N318Y – Sikorsky (Lockheed Martin) S-61A – 2002-03-26 NTSB Accident Report

HISTORY OF FLIGHT

On March 26, 2002, at 1538 Pacific Standard Time, a Sikorsky S-61A twin-engine helicopter, N318Y, was destroyed when it impacted trees and terrain following a loss of control while maneuvering near Lakehead, California. The helicopter was registered to and operated by Croman Corporation of White City, Oregon. The commercial pilot was fatally injured and the commercial co-pilot sustained serious injuries. Visual meteorological conditions prevailed, and a flight plan was not filed for the 14 Code of Federal Regulations Part 133 external load flight. The local flight originated from a staging area near the McCloud River, at 1500.

According to the operator, the helicopter maneuvered for approximately 40 minutes in the area as part of a logging operation prior to the accident. One witness, who was a ground worker, was in radio contact with the pilot. The worker stated that the pilot was supposed to drop chokers near him. A choker is a steel cable that is fitted around a log, and is standard equipment for logging operations. The chokers were attached to a hook that hung from a 200-foot long-line connected to the helicopter. The witness stated that he had radioed to the pilot and informed the pilot that he had moved 100 yards toward Mount Shasta, thus providing the pilot with a reference point to drop the chokers, and the pilot acknowledged. There were no further radio transmissions received from the pilot. The helicopter did not drop the chokers at the point the witness expected, and it over flew the intended drop zone. The witness stated that the helicopter appeared to be gliding, and he observed a tan smoke coming from the engine/transmission area. He stated, "the blades were not moving as fast as they should have been," and "the blades were coning extensively." He added that "the blades were hardly rotating and they were pointing toward the clouds." Two other witnesses reported the main rotor blades had slowed down and they could count each blade. Subsequently, the helicopter "fell like a rock," impacted the ground, and a fire eventually erupted.

The one witness, who described the blade coning, indicated it took him approximately 5 minutes to get to the accident site. While he was making his way to the accident site, the helicopter caught fire (within less than a minute). As the witness approached the helicopter he heard "the engines…running, surging up and down, up and down." Another witness indicated he heard no engine noise when the helicopter over-flew his position.

PERSONNEL INFORMATION

The pilot-in-command (PIC) held a commercial pilot certificate with the following commercial ratings: rotorcraft-helicopter, airplane single- and multi-engine land, and instrument airplane and helicopter. He held a flight instructor certificate with a rotorcraft-helicopter rating; however, it was not current. He was type rated in the Sikorsky SK-58 and SK-61 helicopters (SK is the FAA designator for Sikorsky type ratings). On September 15, 1993, the PIC completed the requirements to act as pilot-in-command for external load operations that are set forth in FAR Part 133 (class A, B, and C). On the morning of the accident the pilot successfully completed a proficiency check in the accident helicopter, in accordance with FAR 61.58. According to a form provided by Croman Corporation, the PIC had accumulated a total of 25,800 flight hours, of which 20,000 hours were in external load operations, and 15,000 hours were in the S-61. He held a second-class medical certificate that was dated January 25, 2002, with the stipulation that he possess glasses that correct for near vision.

The co-pilot was issued a commercial helicopter certificate on November 29, 2000. He also held a flight instructor certificate for helicopters that was issued on March 28, 2001. According to a form provided by Corman Corporation, the co-pilot had accumulated a total of 295.35 flight hours, of which 288.95 hours were in external load operations in the S-61A. He held a second-class medical certificate that was dated April 4, 2001, with the stipulation that he wear corrective lenses.

