N445RH – Robinson Helicopter Company R44 – 1993-07-31 NTSB Accident Report

History of the Flight

On July 31, 1993, at 1349 hours Pacific daylight time, a Robinson R44 helicopter, N445RH, crashed during takeoff at El Monte airport, El Monte, California. The helicopter was being operated as a visual flight rules (VFR) personal flight in the local area when the accident occurred. The helicopter, registered to, and operated by Uni West Aviation Inc., Alhambra, California, was destroyed by impact and post impact fire. The certificated private pilot, and two passengers received fatal injuries. Visual meteorological conditions prevailed.

The helicopter was cleared for takeoff by the El Monte Air Traffic Control Tower (ATCT) and departed from a hover at the approach end of runway 19. Several witness reported that the pilot maintained the runway heading over the centerline. The takeoff and initial climb appeared normal. About 50 to 100 feet above the ground (AGL), and at 50 knots of airspeed, the helicopter rapidly descended to the runway in about a 35 degree nose down attitude with about a 30 degree right bank. The helicopter struck runway 19 about 1,900 feet from the departure end. An intense post crash fire erupted and the fuselage came to rest about 300 feet from the initial point of impact.

The accident occurred during the hours of daylight at latitude 34 degrees 05.16 minutes north and longitude 118 degrees 02.09 minutes west.

Crew Information

The pilot held a private pilot certificate, with a rotorcraft helicopter rating that was issued on January 22, 1991. The most recent third class medical certificate was issued to the pilot on January 13, 1992, and contained no limitations. On the application for the medical certificate, the pilot listed his total accumulated pilot time as 210 hours, with 50 hours accrued in the previous 6 months.

Complete flight records for the pilot were not recovered and the aeronautical experience listed in this report was obtained from a review of the airman FAA records on file in the Airman and Medical Records Center located in Oklahoma City. In addition, partial pilot logbook pages from Robinson Helicopters files and a portion of fire damaged logbook pages were reviewed.

According to the pilot/operator report submitted by the operator, the pilot's total aeronautical experience consisted of 640 hours, of which 60 hours were accrued in the accident aircraft make and model. In the preceding 90 and 30 days prior to the accident, the report lists a total of 50 and 10 hours respectively flown.

Safety Board investigators attempted to validate and reconstruct the pilot's experience. On November 11, 1992, the pilot attended a three day Robinson R22 Safety Course, conducted by Robinson Helicopters. At that time, the pilot indicated that he had accrued 425 hours of flight time. Robinson Helicopters required 500 hours of total time to act as pilot-in-command of an R44. On May 28, 1993, the pilot attended a one day R44 Pilot Check-Out training course at Robinson Helicopters. As verification that the pilot that accrued 500 hours of flight time, the pilot sent a facsimile of two pages of a logbook to Robinson Helicopters that reflected a total time of 525.2 hours. The logbook pages were not dated.

At the conclusion of the R44 training, the instructor pilot from Robinson Helicopters issued the pilot a certificate of training after finding that the pilot's performance was satisfactory to fly the R44 as a 2-place helicopter for at least 50 initial hours. To qualify to carry more than 1 passenger, the pilot needed to return for another flight check after gaining the required R44 flight time. On July 22, 1993, the pilot returned to Robinson Helicopters for his second flight check in the R44. The instructor pilot did not require the pilot to produce any verification that the pilot had accrued an additional 50 hours of R44 flight time. The instructor pilot verbally acknowledged that the pilot was authorized to carry more than one passenger; however, he did not issue a new certificate of training at that time.

The operator produced a copy of a pilot logbook recovered from the helicopter wreckage that was fire damaged. The logbook pages appeared to reflect flight time accumulated by the pilot in the R44. The record did not reflect the dates of various flights; however, the two pages totaled 35.0 hours.

Aircraft Information

The recording hour meter in the helicopter was destroyed. Examination of the fire damaged aircraft maintenance logbook revealed that the helicopter was manufactured on February 12, 1993. A 100-hour inspection was completed by Robinson Helicopters on May 21, 1993, at an hour meter reading of 104.0, 70 hours before the accident. The helicopter was purchased by the operator on May 27, 1993. At that time the helicopter had accrued 106 hours. The operator and a flight test mechanic for Robinson Helicopters reported that just prior to the accident flight, the helicopter had accumulated a total time in service of about 174 flight hours.

