Kalamazoo, MI, USA
C-FFIH
BEECH A36TC
Witnesses and global positioning system data indicated that after takeoff the airplane tracked the runway centerline and climbed approximately 150 feet above ground level before turning to the right and descending. The airplane subsequently impacted the ground, a fence, and a pole about 1,400 feet from the departure end of the runway. A postimpact fire ensued. After acknowledging the takeoff clearance, no further communications were received from the accident airplane. Examination of the airplane’s structure, flight controls, engine controls, and fuel system did not reveal any anomalies consistent with a preimpact failure. The engine manufacturer had issued a Mandatory Service Bulletin (MSB) that was applicable to the accident airplane’s engine. The airplane’s engine was disassembled in order to replace components damaged during the impact sequence. During the disassembly the valve lifters that were the subject of the MSB were examined and found to have pitting. However, the engine was reassembled with the original valve lifters for a functional test run and during the test run the engine produced full rated power and no anomalies were noted. Data retrieved from an engine monitoring system installed on the accident airplane recorded a significant reduction in engine power during the accident takeoff. The monitoring system did not record throttle, mixture, or propeller control positions and the data was consistent with normal power reductions recorded during the landing phase of previous flights. No mechanical reason for the loss of engine power was found.
HISTORY OF FLIGHT On October 27, 2009, at 0842 eastern daylight time, a Beech A36TC, Canadian registration C-FFIH, was destroyed by impact forces and a post-impact fire following takeoff from runway 35 (6,502 feet by 150 feet, grooved asphalt), at the Kalamazoo/Battle Creek International Airport (AZO), Kalamazoo, Michigan. The airplane was piloted by a private pilot and was on an instrument flight rules (IFR) flight plan and was operated under Title 14 Code of Federal Regulations Part 91. The pilot was fatally injured. The flight was originating at the time of the accident and its intended destination was the Muskoka Airport (CYQA), Muskoka, Ontario, Canada. Communications transcripts showed that the pilot contacted AZO ground control at 0835 and was issued an IFR clearance from AZO to CYQA. The pilot was then instructed to taxi to runway 35. At 0840, the pilot contacted the AZO air traffic control tower and informed that he was ready for departure. The AZO tower controller issued a takeoff clearance which the pilot acknowledged. No further communications were received from the accident airplane. The airplane’s flight path was retrieved from the handheld global positioning system (GPS) receiver that was found in the wreckage. The data for the final flight began at 0833 and showed the airplane taxiing from the hangars located on the west side of the airport to runway 35. At 0841:03, the airplane began its takeoff roll. At 0841:40, the airplane was over the runway, about 650 feet from the departure threshold, at an altitude of about 70 feet above the ground (agl). At this point, the airplane had begun a right turn and was climbing. At 0841:57 the airplane reached its maximum altitude of about 156 feet agl. At this point, the airplane’s groundspeed was recorded as 83 knots. Three more data points were recorded with altitudes of 148, 126, and 61 feet agl, respectively. The groundspeeds at these points were 70, 64, and 64 knots respectively. The last recorded position was about 200 feet and 240 degrees from the initial impact point. Witnesses reported seeing the airplane before the crash. The witness descriptions of the airplane’s flight path were consistent with the flight path data retrieved from the GPS. Several witnesses reported that no smoke or flames were coming from the airplane prior to impact. Air traffic control personnel statements indicated that three controllers witnessed the accident sequence. Two of the controllers described hearing a reduction in engine power. One of those controllers stated that he saw a trail of black smoke coming from the airplane prior to impact. The other two controllers saw the airplane in a right turn prior to impact but made no mention of seeing smoke coming from the airplane. PERSONNEL INFORMATION The pilot, age 60, held a Canadian private pilot license with ratings for all non-high performance single and multi-engine land and sea airplanes. Canadian regulations define a high performance airplane as one that is specified in the minimum flight crew document as requiring only one pilot and that has a maximum speed (Vne) of 250 knots indicated airspeed (KIAS) or greater or a stall speed (Vso) of 80 KIAS or greater. The pilot also had a group 3 instrument rating valid for single engine airplanes. That rating was valid until October 1, 2010, provided the pilot had logged 6 hours of instrument flight time and 6 instrument approaches within the preceding 6 months. Records showed that the pilot held a private pilot certificate issued by the Federal Aviation Administration that was based on his Canadian pilot license. That certificate listed ratings for single and