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On December 29, 2000, at approximately 0730 central standard time, a Bell 407 helicopter, N407MM, registered to The Fifth Third Leasing Co., of Cincinnati, Ohio, and operated by Petroleum Helicopters Incorporated (PHI), of Lafayette, Louisiana, was destroyed when it impacted offshore waters in an uncontrolled descent in the Gulf of Mexico, approximately 6 miles south of Gilchrist, Texas. The commercial pilot, who was the sole occupant, sustained fatal injuries. Visual meteorological conditions prevailed and a company flight plan was filed for the Title 14 Code of Federal Regulations Part 91 repositioning flight. The flight departed Cameron, Louisiana, at 0713, and was en route to Galveston, Texas, with an estimated time of arrival (ETA) of 0750 .
According to the operator's refueling logs, the aircraft was fueled to capacity (863 pounds) with Jet-A fuel prior to departing Cameron. The operator's daily communications log and audio tape revealed that the pilot made a routine position report at 0728 on the PHI en route frequency. The pilot reported that he was at 29 degrees, 38 minutes North latitude, and 93 degrees, 57 minutes West longitude. This was the last known contact with the helicopter. At 0751, repeated attempts to contact N407MM by the PHI Sabine Communications Center were made, but to no avail. At 0810, N407MM was reported overdue and the company search plan was activated. For the remainder of the day, search helicopters flew the presumed flight path of N407MM and the adjacent offshore areas with no sightings of the helicopter, pilot, or debris. There were no reported witnesses to the accident.
The pilot was hired by PHI on February 10, 2000, and held a commercial pilot certificate with rotorcraft-helicopter and instrument ratings. His total reported flight time, as of 29 December, 2000, was 1,984 hours, all of which were in light helicopters. His total flight time in the Bell 407 was 58 hours. He had flown 100 operational flight hours in single-engine helicopters in the 90 days prior to the accident.
PHI purchased the helicopter, serial number (S/N) 53060, as a new aircraft in March 1996. It had a total airframe time of 4,429.35 hours at the time of the accident. The helicopter was equipped with an Allison 250C47B turbo shaft engine (S/N CAE847289), and it had a total time of 1,770 hours. According to maintenance records provided by the operator, the last 25-hour inspection was completed on December, 26, 2000 at 4,421.5 hours, and the last 50-hour inspection was completed on December 19, 2000 at 4,407 hours.
On December 24, 2000, at an aircraft total time of 4,418.4 hours, the tail rotor hub and blade assembly was removed due to worn bearings in the pitch change links and hub. A replacement tail rotor hub and blade assembly, S/N 53209, was installed. The hub and yoke components of this assembly had a total time of 37 hours at the time of installation and both blades (which were removed from the previous yoke assembly) had accumulated total times of 2,433.5 hours respectively. On December 28, 2000, at 4,429 hours aircraft total time, the tail rotor mast retaining nut torque check, required 10 to 25 hours after installation of the assembly as per the manufacturer’s maintenance manual, was performed. The entry in the logbook for the torque check stated that there was "no movement" of the retaining nut during the check, and the technician stated in an interview that there was no disassembly required to perform the torque check. A torque check was also done on the #1 and #2 tail rotor drive shaft disc couplings. This torque check was required as per the manufacturer's maintenance manual since the #1 and #2 drive shafts were removed and re-installed during the previous 100-hour inspection.
According to records provided by the FAA, the manufacturer, and the operator, N407MM was in compliance with all applicable airworthiness directives pertinent to the Bell model 407 helicopter. Additionally, all required modifications for the helicopter to operate at 140 knots Vne were verified to have been installed. These modifications included a newly designed tail rotor hub, and an airspeed activated tail rotor pedal stop.
According to PHI operating procedures, the helicopter's transponder should have been turned on with a beacon code of 4345. A military radar source provided radar contact information about a target using code 4345 which was traveling at about 1,000 feet MSL. According to PHI, the target's path corresponded to the planned flight path of N407MM. The first radar contact was at 0725:51 at 29 degrees, 45'33" North latitude, and 93 degrees, 52'04" West longitude. The last radar contact was at 0726:15 at 29 degrees, 40'14" North latitude, and 93 degrees, 51'49" West longitude. No distress calls were monitored on offshore radio frequencies.
WRECKAGE AND IMPACT INFORMATION
The main aircraft wreckage was located in 32 feet of water on January 2, 2000, at grid coordinates 29:25 N longitude 94:32 W latitude, and transported to the PHI facility in Lafayette, Louisiana. These recovered portions included the main fuselage with part of the tail boom (forward section attached to the fuselage), the main rotor drive system and blades, the engine, and the landing gear and skid mounted float assembly. On January 6, 2001, another section of the tail boom (middle section) was recovered after being located floating in the Gulf of Mexico about 75 miles southwest of the main wreckage, and was transported to the PHI facility. On March 28, 2001, the aft portion of the tail boom with the vertical fin attached washed ashore near Sabine Pass, Texas, and was recovered. The tail rotor gearbox housing and gears, tail rotor mast, tail rotor hub and blade assembly, and the aft section of the tail rotor drive shaft were missing. As of the date of this report, these parts have not been recovered.
