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On October 4, 1998, at 1555 hours Pacific daylight time, a Bell 222 helicopter, N213ML, entered a descent and collided with the ground during takeoff at the Las Vegas Motor Speedway, Las Vegas, Nevada. The helicopter was destroyed. The commercial pilot was not injured, and his pilot rated passenger sustained minor injuries. The helicopter was owned and operated by Action Helicopter Service Center, under 14 CFR Part 91 of the Federal Aviation Regulations. Visual meteorological conditions prevailed at the time of the accident, and no flight plan was filed. The personal flight originated from the owner's home at 1300 to attend the race being run at the Motor Speedway, and the helicopter was departing for the return to the residence.
The pilot reported to the Safety Board that he made a normal liftoff to a hover and checked the gauges, which appeared to be normal. He said he then began to takeoff and checked the MGT gauge (measured gas temperature for each engine's operating limits.) He stated that the left No. 1 gauge was "peaked." He pulled in collective to clear the parked cars in the parking lot to reach an open area. After clearing the cars, the helicopter bounced landed, bounced, rolled over, and came to rest on its left side.
There were numerous eyewitnesses to the crash. Witnesses indicated that the helicopter lifted off to the north, "barely missing the palm trees." Some witnesses indicated that the helicopter's attitude pitched forward and side to side, other's stated it was hovering erratically, appeared to lose power, and spun 180 degrees and tilted forward before the main rotor contacted the ground.
The Chief of Security at the track stated that the aircraft wasn't gaining altitude and that he didn't think the aircraft got above the top of the three-story administration building. He said the helicopter was heading northeast towards the handicapped parking lot. He stated he observed the back of the helicopter turn left and then right and its nose started going down like he was going to land in the handicapped lot.
Another eyewitness who was working at the track performing traffic and parking functions said the aircraft ascended about 20 feet before it was blown towards the street away from the administration building. He said the aircraft was moving sideways and it looked like the aircraft was going to land in the handicapped parking lot. He stated that "it appeared that the wind got him as he cleared the building." He said that the winds were from the south and he had observed the signs in the parking lot being blown over to the north.
Another eyewitness said that the aircraft got to the height of the administration building and then appeared to come forward before it began to move around like it was fighting the wind. By then she said the aircraft had drifted toward the road and she watched it "wiggle side to side."
The pilot was sitting in the right seat at the time of the accident. He reported that he had accrued about 5,875 total flight hours, of which 351 hours were in the Bell 222. He reported that he had flown 43 hours in the preceding 90 days, and 26 hours in the preceding 30 days. According to the aircraft logbook, N213ML had flown a total of 11.0 hours during calendar year 1998. The company also owns a sister ship, N212ML, which reportedly flew about the same amount of time. The mechanic on the aircraft reported that the two ships had a combined flight time of less than 25 hours a year. The pilot also has a Bell 206L-4, which he reportedly flew approximately 125 hours a year. The pilot-rated passenger in the left seat said that she was not current in the aircraft, and had assumed no flying duties other than helping the pilot perform the checklist items during normal operations.
Training records for the pilot, obtained from Bell Helicopter Textron in Ft. Worth, Texas, were reviewed. The pilot had completed a biennial flight review in accordance with 14 CFR Part 61.56 on June 6, 1998, in a Bell 206L-4 with the Director of the Bell Helicopter Customer Training Center. The flight time noted on the training sheet was 1.2 hours, with the remark "excellent pilot in all areas of flight." The pilot had flown with the same Bell pilot on November 11, 1997, for 2.0 hours. This flight was also conducted in the Bell 206L-4.
Bell flight records were reviewed for the preceding 2 years from the date of the accident. No record was found that the pilot had flown with any Bell helicopter pilot in any Bell 222 helicopter during this timeframe. Additionally, Flight Safety International, Ft. Worth, was contacted and they have no record of the pilot ever attending their Bell 222 ground school or simulator course.
The Director of the Bell Helicopter Customer Training Center was interviewed and he mentioned that he did fly with the pilot in the Bell 222, but that the flights were "not training flights." According to their training records, the pilot flew his initial aircraft checkout in a Bell 222 on October 29 and 30, 1993. During this checkout totaling 6.4 hours of flight time, the instructor pilot noted "all maneuvers were above average for time in all aircraft."
Safety Board investigators were unable to find any documented evidence that the pilot was proficient in any emergency maneuvers in the Bell 222. After the accident, the Director of the Bell Helicopter Customer Training Center said at the pilot's request, he conducted a short flight manual review followed by a 1.4-hour flight review in the Bell 222 aircraft, and provided a letter to the Federal Aviation Administration (FAA) regarding the pilot's skills. According to the letter, all maneuvers including preflight, starting procedures, normal and maximum performance takeoffs and one engine-out procedure were performed satisfactorily.
