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On May 28, 1999, about 0648 hours Pacific daylight time, a Messerschmitt-Bolkow-Blohm (MBB) BO 105LS A-3 helicopter, N811CE, crashed into the Pacific Ocean approximately 6 miles west-southwest of the Huntington Beach Pier, Huntington Beach, California. The aircraft, operated by Southern California Edison, Ontario, California, was destroyed. The commercial pilot was not recovered and is presumed to have sustained fatal injuries; the two passengers were recovered with the wreckage and suffered fatal injuries in the accident. The flight was conducted under the provisions of 14 CFR Part 91 as a corporate flight to transport two Edison employees to Catalina Island. The flight originated at an Edison Service Center helipad near Irvine, California, about 0625 as a non-stop flight to the Pebbly Beach helipad on Catalina Island, which is about 25 miles offshore. Visual meteorological conditions prevailed at John Wayne airport, near Irvine, at the time of the accident. No flight plan was filed with the Federal Aviation Administration (FAA); however, company flight following was being provided to the pilot.
FAA facility records disclosed that no ATC services or weather briefings were provided to the pilot on the morning of the accident. In both an oral interview and a written statement, the Chief Pilot for the company stated that prior to departure the pilot telephoned the company stations at Rosemead, San Onofre, Irvine, and Westminster for en route weather conditions. The company employee stationed at Catalina Island who was to pickup the passengers said the pilot did not contact him for the weather conditions.
According to weather reports and witness observations from pilots flying along the same general route, the weather in the channel between the Southern California coast and Catalina Island included cloud bases between 500 and 700 feet, and visibility values ranging from 1 to 3 miles. Isolated areas of lower cloud bases and visibility values were observed by the witnesses.
According to Southern California Edison, the pilot departed the Ontario airport at 0521, and flew to the company's general offices in Rosemead and picked up one of the passengers. The pilot had intended to fly to a company helipad in San Onofre to pickup the other passenger; however, the weather along the coast in the San Onofre area included fog and limited visibility. The pilot then asked the second passenger to drive to the Irvine helipad. The flight then left Rosemead and arrived at the Irvine helipad at 0610. After picking up the passenger, the flight departed for Santa Catalina Island. At 0640, the pilot talked to another company pilot on the radio and stated he would be "feet wet in 2 minutes." No further communications were recorded from the helicopter. According to the company chief pilot, the pilot had stated the night before that he intended to cross the coast at the Huntington Beach Pier and proceed direct across the channel to the island.
When the helicopter failed to arrive at Catalina, the company reported it missing and search efforts were initiated by the U.S. Coast Guard. Debris identified from the helicopter was found floating about 3.5 miles off the pier at 1203, along with a fuel slick. The Coast Guard reported the location of the debris field as north latitude 33 degrees 37.2 minutes by west longitude 117 degrees 58.2 minutes.
Recorded radar data covering the area of the accident was examined for the time frame, and a 1200 secondary beacon code target was observed that matched the anticipated flight track of the helicopter from the Irvine helipad to the point of disappearance. The data was supplied by the FAA in the form of a National Track Analysis Program (NTAP) printout from the Los Angeles Air Route Traffic Control Center (ARTCC) long range search radar at San Pedro, California, and a ARTS-IIIA CDR Editor listing from the Southern California Terminal Radar Approach Control, Los Alamitos, California, terminal radar antenna site. The data was then processed in the computer program Radar Viewpoint, by Airways Technology, Inc.
Review of the processed and raw data disclosed that the target was first identified at 0629:16 as it tracked northwest bound over the Interstate 5 freeway from the Irvine area climbing from a Mode C reported altitude of 500 feet to 900 feet. Shortly after reaching 900 feet, it then descended again to 600 feet. At 0636:50, the target turned southbound and tracked over the Santa Ana River until crossing the shoreline near the Huntington Beach pier at 0643:24. The Mode C reported altitude descended to 400 feet as the target crossed the shoreline. Beginning at 0646:24, the Mode C reported altitude climbed to 600 feet over 19 seconds, then turned right at 0646:47 to a northerly heading while the Mode C altitude exhibited a climb from 600 to 900 feet; the computer program calculated the radius of turn at 720 feet at an average ground speed of 69 knots. The Mode C reports then descended to 700 feet as the track began a wide left turn, which reversed to a tight right turn. The last four beacon returns showed a right turn with a 120-foot radius with a constant Mode C report of 700 feet; the average computed ground speed during this turn was 67 knots. The last radar return was a secondary beacon without Mode C at 0648:12, about 6 miles west south west of the Huntington Beach pier at west longitude 118 degrees 05 minutes 58 seconds by north latitude 33 degrees 36 minutes 44 seconds.