The last entry in the maintenance logbook was July 14, 1993, at a hour meter reading of 158.9. On that date the low rotor RPM warning horn unit was replaced and the main rotor blades were re-tracked by a Robinson Helicopter mechanic. The main rotor blade tracking was accomplished by adjustment of a trailing edge tab. No portion of the flight controls, swash plate assembly, or pitch change links were disturbed. The cyclic control assembly installed in the helicopter was revision H.

The fuel system is gravity-fed (no fuel pumps) from the main fuel tank. An auxiliary fuel tank (smaller and mounted higher than the main) drains into the main tank through an inter-connecting line. The tanks are metal and are mounted above a firewall separating the engine from the main transmission and fuel tank area. The main rotor mast from the transmission is mounted vertically between the two fuel tanks.

Fueling records at El Monte airport established that the helicopter was last fueled just prior to departure with the addition of 15.4 gallons of 100LL octane aviation fuel, which completely filled the main fuel tank. A mechanic for Robinson Helicopters reported that the main fuel tank gage indicated full and the auxiliary tank indicated less than 1/8 full.

Meteorological Information

The closest official weather observation station is El Monte, California, which is located at the accident site. At 1349 hours, a surface observation was reporting in part:

Sky condition and ceiling, scattered clouds at 18,000 feet; visibility, 7 miles; wind, 180 degrees at 8 knots; altimeter, 30.00 inHg.

Communications

Review of the air-ground radio communications tapes maintained by the FAA at the El Monte ATCT facility revealed that the aircraft communicated with the local control position. No unusual communications were noted between the local controller and the pilot during the review of the tapes.

Aerodrome and Ground Facilities

The El Monte airport is owned by The County of Los Angeles. The operation of the airport is contracted to Comarco Airport Services Inc. The published elevation of the airport is 296 feet mean sea level.

The airport is equipped a single hard surfaced runway on a 010 to 019 degree magnetic orientation. The runway is 3,995 feet long by 75 feet wide, and is equipped with medium intensity runway lights (MIRL), runway end identifier lights (REIL), and a visual approach slope indicator (VASI) lights. An Automatic Terminal Information Service (ATIS) weather broadcast is provided on a discrete frequency of 118.75 mhz. No formal crash, fire fighting, or rescue services or facilities are located on the airport, nor are any required. A small fire truck is stationed on the airport.

Wreckage and Impact Information

Safety Board investigators examined the wreckage at the accident site on July 31, 1993. The examination of the impact site revealed paint marks and ground scars in the runway surface, oriented parallel to the runway heading, about 9 feet west of the centerline, and about 2,000 feet from the approach end. Examination of the forward ends of the landing gear skid tubes revealed that the helicopter impacted in about a 35 degree nose down attitude and a 30 degree right bank.

Portions of fragmented outboard ends of the main rotors were located scattered along the left and right sides of the runway. The fragments displayed evidence of chordwise scratching, primarily to the underside of the blade fragments. A portion of a coiled electrical cord with an attached push-to-talk button was located on the runway about 90 feet prior to the initial impact point. Additionally, fragments of window plexiglass were also scattered along the sides of the runway just prior to the impact site.

At the impact site, ground impact marks from the main rotor blades were located progressively along the wreckage path at 23 feet, 14 feet, and 10 feet east of the west edge of the runway. The impact marks averaged about 1 inch in depth.

The forward ends of the landing gear skid tubes were separated at the forward cross tube attach points, which is about 3 and 1/2 feet aft of the tip. Both forward ends exhibited longitudinal scratching about 7 to 8 inches long on the underside of the tube, about 6 to 8 inches aft of their respective forward tip. The right side skid tube was separated at the lower attach points of the forward and aft cross tubes. The left side landing gear skid tube was separated at the lower end of the forward cross tube. The aft cross tube was still attached at the lower end of the left skid tube but was separated at the upper end of the lateral cross tube attach point. The separation was in an outward direction and displayed an aft twisting signature. The complete forward cross tube assembly remained intact; however, was separated from the fuselage and from both skid tubes. The lateral cross tubes did not exhibit any downward bending. All of the landing skid segments were located between the impact point and the main rotor mast assembly.

