NTSB Factual Report

HISTORY OF FLIGHT

On June 7, 2004, at 0730 Pacific daylight time, a Schweizer Aircraft Corp. G-164B single-engine airplane, N3626T, collided with terrain following a loss of engine power during takeoff from a private airstrip near Richvale, California. The airplane was operated by AvAg, Inc., Richvale, under the provisions of 14 CFR Part 137 as an aerial application flight. The commercial pilot, sole occupant, was not injured. The airplane was substantially damaged. Visual meteorological conditions prevailed, and a flight plan had not been filed for the local flight, which was departing at the time of the accident.

According to the pilot's written statement, it was the 5th load of the day. The pilot was taking off from the dirt strip, and after the airplane broke ground, the turboprop engine lost power. The pilot landed in a flooded rice field and the airplane impacted a dirt embankment (near a crossroad) that took the landing gear off. The airplane came to rest upright 50 yards off the end of the airstrip.

A follow-up interview with the pilot revealed the airplane was loaded to 6,205 pounds, which included 70 gallons of jet fuel and 300 gallons of herbicide. The pilot reported no anomalies during the takeoff roll; however, as the airplane became airborne the engine began a slow and subtle loss of power. The pilot applied full power and placed the fuel control from ground idle to flight idle. He was unable to dump the load or examine the engine instruments prior to impacting terrain.

The accident was not reported to the National Transportation Safety Board until July 15, 2004.

AIRCRAFT INFORMATION

The Pratt & Whitney PT6A-34 engine (serial number 56841) is orientated on the nose of the Schweizer G-164B with the inlet to the compressor section oriented toward the aft end by the accessory section. When looking at the profile view of the engine, the propeller reduction gearbox is oriented toward the front, followed by the turbine section, combustion section, compressor section, and ending with the accessory gearbox. The airflow through the engine flows into the inlet and moves forward through the compressor section, before making an "S" turn through the diffuser and turns 180 degrees to enter the combustion chamber. The air exits the combustion chamber and turns 180 degrees back toward the nose of the airplane where it then enters the turbine section. The air continues through two turbines before exiting the exhaust, which is located toward the front of the engine near the propeller reduction gearbox. The compressor section consists of three axial stage compressors and one centrifugal impeller, which is all driven by the compressor turbine (N1). Forward of the compressor turbine is the power turbine (N2), which drives the propeller. The two rotating assemblies are not connected together and turn at different speeds and in opposite directions, which is referred to as a "free turbine engine."

Review of the engine maintenance records revealed that it had accumulated 18,197 hours and 25,468 cycles since new. It underwent an overhaul on January 15, 2001, approximately 1,697 hours and 1,137 cycles prior to the accident. At 17,037 hours (August 12, 2002), the power section was repaired for bearing noise. The last inspection recorded was a 100-hour inspection, which was conducted at 18,082 hours on May 15, 2004.

TESTS AND RESEARCH

Following the accident, the operator inspected their fuel truck filters and found no debris or anomalies. They continued to utilize the fuel in the truck for other aircraft following the accident without experiencing additional power losses.

According to a written statement provided by the operator's mechanic, on June 23, 2004, the mechanic removed the fuel control unit (FCU) and high pressure fuel pump (HPFP) and inspected both drive shafts, which were intact. The mechanic noted a "small amount of blue staining on both [the] drive side of HPFP and the splined side of the FCU." He also noted the drive shaft on the FCU had a "very rough" bearing and "a lot of side to side movement." The HPFP inlet and outlet fuel filters were examined and found to be "ok." In addition, the FCU filter bowl contained fuel.

On June 24, 2004, the operator's mechanic removed the oil filter and found it clear of debris. The auxiliary fuel pump was removed and its splined drive was found to be intact. No contamination was found in the oil-to-fuel heater, the firewall fuel filter, the P3 filter, the fuel supply lines, the fuel nozzles, and the remainder of the fuel system.

The engine (along with the removed components) was shipped to Pratt & Whitney Engine Services facility in Los Alamitos, California. On September 2, 2004, an investigator from Pratt & Whitney Canada (PWC) and the Safety Board investigator-in-charge oversaw the disassembly and examination of the engine. All positional references are in relation to viewing the engine from the aft looking forward. Upstream and downstream references are in relation to gas path flow from the compressor inlet to the exhaust.

External review of the engine revealed impact damage to the compressor, turbine, and exhaust segments, and the engine was covered with dried mud. The reduction gearbox chip detector was clean.

External damage to the various segments prevented manual rotation of the compressor section due to case impingement on the compressor axial stages. Visual examination of the compressor's first stage blades revealed no foreign object damage. Due to the extent of engine damage, the combustion chamber was not removed; however, the liner displayed no indications of operational distress.

