NTSB Factual Report

HISTORY OF FLIGHT

On January 5, 2006, about 0945 Hawaiian standard time, a Eurocopter EC130B4 (ECO-star), N11QD, experienced an engine deceleration and loss of main rotor rpm, and made a hard forced landing in Honokohau Valley, near Lahaina, Hawaii. Blue Hawaiian was operating the helicopter with the call sign “Blue 21” under the provisions of 14 Code of Federal Regulations (CFR) Part 135 as a nonscheduled, on-demand tour flight. The helicopter sustained substantial damage. The pilot and five passengers were not injured. Visual meteorological conditions prevailed, and a company visual flight rules (VFR) flight plan had been filed. The local area flight departed Kahului Airport (OGG), Kahului, Hawaii, about 0930.

According to the operator, the flight was on a "complete" island tour, which is 60 minutes in length. When Blue 21 did not return to Kahului Airport at its designated time, Blue Hawaiian personnel attempted to locate the flight. At 1117, they were notified that the subject helicopter had been located below the Jurassic Falls.

In the pilot's written statement to the National Transportation Safety Board, he reported that as he entered Honokohau Valley, he slowed to 30 knots heading upstream to show the passengers the falls. He heard the main rotor warning, checked the rotor tachometer, and saw that it was decreasing. The pilot entered into an autorotation to make a forced landing. He also noted that the rotor rpm (revolutions per minute) was in the "green." However, because there were no available landing sites in his current direction of travel (upstream), he did a 180-degree right turn. He tried to reapply power, but the low rotor horn sounded again. He lowered the collective and "removed twist grip from the flight gate." The pilot reported that he flared for landing. The helicopter came down in trees, with the main rotor blades contacting the treetops.

PERSONNEL INFORMATION

The operator reported that the 56-year-old pilot held an airline transport pilot certificate with ratings for helicopter and airplane single-engine land, multi-engine land, and instrument airplane.

A second-class medical certificate was issued on August 31, 2005, with a limitation that the pilot must possess corrective glasses for near vision.

The operator reported that the pilot had accumulated a total flight time in all aircraft of 13,750 hours. Approximately 200 hours were logged in the last 90 days, and 75 hours in the last 30 days. An estimated 1,703 hours had been accumulated in the make and model helicopter involved in the accident, with a total rotorcraft time of 11,670 hours. A biennial flight review was completed on December 16, 2005.

AIRCRAFT INFORMATION

The helicopter was a Eurocopter EC130B4, serial number 3363. The operator reported that the helicopter had a total airframe time of 4,836 hours at the last 100-hour inspection.

The engine was a Turbomeca Arriel 2B1, serial number 23017. Total time recorded on the engine at the last 100-hour inspection was 3,700 hours, and time since major overhaul was 700 hours.

Fueling records at OGG established that the helicopter was last fueled on January 5, 2006, with the addition of 72 gallons of Jet-A fuel. The operator reported that there were no unresolved maintenance discrepancies against the helicopter prior to departure.

It is noteworthy that this helicopter was reportedly the first EC130B4 delivered to a customer by Eurocopter. Also, the helicopter sustained a lightning strike in August 2004 while being operated by a lessee in Linden, New Jersey. A review of maintenance records from this event revealed that a functional check was performed on the electrical system.

ENGINE SYSTEMS:

ENGINE FUEL CONTROL:

The fuel control, as it relates to the free turbine rpm N2 control, is as follows: when the load varies, the Digital Engine Control Unit (DECU) uses the input parameters (N2, P0, T1, N1, P3, Torque) and the anticipators (collective pitch and yaw) to compute a new setting for the metering unit via the stepper motor in order to bring N2 back to the set point value.

The helicopter is also equipped with an Engine Back-up Control Ancillary Unit (EBCAU) board and associated back-up fuel metering unit electrical actuator. In case of a failure of both DECU channels, the EBCAU is designed to automatically govern the engine. The electronic board is installed on ASU No. 1 and No. 2 boards, and is electrically activated by the DECU. It carries out the simplified back-up governing function using the N2 value.

ENGINE OPERATING CONTROLS:

The engine is controlled via a selector on the instrument panel, a twist grip on the collective pitch lever, and an automatic back-up system: EBCAU (Engine Back-up Ancillary Control Unit).

• Starting selector:

- OFF: engine shut down. The guard is raised.

- ON: the DECU runs the automatic starting sequence. This is the normal setting in flight. In the ON position, the rating of the engine depends on the position of the twist grip (IDLE or FLIGHT).

• EBCAU test selector:

-The "EBCAU TEST" control button is used to switch to the back-up mode to test the engine back-up control system on the ground. When the "EBCAU TEST" button is pressed, the position of the main control actuator controlled by the DECU is frozen. The fuel flow is then monitored by the EBCAU, which controls the back-up control valve. The amber light of the "EBCAU TEST" button comes on and the red "GOV" light illuminates on the warning and caution panel.

