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HISTORY OF FLIGHT
On October 14, 2000, at about 1208 eastern daylight time, a Robinson R-22, N4004J registered to a private owner, operated by Volar Helicopters Inc., as a 14 CFR Part 91 personal flight, crashed on a street in the vicinity of Pembroke Pines, Florida. Visual meteorological conditions prevailed and no flight plan was filed. The helicopter was destroyed. The airline-transport pilot and one passenger were fatally injured. The flight originated from Fort Lauderdale Executive Airport, Fort Lauderdale, Florida, about 30 minutes before the accident.
Two witnesses stated they were stopped at the intersection of Pines Boulevard (Blvd), and SW 155th Street. Both witnesses heard a popping sound as the helicopter approached their location traveling eastbound. The helicopter was estimated to be at about 100 feet. The helicopter yawed from left to right. One witness stated the nose pitched up to a nose-high attitude and started descending rapidly left side low. The helicopter cleared the power lines and crashed. Another witness was driving eastbound on Pines Blvd. and initially observed the helicopter traveling southwest bound turning left towards the east at about 65 feet. The helicopter was observed to assume a nose high, descending left skid low attitude. The helicopter cleared the power lines at SW 155th Street and collided with the westbound lane of Pines Blvd. in the same attitude. As soon as the helicopter collided with the highway, the nose was observed to pitch down, proceeded forward, and turn to its left coming to rest on its left side. A jogger, who was running eastbound on the north side of Pines Blvd., stated he heard a helicopter approaching him from the rear. The rotor blades were popping as they were whipping up the air. He looked up and observed the helicopter about 50 feet over the trees. The helicopter cleared the power lines at SW 155th Avenue, touched down left skid low, went forward, turned counter clockwise and came to a stop on its left side facing the northwest.
Review of records on file with the FAA Airman Certification Branch, Oklahoma City, Oklahoma, revealed the pilot was issued a temporary airman certificate on January 4, 2000, for airline transport pilot with ratings and limitations for airplane single and multiengine land, commercial pilot rotorcraft helicopter. The pilot indicated on her airman certificate and or rating application dated January 4, 2000, that she had accumulated 52 total hours in the Robinson R22 helicopter, of which 6 hours were solo, and 24 hours were as pilot-in-command. Review of her pilot logbook revealed that she had accumulated 9,415.3 total hours in all aircraft of which 7,986.6 hours were as PIC. Her last recorded flight in an airplane was on February 2, 1996. Her first recorded flight in an R22 was on September 28, 1999. She had accumulated 67.1 hours in the R22, of which 48.6 hours were dual, and 18.5 hours were as PIC. She had flown 3.2 total hours in the last 90-days before the accident. Her last flight before the accident was on September 15, 2000, for .7 hours. Her last flight 60 days before the accident was on August 22, 2000, for 1.0 hour. Her last flight 90 days before the accident was on July 27, 2000, for 1.0 hours. Her last recorded autorotation was on June 29, 2000.
The FAA designated examiner that conducted the commercial pilot helicopter add on rating for the deceased pilot stated the oral exam lasted about 3.2 hours and the flight evaluation lasted about 1.4 hours. Upon completion of the evaluation, he informed the pilot that she met the standards for the add on rating, however her autorotations were on the edge, but within tolerance. He recommended that the pilot obtain some additional instruction, to which she agreed. An instructor pilot for Volar Helicopters stated he conducted a check ride with the deceased pilot on June 29, 2000, and she completed emergency procedures pertaining to the R22, which included autorotations that were conducted satisfactory.
The helicopter was a Robinson Helicopter model R22, serial No. 1320, registration No. N4004J, manufactured in 1990. The helicopter is equipped with a Lycoming O-320-B2C, 131 horsepower engine. The engine was removed from helicopter serial No. 0126 on October 28, 1996, and had accumulated 2,010.51 hours. It was overhauled by Textron Lycoming in Williamsport, Pennsylvania, on January 2, 1997, and installed in helicopter serial No.1320 on April 11, 1997. The Hobbs time was 6,007.6. The engine total time was 2,019.2 hours and 8.4 hours since major overhaul (S.M.O.H). The last recorded 100-hour/annual inspections were on September 25, 2000. The Hobbs time was 7,683.4. The engine total time was 3,695.0 hours and 1,684.49 S.M.O.H. The Hobbs time on the helicopter before the accident was 7,740.8, and the Hobbs meter at the accident site was 7,741.3.
