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A Cessna C208 aircraft, registered VH-CYC (CYC), with two pilots on board, was being operated for pilot type endorsement training. Air Traffic Control (ATC) had cleared the pilots to conduct upper level air work between 4,000 and 5,000 ft above mean sea level (AMSL) within a 5 NM radius of Green Island, Queensland. Following the upper level air work, the crew requested, and were granted a clearance for, a simulated engine failure and descent to 2,000 ft.

The pilot in command (PIC) reported that while completing the simulated engine failure training, he had retarded the power lever to the FLIGHT IDLE stop and the fuel condition lever to the LOW IDLE range, setting a value of 55% engine gas generator speed (Ng). The pilot under training then set the glide attitude at the best glide speed (for the operating weight) of about 79 knots indicated airspeed (KIAS). The PIC then instructed the pilot under training to place the propeller into the feathered position, and maintain best glide speed. The PIC reported that he instructed the pilot under training to advance the emergency power lever (EPL) to simulate manual introduction of fuel to the engine.

According to the PIC, he then noticed that there was no engine torque increase, with the engine inter-turbine temperature (ITT or T5) and Ng rapidly decreasing, and a strong smell of fuel in the cockpit. While the pilot under training flew the aircraft, the PIC placed the ignition switch to the ON position and also selected START on the engine starter switch. He then reportedly placed the EPL to the CLOSED position, the propeller to the UNFEATHERED position and the fuel condition lever to the IDLE CUTOFF position to clear the excess fuel from the engine. The PIC reported that they then increased the aircraft airspeed to 120 KIAS, at which point he reintroduced fuel into the engine by advancing the fuel condition lever. He reported that following these actions, the strong fuel smell persisted.

As the aircraft approached 1,500 ft, the PIC broadcast a MAYDAY, informing ATC that they had a 'flameout' of the engine and that they were going to complete a forced landing water ditching near Green Island. While the pilot under training flew the aircraft, the PIC placed the propeller into the feathered position, closed the fuel condition lever to the IDLE CUTOFF position and turned off the starter and ignition switches. They then completed a successful landing in a depth of about 2 m of water near Green Island. The pilots evacuated the aircraft without injury.

The aircraft, which sustained minor damage during the ditching, but subsequent substantial damage due to salt water immersion, was recovered to the mainland. Following examination of all connections and control linkages, the engine was removed for examination under the supervision of the Australian Transport Safety Bureau (ATSB) at the engine manufacturer's overhaul facility. The engine trend monitoring (ETM) data logger was also removed from the aircraft for examination.

Engine information

Manufacturer: Pratt & Whitney (Canada)
Model: PT6A-114
Serial number: 17099
Time since new: 8,473.9 hours
Cycles since new: 15,924 cycles
Time since overhaul: 4,713.4 hours

The general condition of the engine was good except for exfoliation corrosion of the magnesium and aluminium alloy components as a result of salt-water immersion. The first-stage axial compressor blades displayed significant erosion of the blade leading edges at the blade root portion of the airfoil. The erosion of the blades measured a maximum of .250 inch (.635 cm). According to the engine manufacturer's maintenance manual, the erosion limits of the compressor blade at the root was .250 inch without repair. The eight engine thermocouple probes were also examined. Testing of the probes indicated that they all passed the heat response test, but two probes did not pass the insulation test. The engine igniters operated satisfactorily when tested.

The engine fuel control unit (FCU) was removed to another facility for disassembly and examination under ATSB supervision. The examination of the FCU revealed no evidence of any internal component failure or anomaly, which would have prevented normal operation prior to salt-water immersion.

Engine temperature indicating system

The engine temperature indicating system consisted of a bus-bar assembly, eight individual thermocouple probes connected in parallel, a wiring harness incorporating a terminal block and an adjustable trim harness incorporating a T1 thermocouple probe.1 The T1 thermocouple probe was connected in parallel with the T5 wiring harness to bias the T5 signal and give the system a reference point. The engine manufacturer's maintenance manual included a note in the Engine Condition Trend Monitoring (ECTM) Shift Fault Isolation Chart stating that the T5 indication usually decreased when the thermocouple probes were unserviceable. The chart also noted that:

If several probes are broken or damaged, the loop resistance would not necessarily fall outside the allowable tolerance. However, erroneous temperature indications could occur due to reduced sampling.