During the investigation, certified copies of the pilot's records were obtained from the FAA's Aeromedical Certification Division and the Airman Certification Branch. In addition, flight department records obtained from Southern California Edison were reviewed to document the pilot's training history, flight time, and, the recent duty/rest period activity within the week prior to the accident.
The pilot's FAA Airman Record file established that he held a commercial pilot certificate for rotorcraft helicopters with private privileges in single engine land airplanes. The most recent issuance of the pilot certificate was dated May 1, 1982, and was due to successfully passing the required tests for the commercial certificate in rotorcraft. The pilot did not hold an instrument rating for either rotorcraft or airplanes. In addition, he held an Airframe and Powerplant mechanics certificate and an Inspection Authorization.
Review of his Aeromedical Certification Division records disclosed that he held a second-class medical certificate issued November 30, 1998. The medical certificate contained the limitations that correcting lenses be possessed for near and intermediate vision. The pilot's medical record file also revealed that on December 5, 1990, he was issued a Statement of Demonstrated Ability waiver (SODA), number 30DD9025, for deficient right eye vision beyond the standards prescribed in 14 CFR 67.203. According to the most recent medical certificate examination form (FAA Form 8500-8), the pilot's corrected right eye visual acuity was 20/200 for distant and intermediate ranges, and, 20/100 for near ranges. The corrected visual acuity in the left eye was 20/20 in all ranges. Further examination of the records disclosed that the pilot was diagnosed in October 1989 with ischemic optic neuropathy of the right eye. The issuance of the SODA by the FAA Aeromedical Certification Division was based on medical examinations and the pilot's successful completion of a practical demonstrated ability test.
According to company records, the pilot had accrued a total flight time of 4,341 hours, of which 4,112 were in rotorcraft, and 49 in the BO-105. In the preceding 90 and 30 days, the pilot had flown 137 and 35 hours respectively, with 11 hours flown in the BO-105. The pilot's total night experience was determined to be 22 hours, with 2 hours flown in the past 90 days. No instrument flight experience was found in the records supplied. His most recent biennial flight review was accomplished on September 22, 1998, in an American Eurocopter AS-350 as part of a factory conducted recurrent ground and flight training session. The records also disclosed that the pilot completed the Eurocopter factory conducted ground and flight training program in the BO-105 on November 18, 1998. No deficiencies were noted on the training forms for the factory courses. The company chief pilot reported that a desktop simulator was available in the flight department offices for the use of the pilots; no record of the pilot using the device was located.
Company records were examined to determine the pilot's duty and rest periods for the 7 days preceding the accident. The pilot was off duty over May 22 and 23, and reported for work at 0715 on May 24 for a 10-hour shift. He reported for a 12-hour shift at 0730 on May 25, and had a 10-hour shift beginning at 0700 on May 26. On May 27, he worked a 9-hour shift from 0900 to 1800. The pilot reported for duty at 0400 on the day of the accident.
Interviews were conducted with company pilots who interacted with the pilot in the days preceding the accident. All reported that the pilot appeared healthy and upbeat. The company pilot who spoke with the pilot by radio on the day of the accident stated that he sounded normal in all respects. There was no indication from family members or other company pilots that the pilot was taking any prescription or over the counter medications.
The maintenance records for the helicopter were obtained from the operator and reviewed. In addition, information and records were obtained from American Eurocopter concerning original equipment installed during production. The aircraft and registration files maintained at the FAA Registry in Oklahoma City, Oklahoma, were also reviewed.
