Mesa, AZ, USA
N820NA
Eurocopter AS350 B3
The helicopter impacted a taxiway after a experiencing a loss of anti torque control during an emergency landing. The pilot reported a hydraulic system failure and diverted to the nearest airport. During the emergency landing, after slowing the helicopter and descending through 100 feet, the helicopter yawed to the left. The pilot attempted to apply right pedal but was met with more force than he was expecting to overcome in moving the anti torque pedals. The pilot attempted to keep the helicopter in a level attitude; however, he was unable to stop the helicopter from spinning. The helicopter touched down, and rotated on the ground until the right landing gear skid collapsed and the main rotor system impacted the ground. The helicopter came to rest on its right side. The engine continued to operate and the main rotor blades repeatedly impacted the ground until the fire department was able to shutdown the engine by directing a stream of firefighting foam into the engine intake. The pilot turned off the electrical master switch and exited the helicopter. The AS350 B3 is equipped with four servo actuators, three for the main rotor system and one for the tail rotor blade pitch control. The flight manual emergency procedures section notes that without any hydraulic pressure, the anti torque pedal operating force is very high on the helicopter. Post accident examination of the wreckage revealed the hydraulic test switch was depressed. Activation of the hydraulic test switch will dump hydraulic pressure and is used during the preflight inspection to test the hydraulic accumulator pressure. If it is depressed during flight, the hydraulic system will dump pressure and mirror a hydraulic system failure. No other anomalies were noted with the hydraulic system during post accident examinations and testing. The recommended emergency procedures for a hydraulic system failure dictates that the pilot make a flat approach over a clear landing area and land with slight forward speed.
HISTORY OF FLIGHT On November 3, 2003, about 1745 mountain standard time, a Eurocopter AS350 B3 helicopter, N820NA, utilizing call sign Native 9, experienced a loss of control and crashed on a taxiway during an emergency landing at Falcon Field Airport (FFZ), Mesa, Arizona. Native American Air Ambulance, Inc., was operating the helicopter under the provisions of 14 CFR Part 135. The airline transport pilot, two medical crewmembers, and one medical patient were uninjured; the helicopter sustained substantial damage. The aero medical cross-country flight departed Show Low, Arizona, about 1700, and was en route to Scottsdale, Arizona. Day visual meteorological conditions prevailed, and a company flight plan had been filed. The primary wreckage was located at 37 degrees 27.39 minutes north latitude and 111 degrees 43.42 minutes west longitude. The pilot reported he was 3 to 4 miles north of FFZ when the hydraulic warning light illuminated and the associated audio warning activated. He was in communication with Phoenix Sky Harbor International Airport's (PHX) air traffic control tower controllers. He told the PHX controller that he was diverting to FFZ due to a hydraulic problem. The pilot elected to remain on PHX tower frequency, and requested that they notify FFZ tower of his intent to land. The pilot slowed the aircraft down and turned off the hydraulic system using the switch located on the collective lever. The helicopter was approaching FFZ from the northwest, on a heading of 130 degrees. The pilot reported the winds were 250 degrees at 8 knots. He intended, after crossing the runways from the north to the south, to perform a run on landing on the Bravo taxiway and land at the base of the control tower. Prior to crossing the runways (04R-22L and 04L-22R), and after descending below 100 feet agl, the pilot slowed the helicopter, and it started to yaw to the left. The pilot attempted to apply right pedal but was met with more force than he was expecting to overcome. He attempted to keep the helicopter in a level attitude; however, he was unable to stop the helicopter from spinning. The helicopter touched down, and rotated on the ground until the right landing gear skid collapsed. The helicopter came to rest on its right side. Before the pilot could get the engine shutdown or the main rotors stopped, the main rotor blades impacted the ground. The pilot stated that he was unable to conduct an emergency engine shutdown because of the "violent shaking of the helicopter." The engine shutdown after the fire department applied firefighting foam to the engine intake. During a telephone interview with the National Transportation Safety Board investigator-in-charge (IIC), the pilot was asked why he had not declared an emergency during the hydraulic failure. He stated he did not feel it was an emergency; he felt very confident in his abilities to land the helicopter without the hydraulics. PERSONNEL INFORMATION A review of Federal Aviation Administration (FAA) airman records revealed that the pilot held an airline transport pilot certificate with a rating for helicopters. The pilot held a certified flight instructor (CFI) certificate with a rating for helicopters and instrument helicopters. The CFI certificate was issued on January 03, 2003, and expires on January 31, 2005. The pilot held a second-class medical certificate that was issued on April 2, 2003. It had no limitations or waivers. An examination of the pilot's logbook indicated he accumulated an estimated total flight time of 2,288 hours. He logged 171 hours in the last 