N3045R is a 2013 Thrush Aircraft S2R-H80. It is a fixed-wing single-engine aircraft with 1 seat, powered by a P&W Canada PT6A-34AG rated at 750 horsepower. The registration is active, with the registrant based in Marion, KS. Its restricted airworthiness certificate was issued in Jan 2023. It has been registered to its current owner since Feb 2017. This page also carries 1 NTSB event.
- 13 years old
- 1 of 34 Thrush Aircraft S2R-H80
2023
2017
That's everything on file so far — TailWiki checks the FAA registry daily and adds new entries as N3045R's record changes.
On May 5, 2017, about 1450 central daylight time, a Thrush Aircraft Inc. S2R-H80 airplane, N3045R, impacted terrain and a fence during a forced landing near Gypsum, Kansas, following a loss of engine power. The commercial pilot was uninjured. The airplane sustained substantial fuselage damage during the forced landing. The airplane was registered to and operated by Central Ag Air LLC as a Title 14 Code of Federal Regulations Part 137 aerial application flight. Day visual meteorological conditions prevailed in the area about the time of the accident, and the flight was not operated on a flight plan. The local flight originated from the Marion Municipal Airport (43K), near Marion, Kansas, about 1350.
The pilot reported that he departed 43K to perform an aerial application approximately 35 miles to the northwest. He sprayed two fields before reaching another field where he intended to perform the aerial application. During a pass through the field, the engine power decreased "significantly." The pilot reached over and pushed the power lever full forward and the engine did not respond. He continued to make several more passes to get more weight off the airplane and see if the engine would regain power. The engine did not get any better, so the pilot flew the airplane up out of the field and headed back to 43K. The pilot tried to climb but the airplane did not have enough power to climb. The airplane made it about 2 miles and then the smoke started coming out both exhausts and the engine "quit." The pilot observed the surrounding area, which was "big rolling hills and terraced farmland," and decided the best place for a forced landing was an alfalfa field. The airplane touched down in the corner of the alfalfa field and a pasture. The airplane slid to rest in the alfalfa field after sliding through a fence. When the airplane came to a stop, the pilot reached down, turned the master switch off, unbuckled the safety harness, opened the door, and quickly exited the airplane.
The 38-year-old pilot held a Federal Aviation Administration (FAA) commercial pilot certificate with a single and multi-engine airplane rating. He held an FAA second-class medical certificate issued on March 8, 2017, with no limitations. He reported accumulating 5,215 hours total flight time and 1,738 hours of flight time in the same make and model as the accident airplane. The pilot also held an airframe and powerplant mechanic certificate.
N3045R was a full cantilever low-wing, all metal construction monoplane with serial No. H80-140, which was designed for agricultural flying. It has a maximum published gross weight of 10,500 lbs. The airplane was powered by a dual-spool turbopropeller 800-shaft horsepower GE H80-100 engine with serial No. 133001. The engine's gas generator section features a three-stage compressor that is comprised of two axial stages and one centrifugal stage, a reverse flow annular combustor, and a single stage turbine that drives the compressor. The power turbine section, counter rotating to the compressor's turbine, features a single-stage turbine and reduction gearbox (RGB) that drives the propeller. The engine was designed to operate with fuel being supplied to it by an electric fuel pump. However, the airplane's maintenance manual indicated the engine had a 100-hour limitation on operation without an electric fuel pump.
According to the engine manufacturer, the engine build was completed on August 3, 2013, and the engine was fitted to the airplane on September 12, 2013.
A pilot reported a very sensitive throttle while reducing power after takeoff where the torque will drop from 90% to 40% in a quick change even though the throttle is moved slowly and small amount. The squawk was troubleshot as a Fuel Control Unit (FCU) issue and on October 24, 2014, the FCU was replaced. Ground and flight tests were performed, and the airplane was returned to service. The airplane accumulated 494.34 hours of total time and 593 cycles since new at that time.
A logbook endorsement dated June 12, 2015, indicated the airplane had a reported hot wire strike. The RGB bearings and FCU were replaced. The airplane accumulated 634.32 hours of total time and 967 cycles since new at that time.
On April 18, 2016, the propeller governor was replaced at 1,069.87 hours of tachometer time.
An issue with engine starting was reported and on August 25, 2016, the starting and limiting unit (SALM) was replaced.
The engine was reported to start hot and on November 7, 2016, the FCU was changed. The airplane accumulated 1,629.27 hours of total time, 2,345 takeoffs, and 531 engine starts at that time.
The pilot reported that the last inspection on the airplane was an annual inspection completed on February 20, 2017. The airplane accumulated 1,629.3 hours of total flight time at the time of the annual inspection and 1,737.6 hours of total flight time, 2,460 takeoffs, and 553 engine starts at the time of the accident.