AIRCRAFT INFORMATION

The helicopter (serial number 61094) was manufactured in 1962, and was an all-metal, semi-monocoque construction, and was equipped with a 5-blade main rotor system and a 5-blade tail rotor system. The helicopter utilized fixed landing gear without sponsons in place of the original sponson-mounted retractable landing gear. The helicopter was involved in an accident in 1972, which resulted in two fatalities. The FAA listed the helicopter as destroyed as a result of that accident, but it was rebuilt in 1975 by Carson Helicopters, and registered in the restricted category in July 1980, by Croman Corporation. The helicopter was configured for logging operations. The helicopter was listed on Croman's Supplemental Type Certificate, which authorized the use of S-61L/N (20600 series), or S-61R (23000 series) main transmissions. At the time of the accident, the helicopter was utilizing a 23000 series main transmission.

The helicopter was also equipped with two 1500-horsepower General Electric (GE) CT58-140-1 turboshaft engines. The GE CT58-140-1 engine is an axial-flow turboshaft engine incorporating the free turbine principle. Output power is extracted by a free turbine, which is mechanically independent of the gas generator rotor system. The gas generator consists of a ten-stage compressor, annular combustor and two-stage turbine. "Efficient and stall-free operation" is ensured by use of the variable stator principle in the compressor. The inlet guide vanes and stator vanes (in stages 1, 2, and 3) are variable. The turbine section includes a gas turbine section (Ng), which consists of two turbine rotors that provide rotational power to the compressor, and a power turbine (free turbine; Nf), which consists of one turbine rotor that extracts energy from the exit gases and coverts it to shaft horsepower for useful work. Engine performance is monitored by various engine instruments, including the torquemeter indicator (two needles labeled 1 and 2 for the #1 and #2 engine, respectively), and the triple tachometer (three needles labeled 1, 2, and R for #1 power turbine speed, #2 power turbine speed, and main rotor RPM, respectively).

Croman Corporation maintained the helicopter, and on October 24, 2001, the main gearbox was removed from the helicopter and replaced with a serviceable main gearbox. The main gearbox then underwent an overhaul inspection in accordance with the Sikorsky overhaul manual (SA4045-83). On December 18, 2001, the helicopter was removed from service and a major inspection was started in accordance with the Sikorsky Major Inspection Guide (SA 4047-14). During the inspection the overhauled main gearbox was re-installed. On February 26, 2002, at a helicopter total time of 41,761.7 hours, the #1 and #2 engines were replaced with the same engines that were involved in the accident. The #1 engine (left hand engine; serial number 296-018D) was installed at an engine total time of 17,239.7 hours, with a total time since overhaul of 435.4 hours. The #2 engine (right hand engine; serial number 295-057) was installed at an engine total time of 26,764.7 hours, with a total time since overhaul of 5,632.2 hours. The major inspection was completed on March 6, 2002, 76.1 hours prior to the accident. On March 21, 2002, the helicopter underwent a phase C inspection (zones 9 hull, 6 fuel cell installation, 2 power plant; #2 engine). The inspection was accomplished 11.5 hours prior to the accident.

According to maintenance personnel, on the morning of the accident the helicopter flew 5.9 hours. At noon the following helicopter systems/components underwent a visual condition check: the main rotor system, main gearbox, flight controls, #1 engine, and #2 engine. According to the maintenance entry for March 26, 2002, the following Airworthiness Directives (AD) and Service Bulletins (SB) were complied with: AD 79-12-10 (tail rotor blade ultrasonic check), AD 68-20-08 (automatic flight control system functional check), AD 78-12-05 (main landing gear fitting visual inspection), AD 85-18-05-R2 (Blade Inspection Modules - BIM; a safety indicator utilized to check the nitrogen pressure in the blade spar; functional and visual inspection), SB 61B10-18 (tail rotor spindles visual check), SB 61B20-03 (station 243-290 visual check), SB 61B35-66 (main gearbox screen visual check). No anomalies were noted during the checks. The helicopter was refueled and departed. At the time of the accident the helicopter had accumulated approximately 41,837.8 hours.

The five main and tail rotor blades and their respective components were color coded for maintenance purposes. Each blade and its respective components were assigned one of five colors; red, blue, yellow, white and black.

