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On December 18, 1999, at 1617 central standard time, a Globe GC-1B (Swift) airplane, N80951, registered to the Aircraft Components MFG, Inc. of Sulphur Springs, Texas, and operated by the pilot, was substantially damaged when it impacted terrain following a loss of engine power near Edgewood, Texas. The commercial pilot, sole occupant of the airplane, was fatally injured. Visual meteorological conditions prevailed, and a flight plan was not filed for the 14 Code of Federal Regulations Part 91 personal flight. The flight originated from Sulphur Springs, Texas, about 1605, with a destination of the Thompson Field Airport near Canton, Texas.
A pilot reported that while flying 43 miles north of Terrell, he heard a "May Day" call from the pilot of a Swift airplane. The pilot indicated "an engine problem or failure, 4 miles north of Canton going down."
Witnesses reported observing the airplane maneuvering at a low altitude with the propeller not turning. As the airplane made a turn, its nose dropped, and the airplane descended and impacted the ground.
According to FAA records, the pilot was issued a commercial pilot certificate on December 19, 1973, for rotorcraft-helicopter with a helicopter instrument rating. On September 14, 1989, he obtained a private pilot certificate with an airplane single-engine land rating. The pilot held a third class medical certificate, which was issued September 10, 1998. The medical certificate stipulated a limitation to have corrective lenses available for near vision while operating an aircraft.
A review of the pilot's flight logbook revealed that on May 1, 1999, he completed a biennial flight review in a PZL-104 (Wilga) airplane. On July 23, 1999, the pilot received a demonstration flight in the accident aircraft, and on September 19, 1999, he received 3.8 hours flight training in the airplane. The logbook also revealed that as of December 17, 1999, the pilot had logged a total flight time of 2,075.7 hours, of which 1,066 hours were in helicopters and 38.0 hours were logged in the accident airplane.
The 1946-model Globe GC-1B (Swift) was a low wing, single-engine, two-place airplane, which had retractable main landing gear and a fixed tail wheel. It was powered by a WSK "PZL-Rzeszow" Franklin 6A-350-C1R engine rated at 220-horsepower, and a McCauley, two-bladed, constant speed-controllable pitch propeller.
According to an employee of the pilot's company, the pilot had purchased the airplane about three months prior to the accident.
A review of the aircraft's maintenance records revealed that on February 1, 1984, a Franklin 6A-350-C1 engine and McCauley 2A31C21/S84S-6 propeller were installed in accordance with STC SA203NW. The "PZL-Rzeszow" Franklin 6A-350-C1R engine, serial number 911171022, with 1.5 hours since factory new (SFN), was installed in the airplane on August 1, 1998, at a tachometer time of 241.0 hours. The airplane underwent its last annual inspection on September 16, 1999, at a tachometer time of 263.7 hours and a total aircraft time of 2,528.4 hours. At the time of the accident, the engine had accumulated 61.04 hours SFN, and 59.54 hours since it was installed in the airplane. According to a work order, dated November 17, 1999, a magneto check was "strongly recommended" due an "excessive rpm drop." The last maintenance performed on the engine, a replacement of the front oil seal, was done on November 20, 1999.
WRECKAGE IMPACT INFORMATION
The aircraft wreckage was located in a field at latitude 32 degrees 37.818 minutes north and longitude 095 degrees 50.817 minutes west. Examination of the accident site revealed that the airplane impacted the ground on a measured magnetic heading of 230 degrees. The airplane came to rest upright 24 feet from the initial ground scar on a measured magnetic heading of 195 degrees. The empennage was attached to the fuselage but facing 240 degrees. The nose of the airplane was crushed aft, and the firewall was twisted to the right. The cabin roof/canopy was separated from the fuselage. The left wing leading edge was crushed and twisted upward from the wing tip to the landing gear. The flap was separated and laying under the wing. The aileron was partially separated and the landing gear was found in the down position. The right wing leading edge was crushed aft, and the flap was bent under the wing. The landing gear was found in the down position. Flight control continuity was confirmed to all flight control surfaces.
