Fatigue initiation sites were observed at three distinct locations: 1) at both forward and aft edges of the upper side of the No. 2 attachment bolt hole, 2) at both forward and aft edges of the lower side of the No. 2 attachment bolt hole, and 3) at the forward surface of the spar (adjacent to the brass leading edge weight).

The fatigue initiation sites were examined for any notable features indicative of pre-existing damage or defects, but none were found.

The disbondment between the adhesive and one of the fittings (the lower fitting) was examined in greater detail on a chordwise cross section of the blade subassembly. Sealant was observed between the adhesive layer and the fitting up to approximately 0.14 inch in from the edge of the fitting. The debris previously observed on the plan view image appeared as patches of debris on top of and adjacent to the sealant.

The primer along the disbonded interface was examined. White particles were observed within a thin discontinuous layer that was bonded to the adhesive layer. Examination of the particles revealed the presence of strontium, a constituent of the primer that was applied to the fitting prior to the bonding process.

The primer was similarly characterized at the trailing edge of the fitting. As above, the primer was bonded to the adhesive layer. The sealant was observed on top of the primer, and a thin discontinuous layer of debris was observed on top of the sealant. Examination of the debris indicated that, among additional elements, the debris contained titanium, consistent with wear debris from the fitting.

Intact Rotor Blades

The four remaining intact rotor blade assemblies were examined visually for cracks in the paint around the perimeter of the upper and lower fittings. Cracks were observed around the perimeter of the upper and lower fittings on blade SN 085B. No paint cracks were observed on the other three blades.

The root fitting assembly attachment bolts on all of the blade assemblies were removed. On blade SN 085B, removal of the attachment bolts revealed disbonding of the upper root fitting from the blade subassembly, similar to the disbonding observed on blade SN 091B. The lower root fitting for blade SN 085B did not separate from the blade subassembly. Both upper and lower root fittings for the remaining three blades did not separate after removal of the attachment bolts. Examination of blade SN 085B using a stereomicroscope did not reveal evidence of cracks emanating from the root fitting attachment bolt holes.

For further details on the NTSB Materials Laboratory Blade Examination, see the Materials Laboratory Factual Report No. 15-026 in the docket for this investigation.

ADDITIONAL INFORMATION

FAA Air Worthiness Directive (AD) History

The FAA issued AD 96-10-09; it required initial and repetitive inspections at intervals not to exceed 100 hours of time in service of each main rotor blade root for either cracks, paint and sealant cracking, or separation between the lower surface root end fitting and the doubler. The actions specified in this AD were intended to prevent failure of a blade resulting in separation of the blade and subsequent loss of control of the helicopter.

FAA AD 2005-21-02 instructed the operator to determine and record the number of torque events (TE) accumulated on each blade. On or before accumulating an additional 200 TEs or at the end of each day's operations, whichever occurred first, the operator was required to record and update the accumulated TEs total. For each blade that had accumulated 13,720 or more TEs and 750 or more hours TIS, before further flight, unless accomplished previously, the operator was to perform a main rotor blade TE inspection. The AD also required a recurrent main rotor blade TE inspection at intervals not to exceed 200 TEs or 35 hours TIS, whichever occurred first.

A review of maintenance records indicated that the blades on the accident helicopter had accumulated about 1,123 hours and 232,674 TEs since installation.

MD Helicopters Service Bulletin

MD Helicopters service bulletin (SB), SB369E-095R2, described the purported causes of disbondment, the inspection procedure to detect disbondment, and instructions for determining the inspection interval. The cause of the disbondment was purported ...