Saturday, September 28, 2019

Aircraft Structural Failure: Vans RV-3, N177TT; fatal accident occurred June 17, 2017 near Payette Municipal Airport (S75), Idaho

Prescott "Gene" Wilkie


The National Transportation Safety Board did not travel to the scene of this accident. 


Additional Participating Entity:

Federal Aviation Administration / Flight Standards District Office; Boise, Idaho

Aviation Accident Factual Report - National Transportation Safety Board: https://app.ntsb.gov/pdf


Investigation Docket - National Transportation Safety Board: https://dms.ntsb.gov/pubdms 

 
http://registry.faa.gov/N177TT


Location: Payette, ID
Accident Number: WPR17FA128
Date & Time: 06/17/2017, 1145 MDT
Registration: N177TT
Aircraft: PRESCOT E. WILKIE RV-3
Aircraft Damage: Destroyed
Defining Event: Aircraft structural failure
Injuries: 1 Fatal
Flight Conducted Under: Part 91: General Aviation - Personal 

On June 17, 2017, about 1145 mountain daylight time, an experimental, amateur-built Vans RV-3 airplane, N177TT, impacted terrain following an in-flight separation of the left wing while maneuvering near Payette Municipal Airport (S75), Payette, Idaho. The airline transport pilot sustained fatal injuries, and the airplane was destroyed. The airplane was registered to and was being operated by the pilot as a Title 14 Code of Federal Regulations Part 91 personal flight. Visual meteorological conditions prevailed, and no flight plan had been filed for the local flight, which departed S75 about 1140.

A review of a video of the accident sequence revealed that the airplane departed runway 31 and then made a shallow ascending right turn to the north. Seconds later, the airplane made a sweeping left descending turn to align with runway 13. The pilot proceeded to make a high-speed, low-altitude pass over the runway, during which the airplane descended to less than 20 ft above ground level (agl), before the pilot executed a sharp pull-up maneuver at the departure end of the runway. After climbing the airplane several hundred feet, the pilot initiated a left turn to a northerly heading. Several seconds later, the airplane entered a descent, and the left wing separated from the airplane. The airplane impacted flat farmland about 3,025 ft northeast of the departure end of runway 31. 

Pilot Information

Certificate: Airline Transport; Flight Instructor
Age: 74, Male
Airplane Rating(s): Multi-engine Land; Single-engine Land
Seat Occupied: Left
Other Aircraft Rating(s): None
Restraint Used:
Instrument Rating(s): Helicopter
Second Pilot Present: No
Instructor Rating(s): Airplane Single-engine; Helicopter
Toxicology Performed: Yes
Medical Certification: Class 2 With Waivers/Limitations
Last FAA Medical Exam: 04/27/2016
Occupational Pilot: No
Last Flight Review or Equivalent:
Flight Time: 16700 hours (Total, all aircraft) 

The 74-year-old pilot held an airline transport pilot certificate with airplane single engine land, airplane multiengine land, and helicopter ratings. He also held a flight instructor certificate for airplane single engine and helicopter. The pilot was issued a Federal Aviation Administration (FAA) second-class medical certificate on April 27, 2016, with the restriction that he "must wear corrective lenses." At that time, he reported 16,700 total hours of flight experience, about 20 hours of which were in the preceding 6 months.

The pilot also held an FAA mechanic certificate with airframe and powerplant ratings and an inspection authorization.

Aircraft and Owner/Operator Information

Aircraft Make: PRESCOT E. WILKIE
Registration: N177TT
Model/Series: RV-3
Aircraft Category: Airplane
Year of Manufacture: 2006
Amateur Built: Yes
Airworthiness Certificate: Experimental
Serial Number: 10508
Landing Gear Type: Tailwheel
Seats: 1
Date/Type of Last Inspection: 07/01/2010, Continuous Airworthiness
Certified Max Gross Wt.: 2952 lbs
Time Since Last Inspection:
Engines: 1 Reciprocating
Airframe Total Time:
Engine Manufacturer: Lycoming
ELT: Not installed
Engine Model/Series: O-235-N2C
Registered Owner: On file
Rated Power: 118 hp
Operator: On file
Operating Certificate(s) Held: None

The single-seat, single-engine, low-wing, kit airplane was a metal monocoque construction, had a conventional tail, and was equipped with fixed conventional landing gear. The airplane was powered by a four-cylinder Lycoming O-235 engine and was equipped with a two-bladed fixed-pitch wooden propeller. The airplane was modified from a standard RV-3 to have an open cockpit configuration. According to Vans Aircraft, the kit manufacturer, the accident airplane was sold as a kit in 1978.

Most of the airplane was assembled by at least three previous owners. The accident pilot bought the airplane not fully assembled in 1998, and he completed assembling it in 2007. The airplane was granted an FAA airworthiness certificate in the experimental category on May 2, 2007. Standard operating limitations, including requiring an annual condition inspection and prohibiting aerobatic flight, were issued in conjunction with the airworthiness certificate.

In 2014, the airplane was disassembled and shipped in a container from Hawaii to Idaho. The pilot reassembled the airplane in a hangar at S75 before the accident flight. The accident flight was the first flight after the airplane had been reassembled.

Meteorological Information and Flight Plan

Conditions at Accident Site: Visual Conditions
Condition of Light: Day
Observation Facility, Elevation: KONO, 2193 ft msl
Distance from Accident Site: 7 Nautical Miles
Observation Time:1153 MDT 
Direction from Accident Site: 214°
Lowest Cloud Condition: Clear
Visibility: 10 Miles
Lowest Ceiling: None
Visibility (RVR):
Wind Speed/Gusts: Calm /
Turbulence Type Forecast/Actual: / None
Wind Direction:
Turbulence Severity Forecast/Actual:
Altimeter Setting: 30.14 inches Hg
Temperature/Dew Point: 22°C / 5°C
Precipitation and Obscuration: No Obscuration; No Precipitation
Departure Point: Payette, ID (S75)
Type of Flight Plan Filed: None
Destination: Payette, ID (S75)
Type of Clearance: None
Departure Time: 1140 MDT
Type of Airspace: Class G 

At 1153, the recorded weather observation at Ontario Municipal Airport, Ontario, Oregon, located 7 nautical miles northeast of the accident site, reported calm winds, visibility 10 statute miles, sky clear, temperature 22°C, dew point 5°C, and an altimeter setting of 30.14 inches of mercury. The density altitude at the time of the accident was about 2,899 ft.

