NTSB Factual Report

HISTORY OF FLIGHT

On January 4, 2009, at 1409 Central Standard Time (CST), a Sikorsky S-76C++ helicopter, N748P, registered to and operated by PHI, Inc. (PHI), as a 14 CFR Part 135 air taxi flight using day visual flight rules (VFR), crashed into marshy terrain approximately 7 minutes after takeoff and 12 miles southeast of the departure heliport. The helicopter sustained substantial damage. Both pilots and six of the seven passengers were killed, and 1 passenger was critically injured. The helicopter departed Lake Palourde Base Heliport, a PHI base (7LS3), in Amelia, Louisiana, and was en route to the South Timbalier oil platform ST301B to transport workers from two different oil exploration companies. No flight plan was filed with the Federal Aviation Administration (FAA), nor was one required. A company flight following plan was filed with the PHI Communications Center that included weather updates, pertinent advisories, and position reports. The flight was tracked via Outerlink, a satellite based fleet-tracking system used by the PHI communications center based in Lafayette, Louisiana.

The helicopter departed 7LS3 at 1402. The helicopter’s flight track, recorded by the Outerlink system, ended about 7 minutes after departure, at 1409. There were no reports of any distress calls or emergency transmissions from the flight crew on the PHI radio frequencies, or on any monitored air traffic control frequencies.

A search and rescue operation was initiated at 1414 after the US Air Force received a 406 MHz Emergency Locator Transmitter (ELT) distress signal with the helicopter’s unique identifier and location. Notification was made to PHI and the United States Coast Guard. Shortly thereafter, the helicopter wreckage was found partially submerged in a marshy bayou, near the location of the last Outerlink track.

Data and audio recordings retrieved from the helicopter’s combination cockpit voice recorder (CVR) and flight data recorder (FDR) indicated that the helicopter was in level cruise flight at 850 feet mean sea level (msl), traveling at 135 knots indicated air speed, when a loud "bang" occurred. Immediately following the "bang," sounds were recorded consistent with rushing wind, engine power reductions on both engines, and main rotor rpm decay.

AIRCRAFT INFORMATION

General Information

The twin-engine, 14-seat, 2-year-old helicopter was equipped with glass cockpit instrumentation, a combination cockpit voice recorder (CVR) and flight data recorder (FDR), an enhanced ground proximity warning system (EGPWS), solid state quick access recorder (SSQAR), and a VXP vibration recorder. The two Turbomeca Arriel 2S2 turbo shaft engines were equipped with digital engine control units (DECU). All of these devices were recovered and evaluated for recorded information.

Engine Control Quadrant Design

The Sikorsky S-76C++ helicopter has an overhead engine control quadrant that houses two engine fire extinguisher T-handles, two engine power control levers (ECL), two fuel selector valve control levers, and various switches for other essential functions. The fire extinguisher T-handles, which are about 4 inches aft of the captain’s and first officer’s windshield, are normally in the full forward position, and are held in place by a spring-loaded pin that rests in a detent. Force is required to move the handles out of the detent and aft. In the event of an in-flight engine fire indication, the affected engine's fire extinguisher T-handle will illuminate, and the flight crew is trained to pull the illuminated handle full aft. In doing so, a mechanical cam on the T-handle lifts the trigger on the ECL out of a wedge-shaped stop, allowing the handle to move aft, which reduces the fuel flow to the affected engine. Eventually, the fuel flow to the engine is shut off as the fire extinguisher T-handle continues aft and pushes the fuel selector valve to the OFF position. The fire extinguisher system is then automatically armed and ready for the pilots to release the fire extinguishing agent into the appropriate engine compartment.

The S-76C++ engine control quadrant is physically similar to previous models of the S-76 series (S-76A, S-76B, S-76C, S-76C+), in that the ECLs are located in the overhead engine control quadrant. The S-76A, S-76B, and S-76C use push-pull cables to manually control the engine throttle positions on each engine’s hydro-mechanical units. The S-76C+ uses an electronic engine control design with a manual push-pull cable reversionary mode. The ECLs of the S-76C++ series are based on a dual-channel allelectronic engine control design, in that the ECLs are attached to potentiometers that transmit ECL position electronically to each respective electronic engine control unit.

Windscreens

In 2007, about 2 years prior to the accident, PHI removed the original, factory-installed laminated glass windshields in N748P and installed lighter-weight cast acrylic windshields manufactured by Aeronautical Accessories Incorporated (AAI). The Federal Aviation Administration approved use of the replacement windshields under Supplemental Type Certificate SR01340AT, issued to AAI on April 16, 1997. The FAA also issued Parts Manufacturer Approval to AAI on August 3, 1998, for manufacturing of the replacement windshields. The helicopter’s windshields were replaced again in 2008, about 1 year before the accident, due to cracking at the mounting holes.

