C-FICL – Robinson Helicopter Company R44 Raven II – 2006-05-01 NTSB Accident Report

HISTORY OF FLIGHT

On May 1, 2006, about 1430 Pacific daylight time, a Robinson R44 Raven II, Canadian registration C-FICL, experienced an in-flight breakup and impacted desert terrain near Desert Center, California. Zimmer Air Service, Inc., was operating the helicopter under the provisions of 14 CFR Part 91. The Canadian certificated commercial pilot and the non-rated passenger sustained fatal injuries. The helicopter was destroyed. The cross-country delivery flight departed Zamperini Field Airport (TOA), Torrance, California, at 1305, with a planned refueling stop at Blythe Airport (BLH), Blythe, California. Visual meteorological conditions prevailed, and no flight plan had been filed. The approximate global positioning system (GPS) coordinates of the primary wreckage were 33 degrees 38.016 minutes north latitude and 115 degrees 05.224 minutes west longitude.

The pilot had taken delivery of the new helicopter from the Robinson Helicopter Company factory in Torrance on the day of the accident. The pilot and passenger, who were employed by Zimmer Air Service, Inc., departed from TOA with an intended final destination of Blenheim, Ontario, Canada.

Riverside County Sheriff's Department personnel responded to the accident site and interviewed two ground witnesses. The witnesses related that they saw the helicopter just before it impacted the ground. They reported that the tail boom had separated from the fuselage; they did not see the separation sequence.

PERSONNEL INFORMATION

The operator reported that the pilot held a Canadian issued commercial pilot certificate with ratings for airplane single engine land and helicopter, with the following type ratings: BH06; BH47; BH47T; HL12; RH44; S350; SK55; and SK58.

The pilot held a first-class medical certificate issued on December 5, 2005. It had the limitation that the pilot must wear glasses.

The operator reported that the pilot had a total rotorcraft flight time of 7,615 hours. He logged 25 hours in the last 90 days, and 12 hours in the last 30 days. He had an estimated 1,041 hours in this make and model, and had completed a biennial flight review on March 31, 2006.

The operator reported that the passenger did not hold any pilot certificates nor was he attempting to obtain any ratings. The passenger was related to the pilot by marriage and had flown with the accident pilot on previous delivery flights.

AIRCRAFT INFORMATION

The helicopter was a Robinson R44 Raven II, serial number 11209. A review of the helicopter's logbooks revealed that the helicopter had a total airframe time of 4.0 hours when it left the factory. The tachometer read 5.2 hours at the accident scene.

The engine was a Textron Lycoming IO-540-AE1A5, serial number L-31092-48A. Total time recorded on the engine was 5.2 hours.

Fueling records at TOA established that the helicopter was fueled to capacity (48.9 gallons) on May 1, 2006. Examination of the maintenance and flight department records revealed no unresolved maintenance discrepancies against the helicopter prior to departure.

METEOROLOGICAL INFORMATION

The closest official weather observation station was Blythe airport, which was 18 nm east of the accident site. The elevation of the weather observation station was 397 feet mean sea level (msl). A METAR for BLH was issued at 1453. It reported winds from 170 degrees at 8 knots; visibility 10 miles; skies clear; temperature 38 degrees Celsius; dew point 01 degrees Celsius; altimeter 29.74 inHg.

COMMUNICATIONS

The helicopter was not in contact with any ATC facility.

WRECKAGE AND IMPACT INFORMATION

Investigators from the National Transportation Safety Board, the Federal Aviation Administration (FAA), and Robinson Helicopter Company examined the wreckage at the accident scene.

The on-scene examination of the accident site and wreckage revealed that the helicopter and separated components came to rest about 200 yards south of Interstate 10 (I-10) at mile marker 124. The accident site was 16 nm east of Desert Center, and 25 miles west of the city of Blythe. The area was open and flat with low shrubs, typical of desert terrain.

The main wreckage, consisting of the fuselage, engine, transmission, cockpit, most of the right skid tube, a portion of the left skid tube, and the majority of both main rotor blades, came to rest on an orientation of 280 degrees magnetic.

The tail boom was 100 feet west of the main wreckage. The tail boom separated just aft of the main fuselage attach point. The tail rotor blades and tail rotor transmission were attached to the tail boom, with minor ground impact damage noted.

The primary wreckage debris field was about 500 feet long. The primary wreckage debris field consisted of the right side pilot's door, which was in two pieces, and pieces of fiberglass and Plexiglass. A secondary debris field, consisting of papers and light cockpit materials, was another 400 feet downwind from the accident site. The debris path was along a magnetic bearing of 090 degrees.