Airport Information

Airport: Payette (S75)
Runway Surface Type: N/A
Airport Elevation: 2228 ft
Runway Surface Condition:
Runway Used: N/A
IFR Approach: None
Runway Length/Width:
VFR Approach/Landing: None 

S75 is publicly owned and operated by the City of Payette. The airport is not equipped with an operating control tower. Runway 13/31 is 3,000 ft by 50 ft with displaced thresholds at both ends. The airport elevation is 2,228 ft mean sea level.

Wreckage and Impact Information

Crew Injuries: 1 Fatal
Aircraft Damage: Destroyed
Passenger Injuries: N/A
Aircraft Fire:None 
Ground Injuries: N/A
Aircraft Explosion:None 
Total Injuries: 1 Fatal
Latitude, Longitude: 44.102778, -116.896944 (est) 

The main wreckage included the fuselage, engine, empennage, and right wing. The left wing was located about 1,450 ft southeast of the main wreckage, and the left flap was located about 1,150 ft south-southeast of the main wreckage. There was no evidence of a postcrash fire. The wreckage was crushed and deformed consistent with impact. The vertical stabilizer, rudder, horizontal stabilizer, and elevators remained attached to the empennage, and they all displayed extensive impact damage. Rudder control continuity was established from the surface to the rudder pedals.

Control continuity was established from the elevators to the control horn in the empennage, and the elevators moved freely when actuated by hand. The aft elevator control tube was bent in multiple places and fractured at the forward end. The forward elevator control tube was fractured at the aft rod end about 2 ft aft of the control stick. All control tube separations had a dull, grainy appearance consistent with overstress separation.

The right wing had extensive crushing damage and fragmentation. The right main wing spar was mostly intact but damaged, and it was deformed from wing station (WS) 0 to the right-wing tip. The right aileron control tube was separated in multiple places, and the right aileron remained attached to the wing. The aileron control tube fracture surfaces had a dull, grainy appearance consistent with overstress separation. The right flap remained attached to the wing. Right aileron control continuity was established from the fracture points to the right aileron and the control stick.

The left wing separated from the airplane in the wing root area during the accident sequence. The left aileron remained attached to the wing, and the left flap was separated. The left-wing root rib and the inboard portion of the wing were buckled and deformed upward, consistent with wing separation in an upward direction. The left-wing root rib was 0.020-inch thick, and no stiffeners or additional angles were installed on it. No additional wing ribs were installed in the wing walk area outboard of the root rib. The left aileron control tube was fractured at the wing separation point. The aileron control tube fracture surfaces had a dull, grainy appearance consistent with overstress separation. Left aileron control continuity was established from the fracture point to the left aileron and control stick.

The left main wing spar was fractured in the left-wing root area and damaged adjacent to the fracture point. The outboard portion of the left main wing spar was intact and installed in the left wing. The inboard portion of the left main wing spar remained attached to the right main wing spar at WS 0. All 4 four wing fittings were intact and installed at the center portion of the main spar. The wing bolts were numbered starting at 1 adjacent to WS 0 on both the left and right main wing spars and ending at 15.The right upper spar cap at WS 0, from forward to aft, consisted of a 0.250-inch-thick steel wing fitting, seven 0.125-inch-thick aluminum bars, a 0.040-inch-thick aluminum shim, a 0.125-inch-thick aluminum web, a 0.040-inch-thick aluminum aft fuselage bulkhead, and a 0.250-inch-thick steel wing fitting. The 0.040-inch-thick forward fuselage bulkhead only engaged holes 9 and 10 on the right upper spar cap. Bolts were installed in holes 1 to 10 and 12 to 15 with the heads of the bolts on the forward side of the spar cap. Washers were installed underneath the bolt heads in holes 3, 5 to 10, and 13 to 15. No bolt was installed in hole 11. The head markings on the bolts installed in holes 1 to 4 indicated that they were AN5-X(A) bolts. The aft side of the spar cap was inaccessible, so the grip length, drilling of the shank, and type of nut or washer could not be determined. Hole 5 had an NAS1304-27H bolt installed. The aft side of the spar cap was inaccessible, so the drilling of the shank and the type of nut or washer could not be determined. Holes 6 to 10 and hole 12 had AN3HX bolts installed with L, X, and SC marked on the head. Holes 13 to 15 had AN3-XA bolts installed with C, X, and S marked on the head. Self-locking nuts with no washers were installed on the bolts in holes 6 to 10 and 12 to 15. An L-angle stiffener was installed between the upper and lower spar caps at hole 12. No evidence of adhesive was found between the spar cap layers.

The right lower spar cap at WS 0, from forward to aft, consisted of a 0.250-inch- thick steel wing fitting, a 0.040-inch thick aluminum forward fuselage bulkhead, seven 0.125-inch- thick aluminum bars, a 0.040-inch- thick aluminum shim, a 0.125-inch- thick aluminum web, a 0.040-inch- thick aluminum aft fuselage bulkhead, and a 0.250-inch- thick steel wing fitting. Bolts were installed in holes 1 to 9 and 12 to 15. The heads of the bolts were on the forward side of the spar cap except for hole 4, which had the head on the aft side. Washers were installed underneath the bolt heads in holes 2, 3, 5 to 9, and 13- to 15. No bolts were installed in holes 10 or 11, and there was no hole drilled in the forward fuselage bulkhead at the hole 10 location. Holes 1-3 had AN5-X(A) bolts installed based on the head markings. The aft side of the spar cap was inaccessible, so the type of bolt in hole 4 and the grip length, drilling of the shank, and type of nut or washer in the other locations could not be determined. The bolt in hole 4 had a washer and self-locking nut installed. Hole 5 had a NAS1304-27H bolt installed. The aft side of the spar cap was inaccessible, so the drilling of the shank and type of nut or washer could not be determined. Holes 6-9 and hole 12 had AN3HX bolts installed with L, X, and SC marked on the head. Holes 13-15 had AN3-XA bolts installed with C, X, and S marked on the head. Self-locking nuts with no washers were installed on the bolts in holes 6-9 and 12-15. There was an L-angle stiffener installed between the upper and lower spar caps at hole 12. There was no evidence of adhesive between the spar cap layers. There was some evidence of corrosion between the right lower spar cap layers.