The BO 105LS A-3 helicopter, serial number 2036, was manufactured by MBB Helicopters Canada Limited on May 27, 1991. Between the production date and August 12, 1994, it was a MBB factory demonstrator under Canadian registration C-FSBK, and accumulated a total time of 49 hours. It was then de-registered and shipped to Japan, where it again was a demonstrator model for the MBB distributor in that country. The helicopter was sold to the present owner and arrived in the United States on October 12, 1998, with an accumulated total time of 83 hours since new. A U.S. Standard airworthiness certificate in the normal category was issued and the helicopter registered as N811CE.
At the time of the accident, the airframe and both engines had accumulated a total of 233 hours, 939 landing cycles, and 510 engine cycles. The most recent 100-hour inspection was completed on March 26, 1999, 39 hours prior to the accident. A combined manufacturers 25- and 50-hour inspections were accomplished on May 11, 1999. Safety Board investigators reviewed the operator's "Daily Helicopter Flight Report" forms. The last one available was dated May 20, 1999. Two unresolved discrepancies were noted. The first, dated March 12, 1999, noted that the "DAVTRON gage stuck on voltage." The second, dated March 29, 1999, listed the 900 MHz radio as intermittent.
On the morning of the accident, the operator's dispatch records showed that the helicopter departed the company's Ontario flight operations base with 110 gallons of Jet A onboard.
Manufacturer's records disclosed that the helicopter was equipped with the instruments required by 14 CFR 91.207 for flight in instrument conditions; however, it was not certified for IFR flight.
The dual hydraulic power pack installed on the helicopter was a p/n 105-450121, serial number 17-004. Airworthiness Directive 96-08-04 was not applicable to either the 105LS A-3 model helicopter or the hydraulic pack by part number.
A review of radio communications from FAA sources, the frequency used by the company, and other CTAF frequencies used by helicopters in the harbor area, revealed that no distress calls were monitored during the time frame of the accident flight.
A staff Safety Board meteorologist prepared a factual report of the meteorological conditions in the vicinity of the helicopter's flight route. That report is appended to this file.
The 0653 aviation surface weather observation at the Santa Ana airport (close to the departure helipad) was reporting a cloud ceiling of 1,100 feet overcast, with a 5-mile surface visibility. Other coastal airports within 10 miles of the accident site were reporting overcast ceilings ranging from 700 to 1,100 feet, with visibilities between 4 and 5 miles in mist. Pilot reports for the area showed the tops of the stratus clouds ranging from 1,900 to 2,200 feet msl.
A witness report from the destination helipad at Avalon on Catalina Island included an estimated 2-mile visibility in fog with cloud bases of at least 500 feet. Other helicopter pilots who fly passengers on scheduled flights for a commuter operator between the Los Angeles Harbor and Catalina stated that about 0700, the weather in the channel was 700 overcast with 3 miles, but that an area of lower fog and reduced visibility existed toward the southeast in the area offshore of the Huntington Beach Pier.
Two additional witnesses were located on offshore oil platform Ellen, which is about 2.5 miles southwest of the final location of the helicopter. According to the witnesses, they were on deck during the time frame of the accident. The witnesses estimated the cloud bases on the basis of the known height of the oil rig top. Visibility estimates were made on the basis of distances to known objects and geographical features around the platform. They reported that the sky was overcast with bases estimated at 300 to 350 feet msl. The visibility varied according to the cardinal magnetic direction away from the platform and was noted by the witnesses to be: 2 to 2.5 miles to the north; 3 miles to the west; 1 to 2 miles to the south; and 1.5 to 2 miles to the east.
In pertinent part, the Safety Board meteorologist noted that the area was under the influence of a low pressure centered near Needles, California, with a weak inverted trough extending northwestward through central California. High humidity and light winds were found in station plots along the California coastline. Digital brightness data obtained from the GOES-10 satellite for the time frame of the accident showed that the helicopter's radar derived ground track moved from an area with a 61 to 87 brightness count to a 88 to 255 brightness count area when radar contact was lost.
The aviation area forecast for the Pacific Coast was issued by the National Weather Service at 0345 on the morning of the accident. AIRMET Sierra was contained in the area forecast and predicted IFR conditions for the Southern California Coastal waters with broken to overcast ceilings of 1,000 feet; visibilities from 3 to 5 miles in mist and tops of the stratus clouds at 2,000 to 3,000 feet.