Various examinations of the recovered components were conducted under the supervision of the NTSB investigator-in-charge at the PHI facility, Lafayette, Louisiana, on January 3-10, 2001 and at the Bell materials lab, Hurst, Texas, from February 28 to March 16, 2001 and April 2-4, 2001. The following are summaries of these examinations.
The engine was removed from the wreckage and secured to a work stand for examination. The Electronic Control Unit (ECU) was found intact with harnesses connected. The unit exhibited only superficial scarring damage to the cover. During handling, it was determined that the unit contained water. Per consultations with a representative of Chandler Evans Controls (CECO) Division, it was decided to open the unit, drain and flush it with distilled water, and send it to CECO, West Hartford, Connecticut, for download of data. The ECU download did not record any faults or incidents.
The engine gearbox exhibited extensive corrosion damage attributed to salt water immersion. The left side of the engine exhibited extensive impact damage from the diffuser scroll aft to the outer combustion case, which had several indentations. Two impeller blades (about 180 degrees apart) were bent approximately 30-degrees, opposite direction of rotation and the liner exhibited heavy rotational scoring at the impeller blade rotational path. The Hydro-Mechanical Unit (HMU) was found intact and approximately 12 inches of the linkage was attached to the control quadrant. The control lever operated normally through its full range. All fluid, oil, fuel, and air lines were found attached and their fittings were hand tight. The airframe fuel filter was intact, clean, and fuel was present in the filter housing. Approximately 20 cc's of fuel was present in the fuel nozzle line. Both engine chip detectors, upper and lower, were clean of metallic particles.
Much of the airframe was still wet, and mud covered much of the main rotor head when initially examined. Most of the cabin structure (especially the left side) exhibited evidence of hydraulic compression along its longitudinal axis. The nose section was connected to the center fuselage section by tangled wires and the roof section was connected by the engine throttle cable and engine bleed air line, both of which were located in the vertical tunnel (broom closet). The center and right hand nose sections were twisted and broken apart and the nose section was broken away from the forward floor structure. The pilot's seat pan exhibited no downward deformation and the seat back, including the inertial reel was missing. The upper circuit breaker panel was indented and deformed upward and the deformed area had a piece of brownish black hair embedded within the indentation. The rear portion of the fuselage (transition area from the rear bulkhead to the tail boom) was separated from the forward fuselage generally along a vertical line just aft of the fuel bladder cavity. About a 33-inch section of the tail boom remained attached to the rear bulkhead. The roof structure that attached to the engine was missing and the top portion of the rear fuselage from the forward part of the oil cooler was intact rearward to the tail boom attach point. The baggage compartment door was missing.
The tail boom was fractured/separated in two places along its longitudinal axis. The forward most fractures consisted of multiple vertical and horizontal fractures, starting at about 33 inches aft of the tail boom to fuselage attachment point and ending at about 56 inches aft of the attachment point. The fractures on each side of the separation did not match and some tail boom skin was missing between the fractures. Some deformed skin pieces on both sides of the separation were displaced to the right. One piece, that was curled, had what appeared to be white paint smeared onto its surface. The color of the white paint smear was similar to the color of the top side of the main rotor blades. The aft most fractures on the tail boom were between tail boom stations 148 and 154. Within this area, 4 impact marks were observed on the left side outer surface of the boom. The marks cut through the paint layer and appeared to be consistent with tail rotor blade tip edge contact. All four marks were slanted across the boom at slightly different angles.
Main Rotor Hub and Blades
The main rotor hub assembly was first examined at PHI during the initial portion of the investigation. During the first week of March the hub assembly was disassembled and inspected at Bell. No pre-impact discrepancies were observed during any of the examinations. All fractures exhibited evidence overload. According to the manufacturer, the main rotor yoke exhibited damage to each flexure indicative of being powered at impact. Each of the four flexures were splintered and there were separations along their respective bond lines. All flapping up stops and downs tops exhibited evidence of contact. Additionally, one of the pitch horns was fractured near its retention bolts.
All four main rotor blades were reportedly attached to the hub when first recovered. They were detached from the hub for transport from the recovery area. The blade retention bolts were all intact and did not exhibit any anomalies. Two of the blades were bent downward with large sections of their respective after bodies missing. The remaining 2 blades were relatively straight and intact with some scrapes and dents.
Rotating Controls/Main Rotor Fixed Controls
No pre-impact failures were identified within the main rotor rotating controls and main rotor fixed controls. Both inner and outer swash plate rings were intact and the pitch change links exhibited evidence of impact damage. Numerous breaks were observed in the main rotor fixed controls. All breaks and disconnects showed evidence of overload.
