N6677DBeech A36 · 1979

On July 25, 2015, about 1040 eastern daylight time, a Beech A36; N6677D, was substantially damaged during an emergency landing, after a partial loss of power during takeoff at Duchess County Airport (POU), Poughkeepsie, New York. The private pilot received minor injuries, and the passenger was uninjured. Visual meteorological conditions prevailed, and an IFR flight plan was filed for the flight, conducted under the provisions of Title 14 Code of Federal Regulations Part 91, destined for Burlington National Airport (BTV), Burlington, Vermont.

According to the pilot, after arriving at POU from Lehigh Valley International Airport (ABE), Allentown, Pennsylvania, he purchased 15 gallons of fuel and then sumped the tanks in preparation for the next leg of the flight to BTV.

After starting the engine he taxied out, did his engine runup, checked the magnetos, and cycled the propeller. Then during the takeoff from runway 24, he noticed a vibration and unusual noise as he rotated. He believed that it may have been from the wheels and applied the brakes to stop the wheels from rotating and retracted the landing gear but, the vibration continued. He was however unable to continue climbing, as the engine suddenly incurred a partial loss of power.

He checked that the propeller, throttle, and mixture, was full forward but the airplane still would not climb. With the airspeed being low, he knew that he could not make it back to the airport without stalling the airplane.

There were "trees everywhere" and a set of power lines directly ahead of him. He then pulled back on the control wheel, was able to clear the power lines, and then "forced the nose down" to prevent the airplane from stalling, and landed gear up about 300 feet from the power lines.

According to Federal Aviation Administration (FAA) and pilot records, the pilot held a private pilot certificate with ratings for airplane single-engine land, airplane single engine sea, and instrument airplane. His most recent application for a FAA third-class medical certificate was dated August 5, 2014. The pilot reported that he had accrued approximately 8,954 total hours of flight experience, of which 6,311 hours were in the accident airplane make and model.

The accident airplane was a low wing, retractable single engine airplane, of conventional metal construction, equipped with retractable tricycle type landing gear. It was powered by an air cooled, 6-cylinder, horizontally opposed, 285 horsepower Continental IO-520-BB engine driving a McCauley 3-bladed, variable pitch, constant speed propeller.

According to FAA and maintenance records, the airplane was manufactured in 1979. The airplane's most recent annual inspection was completed on December 17, 2014. At the time of the inspection, the airplane had accrued 6,796.6 total hours of operation.

The recorded weather at POU, at 1047, about 7 minutes after the accident, included: calm winds, 10 miles visibility, clear skies, temperature 25° C, dew point 13° C, and an altimeter setting of 30.01 inches of mercury.

The accident airplane was a low wing, retractable single engine airplane, of conventional metal construction, equipped with retractable tricycle type landing gear. It was powered by an air cooled, 6-cylinder, horizontally opposed, 285 horsepower Continental IO-520-BB engine driving a McCauley 3-bladed, variable pitch, constant speed propeller.

According to FAA and maintenance records, the airplane was manufactured in 1979. The airplane's most recent annual inspection was completed on December 17, 2014. At the time of the inspection, the airplane had accrued 6,796.6 total hours of operation.

Examination of the accident site revealed that the airplane had initially touched down on its belly in a grassy area, then traveled over the top of a large rock and came to rest.

Examination of the airplane revealed that it had touched down with the landing gear in the up (stowed) position. All major parts of the airplane were on site, and the airplane had been substantially damaged when it struck and traveled over the rock.

The outboard sections of all three blades of the propeller had been bent backwards, the right side nose wheel door had separated at the hinge line, the aft portion of the lower cowling had been buckled and crushed, the upper cowling had partially opened on impact.

The nose landing gear wheel well had been crushed upward, and both right side engine mounts were fractured. The throttle body had one of the mount lugs fractured and two of the fuel fittings were fractured.

The lower portion of the firewall was buckled, and the fuselage had buckled just aft of the firewall. The wing flaps were in the 15°position, and the inboard portion of the left wing flap was bent. Fuel was observable in both the left, and right, wing tanks.

According to the auxiliary pump manufacturer, fittings and O-rings can be a source of leakage. Care should be taken to ensure that fittings are inspected for cracks, scratches and cross threaded threads. New O-Rings should be installed at the time of fuel pump replacement and if the O-Rings have been on the fittings more than 10 years.

Review of Maintenance Records

Review of maintenance records and the Pilot/Operator Accident/Incident Report submitted by the pilot, indicated that the factory rebuilt engine was installed on the airplane on February 23, 2012, at a recorded tachometer time of 6,103.9 hours.

On June 14, 2013, at 279.1 hours after engine installation, maintenance personnel replaced the airplane's auxiliary fuel pump (fuel boost pump).

On February 6, 2014, (tachometer and engine time not specified), maintenance personnel located in Fort Pierce, Florida, "bled fuel line, checked fuel strainer, cleaned injectors due to rough running engine." The engine was test run and returned to service.

On February 7, 2014, at an unspecified tachometer and engine time, maintenance personnel located in Stuart, Florida, "removed all spark plugs, cleaned, gapped, and rotated upon reinstallation, removed all fuel injectors cleaned inspected and reinstalled, [and performed] aircraft operational and functional check" with no discrepancies noted.

On December 17, 2014, at a recorded tachometer time of 6,796.6 hours, the engine underwent an annual inspection. On December 18, 2014 (same tachometer reading as annual inspection), the spark plugs were removed, cleaned, and reinstalled, and the ignition leads, and magnetos were checked with no anomalies noted.

