On October 14, 1999, at 1631 hours Pacific daylight time, a Bell 206B, N16889, collided with the airport ramp during an attempted landing at the Los Angeles International Airport, Los Angeles, California. The helicopter, operated by TSR Helicopters Inc., Van Nuys, California, sustained substantial damage. The commercial pilot and two passengers were seriously injured. The personal flight, conducted under the provisions of 14 CFR Part 91, originated at the Van Nuys Airport about 1600, and was en route to the Los Angeles Airport to drop off one of the passengers. Visual meteorological conditions prevailed and no flight plan was filed.

A Federal Aviation Administration (FAA) inspector from the Los Angeles Flight Standards District Office interviewed the pilot and passengers in the hospital. The pilot stated that he reported his position to the Los Angeles Air Traffic Control Tower when he was about 1 mile north of the runway 24 complex. He was given instructions to cross the extended centerline of the parallel runways at 1,500 feet. After crossing the runways, and once abeam the elevated helipad, the pilot made a slight left turn away from the pad to allow for more spacing in the descent. He stated that he had performed that particular approach several times and was familiar with the flight profile and procedure.

The pilot reported that to maximize the descent rate, he lowered the collective to flat pitch, and momentarily entered a transient overspeed condition of the N2 turbine; the N2 needle climbed to 103 percent for about 4 seconds. He explained that a transient overspeed in that helicopter allows up to 107 percent rpm overspeed for 15 seconds with torque settings up to 32 percent. The pilot reported that he lowered the nose of the helicopter to stop the N2 increase.

The pilot turned right toward the helipad and descended at a rate of about 1,300 feet per minute until he intercepted a standard approach profile. He reported that as he reapplied power and increased collective pitch, the rotor rpm immediately drooped to 90 percent. The low rotor rpm horn and light came on and the pilot fully lowered the collective. He estimated that his altitude at that time was between 100 to 150 feet agl, and his airspeed was about 60 knots. He declared an emergency to the air traffic control tower and turned left toward an open area between terminals 5 and 6.

The pilot realized he might not have sufficient altitude to clear the terminal buildings, so he pulled collective pitch, which further decayed the rotor rpm but provided additional lift. As he descended to his intended landing point, a twin-engine turboprop airplane turned into the ramp area and the pilot veered to the right to avoid it.

The pilot stated that he fully lowered the collective and realized that he'd have to perform a run-on landing with little or no flare. He pulled all remaining collective pitch before impacting the ground. The helicopter contacted the ground in a tail-low attitude, slid forward approximately 60 feet, and made a 180-degree turn. The right front chin bubble shattered on impact. Both landing gear skids were forced upward through the cabin. The helicopter came to rest on its belly.

The pilot attempted to shut the engine off by rolling the throttle past the idle cutoff position but he was not able to manipulate the throttle. He turned off the fuel shutoff valve, generators, and battery.

The pilot stated that he had not experienced any problems with the helicopter prior to the accident.


According to the FAA airman certification database, the pilot held a commercial pilot certificate with a rotorcraft-helicopter rating. He also held a flight instructor certificate for helicopters. The pilot reported that he had accrued approximately 1,300 hours of helicopter flight time, of which approximately 980 hours were in Bell helicopters. He had completed a 14 CFR Part 135 airman competency/proficiency check on July 21, 1999. At the time of the accident, he was seated in the right front seat.

The rear right seat passenger held an airline transport pilot certificate for rotorcraft-helicopters. He also held a flight instructor certificate for helicopters and instrument helicopters. He reported that he had approximately 4,000 hours of helicopter flight time, of which approximately 1,500 hours were in Bell helicopters.

The left front seat passenger held a Canadian commercial pilot certificate with a helicopter rating. He reported that he had approximately 150 hours of helicopter flight time, of which 60 were in Bell helicopters.


The maintenance records were reviewed. They reflected that the helicopter and engine had undergone a 100/200/300-hour inspection on September 3, 1999. The records did not reveal any unresolved squawks or discrepancies. The helicopter had a total time of 7,897 hours at the time of the accident.


The entire fuselage structure was deformed, with the most severe deformation between STA. 130 and STA. 205. All four doorposts had been yielded outward and exhibited partial separations of the structure. The roof structure had been displaced downward, with the largest displacement occurring in the aft cabin area. The entire underside of the fuselage exhibited deep scoring primarily along the longitudinal axis. The lower wire strike protection system had been deformed upward into the center nose section. The forward windscreens and chin bubbles were broken out.