The left upper spar cap at WS 0, from forward to aft, consisted of a 0.250-inch thick steel wing fitting, seven 0.125-inch thick aluminum bars, a 0.125-inch thick aluminum web, a 0.040-inch thick aluminum aft fuselage bulkhead, and a 0.250-inch thick steel wing fitting. There was no 0.040-inch thick shim installed between spar cap layer 7 and the 0.125-inch thick aluminum web. The 0.040-inch thick forward fuselage bulkhead only engaged holes 9 and 10 on the left upper spar cap. Bolts were installed in holes 1-10 with the heads of the bolts on the forward side of the spar cap. Washers were installed underneath the bolt heads in holes 1-10. There were no bolts installed in holes 11-15 in the area where the left upper spar cap fractured and there was no damage to the holes. None of the missing bolts were recovered. Holes 1, 2, and 4 had AN5-X(A) bolts installed based on the head markings. The aft side of the spar cap was inaccessible, so the grip length, drilling of the shank, and type of nut or washer could not be determined. Hole 3 had a NAS1305-26H bolt installed and hole 5 had a NAS1304-27H bolt installed. The aft side of the spar cap was inaccessible, so the drilling of the shank and type of nut or washer could not be determined. Holes 6-10 had AN3HX bolts installed with L, X, and SC marked on the head. Self-locking nuts with no washers were installed on the bolts in holes 6-10. The L-angle stiffener normally installed between the upper and lower spar caps at hole 12 was separated and remained attached to the wing root rib. The upper hole in the stiffener was elongated. There was no evidence of adhesive between the spar cap layers. There was evidence of substantial corrosion between the left upper spar cap layers. The spar cap layer 1 was fractured about WS 17.25 through hole 15 and the outboard 3 inches was not recovered. Spar cap layers 2-7 were fractured about WS 16 through hole 14 and the fractures were matched with spar cap layers 2-7 on the left outboard wing. The upper spar cap layers were splayed open and exhibited S-bending in the area of fracture and the rivets through the spar cap were popped from about WS 17.5 to WS 22.5. The 0.125-inch thick aluminum web was intact to its production edge about WS 18. All the fractures exhibited a dull, grainy appearance consistent with overstress separation.

The left lower spar cap at WS 0, from forward to aft, consisted of a 0.250-inch thick steel wing fitting, a 0.040-inch thick aluminum forward fuselage bulkhead, seven 0.125-inch thick aluminum bars, a 0.040-inch thick aluminum shim, a 0.125-inch thick aluminum web, a 0.040-inch thick aluminum aft fuselage bulkhead, and a 0.250-inch thick steel wing fitting. Bolts were installed in holes 1-9. The heads of the bolts were on the forward side of the spar cap except for hole 4 which had the head on the aft side. Washers were installed underneath the bolt heads in holes 1-3 and 5-9. There were no bolts installed in holes 10-15 in the area where the left lower spar cap fractured and there was no damage to the holes. None of the missing bolts were recovered. There was no hole drilled in the forward fuselage bulkhead at the hole 10 location. Holes 1-3 had AN5-X(A) bolts installed based on the head markings. The aft side of the spar cap was inaccessible, so the type of bolt in hole 4 and the grip length, drilling of the shank, and type of nut or washer in the other locations could not be determined. The bolt in hole 4 had a washer and self-locking nut installed. Hole 5 had a NAS1304-27H bolt installed. The aft side of the spar cap was inaccessible, so the drilling of the shank and type of nut or washer could not be determined. Holes 6-9 had AN3HX bolts installed with L, X, and SC marked on the head. Self-locking nuts with no washers were installed on the bolts in holes 6, 9, and 10. The L-angle stiffener normally installed between the upper and lower spar caps at hole 12 was separated and remained attached to the wing root rib. There was no evidence of adhesive between the spar cap layers. There was some evidence of corrosion between the left lower spar cap layers. The spar cap layer 1 was intact to the production edge at WS 20. Heads of three rivets at the outboard end were installed but the rivets were fractured. Spar cap layer 2 fractured about WS 10.5 through hole 9, layers 3-6 fractured about WS 13 through hole 11, and layer 7 fractured about WS 10.5 through hole 9. The fractures on layers 2-7 were matched with layers 2-7 on the left outboard wing. The left lower spar cap layers were splayed open and twisted and the rivets through the spar cap were popped from about WS 17.5 to WS 20. The 0.040-inch thick aluminum shim was intact to its production edge at WS 14 and the 0.125-inch thick aluminum web was intact to its production edge at WS 18. All the fractures exhibited a dull, grainy appearance consistent with overstress separation.

The bolts in holes 7, 8, and 9 were removed from the left lower spar cap and examined. There were no nuts present on bolts 7 and 8 and the bolts were fractured through the cotter pin hole in the shank. The threads around the cotter pin hole were also stripped. The self-locking nut remained installed on bolt 9 with 1-2 threads visible. Removal of the nut showed the hole in the shank of the bolt was just visible beyond the material stack up and the nut was engaged in the area of the hole as installed. Examination of several other AN3HX bolts installed on the left and right spar caps showed a similar arrangement. The hole 9 bolt was deformed and measured 1.670 inches long.