Textron Lycoming issued Service Bulletin No. 388B (SB), on May 13, 1992. The SB provides procedures to determine exhaust valve and guide condition, and indicates that helicopter engines should be inspected at 300-hour intervals. The SB states, "To insure positive and trouble free valve train operation, the inspection procedure described in this publication should be accomplished as recommended in the Time of Compliance section of this publication. Failure to comply with the provisions of this publication could result in engine failure due to excessive carbon build up between the valve guide and valve stem resulting in sticking exhaust valves or; broken exhaust valves which result from excessive wear (bell-mouthing) of the exhaust valve guide." Review of the engine logbook revealed that SB No. 388B was conducted on April 30, 1998, at 598.2 S.M.O.H., June 19, 1998, at 694.6 S.M.O.H., and on September 30, 1999, at 1,292.0 S.M.O.H.
An individual who had flown in N4004J stated the helicopter had a history of high oil temperature and low oil pressure, and that the oil cooler had been changed. Review of invoices revealed the following service description and corrective action taken by Cav-Air Inc., maintenance personnel on N4004J:
Invoice No. H3047-2, 06/19/00
4. Engine Oil Temp Very High. Changed oil and swapped temperature sensor and thermostat with another aircraft. Cleaned oil cooler. Test flew and found ok.
Invoice No. H3316-1, 10/18/00
1. Engine Oil Temp Runs At High Green. Date 08-01 Inspected oil screen and installed serviceable oil screen housing o/m 69510. Run up and test flew. Found oil temp at high green. Increased oil pressure and made sure oil level was at 6 qts and OAT was 97 degree. Re-installed original oil filter housing and leak check ok. Aircraft returned to service.
2. Pilot States Engine Oil Runs High Green. Date 08-07 Removed oil screen and found no metal. Removed oil cooler and installed serviceable oil cooler and toped off oil. Test flew ok. Oil temp is still in high green but OAT is 97 degrees. Removed cooler and re-installed original cooler. Ran up and leak checked good. Aircraft returned to service.
3. Engine Oil Temp Runs High Green. Date 08-09 Spoke to Lycoming Rep and they recommend we check the oil pump for leakage. Removed front cover and found oil pump and hardware secure and no evident wear. Re-installed cover and installed magnetos and timed. Run up and leak checked good. Test flew and found oil temp in 3/4 temp green range. Aircraft returned to service.
The nearest weather reporting facility at the time of the accident was Fort Lauderdale Executive Airport, Fort Lauderdale, Florida. At 1553Z, surface weather observation was: 5,500 scattered, visibility 10 miles, temperature 82 degrees Fahrenheit, dew point temperature 57 degrees Fahrenheit, wind from 040 degrees at 8 knots, and altimeter 30.10. Visual meteorological conditions prevailed at the time of the accident.
WRECKAGE AND IMPACT INFORMATION
The wreckage of N4004J was located in the westbound lane on Pines Blvd. about 143 feet east of the intersection at SW 155th Avenue in the vicinity of Pembroke Pines, Florida.
Examination of the crash site revealed N4004J collided with the highway in a nose high, left skid low attitude, and separated both tail rotor blades. The helicopter continued forward and the main rotor blade serial No. 9804C collided with the road 25 feet 2 inches down the crash debris line. The empennage separated from the tailcone at the aft casting 50 feet 7 inches down the crash debris line. The lower vertical stabilizer was displaced to the right about 45 degrees. The tailskid was bent aft and to the right. Scoring was present on the lower and left surface of the tailskid. The horizontal stabilizer was relatively straight. The upper vertical stabilizer was displaced to the left about 35 degrees. The helicopter rotated to the left about its vertical axis and came to rest on its left side on a heading of 330 degrees magnetic 93 feet 3 inches down the crash debris line. The main fuel tank ruptured, and the auxiliary fuel tank was punctured in four places. The forward crosstube and both struts were attached to the right side of the airframe. Both skid toes were separated at the forward struts. The rear crosstube, both rear struts, the aft section of the right skid remained together, and were attached to the steel frame. The left skid was separated at both struts, and the left rear strut was bent forward and upward. The forward bay of the tailcone remained attached to the steel tube frame. The remaining segments of the tailcone were separated at the Nos. 2, 3, 4, and 5 bulkheads, and were held together by the tailrotor driveshaft, control tube, and wires. The tail rotor visual guard was bent aft and to the right. The aft section of the tailcone exhibited evidence of torsional twisting. One tailrotor blade (serial No. 8835E) separated 8 inches from the tailrotor hub center, and the remaining tailrotor blade (serial No. 8829E) separated 7 inches from the tail rotor hub center. There was no evidence of a main rotor blade strike on the tailcone assembly.