METEOROLOGICAL INFORMATION

At 1553, the weather observation facility at the Redding Municipal Airport, Redding, California, (located 30 miles southwest of the accident site) reported clear skies, wind variable at 3 knots, a temperature of 70 degrees Fahrenheit, a dew point of 36 degrees Fahrenheit, and an altimeter setting of 29.97 inches of Mercury.

FLIGHT RECORDERS

A cockpit voice recorder was not installed in the helicopter, nor was one required.

WRECKAGE AND IMPACT INFORMATION

The helicopter came to rest in a densely wooded area of Shasta National Forest. The accident location was recorded by a global positioning satellite (GPS) receiver at 041 degrees 02.289 minutes north latitude and 122 degrees 11.404 minutes west longitude, and at an elevation of 2,240 feet msl. The helicopter came to rest on sloping terrain that varied between 20 and 80 degrees, with the nose oriented on a magnetic heading of 345 degrees. The surrounding terrain consisted of 150-foot, closely set, pine trees. Though the main rotor system has a 62-foot rotating diameter, only two trees sustained any impact damage. The entire 200-foot cable remained attached to the helicopter and came to rest within a 30-foot area of the helicopter.

The cockpit, cabin, and transmission were consumed by fire. The tail boom remained intact; however, it sustained fire damage to the forward and upper sections. The tail boom, from the intermediate gearbox aft, did not sustain fire damage.

The main rotor head (MRH) was examined and the hub assembly, with its sleeve-spindles intact, remained attached to the rotor mast. The red spindle was found rotated approximately 180 degrees from its normal position and, along with the black spindle, could not be moved. The blue spindle could be moved in the lead-lag, flapping, and pitch rotational directions. The yellow and white spindles could not be evaluated for movement due to the weight of the rotor head. All five dampers remained attached at both ends. All of the pitch change rods were fractured, with the exception of the yellow. All but the yellow and blue pitch change horns were burned to some extent (the red and white being burned away in their entirety). The red, blue, yellow, and white droop stops were found retracted. The black droop stop was loose and its position could not be determined. None of the droop stops displayed signs of major impact with their spindles. Two of the bifilar weights were attached to their arms between the red/blue blades and the blue/yellow blades. Two weights were found in the impact area, and one was not recovered.

Four of the five main rotor blades remained attached to the MRH. These were the red, blue, yellow, and black blades. The white rotor blade was fractured near the root and was located next to the right side of the helicopter. No evidence of leading edge impact damage was noted on any of the blades. All of the main rotor blades, with the exception of the yellow, sustained fire damage outboard of the root area. All five BIM indicators sustained fire damage. All five tip weights remained attached to their respective blades.

The stationary scissors, from the main gearbox to the stationary swash plate, was melted at the swash plate attach point. The rotating swash plate and scissors were intact. The main rotor gearbox's forward, aft, and lower magnesium casing were completely consumed by the fire, but 60% of the top cover remained. The main bevel gear and planetary gear system remained attached to the main shaft. All exposed gear teeth appeared to be in good condition. The lower planetary plate was intact. The exposed area of the upper planetary plate was intact. Both high-speed engine inputs (including portions of their sleeve bearings), both input freewheel units (IFWUs), the main bevel pinion shaft (including the helical gear and the rotor brake disk), the through shaft and its freewheel unit, the tail takeoff gear (TTO gear) and shaft, and a number of accessories and accessory drive gears were all recovered from the impact area. It was not possible to verify the main gearbox continuity due to the destruction of the casing. No evidence of distress or excessive wear was noted on any of the gear, gear teeth, bearings, or the rotor brake disk. The IFWUs were retained for further examination.

The #1 tail rotor drive shaft remained attached to the TTO gear but was separated from the #2 coupling flange at the shaft/flange mating surface. The tail rotor drive shaft remained intact aft to the tail rotor head. The intermediate gearbox remained mounted to the base of the vertical pylon and rotational continuity was confirmed. The tail rotor gearbox remained intact and displayed no rotational drag when it was rotated to confirm internal continuity. All five tail rotor pitch change links remained attached and intact. All five tail rotor blades remained intact; however, the white and black blades sustained impact damage.

