N20BH – Bell Textron 407 – 2024-08-03 NTSB Accident Report

HISTORY OF FLIGHTOn August 2, 2024, about 2107 mountain daylight time, a Bell 407 helicopter, N20BH, was substantially damaged when it was involved in an accident near Raft River, Idaho. The pilot and three passengers sustained minor injuries. The helicopter was operated as a public aircraft in support of the USFS.

The helicopter was privately owned, and the flight originated under the Bureau of Land Management’s use of a USFS Exclusive Use flight service contract.

The purpose of the flight was to transport three USFS helitack crew members from Gerlach, Nevada, to their base at Pocatello, Idaho, with intermediate stops at Winnemucca and Elko, Nevada. The pilot reported that the first two legs were uneventful, and that the accident occurred on the final leg.

On the day of the accident, sunset at Pocatello was 2049. The terms of the contract limited operations to day VFR only. The crew planned to arrive at 2120; however, according to the pilot, they were given permission to arrive as late as 2125, which was possible as long as he maintained a groundspeed of 120 knots while enroute.

The pilot stated that, about 55 minutes after departing Elko, the barrier inlet air filter light started to flicker. The filter had been serviced the week prior, and although the helicopter had been flying in a dusty and smoky environment since then, he thought such an activation was premature. A short time later, the filter light transitioned from flickering to on, and the pilot activated the filter bypass system by pressing the “FILTER” light switch. Due to arrival time restraints, the crew then discussed the option of diverting and completing the flight the following morning but chose to continue to Pocatello.

The pilot reported that, a few minutes later, he noticed the engine MGT starting to rise into the yellow band of the gauge. In response, he reduced power and began to descend. However, the MGT continued to rise, and at an altitude between 1,000 and 1,500 ft, it continued to climb past the gauge redline. The pilot decided to make a precautionary power-on landing and began looking for a suitable landing area. As he was maneuvering the helicopter toward a grass field, he determined that the engine was starting to overspeed.

The pilot raised the collective control twice to try to arrest the overspeed, but the engine did not respond, so he decided to perform an autorotation into an adjacent corn field. Due to the corn's unknown height, the pilot was unable to determine the optimal time to initiate the flare, and the helicopter landed hard after dropping from about 5 ft above the ground. The pilot reported that immediately after landing, the engine came back on for about 10 seconds and was shaking the aircraft and crew violently. PERSONNEL INFORMATIONThe pilot reported 3,855 total hours of flight experience, including 69 hours in the 90 days preceding the accident, and 26.8 hours of flight time during the previous 7 days. He had accumulated 908 hours in the Bell 407 and was issued an interagency pilot qualification card by the USFS on June 10, 2024.

On the day of the accident, the pilot was dispatched from the Pacific time zone, where his duty day began at 0700. Crew duty days are limited to 14 hours; a 14-hour duty day starting at 0700 PDT would have concluded by 2200 MDT. WRECKAGE AND IMPACT INFORMATIONThe helicopter came to rest in a cornfield about 41 miles southwest of Pocatello. At the accident site, the surrounding corn stalks were about 6 ft tall. The helicopter was in a level attitude, with the main skids splayed outward such that the cabin belly was resting on the ground.

The main rotor transmission assembly detached from the airframe and had rotated left and dropped down into the cabin compartment. The engine, which sat behind the main transmission, remained attached to the airframe. The engine output drive shaft and associated couplings had failed forward of the engine, resulting in fragments of the surrounding structure and drive components being propelled up into the inlet plenum.

Three of the four main rotor blades remained attached at the hub, and the fourth was partially attached and bent opposite the direction of rotation. Two blades were shattered about midspan, with their remaining fragments located within the immediate vicinity of the accident site.

The tail rotor and associated gearbox had detached and were located about 30 ft from the main wreckage. FLIGHT RECORDERSThe helicopter was equipped with an Appareo Vision 1000 cockpit image recording device mounted to the cabin roof. The unit contained an internal camera and a GPS receiver that captured time, position, altitude, and speed. The unit also had a self-contained real-time inertial measuring unit capable of recording 3-axis acceleration and derived pitch, roll, and yaw data.