The cabin structure lower left quadrant is crushed from the nose extending aft to the vertical firewall. The crush line is at a 45-degree angle about the longitudinal axis relative to the horizontal. The left seat structure is crushed downward and to the left. The right seat sustained lateral displacement to the left. The upper and lower steel frame and engine exhibits distortion from the lower left quadrant consistent with the crushing on the cabin assembly. The governor control assembly (consisting of right engine magneto, B247-5 governor assembly, and the B286-2 governor controller assembly) were removed from the helicopter and forwarded to the manufacturer through an NTSB Regional Office for further analysis.
Examination of the B247-5 governor assembly revealed it had sustained impact damage. The motor was connected to a current limited power supply, and 6.8 volts was applied in both directions. The assembly made noise but the output did not turn. The gearmotor canister was partially removed and the misalignment stress was removed. The assembly was manually held in place while voltage was applied, and the output arm turned. The B247-2 gear motor assembly was removed from the rest of the B247-5 assembly, and voltage was applied. The gear motor assembly rotated about 90 degrees each way with voltage applied. The output stage of the planetary gearbox was removed and revealed that two of the gears had broken teeth, the output drive pins, and output plate were bent, and all but one gearbox assembly screws were broken. Voltage was applied to the gear motor and the motor ran continuously in both directions. The B247-5 gearbox (which includes the friction clutch) was taken to a Robinson Helicopter Company sub assembly department and the friction clutch was tested. The gearbox was mounted in the production test fixture, the output arm was locked in place, the drive bolt was rotated 20 times in each direction, and the friction torque was measured. In the clockwise direction, the friction (drag) was 10 in-lbs. In the counterclockwise direction, the friction (drag) was 9-10 in-lbs. The undamaged portions of the governor assembly appeared to be functioning normally.
The right engine magneto was mounted in a test fixture and the magneto's spark generating performance was normal. An oscilloscope was connected to the tachometer points, and the magneto was run at 2564 RPM. The output of the tachometer points indicated the points were producing a signal that was within the requirements for governor operation.
The B286-2 governor assembly sustained impact damage. From the external deformation of the box, it was apparent that the circuit board was damaged, so an immediate function of the controller was not attempted. The cover of the box was removed, and then the box was partially cut open. The circuit board had been kinked and cracked in several places. A few components had suffered from contact with the distorted outer box, and one integrated circuit chip had its cap broken or popped off. Portions of the circuit were functionally checked by using test probes in various locations on the circuit board. Portions of the circuit were checked, in sequence from input to output, until the portions of the circuit that were non-functional were reached. The error signal portion of the circuit was reading 1.3 percent rpm lower than its calibration setting of 104 percent. The non-functional circuit portions appeared to be linked to the physically damaged components. (For additional information see Extract Robinson Helicopter Accident Report, an attachment to this report.)
Examination of the main rotor system and tail rotor revealed no evidence of a precrash mechanical failure or malfunction. The pitch change link/antirotation linkage for main rotor blade (serial No. 9815C) separated from the upper swash plate arm. The main rotor blade rotated 180 degrees around its pitch change axis. Scoring was present on the lower surface extending inboard 75 inches from the blade tip. The main rotor blade was conned upward about 45 degrees. The lower surface of the main rotor blade exhibited scoring about 30 degrees relative to the chord line. The scores at the inboard extent of the scoring were oriented about 40 degrees from the chordline. Rotor blade (serial No. 9804C) is conned upward 20 degrees. The blade is bent aft and upward, and no scoring was present on the lower surface. The spindle assembly and bearings were removed from both main rotor blades. Main rotor blade serial No. 9804C contained spindle serial No. 8434 and main rotor blade serial No. 9815C contained spindle serial No. 8435. The bearings had evidence of light polishing on the races where the balls ride, but had no brinelling or spalling or damage from the bearing balls. There was no indication of cracking of outer bearing races or brinelling of bearing races. (For additional information see Extract Robinson Helicopter Company Accident Report, an attachment to this report.)
Examination of the airframe, and the flight control assembly revealed no evidence to indicate a precrash mechanical failure or malfunction. All components necessary for flight were present at the crash site. Continuity of the flight control system was confirmed for pitch, roll, and yaw.
