National Forest Service Helicopter Crashes in East Texas

On March 10, 2005, approximately 1354 central standard time, a Bell 206B-3 helicopter, N85BH, sustained substantial damage when it impacted heavily wooded terrain in the Sabine National Forest near Shelbyville, Texas. The airline transport rated pilot and two United Stated Department of Agriculture (USDA) Forest Service (USFS) crewmembers sustained fatal injuries. Visual meteorological conditions prevailed, and the flight/mission was being monitored and conducted in accordance with USFS aviation policies for public use aircraft in fire management operations. The flight departed at 1347 from a field helicopter pad (H1), located approximately 7 miles southeast of the accident site.

On the morning of the accident, the helicopter was assigned to support a prescribed fire within heavily wooded terrain with 100-120 foot high trees near Shelbyville, TX. The prescribed fire was supported by the application of aerial ignition spheres utilizing a cabin mounted plastic sphere dispenser (PSD) machine. According to USFS operating procedures, PSD missions are typically flown at 50-300 feet above the top of vegetation at airspeeds from 20-40 knots. The helicopter was based at Angelina County Airport, Lufkin, Texas. Approximately 0928, after a mission brief, the helicopter, with pilot and two USFS personnel on board, and the re-fueling truck re-positioned to H1 (coordinates North 31 degrees 42.110 minutes West 93 degrees 52.540 minutes) and were met by support equipment and personnel from the Sabine National Forest to conduct a prescribed fire mission.

Approximately 1234 an initial recon of the burn area began, followed by approximately 11 minutes of aerial ignition work on the same flight. Approximately 1300 the PSD machine experienced a sphere jam, and the helicopter returned to H1 to resolve the problem. The helicopter shut down at H1 while the PSD machine problem was resolved. The helicopter then departed H1 at 1347 to resume the mission. According to USFS records from ground personnel on the burn, at 1352, the mission ignition specialist onboard the aircraft reported by radio that the helicopter was commencing firing operations. At 1354, a radio distress call was heard by 7 personnel on the burn. According to USFS personnel, the voice making the distress call appeared to be that of the ignition specialist, not the pilot. The call was, “Mayday, Mayday, Mayday, we are going down.” No further communications were heard from the helicopter.

At 1417, the helicopter wreckage was found at coordinates North 31 degrees 45.425 minutes West 94 degrees 00.244 minutes. Immediate rescue operations commenced. One USFS employee initially survived the crash and was being transported by ambulance to a local hospital. The passenger died during transport due to extensive injuries sustained in the crash.

PERSONNEL INFORMATION

A review of Federal Aviation Administration (FAA) airman records revealed that the pilot held an airline transport pilot certificate with ratings for airplane single engine land, airplane multiengine land, and rotorcraft-helicopter, and a certified flight instructor certificate with ratings for single and multiengine land, rotorcraft-helicopter, and airplane and rotorcraft instrument. He was also an FAA certificated advanced and instrument ground instructor. His most recent FAA first class limited medical certificate was issued on September 7, 2004, with the limitation, “must have available glasses for near vision” At this time, he reported a total of 3,000 flight hours.

The pilot’s personal logbook was obtained. During a review of the logbook, several discrepancies were noted. Rotorcraft time appears to be consistent from the time rotorcraft time is first logged through the page ending February 20, 2004, at which time the pilots logbook shows a total rotorcraft time of 286.1 hours. On the page ending March 2, 2004, in the amount forwarded column, rotorcraft time increases to 1286.1. The page total is 27.3, which was added to time from the previous page for a total time of 1323.4 hours. This was an unaccounted 1,000-hour increase in rotorcraft time from the previous page. The extra 1,000 hours were added to and subtracted from cumulative flight time throughout the remainder of the logbook entries. The last entry in the logbook reflects the pilot’s total time in all aircraft to be 2,187.8 hours, 384.4 of that in rotorcraft, and 115.9 hours in make and model.

AIRCRAFT INFORMATION

The 1980-model Bell 206B-3 helicopter was a single pilot, five place, single engine, light helicopter with a two-blade semi rigid main rotor, and a tail rotor that provided directional control. The helicopter was owned by Brainerd Helicopter Service, Inc, and operated by the United States Forest Service. A review of the aircraft logbooks revealed that the last annual inspection was performed on April 26, 2004, at a tachometer time of 309.27 hours. The tachometer read 2,837.4 at the scene of the accident. The aircraft total time was determined to be 4,565 flight hours at the time of the accident.

The helicopter was equipped with a Rolls Royce 250-C20B engine, serial number (S/N) CAE-840516. Historical records for the engine began on April 12, 1991, with 463.1 hours total time. The engine was converted to a Model 250-C20B (same serial number) on March 15, 1993, with a total time of 907.3 hours. Records continued with periodic inspections (100, 300, 600 and Annual) through February 22, 1999. The February 22, 1999, entry indicated that engine, S/N CAE-270208, was removed due to metal on the upper and lower chip detector plugs. There were no previous indications that engine S/N CAE-270208 had been installed. The total time prior to this entry was reflected as 2,653.2 hours on January 1, 1999, and according to the records, should have been for engine S/N CAE-840516. This total time was carried forward to the questionable entry on February 22, 1999, and subsequently crossed out and replaced with an engine total time of 2,296.2. The airframe log showed an entry for engine, S/N CAE 840516 being installed on February 22, 1999, but did not show engine S/N CAE 270208. The entry appeared to be inserted between two previously entered items on a single line and overlapped the A&P mechanics number from the previous entry. History of engine S/N CAE 840516 could not be determined from the records presented.

Historical records shows the last entry for engine S/N CAE 840516 was on 16 December 2004. Total time was recorded as 4,142.7 hours. The airframe total time was recorded as 4,499.5 hours total time. The daily log indicated a total time of 4,175.6 hours for the engine and 4,563.1 hours for the airframe. The serial number on the component card (FF357738) differed from the serial number on the Canadian Authorized release Certificate (FF37596) for the Bleed Valve.

