ERA11LA150

The pilot was in cruise flight at 1,500 feet over a heavily wooded area when the airplane began to vibrate violently before sustaining a total loss of engine power. The pilot made a forced landing and the airplane collided with trees and the ground. The pilot stated that, about 30 flight hours before this accident, the airplane experienced a similar severe vibration in flight. He shut down the engine and made a successful forced landing to a nearby airport. Examination of the airplane revealed that a triangle-shaped aluminum plate on the rear of the propeller hub had cracked. The pilot installed a new propeller hub and the original propeller blades about 7 months before the accident and began flying again without any vibrations. The total hours on the propeller blades are unknown. A postaccident examination at the accident site revealed that both carburetors had separated from the engine assembly and were hanging by their fuel lines and throttle cables. One composite propeller blade was delaminated with a large section missing. The remaining two propeller blades appeared to be intact without comparable or crash-related damage. The carburetor detachment likely caused the final engine failure and this separation was likely a consequence of the severe vibration. Metallurgical examination of the propeller blades identified that separation of a portion of blade No. 1 initiated as a transverse fracture, originating near the trailing edge of the blade. The transverse fracture was progressive in nature, with the crack propagating in some combination of fatigue and/or stress rupture under continually applied loads. Blade No. 2 also exhibited a pattern of parallel transverse cracks on its aft side. Blade No. 3 was also visibly damaged, with significant delamination or disbonding of the woven carbon-fiber composite skin from the forward side of the blade, along with delamination at the trailing edge. Blade No. 3 also exhibited a pattern of parallel transverse cracks on its aft side. This damage was consistent with bending of the tip of the blade forward under air loading.

On February 17, 2011, at 1530 eastern standard time, an amateur-built experimental Flaglor Kolb MK III airplane, N213, registered to and operated by a private owner, experienced a violent vibration in cruise flight at 1,500 feet above ground level, followed by a total loss of engine power. The pilot performed a forced landing to trees and the ground in the vicinity of Stetson, Maine. The personal flight was operating under the provisions of Title 14 Code of Federal Regulations Part 91, and no flight plan was filed. Visual meteorological conditions prevailed, and the airplane sustained structural damage to the right wing and forward fuselage. The sport pilot reported minor injuries. The flight originated from Old Town, Maine, at 1330, and was en-route to Dexter, Maine. The pilot stated that about 30 flight hours before the accident, the airplane experienced a severe vibration in flight. He shut the engine down and made a successful forced landing to a nearby airport. Examination of the airplane revealed a triangle shaped aluminum plate on the rear of the propeller hub had cracked. The front and rear propeller hubs were replaced by Warp propellers. He installed the propeller hub and same propeller blades on July 10, 2010, and began flying again without any vibrations. The total hours on the propeller blades are unknown. Examination of the crash site revealed the airplane impacted trees and came to rest upright in 3 to 4 feet of snow, on a heading of 360 degrees magnetic. The forward nose section of the cabin area was bent upward and to the right. The forward windscreen, instrument panel, and cabin roof were partially separated. The cabin seats remained intact and the pilot's seat belts and shoulder harness were in use at the time of the accident. The fuel tank was not ruptured and was more than half full. Continuity of the flight control system was confirmed from the cabin area aft to all flight control surfaces. The right main landing gear remained attached to the fuselage and was bent rearward. The right wing remained attached to the wing root. The leading edge of the right wing was damaged, extending about 2 feet outboard of the wing root. The remainder of the leading edge sustained minor damage. The wing strut remained attached to the wing and the fuselage. The right flap and aileron remained attached at all attachment points. The left wing remained attached to the wing root and was not damaged. The wing strut remained attached to the wing and the fuselage. The right flap and aileron remained attached at all attachment points. The tail boom was bent upward. The vertical fin, rudder assembly, horizontal stabilizers and both elevators were not damaged. The engine assembly remained attached to the engine mounts aft of the cabin area. Both carburetors had separated from the engine assembly and were hanging by their fuel lines and throttle cables. One composite propeller blade was de laminated and the remaining two propeller blades were notdamaged. Examination of the remaining engine assembly and accessories revealed no anomalies. The propeller blades labeled blade Nos. 1, 2, and 3, and both propeller hubs were sent to the NTSB Materials Laboratory for further analysis. The fracture and separation of a portion of the after body of blade No. 1 initiated as a transverse fracture, originating near the trailing edge of the blade. The transverse fracture was progressive in nature, with the crack propagating in some combination of fatigue and/or stress rupture under continually applied loads. Blade No. 2 also exhibited a pattern of parallel transverse cracks on its aft side, but with no evidence of delamination. Blade No. 3 was also visibly damaged, with a significant delamination or disbonding of the woven carbon-fiber composite skin from the forward side of the blade, along with a delamination at the trailing edge. Blade No. 3 also exhibited a pattern of parallel transverse cracks on its aft side. This damage was consistent with bending of the tip of the blade to forward under air loading. The hub that was replaced in July 2010 was fractured as a result of a fatigue crack that extended 5.7 inches, with over stress regions at either end that were 0.09 inch and 0.23 inch long. The fatigue crack originated from a stress concentration caused by the sharp indentation of the edge of a faceplate bolted to the aft outer surface of the hub. The Rotax model "E" gear box was forwarded to an authorized repair station and placed in a bonded area until an examination could be conducted by the manufacturer under the supervision of a Federal Aviation Administration inspector. Examination of the "E" gear box was conducted and revealed there were no internal worn parts or signs of vibration producing damage.