NTSB Factual Report

HISTORY OF FLIGHT

On September 6, 2016, about 1340 eastern daylight time, a Sikorsky S-61N, N805AR, was destroyed when it impacted a field after experiencing a dual loss of engine power while in a hover near Palm Bay, Florida. The airline transport pilot, the commercial-rated copilot, and the maintenance crewmember were fatally injured. The helicopter was registered to EP Aviation LLC and was being operated by AAR Airlift Group under the provisions of Title 14 Code of Federal Regulations Part 91 as a post-maintenance flight. Visual meteorological conditions prevailed, and a company visual flight rules flight plan was filed for the local flight that departed Melbourne International Airport (MLB), Melbourne, Florida, at 1324.

According to the operator, the helicopter's fore/aft pitch servo had recently been removed and replaced. Subsequently, three functional check flights (FCF) were required to be completed. Two FCFs were completed uneventfully the day of the accident, and the crewmembers were conducting the final FCF when the accident occurred. One of the maneuvers to be performed during the final FCF was rearward flight at a computed airspeed of 20 knots. According to the cockpit voice recorder (CVR), the flight crew performed two of these rearward FCF maneuvers during the accident flight.

Video taken by a ground witness recorded the helicopter performing the first rearward maneuver about 200 ft above ground level. Correlation of the video to the CVR showed the helicopter flying rearward at 1337:25, when the copilot stated the rear speed was 15 knots. At 1337:40, the helicopter continued to fly rearward as the copilot stated the rear speed was 20 knots. Two thumping sounds were recorded on the CVR at 1337:42 and 1337:44, when the rear speed was about 31 knots, but corresponding sounds could not be identified in the ground witness video. The ground witness video then showed the nose of the helicopter pitch down as the helicopter transitioned from rearward flight to a forward left 90° turn and continued forward in straight and level flight. During the recovery maneuver, an event occurs which caused a 2-second region of overdriven audio to be recorded on the CVR. This occurred during the left pedal turn and while the helicopter was approaching 90° to the rearward flightpath. The helicopter then flew an orbit in increasing altitude, and the video ended.

According to the CVR, the pilot took over control of the helicopter at 1336:41 and the copilot, who was receiving flight training on the day of the accident, identified the thumping sounds as a compressor stall, and the pilot agreed. The pilot then told the maintenance crewmember that they were returning to MLB. At 1338:02, the flight crew discussed the perceived compressor stall and engine exhaust gas temperatures. Then at 1338:26, the pilot told the maintenance crewmember that they were going to try the maneuver again in a different direction relative to the wind (with the wind off the nose), and the maintenance crewmember stated that it was okay with him. At 1339:46, the pilot was recovering from flying rearward when there was a change in background noise, which the maintenance crewmember identified as a compressor stall. The audio for the cockpit area microphone was overdriven again from about 1339:45 to 1339:48. At 1339:48, the copilot stated that the "AFCS is back on" while there was another change in background noise consistent with a decay in drivetrain rpm. The recording ended at 1339:55.

There were no known witnesses to the impact.

PERSONNEL INFORMATION

The pilot in the left seat held an airline transport pilot certificate with a rating for rotorcraft helicopter and commercial privileges in airplane single- and multi-engine land and instrument airplane. In addition, the pilot held a flight instructor certificate with ratings for rotorcraft helicopter and instrument rotorcraft. The pilot's most recent Federal Aviation Administration (FAA) first-class medical certificate was issued on December 5, 2015. According to the operator, the pilot was hired in 2012 and completed all required company training. At the time of the accident, he had accrued a total flight experience of about 6,347 hours of which 5,743 hours were in helicopters and 1,780 of those hours were in the same make and model as the accident helicopter. The pilot had flown 114 hours and 25 hours during the 90-day and 30-day periods preceding the accident, respectively; all of these hours were flown in the same make and model as the accident helicopter.

The copilot in the right seat held a commercial pilot certificate with ratings for rotorcraft helicopter and instrument helicopter. His most recent FAA second-class medical certificate was issued on May 16, 2016. According to the operator, the copilot was hired on July 31, 2016. He had completed company-required ground training and was in the process of completing flight training at the time of the accident. The copilot had accrued a total flight experience of 4,090 hours, all of which were in helicopters. Before the day of the accident, he did not have any flight experience in the same make and model as the accident helicopter but was qualified to act as second-in-command.

