On June 6, 1998, at 0025 hours Pacific daylight time, an Aerospatiale AS350BA, N31621, made a forced landing to a rooftop parking structure 5 miles southwest of the Burbank, California, airport. The forced landing was precipitated by the onset of what the pilot described as "violent vibrations" through the airframe and controls. The aircraft sustained substantial damage and the airline transport pilot, the sole occupant, was not injured. The helicopter was being operated by Jetcopters, Inc., Van Nuys, California, under 14 CFR Part 91, as a positioning flight. The flight originated at Lake Arrowhead, California, approximately 2400, and was en route to the Van Nuys airport at the time of the accident. Night visual meteorological conditions prevailed and no flight plan was filed.

The pilot reported that he was transitioning through the Burbank Class C airspace approximately 2,000 feet agl and 120 knots indicated airspeed. He felt a sudden "violent vibration and oscillation" and the helicopter began to pitch and yaw. The pilot stated that he felt the vibration through all the controls. About 5 seconds after the vibration started, the tail rotor gearbox chip light illuminated. As the pilot slowed the helicopter and began to descend, the aircraft stopped pitching and rolling but the vibration continued. He reported that he made a normal approach to the rooftop parking structure, which was well lit. The pilot landed and fully lowered the collective. As he reduced the throttle, he heard a loud bang and the aircraft hopped a few feet forward. He performed a 2-minute shutdown of the engine and exited the aircraft.


The helicopter had a total airframe time of 1,014 hours and had accumulated approximately 65 hours since being rebuilt. The maintenance records indicate that both tail rotor pitch change links were original equipment. Review of the maintenance records and daily operations records did not reveal any anomalies or unresolved discrepancies.

The pilot reported that he completed a daily inspection of the aircraft on June 5, 1998, and no play was observed in either tail rotor pitch link, nor did he see any cracks. He also reported that he performed a preflight before takeoff on the accident flight and did not note any discrepancies at that time.

In the month that the operator had possession of the helicopter, the pitch links were inspected by company maintenance personnel who reported verbally and in the daily inspection logbooks that no play was noted in the pitch links.


Approximately the last 3 feet of the tail boom, with the vertical fins attached, had separated just aft of the horizontal stabilizer. The separation was 360 degrees in circumference, and the rear section of the tail boom was found attached to the rest of the aircraft solely by a wire bundle and push-pull tube. The tail rotor assembly remained connected to the gearbox, which was still mounted to the tail boom.

One of the two tail rotor blades exhibited a fracture 360 degrees around the shell at the base in an area located approximately 1-inch outboard of the horn counterweight. Visual examination through the crack of the shell revealed that the longitudinal spar was intact. A fragment of the shell had separated and was missing from the outboard end on the trailing edge of the blade. Chordwise scrape marks and marks consistent with impact damage were found on the outboard end of the first tail rotor blade, extending from the outboard end to about 6 inches inboard of the outboard end. The second tail rotor blade displayed minor scrape marks at the outboard leading edge, and a mark about 0.8 square inches near the base.

Both tail rotor pitch change links were found separated at two locations in the link body bearing loop end that retains the bearing assembly. The outer portion of the loop for each pitch change link remained attached to a bearing assembly.

No external damage was noted to any of the three main rotor blades.

No external damage was noted to the engine. The tail rotor drive shaft cover showed rotational scars at the forward tail rotor drive shaft coupling.

Since a tail rotor gearbox chip light was reported a teardown of the tail rotor gearbox was performed. Metal shavings and ferrous and non-ferrous materials were found in the gearbox.


Both pitch change links, the tail rotor blades, and parts of the tail boom were shipped to the Safety Board's Materials Laboratory in Washington, D.C., for further examination and evaluation. The metallurgist's factual report is appended to this file.

Examination of pitch change link "1" revealed fatigue cracking emanating from three adjoining corrosion pits. The corrosion pit measured as deep as 0.004 inches. There was a fracture that contained fatigue cracks that emanated from multiple origins at the inside diameter surface of the loop that is in contact with the bearing assembly. The areas outside the fatigue crack region exhibited features typical of overstress separations. Approximately 180 degrees of the scrim cloth was missing from the raceway of this bearing assembly.

Examination of the inboard and outboard faces of the bearing components revealed severe rubbing damage completely around the inboard face of the outer race. The bearing outer raceway that normally mates with the spherical ball also contained severe wear damage approximately 180 degrees around its circumference, with most of the wear damage located on the outboard side of the raceway. This raceway when fabricated contains a layer of scrim cloth on the raceway surface. The scrim cloth for this bearing assembly was missing. The ball for this assembly exhibited a dark discoloration for the most part all around the hemisphere side next to the head of the bolt. In one area of the ball, discoloration extended almost all around the ball surface. No wear step was noted on the surface of the ball. A large gap existed between the ball and the raceway.

The Eurocopters Maintenance Manual outlines several items to be checked on a daily basis by either a pilot or a mechanic. One of the items is the pitch change links. The manual directs them to be inspected visually and by touch to check their condition and to check for end play. It notes that if play is observed, the pitch link should be measured with a dial gauge (micrometer). According to the manual, a maximum of 0.008 inches of play is allowed in the radial direction, but no limit is specified for the axial direction.

The flat ends of the ball of pitch change link "2" were placed between a vice and a dial micrometer was placed on the surface of the outer race assembly. The micrometer indicated play between 0.003 and 0.004 inches in the radial direction, which was less than the maximum radial play limit of 0.008 inches. When the outer race assembly was pushed and pulled in the axial orientation, the dial indicated between 0.008 and 0.009 inches of play.

The metallurgist indicated that the diameters of both bearing balls appeared to correspond to the specified diameter of 0.7795 inches. He further stated that microhardness testing of a section of pitch change link "1" produced an average hardness greater than the specified minimum hardness for that particular aluminum alloy.

According to the metallurgist, the stereo microscope examination of the fractures from pitch change link "2" and the skin fragments from the tail rotor blades and tail boom exhibited features typical of overstress separation.

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