An anomaly with the insulation of the thermocouple probes would typically manifest itself in abnormally low temperature readings.

Engine trend monitoring data logger

The ETM data logger recorded Ng, bus voltage, engine torque, T5, engine shaft horsepower (SHP), fuel consumption, airframe hours, engine total cycles, starts and duration and system exceedances. This information was electronically stored on a removable data key. Information stored on the data key could then be downloaded into a computer. The pilot reported that the data key was installed into the ETM display unit at the commencement of flight. However, after the aircraft was recovered, the data key was not found.

After preservation, the ETM data logger was shipped to the manufacturer for downloading. The manufacturer successfully recovered nineteen exceedances covering a period from 1 January 2003 to 3 February 2004. The majority of those logged were exceedances of propeller RPM and engine SHP.

Engine emergency power lever

The EPL, which was connected through linkages to the manual over-ride lever on the FCU, governed the fuel supply to the engine should a pneumatic section malfunction occur in the FCU. The EPL permitted the pilot to restore engine power by activating the lever to manually introduce fuel to the engine.

The aircraft manufacturer's Service Kit, SK208-142, provided for the installation of mechanical devices allowing for the installation of copper witness wire to the EPL. If the EPL was been moved from the NORMAL position, the copper witness wire would fracture and provide a physical indication that it had been activated. The installation of SK208-142 was not mandatory for Australian registered aircraft. However, the requirements of SK208-142 had been complied with on the aircraft.

Cessna Alert Bulletin, CAB01-15, included a requirement to ensure the fitment of the copper witness wire to the EPL of all aircraft that had SK208-142 installed. Compliance with the requirements of CAB01-15 was not mandated for Australian or United States (US) registered aircraft. According to the flight crew, no copper witness wire was installed on the aircraft at the time of the occurrence.

The aircraft maintenance manual stated that if the EPL witness wire was broken or missing, a determination was to be made as required by the engine maintenance manual, to assess if the engine limitations had been exceeded.

The aircraft manufacturer's information on the operation of the EPL stated that the use of the EPL was for emergency purposes only, and contained cautions about the use of the EPL for any other purposes. Further information about the aircraft manufacturer's use of the EPL is contained in appendix A.

The engine manufacturer's Service Information Letter (SIL) Number PT6A-053R2 addressed the use of the EPL. Although it also stated that the EPL was for emergency purposes only, it mentioned the use of the EPL for training purposes under supervision to maintain emergency practices proficiency. It included a note which suggested that familiarization training using the EPL be conducted on the ground. Further information on the operation of the EPL is contained in appendix B.

The pilot reported that, based on the reference to familiarization training in the SIL, he considered that the use of the EPL for in-flight familiarization training was acceptable.

Civil Aviation Regulation 1988, Part 50E addressed inconsistent requirements relating to aircraft operation and maintenance. Part 50E noted that by order of priority, the aircraft manufacturer's requirements superseded the requirements of an aircraft component manufacturer such as an engine manufacturer.

Recent engine maintenance

Date Engine hours since new Maintenance
22 January 2002 7,243.2 Remove and replace two thermocouples, hot section inspection, compressor turbine disc and FCU replaced.
9 July 2002 7,704.1 T5 busbar and thermocouples replaced.
25 November 2002 8,110.9 Compressor turbine disc inspected and reinstalled.
18 November 2003 8,431.5 Inspection in accordance with AD/ENG/5 (compressor first-stage)
28 January 2004 8,482.0 Hot section inspection (extension from 1,250 to 1,760 hours)

There were no engine logbook entries concerning engine compressor erosion. Engine compressor washes had been completed on a periodic basis as required. Civil Aviation Safety Authority requirements for the aircraft are contained in appendix C of this report. Engine trend data for the aircraft's engine are contained in appendix D.

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