Examination of the main transmission, overrunning clutch, tail rotor transmission, chip detectors and belt tension actuator revealed no evidence of a precrash mechanical failure or malfunction. Both drive belts were intact and exhibited no evidence of rolling.
Examination of the engine assembly revealed the No. 3 engine cylinder exhaust valve was broken at the stem. The No.3 cylinder with four aluminum fragments, piston, exhaust valve guide pieces, upper and lower exhaust valve spring seat pieces, exhaust valve springs, exhaust valve pieces, exhaust valve stem cap, exhaust push rod, exhaust rocker arm, camshaft, and 8 tappets were forwarded to the NTSB Materials Laboratory for further analysis. The exhaust valve head (portion) and exhaust valve keys were not located or recovered. Examination of the components by the NTSB Materials Laboratory Division revealed:
Cylinder: A hole was observed in the number three cylinder. The hole was adjacent to the exhaust valve stem through hole and had a diameter approximately equal to that of the exhaust valve stem. Material was fractured from the cylinder head around the hole, and the mating fragments. An impression corresponding to contact with the tip of the exhaust valve stem was observed on the inboard side of the largest fragment. Multiple impressions were observed on the inboard surface of the cylinder head (dome area) corresponding to contact with portions of the exhaust valve. Cooling fins adjacent to both the intake and exhaust push rods were damaged.
Piston: Multiple impressions were observed on the No. 3 piston dome. The impressions corresponded to contact with the head and fractured shank portions of the exhaust valve.
Valve Guide: The outboard end and three pieces from the inboard end of the valve guide were submitted separated from the cylinder. The outboard piece measured 1.5 to 1.9 inches in length. The outboard 0.85 inches was sooted and had circumferential scratches. The outboard end was damaged and cracked. The fracture surfaces on the inboard end and on the three pieces were rough and irregular, consistent with overstress fracture.
Lower Exhaust Valve Spring Seat: Two radial fractures were observed approximately corresponding in angle to the fragment fractured from the cylinder head on the lower exhaust valve spring seat. A piece of the spring seat was missing from the outer diameter opposite to the radial fractures.
Exhaust Valve Springs: The exhaust valve springs are deformed at one end, and the outer spring appeared to be undeformed.
Upper Exhaust Valve Spring Seat: A piece of the outer diameter of the upper exhaust valve spring seat was missing and corresponded to contact with the neck of the exhaust rocker arm. Multiple impressions were observed on the outboard surface at the inner diameter. The impressions corresponded to contact with the edge of the exhaust valve stem cap.
The keeper retaining lip at the inboard edge inner diameter (at the exhaust valve keeper location) was fractured and missing. Portions of the fracture surface were obliterated by post-fracture damage. The undamaged areas had an intergranular fracture surface near the edges and a rough, somewhat dimpled appearance between, features consistent with overstress separation. Deformation associated with the fractures was consistent with an inboard motion of the keeper retaining lip relative to the remainder of the exhaust valve spring seat. No evidence of fatigue or any other progressive crack growth was observed.
Exhaust Valve: Two pieces from the stem portion of the exhaust valve were examined. The inboard end of the exhaust valve stem appeared duller and rougher than the outboard end, but no material transfer from the exhaust valve guide was observed. The inboard end of the exhaust valve stem (closer to the head of the valve) was bent and was fractured in two locations. The fracture surfaces at both locations were perpendicular to the longitudinal axis and had a faceted appearance. Examination using scanning electron microscopy showed intergranular fracture surfaces consistent with high temperature overstress failure.
Impressions and damage were observed on the tip of the valve stem. A deformation lip was observed on the inboard side of the valve stem tip (in the keeper contact area) at the edge. Typically, deformation at this location is associated with contact with the keeper and has a radial (in relation to the centerline of the valve stem) and outboard direction. However, the lip was deformed radially and inboard, consistent with the typical deformation pattern (a radial and outboard deformation) followed by an inboard deformation associated with contact with the cylinder head.
Exhaust Valve Cap: The exhaust valve cap appeared undamaged.
Exhaust Rocker Arm: Arc-shaped impressions were observed on the rocker arm neck adjacent to the pad area, corresponding to contact with the upper exhaust valve spring seat. The wear surface on the rocker pad, generally appeared smooth and shiny with small impressions scattered throughout the surface. One area of the wear surface nearest the neck, appeared duller and gray.
Push Rod: The exhaust push rod was bent approximately 7 degrees.