Cockpit and Cabin

The removable cyclic was installed on the left side, which is the normal position for a copilot or passenger. The removable anti torque pedals and collective had not been installed on the left side. The removable pedals were in the baggage compartment. Both the pilot and copilot cyclic grips were detached from their respective mounting positions. The copilot cyclic grip was also crushed inward. The pilot's collective was in a position close to full down, and the friction slider was fractured below the slot. The pilot's pedals were in a position close to full left. The ignition key was in the OFF position, while the throttle twist grip was in a full on position. The cockpit mixture control was in the full rich position; however, the arm at the fuel servo was in the full lean position.

Miscellaneous personal items were recovered, which included luggage, clothing, toiletries, a digital camera, a P.D.A., a handheld GPS, quart bottles of oil, a cabin cover, airport guidebooks, notes, and aircraft charts.

Fuel System

Both fuel caps were in place. Both fuel tanks ruptured with hydraulic bulging in the lower corners. There was a strong odor of fuel in the sand beneath the helicopter and evidence of fluid flow. The finger strainer in the auxiliary tank was free of debris. The fuel line between the auxiliary tank and the main tank separated, and fractured in several places.

The gascolator bowl was dislodged, but remained in place. Investigators removed the gascolator bowl. It was clean inside, but the screen was partially crumpled and inside the bowl. The fitting between the gascolator and the electric fuel pump was broken. The top of the electric fuel pump separated from the lower portion.

The fuel return line between the return pressure valve and the auxiliary tank was kinked, but it was still intact. Investigators were able to blow air through all the components of the vent lines.

Lubrication Systems

The main rotor and tail rotor systems are lubricated using Swepco 90 weight gear oil, which is normally blue in color.

The engine is lubricated using 20-50-weight mineral oil, which is normally a transparent golden color.

The hydraulic system operates using Robinson hydraulic fluid, Part number A257-15, which is normally red in color.

Powerplant

The powerplant was severely damaged. Most of the accessory case fractured. The alternator drive pulley was in contact with the starter ring gear. The pulley exhibited a row of scores on its aft surface. The space of the scoring was approximately in size and dimension (spacing) to the ring gear teeth. The upper sheave's rear surface was in contact with the starter ring gear. There was some scoring at the contact area, with the scores spaced at the approximate size and dimension of the ring gear teeth. The engine oil filler cap was found secured to the filler tube, which had broken off of the engine casing.

Driveline

The drive belts were all intact, with no sign of rolling or chafing. The upper and lower sheaves were displaced longitudinally, and the belts were not in place in the sheave grooves. The belt tension actuator had broken into several pieces. Investigators were unable to locate part of the tube between the upper and lower portion the actuator. The drive motor and worm gear separated, and the attachment lug to the upper actuator fractured. The upper and lower actuator bearings rotated freely. The sprag clutch locked, and free-wheeled normally when manually manipulated. The forward flex coupling remained intact, although distorted.

The main rotor gearbox was open, and fractured into several pieces. The shaft tilted forward, and the gears exhibited blue oil residue. The output shaft and the hydraulic pump drive were free to rotate approximately 45 degrees. However, motion was constrained by interference further up the mast. The condition of the pinion and ring gears appeared normal. The mast tube exhibited a large dent approximately halfway up, and the main rotor shaft appeared to be bent slightly.

Both droop stops and their corresponding retention bolts were in place. Both elastomeric teeter stops had been crushed in the center. The one for the main rotor blade (SN 2608B) was also displaced from its retention bracket; however, both retention brackets remained intact. There was slight circular arc scoring on the main rotor hub just inboard of both main rotor blade pitch change housings. The upper swashplate and the main rotor magnetic pickup mount bracket exhibited a yellow substance, as well as some deformation. The spindle for the main rotor blade (2608B) fractured at the coning bolt, but the side opposite the droop stop "tusk" remained intact around the coning bolt shank. The droop stop tusk and the spindle for the main rotor blade (2628B) were intact.

Both main rotor blades were bent downward with compression buckling on the aft portion of the blade. One main rotor blade (SN 2608B) was separated at two locations that were approximately 27 and 142 inches from the coning hinge bolt. The main blade spar fractured approximately 109 inches from the coning hinge bolt, but the aft portion of the blade remained intact. The lower portion of the blade had a piece of Plexiglas embedded in its lower surface approximately 77 to 88 inches from the coning bolt.

The second main rotor blade (SN 2628B) was bent downward more than 90 degrees at approximately 32 inches from the coning bolt, and was partially separated at 61 inches from the coning bolt. This separation was approximately 7/8 of the chord from the leading edge of the blade to the trailing edge; the remaining attached portion of the blade was cut during the recovery process. The remainder of the blade from this fracture outboard was intact.