Low Rotor Speed Warning Systems

The S-76C++ helicopter's integrated instrument display system (IIDS) provides the flight crew with engine and main rotor system performance information. Three IIDS screens are mounted in the instrument panel; one in front of the captain, one in front of the co-pilot, and one in the center of the instrument panel (the main rotor [Nr] information is only displayed on the pilot's and copilot's IIDS.) The Nr data is provided to the flight crew by a broad colorbar on the right side of the IIDS. The IIDS Nr colorbar is green when the helicopter's Nr is between 106 and 108 percent, yellow when the Nr is between 91 and 105 percent, and red when Nr is 90 percent and below, warning the flight crew of a critical, unsafe flight conditions requiring immediate action. The helicopter was not equipped with an audible alarm or a master warning light to alert the flight crew of a low Nr condition, nor was one required by 14 CFR Part 29.The IIDS also provides a visual caution legend such as "1 out of fly" to the crew any time an engine speed selector is out of the FLY detent with the weight off wheels.

PERSONNEL INFORMATION

A review of the accident flight crew's training records indicated that both pilots had accomplished all required training and had completed emergency initial and recurrent training in ground school and in the Sikorsky S-76C++ simulator.

The 63-year-old captain had approximately 15,373 flight hours when the accident occurred, of which 14,673 were in rotorcraft; 8,549 as pilot-in-command; and 5,423 in the S-76. He held an airline transport pilot certificate for helicopters, and a commercial pilot certificate for fixed-wing airplanes. He also held an instrument rating for helicopters and airplanes. His last FAA flight proficiency check was on October 27, 2008. His first class FAA medical was issued on August 11, 2008, with a restriction that he wear corrective lenses while flying. He had flown 219 hours in helicopters in the preceding 90 days.

The 46-year-old co-pilot had approximately 5,524 flight hours, of which 1,290 were in helicopters, with 962 in the S-76. He held an airline transport pilot certificate for helicopters and a commercial certificate for fixed-wing airplanes. He also had a flight instructor certificate valid for giving instruction in single/multi-engine airplanes and helicopters. His instrument rating was valid for both airplanes and helicopters. His last FAA flight proficiency check was on April 25, 2008, and his last FAA first class medical was issued on February 26, 2008, with a restriction that he wear corrective lenses while flying. He had flown 205 hours in helicopters during the preceding 90 days.

METEOROLOGICAL INFORMATION

The weather conditions reported at Amelia, Louisiana, at 1430 CST were scattered cloud layers at 1,500 feet and 3,500 feet; a broken cloud layer at 10,000 feet; visibility 10 miles; winds at 160 degrees at 6 knots; temperature of 24 degrees Celsius; and a dew point of 19 degrees Celsius.

WRECKAGE AND IMPACT INFORMATION

The majority of the major components were accounted for and recovered from the accident scene. Examination of the accident site indicated that the helicopter impacted on its left side on an approximate heading of 120 degrees. Extensive deformation on the left side of the helicopter was noted and exhibited signatures consistent with hydrodynamic and soft terrain impact. The largest portion of the helicopter came to rest in a marsh area and consisted primarily of the upper deck from above the cockpit area to the aft engine compartment. The corresponding lower fuselage section was adjacent to the upper deck. The two sections remained attached by wiring harnesses only.

The tail boom was separated from the fuselage at the forward attach point (fuselage station 300) and exhibited extensive impact damage. The vertical pylon was partially separated from the tail boom and was deformed to the right side of the aircraft. The left-hand horizontal stabilizer was separated from the tail boom and the right stabilizer was attached but damaged.

The number 2, 3, and 4 tail rotor driveshaft segments, along with their respective hanger bearings, appeared to have been pulled forward during the impact sequence and exhibited minimal rotational scoring/damage. The coupling disk packs were securely attached to each associated coupling and exhibited minimal distortion. The number 4 driveshaft was observed separated approximately six inches forward of the intermediate gearbox attach point. The number 5 driveshaft was securely attached to the intermediate gearbox and the tail rotor gearbox.

The tail rotor system exhibited extensive damage. The tail rotor gearbox output housing and gear separated from the gearbox center housing. The gear teeth appeared normal and did not exhibit any pre-impact anomaly. The blue, yellow, and black tail rotor blades exhibited minimal rotational impact damage. The black and yellow tail rotor blades had fractured just outboard of their respective hub retention plates. The blue blade was securely attached to the hub and the red blade was observed broken with “broom straw” damage approximately seven inches from the root end of the blade. The remaining section of the tail rotor gearbox housing was observed securely attached to the upper vertical pylon. The tail gearbox magnetic chip plug was removed and observed free of ferrous debris.

The four main transmission mounts were securely attached to the deck structure and did not exhibit any pre-impact damage. The transmission rotated freely. Continuity from the two input shafts to the main rotor head and tail takeoff was established. The magnetic chip plugs were removed and observed clean with oil still remaining inside of them. The transmission fluid level was observed to be in its normal state.

