SYLVESTER, GA, USA
N73080
Air Tractor 400
The pilot intentionally operated the engine exceeding the maximum temperature limits during aerial application flights which he was advised against by several different individuals. The pilot stated that during an aerial application flight he first heard a loud sound with a subsequent loss of engine power. He initiated a forced landing in a field and after touchdown with the brakes heavily applied to avoid obstructions ahead, the airplane nosed over. Examination of the engine revealed that 7 of the 58 compressor turbine blades were failed near the midspan or platform. Oxidation was detected at one of the 7 blades and creep and necking was detected at the trailing edge of another of the blades. Additionally, severe burn erosion was noted to the trailing edges of the compressor turbine guide vanes. According to the ASM Metals Handbook, creep is a time dependent strain occurring under stress. Additionally, at elevated temperature, the service life of a metal component subjected to either non-vibratory or vibratory loading is predictably limited.
On October 21, 1996, about 1130 eastern standard time, an Air Tractor, Inc., AT-400A, N73080, registered to a private individual, nosed over during a forced landing in a peanut field about 14 miles north of Sylvester, Georgia. Visual meteorological conditions prevailed at the time and no flight plan was filed for the 14 CFR Part 137 aerial application flight. The airplane was substantially damaged and the commercial-rated pilot sustained minor injuries. The flight originated about 10 minutes earlier from the operator's facilities in Sylvester, Georgia. The pilot stated to an FAA inspector that during a spray pass about 5 feet above ground level at 140 knots, he heard a loud sound and the engine experienced a loss of power. He landed the airplane in a peanut field and after touchdown, he heavily applied the brakes to avoid obstructions. The airplane then nosed over. Examination of the accident site by an FAA inspector revealed fuel soaked into the ground at least 27 inches beneath the surface. The airplane was recovered for further examination of the engine. Examination of the engine revealed that the power section was separated just aft of the exhaust duct. Visual examination only of the compressor section revealed circumferential scoring of the first stage blade tips with the first stage shroud. The engine was sent to the manufacturer's facility for further examination. Examination of the compressor turbine disk revealed that 7 of the 58 blades were fractured at the midspan or near the platform with the remaining blades fractured near the blade tips. Axial contact with the upstream side of the compressor turbine disk to compressor turbine guide vane ring was noted between blades 1 through 27. Also, axial contact was noted to the downstream side of the compressor turbine disk to the power turbine guide vane ring. Severe burning and erosion was noted on the blade tips and on the trailing edges of the blades. The compressor turbine guide vane ring was examined and the vane airfoil segments were found to exhibit severe burning erosion to the airfoil trailing edges. Additionally, molten weld material was noted to the leading edge of several of the vane airfoils which were weld repaired to about 1/2 the width of the vane. Metallurgical examination of the fracture surfaces of the 58 compressor turbine blades revealed that fatigue was not noted. Detailed metallurgical examination of blade Nos. 29-35 revealed that blade No. 29 exhibited evidence of oxidation at the trailing portion of the blade. Additionally, a crack and necking was observed at the trailing edge portion of blade No. 14. According to the ASM Metal's Handbook, creep is a time dependent strain occurring under stress. Also, at elevated temperature, the service life of a metal component subjected to either non-vibratory or vibratory loading is predictably limited. Review of the engine logbook revealed that the engine was signed off on March 15, 1996, as having a hot section inspection completed. The installed compressor turbine disk was removed and a different compressor turbine (CT) disk which had accumulated 6,080.1 hours and 15,221 cycles was installed in the engine at that time. The CT blades were stretch checked and inspected and the CT disk was balanced. Additionally, 12 vanes were replaced in the compressor turbine guide vane ring assembly. The replacement vanes were manufactured under the Federal Aviation Administration Parts Manufacturer Approval (FAA-PMA) system. The logbook entry was signed off as "All work done in accordance with Pratt & Whitney overhaul and maintenance manuals. Note: this engine is sold 'as is' and has no warranty." At the time of the accident the engine had accumulated 384 hours since installation on May 17, 1996, and 7,044.5 hours since overhaul. The pilot had previously reported to two separate individuals which operate facilities that perform 100-hour and Hot-Section inspections to turbine engines, that he had exceeded the engine red line temperature during spray applications on hot days. He had additionally complained of poor performance and was concerned about the rise in the Interstage Turbine Temperature (ITT) during the engine shutdown which is normal. An FAA certificated A & P with IA examined the engine on June 5, 1996, and discovered in part that the throttle was not making full travel, which was corrected. The engine was then started and according to the IA, the engine had acceptable performance. The IA further stated that in the beginning of August, the pilot called him again and stated that when he was motoring the engine after engine shutdown, he could hear a rubbing sound and during operation of his airplane, he was exceeding red line. The IA advised the pilot to not exceed red line and the pilot called again on August 16, 1996, complaining that the engine performance was decreasing. The IA advised the pilot to bring the airplane to his facility which the pilot complied the following day. The IA reported that he split the engine and examined the compressor turbine disk which exhibited evidence of light rubbing but the tip clearance was within limits. Burn erosion was noted at one of the compressor turbine guide ring vane which was replaced. The IA assembled the engine and operated it which revealed it made minimum acceptable performance. He did not sign the engine for return to service due to his belief that the engine performance would soon drop below the minimum acceptable. He reportedly also advised the pilot that a new set of compressor turbine blades would improve the performance but the pilot stated that he did not want to spend any money on the engine. There was no entry in the engine logbook which documents the engine disassembly. The wreckage minus the retained components (compressor turbine and compressor turbine guide vane ring) was released to Mr. Les Sychak on July 23, 1997. The retained components were also released to Mr. Sychak on October 21, 1997.
Was the pilots intentional disregard for operating the engine contrary to the flight manuals for intentionally exceeding the maximum engine temperature, resulting in creep failure of one of the compressor turbine blades.
Source: NTSB Aviation Accident Database
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