The aft spar carry-through was found intact in the main wreckage and consisted of a 0.125-inch-thick L-angle with an additional 0.125-inch-thick bar riveted to the vertical flange. The right-wing aft spar attachment point consisted of a 0.125-inch-thick lug and a 0.125-inch-thick reinforcement plate. The right side of the aft spar carry-through remained attached to the right-wing lug, and the entire carry-through was bent aft. An AN4-6A bolt with no washers and a low-profile nut were in place in the right attachment point. The bolt in the left side of the aft spar carry through was fractured with a portion of the shank remaining in the carry through. The head was fractured from the bolt, and the nut was pulled off the threads. The hole in the left side of the aft spar carry through was elongated in an outboard direction. The left-wing aft spar attachment point was fractured from the wing and was not recovered.

Impact damage was noted throughout the engine, however, all four cylinders remained attached to the engine crankcase. A large hole was noted to the forward section of the crankcase. The hole was about 4 inches in diameter and was consistent with ground impact. The left magneto, starter, carburetor, oil cooler, and alternator separated from the engine and were found within the main wreckage. The engine case was cracked between the push rods on the No. 4 cylinder forward of the cylinder upward toward the case half. The oil sump displayed extensive ground impact damage. Postaccident examination of the engine revealed no evidence of any preimpact mechanical malfunctions or failures that would have prohibited normal operation. 

Medical And Pathological Information

No autopsy was performed.

The FAA Forensic Sciences Laboratory performed toxicological testing on specimens from the pilot. The tests were negative for ethanol and all tested drugs and their metabolites. Tests for carbon monoxide and cyanide were not performed.

Additional Information

After a series of accidents involving RV-3 in-flight wing separations, Vans conducted extensive testing. The testing showed that the initiating failure of the RV-3 wing was the buckling of the upper spar cap and that bonding the spar caps together as a unit during assembly, which was optional, provided better resistance to buckling. The first modification to the RV-3 wing (CN-1) involved strengthening the rear spar attachment points and carry through and modifying the wing root rib. The second modification to the RV-3 wing (CN-2-I or CN-2-II) involved adding stiffening angles to the main spar.

The RV-3 wing has a NACA 23012 airfoil. The original wing spar (referred to as Type I, circa 1973 to 1983) was mathematically stress analyzed to design and ultimate load limits of 6.0 and 9.0 Gs, respectively, at an aerobatic gross weight of 1,050 lbs (the nonaerobatic gross weight is 1,100 lbs). A test in 1982 verified that it met the 9 G ultimate load criteria. The spar consisted of .040-inch aluminum channel web with a buildup of seven 1/8-inch thick by 1-1/4 inch wide bars held together with 1/8-inch AN470 rivets to form the upper and lower caps. As an assembly option, an epoxy adhesive could be used to bond the seven aluminum bars together to form a single unit to facilitate drilling and riveting the unit to the spar web. (It was later discovered that the adhesive provided some interbar shear and column strength, but the bonding process can deteriorate, so it was not considered in the design as contributing to spar strength.)

On March 16, 1981, the FAA issued General Notice TWA 1/40 SVCB, which prohibited aerobatics in the RV-3. The action was permanent and could not be rescinded. Following this action, Vans issued Change Notice 1 (CN-1) to RV-3 owners and builders. Briefly, CN-1 modified the wing by reinforcing the rear spar attachment point and strengthening the wing root rib. FAA and Canadian Ministry of Transport reports on the RV-3 accidents suggested that these areas could have been the primary failure points. When CN-1 was drafted, it appeared that the only means of regaining aerobatic operating authorization would be for individual RV-3 owners to change the airplane's designation to RV-3A. Soon thereafter, the FAA issued another letter stating that RV-3 owners showing compliance with CN-1 could reapply for aerobatic operating limitations. Thus, the RV-3A designation was adopted by some builders but does not signify any definite main spar structural distinction.

In 1984, Vans redesigned the wing spar (referred to as Type II). It incorporated five 3/16-inch thick by 1 1/4-inch wide aluminum bars held together with 3/16-inch diameter AN470 rivets. Although the primary purpose was to simplify assembly and minimize the possibility of assembly errors, the calculated bending strength was slightly increased. Further, two additional ribs were added in the root rib area to increase torsional stiffness, and the rear spar attachment was strengthened.

Despite the changes, the suggested airspeed limitations remained relatively unchanged. Vne (never exceed speed) is 210 mph. Va (maneuvering speed) is 127 mph (down from 132 mph). Vs (stall speed, clean) is 54 mph. According to Vans, because of the high ratio between Vne and Vs, the RV-3 is more susceptible to pilot-induced overstress than most contemporary light airplanes.

In 1996, an RV-3 wing with a Type I spar was again static load tested. During the tests, the wing failed below the 9.0 G load level. A review of test data revealed that the 1982 test had been performed on a wing whose spar had been assembled with the optional epoxy bonding. The 1996 test had been performed on a wing whose spar had been assembled without the epoxy adhesive. Although the epoxy adhesive had not been calculated to add any spar bending strength, it appeared to have added compression buckling strength. As a result of further static load testing on both Type I and Type II wing spars, Vans issued Change Notice 2 (CN-2-I for Type I spars and CN-2-II for Type II spars), which recommended main spar modifications which it deemed necessary for aerobatic strength. CN-2 included a detailed history, explanation, and recommendations, and was sent to all known RV-3 and RV-3A owners and builders. Vans maintained that both CN-1 and CN-2 (I or II) were necessary to achieve the aerobatic strength of the wing.

Loss of Engine Power (Total): Thrush S2R-H80, N3045R; accident occurred May 05, 2017 in Gypsum, Saline County, Kansas

The National Transportation Safety Board did not travel to the scene of this accident.