The engines, still installed side-by-side on the engine bay platform, were examined at the Croman Corporation facility under supervision of the NTSB Investigator-In-Charge.

Left Engine

The left engine was disassembled. The compressor blades (10 stages) were not damaged and no rotational scoring was observed on the compressor shroud. The compressor's variable inlet guide vanes and stators were not damaged. The variable stator vane lever arms were all found in the fully closed position, though their mechanical linkage, from the actuator to the lever arms, was broken. The variable IGVs were also found in the closed position. According to the engine manufacturer, the variable stator vanes will modulate between fully open at about 95% gas generator speed and above (high power), to fully closed at about 64% gas generator speed and below (low power). Engine idle speed is about 56% gas generator speed, during which the vanes are fully closed.

The combustion chamber was intact and the fuel nozzles were not obstructed.

The 1st and 2nd stage turbine rotors were intact. The turbine blades were not damaged, and no rubbing or rotational scoring was noted on the blade tips or on their corresponding shrouds. The power turbine rotor was intact. The power turbine blades were not damaged, and no rubbing or rotational scoring was noted on the blade tips or on their corresponding shrouds. The power turbine exhaust casing was intact and covered with magnesium residue. The power output drive shaft was seized due to magnesium deposits that had forged around the drive shaft.

The accessory drive radial shaft and accessory casing drive gear were intact and fire damaged. The following accessories were heavily damaged by the fire; the dynamic fuel filter, fuel pump, fuel control unit, stator vane actuator, anti-icing valve, starting bleed valve, and the Nf tachometer. The oil cooler, static fuel filter, fuel flow manifold, Nf flex shaft's right angle drive, Ng tachometer, and ignition unit were not found and are presumed to have been destroyed by the fire.

Right Engine

The engine sustained external fire damage along the entire bottom side of the engine, which progressed inward toward the engine core throughout the entire length of the engine. The right engine was disassembled. Re-solidified molten metal slag was found throughout the engine core in the area where extensive fire damage occurred (post-disassembly examination of the slag revealed it had a high content of magnesium and aluminum, with lesser contents of nickel, chromium, molybdenum, zinc, and iron, which Sikorsky advised was generally consistent with the composition of the magnesium alloy casing of the helicopter main transmission gearbox that was consumed by fire).

The compressor blades, for the most part, were not damaged and no rotational scoring was observed on the compressor shroud. Resolidified molten slag was found on the bottom side (engine's 6 o'clock position) of the compressor (stages 5 through 10). Some of the blades in this resolidified molten slag area were found burned and melted. The compressor's variable inlet guide vanes and stators were not damaged. The variable stator vane actuator was in the near retracted position (0.14 inches extended), and the stator vanes appeared to be at or near their closed position with their linkage mechanism intact and connected.

Resolidified molten slag was observed at the front end of the combustor dome between the inner and outer liners, covering two of the fuel nozzle tips at around the 6 o'clock location. Three small holes were burned through the forward end of the outer combustion liner near the resolidified slag. In addition, the combustor casing had two burn holes at its forward end at the 6 o'clock location where the fuel manifold assembly enters through the casing. The steel drain cover below these two holes was partially burnt away and the aluminum fuel manifold block and fuel flow divider were both missing and were assumed burnt away. In the annulus between the outer combustion liner and the combustor casing at this 6 o'clock location, one of the four fuel tubes of the fuel manifold assembly was burnt and completely separated about one inch from its end. This damage was in the left hand manifold tube to the primary nozzles, and aligned with the burn holes at the forward end of the combustion casing and outer liner (the fuel manifold was retained for further examination). The rest of the inner and outer combustion liners displayed no other anomalies that would have precluded their function.