The main rotor mast assembly and main rotor blades, separated from the main rotor transmission and fuselage as one unit and came to rest about 250 feet from the impact point. The blades were attached to the rotor head and both exhibited "S" bending. The blades displayed extensive chordwise scratching, leading edge gouging, trailing edge compression, and tip destruction. The outboard portion of each blade, including the leading edge and tip weights separated from the blade structure. The blades tip weights were located about 800 feet east of the impact site.

The main rotor static mast separated at the base were it attaches to the main rotor transmission gear box. The static mast exhibited aft bending of about 8/32 inches, measured about mid-height between the top of the mast and the separated base. The rotating mast also separated at the base of the transmission and remained inside the static mast. It exhibited bending and torsional twisting signatures. The vertical push-pull tubes remained attached to the swash plate and were attached to the lower support jackshaft assembly. The jackshaft assembly separated at its mounting points on the main rotor gear case.

Examination of the red blade pitch-change link revealed that it separated into two pieces around the circumference of the upper end of the lower link. The lower link was attached to the swashplate. The upper end was attached to the separated portion of the red blade pitch horn.

The blue blade pitch-change link separated into three pieces. The upper link threaded portion was attached to the blue blade pitch horn and exhibited about a 90 degree bend and fracture near its lower end. The remaining portion of the threaded upper link and its corresponding lock nut that was normally threaded into the upper end of the lower link was not initially located. The interior threads of the lower link were undamaged.

The missing upper link threaded segment was located about 30 feet west of the main rotor mast assembly. The lower threaded end of the recovered segment was undamaged. The upper end exhibited bending and a fracture surface that matched the separation of the upper link. The outer surface of the lock nut exhibited gouging and grinding signatures. The lower link, along with the fork assembly, separated from the swashplate attach point.

The swashplate, at the blue blade pitch-change attach point exhibited a fracture at the outboard end bolt hole in an upward direction. The bolt hole was cracked completely through the casting at the bottom (downward) portion of the bolt hole. The teetering stops separated from the mast and were not located. The blade droop stops remained intact.

Indentations produced by swashplate contact with the sliding uniball sleeve were noted. One side exhibited an indentation about 2 inches above the bottom edge of the sleeve. An indentation on the opposite side of the sleeve was found 1 and 3/4 inches above the bottom of the sleeve. According to the manufacturer, the dents correspond to a collective up position of 90.8 percent and 67.9 percent respectively. Disassembly of the rotor mast revealed bending of the rotor mast adjacent to the mast data plate of about 0.300 of an inch in the direction of the red blade attach point.

The vertical main rotor control push-pull tubes were attached to the jackshaft assembly. They were separated at the lower mixing bell crank. The bell crank assembly was separated from its fuselage mounting points and exhibited fire damage. The torque tube from the mixing bell crank forward to the cyclic control tube was attached at both ends; however, the center portion of the tube was destroyed by fire. The push-pull tube from the mixing bell crank to the cyclic control tube was attached by its rod-end bearings at both ends; however, the tube was fractured at both ends of their respective rod ends. The center portion of the push-pull tube was destroyed by fire.

The forward end of the collective control push-pull tube to the collective support assembly was destroyed by fire. The aft end of the collective control push-pull tube was fractured and burned. The engine governor switch located at the forward end of the pilot's collective control was found in the "ON" position. The tail rotor anti-torque push-pull tubes had numerous fractures and fire damage. All of the observed fracture surfaces were oriented on numerous 45 degree angle planes consistent with overload signatures.

The cyclic control assembly was recovered from the burned cockpit area and exhibited impact and fire damage. A fracture was noted below the attached lateral control torque tube and above the forward attach point of the fractured longitudinal control push-pull tube. The area of fracture occurred at the point where the cyclic stick transitioned from a steel tube to a welded steel box structure.

The longitudinal trim motor arm was measured at 15/16 of and inch in the direction of maximum trim force applied on the longitudinal axis. The manufacturer reported that this corresponds to 28 percent of the maximum forward trim setting. The lateral trim motor arm was measured at 1 and 5/16 inches in the direction of right lateral trim. This corresponds to 84 percent of its maximum right trim setting. All of the elastic cord assemblies associated with the trim system were destroyed by fire. The cyclic mounted lead shot pouch utilized as a vibration dampening system was destroyed by fire.