The pilot also stated that he installed a new electric main fuel pump T1800[9]-B8 serial No. 192737 on the airplane on August 4, 2016, and that the Hobbs meter indicated 1,428.6 hours. However, he did not provide a logbook endorsement of that main electric fuel pump. The pilot confirmed that main electric fuel pump accumulated about 309 hours of operation between the installation date and the accident date.
The electric fuel pump manufacturer website describes the 18000 series pump as a slung vane, self-priming, high suction lift, and integral pressure relief valve pump with a built-in bypass valve, which is powered by a permanent magnet motor. The pump is designed for emergency fuel boost, priming injected engines, and primary fuel pump applications where aviation gasoline, jet fuels, and JP-8 is used.
The H80 airplane was fitted with the Electronics International MVP-50T engine monitoring system. The MVP-50T consists of a multifunction glass panel engine monitoring and display system to display engine parameters, along with other data and warnings. The system performs monitoring tasks only. It does not perform any engine or aircraft system controlling functions. The MVP-50T system consists of the glass panel display unit (MVP-50T), the Electronic Data Converter (EDC-33T), and the transducer fuel signal conditioner.
Each wing contains integral wing tanks (wet wing fuel tanks) just outboard of the fuselage. The left wing and right-wing fuel tanks are interconnected through a header tank. The published fuel system capacity is 228 gallons. The aircraft's fuel system is equipped with a 1/4-inch mesh finger strainer installed in the outlet fitting from the header tank. The fuel supply line to the engine is routed from the header tank, located in the fuselage, through a fuel shut off valve, an emergency electric driven fuel pump, and then directed to the main electric driven fuel pump. The fuel supply line exits the main pump and passes through a 10-micron nominal main fuel filter. The fuel line then goes through the fuel pressure sensor, which is located before the firewall. The fuel line is then passed through the forward firewall to the fuel flow meter, and then enters the engine's fuel pump.
The emergency electric driven fuel pump is a backup system to provide continuous fuel pressure in case the main electric fuel pump fails. The main fuel pump and the emergency fuel pump are not to run simultaneously.
The fuel tank vent system is designed to keep the fuel spillage to a minimum. The fuel tanks are vented through tubing connected at both the inboard and outboard ends of the individual fuel tanks to the centrally located vent system in the fuselage. Ram air enters a vent scoop, on the fuselage, under the left wing and pressurizes the vent system to maintain positive pressure on the fuel tanks. The vent system is provided with two quick drains, located on the fuselage under each wing, to drain any fuel that might happened to have migrated in the tanks outboard vent lines. Fuel quantity is displayed individually on the cockpit panel display.
The airplane's flight manual (AFM) limitation section, in part, stated:
FUEL PUMPS: Continuous simultaneous operation of both fuel pumps is prohibited due to high fuel pressure.
The AFM advised pilots that an emergency hopper dump was available. The AFM, in part, stated:
Should circumstances arise that require an emergency landing, the hopper should be dumped by moving the dump lever full forward. Forward pressure on the control stick or forward trim (or both) should be used to prevent excessive nose up pitching moment. Max speed for Hopper dump is 158 MPH.
The AFM remedial action for an 'engine failure in-flight" in part, stated:
Engine failure symptoms could include any or all of the following:
a. Loud noises followed by heavy vibration and loss of power, smoke and/or flame.
b. Rapid loss of power with unusual noises, vibration or sudden increase of ITT.
c. Loss of power following a drop in oil pressure below redline or increase in oil temperature above redline or both.
d. Loss of power following overspeed of gas generator (Ng).
e. Engine explosion and flame & smoke.
If it is clear that the engine has failed, proceed as follows:
f. Hopper Dump DUMP LEVER FORWARD
g. Propeller Lever FEATHERED
h. Fuel Condition Lever CUT OFF
i. Power Lever IDLE POSITION
j. Fuel Valve Lever OFF
k. Fuel Pump Switches OFF
l. LAND AS SOON AS POSSIBLE
The AFM additionally advises on operation with a "failure of automatic fuel scheduling." The remedial action for this failure, in part, stated:
Normally the fuel scheduling is automatic, the engine power lever is positioned by the pilot and the fuel control schedules the fuel in a manner that allows smooth increases and decreases of power while not exceeding any engine limitations during acceleration or deceleration. In the event the automatic fuel scheduling fails, the engine will experience a "low side" failure, sometimes called a "roll back". The engine goes to minimum fuel flow, which is slightly below normal Ng idle speed. Fortunately the GE H-80 is equipped with an emergency governor which will allow the pilot to regain full control of the engine and safely return. While in Emergency Governor, all automatic fuel scheduling is lost and power is controlled by the fuel condition lever. Too rapid of movement of the fuel condition lever while in Emergency Governor can cause ITT exceedences, Ng speed exceedences, and compressor stalls. It's imperative that any power changes while in Emergency Governor be made smoothly and slowly allowing the engine to accelerate normally. Full power is available while in Emergency Governor. In the event automatic fuel scheduling is lost and the engine goes to minimum fuel flow.
a. Raise the switch guard and place Emergency Governor Switch ON.
b. Smoothly advance the ENGINE FUEL CONDITION LEVER until power is restored and a climb is established.
c. LAND AS SOON AS PRACTICABLE.
d. If power is not restored, LAND AS SOON AS POSSIBLE.