METEOROLOGICAL INFORMATION

The automated weather observing system at the A.L. Mangham Junior Regional Airport, near Nacogdoches, Texas, located approximately 30 miles southwest of the accident site, reported wind from 250 degrees at 5 knots, 10 statute miles visibility, a clear sky, temperature 22 degrees Celsius

COMMUNICATIONS

The ignition specialist (FS crewmember in left front seat) was communicating with ground personnel on a U.S. Forest Service Tach channel at the time of the accident. He reported that he was resuming fire operations at 1352. At 1354, the ignition specialist made a “Mayday” call on the frequency. There were no further communications from the aircraft.

WRECKAGE AND IMPACT INFORMATION

The helicopter came to rest on its right side in a heavily wooded area on a heading of 135 degrees at coordinates North 31:45.425 West 94:00.244. The initial point of impact was identified as the tops of 50-foot trees located around the wreckage. Global positioning system (GPS) elevation at the accident site was 386 feet msl.

Fuselage – The right skid was fractured just forward of the forward mounting saddle, and the fuselage floor was fractured aft of the forward cross tube. Impact damage was observed on the right rear portion of the fuselage. Approximately 15 pounds of personal gear was found in the aft baggage compartment. Front right and left doors were not installed. The nose of the helicopter exhibited light crushing and all of the Plexiglas chin bubble was broken out. The right forward door post was bent inward and upward. Damage to the left side of the fuselage was unremarkable. No fuel found in the tank; however, an odor consistent with that of jet fuel was present at the accident site. The fuel boost pump access panels were removed, and the rear boost pump valve locking bar was found loose (approximately 0.065 in.). The forward boost pump valve locking bar was found slightly loose (0.035 in). The fuel quantity indicating float access panel was removed, and the forward fuel quantity indicator lead (“C” post) was found in a loose condition.

Cockpit – All circuit breakers were in and the battery and generator switches were found in the “ON” position. The Hobbs meter indicated a time of 2,837.4 hours. The directional gyro/attitude indicator switch was “ON”. The altimeter was set to a barometric pressure setting of 30.09 inches of Mercury. The directional gyro indicated a 210 degree heading. The artificial horizon displayed a nose-up right-bank attitude. The fuel valve switch was “ON”. The right seat pan was crushed downward approximately two inches. Located at the forward most point of the wreckage debris were two Interstate DCS-33 12-volt batteries, which were reportedly removed from the helicopter battery compartment during rescue operations.

Controls – Control continuity to all flight controls was established. The collective control lever exhibited overload fracture outboard of the collective attachment control collar. Movement at the collective hydraulic actuator bell crank resulted in corresponding movement at the collective attachment control collar, and pre-impact control continuity was confirmed from the collective up to the main rotor system. The cyclic control stick exhibited an overload fracture at the base of the stick forward of the attachment collar. Movement of the control tubes resulted in the movement from the attachment collar up to the hydraulic actuator. A visual inspection confirmed control continuity from the hydraulic actuator to the main rotor pitch change horns. The anti-torque control tube was fractured at the tail rotor control pedal bellcrank. Removal of the broom closet panel revealed that movement of the tail rotor control tube was prevented due to crushing of the surrounding structure. Movement of control tube above the crushing resulted in corresponding movement of the control tube forward of the fracture at the tailboom. The tailboom section of the anti-torque control tube was separated from the tailboom during the impact sequence, and each end exhibited overload fracture. Movement of the control tube at the forward end of the tail rotor resulted in a corresponding pitch change in the tail rotor blades, confirming pre-impact control continuity in the anti-torque pedal system.

Transmission and Main Rotor System – The transmission remained attached to the airframe and the main transmission pylon mounts remained in place and attached to main transmission. The transmission displaced in an aft, downward direction, forcing the K-flex coupling into the isolation mount cover. The forward attachment of the main driveshaft was separated at the K-flex coupling and had multiple overload fractures of the K-flex. Scoring of the roof panel adjacent to the fractured K-flex was observed, and was consistent with circumferential flailing of the drive shaft, indicative that the drive was rotating at impact. There was no notable damage to the main rotor control system, up to the rotating swashplate. One of the two pitch change control tubes, extending from the rotating swashplate to the pitch-change horn, displayed an overload fracture approximately mid-length. One rotor blade, which reportedly came to rest forward of the main wreckage, exhibited trailing edge impact damage that was consistent with the diameter of trees in the immediate vicinity of the impact. The blade fractured just outboard of the doublers. The other blade exhibited trailing edge impact, and chord wise damage. The blade also exhibited chord wise scoring throughout the length of the blade. Both blades displayed damage consistent with a high pitch, slow turning rotor at impact.

Tail Rotor – The tailboom fractured just forward of the tail rotor gear box mounting pad. As a result, the tail rotor control tube fractured in overload as described above. Additionally, the tail rotor drive shaft decoupled from the tail rotor gear box and the aft portion of the tailboom, along with the tail rotor gear box assembly was located approximately 15 feet from the main wreckage. The tail rotor assembly remained intact, though the tail rotor blades displayed both impact and fire damage. The tail rotor pitch change assembly moved freely by hand and resulted in corresponding pitch change of the tail rotor blades. As the tail rotor blades were rotated by hand, the gear box rotated freely without any identifiable grinding or binding. The chord wise accordion bending of the tail rotor blades was consistent with low, or no, power at impact. The forward end of the tail rotor drive shaft exhibited torsional shearing. Deformation of spacers was progressively more identifiable forward of the torsional separation in the tail rotor drive. All hanger bearings rotated freely with no evidence of heat distress. The aft end spline of the tail rotor drive shaft was found decoupled. The torsional shearing of the tail rotor drive shaft was consistent with sudden stoppage forward of the drive shaft fracture point.

Engine – The left side of the engine compartment appeared normal. The left side engine mounts were intact. The Pc pneumatic air line was intact to the PT Governor “T” fitting. All pneumatic system B-nuts were at least finger tight. The PT Governor pointer indicated a high power application. The linkage was intact in the engine compartment, but was separated on the hydraulic deck and at the collective itself. The right side engine compartment door was uniformly crushed into the right side of the engine. The right side engine mounts were bent. The horizontal fire shield was crushed inward from the right side. The right side compressor air discharge tube was crushed from a right side impact. The forward end was partially separated from the compressor scroll. The right side of the outer combustion case was partially crushed from impact. There were no noted separated pneumatic tubes. Some tubes were crushed and deformed from impact damage. All pneumatic system B-nuts were at least finger tight. The 4th stage power turbine wheel would not rotate with attempts at hand rotation.

