ERA13LA204

A total loss of engine power during cruise flight for reasons that could not be determined because postaccident examination and testing did not reveal any mechanical malfunctions or failures that would have precluded normal operation.

The pilot and the pilot-rated passenger departed for the destination airport with about 50 gallons of fuel on board. About 35 minutes into the flight, they heard a "pop." They then scanned the instruments and noted that everything looked good. A short time later, they heard additional "pop" sounds. The pilot then began descending the airplane from 4,500 ft mean sea level (msl) to get under the clouds for a "better visual," and, when the airplane was about 2,000 ft msl, the pilot increased the mixture control to full rich and turned the boost pump on. He then looked at the exhaust gas temperature (EGT) display and noticed that three of the cylinders "had gone cold." The engine then began losing power and running roughly. At this point, he decided to divert, and, as he turned the airplane toward the diversion airport, the engine lost all power. The pilot decided to land on a road. After touching down, he headed the airplane off the road to the right to avoid a car, and the airplane went down the shoulder. The right wing got caught in the grass, which caused the airplane to veer right. The airplane then struck a barbed wire fence and was substantially damaged. Review of data stored in the onboard engine analyzer indicated that the cylinder head temperature and EGT for all of the engine's cylinders dropped over about a 5-minute period in the following order: No. 5, No.3, No. 4, No.1, No. 2, and No. 6, which is typically indicative of a fuel or ignition problem. However, examination and testing of the ignition system did not reveal any damage, unusual wear conditions, or evidence of significant contamination, and all of the ignition components successfully passed new part performance checks. No conditions were found that would have prevented normal ignition operation. Additionally, due to the duration of the loss of engine power and the fact that the ignition system had dual magnetos, it was unlikely that an ignition system problem caused the power loss. Examination of the fuel system revealed that, at the time of the accident, the airplane still had about 40 gallons of clean fuel onboard. Examination and testing of the fuel injection system and engine did not reveal evidence of any preimpact failures or malfunctions. After the examination of the engine and replacement of parts damaged during the accident, the engine was started and an engine test run was conducted with no anomalies noted. According to the airplane's flight manual, modern aircraft engines are extremely durable and seldom fail catastrophically without plenty of advance warning (lowering oil pressure, excessive mechanical noise, rising oil temperature, etc.). However, since none of these conditions were reported by the pilot and postaccident examinations of the ignition, fuel injection, and lubrication systems, as well as the mechanical components of the engine, determined that they were functional at the time of the accident, the reason for the engine failure could not be determined.