WRECKAGE AND IMPACT
The wreckage was located by a side scan sonar search on June 2, 1999, on the ocean floor at a depth of 50 meters. The coordinates of the wreckage were north latitude 33 degrees 36 minutes 46 seconds by west longitude 118 degrees 05 minutes 55 seconds.
Prior to any recovery attempt, the wreckage was videotaped by the salvage firm hired for the retrieval. Safety Board investigators reviewed the video footage and examined the side scan sonar image. All helicopter extremities were present in the immediate vicinity of the main wreckage mass. The tail boom was attached to the fuselage, and the tail rotor blades remained attached to the hub. While damaged, distorted, and fractured at the transition point from the grip arm to the full cord width, all four main rotor blades were attached to the hub.
On June 3, 1999, divers secured a cable to the mast and the helicopter was lifted to the surface. During the process of lifting the wreckage clear of the water, one main rotor blade (subsequently identified as the red blade) separated at the fracture point and sank to the ocean floor. Beyond the one main rotor blade, the only major piece of wreckage not recovered was the pilot's cockpit instrument console containing the flight instruments.
TESTS AND RESEARCH
Following recovery, the wreckage was taken to the facilities of Aircraft Recovery Services, Compton, California. The wreckage was schematically reconstructed and examined by Safety Board investigators over the period from June 7 to June 9.
The fuselage was extensively crushed, both vertically and horizontally. The vertical crush lines were oriented on about a 70-degree angle and extended up to the engine deck, with the structure between the floor subassembly and the engine deck totally disrupted. The horizontal crush lines were delineated in accordion wrinkles extending the full length of the cabin floor structure. The cockpit and cabin windows were missing. The pilot's seat was not recovered.
The tail boom remained attached to the fuselage and was bent downward in a smooth curving arc. The bottoms of the vertical stabilizers were vertically crushed upward. The flexible tail rotor drive shaft couplings at the main transmission and tail rotor transmission ends were torn and separated. The tail rotor drive shaft was axially shifted aft, with all of the hanger bearings bent and shifted aft. The 45- and 90-degree tail rotor gearboxes rotated smoothly, with no deposits found on the chip detector plugs. The red tail rotor blade exhibited trailing edge spanwise compression wrinkles and a 45-degree chordwise fracture near the grip end. The white tail rotor blade exhibited trailing edge spanwise compression wrinkles, a chordwise fracture near the grip end, and leading edge damage near the tip.
All four main rotor blades were fractured at a point just outboard of the transition from the grip arm to the blade's full chord width. All four blades exhibited trailing edge compression damage in the spanwise direction.
The main rotor hub remained attached to the drive shaft and was undamaged. The pitch link pitch horns for each blade were intact. The swash plate assembly was intact and undamaged. The collective and longitudinal control tubes between the swash plate and the dual hydraulic unit were intact. The lateral control tube was bent and fractured near the swash plate end; the fracture had a granular appearance and was on an angled plane to the long axis of the tube. The control tubes and bell cranks between the dual hydraulic unit and the cockpit controls were traced through the airframe. The tubes were fractured at multiple points along their run; the fractures were granular, angular, and accompanied by bend deformation of the respective tubes. The antitorque control tubes and bell cranks were intact through the tail boom to the fuselage; through the forward fuselage areas, multiple fractures were found, which were granular, angular, and accompanied by bend deformation of the respective tube.
The dual hydraulic unit remained attached to the engine deck structure. The support bracket for the wires leading to the override micro switches on top of the unit was fractured at the aft end. Canon plugs 10DB3 and 10DB4 for the collective mechanical override micro switch were sheared at their respective attachments; however, the plug pins were intact. Micro switch 10DB1 was damaged and a corresponding mark was observed on the adjacent micro switch guide assembly. The system No. 1 hydraulic reservoir attachment brackets were distorted but intact. The system No. 2 hydraulic pump drive shaft was bent and the pump pressure line was sheared at the fitting. The selector valve was found on system No. 1; the selector valve linkage was undamaged and smoothly operated from system No. 1 to system No. 2.