Documents and records reviewed by the Safety Board disclosed that Action Helicopters acquired the aircraft on September 6, 1995. At that time, the aircraft had accrued a total time since new of 2,519 hours, with 575 hours since overhaul. The aircraft was powered by two Lycoming LTS101-650C3 engines, which were rated at 600 horsepower each. According to the records, engine number one (left) had a total time of 2,584.8 hours and entries documented compliance with all applicable Airworthiness Directives (AD's). The number 2 engine (right) had a total of 2,580 hours and entries showed compliance with all applicable AD's. There was no evidence of compliance with all recommended engine Service Bulletins (SB), including Service Bulletin LTS101-72-00-189, which would have allowed verification of the availability of the specified One Engine Inoperative (OEI) power. A complete summary of the AD and SB history is appended to this report, and is included in the Test and Teardown Report of Two Model LTS101-650C-3 Turboshaft Engines by Allied Signal Aerospace.
Weather observations were taken at the adjacent Nellis Air Force Base (KLSV), at 1555. The METAR reported winds 150 degrees at 07 knots; overcast clouds at 11,000 feet; broken clouds at 25,000 feet; 40 statute miles visibility; temperature 75 degrees Fahrenheit; dew point 41 degrees Fahrenheit; and an altimeter setting of 29.92 inHg.
The owner of the Las Vegas Motor Speedway stated that the winds were usually gusty and variable depending on where you were located on the property. On October 6, he provided the Safety Board with a diagram of the winds as they were viewed on the track's various video cameras, which were located throughout the property. A diagram of the wind direction is appended to this report. The winds were observed to be generally coming from the south. The wind around the administration building resulted in quartering tailwinds for the departing helicopter.
WRECKAGE AND IMPACT INFORMATION
When the Safety Board investigator arrived at the racetrack on October 5, the helicopter wreckage had been removed from the parking lot and secured in a racetrack hanger facility for examination.
The site was documented by FAA inspectors from the Las Vegas, Nevada Flight Standards District Office, and the Safety Board investigator reviewed the site with the inspectors after arrival.
The accident site was a large flat asphalt parking lot on the grounds of the motor speedway surrounded by widely scattered palm trees, light poles, and speedway outbuildings, with heights ranging from 40 to 60 feet. The parking lot was partially filled with parked automobiles.
The helicopter had been lying on its left side with the nose oriented on a magnetic bearing about 030 degrees. A distance of 208 feet was measured from the witness-identified takeoff point to the wreckage location on a magnetic bearing of 030 degrees. The takeoff point was on the north west side of a three-story administration building with a measured height of 46 feet. The light poles near the front of the building were measured at 59 feet tall. An asphalt road bordered by 10-foot-tall palm trees separated the takeoff point and the wreckage point of rest. The only ground scars observed in the asphalt were within approximately 1.5 rotor disk diameters of the helicopter point of rest. Debris later identified as main rotor blade fragments were found scattered throughout the parking lot at distances about 100 feet from the wreckage.
The landing gear was found extended and undamaged. The tail boom was buckled just forward of the horizontal stabilizer. Both tail rotor blades remained attached to the hub grips; both blades exhibited tip-end leading edge damage, while one was bent toward the face side just outboard of the grip doubler. Stub sections about 15 feet in length of both main rotor blades remained attached to the hub grips. The leading edge tip ends of these sections were curled up and rearward. Both of the remaining main rotor blade stub sections exhibited leading edge damage and compression buckling in a chordwise direction.
Copies of photographs were provided by the Las Vegas Flight Standards District Office (FSDO) personnel, who had arrived on scene prior to the wreckage being moved. The photographs were taken the morning of October 5, 1998. In one photograph, which is appended to this report, the FAA inspector had documented the pilot's side twist grip-style throttle positions for the No's. 1 and 2 engines, respectively. The photograph clearly shows that the No. 1 (left) throttle is rolled to the full on position, while the No. 2 (right) throttle is much less than the full on position.
Safety Board investigators interviewed racetrack personnel and others who initially responded to the downed aircraft. All the personnel interviewed stated that they were focused on disconnecting the battery to the aircraft and turning off the battery power. The mechanic who regularly worked on this aircraft showed the first responders where the battery was located. Another eyewitness who was at the racetrack ran up to the helicopter to see if he could provide aid. He stated he had worked on military aircraft for 20-21 years. He said when he reached the aircraft he could "hear electrical power." He said he leaned inside the aircraft to switch the power off by pulling the emergency bus circuit breaker. He stated he did not notice the throttle position.
None of the individuals who first responded to the helicopter reported moving the twist-grip throttles.
TESTS AND RESEARCH
Engine Tests and Examination
The engines, serial number LE-41234 and LE-41232, were installed in the left and right engine positions, respectively.