The aft end of the tail boom, with the vertical and horizontal stabilizers attached, was located laying on the main rotor mast assembly. The stabilizer assembly had separated from the aft end of the tail boom and the fracture surfaces were oriented on numerous 45 degree angle planes. The lower portion of the vertical stabilizer exhibited a semi-circular indentation and black paint smug on the leading edge, about 6 inches above the lower end.

The horizontal stabilizer exhibited a downward bend about mid-span of about 30 degrees and had paint removal and scraping on the upper surface of the outboard end of the stabilizer. The upper end of the vertical stabilizer exhibited about a 20 degree bend to the left and paint removal and scraping on the right side of the stabilizer. The tail boom exhibited fire damage and melting to the forward attach points, smoke damage along the right side, and scrape marks on the left side at the aft end. The underside of the tail stinger did not exhibit any scraping of the paint.

The tail rotor gear box separated from its mounting flange on the aft tailcone bulkhead. The gearbox was located between the main rotor mast assembly and the fuselage point of rest. The tail rotor blades remained attached to the gear box. Both blades were destroyed and separated about 10 inches outboard of their respective attach points. The tail rotor pitch control push-pull tube was fractured at the forward and aft ends of the pitch control bell crank assembly. The tail rotor drive shaft flex plate exhibited torsional fractures oriented in the direction of rotation at the gear box input spline. The forward end of the tail rotor drive shaft exhibited torsional twisting and fire damage at the tail boom attach point.

The tail rotor drive shaft was intact from the upper sheave aft about 2 feet 4 inches, to about the tail boom attach point were it was damaged by fire. The free wheeling unit located in the upper sheave assembly functioned properly. Remnants of the engine drive belts remained between the upper and lower sheaves and were fire damaged as was the alternator belt. The belt tensioning actuator assembly was extended about one inch. The manufacturer reported that the measurement corresponds to the belts being in the normal tightened range.

The drive shaft from the upper sheave to the main rotor transmission gear box was intact. The forward side of the gear box contained a 1 inch by 2 inch hole in the case. Small portions of metal were present in the beveled gear teeth, preventing 360 degree rotation of the transmission. Disassembly of the gear box revealed no other damage and proper rotation of the transmission was performed.

The engine remained attached to the fuselage frame at its respective attach points. The exhaust tube was crushed in an upward direction at the aft end. The cooling shroud around the lower mounted sheave and cooling fan were fire damaged. Valve and gear train continuity and thumb compression was established by hand rotation of the engine. The suction and pressure oil screens were free of contaminants. Disassembly of the accessory gear case revealed proper lubrication. The oil pump gears rotated normally and exhibited a bright, clean appearance. The engine cylinders were removed and examined. No unusual conditions were noted.

The magnetos were fire damaged. Hand rotation could not produce any spark at any terminal. The wiring harness was destroyed by fire. The spark plugs did not exhibit any unusual combustion signatures. The top and bottom plugs from the numbers 2, 4, and 6 cylinders were oil coated.

The carburetor and air box remained attached to the engine. The fuel valve was intact and found in the "ON" position. The metal floats of the carburetor and the venturi were undamaged. The fuel screen was free of contaminants. The fuel strainer was fire damaged. The throttle and mixture levers at the carburetor were found in an intermediate position.

The post crash fire incinerated most of the cabin/cockpit area. Safety Board investigators, however obtained some instrument readings (see Supplement B for details).

Both fuel tanks sustained severe fire damage and both fuel caps were in place. The upper, inboard longitudinal edge of the auxiliary fuel tank, exhibited a semi-circular dent in an outboard direction that was consistent with the diameter of the main rotor mast.

The Aviation Safety Engineering Branch of the Canadian Aviation Safety Board produced a Light Bulb Filament Impact Dynamics Study in support of aircraft accident investigation. The report is one of several studies of light bulb filaments used as indicator and warning lights in aircraft. Research has provided a basis for determining the condition of the light bulb at the moment of impact. While illuminated, the tungsten filament is ductile from the heat produced when the bulb is on. Impact forces can produce a stretched and uncoiled filament, exhibiting a "snaking" appearance. If the bulb is off, the filament is cold and impact forces produce brittle fractures of the filament, without stretching.

The annunciator bulbs from the instrument panel were examined. The examination revealed that the ENGINE LOW OIL PRESSURE bulb filament was stretched and uncoiled. The "CLUTCH" bulb was damaged during removal and its filament orientation could not be determined. The remaining annunciator bulb filaments exhibited no evidence of stretched or uncoiled filaments.