Additional Participating Entities:
Federal Aviation Administration / Flight Standards District Office; Wichita, Kansas
General Electric; Evendale, Ohio
Thrush Aircraft Inc; Albany, Georgia
Weldon Pumps; Oakwood Village, Ohio
Air Accidents Investigation Institute; Letnany, FN
GE Aviation Czech; Letnany, FN

Aviation Accident Factual Report - National Transportation Safety Board: https://app.ntsb.gov/pdf  


Investigation Docket - National Transportation Safety Board: https://dms.ntsb.gov/pubdms 

http://registry.faa.gov/N3045R


Location: Gypsum, KS
Accident Number: CEN17LA176
Date & Time: 05/05/2017, 1450 CDT
Registration: N3045R
Aircraft: THRUSH AIRCRAFT INC S2R-H80
Aircraft Damage: Substantial
Defining Event: Loss of engine power (total)
Injuries: 1 None
Flight Conducted Under: Part 137: Agricultural 

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.

Pilot Information

Certificate: Commercial
Age: 38, Male
Airplane Rating(s): Multi-engine Land; Single-engine Land
Seat Occupied: Single
Other Aircraft Rating(s): None
Restraint Used: 5-point
Instrument Rating(s): None
Second Pilot Present: No
Instructor Rating(s): None
Toxicology Performed: No
Medical Certification: Class 2 Without Waivers/Limitations
Last FAA Medical Exam: 03/08/2017
Occupational Pilot: Yes
Last Flight Review or Equivalent: 01/07/2017
Flight Time:  5215 hours (Total, all aircraft), 1737.6 hours (Total, this make and model), 5173.9 hours (Pilot In Command, all aircraft), 108 hours (Last 90 days, all aircraft), 108 hours (Last 30 days, all aircraft), 6 hours (Last 24 hours, all aircraft) 

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.

Aircraft and Owner/Operator Information

Aircraft Make: THRUSH AIRCRAFT INC
Registration: N3045R
Model/Series: S2R-H80
Aircraft Category: Airplane
Year of Manufacture:
Amateur Built: No
Airworthiness Certificate: Restricted
Serial Number: H80-140
Landing Gear Type: Tailwheel
Seats: 1
Date/Type of Last Inspection: 02/20/2017, Annual
Certified Max Gross Wt.: 10500 lbs
Time Since Last Inspection:
Engines: 1 Turbo Prop
Airframe Total Time: 1737.6 Hours at time of accident
Engine Manufacturer: GE
ELT: Not installed
Engine Model/Series: H80-100
Registered Owner: Central Ag Air LLC
Rated Power: 800 hp
Operator: Central Ag Air LLC
Operating Certificate(s) Held: Agricultural Aircraft (137)
Operator Does Business As:
Operator Designator Code: 37AG 

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 happen 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:

POWER PLANT INSTRUMENTS

This group [includes a] fuel pressure gauge. ... These readings are displayed on the MVP-50T Glass Panel Engine Monitor. ...

MVP-50T FUEL FLOW

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. ...

ELECTRIC FUEL PUMPS

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.

Meteorological Information and Flight Plan

Conditions at Accident Site: Visual Conditions
Condition of Light: Day
Observation Facility, Elevation: KSLN, 1289 ft msl
Distance from Accident Site: 15 Nautical Miles
Observation Time: 1453 CDT
Direction from Accident Site: 300°
Lowest Cloud Condition: Clear
Visibility:  10 Miles
Lowest Ceiling: None
Visibility (RVR):
Wind Speed/Gusts: 4 knots /
Turbulence Type Forecast/Actual:
Wind Direction: 320°
Turbulence Severity Forecast/Actual:
Altimeter Setting: 30.02 inches Hg
Temperature/Dew Point: 24°C / 6°C
Precipitation and Obscuration: No Obscuration; No Precipitation
Departure Point: MARION, KS (43K)
Type of Flight Plan Filed: None
Destination: MARION, KS (43K)
Type of Clearance: None
Departure Time: 1350 CDT
Type of Airspace: 

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.

Wreckage and Impact Information

Crew Injuries: 1 None
Aircraft Damage: Substantial
Passenger Injuries: N/A
Aircraft Fire: None
Ground Injuries: N/A
Aircraft Explosion: None
Total Injuries: 1 None
Latitude, Longitude: 38.661389, -97.379722 (est) 

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.

Tests And Research

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.

Additional Information

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.

Abnormal Runway Contact: Cirrus SR22, N94LP; fatal accident occurred April 24, 2017 at Meriden Markham Municipal Airport (KMMK), New Haven County, Connecticut

Todd G. Gunther
Investigator In Charge


The National Transportation Safety Board traveled to the scene of this accident.

Additional Participating Entities:

Federal Aviation Administration / Flight Standards District Office; Enfield, Connecticut
Cirrus Aircraft; Duluth, Minnesota
Continental Motors; Mobile, Alabama

Aviation Accident Factual Report - National Transportation Safety Board: https://app.ntsb.gov/pdf


Investigation Docket - National Transportation Safety Board: https://dms.ntsb.gov/pubdms

http://registry.faa.gov/N94LP



Location: Wallingford, CT
Accident Number: ERA17FA167
Date & Time: 04/24/2017, 1825 EDT
Registration: N94LP
Aircraft: CIRRUS DESIGN CORP SR22
Aircraft Damage: Destroyed
Defining Event: Abnormal runway contact
Injuries: 1 Fatal, 1 Serious
Flight Conducted Under: Part 91: General Aviation - Personal 

HISTORY OF FLIGHT

On April 24, 2017, about 1825 eastern daylight time, a Cirrus Design Corporation SR22, N94LP, impacted terrain in Wallingford, Connecticut, following a loss of control during an aborted landing at Meriden Markham Municipal Airport (MMK), Meriden, Connecticut. The private pilot was fatally injured and the passenger was seriously injured. The airplane was destroyed by impact forces and a postcrash fire. The airplane was privately owned and was being operated by the pilot as a Title 14 Code of Federal Regulations Part 91 personal flight. Visual meteorological conditions prevailed, and no flight plan was filed for the local flight.

According to witness statements and security camera video, about 1740, the airplane departed the airport to the east and returned to the airport around 1817. Witnesses described that the airplane was "fast and high" as it approached runway 18. The airplane then flared about 10 ft above the runway before it abruptly descended and touched down about halfway down the runway. The airplane bounced two or three times and became airborne again, then banked about 30° to the left and climbed to airport traffic pattern altitude.

