ERA24LA205

The pilot’s inadequate yaw control of the helicopter while maneuvering at a low altitude and airspeed, which resulted in a hard landing.

During the enroute portion of the flight, the pilot in the left seat of the helicopter requested a transfer of control from the pilot in the right seat so that he could gain proficiency in preparation for obtaining a flight instructor certificate. After a positive transfer of control, the pilot in the left seat successfully conducted a confined-area approach to a 15-ft hover, then departed to search for an area to conduct a second approach. When the helicopter was about 50 ft above ground level (agl) and at an airspeed of about 20 kts, its forward airspeed slowed, and it entered a left turn that the pilot’s control inputs were ineffective in arresting. The turn progressed into a full rotation to the left and the helicopter completed two or three full rotations before impacting the ground, separating the right skid and tail rotor and damaging the tailboom and lower vertical fin. Postaccident examination of the airframe revealed that all damage features observed on the tailboom, aft ring frame, and tail rotor gearbox were consistent with overload structural failures sustained during the hard landing. The separated tail rotor blade did not show significant fragmentation, consistent with unpowered separation of the blade, likely due to impact with the ground after the tail rotor gearbox separated from the airframe. The tail rotor control system was continuous to the point where the tail rotor gearbox had separated from the airframe. One of six attachment bolts for the No. 1 tail rotor driveshaft (TRDS) hanger bearing support was not present in its installed location. The bolt was found lying in the tray area of the TRDS cover, and the respective nut was not found. The five remaining attachment bolts for the No. 1 TRDS hanger bearing support were found loose, consistent with having been inadequately torqued when installed during maintenance activities performed about 1 month before the accident. However, the tail rotor driveshafts showed no evidence of anomalous wear or preimpact fracture; thus, it is unlikely that the improper maintenance of the No. 1 TRDS hanger bearing support attachment hardware was a factor in this accident. No evidence of a pre-accident mechanical malfunction of the helicopter, control systems, or components that would have precluded normal operation was identified. Although the pilot stated that his control inputs were ineffective to counter the helicopter’s uncommanded left turn, data recovered from the engine data recorder (EDR) showed that the tail rotor potentiometer (XPA) value peaked only at 58.82% (consistent with a pedal position value of only 74.5%) and for only 1 second. This indicated that the pilot did not push the right pedal to its fullest extent and did not push it long enough for the pedal input to effectively counter the uncommanded left turn. According to a safety notice published by the helicopter manufacturer, all types of single-rotor helicopters can be susceptible to unanticipated left yaw at low speed. The notice indicated that unanticipated yaw has historically been referred to as “loss of tail rotor effectiveness,” but noted that the term “wrongly implies that tail rotor efficiency is reduced.” The notice stated that, when an unanticipated left yaw situation is encountered, “swift corrective action is needed in response, otherwise loss of control and possible accident may result.” According to the notice, “the key feature of an unanticipated left yaw recovery is large amplitude right pedal input. Recovery may not be immediate but will occur if the pilot persists in maintaining right pedal.” The helicopter was equipped with a cockpit image recorder that could have provided additional data to better understand the sequence of events leading up to the accident. However, it did not record the accident flight or any other flights in the previous 90 days and would not record when examined after the accident. The required annual functional check of the system, which maintenance records indicated was performed about 1 month before the accident, should have detected and rectified the malfunction before returning the helicopter to service.