The 1st and 2nd stage turbine rotors were intact; however, were heat damaged. The 1st stage turbine rotor blades were missing 37 mils from the tip inboard, and displayed leading edge and tip burning. Its shroud exhibited resolidified molten deposits consistent with the blade tips all around its leading edge. In addition, the stage 1 shroud displayed rubbing and heat distress all around. The 2nd stage turbine rotor blades were missing 10 mils from the tip inboard and its shroud exhibited deposits consistent with the tips shedding their material while rotating. The power turbine rotor was intact and there was no evidence of rub marks or rotational scoring on the blades or on the blades corresponding shroud. The power turbine exhaust casing displayed heat distress aft of the power turbine rotor and a hole exposed the power turbine rotor. The exhaust casing also was covered with magnesium residue. The power output drive shaft was seized due to magnesium deposits that had re-solidified around the drive shaft.

The accessory drive radial shaft was intact and in place; however, the remainder of the accessory drive gears were not identified. The following accessories were heavily damaged by the fire: the starter, starter bleed valve, stator vane actuator, anti-ice valve, and the oil-air separator. No other accessories were identified.

The cockpit left side triple tachometer, left side torque indicator, and right side torque indicator were visually examined. The triple tachometer's #1 Nf and Nr needles were matched, and the #2 Nf needle was indicating slightly lower. The instrument's numbers were not legible. The left and right torquemeter depicted the #1 needle at approximately 13%, while the left torquemeter indicated the #2 needle at 140%. The right torquemeter's #2 needle was burned away.

TESTS AND RESEARCH

The fuel manifold from the right engine was submitted to the NTSB Materials Laboratory for examination. There was extensive melting of the assembly where the fuel transmission tubes passed into the combustion chamber of the engine. One of the tubes showed localized melting in that area causing a portion of the tube wall to disintegrate and result in an opening. This area of localized melting, with the manifold installed on the engine, coincided with the holes melted into the combustion casing. Visual and binocular microscope examination of the fuel transmission tubes and other parts of the assembly revealed no evidence of cracks.

The two IFWUs were examined at the Sikorsky Materials and Process Technology Laboratory under the supervision of the NTSB IIC. The units were examined radiographically in an attempt to determine the relative positions of internal components without disturbance; however, the densities of the units' steel precluded practical imaging of the free wheel unit components. The two free wheel units were sectioned open by Electric Discharge Machining (EDM) to allow direct examination of the internal components. Due to the fire-damaged condition of the components, no assembly dimensional inspections were attempted, except for the spring-loaded cage return pin extensions. According to Sikorsky, the highest confidence measurements taken from the #1 and #2 units did not indicate any unusual or improper operating conditions.

The half sections of the IFWUs (which included the gear housing, cam flats, and the roller cage) were cleaned per Sikorsky's M&P Specification SS8460, using a vapor blast process to remove the surface oxidation and debris deposits that resulted from the heat and fire damage. The roller diameters were checked and compared to the manufacturer's specifications (0.4138 - 0.4140 inches). The roller diameters ranged between 0.4125 - 0.4135 inches. Sikorsky opined that material loss could have occurred as a result of surface oxidation from the fire and the cleaning process.

#1 (Left) IFWU

The gear housing roller path, cam flats, roller, and cage showed no visible wear depression, evidence of skidding, impact marks, or any other indications of improper operation. In addition, there was no evidence of melting. The cage was magnetic particle inspected and found to be uncracked. In addition, the Oilite bushings were found unworn and uncracked. Hardness tests of the gear housing, cam, and randomly selected rollers revealed only the housing's inner diameter roller path met the manufacturer's specifications. According to Sikorsky, the components sustained some "fire-related tempering", which effected the component hardness.

