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N1109V accident description

Hawaii map... Hawaii list
Crash location 21.950000°N, 159.550000°W
Nearest city Hanapepe, HI
22.020489°N, 159.617765°W
6.5 miles away
Tail number N1109V
Accident date 03 Jun 2001
Aircraft type Hughes 369D
Additional details: None

NTSB Factual Report


On June 3, 2001, approximately 1640 Hawaiian standard time, a Hughes 369D helicopter, N1109V, was substantially damaged during a forced landing following a loss of engine power near Hanapepe, on the island of Kauai, Hawaii. The helicopter was registered to and being operated by Smokey Mountain Helicopters, Inc., under the provisions of 14 CFR Part 135. The commercial pilot and three passengers were not injured. The medical evacuation flight departed Port Allen Airport, Hanapepe, about 1620. The flight was en route to the Hanakoa Valley, which is along the Na Pali Coast on the northwest side of Kauai. Visual meteorological conditions prevailed, and a company flight plan had been filed.

According to the pilot's written statement, upon receiving the mission call from search and rescue personnel, he conducted a preflight inspection and refueled the helicopter. At 1620, the pilot activated a company flight plan via the radio. He took off from Port Allen Airport and flew to Kukui Grove landing area to pickup the rescue team.

On the ground at Kukui Grove, the search and rescue personnel briefed the pilot about the rescue mission, which was to rescue an injured person. Three firefighters and a rescue basket boarded the helicopter, and the helicopter departed for Hanakoa Valley.

Approximately 5 minutes after departing the Kukui Grove landing area, the engine out warning system (light and horn) came on. The pilot lowered the collective lever approximately 1/2 of its travel, corrected for the left yaw with the right antitorque pedal, and verified that the throttle was full open. The pilot attempted manual control of the governor without any result. The pilot elected to perform an autorotation emergency landing.

The northwest side of the Na Pali Coast consists of steep, rugged, and thickly vegetated terrain rising out from the ocean. The pilot could not find a suitable landing spot in the "entire area" and decided to place the helicopter in "the lower portion of a small creek to avoid possible rollover after touchdown." As the helicopter neared the treetops, the pilot began to flare. The pilot reported that the tail rotor stinger touched the ground first. The pilot then leveled the helicopter and pulled all the way up on the collective. The helicopter touched down in a "level attitude and no forward motion."

The helicopter came to rest upright, and the pilot shutoff the fuel control valve. The pilot reported that the main rotor blades came to a complete stop within a few seconds after touchdown, and he then gave the firefighters the clearance to exit the helicopter. The hot exhaust stack was touching some ferns and ignited a fire, which was extinguished by one of the firefighters with the cockpit fire extinguisher.

The helicopter sustained substantial damage to its tail boom, which was bent and twisted forward approximately 110 degrees just aft of the helicopter registration number. The tail rotor blades and main rotor blades remained intact and attached to their hubs; however, the blades did sustain some bending damage. The bottom side of the fuselage displayed a crack from which fuel was leaking.


The engine, a Rolls-Royce Allison 250-C20B (serial number CEA 833407), was a two-shaft turboshaft engine. It had a combination compressor (6-stage axial and 1-stage centrifugal). It had a reverse-flow annular combustor, a two-stage high-pressure turbine (also known as the gas producer turbine or N1 turbine) that drives the compressor section. It also had a two-stage low-pressure turbine (also called the power turbine or N2 turbine) that drives the power output to the helicopter transmission. Airflow through the engine enters the inlet, and then flows into the compressor section. The air then enters two external air transfer tubes, which duct the compressed air to the combustor, located at the rear of the engine. The gas airflow then turns 180 degrees forward and passes through the two-stage gas producer turbine (N1), and the two-stage power turbine (N2). Finally, the gases are directed out of the exhaust duct and upward through the outlets.


The helicopter was transported to the Port Allen Airport where the helicopter and engine were examined on June 6-7, 2001, by a Federal Aviation Administration (FAA) inspector, a representative from Boeing, and one from Rolls-Royce Corporation. According to a report provided by Rolls-Royce, the engine fuel lines passed a vacuum leak check, and they found all lines hooked up and tight. No fuel or oil filters exhibited any contamination, and the oil chip detectors were clean. The engine's N2 rotor system was very tight and hard to turn manually. The engine's N1 rotor system could not be turned by hand. Investigators shipped the engine to a Rolls-Royce Allison authorized service center in Scottsdale, Arizona, for further inspection.

On July 2-3, 2001, the National Transportation Safety Board investigator-in-charge (IIC) and representatives from Boeing and Rolls-Royce reexamined the engine. Removal of the accessory section revealed no anomalies. Upon removal of the turbine section, investigators noted that the compressor sections rotated freely. The turbine to compressor coupling adapter displayed minor coking. They removed the thermocouple harness and three of the four elements were broken, and all four thermocouple housings exhibited heavy rub damage.

The No. 8 bearing is the aft most bearing in the engine, and it is the first bearing encountering hot exhaust air from the combustor. Technicians applied an air pressure of 50 psi to the No. 8 bearing sump and observed no leaks. The No. 8 gearing sump cover, bearing sump cavity, and bearing displayed heavy coking. In addition, the gas producer turbine's No. 1 turbine nozzle shield displayed heavy oil coking. Removal of the gas producer turbine revealed that approximately half of the gas producer's 1st stage turbine blade airfoils were reduced in dimension. Approximately half of the gas producer turbine's 2nd stage turbine blade airfoils were also damaged. The gas producer turbine nozzle also displayed damage to approximately 7 stator vanes, and the outside diameter of the nozzle displayed two cracks. Both the No. 6 and No. 7 bearings rotated roughly due to oil coking; however, an oil flow check of the No. 6 and No. 7 bearing oil supply tube did not reveal any anomalies.