After power is restored and you have climbed to a safe altitude, reduce the engine power lever to idle and continue the flight to a safe place to land using the Fuel Condition Lever as you would normally use the Engine Power Lever.
The airplane's maintenance manual, in part, stated:
This group [includes a] fuel pressure gauge. ... These readings are displayed on the MVP-50T Glass Panel Engine Monitor. ...
The S2R-H80 aircraft is equipped with a MVP-50T glass panel display that reads fuel pressure and flow rate. The fuel flow transducer is installed in the fuel line between the engine's FCU and the fuel filter. ...
The two electrical fuel pumps are installed in the fuel system. Two, two-position switches labeled MAIN ELECTRIC FUEL PUMP and EMERGENCY ELECTRIC FUEL PUMP on the start panel electrically control each pump. The emergency pump switch has a red guard cover. Both the pumps ... provide a fuel pressure of 12.5 to 34 PSI. These pumps provide positive fuel pressure continuously during starting and engine operation. ...
TRACKING HOURS THE ENGINE HAS RUN WITHOUT INLET FUEL PRESSURE
The S2R-H80 has two electric fuel pumps to provide fuel to the engine under the required pressure. The Main Fuel Pump is supposed to be operating whenever the engine is running. If it fails, as indicated on the MVP-50T by a low fuel pressure alarm, its switch is to be placed off and the Emergency Fuel Pump is to be turned on. This provides the fuel pressure the engine requires and would normally be only a matter of a minute or two without fuel pressure. Accumulating significant time on the engine without inlet fuel pressure would require either a dual fuel pump failure or a pilot who neglects to turn the Emergency Fuel Pump on when the Main Fuel Pump fails. … The MVP-50T records the engine operating parameters, which gives the mechanic a way to determine how long the engine has run without fuel pressure recently.
At 1453, the recorded weather at the Salina Regional Airport, near Salina, Kansas, was, wind 320° at 4 knots, visibility 10 statute miles, sky condition clear, temperature 24° C, dew point 6° C, altimeter 30.02 inches of mercury.
N3045R was a full cantilever low-wing, all metal construction monoplane with serial No. H80-140, which was designed for agricultural flying. It has a maximum published gross weight of 10,500 lbs. The airplane was powered by a dual-spool turbopropeller 800-shaft horsepower GE H80-100 engine with serial No. 133001. The engine's gas generator section features a three-stage compressor that is comprised of two axial stages and one centrifugal stage, a reverse flow annular combustor, and a single stage turbine that drives the compressor. The power turbine section, counter rotating to the compressor's turbine, features a single-stage turbine and reduction gearbox (RGB) that drives the propeller. The engine was designed to operate with fuel being supplied to it by an electric fuel pump. However, the airplane's maintenance manual indicated the engine had a 100-hour limitation on operation without an electric fuel pump.
According to the engine manufacturer, the engine build was completed on August 3, 2013, and the engine was fitted to the airplane on September 12, 2013.
A pilot reported a very sensitive throttle while reducing power after takeoff where the torque will drop from 90% to 40% in a quick change even though the throttle is moved slowly and small amount. The squawk was troubleshot as a Fuel Control Unit (FCU) issue and on October 24, 2014, the FCU was replaced. Ground and flight tests were performed, and the airplane was returned to service. The airplane accumulated 494.34 hours of total time and 593 cycles since new at that time.
A logbook endorsement dated June 12, 2015, indicated the airplane had a reported hot wire strike. The RGB bearings and FCU were replaced. The airplane accumulated 634.32 hours of total time and 967 cycles since new at that time.
On April 18, 2016, the propeller governor was replaced at 1,069.87 hours of tachometer time.
An issue with engine starting was reported and on August 25, 2016, the starting and limiting unit (SALM) was replaced.
The engine was reported to start hot and on November 7, 2016, the FCU was changed. The airplane accumulated 1,629.27 hours of total time, 2,345 takeoffs, and 531 engine starts at that time.
The pilot reported that the last inspection on the airplane was an annual inspection completed on February 20, 2017. The airplane accumulated 1,629.3 hours of total flight time at the time of the annual inspection and 1,737.6 hours of total flight time, 2,460 takeoffs, and 553 engine starts at the time of the accident.