The main rotor blade system exhibited impact damage consistent with low-speed rotation. The yellow and black main rotor blades were observed attached to the hub and predominately intact with some impact damage. The red blade had separated approximately 27 inches from the root end of the blade, and the remaining portion of the blade was recovered. The blue blade exhibited two separations, one at approximately 40 inches from the root and another about 12 feet from the root. With the exception of a small tip portion, approximately 8 feet of the blue blade was not recovered.

The main rotor hub was observed securely attached to the main rotor shaft and exhibited substantial impact damage. The drive links and swashplate were intact and did not exhibit pre-impact damage. The four pitch control rods were observed securely attached and undamaged. The three main rotor servo actuators and associated hydraulic lines were securely attached and did not exhibit any pre-impact anomalies. The three primary servo actuators all displayed a part number of 76650-09805-111 and had experienced a recorded 1,104 hours of operations since overhaul, with an approximate total time of 3,400 hours.

The engines were mounted in the airframe engine compartment. The No. 1 (left) engine exhibited significant deformation of the left side and both engines were deformed from their respective mounts in a left-to-right direction. There was no evidence of fire, fuel leaks, or oil leaks.

The No. 1 and No. 2 axial compressor wheels rotated easily. There was evidence of some ingestion of mud and debris. The axial compressor wheel and blades were intact with some tip bending. The power turbine wheel rotated easily. The power turbine wheel and blades were intact and there were signs of blade rub on the bottom of the housing, consistent with a hard landing or impact. The short shafts were observed pulled out of the engine output coupling for both engines but securely attached to the transmission inputs. The flexible couplings and triangular flange exhibited minimal deformation.

Complete control continuity could not be established from the cockpit aft to the mixing unit due to impact damage and crush deformation of the airframe. Control continuity was established from the mixing unit to the flight control servos to the main rotor blades. No pre-impact anomalies were observed. All hydraulic fluid reservoirs were found to be full of hydraulic fluid with no evidence of leakage noted.

FLIGHT RECORDER INFORMATION

Data from the Penny and Giles combination FDR and CVR were analyzed at the NTSB's Recorders Laboratory with download assistance from the manufacturer's facility in England and the US Army Safety Center in Fort Rucker, Alabama. Both recorders captured the entire accident flight.

The CVR recorded the sound of a bang and a loud air noise followed by a substantial increase in the background noise level that was recorded on both intercom microphones and the cockpit area microphone. Less than a second after the bang and loud air noise, the CVR captured the sound of decreasing rotor and engine rpm. Seventeen seconds later, the recording ended.

The non-volatile memory (NVM) from the engines' digital Engine Electronic Control Units (EECUs) was successfully downloaded and no faults were recorded.

TESTS AND RESEARCH

Engine Examinations

On January 22, 2009, the No. 2 engine, a Turbomeca Arriel 2S2, SN 21010, was disassembled under NTSB supervision. Other than impact damage, no anomalies were noted. The engine’s hydromechanical unit (HMU) was removed and examined. It was determined that it could be run on a test bench. Prior to running the HMU, the position of the resolver and the manual microswitch were determined. The resolver was at 59.33 degrees, which equates to a fuel flow of about 137 pounds per hour and an N1 of about 86.8 percent. The manual microswitch was found to be in the open or neutral position, indicating that the HMU was in the automatic mode. The HMU was then run on the test bench with no significant out-of-limits noted; however, there was a fuel leak observed that appeared to be from the varilip seal. As fuel pressure increased, the fuel leak decreased.

On January 23, 2009, the No. 1 engine, a Turbomeca Arriel 2S2, SN 21022, was disassembled. Other than impact damage, no anomalies were noted. The engine’s HMU was removed and examined. Impact damage to the unit precluded it from being run on the test bench. The position of the resolver and the manual microswitch were determined. The resolver was at 28.38 degrees, which equates to a fuel flow of about 250 pounds per hour and an N1 of about 98.0 percent. The manual microswitch was found to be in the open or neutral position, indicating that the HMU was in the automatic mode.

Flight Computer Memory Download and Testing

The FZ-706 Digital Flight Computer, P/N 7015480-903, S/N 05061626, was connected to test equipment and the error codes were successful

NTSB Probable Cause

(1) the sudden loss of power to both engines that resulted from impact with a bird (red-tailed hawk), which fractured the windshield and interfered with engine fuel controls, and (2) the subsequent disorientation of the flight crewmembers, which left them unable to recover from the loss of power. Contributing to the accident were (1) the lack of Federal Aviation Administration regulations and guidance, at the time the helicopter was certificated, requiring helicopter windshields to be resistant to bird strikes; (2) the lack of protections that would prevent the T handles from inadvertently dislodging out of their detents; and (3) the lack of a master warning light and audible system to alert the flight crew of a low-rotor-speed condition.