NTSB Factual Report

HISTORY OF FLIGHT

On September 29, 2001, at 1423 Hawaiian standard time, a Bell 206B single engine helicopter, N206KS, experienced a loss of engine power and made a forced landing about 5 miles south of Hilo, Hawaii. The helicopter was registered to and operated by K&S Helicopters, Inc., Hilo, as an on-demand sightseeing flight under the provisions of 14 CFR Part 135. The commercial pilot and two passengers were not injured; however, the remaining two passengers sustained minor injuries. The helicopter was destroyed in the collision sequence and post-accident fire. The local flight departed Hilo about 1420. Visual meteorological conditions prevailed, and a company VFR (visual flight rules) flight plan had been filed. The accident site was located at 19 degrees 38.425 minutes north latitude and 152 degrees 02.088 minutes west longitude.

The Federal Aviation Administration (FAA) inspector, who traveled to the accident site, interviewed the pilot and examined the wreckage on scene. The pilot reported that she heard a bang followed by illumination of the engine out warning light, and the engine began smoking. She made a mayday call, shut off the fuel valve, performed an autorotation, and landed in high grass. The pilot thought that the grass was 2-to 3-feet tall; however, the grass was actually 6 feet tall. She and the passengers evacuated the helicopter, and a couple of the passengers received minor burns from molten plastic or metal that was dripping from the engine deck, which was on fire. The pilot and passengers experienced difficulties as they attempted to gain distance from the helicopter in the tall grass; however, they all managed to obtain a safe distance from the burning helicopter.

PERSONNEL INFORMATION

The pilot held commercial helicopter and single engine airplane pilot certificates with instrument airplane and helicopter ratings. She also held an instructor rating for helicopters. She was issued a second-class medical certificate on November 6, 2000, with a limitation that she wear corrective lenses. The pilot obtained her last biennial flight review on June 29, 2001, in a Bell 206 helicopter. According to the Pilot/Operator Aircraft Accident Report (NTSB form 6120.1/2), the pilot had accumulated 2,955 hours of total aircraft flight time, of which 2,795 hours were obtained in rotorcraft, and 1,255 hours were obtained in the same make and model as the accident helicopter.

AIRCRAFT INFORMATION

The single engine helicopter utilized a 420 shaft horsepower Allison 250-C20B turboshaft engine (serial number CAE831174). The engine is a two-shaft turboshaft engine with a combination compressor (6-stage axial compressor attached to a 1-stage centrifugal compressor). It has a reverse-flow annular combustor, a two-stage high-pressure turbine (also referred to as the gas producer turbine or the N1 turbine), and a two-stage low-pressure turbine (also referred to as the power turbine or the N2 turbine). The gas path of the Allison 250-C20B engine flows into the inlet, through the compressor, and then is taken to the combustor at the rear of the engine through two external transfer tubes. The gases then turn (180 degrees) forward and flow through the gas producer turbine and the power turbine. Finally, the gasses are directed out the exhaust ducting.

The flow of hot gases coming from the combustor is such that the first stage turbine nozzle shield is the first thing encountered, behind which, is the No. 8 bearing sump nut. The No. 8 bearing sump nut resembles a cover or cap; not what one would picture as a typical nut. Behind the bearing sump nut is the No. 8 bearing thrust nut, No. 8 bearing, and No. 8 bearing thrust plate.

A review of the helicopter's maintenance records revealed the helicopter underwent its last annual inspection on July 1, 2001, at an engine total time of 5,190.0 hours, and underwent its last 100-hour inspection on September 20, 2001, at an engine total time of 5,441.5 hours. A review of the daily flight logs revealed that during the period of September 18 through September 27, the helicopter had been flown for 11.1 hours, and 12.5 quarts of oil were added. On September 27, the helicopter flew for 0.6 hours with 1 quart of oil consumed. The engine operation and maintenance manual prescribes a maximum oil consumption of 1 quart of oil per flight hour.

The following is the oil addition information and flight times (at the beginning of each day) taken from the flight logs:

DATE: TOTAL TIME AT START OF DAY: OIL ADDED:

September 18, 2001 5,438.2 Hours 2 quarts

September 22, 2001 5,444.0 Hours 1 quart

September 24, 2001 5,447.5 Hours 5 quarts

September 25, 2001 5451.0 Hours 3.5 quarts

September 26, 2001 5453.6 Hours 0.5 quarts

September 27, 2001 5454.7 Hours 1 quart

On September 8, 2001, at a helicopter total time of 5,426.1 hours, a flight log maintenance entry indicated that the turbine (serial number CAT 37394) was removed for overhaul and No. 1 and No. 2 wheel change. Another turbine (serial number CAT 36018) was installed at a turbine total time of 4,410.1 hours. That same entry indicated that the installed No. 3 and No. 4 turbine wheels had accumulated 74.0 hours since new, and the No. 1 and No. 2 turbine wheels had accumulated 0.0 hours since new. After this turbine work was complete, the helicopter was flown on a 0.2-hour check flight and was signed off with no discrepancies noted. The helicopter accumulated approximately 30 hours after this maintenance work was performed.