The power turbine section was not disassembled during the July 2nd/3rd disassembly. Investigators shipped the engine to Rolls-Royce's Indianapolis facility for further examination.

On July 27, 2001, Rolls-Royce technicians reexamined the gas producer section of the engine under the supervision of an FAA inspector. According to a Metallurgical Investigation Report, the gas producer turbine's 1st and 2nd stage turbine wheels would not rotate in the 2nd stage turbine nozzle assembly. A close examination of the 1st stage turbine wheel revealed all of the airfoils were broken off 0.2 - 0.4 inches above the wheel rim at the leading edge and 0.1 - 0.3 inches above the rim at the trailing edge. One airfoil was broken off notably closer to the rim than the rest, and according to Rolls-Royce, "was likely the primary failed airfoil." Technicians identified that airfoil and referenced it as airfoil No. 1.

Closer examination of airfoil No. 1 and adjoining airfoils revealed that they exhibited a dark green area on the pressure surface, adjacent to the blade root radius. The area was approximately 0.2 by 0.5 inches and exhibited metal loss. According to their metallurgical report, the appearance and location of the metal loss area was typical of "hot corrosion attack." A fractographic view of airfoil No. 1 displayed a hot corrosion area that thinned the airfoil approximately 70 percent of its chord width. A Scanning Electron Microscope (SEM) examination of the remaining fracture surface revealed "features indicative of tensile overload." There was no evidence of fatigue progression.

Another airfoil (No. 2 airfoil), which was adjacent to the No. 1 airfoil (next airfoil in the clockwise direction when viewing the leading edge of the airfoils), was examined. Metallographic examination of airfoil No. 2 revealed the wall thickness was "severely reduced," and a closer examination of the reduced area revealed "spherical particles ahead of the hot corrosion front, which is indicative of sulfidation." Examination of the non-pressure side of the airfoil revealed an "etching zone at the outer airfoil, indicative of overheating."

Examination of radial section of a randomly selected airfoil on the 2nd stage turbine wheel revealed voids in the airfoil, indicative of incipient melting.

Hardness and material tests conducted on the turbine wheel indicated it met its engineering requirements.

Technicians examined the gas producer thermocouple assembly (for the Turbine Outlet Temperature, TOT gauge), and noted that three of the four thermocouple tips were broken. The thermocouple assembly was not tested. According to Aircraft Gas Turbine Powerplants, a false low TOT indication can be attributed to an open thermocouple hot junction; however, it was not possible to determine when the thermocouple tips broke. On July 25, 2001, the IIC observed a test, which was in accordance with the Barfield Indicator Test Procedure, of the TOT gauge. Technicians compared the accident helicopter's TOT gauge to a test gauge, which was subjected to temperatures between 400 and 1,000 degrees Celsius, in 100-degree increments. The accident helicopter's gauge indicated a minimum difference of zero degrees (at 600 degrees) and a maximum difference of 30 degrees Celsius (at 900 degrees).

Maintenance Information

Review of the helicopter's maintenance records revealed that the engine's turbine section (part number 6898735, serial number CAT-36444) had been removed on February 15, 1999, at an engine total time of 1,769 hours, because the 1st and 2nd stage wheels were due for an overhaul. The manufacturer recommended overhaul period for the turbine wheels is 1,775 hours. The turbine section was overhauled and reinstalled on the engine.

On April 16, 1999, at an engine total time of 1,978.5 hours, maintenance personnel removed and replaced the TOT gauge because of its inaccuracy. The removed TOT gauge was part number 369A4521-7, and the serial number was 1040.

On August 9, 2000, at an engine total time of 3,500 hours, maintenance personnel removed the compressor and turbine wheels for overhaul. They installed a different turbine section (part number 6898735, serial number 24353). The total time on the installed turbine section was 1,055.7 hours with a "TSO [time since overhaul] = NEW."

On August 25, 2000, maintenance personnel removed the turbine at an engine total time of 3,558.5 hours "due to No. 1 wheel rubbing." According to the maintenance endorsement, they reinstalled the turbine section (serial number 24353) on the engine on September 1, 2000, after No. 1 wheel replacement.

On November 23, 2000, at an engine total time of 3,777.7 hours, maintenance personnel removed the engine's combustor case due to a crack in the right armpit web. They installed a replacement case with zero time accumulated since overhaul.

On February 11, 2001, at an engine total time of 4,052.2 hours, maintenance personnel removed the TOT gauge "due to excessive instrument error," and they reinstalled the previously removed TOT gauge (serial number 1040). The 1040 TOT gauge was overhauled and given an airworthiness approval tag on November 20, 2000, and had not accumulated any time since its overhaul.

The helicopter underwent its last 300-hour inspection on June 2, 2001, at an engine total time of 4,449.7 hours. Review of the aircraft daily log sheet, a component log sheet, and the maintenance records revealed that the engine had accumulated a total time of 4,450.4 hours as of the day of the accident. According to the daily log sheet and the component log, the gas producer's 2nd stage turbine wheel had exceeded the recommended overhaul period of 1,775 hours by approximately 231 hours.

NTSB Probable Cause

a loss of engine power while in cruise flight resulting from the failure and liberation of several turbine blades due to hot corrosion effects which weakened the blades. The hot corrosion damage was the result of multiple turbine over temperature events which occurred over several flights. Contributing factors were the lack of suitable terrain, and high vegetation.

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