N911FKAgustawestland Philadelphia AW139 · 2025

On December 15, 2017, about 0650 eastern standard time, a Sikorsky S-76A, N911FK, was substantially damaged when it was involved in an accident near Islamorada, Florida. The pilot, flight nurse, and paramedic were not injured. The helicopter was operated as a Title 14 Code of Federal Regulations Part 135 on-demand air ambulance flight.

According to the operator, the flight departed Florida Keys Marathon International Airport (MTH), Marathon, Florida, about 0637 for an air medical trauma patient pickup; the pilot landed at a presurveyed landing zone about 0650 to pick up the patient. The pilot reported that, after landing, he positioned the engine controls to idle and the main rotor speed (Nr) was less than 60% before he applied the rotor brake to stop the rotors. After the rotors stopped, the paramedic and flight nurse exited the helicopter to attend to the patient.

The pilot indicated that after the medical crew departed the helicopter, he noticed that the No. 1 engine temperature was fluctuating with an increase in the inlet turbine temperature, so he shut down the No. 1 engine. The operator reported that, about the same time, the flight nurse noticed dark smoke coming out of the main rotor gearbox cowling area of the helicopter. The flight nurse reported that the smoke was accompanied by sparks, which shortly after turned into flames; she and the paramedic ran toward the helicopter and signaled the pilot about the fire.

The pilot reported that he noticed the flight nurse waving her arms and warning of a fire but did not observe any cockpit indications of a fire; however, he shut down the No. 2 engine and observed flames when he partially exited the helicopter. He discharged both engine fire bottles and then exited the helicopter. He stated that the fire continued to burn until fire department personnel extinguished it.

The pilot additionally reported that the area around the rotor brake assembly generally accumulated a lot of slung grease and oil.

The helicopter was equipped with two Safran Helicopter Engines Arriel 1S1 turboshaft engines (in accordance with supplemental type certificate No. SH568NE) driving a four-blade main rotor system. The No. 1 engine had accumulated about 10,412.5 hours total time since new (TSN) and 383 hours total time since overhaul. The No. 2 engine had accumulated about 5,101 hours TSN and 228.3 hours total time since overhaul. A manually activated rotor brake system was installed in the helicopter.

The helicopter remained largely intact as the fire was in the area immediately adjacent to the rotor hub assembly and main transmission compartment; the helicopter sustained thermal damage to the main gearbox compartment and engine inlets. Further postaccident examination of the helicopter revealed heavy soot deposits in the engine's transmission compartment and debris on the floor of the compartment consistent with thermal damage deposits from the surrounding compartment. There was soot on the rotor head assembly to the pitch change links, dampers, and pitch horns. The upper surfaces of the engine cowling exhibited two areas of heavy soot deposits aligned between the Nos. 1 and 2 oil cooler ducts and the main gearbox oil cooler duct. The left and right intake fairing, engine cowling, internal ducting for both engine oil coolers, and rotor brake cowl exhibited thermal damage.

The airframe oil supply lines from the oil tank were intact for both engines. The airframe oil return lines to the oil tanks for both engines were thermally damaged and fractured. The No. 1 engine oil filler cap was removed, and there was no visible evidence of oil near the filler. The No. 2 engine oil filler cap was removed, and there was visible evidence of an oil level near the filler.

No. 1 Engine Examination and Disassembly

The exterior of the No. 1 engine did not exhibit evidence of thermal damage. The compressor turned freely when manually rotated. The transmission input shaft turned freely in the freewheeling direction when manually rotated, and turning the transmission input shaft in the driving direction resulted in rotation of the main rotor head. A check of the rigging between the No. 1 engine control lever (ECL) in the cockpit and the fuel control unit (FCU) for the No. 1 engine revealed no anomalies and indicated an FCU throttle position of about 27° when the ECL was set to idle.

The No. 1 first-stage compressor did not exhibit evidence of foreign object debris (FOD) damage. The free (power) turbine exhibited no evidence of blade shedding. A borescope inspection of the first- and second-stage gas generator turbines revealed no anomalies or damage.

Disassembly of the No. 1 engine revealed no evidence of cracks or fractures to the external lines removed from the engine or damage noted to the thermocouples, power transmission shaft, or reduction gearbox. All power turbine blades were present and free of damage, and the power turbine rotated freely with no binding. The combustion liner had no damage or debris. The first-stage (axial) compressor exhibited black coloration but did not exhibit anomalous damage. The second-stage (centrifugal) compressor had an even coating of soot. The first- and second-stage nozzle guide vanes exhibited no evidence of anomalous damage. Soot deposits were observed on the first-stage nozzle guide vane surfaces, and the second-stage nozzle guide vane surfaces exhibited bluing. Continuity was confirmed between the accessory drive and the starter-generator pad. The drained oil was tan in color.

The No. 1 FCU was removed and bench tested. The FCU was brought from normal start to flight idle with no anomalies detected.