The pilot also stated that he installed a new electric main fuel pump T1800[9]-B8 serial No. 192737 on the airplane on August 4, 2016, and that the Hobbs meter indicated 1,428.6 hours. However, he did not provide a logbook endorsement of that main electric fuel pump. The pilot confirmed that main electric fuel pump accumulated about 309 hours of operation between the installation date and the accident date.
The electric fuel pump manufacturer website describes the 18000 series pump as a slung vane, self-priming, high suction lift, and integral pressure relief valve pump with a built-in bypass valve, which is powered by a permanent magnet motor. The pump is designed for emergency fuel boost, priming injected engines, and primary fuel pump applications where aviation gasoline, jet fuels, and JP-8 is used.
The H80 airplane was fitted with the Electronics International MVP-50T engine monitoring system. The MVP-50T consists of a multifunction glass panel engine monitoring and display system to display engine parameters, along with other data and warnings. The system performs monitoring tasks only. It does not perform any engine or aircraft system controlling functions. The MVP-50T system consists of the glass panel display unit (MVP-50T), the Electronic Data Converter (EDC-33T), and the transducer fuel signal conditioner.
Each wing contains integral wing tanks (wet wing fuel tanks) just outboard of the fuselage. The left wing and right-wing fuel tanks are interconnected through a header tank. The published fuel system capacity is 228 gallons. The aircraft's fuel system is equipped with a 1/4-inch mesh finger strainer installed in the outlet fitting from the header tank. The fuel supply line to the engine is routed from the header tank, located in the fuselage, through a fuel shut off valve, an emergency electric driven fuel pump, and then directed to the main electric driven fuel pump. The fuel supply line exits the main pump and passes through a 10-micron nominal main fuel filter. The fuel line then goes through the fuel pressure sensor, which is located before the firewall. The fuel line is then passed through the forward firewall to the fuel flow meter, and then enters the engine's fuel pump.
The emergency electric driven fuel pump is a backup system to provide continuous fuel pressure in case the main electric fuel pump fails. The main fuel pump and the emergency fuel pump are not to run simultaneously.
The fuel tank vent system is designed to keep the fuel spillage to a minimum. The fuel tanks are vented through tubing connected at both the inboard and outboard ends of the individual fuel tanks to the centrally located vent system in the fuselage. Ram air enters a vent scoop, on the fuselage, under the left wing and pressurizes the vent system to maintain positive pressure on the fuel tanks. The vent system is provided with two quick drains, located on the fuselage under each wing, to drain any fuel that might happened to have migrated in the tanks outboard vent lines. Fuel quantity is displayed individually on the cockpit panel display.
The airplane's flight manual (AFM) limitation section, in part, stated:
FUEL PUMPS: Continuous simultaneous operation of both fuel pumps is prohibited due to high fuel pressure.
The AFM advised pilots that an emergency hopper dump was available. The AFM, in part, stated:
Should circumstances arise that require an emergency landing, the hopper should be dumped by moving the dump lever full forward. Forward pressure on the control stick or forward trim (or both) should be used to prevent excessive nose up pitching moment. Max speed for Hopper dump is 158 MPH.
The AFM remedial action for an 'engine failure in-flight" in part, stated:
Engine failure symptoms could include any or all of the following:
a. Loud noises followed by heavy vibration and loss of power, smoke and/or flame.
b. Rapid loss of power with unusual noises, vibration or sudden increase of ITT.
c. Loss of power following a drop in oil pressure below redline or increase in oil temperature above redline or both.
d. Loss of power following overspeed of gas generator (Ng).
e. Engine explosion and flame & smoke.
If it is clear that the engine has failed, proceed as follows:
f. Hopper Dump DUMP LEVER FORWARD
g. Propeller Lever FEATHERED
h. Fuel Condition Lever CUT OFF
i. Power Lever IDLE POSITION
j. Fuel Valve Lever OFF
k. Fuel Pump Switches OFF
l. LAND AS SOON AS POSSIBLE
The AFM additionally advises on operation with a "failure of automatic fuel scheduling." The remedial action for this failure, in part, stated:
Normally the fuel scheduling is automatic, the engine power lever is positioned by the pilot and the fuel control schedules the fuel in a manner that allows smooth increases and decreases of power while not exceeding any engine limitations during acceleration or deceleration. In the event the automatic fuel scheduling fails, the engine will experience a "low side" failure, sometimes called a "roll back". The engine goes to minimum fuel flow, which is slightly below normal Ng idle speed. Fortunately the GE H-80 is equipped with an emergency governor which will allow the pilot to regain full control of the engine and safely return. While in Emergency Governor, all automatic fuel scheduling is lost and power is controlled by the fuel condition lever. Too rapid of movement of the fuel condition lever while in Emergency Governor can cause ITT exceedences, Ng speed exceedences, and compressor stalls. It's imperative that any power changes while in Emergency Governor be made smoothly and slowly allowing the engine to accelerate normally. Full power is available while in Emergency Governor. In the event automatic fuel scheduling is lost and the engine goes to minimum fuel flow.
a. Raise the switch guard and place Emergency Governor Switch ON.
b. Smoothly advance the ENGINE FUEL CONDITION LEVER until power is restored and a climb is established.
c. LAND AS SOON AS PRACTICABLE.
d. If power is not restored, LAND AS SOON AS POSSIBLE.