The airplane was not equipped with a cockpit voice recorder or flight data recorder, nor was it required to be under the CFRs. It was however equipped with a J. P. Instruments EDM-700 which is a panel mounted instrument that the pilot can monitor and records up to 24 parameters related to engine operations. The unit contained non-volatile memory for data storage of the parameters recorded and calculated. The rate at which the data was stored and selected by the operator was from 2 to 500 seconds per sample. The memory could store up to 20 hours of data at a 6 second sample rate. The EDM recording contained data from 35 power cycles. The event flight was logged as "Flt79" and was the second to last recording available upon download. Data retrieved from the unit included cylinder head temperature (CHT), and exhaust gas temperature (EGT). The unit had not been set up to record fuel flow. Review of the data indicated that a drop in CHT and EGT for all of the engines cylinders occurred over an approximate 5 minute period. The drop first occurred on Cylinder No. 5, then No.3, then No. 4, then No.1, then No. 2, and finally on No. 6. Fuel System The fuel system consisted of two 33 gallon wing tanks. There was no provision for cross feed as fuel was used from both wing tanks simultaneously. A 4 gallon fuel sump tank was located behind the rear seat to assure fuel supply to the engine in normal flight attitudes. Each main tank and the sump tank were vented. A mechanical engine-driven fuel pump transferred fuel from the sump tank to the fuel injection system. An auxiliary electric fuel pump provided backup for the engine-driven pump. Fuel pressure was indicated on a gauge in the cockpit. The electric pump could be turned on if the engine-driven pump failed which would be noted by a loss of fuel pressure. The electric fuel pump would also be used to provide fuel pressure redundancy during low altitude operations, such as takeoff and landing. There is one fuel drain on the airplane, under the sump tank. Examination of the fuel system by an FAA inspector did not reveal evidence of any preimpact failure or malfunction of the fuel system. Approximately 35 gallons of fuel was discovered in the left wing tank and 5 gallons in the right fuel tank. Fuel drained from the system through the fuel drain located under the sump tank was clean, and did not exhibit evidence of water or sediment being present. Fuel Injection System The fuel injection system operated by measuring the airflow through the throttle body of the servo valve regulator controls, and used this measurement to operate the servo valve within the control. The regulated fuel pressure established by the servo valve was then used to control the distributor valve assembly, which then scheduled fuel flow in proportion to the air flow. Examination of the fuel injection system revealed no evidence of any preimpact failure or malfunction, and airflow tests of the fuel servo did not reveal any anomalies. Internal examination of the fuel servo also revealed that it was in new condition, and no contamination was present. It was later overhauled, recalibrated, and airflow tested a second time for return to service. Ignition Booster System The engine was equipped with dual magnetos and a SlickSTART ignition booster system which integrated solid state electronics with the engine's conventional ignition hardware. Examination of the SlickSTART ignition booster revealed that the unit's housing and mounting bracket were clean and undamaged and the exposed encapsulation compound was complete with no voids, cracks, or separation from the terminals or enclosure. Further examination revealed that, terminals No. 2 and No. 3 were capped with insulated slip-on terminals. The wire wells of the mated terminals were not deformed, indicating that no conductors had ever been crimped in place. Terminal No. 3 which was supposed to be connected to the retard breaker contacts in the left magneto was not connected to them, indicating that the retard breaker in the left magneto was inactive during engine starting. Slick Service Bulletin SB1-06A, required that terminal No. 2 be removed from the SlickSTART ignition booster, and that a notation be made on the unit's identification label indicating that the unit was in compliance with SB1-06A. Examination of the identification label revealed that the notation was missing. Terminal No. 2 being capped by an insulated slip-on terminal indicated however that even though no notation had been made on the identification label and terminal No. 2 had not been removed from the SlickSTART ignition booster, that Terminal No. 2 was insulated and not being used. Testing of the SlickSTART unit was also performed, in accordance with Champion Aerospace's new part acceptance test procedures. During the testing the unit operated properly, and passed all tests. Left Magneto Examination of the left Magneto revealed that there was no inspection lacquer on the cover screws indicating that the magneto had been internally inspected or serviced at some time subsequent to being manufactured. The magneto exterior was clean and undamaged. The distributor towers were clean, with no evidence of contamination or arcing. There was minimal evidence of star lock washer deformation under the retard breaker connection nut, which was also indicative that terminal No. 3 of the SlickSTART ignition booster had not been connected to the retard breaker contacts. A small amount of crusty-oily residue was observed on the rotor shaft between the drive taper and oil seal. The vent bushing in the lower housing adjacent to the rotor was not installed being previously removed when the magneto had a tachometer pickup installed in this location when the magneto was in the airplane. There was evidence of slight contact by the tachometer pickup with the rotor pole-pieces where there was an approximately ¼ inch wide area of rotational polishing on each of the two pole pieces, closest to the distributer end of the magneto. There was no measurable depth to the polished areas and no evidence of contamination visible through the vent hole. The rotor shaft rotated normally. Internal examination revealed that the inside of the magneto looked factory-new. Both sets of points would open and close and the contacts were clean. The main breaker contacts displayed a frosty gray color on the edges which was indicative of normal operation. The cam was in good condition. The carbon brush appeared new, and there was no indication of carbon dust or other contamination inside the magneto. Testing of the left magneto on a motorized production test stand per Champion Aerospace part acceptance test requirements did not reveal any anomalies. The left magneto operated properly, and passed all tests. Right Magneto Examination of the right magneto revealed that there was no inspection lacquer on the cover screws indicating that the magneto had been internally inspected or serviced at some time subsequent to being manufactured. The magneto exterior was clean and undamaged. The distributor towers were clean, with no evidence of contamination or arcing. A moderate amount of crusty-oily residue was observed on the rotor shaft between the drive taper and oil seal. The rotor shaft rotated normally. Internal examination revealed that the inside of the magneto looked factory-new. The breaker points would open and close and the contacts were very clean and exhibited a frosty gray color on the edges which was indicative of normal operation. The cam was in good condition. The carbon brush appeared new, and there was no indication of carbon dust or other contamination inside the magneto. Testing of the right magneto on a motorized production test stand per Champion Aerospace part acceptance test requirements did not reveal any anomalies. The right magneto operated properly, and passed all tests. Engine Exami