PSD Machine – The PSD machine was found inside the aircraft still strapped in its installed position. Some slight denting damage was found on the hopper and the plastic lid was broken off from the hopper at the hinges. Plastic spheres had been ejected from the hopper and were strewn within the cabin and outside the cabin around the aircraft. Several nylon fabric bags containing additional plastic spheres were still securely attached to their tether and intact. In each of the two slipper blocks on the same side as the PSD machine controls, slipper blocks 1 and 2, a partially burned plastic sphere was found. That is, one plastic sphere in slipper block 1 and one plastic sphere in slipper block two. The other two slipper blocks, 3 and 4, were empty. Black ash residue was found on the outside of the lower portion of the feed chutes above blocks 1 and 2. The inside feeder control lever was in the up position. This lever in the up position allows the plastic spheres to feed into the two inner slipper blocks; slipper blocks 2 and 3. The outside feeder control lever was in the down position. In the down position, plastic spheres are restricted from entering blocks 1 and 4. The power control toggle switch was found in the on position. The speed control switch was found in the slow position. It is unknown whether the emergency water was used. The power cord from the machine to the hopper was disconnected. The main power cord from the machine to the helicopter was disconnected at the cannon plug. It is unknown whether these plugs where disconnected on prior to impact, during impact, or by the rescue personnel first to arrive at the accident site.

MEDICAL AND PATHOLOGICAL INFORMATION

An autopsy was performed on the pilot Dr. Brown, Forensic Pathologist, Jefferson County, Texas. Toxicological tests will be conducted at the FAA’s Civil Aeromedical Institute (CAMI), Oklahoma City, Oklahoma.

TESTS AND RESEARCH

The wreckage was recovered to Air Salvage of Dallas, Lancaster, Texas, on March 12, 2005, for further examinations. On March 16, 2005, representatives from the NTSB, USFS, Bell Helicopter, and Rolls-Royce Engines convened at Air Salvage of Dallas to examine the wreckage.

Airframe – When the airframe fuel filter was removed a small amount of retained debris was noted and the filter was clean. External power was applied to the helicopter to check gauges, warning horns, enunciator lights, and fuel boost pumps. When power was applied; the fuel gauge reading was +100 gallons, enunciator lights illuminated for the Fuel Pump, Tail Rotor Chip, Rotor Low, and Engine-Out. The low-rotor warning horn was audible. When power was applied to the airframe fuel boost pumps, no audible indication of pump operation was noted. The pumps were then removed and bench tested. They appeared to be operating, and pressure and volume were not verified.

Engine – The fuel flow control, which should be set on the “low” setting, was set to an intermediate position (an etched setting between “low” and “high”). Fuel from the fire shield to the fuel nozzle could not be verified due to premature removal of the fuel line. Both chip detectors were removed, and a slight amount of fuzz was found on one of the chip detectors. The tach generators were removed. N1 rotated freely and smoothly with continuity to N1 drive train. The engine fuel filter was removed and was found clean and full with clear and bright fuel. The oil filter was removed from the accessory gearbox and the oil was normal in color with no burning indications. The bleed valve was found in an open position and closed fully when actuated by hand and returned to an open position when released. The fuel nozzle was removed and its screen was intact and free of debris.

The engine was shipped to Aeromaritime in Mesa, Arizona, for disassembly and further examination on April 7, 2005. After the shipping container was opened a component inventory and part numbers were verified. Pneumatic system leak checks were conducted and a slight leak was found at the PC line filter. The right side shoulder of outer combustion case (OCC) was cut away, and the combustion liner and case appeared normal. The right side compressor air discharge tube, all pneumatic lines, oil lines, fuel lines, and attachments were removed and inspected. After removal of the power turbine governor and fuel control, the drive shaft and splines were found intact. The first stage nozzle shield was intact and first stage nozzle appeared normal. The first stage wheel appeared normal with no visible damage to blades. The second stage wheel appeared normal with no visible damage to blades. The gas producer rotor rotated free and smooth by hand. Turbine shafting (pinion gear coupling, turbine to compressor coupling, power turbine inner and outer shafts) were intact when removed. The power turbine rotor could not be rotated by hand due to impact damage. The power turbine rotor was then removed from the exhaust collector, after which, the power turbine was then able to rotated by hand. Rotational scoring on the fourth stage nozzle was noted corresponding to the tip path planes of the third and fourth stage wheels. A drive spline was inserted into the N1 drive train of the accessory gear box, manually turned, and fuel pump pumped fuel from outlet. No anomalies were noted on the N1 side of the accessory gear box. Both the accessory gear box and compressor rotated free and smooth by hand and there was no noted damage. In summary, no conditions were found that would have precluded the engine from normal operation. The fuel control and power turbine governor were packaged and sent to Honeywell, South Bend, Indiana for bench testing.

Functional testing of the fuel control unit did not reveal conditions that would have precluded normal operation. Functional testing of the power turbine governor found out of limits repeatability on the initial test run. The repeatability improved on each subsequent test run and after the fourth test run was within test specifications. Disassembly of the unit disclosed wear material (Teflon) on the spool valve assembly where it interfaces with the Teflon tube. Signs of vibration were evident on the spool bearing and flyweights. No other test points and part inspections revealed anomalies.

ADDITIONAL INFORMATION

Aerial Ignition Information as provided by the USFS:

Prescribed fire is a method of reducing the build-up of live and/or dead organic material in managed forest or range environments. This reduction in biomass has general short and long term benefits in that it may reduce the risk of uncontrolled wildfires, remove or prevent the establishment of undesired plant species, improve the health of established desired trees and plants, and improve wildlife habitat. Prescribed burning operations are performed in a variety of manners including hand ignition and aerial ignition which involves the application of burning material, generally fuel of some nature, to designated areas under specified and desirable meteorological and fuel conditions.