After power is restored and you have climbed to a safe altitude, reduce the engine power lever to idle and continue the flight to a safe place to land using the Fuel Condition Lever as you would normally use the Engine Power Lever.
The airplane's maintenance manual, in part, stated:
This group [includes a] fuel pressure gauge. ... These readings are displayed on the MVP-50T Glass Panel Engine Monitor. ...
The S2R-H80 aircraft is equipped with a MVP-50T glass panel display that reads fuel pressure and flow rate. The fuel flow transducer is installed in the fuel line between the engine's FCU and the fuel filter. ...
The two electrical fuel pumps are installed in the fuel system. Two, two-position switches labeled MAIN ELECTRIC FUEL PUMP and EMERGENCY ELECTRIC FUEL PUMP on the start panel electrically control each pump. The emergency pump switch has a red guard cover. Both the pumps ... provide a fuel pressure of 12.5 to 34 PSI. These pumps provide positive fuel pressure continuously during starting and engine operation. ...
TRACKING HOURS THE ENGINE HAS RUN WITHOUT INLET FUEL PRESSURE
The S2R-H80 has two electric fuel pumps to provide fuel to the engine under the required pressure. The Main Fuel Pump is supposed to be operating whenever the engine is running. If it fails, as indicated on the MVP-50T by a low fuel pressure alarm, its switch is to be placed off and the Emergency Fuel Pump is to be turned on. This provides the fuel pressure the engine requires and would normally be only a matter of a minute or two without fuel pressure. Accumulating significant time on the engine without inlet fuel pressure would require either a dual fuel pump failure or a pilot who neglects to turn the Emergency Fuel Pump on when the Main Fuel Pump fails. … The MVP-50T records the engine operating parameters, which gives the mechanic a way to determine how long the engine has run without fuel pressure recently.
The airplane came to rest in a field. FAA inspectors examined the wreckage and documented the site. A review of images confirmed that the airplane had impacted a fence line. A linear ground scar is visible between the fence and the resting airplane. The airplane exhibited substantial left-wing damage consistent with the wing impacting a fence post.
The pilot reported that there is a fuel sump in the bottom of the fuel farm holding tank to help catch water if any is present and it is sumped each morning before operation. The filtration system was two Facet model VF-21SB filters that were in line before supplying fuel to the airplane. The filters also absorb water and will not let water through the fuel system. Additionally, the pilot stated that he would visually inspect the airplane before each take off. He had flown several flights earlier in the morning and the airplane operated perfectly. The pilot reported that during the accident flight he did not dump the chemical load and he did not use the emergency governor procedures following the loss of engine power. The pilot stated in his safety recommendation, "I do not know if I could have done much more than what I did because of the terrain that I was in, but my recommendation to anyone flying an H80, if you have any loss of power, try and find a suitable place to sit the plane down if you have a chance. Do not wait for the engine to quit."
According to the airplane manufacturer, flight training is supplied to operators of its airplanes. Pilots were given ground training on the use of the backup fuel control. However, they were not given flight training on use of the backup. The majority of agricultural aircraft are single pilot aircraft, as such inflight training is not possible. Dual seat trainers do exist, but they are uncommon. Subsequent to the accident the airplane manufacturer purchased a simulator.
The airplane manufacturer's maintenance manual directs mechanics, in reference to the main electric fuel pump removal, to "detach drain tube at pump." However, the manual does not depict nor direct the mechanic on the proper installation and/or reinstallation of the main electric fuel pump's drain tube.