The dual hydraulic unit was removed from the helicopter and sent to the Eurocopter factory in Fort Erie, Ontario, Canada, for detailed examination and functional testing. The examination and tests were conducted on August 18, 1999, under the supervision of an aviation investigator from the Transport Safety Board of Canada (TSB). A complete report of the examination is appended to this file.
According to the TSB investigator's report, the fluid from system Nos. 1 and 2 was free of contamination, and the reservoirs were at the factory recommended servicing levels. In order to facilitate the functional testing, the following work was accomplished on the units: 1) the reservoirs and lines were flushed and new hydraulic fluid added to capacity; 2) the bent drive shaft for the system No. 2 pump was replaced with a new component; 3) the broken pump pressure line fitting was replaced; and 4) the sheared micro switch canon plugs were re-mated to their respective switches. The unit was then functionally tested in accordance with the factory acceptance protocol.
The system No. 1 hydraulic pump operated within limits at 103 bars.
The system No. 2 hydraulic pump exceeded normal operating pressures at 145 bars and this was traced to the damaged micro switch 10DB1 (see above), which failed to operate in the down position.
The change over system operated smoothly and within limits. Pressure switches operated properly. System simulated load checks showed normal operation. Symmetry checks of the two systems showed that the longitudinal, lateral, and collective axis were all within limits. System leak checks were well within the allowable limits for serviceable units.
On September 6, 2000, Safety Board investigators traveled to the Eurocopter facility in Grand Prairie, Texas, to explore the effect of the fractured canon plugs noted above on the operation of the dual hydraulic unit. A BO 105LS A-3 helicopter with an identical hydraulic unit to the accident helicopter was used for the tests. Following the normal checklist, the fractured canon plugs were observed to be clearly visible on the preflight inspection. With electrical power turned on and the plugs separated, the "Hydraulic System 1 to 2" warning light illuminated on the enunciator panel. The plugs were reconnected and the hydraulic system was then powered on the ground by the use of a hydraulic mule. The plugs were then separated to simulate an in-flight failure of the plugs. No effect was observed on system operation. The No. 1 system pressure was manually reduced to simulate a failure of that system; the hydraulic unit switched to the No. 2 system and the aforementioned cockpit warning light illuminated.
The cockpit engine fire extinguisher bottle switches were in the off and guarded positions. The discharge disks on the bottles were intact. Bottle No. 1 had no pressure. Bottle No. 2 had a pressure of 30 psi.
The electrically actuated fuel shutoff valve switches in the cockpit were broken, with the overlying guards destroyed. The left shutoff valve was found in the full open position. The right shutoff valve was found with the ball cock opening about 90 percent closed. According to the valve manufacturer, under nominal voltage and temperature conditions, the valves have a cycle time from full open to full closed of less than 1 second.
The engine to transmission drive shafts for both engines were torsionally twisted (greater than 360 degrees) to failure in the clockwise direction as viewed from the rear. The driving direction of the shafts are clockwise. The flexible bendix coupling for the No. 2 engine drive shaft was torn about 80 percent around its circumference.
The Rolls Royce Allison 250-C28 engines were examined externally, with no damage evident. Serial number 280100 was installed in the No. 1 position and 280101 was installed in the No. 2 position. The governor indices for both engines was pointing to 90 percent. The fuel control indices for both engines was at 10 percent.
Both engines were removed from the airframe and transported to a factory authorized overhaul facility for teardown examination. No internal failures were noted on either engine. All bearings were oiled and free to move. Rotational scoring was found on the compressor impeller, the compressor shroud, and the 3rd and 4th stage wheel blade paths. A complete report is appended to this file.
The fuel cells were ruptured. Disassembly of the fuel boost pumps revealed that the shafts rotated, the impellors were intact, and the housing walls were not scored.
MEDICAL AND PATHOLOGICAL INFORMATION
The pilot was not recovered and no suitable specimens remained from the passengers to conduct toxicological tests for carbon monoxide.
With the exception of the dual hydraulic power pack and both left and right fuel shutoff valves, the remaining wreckage was released to Southern California Edison on June 9, 1999. At the conclusion of the hydraulic assembly examination on August 18, 1999, that component was shipped directly from the factory to the Edison facilities. The fuel valves were returned to the company on October 26, 2000.