multi-engine land airplane ratings. The pilot also held a Canadian issued category 3 medical certificate. The examination date was listed as July 17, 2008, and the certificate was valid for 24 calendar months from the date of examination when used for flights outside of Canada. The limitations section listed that the pilot have glasses available. The pilot’s flight logbook was not recovered during the investigation. AIRCRAFT INFORMATION The airplane was a 1981 Beechcraft model A36TC, Bonanza, serial number EA-258. The airplane was a single engine low-wing monoplane with retractable landing gear. It was configured to seat 6 occupants including the pilot, and was constructed primarily of aluminum. The airplane was powered by a Teledyne Continental Motors (TCM) model TSIO-520-UB engine rated to produce 300 horsepower. FAA records indicated that the airplane was involved in an off-airport landing due to a loss of engine power in July 2009. The airframe was not damaged during the off-airport landing. The airplane was disassembled and transported to AZO subsequent to the off-airport landing. Maintenance records indicated that a 100-hour inspection was completed on September 28, 2009, at an airframe time of 3,167 hours. The entries in the maintenance records indicated that the wings and horizontal stabilizer were re-installed, and a newly manufactured engine, serial number 1000903, was installed. An entry in the records indicated that the airplane had also undergone an annual inspection as required by Canadian Aviation Regulations (CAR) 625, on September 28, 2009. The most recent maintenance was performed on October 26, 2009, and included an oil and filter change, replacement of the throttle control, re-clocking of the control yoke chains, straightening of the right aileron trim tab, and brake work. METEOROLOGICAL INFORMATION The AZO weather conditions at 0853 were: Winds 360 degrees at 4 knots; 6 statute miles visibility; few clouds at 1,000 feet above ground level (agl), overcast clouds at 2,000 feet agl; temperature 9 degrees Celsius; dew point 7 degrees Celsius; altimeter setting 30.02 inches of mercury. AIRPORT INFORMATION AZO was a public use airport with an air traffic control tower that operated on a part time basis. At the time of the accident the control tower was in operation. The airport had 3 runways, 17/35 (6,502 feet by 150 feet), 05/23 (3,438 feet by 100 feet), and 09/27 (2,800 feet by 60 feet). At the time of the accident runway 35 was in use for arriving and departing aircraft. WRECKAGE AND IMPACT INFORMATION The airplane impacted the ground, an airport perimeter fence, and a pole, about 1,400 feet and 27 degrees from the departure end of runway 35. The airplane’s fuel tanks were ruptured during the impact and fuel was released and ignited resulting in fire damage to several vehicles in a parking lot on the north side of the airport. The airplane came to rest upright and facing north about 200 feet from the initial impact point. The airplane had sustained substantial fire damage to the entire cabin section of the fuselage, the inboard left wing, and the right wing. The left wing remained attached to the fuselage and was intact with minimal impact damage. The outboard 1/3 of the right wing was separated from the inboard portion. The inboard portion of the right wing remained attached to the fuselage. The left aileron and both flaps remained attached to their respective wing surfaces. The right aileron remained attached to the separated portion of the right wing by its inboard hinge. The remaining right aileron hinges were separated from the wing. The aft fuselage and empennage remained attached to the cabin section of the fuselage. The tail surfaces were intact except the right elevator counterbalance weight was partially separated from the elevator. The cabin section roof was almost entirely consumed by the post impact fire. The interior of the cockpit area including the instrument panel was partially consumed by fire. Continuity of the primary flight control cables was verified from the cockpit controls to the respective surfaces. The left and right elevator trim tabs were found to be at 0 degrees and 3 degrees tab down respectively. The flap control handle and flap actuators were found to be in the up position. Aileron and elevator autopilot servo capstans rotated freely in both directions. The landing gear switch and landing gear actuators were found in the up position. The mixture and propeller controls were in the forward position and the throttle control was about 2.6 inches from its full forward position. The engine controls remained attached to their respective control arms within the engine compartment. The fuel selector handle and valve were found positioned for the right wing main fuel tank position. The fuel selector was removed from the airplane and disassembled. The fuel port was verified to be open and the filter bowl contained fuel. A test of the fuel within the bowl using water detecting paste was negative for the presence of water. The fuel auxiliary pump switch was found in the “HI” position. The fuel pump was removed from the airframe and a functional test performed. It was able to pump fuel during this test. The engine had sustained impact damage and