AIRCRAFT INFORMATION

The 41-seat capacity, tricycle-gear helicopter, serial number 61717, was manufactured in 1974. It was powered by two 1,500-horsepower General Electric CT58-140-2 turboshaft engines. The helicopter was maintained under a continuous airworthiness program. Its most recent inspection was a phase five check, which was completed on August 25, 2016. At that time, the airframe had accumulated 40,296.2 total hours of operation. The No. 1 engine had accumulated 708.9 hours since major overhaul (23,235 hours since new), and the No. 2 engine had accumulated 4,520.2 hours since major overhaul (26,259 hours since new). Following the phase five check, the helicopter had flown about 1.2 hours during the two previous FCFs. At the end of the second flight, prior to the accident flight, the flight crew reported total fuel onboard was 2,200 lbs.

METEOROLOGICAL INFORMATION

At 1353, the recorded weather at MLB, which was located about 8 miles north of the accident site, included wind from 070° at 11 knots, visibility 10 miles, and few clouds at 5,000 ft.

WRECKAGE AND IMPACT INFORMATION

The helicopter came to rest upright in a field with no debris path noted. The wreckage was oriented on a magnetic heading of about 190°. A postcrash fire consumed the cockpit and cabin. The tail boom transition section exhibited partial thermal damage, and the tail boom remained intact. The five main rotor blades and the five tail rotor blades remained attached to their respective rotor hubs. The main and tail rotor blades exhibited signatures consistent with low rotational energy at ground impact. Four of the five main rotor blades exhibited partial thermal damage, and one main rotor blade exhibited thermal damage along its entire span. One tail rotor blade was fractured about 1 ft outboard of the attachment bolt; the outboard section of the separated blade was found on the ground next to the tailrotor. Another tail rotor blade was partially separated about 1 ft outboard of the attachment bolt, and its tip was embedded in the ground. Drivetrain continuity was confirmed between the main transmission and the tail rotor gearbox.

Both engines remained attached to the airframe and exhibited fire damage. Examination of the engines revealed that the first stage compressor blades of both engines exhibited little or no leading-edge damage. Both engine fuel control units were found with their respective control shafts in the "FLIGHT" position. Both engines were separated from the main gearbox at the aft end of the high-speed shaft. A boresope inspection of the helicopter's main gearbox was performed. No evidence of thermal damage was observed on the internal components of the main gearbox, and the interior coatings appeared in good condition. The gear teeth exhibited normal wear patterns consistent with typical service wear. No anomalous damage was observed with the gears and bearings. Additionally, the main rotor drive system was evaluated for continuity and all components operated normally. Manual rotation of the rotor brake disk resulted in corresponding movement of the main rotor head, tail takeoff pinion, No. 1 tail rotor drive shaft, and the two input pinion splined couplings (normally attached to the aft end of the high-speed shafts).

Freewheeling Units (FWU)

The left and right FWUs were subsequently examined at the manufacturer's facility. The ramp-roller clutch engaged on the right FWU when rotated in the drive direction and disengaged when rotated in the freewheeling direction. No anomalous damage was observed on the right FWU.

When the left FWU was rotated, the ramp-roller clutch remained disengaged and freewheeled in both directions of rotation. The splined nut was removed and exhibited no anomalous damage. Additionally, the ramps, roller elements, bearing cage, and bearing outer race did not exhibit anomalous damage. Further examination of the left FWU revealed that both bearing cage pins were observed to be in the retracted position. The bearing cage pin housings were labeled "A" and "B" for identification purposes. When both pin assemblies were removed from the bearing cage, each pin remained contained (and retracted) within its sleeve. Circumferential scoring was observed within the inner diameter of the bores that housed the bearing cage pin sleeves. Pin "A" was removed from the sleeve by pushing an Allen key through the hole at the rear end of the sleeve. Resistance was felt when pushing the pin out of the sleeve. Once the pin was pushed out, the spring exited the sleeve. The pin and spring were reinstalled into the sleeve, and subsequent compression of the pin resulted in normal operation of the pin-and-spring mechanism. Pin "B" was not disassembled and remained in its compressed position for X-ray examination. Circular contact marks were seen on the bearing cage tangs, consistent with bearing cage pin contact with the tangs during impact forces. The circular contact marks were more pronounced compared to those observed on the right FWU.

No. 1 Engine

A teardown examination of both engines was performed at an independent helicopter facility.

Application of 25 foot-pounds (ft-lbs) of torque on the No. 1 engine starter dog (jaws) did not result in any movement of the engine core. Removal of the stator vane actuator (SVA) pilot valve cover revealed that the feedback cable remained connected to the pilot valve arm. The pilot valve exhibited no anomalous damage. The SVA pilot valve internal housing exhibited witness marks consistent with contact with the helical exterior of the feedback cable. The feedback cable could not be manually actuated and was found in a position consistent with the SVA in the closed position. The SVA housing was melted off from its piston, and remnants of the melted actuator housing were observed below its normally installed location.