Tappets: No excessive wear or damage was observed for any of the tappets.
Camshaft: No excessive wear was observed on any of the cam lobes. Two areas of damage were observed near the apex of the lobe of the cam for the number one and two intake valves. The damage appeared to be limited to near the surface. According to the manufacturer, this type of damage is typical for a high-time engine and may result from a cold start. No other areas of damage were observed on the other cams. (For additional information see NTSB Materials Laboratory Factual Report No. 01-025 an attachment to this report.)
The hydraulic plunger assemblies were forwarded through the FAA for examination by Textron Lycoming. All of the hydraulic plungers tested within the required Textron Lycoming specifications. (For additional information see Textron Lycoming letter dated March 19, 2001, an attachment to this report.)
MEDICAL AND PATHOLOGICAL INFORMATION
Dr. Linda Rush, Associate Medical Examiner, District 17, Fort Lauderdale, Florida,conducted postmortem examination of the pilot, on October 15, 2000. The cause of death was multiple blunt traumas. The Forensic Toxicology Research Section, Federal Aviation Administration, and Oklahoma City, Oklahoma performed postmortem toxicology of specimens from the pilot. The studies were negative for basis, acidic, and neutral drugs. Atropine a prescription drug routinely used during emergency resuscitation to help restore or control heart function was detected in the blood and liver. Ephedrine an asthma medication was detected in the blood. Phenylpropanolamine a decongestant and weight loss product was detected in the blood.
Dr. Linda Rush, Associate Medical Examiner, District 17, Fort Lauderdale, Florida conducted postmortem examination of the passenger, on October 15, 2000. The cause of death was multiple blunt traumas. No toxicology of specimens from the passenger was requested.
TEST AND RESEARCH
Review of the Robinson R22 Pilot's Operating Handbook, Chapter 10 Safety Tips, Safety Notice SN-29, AIRPLANE PILOTS HIGH RISK WHEN FLYING HELICOPTERS states, "There have been a high number of fatal accidents involving experienced pilots who have many hours in airplanes but with only limited experience flying helicopters.
The ingrained reactions of an experienced airplane pilot can be deadly when flying a helicopter. The airplane pilot may fly the helicopter well when doing normal maneuvers under ordinary conditions when there is time to think about the proper control response. But when required to react suddenly under unexpected circumstances, he may revert to his airplane actions and commit a fatal error. Under those conditions, his hands and feet move purely by reaction without conscious thought. Those reactions may well be based on his greater experience, ie. the reactions developed flying airplanes....
To stay alive in the helicopter, the experienced airplane pilot must devote considerable time and effort to developing safe helicopter reactions. The helicopter reactions must be stronger and take precedence over the pilot's airplane reactions because everything happens faster in a helicopter. The pilot does not have time to realize he made the wrong move, think about it, and correct it. It's too late; the rotor has already stalled or a blade has struck the airframe and there is no chance of recovery. To develop safe helicopter reactions, the airplane pilot must practice each procedure over and over again with a competent instructor until his hands and feet will always make the right move without requiring conscious thought."
Review of ROBINSON MODEL R22 SECTION 3 EMERGENCY PROCEDURES Page 3-2, POWER FAILURE BETWEEN 8 FEET AND 500 FEET AGL states:
"1. Takeoff operation should be conducted per the Height-Velocity Diagram in Section 5.
2. If power failure occurs, lower collective immediately to maintain rotor RPM.
3. Adjust collective to keep RPM in the green arc or apply full down collective if lightweight prevents attaining above 97%.
4. Maintain airspeed until ground is approached, then begin cyclic flare to reduce rate of descent and forward speed.
5. At about 8 feet AGL, apply forward cyclic to level ship and raise collective just before touchdown to cushion landing. Touch down with skids level and nose straight ahead."
The wreckage of N4004J was released to Mr. Tony Hicks, Director of Operations, Volar Helicopter Inc., on October 16, 2000. The components retained for further analysis by the NTSB Materials Laboratory were released to Mr. Craig T. Walker, President, Marco Flite Services Inc on May 14, 2001.
A staff support specialist for the Miami Terminal Radar Approach Control stated, the base of the radar coverage in the vicinity of the crash site was between 500 to 600 feet. The area was interrogated from 1600Z to 1615Z. Two aircraft entered and exited this area squawking transponder code 1200. No other aircraft were observed on radar working in this area. (See FAA inspector statement, an attachment to this report.)