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On December 28, 2004, at 1135 central standard time, a Piper PA-23-250 (Aztec), N5398M, piloted by a commercial pilot, was substantially damaged during an in-flight collision with terrain at the Gogebic-Iron County Airport (IWD), Ironwood, Michigan. The pilot had reported an engine problem about 20 minutes before the accident while en route to IWD. The personal flight was operating under the provisions of 14 CFR Part 91. The flight was on an instrument flight rules (IFR) flight plan until the pilot canceled IFR on approach to IWD and proceeded under visual flight rules (VFR). Visual meteorological conditions prevailed at IWD. The pilot and four passengers sustained fatal injuries. The flight departed Menominee-Marinette Twin County Airport (MNM), Menominee, Michigan, about 1010.
An individual representing the accident aircraft contacted Green Bay Automated Flight Service Station (AFSS) at 0814 and obtained a pre-flight weather briefing from MNM to IWD. The pilot called Green Bay AFSS again at 0903 and requested an updated weather briefing. At the conclusion of that briefing the pilot filed an IFR flight plan from MNM to IWD.
The Federal Aviation Administration (FAA) provided transcripts of air traffic control (ATC) transmissions with the accident aircraft and radar track data for the flight. Radar track data was processed and plotted by the NTSB. The transcripts and plot of the radar track data are included with the docket information associated with the accident file.
The pilot established contact with Minneapolis Air Route Traffic Control Center (ARTCC) at 1021. The flight was approximately 10 nautical miles (nm) northwest of MNM and climbing through 5,000 feet mean sea level (msl) at the time. The flight climbed to 8,000 feet msl en route to IWD. Communications were routine during this portion of the flight. The pilot was informed that he could expect radar vectors for the Instrument Landing System (ILS) approach to runway 27 at IWD
At 1104, ATC instructed the flight to descend at pilot's discretion and maintain 3,500 feet msl. The pilot acknowledged this clearance.
At 1114:58 (HHMM:SS) the pilot transmitted: "I've got single engine problems here mayday mayday." The controller replied that the closest airport was IWD, the intended destination, and advised a heading of 310 degrees from the aircraft's present position.
At 1119:36 the controller instructed the pilot to turn left to 290 degrees to intercept the localizer on the ILS approach. Twenty-seven seconds later the pilot inquired about the aircraft's heading. The controller replied "you just took a hard left again on that twenty degree turn" and provided a 310-degree heading for the localizer intercept. The pilot stated that he had a "vacuum failure," however, he noted that he was able to maintain a heading. The aircraft subsequently descended below the level of radar coverage at 1120:35.
The pilot reported intercepting the ILS localizer at 1122:04. The pilot also stated, "[I] am single engine and one propeller feathered." At 1127:07, the pilot reported that the runway was in sight. At 1128:00, he reported a "complete loss of hydraulics" to the center controller. The pilot added that he was planning to circle to the left and hand pump the landing gear down. No further transmissions were received from the accident aircraft.
The plot of radar track data for the accident flight depicted the airplane approaching IWD from the southeast. The plot showed the airplane on a northwest ground track passing 6,700 feet msl at 1106:45 and descending until it reached 3,500 feet msl at 1113:46.
At 1114:58, when the pilot initially indicated that he had an engine problem, the radar track indicated that the aircraft was at 3,500 feet msl on an approximate magnetic course of 317 degrees. The aircraft then entered a descending left turn, reaching an altitude of 2,500 feet msl on a 169-degree magnetic course at 1115:22, 24 seconds later. The calculated average descent rate during this time frame was 2,500 feet-per-minute. The calculated average rate of course change during this time was approximately 7 degrees-per-second. A standard rate turn is defined at 3 degrees-per-second.
The track data indicated that the aircraft then entered a climbing right turn, reaching 3,500 feet msl at 1118:10 on an approximate course of 316 degrees. The aircraft subsequently turned left and descended again. Final radar contact was at 1120:35, 10.8 nm east of IWD at 2,400 feet msl.