The intermediate flex coupling was intact, but distorted. The tail cone separated just aft of the fuselage attach point. The tail rotor damper bearing rotated freely. The manufacturer represented that it felt a little gritty. The stainless steel friction arm separated, and the mounting bracket separated from the tail cone.

The aft flex coupling was intact. The tail rotor gearbox was intact; with the exception of one mount tab. It contained blue oil, and it rotated at least one full turn. The pitch change slider bearing rotated freely. The hub and tail rotor blades were intact.

Flight Controls

Investigators examined the flight control systems and located numerous discontinuities throughout the flight control assemblies. Examinations of the fracture surfaces associated with the flight control discontinuities were all observed to be consistent with overload. According to the manufacturer, no evidence of preimpact failures of the flight control systems was found.

Airframe

The fuselage exhibited severe crushing and impact damage. The right front door and the cabin immediately above and below the door, exhibited a vertical cut that extended laterally inboard approximately 18 inches. The rain gutter exhibited a yellow material above the cut in the right front door. The door handle opening on the right front door had a small piece of what appeared to be main rotor blade skin lodged in it. Main rotor blade (SN 2608B) was found extending through the mast fairing, aft of the mast tube, approximately halfway up.

The tail cone separated in the forward portion of the second bay. At the separation, the tail cone was partially flattened out horizontally. The tail cone was intact from the separation point aft, and exhibited a small dent adjacent to the tip of the tail rotor. The empennage, tail rotor visual guard, and tail skid "stinger" were intact. The upper and lower stabilizers exhibited minor buckling.

The belly of the aircraft, as well as the tail cone and empennage, exhibited a thin coating of engine oil. The oil film was denser at the forward portion of the tail cone, becoming less dense moving rearward. The Robinson representative stated it was not uncommon to see an excess of oil being discharged out the breather tube due to overfill at the factory.

Most of the landing gear remained attached to the airframe. The right skid toe was approximately 140 feet northeast of the main wreckage; it exhibited a large dent on the right side near the tip. Both cross tubes remained attached to their respective struts, and were fairly straight. The left skid separated at the attachment points for both left struts.

Wreckage Recovery

On May 2, 2006, the wreckage was recovered from the accident site for further investigation. On May 12, 2006, the FAA accident coordinator returned to the accident site, and recovered a portion of the main rotor pitch horn and pitch link.

MEDICAL AND PATHOLOGICAL INFORMATION

The Riverside County Coroner completed an autopsy. The FAA Bioaeronautical Sciences Research Laboratory, Oklahoma City, Oklahoma, performed toxicological testing of specimens of the pilot. Analysis of the specimens contained no findings for tested drugs.

The report contained the following findings for volatiles: 16 mg/dL, ethanol detected in muscle. The report stated that the ethanol found in this case was from sources other than ingestion of ethanol.

TESTS AND RESEARCH

Investigators examined the wreckage at Aircraft Recovery Service, Littlerock, California, on May 3, 2006. The FAA, Robinson Helicopter Company, and Lycoming were parties to the investigation and assisted in the examination.

Investigators removed the engine. They slung it from a hoist, and disassembled the engine.

Investigators removed the top spark plugs. All spark plugs were clean with no mechanical deformation. The spark plug electrodes were gray, which corresponded to normal operation according to the Champion Aviation Check-A-Plug AV-27 Chart.

Investigators could not rotate the crankshaft due to impact damage to the engine case. Mechanical continuity of the rotating group and internal mechanisms was established visually during the disassembly and examination of the engine. According to the manufacturer's investigator, no evidence of lubrication deprivation or contamination, or any evidence of signatures or conditions indicated any preimpact catastrophic mechanical malfunction were observed. The camshaft was intact, and each cam lobe appeared normal in shape. The accessory gears, including the crankshaft gear, bolt, and dowel, were intact, and remained undamaged by any preimpact malfunction. The combustion chamber of each cylinder remained undamaged. The gas path and combustion signatures, as viewed from the spark plugs, combustion chambers, and exhaust components, displayed coloration consistent with normal operation. The manufacturer's investigator also indicated that the valve train was observed to be undamaged by any preimpact mechanical malfunction.

The magnetos sustained substantial damage. The right magneto could not be tested due to the sustained damage; however, investigators were able to obtain spark from the left magneto. The ignition harness sustained varying degrees of damage and could not be tested. However, the ignition harness was attached at each magneto and respective spark plug.