The left engine was examined on site prior to the engine being removed from the helicopter. According to the Allied Signal engine representative who examined the engines under supervision of the Safety Board investigator, the engine and its external components were intact and appeared to be undamaged prior to removal from the helicopter. The power turbine (NP) and gas generator (NG) spools rotated freely. A leak check of the fuel control pneumatic system was conducted in accordance with the engine maintenance manual prior to removal. No leakage was observed. The fuel control throttle angle was observed to be approximately 0.125 inches from the fuel control maximum throttle stop when the cockpit throttle was advanced to the maximum position. The upper inlet scroll and T1 sensor appeared to be undamaged, and were removed from the engine to facilitate removal of the engine from the helicopter. There was no fire damage.
The left engine was installed in test cell C-903 at the AE Test Laboratory in Phoenix, Arizona, on October 15, 1998. The required prestart checks were completed. The engine operated satisfactorily during all tests and produced rated power. Additionally, the Allied Signal representative stated that testing of the left engine did not identify any operating condition that would have resulted in the excessive MBT reported by the pilot.
The left engine was partially disassembled and the power section was inspected for evidence of excessive MGT at the AE Development Assembly facility in Phoenix on January 14, 1999. The Allied Signal representative stated that no physical evidence of the excessive TOT reported by the pilot was observed on the left engine.
The right engine was also examined prior to being removed from the helicopter. The engine and external components were intact and appeared to be undamaged prior to removal from the aircraft. There was no fire damage. The NP and NG spools rotated freely. A leak check of the fuel control pneumatic system was conducted and bubbles indicating leakage were observed at the pneumatic connections to the T1 sensor and the fuel control orifice during the leak check. These connections were not disturbed and subsequent test cell runs of the engine were conducted with this condition present.
The right engine was installed in test cell C-903 at the AE Test Laboratory in Phoenix on October 14, 1998. According to the Allied Signal representative, the right engine operated satisfactorily during all tests and produced rated power.
No conditions were identified on either engine, which would have interfered with satisfactory operation. A complete copy of the Test and Teardown report of the engines, report No. 21-10548, is on file at Allied Signal Aerospace, Phoenix.
The gearbox chip detector was undamaged and uncontaminated when examined. There was residual oil on the gearbox chip detector. The rear bearing chip detector was undamaged and uncontaminated. There was residual oil on the rear bearing chip detector.
Fuel Sample Tests
Fuel samples were obtained from both left and right fuel tanks. The AE Chemical Laboratory in Phoenix analyzed the fuel samples on October 21, 1999, at the request of the Safety Board. No anomalies were observed.
Engine Component Examinations
The left engine fuel manifold (P/N 4-301-236-02, S/N 134) was intact. The fuel manifold was installed on a test stand at the AE test laboratory in Phoenix on November 16, 1998, and was flushed to remove any contaminants. The collected fluid samples were drained onto filter paper, and the filter paper examined for contaminants. No significant amount of contaminants was observed.
The fuel manifold was functionally tested at the Allied Signal facility, and the following variations from the test specification requirements were observed by Allied Signal and FAA personnel:
1. The fuel manifold inlet pressure was below the minimum test specification values at all five fuel flow valves.
2. Uneven distribution of fuel flow was observed on three fuel nozzles at the 50 pph fuel flow point. Uneven distribution of fuel flow was observed on one fuel nozzle at the 440-pph-fuel flow point.
3. Poor fuel atomization was observed on all eight fuel nozzles at the 50-pph-fuel flow point. Streaks were observed on all eight fuel nozzles at the 400-pph fuel flow point.
The test results confirmed that the low fuel manifold inlet pressure resulted in poor fuel atomization at low fuel flows. The fuel manifold was removed from the test stand and returned to the supplier, Delavan, Inc., for further investigation.
The teardown and inspection of the fuel manifold was conducted at Delavan, Inc., in West Des Moines, Iowa, on December 16, 1998 under the supervision of the FAA. The tests revealed that the manifold was pressure tested with no external leakage detected. Additionally, all nozzles were removed to inspect the primary seals, P/N ASE 2-300-235-01, and the secondary seals, P/N ASE 1-300-369. Delavan determined that of the eight primary seals, three seals were dissimilar from the other seals and had visual evidence of distortion on their internal diameter. All of the secondary seals appeared deteriorated, as compared to new, based on a visual inspection, but did not leak. Nozzle No. 8 was inspected for blockage in the primary circuits. Two of the four slots were partially blocked by foreign material. The material was removed and the nozzle was reassembled for test purposes.