The airplane and engine manufacturer were asked if there were any differences in the way the Thrush airplanes equipped with GE engines and Thrush airplanes equipped with Pratt & Whitney engines were manufactured. A Thrush representative reported that the GE installation requires an electric fuel pump for engine operation while the Pratt and Whitney installation uses a mechanical fuel pump for engine operation and the electric fuel pump for starting. A GE representative responded that the GE H80 engine driven pump is capable of operation without an electric boost pump (positive fuel pressure is still required); without positive inlet fuel pressure the engine driven pump can cavitate which can result in engine power degradation. The engine driven fuel pump has been shown to operate for 100 hrs before cavitation damage has shown performance degradation. However, its suction capability corresponds to pressure loses across the airplane fuel system. The GE installation has two boost pumps installed in series and has sharp bends and changes of cross-sectional area in the fuel system plumbing. GE does not know what the pressure loss is across boost pumps as well as the whole airplane fuel system. According to GE, the engine driven fuel pump was never meant to be used as a prime fuel delivery unit. The H80 specifications requires a nominal inlet pressure range for engine driven pump, which is between 150 kPa Abs (21.75 psi absolute) to 300kPa Abs (43.5 psi absolute) for a fuel flow of 300kg/hr (661.3 lb/hr). To ensure sufficient fuel flow at maximum power, the engine is required to have a minimum fuel flow of 300kg/hr at minimum pressure of 150 kPa Abs delivered to it. GE advised that measurement of fuel pressure by itself is not sufficient to assess that the required fuel flow is delivered to the engine. However, this airplane/GE engine configuration has been FAA certified.
Subsequent to the accident, the airplane manufacturer issued Service Bulletin (SB) No. SB-AG-72, titled, "S2R-H8O FUEL PUMP FITTINGS & LINES IMPROVEMENT AND RELOCATION." According to the SB, it provides instructions and parts for an improved fuel pump system including new hardware, fuel lines, and location of pumps. The SB collocates both the main and emergency electric fuel pumps onto one support plate and depicts a tee to combine both their drain ports to a single descending drain line. An airplane manufacturer's representative additionally reported that "the purpose for this bulletin is to provide ease of replacement of fuel pumps in the field. This is to address field complaints about inadequate, for our market, longevity of the pumps and the complicated procedures for their replacement. This bulletin reflects a design change in place in production."
Similar electric fuel pumps are made that also have the capability to monitor brush wear or are fitted with brushless direct current motors that have a longer mean time between failure. However, at the time of the accident, the airplane's manufacturer had not specified these type of electric fuel pumps, nor applied for FAA approval of them, directly or as a suitable substitute. Subsequent to the accident, the airplane manufacturer advised that they were in the initial testing stage of a brushless pump.
The airplane wreckage was recovered to an airplane repair station where it was subsequently examined by the engine manufacturer under FAA supervision. No visual preimpact anomalies were detected during the examination. The MVP-50T unit was downloaded, the engine was removed and shipped to an engine repair station for further detailed examination along with both the main and emergency electric fuel pumps.
The engine was subsequently received in a sealed shipping container at the engine repair station where it was examined by the engine manufacturer under the supervision of a National Transportation Safety Board Senior Engine Investigator and the FAA. Upon removal from the shipping container, the engine was complete from the propeller shaft flange to the fuel control and starter-generator mounted on the accessory gear box (AGB). A close up visual inspection revealed that the engine did not have any fire damage, uncontainments, or case ruptures.
Power turbine to propeller hub continuity was established and a sound, consistent with a turbine rotation, could be heard from the engine exhaust when the starter-generator fan was rotated.
The AGB housing was intact, did not have any damage, and its gear train rotated when the AGB drive shaft was rotated. A disassembly examination revealed that the interior surface of the AGB did not have any debris and no preimpact anomalies were detected.
The fuel pump, part No. LUN 6290.04-8 serial No. 132016, was removed from the engine. The fuel control unit, part No. LUN 6590.07-8 serial No. 124004, was also removed from the engine and a clear fluid drained out that had an odor consistent with jet fuel. The fuel control unit input shaft rotated freely and smoothly. The throttle lever and fuel shutoff levers moved freely and smoothly throughout their respective full ranges of travel.
An examination of the engine oil system did not reveal any foreign debris, its screens were clean, and no anomalies were found.
The compressors did not exhibit any damage during visual and borescope examinations and the gas generator ball and roller bearings were intact, wet with oil, and did not have any rotational damage.
The air bleed valve was found in the closed position. However, the valve could be moved freely from the open to closed to open position without any binding.
The hot section's main shaft was intact, the deflector was in place, and they did not exhibit any damage. The fuel manifold and the slinger ring were intact. The slinger ring did not have any circumferential rub marks on its inner diameter and the ring did not exhibit any thermal distress. The outer combustion liner was intact and did not have any thermal distress or cracking. There was no cracking around any of the stub tubes. The inner combustion liner was intact and did not have any thermal distress or cracking. The thermal barrier coating was in place. However, the thermal barrier coating was flaking off in several places at the forward edge of the inner liner consistent with an engine in service this length of time.
The gas generator nozzle guide vane ring was intact and all of its vanes were in place. The vanes' airfoils did not show any damage or thermal distress. However, there were several airfoils that had some spots, consistent with metal spatter, on the convex surface of the airfoil.