The oil filter was detached and crushed. The oil suction screen and oil filter were uncontaminated by any preimpact debris. The manufacturer's investigator also did not observe any evidence of preimpact lubrication system contamination.

The fuel injector servo was partially displaced from the engine with a portion of the servo remaining secured at its mounting pad. It exhibited fracture surface signatures that the manufacturer's investigator stated were consistent with overload. The fuel injection servo and induction system were examined and observed to be free from obstruction. The throttle and mixture controls were securely attached to their respective control arms of the servo. The servo fuel inlet screen was not recovered.

The fuel flow divider remained secure at the mounting bracket. Fuel lines remained secure at each flow divider fitting, and fuel injectors remained secure at each cylinder. The fuel injection nozzles remained secure at each cylinder with the respective fuel line attached. The flow divider was disassembled. There was no evidence of internal mechanical malfunction or obstruction to the fuel flow. The diaphragm remained intact and undamaged. The fuel ejection nozzles were also removed and examined. The nozzles were free of visible contamination and obstruction flow.

The fuel pump was displaced from the engine and fractured. The portion of the mounting flange remained secure at the mounting pad. It exhibited fracture surface signatures, which the manufacturer's representative stated were consistent with overload. He opened the fuel pump for examination. The fuel pump remained free of internal mechanical malfunction and obstruction to the flow.

The warning light panel was examined; the low rotor rpm, the oil pressure - engine; the governor OFF, and the auxiliary fuel pump lights each had a stretched coil. The low fuel, clutch, tail rotor chip light, main rotor chip, main rotor temperature, rotor brake, carbon monoxide, engine fire, and fuel filter lights each had a tight coil. With regard to the alternator, the filament appeared to be excessively long, but no stretch was observed.

The detached pitch link and the portion of the main rotor pitch horn was sent to the Safety Board Materials Laboratory, Washington, D.C., for further examination. The Materials Laboratory examined the pitch link, and noted that a portion of the pitch link thread adjacent to the fracture face displayed "cup and cone" features consistent with a bending overload event.

Examination of the fracture face of the portion of the main rotor pitch horn that was still attached to the pitch link displayed a clean surface with no indications of a pre-existing condition and 45-degree shear lips, consistent with an overload event.

A damaged handheld Garmin GPS, model number 296, serial number 670199944, was recovered at the accident site. The unit was sent to Garmin International to attempt to recover any flight data. Garmin personnel examined the unit but were unsuccessful in extracting any data.

Searches of all available radar sources in the area of the accident were unsuccessful in recovering any data of the accident flight.

On August 17, 2006, a Safety Board investigator took the spindle and blade root of the main rotor blade (SN 2608B) to Robinson Helicopter Company for hardness testing. The spindle drawing calls out a hardness requirement of HRC 36.5-41.0. Tested values ranged from 36.6 to 37.7. The investigator and Robinson Helicopter Company personnel also tested the hardness of the separated blade pitch horn. A review of the manufacturer's drawing for the pitch horn revealed no specified hardness requirement for the component. The hardness was measured using a scale and tip size not recommended for forged aluminum, so the hardness readings were not accurate and contained broad scatter results.

ADDITIONAL INFORMATION

On April 2, 1996, the Safety Board adopted a Special Investigation Report, "Robinson Helicopter Company R22 Loss of Main Rotor Control Accidents," (NTSB/SIR-96/03). The Safety Board's special investigation initially focused on R22 accidents in which the main rotor blade diverged from its normal path and struck the helicopter. When similar R44 accidents occurred, the special investigation was expanded to include those accidents.

In the findings of the investigation the report states, "...between 1981 and 1985, 31 R22 accidents and three R44 accidents (domestic and foreign) have involved a loss of main rotor control. The investigations of these accidents did not identify any precipitating progressive mechanical failures or material defects. The cause of the loss of main rotor control in many of the accidents most likely stems from a large, abrupt pilot control input to a helicopter that is highly responsive to cyclic control inputs."

The SIR identified the following characteristics that were common to the R22 and R44 accidents identified above.

1. An in-flight breakup or main rotor blade contact to the airframe occurred in flight before collision with any object or terrain.

2. There was no evidence of an initiating airframe or engine component malfunction.

3. Flight into adverse weather such as low visibility or ceilings was not involved.

4. Pilot impairment from drugs or alcohol was not involved.

In the Robinson Model R44 II Pilot Operating Handbook, under normal procedures, Section 4, page 4-5 it states:

CAUTION

Remove left seat controls if person in

that seat is not a rated helicopter pilot.

The FAA approved this POH change on June 13, 2005.

The Safety Board investigator released the wreckage to the owner's representative on August 29, 2006.