The turbine section was disassembled and examined. According to the PWC investigator, the compressor turbine guide vane airfoils displayed "normal operational erosion to their leading edges." The downstream side of inner drum displayed circumferential rubbing. The compressor turbine shroud displayed heavy circumferential rubbing and scoring, with frictional heat discoloration around the right-hand circumference due to radial contact with the compressor turbine blade tips. The compressor turbine blade tips also displayed circumferential rubbing, with frictional heat discoloration, consistent with the damage sustained by its shroud. The turbine disc on the upstream side displayed circumferential rubbing damage and heat discoloration. The downstream side had circumferential rubbing and machining that measured approximately 1/32 inches deep, along with material transfer with the power turbine guide vane ring.

The power turbine guide vane airfoils were intact, and its inner drum displayed heavy rubbing around the right hand circumference, with heavy frictional heat discoloration and material transfer consistent with the downstream side of compressor turbine (mentioned in previous paragraph). The power turbine shroud also displayed heavy circumferential rubbing. The power turbine airfoils were displaced forward in their fir-tree retention slots due to the compression of the exhaust duct. The blade tips displayed rubbing damage consistent with the shroud damage.

The propeller reduction gearbox was examined and upon case separation the gears rotated freely in place (exhaust duct was bent and deformed and precluded rotation of the aft end of the reduction gear box gears). No anomalies were noted with the gears or their teeth.

The accessory gearbox gears were rotated through manual rotation of the starter/generator drive shaft. The gears rotated freely and oil was pumped out of the oil pump housing. The bleed air valve was removed and tested with a thumb seal. The gasket appeared to in good condition and the valve worked properly.

The oil-to-fuel heater, HPFP, FCU, start control, propeller governor, overspeed governor, and compressor bleed valve were shipped to Pratt & Whitney Canada for further testing and evaluation.

On September 14, 2004, the aforementioned components were reviewed at Pratt & Whitney Canada's facility in Longueuil, Quebec, Canada. The hydro-pneumatic FCU's (Allied Signal part number 2524440-7-6, serial number A59527) exterior condition was unremarkable and its throttle lever rotated easily by hand. The drive splines, as well as the coupling inner and outer splines, were observed to be in a satisfactory condition. Manual rotation of the drive resulted in a rough running bearing consistent with a lack of lubrication in the drive bearings. Additional manipulation of the drive revealed a degree of radial looseness as noted by the operator's mechanic; however, no axial looseness was noted. Due to the poor condition of the drive, the decision was made to forgo functional testing and to conduct a teardown examination of the unit.

Internal examination of the FCU revealed the throttle lever cam follower displayed a worn grove and a greasy coating and small debris particles in the drive body. The flyweights were slightly stiff. Upon removal of the drive shaft from the drive body, one of the inner drive bearing balls fell out of position. The ball bearing inner race was visibly misaligned, the bearing seal was displaced, the ball bearings were loose, and the non-metallic bearing cage was missing. In addition, only grease residue was present in the bearing. Material analysis of the inner bearing identified fragmentation of the cage and spalling of the balls and inner raceway. Some orange colored particles collected from the drive and flow body inner cavities are believed to be fragmented portions of the inner bearing cage.

Operation of the fuel pump revealed that during the low-speed tests of 810 revolutions per minute (rpm) the unit pumped a capacity of 100 pounds per hour (pph), which was below the 140 pph minimum acceptance limit. The high-speed test of 6,350 rpm produced an output capacity of 1,243 pph, which was above the minimum acceptance limit of 1,181 pph.

The propeller governor underwent functional testing. The results revealed the pump had a few deviations from its testing tolerances; however, PWC personnel did not consider the deviations to be excessive. The overspeed governor was also placed on a test stand for functional testing. No anomalies were noted that would have resulted in the loss of engine power. The oil-to-fuel heater was functionally tested revealed no cross leakage between the fuel and oil units. The bleed air valve was tested. The valve passed the calibrated pressure check, but failed the pressure loss (leak) test. Disassembly of the bleed valve identified two small perforations in the bleed valve diaphragm; however, they would not have affected the valve's calibration. The start control unit was also tested and found to be slightly out of tolerance, but the deviations were not considered significant.

ADDITIONAL INFORMATION

According to a PT6A-34 training manual, under the section titled troubleshooting, a limited power range could be caused by an internal FCU problem. Additionally, fluctuations in power could be a result of a FCU Ng governor bearing problem.

Review of the engine maintenance records revealed the FCU (serial number A59527) was installed on the engine during the January 2001 overhaul. According to the maintenance entry, the unit had not accumulated any cycles at the time of installation. No additional maintenance entries were noted for the unit.

According to Pratt & Whitney Canada's maintenance manual, the interval for a FCU drive body inspection or drive shaft bearing replacement should be at the midlife of the engine's time between overhaul, which would be 1,750 hours (not exceeding 3,000 hours or 6 years). The unit accumulated a total of 1,697 hours prior to the accident.

NTSB Probable Cause

the loss of engine power during takeoff due to the engine's deteriorated fuel control unit's inner drive shaft bearing. The reason for the bearing's deteriorated condition could not be determined.