The gas generator turbine disk was intact and all of the gas generator blades were in place with their tab locks. The gas generator blades did not exhibit any rub marks or damage to the airfoils. Additionally, the gas generator turbine blade shroud did not have any circumferential rub marks. The rear shaft cover was intact and did not have any rub marks on its labyrinth seals.
The power turbine nozzle guide vane ring was in place with all of its airfoils in place. There was no damage to any of the power turbine guide vane airfoils. However, the power turbine nozzle guide vane ring had two parallel rub marks that corresponded to the power turbine blade tip shroud knife edges consistent with an engine in service this length of time.
The power turbine disk was intact and all of the power turbine blades were in place in the disk with their lock tabs in place. There was no damage to any of the power turbine blades' airfoils. The power turbine blades' shroud knife edges were in place. The dimensional check of the power turbine blade shroud knife edges showed that they were within specified operational limits.
The left and right exhaust case halves and the power turbine shaft housing did not exhibit any damage. The power turbine shaft was intact and did not have any damage.
The reduction gearbox housing did not exhibit any damage its interior surface did not have any debris. The propeller shaft was intact. The propeller flange was flat and its bolt holes did not have any imprints of threads on their inner surfaces. The propeller flange rear face fillet radius to shaft edge had two spots consistent with tool marks that straddled each bolt hole location.
The ring gear was intact. There was an approximately 90° arc of teeth that had small areas of spalling on the gear teeth that corresponded to the areas of spalling on its satellite gear teeth.
The satellite gears were intact and did not have any deformation. The satellite gear teeth that engaged with the ring gear had small areas of spalling that corresponded to the spalling on the ring gear.
The quill shaft was intact and its gear teeth did not have any damage. The quill shaft bearing was intact, wet with oil, and did not have any rotational damage.
The reduction gear box magnetic chip detector did not have any debris. The reduction gear box pressure oil screen was clean. However, several soft, non-metallic flakes were observed on the reduction gear box scavenge screen.
The removed FCU and engine driven fuel pump along with the SALM unit were shipped to the engine manufacturer and were tested under the supervision of the Czech Republic Air Accidents Investigation Institute. The FCU was tested and was found to be operational where it met acceptance test performance requirements except for minimum starting fuel delivery. The fuel pump was tested and found to be operational and met all acceptance test performance requirements. The SALM unit LUN 2272.2 serial No. HG0019 was tested and it met acceptance test requirements.
The downloaded data from the MVP-50T indicated that there were no anomalies during the earlier flights during the day. Flight No. 6, the accident flight started about 405 minutes in the recorded data. No anomalies were observed up to 432 minutes into the flight. At that point, an instability in the gas generator turbine speed (Ng) was recorded it the range from 87% to 98%. Torque and interstage turbine temperature indications accordingly correspond to the Ng speed fluctuations. This anomaly lasted about 10 minutes until the 442.5 minute mark. Within that timeframe, the Electro Hydraulic Transducer (EHT) voltage level was observed between 12.3 to 14.2 volts DC. This voltage indication is consistent with no EHT fuel intervention. According to the engine manufacturer, at the 442.5 minute mark in the data, engine operating indications consistent with a fuel interruption were recorded. The Ng speed subsequently dropped below ground idle. However, the airplane manufacturer reported that the only time during the entire day that the fuel pressure was lower than the normal operating lower limit (12.5 psi) was prior to engine start and after the first impact.
Both electric fuel pumps were subsequently shipped to their manufacturer for detailed inspection. The emergency fuel boost pump, 18009-B8, serial No. 176388, was visually inspected and found to have shipping plugs installed into its fluid ports and the tamper evident seals appeared to be intact. Upon removal of the shipping plugs, pieces of red colored debris were noticed in both inlet and outlet ports. The shipping plugs were a similar color as the debris. The emergency fuel pump was intact and according to the manufacturer, exhibited a nearly new condition. An electrical meter revealed that both the positive and negative hook-up wires were not shorted to the emergency pump's motor case. Momentary power was applied to the pump and it ruled out that the pump had either a bind or an open circuit.
The emergency electric fuel boost pump was installed into an applicable standard acceptance test set-up. With all test set-up valves full open direct current power at 28 volts was supplied and the pump primed. While the still operating the outlet needle valve was adjusted to achieve 12.5 pounds per square inch gauge (PSIG) and the flow measured at 126.6 GPH while drawing 2.2 amperes of current. The outlet ball valve was fully closed long enough to record a settled 25.0 PSIG while drawing 3.0 amperes current. The pump met the standard acceptance criteria of 118 GPH minimum flow at 12.5 PSIG and of the maximum full relief pressure of 29.0 PSIG.