According to the Delavan representative, it was noted during the disassembly process that three primary seals were distorted. The dimensional features of the seals were also recorded. The three distorted seals were dimensionally different from the undistorted seals. The distorted seals did not meet the requirement of the design and product specifications of the fuel manifold as specified in drawing ASE P/N 2-300-235-01. According to the written report supplied by Delavan, there are no Delavan drawings or part numbers that correspond to ASE 1-300-366-01 or -03. Delavan representatives noted that the measurements of the distorted seal did appear similar to other ASE seals used on the T53 manifold. Delavan had not serviced the fuel manifold before the accident, and a review of the maintenance records is unclear about the origin of the distorted seals.
The tests confirmed the Allied Signal result of low fuel manifold inlet pressure at the fuel flow rates specified in PS978, the Delavan LTS 101 test requirement. The manifold was then retested with eight new primary seals. With the new seals installed, Delavan found the pressure differential across the manifold met all but one of the test point requirements of PS978. The one test point not met gave a higher than required differential pressure.
Section 5 of the FAA approved Bell helicopter model 222 Flight Manual states that the helicopter is certified under FAR Part 29, Category B provisions. According to these provisions, if an engine failure occurs during takeoff, continued takeoff and climbout capability is not assured. The Category B takeoff profile assures the capability to safely land (on a smooth level surface) from any point in the takeoff profile should an engine failure occur. The takeoff performance data chart in the flight manual utilizes a takeoff profile clear of the height-velocity (H-V) diagram. The chart specifies that to achieve published distance, the takeoff should be initiated from a 4-foot wheel height in a stabilized hover; start nose down pitch rotation and simultaneously increase power smoothly to 90 percent mast torque so that the aircraft accelerates along a flight path within the takeoff corridor as defined by the H-V diagram.
An attempt was made to replicate the accident conditions in a Bell 222 simulator at the Flight Safety International training facility in Hurst, Texas, on April 22, 1999. At that time, the flight safety instructor (FSI) pilot set up the configuration of the simulator using Las Vegas International Airport's field elevation of 2,179 feet msl, a gross weight of 7,200 pounds, and an outside air temperature of 24 degrees Celsius or 75 degrees Fahrenheit. The FSI instructor sat in the right seat, with a Bell Helicopter accident investigator in the left seat. A Safety Board investigator sat in the simulator instructor seat and monitored the scenarios.
The first scenario involved having the FSI pilot bring the aircraft to a 50-foot out-of-ground effect hover, drop the nose of the aircraft, and fly out with both engines operating at 100 percent N1 speed. The aircraft was positioned on a heading of 255 with the winds from 074 degrees at 15 knots. The wind speed and direction included a turbulence and wind profile selected on the computer to simulate variable wind conditions. The No. 1 throttle was split on the ground to an indicated 150-degree difference in the indicated MGT. The takeoff profile at 50 feet agl indicated that the No. 1 engine was operating at 620 degrees MGT, and the No. 2 engine was operating at 900 degrees MGT.
The second scenario involved having the No. 2 throttle split on the ground to indicate an approximate 150-degree difference between the MGT of each engine. Using the triple tachometer as a guide, the No. 2 engine N1 needle was located just off the 100 percent stop. The FSI pilot lifted off to a 50-foot OGE hover. The No. 1 engine indicated 820 degrees MGT, and the No. 2 engine indicated 600 degrees MGT. An audible rotor horn was heard and the rotor rpm dropped to 92 percent. Additionally, the rate-of-climb gage indicated a 1,000 foot-per-minute rate of descent.
The third scenario involved having the No. 1 throttle split on the ground to indicate an approximate 150-degree difference between the MGT of each engine. The No. 1 needle N1 needle was located just off the 100 percent stop. The FSI pilot lifted off to a 50-foot OGE hover. The No. 1 engine indicated 725 degrees MGT, and the No. 2 engine indicated 800 degrees MGT. During the takeoff, the airspeed indicator indicated 31 knots forward airspeed, but the rate-of-climb gauge indicated a 1,700 foot-per-minute rate of descent. Rotor rpm stayed at 100 percent, but the FSI pilot indicated that he had run out of aft right cyclic control authority during this scenario.
Safety Board investigators requested that Bell Helicopter Textron conduct an engineering analysis to calculate the power requirements which would have been needed to successfully complete the 50-foot out-of-ground-effect (OGE) hovering takeoff using a 7,200 pound gross weight, pressure altitude of 2,200 feet, and 24-degree Celsius temperature. The engineers were unable to specify the effects of taking off with a downwind condition, other than to state "it would aggravate the situation." The combined shaft horsepower (SHP) for both normally operating engines was calculated to be 988 SHP, with each engine capable of producing 494 SHP. Bell Helicopter Textron personnel concluded that in a single engine situation, insufficient power would be available to hover, and in a situation with the power output of one engine reduced for any reason, by 150 SHP, it would be impossible to maintain an OGE hover.
The aircraft was released to the insurance company, representing the registered owner, on August 11, 1999.