WPR17FA179

A loss of engine power for reasons that could not be determined based on available evidence.

The pilot departed on a visual flight rules cross-country flight in his turboprop-equipped experimental airplane. Onboard data indicated that, upon reaching about 6,500 ft mean sea level, about 5 minutes after takeoff, the airplane pitched down about 25° and entered an approximate 35° left bank; the engine power reduced to idle with a simultaneous decrease in fuel flow. Several seconds later, the voltage began to slowly decrease, consistent with the generator turning off. The airplane continued to descend toward the nearest airport and ultimately impacted powerlines and terrain about 1.3 nautical miles from the approach end of the runway. The pilot made no radio transmissions that indicated that he was experiencing an emergency. Examination of the fuel system and engine revealed no mechanical malfunctions or failures, and the reason for the reduction of engine power could not be determined. The propeller was found in a low pitch position. The propeller was not equipped with a feathering pump; therefore, unless the propeller control was moved to the feather position immediately following a loss of engine power, the speed of the counterweights and the aerodynamic loads on the blades would quickly degrade to a point where feathering would become impossible. It is likely that, had the pilot immediately placed the propeller control into the feather position following the loss of engine power, significantly less drag would have been produced by the propeller and the airplane would likely have been able to reach the intended runway. The first responders reported that the pilot stopped breathing shortly after being extricated from the wreckage. Given that the pilot diverted the airplane and initially survived the impact, it is unlikely that he experienced sudden incapacitation during the flight.

The pilot's personal flight records were not recovered. On his last application for a medical certificate, the pilot reported 7,000 total hours of flight experience. The accident site was in the back yard of a residence located at the corner of two streets. Powerlines were located 190 ft north of the wreckage with two support structures (wood poles) on both sides of the street about 73 ft apart. A powerline about 190 ft from the main wreckage exhibited several bends in the center area consistent with airplane impact. (See Figure 2) Figure 2: Accident Site The airplane came to rest upright on level terrain oriented on an approximate 100° magnetic heading. The main wreckage, which comprised a majority of the airframe and engine, was located about 1.3 nautical miles (nm) from the approach end of runway 16 at MCC. The airframe remained relatively intact and the cockpit sustained minor crush damage. The throttle was between the mid- and full-throttle positions; the propeller lever was at feather and the condition lever was mid-range. The throttle quadrant retention clip was displaced and the throttle lever was slightly bent. It could not be determined if the clip displacement was the result of cockpit deformation at impact. The generator switch was selected to "on," and the generator circuit breaker was out. The air conditioner switch was "on." Fuel System According to the recorded airplane data, the fuel pressure remained relatively stable through the flight. After the accident, the left and right fuel pumps were selected to "auto" and the center pump was selected to "off." During post-accident examination, the fuel line was detached from the FCU and the system was activated. Upon activation of the boost pump, about 20 gallons of fluid was pumped from the header tank at a flow rate of about 70 gph, consistent with normal operation. An external examination revealed that the FCU had sustained impact damage, and whether there was air in the system could not be determined. Functional testing of the FCU and hydromechanical system revealed no anomalies that would have prevented normal operation. The mechanical fuel pump was removed and disassembled, revealing that its shaft was intact. There was no evidence of excessive wear or pre-impact damage that would have prevented normal operation. Engine Examination The engine was removed and placed in a test cell for operational testing. The engine started normally. The propeller was cycled twice from fine to feather pitch to purge the propeller piston cavity of air. After an initial idle period, the engine power was increased to 85% N1. Acceleration and deceleration behavior were acceptable, with no indications of hesitation, stall, or flameout. The examination revealed no evidence of pre-impact mechanical malfunction or failure that would have precluded normal operation. Propeller Examination The propeller was generally intact and clean. The spinner was not present. When one blade was rotated around its span axis, all other blades rotated in unison, consistent with internal mechanical integrity. There was no evidence of positive blade twisting along the span axis, which is normally observed when the propeller is being driven with engine power at impact. The leading edges of all the blades, although abraded from normal use, displayed no evidence of soft or hard body impact damage. A geometric analysis of shear damage to the bearing ring concluded that the blades were at the minimum flight angle (low pitch stop) of about 18° to 20°, or fine pitch. The emergency condition of the blades during an engine problem should be feather, or 90°; the dual-acting propeller was not equipped with an internal feathering spring. Propeller Governor An external visual examination revealed that the propeller governor was undamaged. Oil was seen exiting the mounting flange passages. The accident airplane was not equipped with an emergency electric propeller feathering pump. According to Lancair, there was no requirement or installation guidance for such a system. The airplane was equipped with a VR Avionics Turbine Starter Limiting/Monitoring System (TSLM). It was designed to act as a start sequence controller, an engine protection limiter, and an engine monitor/recorder. Data from the unit indicated that the pilot did not attempt to restart the engine during the flight. The Lancair IV-TP was an experimental, amateur-built airplane constructed mainly of composite materials The pilot purchased the kit directly from Lancair International Inc., in November 1995. The airplane received a special airworthiness certificate in the experimental category in September 2004. The logbook entry on August 4, 2016, stated that "a list of discrepancies and unairworthy items" were provided to the airplane's owner. This list, and the mechanic who performed the inspection, could not be located after the accident. The pilot recorded that he replaced the starter/generator in July 2016 at a total time of 868.0 hours. Engine and Propeller The Diemech M601D engine is a two-spool engine comprising a gas generator that drives a power turbine, which drives a reduction gearbox. The gas generator compressor consists of two axial flow stages and one centrifugal stage. Inlet air enters the compressor section radially just forward of the accessory section and travels forward through the compressor. The exiting compressor air enters an annular combustor to mix with fuel for the combustion process. The gas generator turbine nozzles then direct the expanded flowpath gases to the gas generator turbine, which directs the exiting gases to the power turbine for the final power extraction before exiting the engine forward of the compressor inlet. The power turbine drives the propeller system by means of the reduction gearbox. The accessory gearbox, which is located on the aft end of the engine drives all engine accessories by a direct shaft coming from the compressor spool. Typical engine accessories are the main fuel pump, fuel control unit, starter/generator, hydraulic pump, and the propeller governor, which is driven by the reduction gearbox located at the front of the engine. The oil system is a circulatory pressure system with an integral oil tank incorporated into the accessory gearbox. This system provides lubrication for all areas of the engine and oil pressure for the torque meter and propeller pitch control. The powerplant was controlled by three sets of levers. The power lever controlled the power output of the engine and the propeller blade angles in Beta and reverse. The propeller lever controlled the propeller speed via the primary propeller governor and emergency propeller feathering. The condition lever actuated the fuel shutoff valve and, if an emergency circuit was on, controlled engine power. The propeller was equipped with an overspeed governor on the cylinder front face, which featured an internal spring-loaded weighted valve. Centrifugal forces of the propeller rpm act on the weighted valve, and once the spring pressure is overcome, the valve opens, allowing oil from the low pitch area in the hub to the drain until the rpm decreases to correspond with speed setting. The airplane was not equipped with an emergency electric propeller feather pump; thus, emergency feathering could be activated by moving the propeller lever onto the feather stop. However, this required the gas generator portion of the engine to still be operating, since the gas generator drove the main oil pump and provided oil pressure to the propeller for pitch control and emergency feathering. However, when an engine loses power for any reason, the gas generator section quickly stops rotating and there is no engine oil pressure available to feather the propeller. With no engine power to turn the propeller, it will quickly stop rotating, making the counterweights and aerodynamic pressure on the rotatin