The unit was sent to the NTSB Vehicle Recorders Division for data extraction and a video group reviewed the data. Review of the video showed that the pilot was seated in the right seat, with a crew member in the left seat. The recording began about 7 minutes before takeoff after an uneventful engine start. The helicopter departed at 1947:41. Engine gauge indications for the next 10 minutes as the helicopter flew northeast were within the normal operating parameters, with MGT remaining just below the 727° C maximum continuous operating range.

At 2005:47, the MGT gauge began to show an increase in temperature into the takeoff range and the display began to flash, accompanied a short time later by a brief illumination of the “CHECK INSTR” light on the annunciator panel. MGT then briefly dropped back to 727° C, after which it began to climb. About 30 seconds later it had reached 736° C. The CHECK INSTR light then came back on and remained illuminated for the rest of the flight. The pilot then pointed to the MGT gauge, as it dropped down to 710° C.

The flight proceeded while making heading and altitude changes to avoid terrain, and after 30 minutes, the barrier filter alert could be seen reflected in the passenger’s helmet. The pilot reached for the overhead panel, and at 2102:29, he turned on the instrument panel lights as the sky grew darker.

At 2106:16, the MGT radial and numeric indicator began to display missing and random segments, and the temperature began to climb. (see figure 1.) The HYD SYSTEM light briefly illuminated for 1 second, and there was no corresponding change in the other engine parameters. About 20 seconds later, MGT had reached 840° C, and the vertical speed indicator was showing a 1,200 ft per minute (fpm) descent rate. The airspeed slowed to 115 kts, torque dropped to 43.6%, and gas generator speed (NG) dropped to 91.7%, while power turbine speed (NP) and main rotor speed (NR) remained constant at 100%.

The MGT gauge then appeared to perform a built-in-test, or BIT check, where all segments of the display illuminated. (see figure 1.) By the time the unit had come back online, it was showing a temperature of just below 500° C on both the segmented display and radial indicator.

Figure 1 – MGT rising with missing segments (left). BIT check (right)

By this time, the helicopter had descended to 5,000 ft mean sea level (780 ft above ground level [agl]) and was descending at a rate of 750 fpm, at an airspeed of 97 kts. The torque gauge indicated 6.4%, and both NP and NR were still at 100%.

About 5 seconds later, at 2105:53 with NP remaining at 100%, NR increased to 103% and both the ambient noise pitch and volume increased consistent with increasing rotor rpm, followed by the illumination of the rpm light. At this time, NR was 107% and NP was 100%, with MGT indicating 475° C.

By the time the helicopter had descended to about 120 ft agl, the rotor rpm light illuminated accompanied by the aural low rotor rpm tone. NR was at 94% as the helicopter began to pitch up in a manner consistent with it beginning a flare, and a few seconds later the helicopter violently impacted the ground. TESTS AND RESEARCHMGT Indicator Gauge

The helicopter’s MGT gauge screen was composed of a segmented radial temperature indicator with color-coded bands, along with a series of seven-segment displays that indicated the temperature numerically along with the exceedance status. (see figure 2.) On power-up or under conditions where the unit sensed a software or hardware anomaly, it was designed to perform a BIT, which included an illumination of all display segments.

Figure 2 - MGT Gauge

The gauge, in correspondence with the flight manual’s engine operating guidance, indicated a continuous operating temperature of 100 to 727° C with a green band.

The 5-minute takeoff range temperature was between 727° and 779° C and denoted by a yellow band. The maximum takeoff temperature was 779° C and indicated by a red bar.

The transient, 12-second range was between 780° and 843° C with the upper limit bound with a red triangle, and the maximum start and shutdown range, which was not to exceed 10 seconds above 843° C or 1 second at 927° C, was indicated by a red circle.

The design of the unit was such that if the temperature was in the yellow takeoff region (727° to 779° C) for 4 ½ minutes the segments of the display would flash as a warning. If this temperature condition continued for an additional 30 seconds, the numeric display would show an “E” indicating that an exceedance had been logged in non-volatile memory (NVM). An amber CHECK INSTR caution light would also illuminate on the annunciator panel if an MGT exceedance was about to or had occurred. An exceedance would also be logged into if the temperature climbed into the transient range (780° to 843° C) for more than 10 seconds, or if it ever entered the transient upper limit.

Once the exceedance flag was triggered, the “E” indication remained on the screen until a mechanic reset it using specialized software.