The pilot's second landing approach appeared to be slower, but the airplane was again high. The airplane flared about 10 ft above the runway, abruptly descended, and touched down about halfway down the runway. It bounced two or three times; the pilot then initiated a go-around. One witness described that, during the subsequent climb, the airplane entered a 40° nose-up attitude and it sounded as if the airplane was "hanging on its prop." About 75 ft above the ground, the airplane rolled into a steep left descending turn. It then impacted the ground, cartwheeled, impacted the airport perimeter (security) fence, slid across the ground while continuing to turn to the left, came to rest, and caught fire.

According to the passenger, who was the pilot's son, the accident flight was his father's first flight in the airplane without an instructor and was a proficiency flight in preparation for an upcoming trip to North Carolina. The passenger stated that he did not handle the flight controls during the accident flight and that there were no unusual noises or issues with the airplane. During the pilot's first landing attempt, which was supposed to have been a full-stop landing, the pilot said "oops," commenced a go-around, then said, "let's try it again." During the second landing attempt, the airplane bounced "a couple of times" and the bounces were "pretty high."



PERSONNEL INFORMATION

According to Federal Aviation Administration (FAA) and pilot records, the pilot held a private pilot certificate with ratings for airplane single-engine land and instrument airplane. His most recent FAA third-class medical certificate was issued on February 1, 2017. On that date, he reported about 1,200 total hours of flight experience.

The pilot had flown out of MMK for several years. He previously owned a Piper PA-28-180, which he recently sold, and had purchased the accident airplane about 3 weeks before the accident. After the purchase of the airplane, he had taken transition training from a local flight instructor who also owned an SR22. The pilot received ground instruction from the flight instructor as well as 2 hours of dual instruction in the flight instructor's SR22, and 8.5 hours of dual instruction in the accident airplane. During that time, the pilot performed 12 landings.

The flight instructor stated that he used the Cirrus Transition Training Manual as a guide for the accident pilot's training, instructed him in the use of the airplane's avionics, and had taught him to use more right rudder input during climb. He endorsed the pilot for operation of high-performance airplanes (airplanes equipped with engines producing 200 horsepower or greater) on April 23, 2017, the day before the accident. Review of pilot records revealed that the pilot's most recent flight review occurred on October 30, 2014.



AIRCRAFT INFORMATION

The accident airplane was a low-wing, fully cantilevered, single-engine monoplane of composite construction. It was equipped with fixed tricycle configuration landing gear, with a castering nose wheel, and steering was accomplished via differential braking on the main wheels. It was also equipped with a ballistic recovery system known as the Cirrus Airframe Parachute System (CAPS), which could, under certain conditions, lower the entire airplane to the ground in an emergency. It was powered by a fuel-injected, horizontally opposed, air-cooled, 310-horsepower, Continental IO-550-N27B engine, driving a constant-speed, variable pitch Hartzell three-bladed propeller.

According to FAA and airplane maintenance records, the airplane was manufactured in 2005. The airplane's most recent annual inspection was completed on March 13, 2017. At the time of the inspection, the airplane had accrued about 1,229 total hours of operation.

The four-seat cabin included a composite roll cage within the fuselage structure to provide roll protection for the cabin occupants and was accessed through doors on either side of the fuselage. The seats were equipped with 4-point, integrated seat belt and shoulder harness assemblies with inertia reels, and seat bottoms with an integral aluminum honeycomb core designed to crush under impact to absorb downward loads. The Avidyne Entegra integrated aircraft instrumentation system comprised a primary flight display (PFD) and multi-function display (MFD).

The flight controls for ailerons, elevator, and rudder were conventional in design. The control surfaces were pilot-controlled through either of two single-handed side-control yokes mounted beneath the instrument panel. Roll and pitch trim were available through an electric button on the top of each side-control yoke. The yaw trim system employed a ground-adjustable trim tab. Neutral rudder position was held by a ground-adjustable spring cartridge that was bolted to the left rudder pedal torque tube and center console assembly, which provided a centering force regardless of the direction of control surface deflection.

METEOROLOGICAL INFORMATION

The recorded weather conditions at MMK at 1833 included wind from 180° at 5 knots, 10 statute miles visibility, few clouds at 300 ft, an overcast ceiling at 12,000 ft, temperature 16°C, dew point 2°C, and an altimeter setting of 30.15 inches of mercury.



AIRPORT INFORMATION

According to FAA Chart Supplements, MMK was owned by the City of Meriden, Connecticut, and was classified by the FAA as a non-towered, public use airport. The airport elevation was 103 ft mean sea level and there was one runway oriented in a 18/36 configuration. Runway 18 was asphalt and was in good condition; it measured 3,100 ft long by 75 ft wide.

WRECKAGE AND IMPACT INFORMATION

Runway Examination

Examination of runway 18 revealed black tire marks in an S-shaped (sinusoidal) pattern co-located with white paint transfer marks on the surface of the runway pavement. The tire marks and paint transfer marks were discovered in two locations about 1,350 ft from the beginning of runway 18. Both the tire marks and paint transfer marks were consistent with nose wheel shimmy and nose wheel pant contact.

Accident Site Examination

Examination of the accident site revealed that the airplane first made ground contact with the left wingtip. After cartwheeling and subsequently impacting and breaching a 30-ft section of the 8-ft-tall airport security fence, the airplane slid along a public roadway on an approximate 078° magnetic heading. About 115 ft from the initial impact point, the airplane came to rest in the northbound travel lane against an earthen berm. Most of the airplane was then consumed by a postcrash fire.

A 115-ft-long and 62-ft-wide debris path extended from the initial impact point to the main wreckage. It contained the propeller, which was found buried beneath the shoulder of the southbound travel lane about 37 ft from the initial impact point; the engine cowling, which came to rest about 52 ft from the initial impact point; the left wing tip and a portion of the outer left wing panel, which came to rest about 81 ft from the initial impact point; and the top rail of the breached section of airport security fence, which came to rest about 92 ft from the initial impact point. It also contained smaller components of the airplane and portions of the airplane structure.