On May 11, 2015, at a tachometer time of 6,881 hours, the engine underwent an oil and oil filter change.

According to the pilot, the engine had accumulated 790 hours at the time of the accident.

Engine Test Run

On November 2, 2015, in order to help determine why in the maintenance records the engine had been reported to have been running rough and why the loss of engine power had occurred, an engine test run was performed.

Prior to the test run several airframe related items were removed in preparation for operation in the test cell. The removed items included:

- Both fractured Engine mounts (right front and right rear)

- Fuel fittings (the throttle body outlet and mixture return)

- Cooling Baffles

- Propeller Governor

The following substitute or repaired parts were then installed for engine operation:

- Engine Mount legs

- Fuel fittings

The magneto-to-engine timing which was specified to be 22° before top dead center (BTDC) was then checked, with the following results:

- Left Magneto (21°BTDC)

- Right Magneto (20°BTDC)

The engine was not disassembled prior to the engine run. The crankshaft end-play was measured 0.010" and the run-out was 0.001".

The engine was prepared for operation by installing the appropriate thermocouples, pressure lines and test pads for monitoring purposes. The engine was then moved to a test cell, mounted for operation, and then fitted with a club type propeller for testing.

The engine experienced a normal start on the first attempt without hesitation or stumbling in observed RPM. The engine RPM was advanced in steps for warm-up in preparation for full power operation. The engine throttle was advanced to 1200 RPM and held for five (5) minutes to stabilize. The engine throttle was advanced to 1600 RPM and held for five (5) minutes to stabilize. The engine throttle was advanced to 2450 RPM and held for five (5) minutes to stabilize. The engine throttle was advanced to the fully open position and held for five (5) minutes to stabilize. The engine throttle was rapidly advanced from idle to full throttle five times where it performed normally without any hesitation, stumbling or interruption in power.

Throughout the test phase, the engine accelerated normally without any hesitation, stumbling or interruption in power and demonstrated the ability to produce rated horsepower.

After the test run with the engine still hot, a cylinder leakage test was performed in accordance with the latest revision of CMI Service Bulletin SB03-3 with the following results (master orifice reading – 43 PSI):

- Cylinder No. 1 - 17/80 PSI (rings)

- Cylinder No. 3 - 56/80 PSI (rings)

- Cylinder No. 5 - 60/80 PSI (rings)

- Cylinder No. 2 - 72/80 PSI (rings)

- Cylinder No. 4 - 36/80 PSI (rings)

- Cylinder No. 6 - 60/80 PSI (rings)

(*) – Leakage Source

Airplane Fuel System

Airplane fuel systems are designed to provide an uninterrupted flow of clean fuel from the fuel tanks to the engine. The fuel must be available to the engine under all conditions of engine power, altitude, attitude, and during all approved flight maneuvers. Two common classifications apply to fuel systems: gravity-feed and fuel-pump systems.

Low- and mid-wing single reciprocating engine airplanes cannot utilize gravity-feed fuel systems because the fuel tanks are not located above the engine. Instead, one or more pumps are used to move the fuel from the tanks to the engine.

In a low wing airplane, with a fuel injection system such as the Continental system, fuel pressurized by an engine-driven pump is metered as a function of engine rpm. It is first delivered from the fuel tanks (one for each wing), To a three-way selector valve (LEFT, RIGHT, or OFF). The selector valve also acts simultaneously as a diverter of air that has been separated out of the fuel in the engine-driven fuel pump and returned to the valve. It routes the air to the vent space above the fuel in the selected tank.

An electric auxiliary fuel pump draws fuel through the selector valve. It forces the fuel through the strainer, making it available for the engine-driven fuel pump. The electric auxiliary pump also supplies fuel pressure while starting, is used to prevent vapor lock, and is also used as a backup should the engine-driven pump fail and does not need to be operating to allow the engine-driven fuel pump access to the fuel.

The engine driven pump supplies a higher-than needed volume of fuel under pressure to the fuel control. Excess fuel is returned to the pump, which pumps it through the selector valve into the appropriate tank. Fuel vapor is also returned to tanks by the pump. The fuel control unit meters the fuel according to engine rpm and mixture control inputs from the cockpit and then supplies it to the fuel manifold and injectors, which spray fuel without any air mixed in directly into the cylinders, to provide a measured, continuous spray and smooth engine operation.

Auxiliary Fuel Pump

The auxiliary fuel pump was controlled by an ON-OFF toggle switch on the control console. It provided pressure for starting and emergency operation. Immediately after starting, the auxiliary fuel pump could be used to purge the system of vapor caused by extremally high ambient temperature or start with the engine hot. The auxiliary fuel pump provided for near maximum engine performance should the engine driven pump fail.

On May 4, 2016, the auxiliary fuel pump was tested.

It was noted that the fuel pump was intact and had not been disassembled from the time of manufacture.

During the testing, it took 5-7 seconds for the pump to self-prime and to start pumping.

Fuel pump requirements were a minimum of 42 gph at 16 psi, with a maximum amperage draw of 3 amps at 28 volts dc.

The pump during testing produced 53 gph at 16 psi, and amperage draw was 2.1 amps at 28 volts dc. indicating that the pump was operating within specifications.

During the testing however, it was discovered that the discharge fitting was leaking. Disassembly of the discharge fitting revealed that it appeared that it had had been installed to the proper torque, but it was observed that there were deep scratches on the O-Ring boss on the pump side of the fitting, and a piece of metal was found imbedded in the O-Ring.

The fittings were then removed, and the pump was then retested with another set of fittings. No leaks were found with the pump and /or the replacement fittings.