MGT Indicator Gauge Examination

The unit was examined and tested at an FAA-certified overhaul facility, in accordance with an approved test protocol, and no anomalies were noted. It was found that by lowering the input supply voltage to around 6 volts direct current (VDC), the unit would exhibit the increasing MGT behavior observed during the accident sequence. Once it had dropped below 6VDC, the unit would display random characters in a manner almost identical to that observed before the accident. With the restoration of supply power to the nominal 28VDC, the unit restarted and performed a BIT check, consistent with that observed during the accident flight. (see figure 3)

Figure 3 – Increasing MGT (left), increasing with random segments (center), BIT check (right).

The unit’s memory was accessed using the standard maintenance software, and one exceedance had been recorded (the unit can record 50 exceedances).

This exceedance was for 737° C and lasted 5 seconds. It was recorded on the day of the accident and appeared to be consistent with when the CHECK INSTR light came on and stayed on after the overtemperature indication during the initial cruise phase of the flight.

MGT Gauge Airframe Installation

The MGT indication system comprised the indicator mounted to the center instrument panel, which was supplied with electrical power via a 3-amp circuit breaker mounted to the overhead panel. The MGT signal was captured by 3 thermocouples connected in parallel and mounted to the engine exhaust section. The thermocouples provided data to both the MGT instrument and the ECU.

On the airframe, the MGT circuit breaker was found in the closed (pressed) position. The circuit breaker performed appropriately when tested, and the entire electrical supply and ground wiring for the MGT system were examined and tested. No short or open circuits were observed, all cannon plugs and connector fittings were intact and free of corrosion, and there was no evidence of electrical arcing, thermal damage, or exposed conductors through cable chafing.

Engine Examination

Examination of the Rolls-Royce M250-C47B engine revealed evidence of hard body ingestion and engine uncontainment. All impeller inducer blades exhibited impact damage, tears/rips, missing material, and distortion, while the exducer blades exhibited shiny contact wear consistent with contact with the compressor shroud housing. The engine exhaust duct, exhaust collector support, horizontal firewall shield assembly, and right compressor discharge tube exhibited varying degrees of pockmarks and exit holes. Looking through the exit hole in the exhaust collector support, the stage 4 turbine wheel was observed, with all blades fractured transversely at their roots.

The engine was examined and disassembled at the facilities of Rolls-Royce under the supervision of engineers from the NTSB.

Examination revealed continuity between the compressor and accessory gearbox, but not with the gas producer turbine.

The tie bolt, which secures the gas producer stages 1 and 2 wheels together, was found fractured, as was the turbine-to-compressor coupling shaft that connects the gas producer turbine to the accessory gearbox. The No. 6 roller bearing inner race and power turbine coupling shaft, which connects to the power turbine-to-pinion gear shaft, was also fractured. The turbine shaft-to-pinion gear coupling was fractured and no longer coupled to the inside spur teeth of the helical gear, but remained seized within the helical gear bore. All the blades on stages 1, 2, and 3 turbine wheels were present and fractured transversely across the airfoil; the stage 4 turbine wheel blades were fractured at the platform. Both the gas producer and power turbine exhibited extensive rotational damage throughout, consistent with stationary and rotating parts/components circumferentially and longitudinally contacting one another.

The gas producer turbine and power turbine components were submitted for material analysis. No parts were found with any pre-existing anomalies or fatigue fractures; all damage appeared consistent with secondary damage following loss of axial and radial control of the components during the ground impact sequence.

Engine Control Unit

The engine was equipped with a dual-channel ECU, manufactured by Triumph Engine Control Systems.

The ECU included an incident recording feature, which was intended to provide data in support of engine troubleshooting and maintenance. The system was designed to activate after a successful engine start, then continuously record data every 1.2 seconds over a 12-second loop, resulting in a stream of the last 10 sets of data. When an incident was detected, the 10 sets of data recorded pre-incident were saved, directly followed by an additional 40 sets of post-incident data. Each data set included critical parameters such as engine history, discrete words, fault words, and key control inputs.

The unit sustained impact damage during the accident sequence; however, the integrity of its recording components was established. The data extract revealed no events that would have met the threshold to trigger the incident recorder. Four undated limit peak events were recorded, none of which appeared to be related to the accident flight.