During aerial application of fuel, there are two primary methods of fire application: helitorch and Premo Mark III Aerial Ignition System. A helitorch utilizes gelled gasoline and is pumped from a barrel suspended beneath the helicopter. The Premo Mark III Aerial Ignition Device utilizes a small polystyrene ball, 32 mm in diameter, known more commonly as a plastic sphere, containing approximately 3.0 grams of Potassium Permanganate (KMnO) 99% reagent: an oxidizer used in a crystallized/powder form. When the KMnO comes in contact with Ethylene Glycol (anti-freeze), a combustive exothermic reaction occurs.

The Premo Mark III Aerial Ignition Device, often times referred to as a PSD (plastic sphere dispenser) machine or Ping-Pong ball machine, achieves the chemical reaction by physically injecting the plastic sphere with the ethylene glycol. The combustive reaction takes approximately 15 to 30 seconds to occur. During that time, prior to combustion, the machine ejects the ball, or essentially drops it. The PSD machine requires 24 volts DC power to operate.

For prescribed fire operations, the PSD machine is secured in the cabin of a helicopter. A nylon web strap runs from one side of the machine out the left side of the aircraft under the belly of the helicopter and then back into the aircraft right side attaching to the opposite side of the PSD machine. Normal configuration requires the removal of the right aft door (on Bell Helicopters) allowing the machine to extend over the door sill to drop the plastic spheres to the ground. The PSD machine comprises a hopper which holds approximately 450 plastic spheres; 4 chutes which funnel the plastic spheres to the slipper blocks; 4 needles which inject the plastic spheres with the ethylene glycol in a timed sequential order in each of the slipper blocks; two feed control levers; a 9 liter ethylene glycol tank; a 3.2 liter emergency water tank; a 2 amp drive motor; and a 2 amp glycol pump. Total weight of machine wet is approximately 98.0 lbs.

The feed control levers allow for the use of either 2 chutes or 4 chutes thus managing the quantity of balls injected and dropped from the machine. The PSD machine also has a slow and high speed controlling the rate at which balls are fed into the slipper blocks. During normal operations, spacing of the balls is achieved by a combination of the speed setting of the PSD machine, number of chutes used (1, 2 or 4 (using one chute requires the installation of a spacing kit which blocks off one of the chutes)), and helicopter airspeed. Normal PSD operations require helicopter flight below 500 ft. AGL and less than 50 mph. Optimum airspeed for application is 25-35 mph. Hovering out of ground effect (HOGE) often occurs. Application of the plastic spheres is generally performed in strips with the intent of allowing the fire to spread in a ‘backing’ manner.

It is not uncommon for a plastic sphere to become jammed or lodged in the PSD machine during operation. A jammed ball, if left unattended, could potentially ignite in the machine and then spread fire to the other plastic spheres in the machine and hopper. Operators are trained to respond at the first sign of a jam or smoke. Water can be injected into the slipper blocks from the water reservoir with a push of a button. Additional water is carried on board as a back up. With the first sign of smoke the pilot is alerted and on agreement with the operator will seek a landing sight to remove the PSD machine if necessary. If necessary, the PSD machine can be cut free from its restraining strap and dropped from the helicopter. However, development of a fire is rather slow and the resolution of any smoke or fire related problem is generally accomplished with the application of water.

Fueling History of the Accident Helicopter:

On March 6, 2005, Helicopter N85BH fueled at Angelina Co. Airport from the airports fuel truck. At approximately 0900 the helicopter took on 71 gallons of Jet-A fuel. The fuel serviceman stated that he thought that it was a “topped-off”. If that was the case, the helicopter would have had approximately 91 gallons on board before the first flight that morning. During the course of the day, the helicopter logged 4.3 flight hours. It was fueled two additional times from its own service truck for a total of 56.4 gallons, according to the fuel logs. The fuel burn rate for a 206B-3 is specified in the contract at 27 gallons per hour. 4.3 flight hours at 27 gallons per hour equals 116.1 gallons of fuel consumed during the day. The starting fuel quantity was approximately 91 gallons, plus the two fuelings equaling 56.4 gallons for a total of 147.4 gallons. Total gallons pumped, 147.4 gallons, minus the fuel consumed, 116.1 gallons, the result is 31.3 gallons remaining in the fuel tank at days end. The next fueling was on March 9, 2005. A total of 15 gallons was pumped bringing the on board total fuel to approximately 46.1 gallons. The fuel was dispensed from one of Angelina County Airport fuel trucks. On March 10, 2005, the helicopter flew from Angelina Co. Airport to H1 near the project area. Flight time from Angelina Co. Airport to H1 was approximately 30 minutes, consuming approximately 13.5 gallons. Fuel on board after flight would have been an estimated 32.6 gallons. At H1, before initiating the prescribed fire mission, N85BH took on 10 gallons of fuel from its service truck bringing the total to approximately 42.6 gallons. N85BH then flew a partial fuel cycle lasting approximately 46 minutes performing aerial ignition. Fuel consumed would have been approximately 20.7 gallons leaving 21.9 gallons on board. N85BH then took on an additional 20 gallons of fuel bringing the total on board fuel to approximately 41.9 gallons. N85BH then departed H1at 1347 after fueling returned to the burn area to resume ignition operations. N85BH was reported down 11 minutes later, at 1358. Fuel consumed during that 11 minutes would have been approximately 5 gallons. Usable fuel on board at the time of accident should have been approximately 36.9 gallons.

The helicopter wreckage and components were released to the owner after examinations were completed.

