History of Flight
On July 19, 2023, about 0830 mountain daylight time, an Aerospatiale AS-355E helicopter, registration N102UM, was substantially damaged during a forced landing near Farmington, Utah. The pilot and four passengers were not injured. The flight was conducted under Title 14 Code of Federal Regulations Part 135 as an on-demand air taxi operation.
The helicopter had been dispatched to transport a vegetation removal crew to power distribution lines in the Francis Peak area of the Wasatch Mountain Range. The pilot departed from the helicopter's base at Ogden-Hinckley Airport (OGD) about 0800 and flew directly to a staging area in the foothills. After landing, the pilot briefed the crew and the flight departed for the work site.
During the climb-out to the work area, the pilot observed that main rotor speed was beginning to decay. As he maneuvered for landing, he noticed that engine speed and temperature readings were not properly matched between the two engines, with the No. 2 engine showing an 8-10% reduction in torque. The pilot performed a series of engine adjustments using the trim system, maneuvered away from nearby power lines, and increased the helicopter's speed; main rotor speed then returned to normal.
He began the landing approach again. However, when the helicopter descended to within about 20 to 30 ft of the landing zone (elevation 8,600 ft), main rotor speed again began to decay. The pilot turned the helicopter away from the landing zone to maintain clearance from the mountain. The helicopter began to descend, and the main rotor blades started cutting through surrounding trees. The helicopter landed on its belly and rolled over. The pilot and crew egressed through the left door. The engines were still operating after the accident, and the pilot shut them down and secured the electrical system.
Aircraft Information
The helicopter was manufactured in 1981 and powered by two Rolls-Royce/Allison 250 C20F turboshaft engines mounted at the top of the fuselage to the rear of the main transmission gearbox.
The engines were controlled mechanically through a cable-driven collective anticipator system. An electro-mechanical “engine trim” assembly, operated by a rocker switch on the collective lever grip, allowed the pilot to adjust engine power balance. The system comprised mixing bellcranks on the main transmission deck and allowed differential adjustment—if the pilot commanded an increase in power on one engine, it would reduce power on the other.
The flight manual stated that under normal flight conditions, the trim system is used to synchronize engines during takeoff, climb, and hover, and for power checks. No specific instructions existed for use during a loss of engine power, but the emergency procedures section noted: “The procedures outlined in this section deal with the common types of emergencies; however, the actions taken in each actual emergency must relate to the complete situation.”
The emergency procedures and height-velocity diagram indicated the helicopter could be flown safely after a single-engine failure during takeoff at altitudes up to 7,000 ft and weight at or below 4,740 lbs. The helicopter's maximum permissible weight (ambient-dependent) was 5,291 lbs; at the time of the accident, weight was 4,963 lbs.
Maintenance
The compressor section of engine No. 2 had been replaced 12.1 flight hours before the accident. Both the pilot and maintenance personnel reported difficulty synchronizing the engines after the replacement. The problems were intermittent; troubleshooting included trim adjustments and test flights, eventually matching the engines within manufacturer’s power specifications. During troubleshooting, the pilot noted the bleed valve was approaching overhaul time but decided to leave it in place because the other engine operated correctly.
Following the compressor change, the engine emitted a loud “howling” sound. The sound was not audible inside the cabin due to ambient noise but was pronounced from outside during ground and flight operations. It did not appear to affect engine performance. The mechanic attributed the sound to the newer overhauled compressor blades. The helicopter continued flying multiple missions with the sound present.
Postaccident Examination
Postaccident examination of the engine trim system indicated it was functioning appropriately. Both engines were removed and examined at a Rolls-Royce-approved overhaul facility under NTSB oversight.
Engine No. 1 was intact with minor foreign object damage to first stage compressor blades but no catastrophic failure. All components, including fuel control, fuel pump, nozzle, power turbine governor, and compressor bleed valve, were within serviceable limits.
Engine No. 2 was similarly intact with similar foreign object damage. The core was disassembled; fuel control, pump, nozzle, and governor tested within limits. Examination of the compressor section revealed a prominent groove in the inner liner at the No. 5 and 6 blade tip area on one case half, with shallower rub signatures on the other. Stator vanes showed tip smearing opposite rotation direction, with corresponding rub marks.
During testing, the compressor bleed valve intermittently hung up when transitioning from open to closed. It did not meet interstage pressure reduction specifications, but the diaphragm leakage test was normal. The valve was disassembled; a 1/16th-inch hole was found in the rolling section of the diaphragm. The diaphragm was an FAA PMA component (6874725HT) manufactured by HYE-Tech LLC, with a service life of 1,500 hours; records showed 76.6 hours remained at the time of the accident.
The bleed valve assembly was sent to the NTSB Materials Laboratory. Fourier-transform infrared spectroscopy indicated the diaphragm material composition was consistent with Rolls-Royce specifications. Scanning electron microscopy revealed elongated fiber ends, knife-edge separation areas, and severe mechanical damage consistent with fracture faces contacting each other. Non-damaged fiber faces showed shear or cup-and-cone features typical of overstress separation. The elastomer showed no fatigue cracking.
The compressor assembly was examined at Rolls-Royce facilities. Rub was confirmed on stage 4-6 axial compressor blade tips and associated case tracks, consistent with rotor imbalance. Measured imbalance was 5.1 times the allowable limit forward and 8.4 times at the aft end. Metal-to-metal contact was consistent with imbalance causing rotor orbit. Circumferential rub on the compressor impeller and uneven wear on the abradable shroud were observed; build clearance between impeller and shroud was 55% larger than maximum specification, indicating reduced compressor efficiency.