Medical and Pathological Information

A post mortem examination of the pilot and passengers was conducted by the Los Angeles County Coroner's Office on August 1, 1993. No pre-existing conditions were noted during the autopsy which would have adversely affected the pilot's abilities to pilot an aircraft.

Toxicological examinations of the pilot were conducted by the FAA's Civil Aeromedical Institute (CAMI) on November 2, 1993 and was negative for all screened drugs and alcohol.

Fire Aspects

Witnesses reported that flame and smoke was observed generally at the bottom of the fuselage as the wreckage was still sliding to a stop from the point of impact. An intense post crash fire immediately engulfed the wreckage at the point of rest.

The County of Los Angeles provides El Monte airport with an Ansel 440, fire fighting vehicle. The vehicle is equipped with 110 gallons of aqueous film forming foam (AFFF), and 450 pounds of Purple K dry chemical extinguishing agent. Both systems are independently charged by a nitrogen gas cylinder. A single hose reel assembly is provided with independent internal hoses for each system and separate nozzles for each fire suppression agent. The systems can be charged by pushing a plunger, lifting a lever, or opening a valve on the nitrogen bottle. A Comarco Inc. employee was on duty and responded to the accident site with the fire truck.

Witnesses indicated that after the fire truck arrived at the scene, there appeared to be some confusion about activation of the nitrogen system and a delay in application of the fire suppression agents.

Tests and Research

Following the accident, fuel samples from the nozzle of Comarco Inc.'s fuel truck were obtained and tested by the Exxon Company. They reported that the sample was clear and bright and visually free of water, sediment and suspended matter. The sample met the standards for aviation gasoline.

The main rotor swashplate assembly, including the red and blue pitch- change links were submitted to the National Transportation Safety Board Materials Laboratory for examination. All of the fracture surfaces exhibited evidence of overstress separations. Hardness testing of the submitted components conformed to design specified materials.

On January 26, 1994, the operator submitted the aircraft's cyclic control tube assembly to the National Transportation Safety Board Materials Laboratory for examination. The examination revealed fractures through two flat side plates of the lower cyclic assembly without adjacent material deformation. After removal of soot and oxidation, it was noted that several of the fracture planes were flat and intersected the surface at a 90 degree angle consistent with a brittle fracture mechanism such as fatigue cracking. Optical and scanning electron microscope examinations revealed fracture surfaces that exhibited a river pattern consistent with fatigue cracking.

Additional Information

In addition to the persons listed on page 5 of this report, the following persons participated in this investigation:

Frank Robinson, Robinson Helicopters Inc., Torrance, CA. Ronald Hamilton, Robinson Helicopters Inc., Torrance, CA. Kurt Robinson, Robinson Helicopters Inc., Torrance, CA. Donald Skunberg, FAA-WP-LAX-FSDO, Los Angeles, CA. Lirio Liu, FAA Aircraft Certification Office, 3229 E. Spring St., Long Beach, CA., 90806.

Robinson Helicopters reported that following the accident, they were in possession of 3 R44 aircraft and examined the cyclic control assemblies of the 3 helicopters. The results of a visual examination revealed:

Aircraft serial number 001; total time on aircraft, 120 hours; cyclic control assembly C320-1 revision G, free of cracks; torque tube C319-1 revision F, free of cracks. Aircraft serial number 002; total time, 258 hours; torque tube C319-1 revision F, free of cracks. Aircraft serial number 004; total time, 796 hours; cyclic control assembly C320- 1 revision G, free of cracks.

Robinson Helicopters also reported that following the accident, more than a dozen different parts in the cyclic control system were redesigned. The initial effort to strengthen 5 accident damaged control system components resulted in them individually being considerably stronger than the rest of the system. Continued fatigue testing and redesign resulted in the entire cyclic system being strengthened. The following parts were redesigned:

Collective fork; cyclic pivot assembly; 2 yokes; torque tube; cyclic stick assembly; jackshaft assembly; cyclic control system; 4 jackshaft support points.

Wreckage Release

The Safety Board initially released the wreckage, located at Lynn's Aircraft, El Monte, California, to the owner's representatives on September 21, 1993. The swashplate assembly and pitch-change links were retained by the Safety Board for examination until its release on November 10, 1993.