Contact a Helicopter Lawyer

If you have been injured or a loved one has been killed in a helicopter crash, then call us 24/7 for an immediate consultation to discuss the details of the accident and learn what we can do to help protect your legal rights. Whether the accident was caused by negligence on the part of the helicopter owner, hospital or corporation, the manufacturer or due to lack of training, poor maintenance, pilot or operator error, tail rotor failure, sudden loss of power, defective electronics or engine failure or flying in bad weather conditions, we can investigate the case and provide you the answers you need. Call Toll Free 1-800-883-9858 and talk to a Board Certified Trial Lawyer with over 30 years of legal experience or fill out our online form by clicking below:

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Tail Rotor Gearbox Failure Forces Water Landing

On August 19, 2004, approximately 0705 central daylight time, a Bell 412 twin-engine helicopter, N22347, sustained minor damage during a forced landing following a loss of tail rotor control near South Pass 65, an offshore platform located in the Gulf of Mexico. The helicopter was registered to and operated by Petroleum Helicopters Inc. (PHI), of Lafayette, Louisiana. The airline transport pilot-in-command, commercial co-pilot, and seven passengers were not injured. Visual meteorological conditions prevailed, and a company visual flight rules (VFR) flight plan was filed for the 14 Code of Federal Regulations Part 135 on-demand air taxi flight. The cross-country flight originated from Boothville, Louisiana, at 0635, destined for offshore platform Viosca Knoll 989.

The 11,750-hour co-pilot, who was in the right seat and piloting the helicopter, reported in the Pilot/Operator Aircraft Accident Report (NTSB Form 6120.1/2) that while in cruise flight he heard a loud bang followed by an uncontrolled 30-degree yaw to the right and a 15-20 degree “nose tuck.” The co-pilot stated that he attempted to correct the situation by lowering the collective and “trimming the yaw with the pedals;” however, the helicopter failed to respond to the inputs. Subsequently, the co-pilot reduced power and initiated an autorotation to the water. Prior to touchdown, the co-pilot successfully deployed the emergency floats. At an altitude approximately 10 feet above the water, the co-pilot “pulled pitch” until the helicopter settled onto the water. Both pilots and the passengers evacuated the helicopter into an inflatable life raft.

Tail Rotor Blade & Assembly Not Attached After Crash

When the helicopter was located in the water, the tail rotor blade assembly and a section of the 90-degree tail rotor gearbox were not attached to the fuselage. After recovery from the water, the helicopter and separated components were transported by truck to the PHI facilities near Lafayette, Louisiana. Examination of the helicopter was conducted by personnel from the Federal Aviation Administration (FAA), PHI, and Bell Helicopter.

The cockpit voice recorder (CVR) was removed from the wreckage and forwarded to the NTSB laboratories in Washington, D.C., for review. The 90-degree input quill and pieces of the 90-degree tail rotor gearbox assembly, tail rotor link assembly, and the upper and lower hinge of the lower access door were sealed in a box and sent to the engineering laboratories of Bell Helicopter for further examination.

The review of the CVR from the NTSB laboratories did not reveal any additional information or significant findings relative to the accident.

On September 23, 2004, at the field investigations laboratory of Bell Helicopter, Fort Worth, Texas, the examination of the remaining section of the tail rotor gearbox and access door hinges was conducted under the supervision of the NTSB investigator-in-charge, FAA, PHI, and Bell Helicopter. The examination revealed that the input quill exhibited no damage other than saltwater corrosion. The 90-degree gearbox case assembly fracture surfaces were also corroded from saltwater, which hindered determination of the fracture mode, but the surfaces that could be determined were consistent with an overload condition. The other fractured parts that were sent for examination were also determined to be caused by an overload condition.

Conclusions as to PHI Helicopter Accident

The NTSB could not deteremine the reason for the failure of the tail rotor gearbox.

Links to Other PHI Helicopter Crashes:

PHI Helicopter Crash in Louisiana – Jan 2009

Contact a Helicopter Lawyer

If you have been injured or a loved one has been killed in a helicopter crash, then call us 24/7 for an immediate consultation to discuss the details of the accident and learn what we can do to help protect your legal rights. Whether the accident was caused by negligence on the part of the helicopter owner, hospital or corporation, the manufacturer or due to lack of training, poor maintenance, pilot or operator error, tail rotor failure, sudden loss of power, defective electronics or engine failure or flying in bad weather conditions, we can investigate the case and provide you the answers you need. Call Toll Free 1-800-883-9858 and talk to a Board Certified Trial Lawyer with over 30 years of legal experience or fill out our online form by clicking below:

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FTW04IA217 NTSB Report

On August 19, 2004, approximately 0705 central daylight time, a Bell 412 twin-engine helicopter, N22347, sustained minor damage during a forced landing following a loss of tail rotor control near South Pass 65, an offshore platform located in the Gulf of Mexico. The helicopter was registered to and operated by Petroleum Helicopters Inc. (PHI), of Lafayette, Louisiana. The airline transport pilot-in-command, commercial co-pilot, and seven passengers were not injured. Visual meteorological conditions prevailed, and a company visual flight rules (VFR) flight plan was filed for the 14 Code of Federal Regulations Part 135 on-demand air taxi flight. The cross-country flight originated from Boothville, Louisiana, at 0635, destined for offshore platform Viosca Knoll 989.

The 11,750-hour co-pilot, who was in the right seat and piloting the helicopter, reported in the Pilot/Operator Aircraft Accident Report (NTSB Form 6120.1/2) that while in cruise flight he heard a loud bang followed by an uncontrolled 30-degree yaw to the right and a 15-20 degree “nose tuck.” The co-pilot stated that he attempted to correct the situation by lowering the collective and “trimming the yaw with the pedals;” however, the helicopter failed to respond to the inputs. Subsequently, the co-pilot reduced power and initiated an autorotation to the water. Prior to touchdown, the co-pilot successfully deployed the emergency floats. At an altitude approximately 10 feet above the water, the co-pilot “pulled pitch” until the helicopter settled onto the water. Both pilots and the passengers evacuated the helicopter into an inflatable life raft.

When the helicopter was located in the water, the tail rotor blade assembly and a section of the 90-degree tail rotor gearbox were not attached to the fuselage. After recovery from the water, the helicopter and separated components were transported by truck to the PHI facilities near Lafayette, Louisiana. Examination of the helicopter was conducted by personnel from the Federal Aviation Administration (FAA), PHI, and Bell Helicopter.

The cockpit voice recorder (CVR) was removed from the wreckage and forwarded to the NTSB laboratories in Washington, D.C., for review. The 90-degree input quill and pieces of the 90-degree tail rotor gearbox assembly, tail rotor link assembly, and the upper and lower hinge of the lower access door were sealed in a box and sent to the engineering laboratories of Bell Helicopter for further examination.