On May 1, 2024, about 1600 eastern daylight time, an Airbus Helicopters AS350B3e helicopter, N413JM, was substantially damaged when it was involved in an accident near Plant City, Florida. The commercial pilot and the private pilot were not injured. The helicopter was operated by the Hillsborough County Sheriff’s Office (HCSO) as a public aircraft. The private pilot in the right seat departed the helicopter from Lakeland Linder International Airport (LAL), Lakeland, Florida, about 1545 en route to the HCSO base at Tampa Executive Airport (VDF), Tampa, Florida. During the flight, the commercial pilot in the left seat requested a transfer of control so that he could gain proficiency in preparation for obtaining a flight instructor certificate. After a positive transfer of control, the commercial pilot successfully conducted a confined-area approach to a 15-ft hover, then departed to search for an area to conduct a second approach. When the helicopter was about 50 ft agl at an airspeed of about 20 kts, its forward airspeed slowed, and it entered a left turn that the pilot’s control inputs were ineffective in arresting. The turn progressed into a left spin, and the helicopter completed two or three full rotations before impacting the ground. The commercial pilot stated that he kept the helicopter level on impact, the main rotor did not contact the ground, and he subsequently placed the engine into an idle state before he shut it down. On-scene photographs provided by the HCSO showed the helicopter upright and tilted to the right in a field. The tailboom skin showed wrinkling near where the tailboom attached to the fuselage, the lower vertical fin showed deformation, and the tail rotor gearbox was separated and on the ground near its normally installed location. One tail rotor blade (designated the “blue” blade) remained attached and exhibited no fragmentation, and the other tail rotor blade (designated the “red” blade) was found separated on the ground about 40 ft from the helicopter. One HCSO-provided photograph showed that the No. 1 TRDS hanger bearing support (near the trailing edge of the horizontal stabilizer) was partially disconnected from its support mount; the left attachment bolt and nut were not in their installed location, and the left attachment bolt was lying within the tray area of the TRDS cover. HCSO personnel reported that the right attachment bolt and nut remained installed but “finger tight” with “little to no threads” visible beyond the nut and that the attaching hardware for the four remaining TRDS hanger bearing support mounts also had low torque. The six attachment bolts and nuts for the tail rotor control rod support remained installed on the No. 1 TRDS bearing support mount. Examination of the helicopter revealed that the right skid had separated from the landing gear at the forward and aft cross-tubes. The aft ring frame for the vertical fin and tail rotor gearbox was found partially separated from the tailboom and rotated to the right about 45°. The aft ring frame exhibited overload signatures on the forward-upper portion (where the tail rotor gearbox mount attaches), and the rivets that attach the ring frame to the tailboom skin were found sheared with no evidence of working loose. The upper vertical fin did not exhibit anomalous damage. The deformation of the lower vertical fin was consistent with ground impact damage. The tail gearbox mounting bolts remained installed on the gearbox but were found separated from the aft ring frame. Continuity of drive to the main rotor was confirmed when rotating the TRDS. The fracture features of the separated tail rotor gearbox were consistent with overload and smearing during impact. The tail rotor gearbox output shaft and pitch change spider exhibited smooth rotation with no evidence of binding when rotated manually. Continuity of control from the cockpit pedals to the fractured aft end of the tail rotor control rod was confirmed. Rotational scoring was present on the fractured aft end of the tail rotor control rod, consistent with contact with the aft-most flexible coupling at the tail rotor gearbox input flange. The pedal position in the cockpit showed the left pedal slightly forward of neutral. The “blue” tail rotor blade pitch change link was fractured near its lower connection to the pitch change spider, but the attaching hardware on both ends remained installed. The “red” tail rotor blade pitch change link was intact and remained attached at both ends, but the pitch change horn was separated from the “red” tail rotor blade. The pitch change spider could be moved longitudinally along the output shaft and did not exhibit evidence of binding. The “blue” tail rotor blade had coning deformation to the helicopter-left direction. The “red” tail rotor blade had a fractured spar, its outboard tab was deformed inward, and its inboard skin had a V-shaped crack. The helicopter was equipped with an EDR capable of recording data for various engine and flight control parameters, including engine control selector position and tail rotor potentiometer (XPA). The electronic engine control unit used the XPA parameter to anticipate the need for additional engine power when the pedal position value was increased above 68.3%. For pedal position values, 0% represented full left pedal forward, 50% represented neutral pedals, and 100% represented full right pedal forward. The recovered EDR data, which captured about 21 minutes 39 seconds of the accident flight, from when the EDR was powered on until after engine shutdown after the accident, showed no record of faults. The data showed that, about 20 minutes 51 seconds from the start of the recorded data, a peak XPA value of about 58.82%, representing an equivalent pedal position value of about 74.5% (right pedal application) was recorded for about 1 second. About 14 seconds later, or about 21 minutes 5 seconds from the start of the recorded data, the engine control selector was moved from flight to idle, then about 9 seconds later, the engine control selector was moved from idle to stop. The helicopter was equipped with an Appareo Vision 1000 cockpit image recorder. Recovery of data from the unit’s internal memory and its SD card revealed that the last recorded file on each was dated February 4, 2021. A functional test of the unit revealed that the status indicator light remained solid blue and did not change past this status light indicator state. A solid blue status indicator light occurred when an embedded software update file was being processed. A solid green status indicator light would occur when the unit was receiving power and recording data. A review of helicopter maintenance logs revealed that, on August 17, 2023, at an airframe total time (ATT) of 1970.2 hours, numerous maintenance items were completed, and a work order contained five separate entries for replacement of the five TRDS hanger bearings. A logbook entry dated April 4, 2024, at an ATT of 1,970.2 hours, documented the completion of numerous maintenance items, including airworthiness directives, scheduled inspections, the replacement of the TRDS hanger bearings, and a functional check of the Appareo Vision 1000 cockpit image recorder. Per the 12-month inspection checklist contained in the AS350 Master Servicing Manual, the functional check of the Appareo Vision 1000 must be performed in accordance with the AS350 Aircraft Maintenance Manual, the procedures for which specify powering up the unit, allowing it to record imagery, then verifying the recording by removing the SD and accessing its continents using playback software. A logbook entry dated April 5, 2024, at an ATT of 1,972.2 hours, documented the completion of a main rotor track and balance and a tail rotor balance with no noted anomaly. The most recent logbook entry was dated April 30, 2024, at an ATT of 1,982.2 hours, documented the painting of the helicopt