After removal of the accessory gearbox (AGB), application of 25 ft-lbs of torque on the starter dog did not result in any movement of the engine core. Attempting to manually rotate the AGB using the radial drive shaft was unsuccessful. The centrifugal fuel purifier (CFP) was removed from the AGB, and its external drive splines exhibited no anomalous damage. After removal of the CFP, manual rotation of the AGB resulted in minimal rotation with evidence of binding. The lubrication pump was removed, and both its internal and external splines exhibited no anomalous damage. The lubrication pump was not able to be manually rotated. The fuel control was removed, and the splined connection between the fuel control and fuel pump exhibited no anomalous damage. Thermally degraded O-rings were observed when the fuel control was removed from the fuel pump. After removal of the fuel control, manual rotation of the AGB resulted in limited rotation and a corresponding rotation of the drive splines to the lubrication pump and the CFP but not to the fuel pump-to-fuel control drive splines. During removal of the fuel pump from the AGB, one of the six studs separated and was retained in its mounting flange. The fuel pump could not be manually rotated. After fuel pump removal, manual rotation of the AGB was successful and resulted in a corresponding rotation of the drive splines for the fuel pump, lubrication pump, and CFP. The fuel pump external splines exhibited no anomalous damage. The front frame accessory drive gears did not exhibit anomalous damage. At this point in the engine disassembly, manual rotation of the starter dog was successful and resulted in corresponding rotation of the accessory drive internal spline (normally mating to the AGB radial drive shaft).

The second stage turbine rotor was removed, and its blades exhibited no anomalous damage or erosion. All blades were present, and both sides of the blades exhibited evidence of soot deposits. The first stage turbine rotor was removed, and its blades exhibited no anomalous damage or erosion. All blades were present, and both sides of the blades exhibited evidence of soot deposits with heavier soot deposits observed on the blades around the 12 o'clock position. Separation of the compressor case halves revealed all 10 stages of the compressor rotors, after which the compressor spool (with the front frame and rear frame still attached) could be manually rotated. All compressor blade surfaces exhibited soot deposits. The compressor blades did not exhibit evidence of damage consistent with hard or soft body foreign object debris ingestion or severe erosion. The six compressor stator stages exhibited no evidence of anomalous damage. Soot deposits were observed on the inlet guide vanes (IGV), all three stages of variable guide vanes (VGV), and the six stages of stators. From the 3 o'clock to the 6 o'clock region, the VGV tips were scored in the area between the second, third, and fourth stages. Impressions of the stator vanes were observed on the soot deposits on the compressor spool surfaces. Very light impressions of the last two stages of the VGV were observed on the soot deposits on top of the spool surfaces but impressions of the first stage of VGV and IGV were not observed on the soot deposits on the compressor spool surfaces. The power turbine blades did not exhibit evidence of anomalous damage or severe erosion; the blades exhibited discoloration, and the surfaces exhibited a rough appearance in areas that were not coated with soot.

The fuel pump cover was removed, revealing a thermally degraded gasket underneath the cover. The front coupling external and internal splines exhibited no evidence of wear. The internal splines exhibited a rust-colored residue on a portion of its outer diameter. The front (muff) coupling adapter splines exhibited no anomalous damage. Removal of the front cover revealed white-colored deposits consistent with corrosion; however, all components were subject to postcrash firefighting efforts and then stored in an open hangar before examination. A thermally degraded gasket was observed underneath the front cover. Two small, metallic globules were found underneath the front cover. A plug covering a port normally used for testing was removed, and a thermally degraded O-ring was found underneath the head of the plug.

The rear cover was removed, revealing corrosion and thermal discoloration within the interior surfaces of the pump. Red-colored deposits were seen on the boost drive gear outer surfaces and about half of the gear web surfaces. The three shear posts were intact. The internal splines of the rear coupling, connected to the boost drive gear, exhibited no anomalous damage. The rear coupling was removed, and the retaining ring was still in its installed position. The driveshaft splines mating to the rear coupling exhibited no anomalous damage. Red-colored deposits were observed on the contact surface of the boost driven gear teeth, but none were observed on the non-contact side of the gear teeth. The booster driven gear bearing support contained white-colored deposits consistent with corrosion. The pump drive gear and driven gear teeth contained re

NTSB Probable Cause

A dual loss of engine power for undetermined reasons after the pilot's improper decision to attempt another maneuver after recovering from a perceived compressor stall, rather than returning to the airport.