The review of the CVR from the NTSB laboratories did not reveal any additional information or significant findings relative to the accident.

On September 23, 2004, at the field investigations laboratory of Bell Helicopter, Fort Worth, Texas, the examination of the remaining section of the tail rotor gearbox and access door hinges was conducted under the supervision of the NTSB investigator-in-charge, FAA, PHI, and Bell Helicopter. The examination revealed that the input quill exhibited no damage other than saltwater corrosion. The 90-degree gearbox case assembly fracture surfaces were also corroded from saltwater, which hindered determination of the fracture mode, but the surfaces that could be determined were consistent with an overload condition. The other fractured parts that were sent for examination were also determined to be caused by an overload condition.

The reason for the failure of the tail rotor gearbox could not be determined.

SOURCE: NTSB

Contact a Helicopter Lawyer

If you have been injured or a loved one has been killed in a helicopter crash, then call us 24/7 for an immediate consultation to discuss the details of the accident and learn what we can do to help protect your legal rights. Whether the accident was caused by negligence on the part of the helicopter owner, hospital or corporation, the manufacturer or due to lack of training, poor maintenance, pilot or operator error, tail rotor failure, sudden loss of power, defective electronics or engine failure or flying in bad weather conditions, we can investigate the case and provide you the answers you need. Call Toll Free 1-800-883-9858 and talk to a Board Certified Trial Lawyer with over 30 years of legal experience or fill out our online form by clicking below:

Legal Help: Contact a Helicopter Lawyer

US District Judge Florence-Marie Cooper determined that two tower controllers Edward Weber and Cynthia Issa made a series of negligent decisions that led to the cause of a Robinson R44 Helicopter crash killing pilots Robert Bailey and Brett Boyd and severely injuring Gavin Heyworth, a stident plot. The Federal Aviation Administration (FAA) agreed to pay $4.5 million in damages to Heyworth of the November 6, 2003 helicopter crash at Torrance Municipal Airport (TOA). Gavin Heyworth was taking a solo instructional flight in a Robinson R22 at the time of the accident. Heyworth survived the crash, but suffered severe and life changing injuries.

In the original NTSB report, the Board found the following took place. On November 6, 2003, at 1528 Pacific standard time, a Robinson R22 Beta II, N206TV, and a Robinson R44, N442RH, collided in midair while in the traffic pattern at Zamperini Field, Torrance, California. Pacific Coast Helicopters was operating the R22 under the provisions of 14 CFR Part 91. Robinson Helicopter Company was operating the R44 under the provisions of 14 CFR Part 91. The solo student pilot in the R22 sustained serious injuries. The certified flight instructor (CFI) and the private pilot undergoing instruction (PUI) in the R44 sustained fatal injuries. Both helicopters were destroyed; a post crash fire partially consumed the R44. The R22 departed on a local instructional flight about 1442. The R44 departed on a local instructional flight about 1449. Visual meteorological conditions prevailed, and no flight plans had been filed. The R22 came to rest between runways 29R and 29L; approximate global positioning system (GPS) coordinates of the primary wreckage were 33 degrees 48.275 minutes north latitude and 118 degrees 20.536 minutes west longitude. The R44 came to rest on the departure end of runway 29L; approximate global positioning system (GPS) coordinates of the primary wreckage were 33 degrees 48.277 minutes north latitude and 118 degrees 20.584 minutes west longitude.

The instructor for the solo student had been watching him during his flight. The student flew the R22 from its parking area between taxiways D and E to a helipad north of runway 29R. The student practiced on the helipad, and then completed several touch-and-go landings to the helipad. He requested a return to his parking area. Upon hearing this request, the instructor turned the volume of his radio down, and turned away to talk to a bystander.

One witness reported that the R44 was speeding up and increasing in altitude as it took off straight ahead on runway 29L. He first observed the R22 when it was over runway 29R, or slightly north of it. The R22 was starting to descend as it was transiting across the left runway to the southwest, and appeared to be heading toward its landing area.

Other witnesses pointed out that the R22 was above the R44. The R44 seemed to increase its climb rate just before the collision. The two helicopters collided about 50 feet in the air over runway 29L. The R22 spun left several times before it contacted the ground.

A National Transportation Safety Board specialist interviewed the controllers, and obtained recorded radar data. He prepared a factual report, and pertinent parts follow.

Because of technical difficulties with the recordings of the ATC voice channels, times in this report prior to 1523:02 are based on draft transcripts provided early in the investigation. Times after that are valid times.

The R22 pilot first called the LC1 controller at 1442 requesting to fly from the Pacific Coast Helicopters parking area to the North Pad. He did not indicate that he was a student pilot; the controller did not think that he was a student, because his radio technique was good. He flew to the North Pad, which is a helicopter-only practice landing point that is at midfield on the north side of runway 29R.

Pilots operating at the North Pad typically fly right closed traffic patterns at 600 feet msl. They are required to keep their pattern within the lateral confines of the runway 29R displaced thresholds. They are required to contact the LC1 controller for each circuit around the pattern, or if they wish to extend their pattern beyond the 29R threshold limits.

The R44 pilot contacted the LC1 controller at 1449, and requested a northeast departure from the “antennae site,” which is at the intersection of the ramp area and taxiway G. The LC1 controller cleared him for takeoff from runway 29R, and the pilot departed the airport area to the northeast. The R44 pilot returned at 1505; he reported 6 miles north of the airport, and requested to operate on the North Pad. The controller advised him that the pad was in use (by the R22), and asked the pilot if he wanted to use the runway instead. The pilot accepted, and the controller instructed him to report a 2-mile right base entry. At 1507, the controller provided a traffic advisory of a departing helicopter, cleared him for the option on runway 29R, and told him to enter right closed traffic. The pilot continued routine traffic pattern operations until 1525, including landings on runway 29L.

