On September 18, 2003, about 0905 eastern daylight time, an Air Tractor AT-502, N1528T, registered to and operated by Southeastern Aerial Crop Service, collided with trees during an aborted takeoff from the Southeastern Airport, Fort Pierce, Florida. 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 destroyed by impact and a postcrash fire, and the commercial-rated pilot, the sole occupant, was not injured. The flight was originating at the time of the occurrence. The pilot stated that he had sprayed five hopper loads of chemical previously that morning, and after landing following the fifth load application, 450 gallons of sulfur and water were loaded into the hopper. He began his takeoff roll with 10 degrees of flaps extended and slowly increased power during the takeoff roll paying attention not to exceed the prescribed torque value. When the airplane had just rotated, he lowered an additional 10 degrees of flaps and at that time, he heard a loud explosion, a momentary vibration, and the torque dropped to zero. The airplane settled to the runway, and he applied braking and maneuvered the airplane to the left to travel between two rows of grapefruit trees that were located past the departure end of the runway. The fuselage of the airplane traveled between the rows of trees, and at some point, the left wing contacted a tree, causing the airplane to veer to the left. The airplane came to rest with the tail elevated; the pilot evacuated the airplane and observed a fire initially near the forward part of the engine. He then observed a fire near the leading edge of the left wing. Postcrash disassembly of the engine assembly was performed at an engine overhaul facility. The engine had sustained postcrash fire damage and the nose reduction gear box case had been consumed by fire. All gears from the reduction gear box had no damage. The compressor would not rotate. Disassembly of the compressor showed all blades and disk were intact. The combustion area also was intact. Disassembly of the turbine section showed the No. 2 bearing housing had one of four bolts loose and there was damage to the bolt threads. Impact damage was present on the power turbine disk and blades, power turbine stator, compressor turbine disk and blades, and compressor stator. The accessory drive gearbox would not rotate. All gear teeth were intact. All accessories had sustained fire damage and were examined to the extent possible. The compressor turbine disk and blades, power turbine disk and blades, compressor turbine and power turbine stators, small exit duct, the compressor turbine stator lock plate, and the No. 2 bearing housing cover and bolts were forwarded to the engine manufacturer, Pratt and Whitney Canada, for examination under supervision of the Transportation Safety Board of Canada. The examination showed the power turbine blades were fractured at various locations within the airfoil. The fracture surfaces showed characteristics of tensile overload. The power turbine stator revealed significant impact damage. The compressor turbine disk exhibited two blades fractured near the platform, the remaining blades exhibiting severe impact damage. Fatigue was observed on the fracture surfaces of the two blades which had fractured near the platform. Corrosion was observed on the side of the compressor turbine blade platforms from these two fractured blades. The compressor turbine stator showed impact damage. The remaining components suffered from secondary damage. The corroded platform area of the compressor turbine blades contained foreign elements mostly composed of phosphate and lead oxide. The corrosion observed had caused a volume change which closed the required gap between blade platforms. According to Pratt and Whitney personnel this would have an adverse effect on the blade airfoils by altering the natural vibration frequency of the blades. Pratt and Whitney personnel concluded that the distress of the compressor turbine disk was initiated by the fracture of one blade in fatigue from multiple origins. The damage to stators, the power turbine disk, and remaining components was considered secondary to the compressor blade fracture. No burning damage was found on the stators. The fatigue fracture of the compressor turbine blades likely resulted from a loss of gap between the blade and platforms as a result of the observed corrosion. Logbook records show the engine was last inspected on August 1, 2003, 89.6 flight hours before the accident, when it received a 100 hour/Annual inspection. At the time of the accident the engine had accumulated 9,668.4 total flight hours, and a hot section inspection was last performed on July 8, 2002, 308.6 flight hours before the accident. The compressor turbine blades and power turbine blades were installed new in the engine on July 30, 1999, 2,222.2 flight hours before the accident.