Airplane Examination

Examination of the airplane revealed no evidence of any preimpact failure or malfunction of the airplane or flight controls.

The fuselage came to rest upright and was mostly consumed by fire. The empennage was separated from the aft fuselage, inverted, and displayed impact and fire damage.

The outboard section of the left wing and the left wing tip separated during the impact sequence. The remaining portion of the left wing remained in its mounting location and exhibited impact and fire damage. The left aileron was almost completely consumed by fire. Pooled aluminum was located on the ground aft of the wing at the mounting location of the left wing flap along with the remains of a flap hinge.

The right wing exhibited impact and fire damage. The inboard third of the right aileron was consumed by fire. Pooled aluminum was located on the ground aft of the wing at the mounting location of the right wing flap along with the remains of a flap hinge.

Aileron control cable continuity was verified from the remains of the cabin to the ailerons. The flap actuator was fully extended, consistent with the wing flaps in the retracted position.

The horizontal stabilizer remained attached to the empennage and exhibited impact and fire damage. The right elevator was mostly consumed by fire, with the outboard portion and elevator tip still present. The left elevator was mostly consumed by fire, with the leading edge and tip still present. Elevator control cable continuity was verified from the remains of the cabin to the elevators.

The vertical stabilizer was impact and fire damaged and remained attached to the empennage. The rudder also exhibited impact and fire damage. Rudder control cable continuity was verified from the remains of the cabin to the rudder. The pitch trim motor position could not be determined due to fire damage.



Propeller Examination

Examination of the three-bladed propeller revealed no evidence of any preimpact malfunction or failure.

The propeller remained attached to the propeller flange, but separated from the crankshaft, which fractured just aft of the propeller flange. The crankshaft fracture surface displayed 45° shear angles and a cupped appearance with blue-black discoloration in a smeared area. All three blades remained attached to the propeller hub; however, one propeller blade tip separated during the impact sequence and one propeller blade rotated 180° in the propeller hub. All three propeller blades exhibited chordwise scratching and leading-edge gouging; the gouges matched the spacing of the chain links of the airport security fence. The propeller governor remained secured to the front left side of the engine and the propeller control cable remained secured to the propeller control lever.



Engine Examination

Examination of the engine revealed no evidence of any preimpact failure or malfunction of the engine.

The engine had remained attached to the firewall via the engine control cables, the main fuel line, and the electrical wires and cables. There were no pre-accident anomalies noted with the induction system. The exhaust system components remained attached to the engine with no signs of pre-accident anomalies noted. The exhaust mufflers and shrouds sustained deformation damage.

The ignition harness remained secured to each magneto and each terminal remained secured to its respective spark plug. During crankshaft rotation, and audible snap and spark was observed from both magnetos. No pre-accident anomalies were noted with either of the magnetos.

The spark plugs had remained secured to their respective cylinders. The top spark plugs were removed and displayed normal wear with lean operation signatures and no signs of carbon or lead fouling. The bottom spark plugs were observed during the borescope inspection with no signs of lead or carbon fouling noted.

The engine-driven fuel pump remained secured to the backside of the engine Manual rotation of the drive coupling resulted in rotation of the drive shaft. No pre-accident anomalies were noted with the internal components.

The fuel lines to and from the throttle body/fuel metering unit remained secured and fuel was observed in the lines between the fuel flow transducer and the fuel metering unit. Manual rotation of the throttle lever resulted in a coinciding rotation of the drive shaft. No pre-accident anomalies were noted with the unit.

The fuel manifold valve remained secured to the engine. All fuel injection lines remained secured to the manifold valve body and the torque putty was intact. Fuel was noted within the manifold valve. The screen was not obstructed. The diaphragm remained intact and pliable and was still attached to the plunger. No pre-accident anomalies were noted with the unit.

The injector lines remained secured to the nozzles. Each nozzle was removed and inspected and no obstructions were noted. Light was visible through each nozzle jet, except for the No. 1 nozzle jet, due to bending damage.

The oil sump sustained impact deformation and puncture damage. Oil was observed leaking from the oil sump and the oil pump remained secured to the backside of the engine. The oil filter sustained thermal damage and was dented, and the internal components were charred. The oil cooler remained secured on the back left side of the engine. There were no signs of lubrication distress on the observed engine components, and no pre-accident anomalies were noted.

All six cylinders remained attached to the engine and borescope examination of the internal components revealed no preaccident anomalies. All six cylinders also produced thumb compression and suction during rotation of the drivetrain, and valve functionality was confirmed.

The crankcase remained intact with no external signs of operational distress. There were no pre-accident anomalies noted with the crankcase.

The crankshaft was fractured aft of the propeller flange. Crankshaft continuity was confirmed to the front and out to each connecting rod during manual rotation from the accessory end. No pre-accident anomalies were noted.

Camshaft continuity was confirmed during manual rotation of the upper right accessory drive gear. The rockers and valve springs functioned during the continuity test and no pre-accident anomalies were noted. The Nos. 5 and 6 pushrods displayed impact-related deformation damage.




MEDICAL AND PATHOLOGICAL

The Office of the Chief Medical Examiner, Farmington, Connecticut, performed an autopsy on the pilot. The cause of death was blunt injuries of head and trunk with fractures and aortic laceration.

The FAA Forensic Sciences Laboratory conducted toxicological testing on specimens from the pilot. The toxicological testing results for the pilot were negative for carbon monoxide, and ethanol. Acetaminophen, a common over-the-counter analgesic/antipyretic, was detected in urine; it is not impairing.

ADDITIONAL INFORMATION

Cirrus Aircraft Guidance

According to the Cirrus Design SR22 Pilot Operating Handbook and Airplane Flight Manual (Section 4, Normal Procedures);

Normal landings are made with full flaps with power on or off. Surface winds and air turbulence are usually the primary factors in determining the most comfortable approach speeds. Actual touchdown should be made with power off and on the main wheels first to reduce the landing speed and subsequent need for braking. Gently lower the nose wheel to the runway after airplane speed has diminished. This is especially important for rough or soft field landings.