The accident site was located in the side yard of a residence south of the IWD airport. The owner was home at the time. He reported hearing a loud noise and looked out a window to investigate. When he saw the accident site he contacted the sheriff's department.
The pilot held a commercial pilot certificate with airplane single and multi-engine land and instrument airplane ratings. He also held a flight instructor certificate with an airplane single-engine rating and a ground instructor certificate with advanced and instrument ratings.
The pilot was issued a second class airman medical certificate on October 9, 2004, with a restriction that corrective lenses must be worn.
The pilot's logbook was reviewed. The last logged flight was on November 30, 2004. According to the logbook, the pilot had accumulated 950.4 hours total flight time. He had logged 313.7 hours in multi-engine airplanes, with 11.5 hours logged within the previous 90 days. The pilot's most recent flight review was completed on February 26, 2003.
Within the previous 6 months, the pilot had logged 8.0 hours actual instrument flight time and 5.7 hours simulated instrument flight time. He had logged 20 approaches on flights where actual or simulated instrument flight time was also noted. An additional 24 approaches were logged without any corresponding instrument time for the particular flight.
The accident airplane was a twin-engine 1977 Piper PA-23-250 Aztec, serial number 27-7754135. The airplane was a six-place, low-wing configuration with retractable tricycle landing gear. The airplane was powered by Lycoming IO-540-C4B5 six-cylinder, fuel-injected, reciprocating engines. Each engine was rated at 250 horsepower.
The engines provided power through Hartzell HC-E2YK-2RBS two-bladed propeller assemblies. The propellers were a hydraulically operated, constant speed design with feathering capability. The propeller blades were moved to lower pitch settings by oil pressure from the propeller governor. The blades were moved to a higher pitch setting by spring and air charge in the hub.
The maintenance logbooks were reviewed. The most recent annual inspection was completed on December 11, 2004, at a total airframe time of 7,210.5 hours. The engine and propeller logbooks noted a recording hour meter of 41.9 hours at that inspection. The hour meter read 48.6 hours at the accident site, indicating 6.7 hours had accumulated since the annual inspection.
The right engine assembly, serial number RL-9402-48, was overhauled on April 24, 1997, and installed on the accident aircraft on December 7, 1997. The logbook noted that 6 "CermiNil" remanufactured cylinders had been installed on the engine. At the time of the last annual inspection, the engine had accumulated 343.8 hours since overhaul. Total time for the engine was not noted.
The left engine assembly, serial number RL-11740-48, was overhauled on April 24, 1997, and installed on the accident aircraft on December 7, 1997. According to the logbook, 6 "CermiNil" remanufactured cylinders were also installed on this engine during the overhaul. Time since overhaul was 343.8 hours and total engine time was not noted in the log.
The IWD Automated Weather Observing System (AWOS) recorded at 1135: Wind from 260 degrees at 12 knots; visibility 5 statute miles; overcast clouds at 1,100 feet above ground level (agl); temperature -04 degrees Celsius; dew point -08 degrees Celsius; and altimeter 29.89 inches of mercury.
WRECKAGE AND IMPACT INFORMATION
The accident site was located in the side yard of a residence adjacent to the IWD airport. Coordinates of the site were determined to be 46 degrees 31 minutes 17 seconds north latitude, 090 degrees 07 minutes 42 seconds west longitude using a handheld global positioning system (GPS) receiver. A line of trees about 40 feet in height was located approximately 19 feet west of the accident site. The tree line was oriented in a north-south direction and extended from the roadway south of the site to a point west of the accident site. No evidence of a tree strike was observed. Snow on the tree branches also appeared undisturbed.
A ground impact mark was located about 30 feet south of the wreckage location. The mark consisted of an elliptical depression in the snow about 20 feet long and 5 feet wide. Within the boundary of the depression, there were two gouges extending about 2 feet into the ground. They were 12 feet apart and measured approximately 4 feet by 4 feet in length and width. They were centered on the width of the depression and located symmetrically to the length.