At 1523:14, the R22 pilot requested a North Pad takeoff and landing at PCH parking. PCH parking referred to the parking area used by Pacific Coast Helicopters. It is west of the tower, on the ramp between taxiways D and E. The controller instructed him to hold, and the pilot acknowledged holding. At 1524:33, the controller advised him that he could proceed in right traffic to the North Pad after a Cessna passed off his left. At 1524:56, the R22 pilot transmitted, ” takeoff and land PCH parking.” At 1524:59, the LC1 controller responded, “Helicopter six tango victor fly westbound.” Between 1525:18 and 1525:52, there was some confusion caused by the pilot of a departing helicopter (29M) who incorrectly used the call sign 2RH when requesting departure from the ramp area. The controller resolved the confusion.

At 1526:01, the controller cleared the pilot of the R44 to, “make your base your discretion two niner left cleared for the option”, and at 1526:15, in the same transmission, continued, “helicopter six tango victor make a right turn to the downwind.” At 1526:19, the R22 pilot acknowledged, but only with his call sign. At 1526:32, the controller again cleared the R44 for the option on runway 29L, and the pilot acknowledged.

At 1526:59, the controller advised the pilot of the R22, “ah you’re gonna cross midfield as soon as I get a chance.” At 1527:17, the controller instructed the R22 pilot to, “turn right,” and the pilot acknowledged with his call sign. At 1527:49, the controller transmitted, “Helicopter six tango victor runway two niner right cleared to land.” At 1527:53, the R22 pilot acknowledged with his call sign. At 1527:54, the controller transmitted, “turn right helicopter six tango victor runway two niner right cleared to land.” There was no communication from the R22 pilot. At 1528:12, the LC1 controller advised the R44 pilot, “robinson two romeo hotel caution for the helo oh.”

A review of recorded radar data showed a target that turned off the right downwind leg, crossed runway 29R, and approached runway 29L in the immediate area of the accident. The last target for this track was at 1528:10, approximately 2 seconds before the collision. A plot of this track on a street map indicated that it was perpendicular to the runways at 1527:49, and the target was between Lomita Boulevard and Skypark Drive. At 1527:54, this target was still approaching Skypark Drive and north of runway 29R. After crossing Skypark about 5 seconds later, the target appeared to turn toward the southwest, and the last two targets were approaching runway 29L at a shallow angle. Another target turned from right downwind to base to final for runway 29L. Its last target appeared at 1527:15; its track lined up with runway 29L, and was westbound abeam the approach end of runway 29R.

PERSONNEL INFORMATION

R22 Pilot

A review of Federal Aviation Administration (FAA) airmen records revealed that the R22 pilot held a student pilot certificate, and a first-class medical certificate issued in September 2003.

An examination of the student pilot’s logbook indicated that his first flight occurred on September 7, 2003. He had an estimated total flight time of 32 hours. He logged 16 hours in the last 30 days. He had solo time on two previous flights that totaled about 1.5 hours.

R44 CFI

A review of FAA airman records revealed that the pilot held a commercial pilot certificate with ratings for rotorcraft helicopter and instrument helicopter. He had a mechanic certificate with ratings for airframe and powerplant. He had a second-class medical certificate issued on October 3, 2003. It had no limitations or waivers.

No personal flight records were located for the CFI. The FAA indicated that the pilot reported that he had a total time of 8,900 hours on his last medical application.

R44 PUI

A review of FAA airman records revealed that the pilot held a private pilot certificate with ratings for airplane single engine land, multiengine land, and instrument airplane; he also had a helicopter rating. He held a second-class medical certificate issued on January 16, 2003. It had the limitation that the pilot must wear corrective lenses.

No personal flight records were located for the PUI. The FAA indicated that the pilot reported that he had a total time of 370 hours on his last medical application. An application for the Robinson safety course indicated that he had 52 hours in rotorcraft; all were in this make and model.

AIRCRAFT INFORMATION

R22

The helicopter was a Robinson R22 Beta II, serial number 2753. A review of the helicopter’s logbooks revealed that it had a total airframe time of 2,974.6 hours. The logbooks contained an entry for an annual inspection dated February 1, 2003. A 100-hour inspection occurred on October 30, 2003, and the helicopter accumulated 23.9 hours since its completion. The Hobbs hour meter read 2,974.6 at the accident site. The time since an airframe overhaul was 783.7 hours.

The engine was a Textron Lycoming O-360-J2A, serial number L-32698-36A. Total time recorded on the engine was 1,976.2 hours, and time since major overhaul was 789.3 hours.

R44

The helicopter was a Robinson R44, serial number 0002. A review of the helicopter’s logbooks revealed that the helicopter had a total airframe time of 1,046.5 hours. The logbooks contained an entry for an annual inspection dated April 3, 2003. It had a 100-hour inspection on July 15, 2003. It accumulated 74.7 hours since that inspection.

The engine was a Textron Lycoming O-540-F1B5, serial number L-25143-40A. Total time on the engine was 2,672.5 hours, and time since major overhaul was 479.1 hours.

COMMUNICATIONS

Both helicopters were in contact with the Torrance airport traffic control tower (ATCT) on frequency 135.6.

AIRPORT INFORMATION

The Airport/ Facility Directory, Southwest U. S., indicated that runway 29L was 3,000 feet long and 75 feet wide. The runway surface was asphalt. Runway 29R was 5,001 feet long and 150 feet wide. The runway surface was asphalt and concrete.

WRECKAGE AND IMPACT INFORMATION

The FAA and Robinson were parties to the investigation. Investigators from the Safety Board and the parties examined the wreckage at the accident scene.

The debris field was on runway 29L, and extended over 500 feet. The first identified debris (FID) pieces were shards of Plexiglass and R44 main rotor blade. A section of one R44 main rotor blade was 135 feet from the FID, and another piece was about 145 feet. The tip cap from this blade was at 174 feet.

A piece of R22 lower frame was 234 feet from the FID. The main wreckage of the R22 was at 270 feet, and about 70 feet north of the runway. About the same distance and another 120 north was a piece of R44 spar. Another piece of R44 spar was on the runway at 412 feet.

The main wreckage of the R44 came to rest inverted at 495 feet and about 25 feet right of the runway centerline. The left skid of the R44 was at 515 feet, and just off the right edge of the runway.