The engine was displaced downward and partially separated from the fuselage,. The engine sustained impact damage to the accessory case. The crankshaft was rotated by hand. There was continuity to the accessory gears, valve action on all cylinders, and thumb compression on all cylinders. The starter, alternator, and both magnetos were separated from the engine. Both magnetos sparked when rotated by hand. The engine driven fuel pump was separated from the engine. The engine driven fuel pump was disassembled, and both the inlet and outlet valve assemblies were found separated into pieces. The fuel pump was sent to the NTSB Materials Laboratory in Washington, D.C., for further examination. Fuel was found in the electric fuel boost pump. The fuel boost pump was tested and it operated. The fuel boost pump switch was found in the off position; however, the switch appeared to be broken.
The propeller was separated from the crankshaft at the propeller flange. The propeller was found imbedded in the ground at the initial ground scar. One blade was bent aft with span wise scratches. The other blade had a forward bend with some chord wise scratching near the hub.
MEDICAL AND PATHOLOGICAL INFORMATION
The Southwestern Institute Of Forensic Sciences in Dallas, Texas, conducted an autopsy of the pilot. Toxicological testing was performed by the FAA Civil Aeromedical Institute's (CAMI) Forensic Toxicology and Accident Research Center at Oklahoma City, Oklahoma. The toxicological tests were negative for alcohol and drugs.
TESTS AND RESEARCH
Examination of the engine driven fuel pump at the NTSB Materials Laboratory revealed that the upper surfaces of both the inlet and outlet valve seats were worn, appearing smooth and shiny with deformation.
The edges of the inlet valve stem were rounded and worn. The riveted (lower) end of the inlet valve stem shaft was missing from the remainder of the valve stem, and any possible fracture surfaces were completely obliterated by mechanical deformation and wear. "Relatively extensive" porosity was observed in a cross section of the inlet valve stem. Cracks were observed at the corners of the groove for the valve spring. A portion of the fracture surface of a corner crack was exposed in the laboratory. The crack fracture surface was relatively smooth, and at high magnification using scanning electron microscopy (SEM), striations were observed on the fracture surface, features typical of fatigue. By contrast, the remainder of the fracture surface that was separated in the laboratory had a ductile dimple morphology. The inlet valve spring was deformed. The edges of the inlet rubber flapper were rounded and worn. Wear was also observed on the lower side of the diaphragm body just above the outlet valve seat.
The outlet valve stem was separated in the shaft at the valve seat perpendicular to the valve stem axis. The outlet valve stem shaft was cut approximately 0.05 inch below the fracture surface, and then the lower end of the outlet valve stem was sectioned longitudinally. The pattern of voids on the outlet valve stem was similar to that of the inlet valve stem. No cracks were observed at the corners of the spring groove. The outlet valve stem fracture surface (lower surface) was viewed using SEM. Across the fracture surface, voids and detached solidification boundaries were observed. A black circular feature that was found near the center was a relatively large void. Between the voids and detached solidification boundaries, the fracture surface was relatively flat, and had features typical of fatigue. Faint striations were observed in several locations on the fracture surface. The outlet valve stem and outlet rubber flapper did not show signs of wear, and the outlet valve spring appeared undeformed.
According to a representative of WSK "PZL-Rzeszow," the manufacturer of the pump was unknown, and the manufacturing specifications were also unknown. Specifically the representative stated that WSK "PZL-Rzeszow" "do[es] not know characteristics of AC4886 fuel pump." An aluminum tab was attached to one of the pump's flange screws, and was marked "4886 262N." According to a representative of Dana Corporation, the marking is consistent with that used by Blackstone, formerly a subsidiary of the Dana Corporation. According to the representative of the Dana Corporation, the type 4886 (also known as a type AC4886 after the original design company) was designed for use in the Chevrolet Corvair. The pump was intended strictly for automotive applications, and was not intended for use in aviation. The pump is designed to operate at a minimum of 5.4 pounds per square inch and a maximum of 6.9 pounds per square inch static pressure at the pump outlet when the pump is operated at 1,800 revolutions per minute. According to a representative at WSK "PZL-Rzeszow," the operating limits for fuel pressure in Franklin model 6A-350-C1R engines are 0.5 pounds per square inch minimum to 8.5 pounds per square inch maximum.
The FAA issued "Type Certification Data Sheet No. E9EA" for the WSK "PZL-Rzeszow" Franklin 6A-350-C1R engine on December 8, 1994.
The airplane was released to the owner's representative.