• "Forced idle" Microswitch:

In autorotation training, the pilot twists the grip to move it out of the "FLT" detent to the "IDLE" setting and to activate a microswitch. The DECU then adjusts the engine to the idle rpm. Moving the grip back to the "FLT" detent resets the DECU in "flight" mode.

• Description of warning lights:

- Red "GOV" light, indicating a major engine fuel control system failure with seizure of the metering unit or the EBCAU on ground test.

- Amber "GOV" light, indicating a minor failure resulting in degraded engine fuel control. When flashing, the light indicates a failure not affecting the engine fuel control system, such as loss of redundancy.

- Red "TWT GRIP" light, indicating the twist grip is no longer in the "FLT" detent.

COLLECTIVE ANTICIPATOR SYSTEM:

The controlling function adjusts the gas generator speed to balance the power supplied with the power needed and in order to maintain the constant N2 (or NR) speed. In addition, the collective anticipator system supplies an immediate load variation signal from the collective potentiometer (XPC) to the DECU allowing it to react more rapidly to demand changes.

The DECU uses a voltage signal from the collective anticipator (or collective potentiometer (XPC)) to anticipate power requirements induced by pitch changes. The DECU supplies the XPC potentiometer with a 10V power supply. The XPC potentiometer then returns a modified voltage to the DECU which is proportional to the collective pitch position.

If there is a collective anticipator failure, the DECU is designed to switch to the proportional/integral mode to maintain an NR speed equal to the N2 set point whatever the pitch requirement. This should result in the illumination of the amber GOV light on the caution warning panel and a VEMD Test Code 122.

There are three conditions designed to trigger a collective-pitch anticipator failure (VEMD code 122) in the event of XPC signal drift:

• XPC signal value in flight decrease below 5% (out of range) resulting NR decrease below 360 RPM

• XPC signal value increase above 95% (out of range)

• XPC gradient test (350°/s = 7% / 20ms)

If XPC signal remains within range (5%; 95%) and gradient is lower than 350°/s, no failure will be detected.

METEOROLOGICAL CONDITIONS

The closest official weather observation station was Kahului Airport (OGG), which was located 10 nautical miles (nm) on a magnetic bearing of 285 degrees from the accident site. An aviation routine weather report (METAR) for OGG was issued at 0854 local time and reported: winds from 20 degrees at 15 knots; visibility 10 miles; skies 1,000 feet scattered; temperature 22.8 degrees Celsius; dew point 15.6 degrees Celsius; altimeter 30.17 inHg.

WRECKAGE AND IMPACT INFORMATION

The accident scene was in a very remote location. The helicopter was recovered to a hangar at OGG and stored for further examination.

TESTS AND RESEARCH

INTIAL AIRCRAFT EXAMINATION

After the aircraft was recovered, the NTSB and FAA examined the wreckage at OGG and performed the following examination. The helicopter was equipped with on-board video recorders, which were recovered and shipped to the Safety Board Vehicle Recorders laboratory for further examination. Power was applied to the helicopter. Investigators went through all Vehicle Engine and Management Display (VEMD) maintenance pages and documented the VEMD screen findings. Investigators removed the VEMD, Digital Engine Control Unit (DECU), engine and (ASU) cards for further examination.

On January 18, 2006, the accident engine was installed into a test cell and run at Turbomeca USA facility in Grand Prairie, Texas. No abnormalities were identified that would have explained the reported loss of engine power.

VEMD & DECU EXAMINATION

On January 24, 2006, the VEMD was examined at the manufacturer's facility with the oversight of Bureau d’Enquetes et d’Analysis (BEA). Flight duration recorded on the VEMD was 13 minutes 18.5 seconds. Neither failures nor over-limits were recorded by the VEMD. NG cycles (generator) and NF cycles (free turbine) were not recorded by the VEMD because the recording process was interrupted before the end of the flight.

On February 16, 2006, representatives of Turbomeca and BEA examined the DECU at Turbomeca factory. The DECU recorded a power up duration of 14 minutes 21 seconds. The DECU recorded a single failure block. Within this block, three electrical failures were recorded. Power discrete input selector failures were recorded 10 times in the whole DECU records. The previous selector failure was recorded in the DECU, identified as the VEMD flight number 2804. The recorded duration of that flight was 87 seconds.

According to BEA personnel, the VEMD and DECU data analysis seems to indicate that the failures recorded by the DECU occurred simultaneously with the VEMD power supply loss.

SUBSEQUENT AIRFRAME EXAMINATION

Investigators conducted an examination between June 20 and June 25, 2006 of the wreckage which had been stored in a T-hangar at Kahului Airport (OGG).