The last piece of debris was the main rotor of the R44 at 525 feet and near the runway centerline. One entire blade was present; it exhibited a wavy appearance, and buckled in several places. It had a scrape mark that was 2 inches wide across half of the blade chord (from the leading edge) and 36 inches from the tip. This scrape was dimensionally similar to the aft right strut of the R22. The second blade fractured and separated about 2.5 feet from the rotor hub; the fracture ran chordwise from the leading edge back to the trailing edge doubler, with the doubler remaining intact. The blade bent aft about 120 degrees at this point. The next 4.5 feet of blade exhibited some buckling. The rest of the honeycomb/skin section of the blade separated, as did the outboard 4.5 feet of spar. Two pieces of this blade, one being about 1.5 feet with the tip weights (3.5 pounds), and the other being the tip cap, were in the R22’s engine compartment behind the left pilots seat.

The R22 exhibited rotational scoring on the fan shroud, on the fan itself, and the tail rotor drive shaft twisted.

MEDICAL AND PATHOLOGICAL INFORMATION

The Los Angeles County Coroner completed autopsies of both pilots in the R44. The FAA Bioaeronautical Sciences Research Laboratory, Oklahoma City, Oklahoma, performed toxicological testing of specimens of the pilots.

Analysis of the specimens for the CFI contained no findings for tested drugs in the liver. They did not perform tests for carbon monoxide or cyanide. The report contained the following findings for volatiles: no ethanol detected in muscle; 28 (mg/dL, mg/hg) ethanol detected in the brain; 11 (mg/dL, mg/hg) methanol detected in muscle; 212 (mg/dL, mg/hg) methanol detected in the brain; and 30 (mg/dL, mg/hg) of 2-butanol detected in the brain. The report stated that the ethanol found in this case might potentially be from postmortem ethanol formation, and not from the ingestion of ethanol.

Analysis of the specimens for the PUI contained no findings for carbon monoxide, cyanide, volatiles, and tested drugs.

TESTS AND RESEARCH

The aft strut of the right skid of the R22 separated about 1-foot from the bottom of the skid, with the skid placed in its approximate installed orientation. The fracture surface was relatively flat, and the round tubing bent inboard. The right side of the engine exhibited crush damage to the mid portion of the rocker covers and valves that was similar in dimension to an R44 rotor blade. The narrow band of damage continued around to the accessories between the engine and cabin. Tubing in this area and the back of the front seat exhibited fractures across a similar plane. A section of rotor blade from the R44 was imbedded in the lower left side of the R22 near the battery, which was behind the front left seat.

Robinson personnel provided the following information.

The radius of the R44 rotor, from the centerline of the hub to the tip is 198 inches. The distance between the aft strut and forward strut of the R22 is 50.5 inches (centerline to centerline). The distance between the left and right struts on the R22 (at the level of the first contact point) is 55 inches centerline to centerline).

Based on unique impact markings and color transfers, the Robinson air safety investigators opined that first R44 blade hit the center of the aft right strut of the R22, 36 inches from the tip of the blade, cutting through the strut without hitting anything else. The second R44 blade hit the R22 across the valve covers of the engine, 29 inches higher then the first hit. The blade fractured and separated about 48 inches from the tip. This blade also made contact with the engine cooling fan and scroll. The tip of this blade struck the cabin of the R22, 48 inches forward of the aft strut (12 inches of movement from the first blade hit to second blade hit).

This information could not yield a collision angle; however, it placed the R22 above, slightly forward of the R44, and on a similar course.

Visibility Study

The IIC and Robinson investigators examined exemplary helicopters at the Robinson factory. Pilots of the approximate height of the R44 pilots sat in an R44. Looking forward, they could see as high as the outboard 8 feet of the main rotor blade. Looking to the right, they could see eye level and no higher. They could see nothing aft. A pilot of similar height to the R22 pilot sat in the right seat of an R22. He could see about 10 degrees aft with the left door not installed. With it installed, he could only see abeam his seat, and no higher than eye level. The accident R22 had the door installed.

There were several procedures for helicopters to return from the helipad to the parking ramp. One controller stated that helicopters could travel directly across both runways, and land on the ramp if there was no conflicting traffic. The next method was to land on runway 29R; then the controller would clear the pilot to hover taxi across the runways to the ramp. A third method was to have the pilot cross both runways at midfield, land on taxiway A, and then taxi to the ramp.

The student pilot’s flight instructor described the procedures for returning to the ramp from the North Pad. One method was to request a direct air taxi crossing both runways and taxiway A. A second method was to depart west on the up wind. After reaching pattern altitude at the end of runway 29R, the controller would clear the pilot for a left turn to the south. After crossing Airport Drive, the pilot could turn downwind, fly east to the east “tees,” turn left to base, and then turn left to final for taxiway A or runway 29L. A third method was to take off to the west to pattern altitude, make a right crosswind turn, and then turn to a right downwind until abeam the east end of the 29R threshold. Here the pilot would turn right base and right final for either taxiway A or 29L. A fourth method was to take off to the west, make a right crosswind, make a right downwind to midfield, make a right turn to cross both runways at midfield, turn left to a left downwind for 29L, and then left base and final to either taxiway A or 29L.

The LC1 controller stated that he intended to have the R22 pilot depart the pad westbound along runway 29R, turn right to the downwind, and then turn right and land on 29R. He would then clear the pilot to hover taxi along taxiway C to the ramp. As the R22 passed the North Pad area as it traversed eastbound on the downwind leg, he instructed the pilot to turn right. There was no response, and the helicopter did not turn. A few seconds later, he cleared the pilot to turn right and land on 29R. He did not hear an acknowledgement, so he repeated the instruction. The R22 turned at that point. The controller saw the R22 north of the approach end of 29R with the nose pointed roughly at the tower. He then looked away to his radar display as he worked another aircraft. After talking to that aircraft, he decided to advise the R44 that the R22 would be landing on 29R abeam the R44 as it departed. He looked back just as the helicopters collided.

The Safety Board investigator-in-charge (IIC) released the R22 wreckage to the owner’s representative on February 12, 2004. The IIC released the R44 wreckage to the owner’s representative on April 30, 2007.

**This narrative was modified on May 14, 2007.

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Legal Help: Contact a Helicopter Lawyer