The landing gear was a high skid configuration with after-market auxiliary steps installed. Neither skid displayed any indication of having spread upon impact; they were not spread outward, but had been driven upward through the fuselage. The left side aft landing gear cross tube and supporting structure penetrated the fuel bladder, reportedly causing a substantial fuel spill. The right side aft cross tube penetrated into the fuselage up to the base of the fuel filler port. The forward cross tube separated from the fuselage attach point on the right side of the aircraft.

One of the main rotor blades exhibited a dent approximately 2 inches in diameter about 6 feet outboard from the root end. The outboard tip end of the blade exhibited minor chordwise scoring. The other main rotor blade displayed chordwise scoring, as well as red and white paint marks.

The tail boom exhibited upward compression buckles and the lower end of the vertical fin displayed upward crushing and had been bent to the right. Both tail rotor blades displayed compression buckling. One tail rotor blade separated approximately 6 inches from the root end of the blade. The separated blade exhibited chordwise striations. Tail rotor pitch control was verified by manually exercising the pitch control tube inside the tail boom. The tail rotor gearbox rotated freely. The magnetic chip detector plug was absent of debris and the gearbox contained oil.

The main drive shaft was securely attached at both the engine output side and the transmission side. The coupling temperature indicators appeared normal and exhibited no indication of an over-temp condition. The first tail rotor drive shaft exhibited a rotational score and had been separated approximately midspan.

The transmission and mast assembly did not exhibit any external damage. The transmission drive system and tail rotor output rotated freely during the examination. Rotation of the drive shaft in the freewheel direction established normal operation of the freewheeling unit. The lower transmission chip plug was examined and a small amount of ferrous debris was observed.

The aircraft was equipped with dual flight controls. The collective was in the full down position and the throttle was in the fully open position. The right anti-torque pedal support was separated. The left side anti-torque pedals had approximately 1.5 inches of right pedal input. The cabin floor was deformed upward and a portion of the separated cross tube was observed penetrating into the cabin area on the left side.

The right front seat exhibited significant deformation of the seat frame and seat pan structure. The left front seat pan exhibited deformation primarily on the left outboard seat portion. The right aft passenger seat, occupied by one of the passengers at the time of the accident, did not display any abnormalities. All seat restraint systems were intact and appeared to latch properly. The filaments for the caution and warning light bulbs were examined. The light bulb filament for the "engine out" warning was intact was not stretched. The "low rotor rpm" light bulb filament was partially stretched.


The engine was securely mounted to the airframe and exhibited no external damage.

The fuel cell displayed extensive damage. The fuel bladder was torn open along the backside. The airframe fuel filter was removed and examined. The filter bowl was full of fuel. A small amount of particulate was noted at the bottom of the bowl. The fuel boost pumps were examined with no noted discrepancies.

All connections for the oil supply and scavenge lines were secure and there was evidence of oil in the lines at the connections. The oil tank contained oil. All pneumatic lines were intact and secure. Fuel was present in all the fuel lines and a vacuum check was performed on the fuel system. The system maintained vacuum for the duration of the 2-minute test and no leaks were detected.

The N1 and N2 sections rotated freely. The compressor rotated freely.

The engine was removed and run in a test cell. Due to a breakdown of the test cell, the engine run could not be completed. Immediately before the test cell failed, an uncommanded decrease in rpm was observed.

The fuel control unit and power turbine governor were removed and taken to Allied Signal for additional testing. The static bench testing of the fuel control unit and power turbine governor indicated that both components were slightly out of calibration. According to Rolls Royce, the irregularities noted on the power turbine governor were consistent with a high time governor and/or impact or vibration damage. The governor had 246.5 hours remaining before the next overhaul. The Rolls Royce investigator reported that the discrepancies noted might result in a 2 percent increase or decrease of N2 rpm with a corresponding reduction or increase of the collective lever. He stated that the fuel control discrepancies were insignificant deviations.

The engine was then transported to Scottsdale, Arizona, where a complete and normal run was accomplished. During the test run, several aggressive accelerations/decelerations were conducted with no discrepancies noted. The test included: initial idle; seal run in to maximum power; five power point checks; governor droop; and a final idle setting. The engine performance met all required specifications. A copy of the performance data is appended to this file.

The power turbine governor was bench tested again by Allied Signal.


A Pilot/Operator Aircraft Accident Report, Form 6120.1/2, was sent to the pilot three times; once on October 20, 1999, again on December 9, 1999, and again on January 10, 2001. The pilot did not respond.

The wreckage was released to a representative of the owner on February 17, 2000.

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