The airplane came to rest on an approximate heading of 170 degrees magnetic. The wreckage was resting on the nose and the wings. The empennage was oriented vertically, approximately 65 degrees relative to the terrain.
The nose structure and engines were dislocated upward relative to the airframe. The wings exhibited leading edge crushing along the entire length. The empennage was buckled forward of the vertical stabilizer. Ice accretion was present on the leading edges of the wings, stabilator, vertical stabilizer, and antennas.
The flight control surfaces remained attached to the airframe. Control continuity was verified from the rudder and stabilator to the cabin area and from the ailerons to the cabin area. Stabilator trim continuity was confirmed. The bracket connecting the right stabilator half to the control rod had separated from the stabilator. The failure appearance was consistent with overload. The stabilator trim jackscrew position was measured at 0.2 inches, consistent with a slight nose down trim condition.
The wing flaps remained attached to the airframe, however the wing structure common to the hinges had separated at several locations. Appearance of the failure areas was consistent with overload. The wing flap hydraulic actuator was extended consistent with a flaps up configuration.
The main landing gear was observed in the extended position. The down lock linkage was intact and in the extended position.
The right engine examination did not reveal any anomalies associated with a pre-impact failure. Internal engine and accessory section continuity was verified via crankshaft rotation. The magnetos provided spark at each ignition lead when rotated. The spark plug electrodes appeared light gray consistent with normal wear. The fuel divider was disassembled. The diaphragm was undamaged and fluid consistent in texture and odor to aviation fuel was present. The right vacuum pump was disassembled. The rotor and vanes were intact.
The right propeller exhibited S-shaped bending, leading edge gouges, and chordwise scratching. The blades remained attached to the hub.
The left engine examination revealed that the No. 4 cylinder was fractured completely in the threaded area of the cylinder head. The cylinder head section was separated from the cylinder body approximately 1 inch, but was otherwise resting in position within the engine assembly. The valve lifter rods remained attached to the cylinder head section and were separated from the crankcase housing.
Internal engine and accessory section continuity was observed via crankshaft rotation. The No. 4 piston traveled freely within the remaining cylinder base section. Intake and exhaust valves functioned during crankcase rotation with the exception of the No. 4 cylinder.
The left engine magnetos produced a spark at each ignition lead when rotated. The spark plugs were light gray in appearance consistent with normal wear. The fuel divider was disassembled. The diaphragm was undamaged and fluid consistent in texture and odor to aviation fuel was present.
The left engine propeller blades remained attached to the hub. One blade exhibited a shallow S-bend over the length of the blade. Leading edge gouges and chordwise scratches were present. The other blade was bent aft approximately 45-degrees at a point about 1/3 span. No leading edge damage to this blade was observed. The propeller governor was intact and secured to the crankcase. The propeller pitch control linkage was attached to the control arm on the governor unit. The governor unit was removed for further testing.
The left vacuum pump housing was damaged on one side. The pump was disassembled. The rotor was cracked near the center but was not separated. The rotor vane slots did not appear to be distorted. The individual vanes were intact and were located within their respective rotor guide slots.
MEDICAL AND PATHOLOGICAL INFORMATION
An autopsy of the pilot was conducted at Grand View Hospital in Ironwood, Michigan, on December 29, 2004.
The FAA Civil Aerospace Medical Institute toxicology report for the pilot was negative for all substances tested.
TESTS AND RESEARCH
The left engine propeller governor was bench tested by the propeller manufacturer under supervision of the NTSB. The control arm was bent and exhibited some roughness when actuated. The feather speed was measured at 1,640 revolutions per minute (RPM). The design specification required 1,675 - 1,700 RPM. The maximum speed was measured at 2,348 RPM. The specification required 2,320 - 2,340 RPM. The internal leakage was measured at 10 ounces per minute (oz/min). The specification required internal leakage to be less than 8 oz/min. The relief pressure and pump capacity were within specifications. The manufacturer stated that the out-of-specification conditions noted above would not have precluded normal operation.