A disassembly examination revealed the emergency fuel pump's relief valve showed very little wear consistent with this model of pump that would have less than 50 hours operation. No anomalies were detected in the pumping mechanism. Disassembly of the emergency pump's motor showed that it was in a condition consistent with a pump that has less than 50 hours of operation.
The main electric fuel boost pump, T18009-B8, serial No. 192737, was visually inspected and found to have shipping plugs installed into its fluid ports and the tamper evident seals appeared to be intact. The main pump's grounding ring terminal connector exhibited a separation consistent with being cut off by side cutters. The main pump's positive hook-up wire was connected through a non-original equipment manufacturer's terminal to both a white wire and a green wire. An electrical meter revealed that both the positive and negative hook-up wires were not shorted to the main pump's motor case. Momentary power was applied to the pump and it ruled out that the pump had either a bind or an open circuit.
The main fuel boost pump was installed into an applicable standard acceptance test set-up. With all test set-up valves full open direct current power at 28 volts was supplied and the pump primed. The fuel pump's tone/pitch was found to vary. The pump manufacturer reported that a varying tone could be the result of a few reasons to include a known characteristic of this type of fuel pump motor's brushes with minimal percentage of service life remaining. While still operating, the outlet needle valve was adjusted to achieve 12.5 PSIG and the flow measured at 109.2 GPH while drawing 2.3 amperes of current. The outlet ball valve was fully closed long enough to record a settled 29.2 PSIG while drawing 3.0 amperes current. It should be noted that while the pump was still operational it did not meet the standard acceptance criteria of 118 GPH minimum flow at 12.5 PSIG or the maximum full relief pressure of 29.0 PSIG. However, the pump did meet the engine manufacturer's minimum fuel flow requirement of 98.95 GPH at a minimum pressure of 7.05 PSIG. The pump was powered straight to its factory supplied hook-up wires to rule out poor connections as root cause of above noted conditions. The pump was found to operate the same, which ruled out the wiring.
A disassembly examination revealed the main fuel pump's relief valve showed standard wear consistent with this model of pump that would have hundreds of hours operation. The pump's motor end frame exhibited some abrasion. However, no anomalies were detected with the pumping mechanism.
Disassembly of the main pump's motor showed that the negative commutation brush had minimal percentage of service life remaining, having worn thru the embedded braided copper shunt wire leaving little guidance/support via the brass brush tube. This condition requires current to be routed thru the brass brush tube and the conductive printed circuit board copper pad instead of the braided copper shunt wire. According to the pump manufacturer, the reduced guidance/support of a brush, worn at this stage, allows the remaining brush to "track" the commutator grooves similar to a needle on a record and this tracking presents as the "tone/pitch" variation observed earlier. With the negative brush's coil spring getting closer to the commutation event, it is subjected to more heat than what a new brush would encounter. The negative brush spring exhibited an anticipated "compression set" on approximately one or two coils from this heat.
The interior portion of the main electric fuel pump motor, to include its armature and end frame were coated with a dark, wet substance. A sample of the substance along with an exemplar brush were sent to the NTSB Materials Laboratory.
A senior chemist received the sample for testing and reported that the black, powdery solid was examined using a Fourier-Transform Infrared (FT-IR) spectrometer. The resulting spectrum showed no significant peaks for any organic (carbon-based) functional group. The material was then examined using an x-ray fluorescence (XRF) alloy analyzer. The testing results showed that the material was comprised of copper (Cu) with a small amount (~0.1%) of molybdenum (Mo). The XRF results for the exemplar brush head were the same as the black solid. The unknown solid contained material from a brush head.
According to a maintenance manual figure of the airplane's fuel system, the emergency fuel pump is mounted in a vertical orientation where the electric motor is mounded above its pump housing. The emergency fuel pump motor end frame is equipped with two threaded drain holes and is depicted with a directly downward draining drain tube. The main fuel pump is mounted in a horizontal orientation where the electric motor is mounted adjacent to its pump housing. The main fuel pump end frame is also equipped with two threaded drain holes. A picture of the accident airplane showed that it was plumbed with a drain tube that was routed above the pump before it had a descending path below the pump.
The recorded data showed that there was stabilizing time between flights before flight No. 5. There was no recorded refueling stabilizing time between flight No. 5 and No. 6 consistent with refueling conducted immediately prior the airplane takeoff. The AFM indicated that there shall be "reasonable" time to allow stabilizing of fuel in both tanks.
The airplane manufacturer was asked to review the MVP-50T data and report how much time the accident engine had accumulated while operating without fuel supplied under electric pump pressure. The airplane manufacturer's representative replied that he reviewed the recorded data from the accident airplane and that MVP data indicated less than 1 minute of total engine operation without electric fuel pump operation. He additionally reported that a master warning indication occurred near the end of the accident flight. It was not a low fuel pressure alarm. However, it was a stall warning indication.
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