#2 (Right) IFWU

The internal components from the #2 IFWU visually appeared to be in similar condition to that of the #1 IFWU's, with the exception of the Oilite bushing, which apparently melted away (bronze-colored re-solidified metal was found adhering to the cam flats and was determined to likely be remnants of the Oilite bushing). None of the steel components appeared melted. The conditions of all of the components indicated "normal operation" prior to the fire. Magnetic particle inspection of the cleaned cage revealed no cracks. Microstructural appearances of the #2 IFWU components revealed visual differences in the #2 rollers, housing inside diameter roller path, and cam flat when compared to the microstructural appearance for the same components in the #1 IFWU. Sikorsky determined this was likely the result of the #2 IFWU experiencing higher temperatures during the post-accident fire than the #1 IFWU.

No anomalies were noted with the IFWUs that would have led to a "spit-out" condition or other unusual operating conditions.

MEDICAL AND PATHOLOGICAL INFORMATION

An autopsy on the senior pilot was conducted at the Shasta County Coroner's Office. According to the autopsy report, the pilot died as a result of multiple traumatic injuries and blunt force trauma.

Toxicological tests for volatiles and drugs were performed on the first pilot. According the toxicological report, there was no ethanol detected in the pilot's kidney and muscle. An unquantified amount of sertraline was detected in the pilot's kidney and liver. Likewise, an unquantified amount of desmethylsertraline (a metabolite of sertraline) was detected in the pilot's kidney and liver. Sertraline is an anti-depressive medication also known as Zoloft. It use is not approved by the FAA.

The co-pilot sustained serious burn injuries and could not remember the events leading to the accident. He was found outside the helicopter, but could not recall how he got there.

ADDITIONAL INFORMATION

An Airworthiness Directive (AD) existed that required the replacement of the fuel flow dividers with specific serial numbers (or unreadable numbers). The AD (2001-12-07) indicated that the suspect fuel flow dividers could be susceptible to fuel leaks that could cause an engine fire. Review of the engine service records revealed that during the left engine's last major overhaul (June 2001) the serial number of its fuel flow divider was checked and found to be not applicable based on its serial number. The right engine compliance with the AD had been done during various times of the engines' life. Though the serial number is unknown, based on the right engine's fuel flow divider part number, it was determined that it did not utilize the affected components and therefore did not need to comply with the AD.

Review of the helicopter's flight manual revealed a section under emergency procedures regarding "Engine Failure". The section discussed main rotor decay rate, which is affected by the "engine torque required at the time of failure." According to the manual, a 5% per second decay rate can be expected when a failure occurs in the flight regimes requiring 74.5% torque. All of the emergency procedures for single-engine loss of power (regardless of phase of flight) dictate that the flight crew place the speed selectors in the full forward position and they "contain Nr." Another section discusses dual-engine failures, and indicates that, "failure of both engines requires immediate action if a safe power-off landing is to be accomplished. The altitude and airspeed at which a two-engine failure occurs will dictate the action to be taken. Immediately upon a two-engine failure, rotary wing rpm will decay and the helicopter will yaw to the left. Collective pitch must be immediately reduced to prevent excessive loss of rotary wing rpm…"

A warning accompanies this section indicating that, "within 3 to 5 seconds, rotor rpm will decay to an unrecoverable state with resultant loss of control unless autorotation is entered immediately. The section titled "Dual-Engine Failure While Hovering At A Low Altitude" indicates, "settling will be so rapid if both engines fail at low altitude that little can be done to avoid a hard landing. The landing can be cushioned somewhat by increasing collective pitch as the helicopter settles to the ground. Do not reduce collective pitch as in normal procedures in the event of both engines failing at higher altitude. In this case, a reduction of pitch would cause the helicopter to settle more rapidly."

Another section titled, "Landing In Trees," indicates "a power-off landing into a heavily wooded area can be made by executing a normal autorotative approach and full flare. The flare should be executed so as to reach zero rate-of-descent and zero ground speed as close to the top of the trees as possible. As the helicopter settles, increase collective pitch to maximum and allow the helicopter to descend vertically through the trees."