90 days and 19 hours in the last 30 days. He had an estimated 398 hours in this make and model. He completed a 14 CFR Part 135 airman competency check ride on October 21, 2003. The pilot was conducting the third dispatched flight of the day on his first day of employment with Native Air. AIRCRAFT INFORMATION The helicopter was a single engine multipurpose Eurocopter AS350 B3, serial number 3602. A review of the helicopter's logbooks revealed a total airframe time of 1,097 hours at the last 100-hour inspection. The Hobbs hour meter read 1,132.0 at the accident scene. The total time on the helicopter at the time of the accident was 1,143.3 hours. The helicopter's engine was a Turbomeca Arriel 2B1 engine, serial number 22342. Total time on the engine at the last 100-hour inspection was 1,097.6 hours. Examination of the maintenance and flight department records revealed no unresolved maintenance discrepancies against the helicopter prior to departure. All four hydraulic servo accumulators were serviced on July 21, 2003, at 814.1 hours. On the day of the accident, the pilot reported "sluggish" cyclic controls during the hydraulic system ground check. The three main rotor accumulators were serviced with nitrogen to 15 bars at 74 degrees. The maintenance was completed at the aircraft's total time of 1140.2 hours. The pilot completed weight and balance calculations prior to the flight. The total weight was calculated to be 4,938.0 pounds with a center of gravity (cg) of 130.8 inches. The maximum gross weight for the helicopter was 4,961 pounds with forward and aft cg limits of 126.4 inches and 134.7 inches, respectively. HYDRAULIC SYSTEM The AS350 B3 is equipped with four servo actuators, three for the main rotor system and one for the tailrotor blade pitch control. The Eurocopter AS350 Instruction Manual states: The helicopter can be controlled without servo actuators, but this requires the pilot to apply non-negligible forces that are difficult to gauge. These control loads are absorbed by hydraulic servo actuators so that the pilot can fly the helicopter PRECISELY and EFFORTLESSLY. In case of loss of hydraulic pressure, accumulators in the main rotor servo actuators provide a small energy reserve, giving the pilot time to reconfigure in the safety configuration. The B1, B2 and B3 versions are fitted with a yaw load compensator. YAW LOAD COMPENSATOR ON VERSIONS B1, B2 and B3 Without any hydraulic pressure, the pedal operating force is very high on the B1, B2 and B3 versions. This is why a hydraulic device or "load compensator" has been mounted in parallel with the tail rotor servo actuator. In case of hydraulic failure (pump inoperative, leakage, etc.) ,the accumulator is kept charged by -non-return valve in the pressure circuit -pressure relief valve set at 55 bar (N.B. The system's operating pressure is 40 bar) [1 bar equals 14.7 PSI] -solenoid valve always closed and opened by the pilot The AS350 B3 Flight Manual states: A tail rotor servo-control mounted on the tail boom actuates a rod which controls the tail rotor spider bellcrank. In the event of a hydraulic system failure, a load compensating servo in the tail rotor linkage limits the yaw pedal operating load. The hydraulic accumulator that supplies the compensation system may be depressurized by means of a HYD. TEST pushbutton. The pilot is informed of hydraulic system fault conditions by a red "HYD" low-pressure warning light on the Warning-Caution-Advisory Panel, and by an aural alarm, both of which are actuated by the pressure switch on the regulator unit. The gong sounds to warn of: -Rotor speed (NR) between approx. 250 and 360 r.p.m. (continuous sound). -Rotor speed above 410 r.p.m. (intermittent sound). -Hydraulic pressure drop (below 30 bars). It is operative only if the "HORN" pushbutton is pushed in. When this push-button is out, at nominal rotor speed, the HORN light of the warning-caution-advisory panel is on. Alarm procedure (if HORN sounds): -If the HYD warning light is on: The malfunction is in the hydraulic system. The switch on the collective pitch lever can be used to cut off all hydraulic power by opening the three solenoid valves on the main rotor servo-controls to depressurize the system. A push-button [HYDR TEST] on the control console is used: -to test the hydraulic accumulators by opening the regulator unit solenoid valve -to depressurize the tail rotor load compensating servo. The Flight Manual lists the Hydraulic Failure and Tail Rotor emergency procedures as: Hydraulic System Failures In cruise flight: Reduce speed, entering into a side slip if necessary, then cut off hydraulic pressure by actuating the switch situated on the collective pitch control lever. Section 3.3 of the manual then advises the pilot to make a flat approach over a clear landing area and land with slight forward airspeed. Tail Rotor Drive Failure Loss of tail rotor in power on flight results in a yaw moment to the left: the extent of such rotation will depend on the power and speed configuration at the time the failure occurs. Failure of the Tail Rotor in Hover or at Low Speed O.G.E. (Out of Ground Effect): reduce collective pitch moderately, to reduce yaw torque, and simultaneously start to pick up speed. Failure in Forward Flight -In forward flight reduce the power as much as possible and maintain forward speed (weathercock effect), select a suitable landing area for a steep approach at a hover enabling a reasonable coordinated flight. -On final approach, shut down the engine and make an autorotative landing at the lowest possible speed. Tail Rotor Control Failure -Set