In a balked landing (go around) climb, disengage autopilot, apply full power, then reduce the flap setting to 50%. If obstacles must be cleared during the go around, climb at 75-80 KIAS with 50% flaps. After clearing any obstacles, retract the flaps and accelerate to the normal flaps up climb speed.

FAA Guidance

According to the Airplane Flying Handbook (FAA-H-8083-3B):

A stabilized descent angle is controlled throughout the approach so that the airplane lands in the center of the first third of the runway…The objective of a good, stabilized final approach is to descend at an angle and airspeed that permits the airplane to reach the desired touchdown point at an airspeed that results in minimum floating just before touchdown; in essence, a semi-stalled condition. To accomplish this, it is essential that both the descent angle and the airspeed be accurately controlled.

Regarding bouncing on touchdown, the handbook states:

When a bounce is severe, the safest procedure is to execute a go-around immediately. An attempt should not be made to salvage the landing. Full power should be applied while simultaneously maintaining directional control and lowering the nose to a safe climb attitude.

The handbook also states that, whenever landing conditions are not satisfactory, a go-around is warranted:

The assumption that an aborted landing is invariably the consequence of a poor approach, which in turn is due to insufficient experience or skill, is a fallacy. The go-around is not strictly an emergency procedure. It is a normal maneuver that is also used in an emergency situation….The earlier a condition that warrants a go-around is recognized, the safer the go-round/rejected landing is. The go-around maneuver is not inherently dangerous in itself. It becomes dangerous only when delayed unduly or executed improperly.

Attitude is always critical when close to the ground, and when power is added, a deliberate effort on the part of the pilot is required to keep the nose from pitching up prematurely. The airplane executing a go-around must be maintained in an attitude that permits a buildup of airspeed well beyond the stall point before any effort is made to gain altitude or to execute a turn. Raising the nose too early could result in a stall from which the airplane could not be recovered if the go-around is performed at a low altitude.

A concern for quickly regaining altitude during a go-around produces a natural tendency to pull the nose up. A pilot executing a go-around must accept the fact that an airplane cannot climb until it can fly, and it cannot fly below stall speed. In some circumstances, it is desirable to lower the nose briefly to gain airspeed. As soon as the appropriate climb airspeed and pitch attitude are attained, "rough trim" the airplane to relieve any adverse control pressures. More precise trim adjustments can be made when flight conditions have stabilized. 



Pilot Information

Certificate: Private
Age: 56, Male
Airplane Rating(s): Single-engine Land
Seat Occupied: Left
Other Aircraft Rating(s): None
Restraint Used: 4-point
Instrument Rating(s): Airplane
Second Pilot Present: No
Instructor Rating(s): None
Toxicology Performed: Yes
Medical Certification: Class 3 Without Waivers/Limitations
Last FAA Medical Exam: 02/01/2017
Occupational Pilot: No
Last Flight Review or Equivalent: 10/30/2014
Flight Time:  (Estimated) 1217.1 hours (Total, all aircraft), 10.2 hours (Total, this make and model), 1139.4 hours (Pilot In Command, all aircraft), 10.2 hours (Last 90 days, all aircraft), 8.4 hours (Last 30 days, all aircraft), 1.8 hours (Last 24 hours, all aircraft)

Aircraft and Owner/Operator Information

Aircraft Make: CIRRUS DESIGN CORP
Registration: N94LP
Model/Series: SR22
Aircraft Category: Airplane
Year of Manufacture:
Amateur Built: No
Airworthiness Certificate: Normal
Serial Number: 1484
Landing Gear Type: Tricycle
Seats: 4
Date/Type of Last Inspection: 03/13/2017, Annual
Certified Max Gross Wt.: 3400 lbs
Time Since Last Inspection:
Engines: 1 Reciprocating
Airframe Total Time: 1229 Hours as of last inspection
Engine Manufacturer: CONT MOTOR
ELT: C91A installed, not activated
Engine Model/Series: IO-550-N27B
Registered Owner: On file
Rated Power: 310 hp
Operator: On file
Operating Certificate(s) Held: None

Meteorological Information and Flight Plan

Conditions at Accident Site: Visual Conditions
Condition of Light: Day
Observation Facility, Elevation: MMK, 103 ft msl
Distance from Accident Site: 0 Nautical Miles
Observation Time: 1833 EDT
Direction from Accident Site: 283°
Lowest Cloud Condition: Few / 300 ft agl
Visibility:  10 Miles
Lowest Ceiling: Overcast / 12000 ft agl
Visibility (RVR):
Wind Speed/Gusts: 5 knots /
Turbulence Type Forecast/Actual: / None
Wind Direction: 180°
Turbulence Severity Forecast/Actual: / N/A
Altimeter Setting: 30.15 inches Hg
Temperature/Dew Point: 16°C / 2°C
Precipitation and Obscuration: No Obscuration; No Precipitation
Departure Point: Wallingford, CT (MMK)
Type of Flight Plan Filed:None 
Destination: Wallingford, CT (MMK)
Type of Clearance: None
Departure Time: 1740 EDT
Type of Airspace: Class G 

Airport Information

Airport: MERIDEN MARKHAM MUNI (MMK)
Runway Surface Type: Asphalt
Airport Elevation: 103 ft
Runway Surface Condition: Dry
Runway Used: 18
IFR Approach: None
Runway Length/Width: 3100 ft / 75 ft
VFR Approach/Landing: Go Around; Traffic Pattern

Wreckage and Impact Information

Crew Injuries: 1 Fatal
Aircraft Damage: Destroyed
Passenger Injuries: 1 Serious
Aircraft Fire: On-Ground
Ground Injuries:N/A 
Aircraft Explosion: None
Total Injuries: 1 Fatal, 1 Serious
Latitude, Longitude: 41.508333, -72.827222