Airframe

The airframe maintained its structural integrity. The front left and right seats had attenuated, with greater attenuation on the left side. Damage to the collective micro switch assembly underneath the pilot’s collective was present. There was greater overall damage to the left side of the fuselage than the right. Both left and right canopy windscreens of the cabin were broken. The center windscreens (upper and lower) were intact. The doors were removed. The fuel cutoff handle was in the cutoff (pulled) position. The collective lever was broken from its mount. The twist grip was jammed just slightly low of the FLI position. The microswitch assembly underneath the floor had been damaged.

Continuity was verified for cyclic and anti-torque controls.

The tail boom remained attached to the fuselage. The section of the tail boom that houses the aircraft battery sustained impact damage consistent with the dimensions of the main rotor blade impacting the tail boom. The battery was ejected from the battery compartment.

All tail rotor drive shaft hanger bearings remained attached but appeared to be shifted slightly aft. The forward section of the tail rotor drive shaft had been removed. Neither the shaft nor the flexible couplings on each end of the shaft exhibited torsional damage. The intermediate and rear shafts remained intact and rotated freely, which resulted in rotation of the Fenestron assembly. There was a mark within the Fenestron shroud consistent with impact from the tip of the Fenestron blade in a forward and slightly downward direction.

The landing gear cross tubes were broken. The skids, which had been removed, were relatively straight.

The yellow and blue star arms of the main rotor head were broken at a 45-degree angle. All three of the blade sleeves remained attached to the hub. The yellow and blue frequency adapters had detached from their respective star arms. The blade sleeves were trailing their normal location. According to the manufacturer, all three main rotor blades exhibited damage consistent with low rotor rpm. The blue blade exhibited chord wise bending.

Electrical System

Upon inspection of the electrical system, all harnesses and cables appeared intact with the exception of the collective microswitch assembly. The helicopter was equipped with additional video recording equipment in the right side baggage compartment in the vicinity of the 67K relay and DECU. Wiring for the equipment was secured to the existing electrical and radio bundles without any segregation precaution.

Electrical continuity tests were performed and continuity was confirmed on the following circuit systems using Eurocopter wiring diagrams: Starting Arriel 2B1 wiring; Arriel 2B1 engine control wiring; VEMD wiring; dumping VEMD wiring; and power ASU wiring. Because of the damage to the collective lever/twist grip assembly, it was not possible to perform a continuity test on the collective grip wiring system.

An insulation test was performed on the engine related electrical systems using a 50-volt Ohmmeter, with no anomalies noted.

Following the electrical system tests, several components and aircraft wiring related to engine starting and regulation were removed. These included the on/off switch, 67K connector (controls engine shut down), and wiring to the collective lever (controls flight/idle). The individual wires were inspected. One CAPTON wire appeared to be worn through the outer layer of insulation (opaque). However, the wire was tested for insulation and was found that a clear layer of insulation remained. No evidence of electrical arching was observed on any of the wires inspected.

Aircraft Equipment Functional Test

Investigators accomplished a functional test of the aircraft by following Eurocopter Technical Ground Test (Document No. 350A045224FA) with accompanying test equipment normally utilized by Eurocopter when aircraft leave the production line for initial delivery.

The VEMD, engine and DECU had been removed for analysis and were not present at the examination. Initially, the test equipment was installed on the helicopter in absence of these components. The helicopter was powered using an external power unit. No significant anomalies were noted.

A spare VEMD, DECU, and other engine sensor/equipment necessary for the test were installed on the helicopter. A spare collective lever/twist grip assembly was installed as well. The following functional tests were performed: 10V check of test equipment; 20V functional test of VEMD-ASU; 30V VEMD test; 40V engine parameters and alarms (low oil pressure, fuel and fire detection alarms); 50V engine parameters and alarms (temperature, oil pressure, T4, NTL (NF), NG, OAT, and engine torque, of which step 5 was not performed as a frequency generator was not available); 60V engine parameters and alarms (IPL alarm); 70V regulation starting (twist grip); 80V regulation starting (collective and trim potentiometer. Note: The collective lever and microswitches sustained damage in the accident sequence, therefore, a collective lever supplied by Blue Hawaiian was used to perform the test); 90V regulation starting (alarms and indications); 100V regulation starting (regulation); 110V regulation starting (ventilation/CRANK switch); 120V regulation starting (emergency command); 130V regulation starting (starting); and 140V parameters (NR indications and alarms).

The Nr magnetic pick-up sensor was tested for resistance. Resistance for the coil (pins 2 and 3) that provides a signal to the 39 Delta P1 (ASU 1/Alarm) measured 1.28 ohms. Accor

NTSB Probable Cause

An intermittent electrical continuity failure and short of the DECU-XPC wiring harness, which resulted in an uncommanded and an initially undetected engine deceleration and a resulting loss of main rotor rpm. Also causal was the manufacturer's inadequate installation of the wiring harness at the time of manufacture.