On the September 18, 2001, flight log (helicopter total time of 5,441.5 hours), a maintenance entry, dated September 20, 2001, indicated, "engine seem to consume more oil than normal (not smoking). Removed turbine, replaced No. 5 carbon seal, installed turbine." This work took place approximately 14 hours prior to the accident.

Review of the turbine assembly inspection - maintenance - overhaul records for serial number CAT 36018, revealed that maintenance personnel replaced the No. 1 and No. 2 turbine wheels with new wheels and a new tie bolt (part number E6898784, serial number A63-128) on September 1, 2001. On September 16, 2001, they replaced the "No. 3 and No. 4 turbine wheel package and carbon seal" for "smoking and high operating temp for trouble shooting." On September 27, 2001, they "replaced sump cover for trouble shooting oil consumption 1 quart."

The FAA inspector, who traveled to the accident site, spoke with the mechanic who performed the aforementioned work. The inspector asked if there had been any recent maintenance performed on the accident helicopter. The mechanic reported on the replacement of the turbine wheels, and approximately 4 hours prior to the accident, he replaced a turbine oil seal because of smoking and high oil consumption. The FAA inspector was told that a test flight was conducted after the oil seal replacement, a relocation flight was conducted, and the helicopter was returned to service. The helicopter then underwent a tour flight earlier on the day of the accident, and was on its second tour flight since the seal replacement when the accident occurred.

On the morning of the accident, the flight log listed the helicopter total time as 5,455.3 hours.

WRECKAGE AND IMPACT INFORMATION

The FAA inspector observed that the tail boom separated from the fuselage at a point just aft of the baggage compartment. The rearmost portion of the tail boom, aft of the horizontal stabilizer, separated from the forward portion and was within 10 feet of the main wreckage; the middle section was about 50 yards away. Visual external examination of the thermally damaged engine disclosed that the first stage turbine wheel was missing. The FAA inspector found a section of what he thought to be turbine wheel rim in the wreckage, in an area that was formerly the baggage compartment.

The engine and the recovered turbine wheel section were removed from the helicopter and were shipped to Rolls-Royce Corporation for further examination.

TESTS AND RESEARCH

On December 13, 2001, technicians examined the engine at the Rolls-Royce facility under the auspices of an FAA inspector. Both Rolls-Royce personnel and the FAA inspector made the following observations.

The engine hardware was severely fire damaged and the accessory gearbox was consumed by fire. The examiners noted failures of the accessory gears that were shipped with the engine. They noted no damage other than fire damage to the compressor section, and therefore they did not disassemble the compressor section. They conducted a four PSIA pressure/leak check was conducted on the No. 8 sump, and noted a "major leak" around the nut.

The second stage turbine wheel had 360-degree circumferential damage to the blade leading edges. The third stage turbine wheel was intact and had material missing from the leading edges of some of its blades at the midspan area. The fourth stage turbine wheel remained intact.

The first and second stage turbine nozzles were "severely damaged" with missing vanes and outer ring and diaphragm damage. The third stage nozzle was intact and some of its vanes sustained denting damage. The fourth stage nozzle was intact and displayed no damage.

The turbine shafting remained intact; however, the gas producer tie bolt was fractured at the first stage wheel area. The gas producer tie bolt spanner nut (also known as the thrust nut) was not installed when the No. 8 sump nut was removed. The thrust nut was free in the sump. This thrust nut is a staked (safety notched) nut with reverse threading, and according to Rolls-Royce, cannot back off during engine operation. The No. 8 bearing outer race fractured into multiple pieces. The No. 6 and No. 7 bearings remained intact, but were dry.

The FAA inspector, who oversaw the engine examination at the Rolls-Royce facility, sent the following items to the Safety Board Materials Laboratory, Washington D.C. for further examination:

- No. 8 bearing

- No. 8 sump nut

- No. 8 bearing oil supply line

- No. 8 bearing tie bolt spanner nut (thrust nut)

- No. 8 pressure oil tube (otherwise referred to as the "J" tube)

- No. 8 bearing thrust plate

- Bearing No. 6 & No. 7 thimble oil screens

- Aft end of the gas producer tie bolt

- Gas producer turbine aft rotating seal

- first stage turbine wheel, shaft piece

- first stage wheel rim chunk

- first stage nozzle

- second stage turbine wheel

A Safety Board Material's Laboratory specialist prepared a factual report. The following paragraphs contain pertinent parts of the report.