No. 2 Engine Examination and Disassembly

The exterior of the No. 2 engine did not exhibit evidence of thermal damage. The compressor and the transmission input shaft could not be manually rotated in either the freewheeling or the driving direction. The exterior surfaces of the transmission input shaft exhibited a matted appearance consistent with exposure to fire. A check of the rigging between the No. 2 ECL in the cockpit and the FCU for the No. 2 engine revealed no anomalies and indicated an FCU throttle position of about 27° when the ECL was set to idle.

The No. 2 first-stage compressor did not exhibit evidence of FOD damage. The power turbine exhibited no evidence of blade shedding. A borescope inspection of the first- and second-stage gas generator turbines revealed damage to the turbine blades.

Disassembly of the No. 2 engine revealed that the power transmission shaft and the gas generator could not be manually rotated. All three thermocouples exhibited evidence of thermal damage. The reduction gearbox was removed, and the gearbox and power transmission shaft rotated freely with no evidence of binding, but the power turbine could not be rotated. The reduction gearbox splined nut did not show misalignment. The overspeed sensor and muff coupling showed no anomalous damage. The power turbine was removed and could be manually rotated, during which the turbine bearings made a chattering sound. The turbine inlet stator vanes had a dull, gray coating flaking off their surfaces. The axial compressor could be manually rotated, but its movement was not smooth. Soot deposits were on the axial compressor and centrifugal compressor surfaces. Rub marks were on the interior of the centrifugal compressor cover and the centrifugal compressor blades matching the rub mark location on the cover. The tips were missing on the second-stage gas generator turbine blades, and the second-stage nozzle guide vanes exhibited partial melting. The first-stage gas generator turbine blades exhibited missing surface coating near the blade tip ends. Continuity of drive was confirmed between the accessory drive and the starter-generator pad. The drained oil was tan in color.

The No. 2 FCU was removed and bench tested. The FCU was brought from normal start to flight idle, and the Np rpm, fuel flow, and gas generator speed (Ng) were normal in both the governed and nongoverned range; no anomalies were detected.

Rotor Brake Assembly

Raised material was observed on the rotor brake disk surface and had the appearance of brake pad (puck) material fused to the rotor brake disk surface. Soot deposits were observed on the exterior of the No. 1 tail rotor drive shaft. A hydraulic fluid level could not be seen through the sight gauge for the rotor brake accumulator. Hydraulic pressure was applied to both rotor brake calipers using a manual handheld hydraulic pump. When about 200 pounds per square inch (psi) was applied, the pistons extended, and no leaks were observed.

The left and right brake caliper assemblies were examined further and tested with hydraulic pressure. Tests were performed after bleeding air from the system, then increasing hydraulic pressure to 230 psi. The forward and aft pucks (both left and right brake calipers) extended as pressure was applied. When pressure was reduced, the forward pucks did not automatically retract but were manually retracted without difficulty. A slight hydraulic leak existed at the forward seal subassembly of the right brake caliper but did not affect the action of the brake pucks.

The caliper was disassembled. The O-ring was in good condition, and there was no evidence of damage or irregular tolerances between the piston and bore. The spring and pin were disassembled from the forward and aft puck assemblies and revealed no anomalous damage.

The accumulator and master cylinder were examined and bench tested with the application of hydraulic pressure and revealed no anomalies.

The S-76A++ rotorcraft flight manual (RFM) No. SA 4047-76-1 in the accident helicopter contained revision 5 of rotorcraft flight manual supplement (RFMS) No. 29B that was applicable to helicopters equipped with the Arriel 1S1 engines installed in in the accident helicopter. The RFMS contained rotor brake limitations for when the rotor was either stopped or turning. The rotor turning limitation stated, in part, “Rotor brake application limited to one (or two) engine(s) operating at idle or both engines shut down” and “Maximum rotor speed for normal rotor brake application is 65% Nr.”

A decal on the rotor brake lever of the accident helicopter contained the following caution in red lettering: “NORMAL ROTOR BRAKE STOPS ARE AUTHORIZED ONLY AT 65% NR OR LESS WITH BOTH ENGINES SHUT OFF OR ONE ENGINE AT IDLE.”

The RFM, Part 1, Section III, Emergency Procedures, contained the following statement about the possibility of a rotor brake fire: “The ROTOR BRAKE caution light indicates that pressure has been applied to the rotor brake system. If this pressure results in a rotor brake puck dragging against brake disk, a rotor brake fire might occur.”

According to the Trauma Star Part 135 General Operations Manual, Section S, Helicopter Emergency Medical Services (HEMS), S-26; Ground Operations at Off Airport Landing Sites,

A. Once the aircraft has landed, the pilot will apply the rotor brake and stop the aircraft rotor system. The pilot will then advise the medical crew when it is safe to exit the aircraft. Trained medical crew will supervise the preparation of the patient for loading into the aircraft, including ensuring the patient is adequately secured to the litter, removal of any unsecured sheets or blankets from the litter assembly, etc. Ambulances may bring the patient to the aircraft but will not approach within 30 feet of the aircraft or rotor system.