I.A.S 70 knots (130 km/hr), in level flight. -Press the hyd accumulator test push button (this cuts off hydraulic power to the yaw servo control and depressurizes the load-compensating servo accumulator). After 5 seconds, reset the test button to the normal position. -Make a shallow approach to a clear landing area with a slight sideslip to the left. Perform a run-on landing: the sideslip will be reduced progressively as power is applied. METEOROLOGICAL INFORMATION The closest official weather observation station was Falcon Field Airport. The accident occurred at the airport. The elevation of the weather observation station was 1,394 feet msl. An aviation routine weather report (METAR) for FFZ was issued at 1646, and reported the wind from 220 degrees at 10 knots; visibility 40 statute miles; few clouds at 10,000 feet; temperature 18 degrees Celsius; and altimeter setting 29.96 inches of Hg. WRECKAGE AND IMPACT INFORMATION The Mesa Airport Fire Department and Police Department personnel secured the accident site, and photographically documented the position and condition of the wreckage and the cockpit switch positions prior to the arrival of the FAA. Inspectors from the FAA responded to the scene and examined and documented the wreckage prior to the recovery process. The following buttons located in the center console of the helicopter were found in the activated position (depressed): Generator; Anti Collision Light; Position Lights; Hydraulic Test; Horn Arm; Pitot Heat; Instrument Lights 1 & 2; and Attitude Gyro. The avionics master switch located on the instrument panel was also in the on position. The pilot did not remember activating the hydraulic test button or know how the hydraulic test button became activated. The battery master switch was in the off position. The pilot reported that he had turned off the battery master switch prior to exiting the helicopter. TESTS AND RESEARCH The FAA, American Eurocopter (AEC), and Native Air were parties to the investigation. The IIC and the party members examined the wreckage at Air Transport, Phoenix, Arizona, on December 2, 2003. The hydraulic system switches were activated, and both switches operated as normal, per the AEC representative. The hydraulic test button was activated; the manifold solenoid was heard and felt to be activating. The hydraulic isolation switch was activated; the three main servo solenoids were heard and felt to be activating. The Servo Accumulators were pressure tested with the following results: Right Lateral Servo Ser# 1706 190 PSI Forward Pitch Servo Ser# 218 190 PSI Left Lateral Servo Ser# 309 190 PSI Tail Rotor Servo Ser# 978 160 PSI The servo accumulators are serviced with nitrogen and pressurized to 15 bars of pressure (220 PSI) [1 bar = 14.7 PSI] at 20 degrees centigrade. The AEC factory representative stated pressures recorded from the accident accumulators were low but the accumulators would be functional during a hydraulic failure. Flight control continuity was established from the cockpit to the main rotor and tail rotor pitch change links. The hydraulic system drive belt was found intact and was removed from the hydraulic pulley by rotating the pulley. The interface between the S40 coupling and the hydraulic drive spline was undamaged. The chip detector on the hydraulic pump was removed and was found clean and free of debris. Investigators removed the hydraulic pump from the mounting bracket. The splines were intact and undamaged; the o-ring was present and undamaged. The S-40 coupling assembly and bracket were removed. The splines were inspected and were found intact and undamaged. Grease was present and viable, extruding from the face of the S-40 coupling. The hydraulic fluid from the reservoir was drained into a clean bucket. No visible contamination was noted. The hydraulic pump's filter screen was inspected. A nominal amount of debris was noted on the screen, though it was not blocked. The hydraulic pump was disassembled and inspected. No sign of any anomalies was evident, and the gears and case were pristine. Electrical power was applied to the aircraft system by using the helicopter's battery. Investigators observed the HYD light illuminate on the pilot caution panel. The electrical connection for the hydraulic pressure switch and the HYD warning light circuit were tested by manually disconnecting at the hydraulic pressure switch, and the HYD warning light extinguished. The hydraulic system was removed from the aircraft and sent to the manufacturer in France for further testing. On February 17, 2004, investigators from the Safety Board, American Eurocopter (AEC), and the French Bureau Enquetes Accidents (BEA) met at the Eurocopter factory in Marignane, France. The hydraulic components removed from N820NA, S/N 3602, were tested to new component standards. The four servos were each installed on a test stand and tested. All were serviceable and operated within manufacturer's specifications. The hydraulic pump was installed on a test stand and was tested. The pump was found serviceable and operated within manufacturer's specifications. The hydraulic distribution block was tested and found serviceable and operated within the manufacturer's specifications. ADDITIONAL INFORMATION The IIC released the wreckage and returned the tested components to the owner's representative.
The pilot's inadvertent activation of the hydraulic test switch, which resulted in a loss of the hydraulic system pressure, and his failure to maintain directional control.
Source: NTSB Aviation Accident Database
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