First Stage Nozzle Examination

The nozzle was missing portions of its outer flange and displayed scored surfaces, cracked vanes, and a crack at one of the cooling holes. The specialist examined cooling holes after radially cutting a few samples out of the nozzle. One cooling hole displayed a deposit in and around the hole, and a crack emanating from the inner surface of the hole. All cooling hole areas displayed a pale green deposit on the rear side of the nozzle and displayed a patch of discoloration with a cream colored deposit on the forward face of the nozzle. The specialist dissected the cooling holes and examined them under a Scanning Electron Microscope (SEM). The one crack noted in one of the cooling holes only extended through the surface deposits. Examination of the cooling holes via an Energy Dispersive Spectrometer (EDS), resulted in a spectrum displaying major peaks of nickel, chromium, and iron, with small peaks of sulfur, silicon, and phosphorus. A closer examination of the sectioned nozzle revealed that the cooling hole section of the nozzle was manufactured by brazing a curved section with the cooling holes to an "L" shaped piece. Some of the braze material had flowed along the curved section surface, past the cooling holes, suggesting that the braze material was the light green deposit noted along the cooling holes.

Second Stage Turbine Wheel Examination

Examination of the second stage turbine wheel revealed no significant damage to the blades. The tie bolt head and a portion of the tie bolt shank were retained in the turbine wheel's central bore, and there were circumferentially oriented score marks on the rear face of the disc. The driving dogs, which provide the engagement to the first stage turbine wheel, were damaged.

Tie Bolt Examination

The fracture face of the tie bolt displayed mechanical damage that obliterated its fracture features. The fact that the fractured tie bolt was retained in the second stage turbine wheel central bore indicates it had been subjected to severe bending. The end of the tie bolt still had the nut engaged on its thread. The shank of the tie bolt was bent and displayed an arced impression.

No. 8 Bearing Sump Nut and Thrust Nut Examination

The sump nut displayed a puncture mark from the inside out, or from the front of the engine to the rear. Closer examination of the inside (forward facing) surface of the sump nut revealed shiny marks describing portions of a circle around the penetration hole.

The thrust nut retains the No. 8 bearing and seal on the shaft portion of the first stage turbine wheel. It is prevented from rotating on the shaft by deformation of a thin cylindrical portion of the nut into a groove on the shaft (previously described by Rolls-Royce as staking or safety notching). The thrust nut was a uniform gray color with shiny sections on the thin cylindrical portion, which faces the sump nut. The diameter of this thin cylindrical portion approximately matched the diameter of the circular imprint found on the inside face of the sump nut. The thin cylindrical portion also displayed an area where it had been deformed radially inwards, and two areas where the material had been milled out (possibly from previous deformations). After conducting the examination of the first stage turbine wheel shaft, the specialist gently screwed the thrust nut onto the shaft to reveal that rotation of the nut halted when the inward deformation on the nut reached the end of the shaft.

No. 8 Bearing Examination

The No. 8 bearing components consisted of the inner race, the cage, two flat metal rings, the outer race pieces, and the ball bearings. The inner race was intact and uniformly discolored. Examination of the inner race revealed eleven marks on the ball groove ranging from light marks across the groove surface to smaller localized marks consistent with being impacted by a hard sphere such as a bearing ball (a process commonly known as "Brinelling"). Apart from the small marks, the Brinell damage, and the onset of corrosion, the ball groove did not display any other surface damage.

Examination of the bearing cage revealed mechanical damage in the form of scratches and smearing at opposing locations (scratches on opposite sides of the cage from each other).

The 11 ball bearings were intact and evenly discolored. Seven of the balls displayed individual and combinations of features described as a line, an impression, and a flat circle on their surface.

The outer race fractured into eight pieces. Original bearing surfaces were a uniform blue/gray color. The fracture faces, with the exception of two, were uniformly blue/gray in color, oriented at varying angles, and displayed features consistent with an overload event. The other two fractures' surfaces were oriented along a manufactured slot, which allowed the spreading apart of the outer race to facilitate the insertion of the balls. The ball groove on the outer race displayed eleven evenly spaced marks of varying clarity on its surface. A group of four evenly spaced indentations were along one edge, and they were adjacent to four of the marks. Apart from the marks and the onset of corrosion, the ball grooves did not display any other surface damage.

No. 8 Bearing Thrust Plate Examination

Examination of the No. 8 bearing thrust plate, revealed it was evenly discolored and displayed the onset of corrosion. A small portion of the inside edge had been deformed slightly out of plane.

No. 8 Bearing Oil Supply Tube

The No. 8 bearing oil supply tube displayed no fractures, and air passed through it with no indicat

NTSB Probable Cause

The failure of the mechanic to secure the No. 8 bearing thrust nut during maintenance work, which resulted in the total failure of the first stage turbine wheel shaft. Also causal was his misjudgment of his height during the flare resulting in a hard landing. A contributing factor was the high vegetation in the forced landing field, which resulted in a hard landing.