The four hydraulic actuators were marked of their respective piston positions and removed from the aircraft on January 5, 2001. During functional tests conducted at PHI on January 8, 2001, the center, right, and tail rotor servos functioned normally. The left servo was reported to be slightly slow when retracted. Hydraulic fluid from each actuator was captured in a bottle and sent to HR Textron for examination. Additionally, the actuators were subsequently tested and inspected at HR Textron and a report issued. The hydraulic pump was inspected and run at Bell Helicopter under the direction of the NTSB and no anomalies were observed. No pre-impact discrepancies were observed in the hydraulics system.
Prior to recovery, the wreckage was reported to be situated upside down on the bottom of the gulf. The wreckage was recovered from the water with the landing gear coming out first and upside down. The landing gear was essentially intact when recovered and the skid mounted float bags were somewhat still folded in their respective covers. Both float inflation bottles were charged. The fore and aft cross tubes were both bent inwards on both the left and right sides.
No pre-impact discrepancies were observed in the fuel system. All fuel bladders were recovered but were essentially found shredded. Fuel was present in the fuel line connected to the engine fuel nozzle.
Main Rotor Drive System
The main transmission, freewheeling unit and main drive shaft were disassembled and examined at Bell beginning on February 28, 2001. No pre-impact discrepancies or anomalies were found during the disassembly and examinations.
Tail Rotor Drive System
Evidence of shaft to tail boom contact was present underneath the aft end of the #6 shaft Thomas coupling. Indentations matching the coupling disk pack were observed on the top surface of the tail boom. Also, there was evidence of contact between the #6 coupling adapter and a rivet on top of the tail boom. The #1 steel tail rotor drive shaft was fractured approximately 6 inches from the forward end. The fracture was consistent with torsion overload. The aft section of the shaft was not recovered. The #4 shaft was found severed, with no local evidence of flailing. X-ray examination of the #6 hangar bearing revealed that the balls inside the cage were bunched up to one side. A portion of the aluminum splined adapter from the #7 shaft remained within the #6 coupling adapter. There was evidence of heat on the #7 splined adapter. Smeared aluminum from the #7 splined adapter was found deposited on the inner diameter of the inner race. A curved impression was found indented on the aft surface of the #6 coupling adapter. This impression matched the forward face of the #6 hanger bearing housing.
Tail Rotor Control System
Segments of the tail rotor flight control linkages from the anti-torque pedals rearward were found damaged and separated resultant from deformation of the airframe at impact. A tail rotor pitch control rod end bearing was found missing from the rod assembly located in the vicinity of the tail rotor gearbox support. The rod end jam nut was still intact on the rod assembly. This rod end bearing is a tail rotor pitch control rigging point and is normally constrained between two lever assemblies and attached by a bolt and nut and secured by a cotter. In addition to the missing rod end, the bolt, nut, and cotter pin used to secure the rod end to the lever assemblies were missing. Of the other three tail rotor rigging points, two exhibited loose jam nuts.
Examination of the rod revealed that it was bent slightly upwards. The rod run out was approximately 0.04-inches. Scrape marks were located on the top of the rod approximately in the center. The rod assembly in normal operation is situated crosswise underneath the control tube.
The amount of corrosive deposits on the rod threads under the missing rod end (outboard end) was similar to that of the exposed threads on the inboard side of the loose jam nut. No deposits were observed on the rod threads where the loose jam nut had been located. Additionally, the rod threads were not "rounded" or worn and did not exhibit mechanical damage. Compression and tension tests of two new rod assemblies were done to examine the possibility that the missing rod end fractured through its threads without damaging the rod's threads. In neither test did the rod end fracture through the threads.
The manufacturer stated that a tail rotor linkage disconnect in steady state cruise, may result in no immediate effect. However, as the main rotor and tail rotor torque changes, the aircraft attitude would change.
MEDICAL AND PATHOLOGICAL INFORMATION
The 38 year old pilot was recovered on February 5, 2001, in shallow coastal water near Galveston, Texas. An autopsy and toxicological tests were conducted. The Civil Aero Medical Institute (CAMI) in Oklahoma City, Oklahoma, performed these tests. No drugs were detected but 42 (mg/dl, mg/hg) ethanol was detected in muscle tissue. CAMI attributed this to post mortem putrefaction. The County of Galveston Medical Examiners Office, Texas City, Texas, performed the autopsy on February 6, 2001. The primary cause of death was "craniofacial trauma." Contributing to the cause of death was "acute occlusion to left anterior descending coronary artery of the heart."
TEST AND RESEARCH
The Pedal Restrictor Control Unit (PRCU) was examined at Bell Helicopter Textron Canada in Mirabel, Quebec, under the direction of TSB of Canada. Upon initial testing, the unit did not respond to airspeed pressure inputs, although the circuit board was verified operational, it was found that the internal airspeed detection pressure sensor was not functioning. In later testing, the pressure sensor was replaced and the unit functioned correctly. The manufacturer stated that the pressure sensor was likely damaged from salt water immersion or impact. Additionally, the manufacturer stated that the overall wreckage damage indicates that there was little forward airspeed at impact. Since it was observed that the PRCU cam stop did not exhibit damage on its surface, it was most likely retracted at impact
The wreckage was released to the owner.