sooting from the post-impact fire. The forward engine mounts were fractured and the rear mounts remained intact. The engine crankshaft flange was separated and the separated portion remained attached to the propeller. The propeller was found about 50 feet from the main wreckage of the airplane. The engine was removed from the airframe. The engine driven vacuum pump was removed to facilitate rotation using a tool inserted into the accessory drive. The vacuum pump was operated by hand and suction and exhaust were noted. The upper spark plugs were removed and visually inspected. No abnormal wear indications were noted. A borescope examination of all of the cylinders was performed and no foreign materials or abnormal signatures were noted. The engine was rotated using the accessory drive and thumb compression, suction, and valve action was confirmed on all cylinders. Subsequent to the field examination of the engine it was crated and shipped to the manufacturer’s facility for further testing. MEDICAL AND PATHOLOGICAL INFORMATION An autopsy was performed on October 28, 2009, at Sparrow Hospital, Lansing, Michigan. The cause of death was listed as thermal injuries and inhalation of products of combustion. A Final Forensic Toxicology Fatal Accident Report prepared by the FAA listed the following findings: 30 percent Carbon monoxide detected in blood 2.02 (ug/ml) Cyanide detected in blood No ethanol detected in vitreous Quinine detected in urine TESTS AND RESEARCH The engine was shipped to the manufacturer for examination and a possible operational test run. The NTSB investigator in charge (IIC) was present for these tests. Upon the IIC’s arrival at the manufacturer’s facility, the engine was uncrated and evaluated. The crankshaft flange had separated from the remainder of the crankshaft during the accident sequence. The break was aft of the forward most portion of the crankcase preventing welding of a substitute flange to facilitate an operational test. In addition to the crankshaft break, the forward induction tubes exhibited impact damage. The engine was disassembled to allow replacement of the damaged crankshaft. When the engine was disassembled, pitting was noted on the intake valve lifter faces for the number 4, 5 and 6 cylinders. The engine was reassembled using original components with the exception of the crankshaft, engine mounts, and damaged induction and exhaust tubes. The original camshaft and valve lifters were reinstalled for the functional test. The magnetos were timed according to manufacturer specifications. No adjustments were made to the fuel system or turbocharger controls. The engine was installed in a test cell at the manufacturer’s facility and an operational test was performed. The engine started normally without hesitation and was warmed to operating temperature by increasing the rotational speed in stages. After warm-up, the engine was advanced to full throttle and produced rated power. The engine was then subjected to 6 rapid throttle advancements from idle to full throttle. The engine performed normally without hesitation throughout the test. TCM issued a Mandatory Service Bulletin, number MSB09-8, on November 3, 2009, concerning inspection and removal of certain hydraulic lifters installed in various engines due to accelerated wear. The scope of the service bulletin included the accident airplane’s engine. TCM subsequently issued SB MSB09-8A which expanded the effectivity of the previous service bulletin. The FAA issued Airworthiness Directive 2009-24-52, on November 18, 2009 requiring operators to determine if the subject valve lifters indicated in TCM SB MSB09-8 were installed in various model engines and if installed to remove said lifters before further flight. This AD was subsequently modified to expand the effectivity as indicated in TCM SB MSB09-8A, on June 1, 2010. The airplane was equipped with a Xerion Auracle II 6-cylinder engine monitoring system. Downloaded data included many flights including the accident flight. Parameters recorded by the system included exhaust gas temperature, cylinder head temperature, oil temperature, oil pressure, fuel flow, manifold air pressure, engine horsepower, and rpm. The system did not record throttle, mixture, or propeller control positions. Review of the data for the accident flight confirmed that there was a reduction in the horsepower output of the engine. During the accident flight, engine horsepower increased to over 300 horsepower, and remained at this level for approximately 20 seconds. The engine power then reduced to about 275 horsepower for another 10 seconds. At this point, the engine power dropped off significantly falling below 50 horsepower about 10 seconds later. The engine power rose briefly to about 140 horsepower before again dropping to approximately 75 horsepower. Changes in fuel flow and engine manifold air pressure were recorded which corresponded with the engine’s recorded power output. The recorded power reduction during the accident flight was consistent with power reductions recorded during the landing phase of previous flights.
The loss of engine power for an undetermined reason.
Source: NTSB Aviation Accident Database
Aviation Accidents App
In-Depth Access to Aviation Accident Reports