The left propeller assembly was torn down. The blade pitch change mechanism was intact. The blades were initially at low pitch and moved to the feather position when the low pitch stop screw was removed. The start lock mechanism was intact. Lack of damage to the start lock mechanism was consistent with the lock not being engaged at the time of impact.
According to the manufacturer, the quantity of oil drained from the propeller hub cylinder would not have been present had the propeller been feathered at the time of impact. Additionally, no witness marks due to the pitch change mechanism were observed at or near the feather position.
The No. 4 cylinder from the left engine was examined at the NTSB Materials Laboratory. The cylinder head on the left engine had separated between the second and third cooling fins leaving a portion of the head still attached to the barrel. The separation had exposed two barrel threads.
The fracture face on the cylinder head still attached to the barrel exhibited curved crack arrest marks consistent with fatigue cracking that originated at the head-to-barrel interface. The fatigue cracking had penetrated completely through the head over more than one-half of the cylinder head's circumference before final fracture. The fracture initiated in the root of the thread.
Dimensional and visual inspection of the threads revealed no significant variation from specification requirements.
The No. 4 cylinder on the left engine was part number EC75362CN and marked as serial number 02. Engine Components Inc. (ECi) overhauled the cylinder for installation on the engine. Documents indicated that the cylinder overhaul included CermiNil plating of the barrel. Weld repair and an improved fatigue resistance process was applied to the cylinder head.
CermiNil plating is a nickel-based coating infused with silicon carbide particles intended to increase wear resistance, speed break-in and improve oil adhesion. The heat-treating process is intended to restore the cylinder heads to the appropriate solution heat-treated and over-aged condition for continued use.
Information provided by the aircraft manufacturer stated that the landing gear and flaps were hydraulically operated. On the accident aircraft, hydraulic pressure was supplied by one engine-driven pump mounted on the left engine. In this installation the emergency hydraulic hand pump must be used to supply hydraulic pressure in the event that the left engine was inoperative. In addition, a carbon dioxide (CO2) powered emergency landing gear extension system was also installed. Emergency landing gear extension may be accomplished by using the emergency hydraulic hand pump or the CO2 system.
Emergency procedures listed in the aircraft information manual state that in the event of a single-engine landing the pilot should determine if hydraulic pressure is available by moving the landing gear control handle to the UP position with the landing gear retracted. If hydraulic pressure is available, the handle will return to the NEUTRAL position. The manual noted that this check should be conducted prior to entering the traffic pattern to allow sufficient time to use the emergency gear extension procedure.
The FAA Airplane Flying Handbook (FAA-H-8083-3) provided information on multi-engine flight characteristics. Concerning flight with one engine inoperative, the handbook notes that flight at any airspeed less than the minimum control airspeed (Vmc) will not provide sufficient airflow over the rudder to overcome the asymmetrical yawing forces caused by takeoff power applied on one engine and a windmilling propeller on the other engine.
The handbook states: "When one engine fails, the pilot must overcome the asymmetrical thrust (except on airplanes with centerline thrust) created by the operating engine by setting up a counteracting moment with the rudder. When the rudder is fully deflected, its yawing power will depend on the velocity of airflow across the rudder, which in turn is dependent on the airspeed. As the airplane decelerates, it will reach a speed below which the rudder moment will no longer balance the thrust moment and directional control will be lost."
The minimum control airspeed is dependent upon several factors. One of these factors is the bank angle of the airplane. The handbook noted that banking toward the operating engine decreases the minimum control airspeed, whereas banking toward the inoperative engine increases the minimum control airspeed. The minimum control airspeed increases approximately 3 knots per degree of bank angle.
The airframe, right engine, and right propeller were released at the conclusion of the on-scene investigation. The left engine and left propeller assembly were retained for further examination and subsequently released on July 28, 2005. A representative of the insurance company acknowledged release of the aircraft.
The Federal Aviation Administration, New Piper Aircraft, Lycoming Engines, Hartzell Propeller Inc. and Engine Components Inc. were parties to the investigation.