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On June 30, 2000, about 0705 central daylight time, a Weatherly 201C, N9228W, piloted by a commercial pilot, was destroyed by impact with terrain while maneuvering near Tolna, North Dakota. The right wing was found detached. A fire occurred in the fuselage and left wing. The aerial application flight was operating under 14 CFR Part 137. Visual meteorological conditions prevailed at the time of the accident. No flight plan was on file. The pilot sustained fatal injuries. The local flight originated from Breckheimer Airport near Tolna, North Dakota, at time unknown.
A witness stated that at 0705, he heard an airplane sputter and then went dead. Approximately five minutes later he saw smoke. He met another witness who went to a neighbor's farm to call 911. He proceeded to the crash site to look for the pilot. The wreckage was observed to be on fire. The witness and the neighbor came to the crash site and all three looked in the field for the pilot.
The pilot held a commercial pilot certificate. He held a second class medical certificate with limitations for corrective lenses. The operator's report had 10,300 hours listed as his total flight time.
The airplane was a Weatherly 201C, N9228W, serial number 1036. Its restricted category airworthiness certificate was issued May 6, 1997. Its annual inspection was completed on May 1, 2000. The operator's reports listed 343 hours since the airplane's last annual inspection and "5,400+" hours as its total time.
At 0635, the Devils Lake Municipal Airport, near Devils Lake, North Dakota, weather observation was: Wind 200 degrees at 8 knots; visibility 10 statute miles; sky condition clear; temperature 17 degrees C; dew point 14 degrees C; altimeter 29.83 inches of mercury.
WRECKAGE AND IMPACT INFORMATION
The North Dakota Highway Patrol report stated, "The wing came off the plane and the plane then crashed approx. 396' west of the township road in a field of flax. The plane came to rest and was burning approx 471' west of the township road." The right wing came to rest on the township road. The pilot was found fatally injured within the main wreckage. (See appended report and photographs.)
MEDICAL AND PATHOLOGICAL INFORMATION
An autopsy was performed on the pilot by the North Dakota Department of Health, State Medical Examiner on July 1, 2000.
The Federal Aviation Administration Civil Aeromedical Institute prepared a Final Forensic Toxicology Accident Report. The report indicated 0.353 (ug/ml, ug/g) morphine detected in urine and 46.404 (ug/ml, ug/g) salicylate detected in urine.
A post impact ground fire occurred.
TESTS AND RESEARCH
Sections of the separated right wing and its mating surface were cut away and shipped to the National Transportation Safety Board's Materials Laboratory for detailed examination. An excerpt of the materials laboratory factual report stated the following:
C. DETAILS OF THE EXAMINATION Three pieces of the forward wing spar labeled "1" to "3" in figure 1 were submitted for examination. The two inboard pieces, "1" and "2", were darker in appearance, consistent with exposure to fire. The web between the upper spar cap in piece "1" and the lower spar cap in piece "2" was deformed roughly into an "S" shape and was fractured. Wing attachment fittings, indicated by arrows "A" in figure 1, were attached to the upper and lower spar caps. Pieces 1 and 2 had black soot deposits typical of exposure to a fire. The outboard piece of the wing spar, piece "3", was green in color, with no evidence of exposure to a fire. The fracture surface on piece "3" mated with the outboard fracture surface on piece "2". The outboard end of piece "3" was cut, apparently to facilitate shipping to the materials laboratory.
A view of the fracture surface on piece "3" is shown in figure 2. The fracture surface intersected the outboard hole for the lower wing attachment fitting ("A" on piece 2 in figure 1) and several rivet holes in the web. Most of the fracture through the lower spar cap and the lower portion of the web, shown at higher magnification in figure 3, was on a flat transverse plane and contained crack arrest lines, features typical of fatigue cracking. The crack arrest lines emanated from the outboard hole for the lower wing attachment fitting and extended to the locations indicated by dashed lines in figure 3. For the lower spar cap, the fatigue cracking extended from the bolt hole (1) in the downward / forward direction to the dashed line position near the forward end of the lower leg of the cap and (2) in the upward direction completely through the aft leg of the cap. For the web, the fatigue cracking extended from the bolt hole (1) downward to the lower edge of the web and (2) upward to the next rivet hole (the position indicated by the upper dashed line in figure 3).
The outboard hole for the wing attachment fitting and adjacent portions of the fracture are shown at higher magnification in figure 4. Corrosion was observed in the hole and on portions of the spar cap fracture surface. Circumferential grooves consistent with machining marks were also observed on the hole surface both in the cap and web. The upper and lower sides of the spar cap hole and adjacent fracture surfaces are shown at higher magnification using scanning electron microscopy (SEM) in figures 5 and 6. Arrows "O" indicate fatigue origins initiating from corrosion pits on the hole surface.
Fatigue striations were observed on the fracture surface of the fatigue region, as shown in figure 7. The number of striations were counted on the lower spar cap fracture surface at locations along the crack path between the origin at the lower side of the outboard wing attachment fitting hole and the lower / forward fatigue boundary. The resulting data is shown in figure 8. A curve of the form dn/da = C / a 0.5 was fit to the data as shown in figure 7, where dn/da are the number of striations per inch, C is a constant, and a is the crack length from the origin. By integrating under the curve, the total number of striations was estimated for a given change in crack length. For a crack extending from 0.063 inches to failure, the estimated number of striations was 67,208 [+ or -] 13,442.
Reportedly, the accident aircraft was used in an agricultural spraying application. Typically during a spraying pass, a loading cycle is placed on the wings twice per pass (once to level out at the start of the pass, and once to lift up at the end of the pass). For the purposes of estimating the number of loading cycles per hour of operation, it was assumed that the typical number of passes during a mission for a Weatherly 201C is 10, and that the average mission time is 0.34 hours. Adding a loading cycle for landing and takeoff, the loading rate is 60 cycles per hour for the typical Weatherly 201C used in an agricultural spraying application. This data was combined with the striation data from the accident aircraft, where one striation correlates to one cycle, to estimate the time to failure. The time to failure from a 0.063-inch crack was estimated to be 1120 [+ or -] 224 hours.
Hardness and conductivity of the lower spar cap were measured. The hardness was 80.8 HRB and the conductivity was 29.0 percent IACS. According to specifications for the spar cap, the material is aluminum 2024-T3511. The typical values of hardness and conductivity for aluminum 2024-T351 are 74.5 HRB and 29.7 percent IACS.
The remaining fracture surfaces (outside of the previously described fatigue regions) formed slant angles in the thinner members and had rough irregular surfaces in the thicker members, consistent with overstress fracture. Slight upward deformation was observed on the upper spar cap on piece "1" near the outboard fracture surface and at the upper end of piece "3", indicating that the outer portion of the wing separated in an upward motion relative to the inboard portion.
Subsequent to the accident the manufacturer issued a mandatory service bulletin. Bulletin excerpts stated the following:
Required Action Remove the right and left outer wings from the aircraft and remove the lower hinge strap on the front spar. Inspect the spar cap for cracks at the bolts hole penetrations giving special attention to the farthest outboard hole where the highest stress is experienced. Visual inspection with a strong light source and an inspection mirror should be adequate; but if a dye penetrant inspection by qualified personnel is available, it may be more desirable and prudent to use this method. If cracks are found the spar cap must be replaced prior to flight.
Frequency of Required Inspection The above described inspection should be accomplished before the next flight. Subsequent inspections must be accomplished at intervals no greater than 500 hours of flight. (See appended report.)
A party to the investigation was the Federal Aviation Administration.