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N186JC accident description

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Tail numberN186JC
Accident dateJune 19, 1999
Aircraft typeCanadair F-86E
LocationMount Holly, NJ
Additional details: None

NTSB description


On June 19, 1999, about 1340 eastern daylight time, a Canadair F-86E Sabre on an experimental certificate, N186JC, was destroyed when it impacted terrain while conducting a low pass at the South Jersey Regional Airport, Mount Holly, New Jersey. The certificated commercial pilot was fatally injured. Visual meteorological conditions prevailed, and no flight plan was filed for the personal flight conducted under 14 CFR Part 91.

A witness, approximately 2,000 feet from the accident site, saw the airplane depart, and perform several passes before attempting a "slow" pass to the west at 200 feet agl. As the airplane passed in front of the witness, it slowed, and started to sink. The airplane's pitch attitude increased, and the witness heard an increase in power, followed by a bang. He described the bang as consistent with a "compressor stall."

Examination of photographs taken of the accident sequence, showed the airplane's landing gear extended, speed-brakes deployed, and flaps down. In the first photograph, the airplane was nose up approximately 15 degrees, and wings level. In the second photograph, the airplane was approximately 30 feet agl, banking right approximately 90 degrees, and nose down about 10 degrees. In the third photograph, the right wing and nose of the airplane had impacted the ground. During the impact, the airplane was in a right bank approximately 120 degrees, and nose down approximately 45 degrees.

The accident happened during the hours of daylight. The wreckage was located 39 degrees, 56.52 minutes north latitude, 74 degrees, 50.93 minutes west longitude, and about 45 feet elevation.


The pilot held a commercial pilot certificate with ratings for airplane single engine land, airplane multi engine land, and airplane instrument. The pilot's logbook was not recovered. According to the Pilot/Operator Report of Aircraft Accident form submitted by the owner of the airplane, the pilot had a total of approximately 7,000 hours of flight experience, with 50 hours in the accident airplane's make and model. In addition, the pilot completed a biennial flight review on July 1, 1997.


On April 10, 1999, the pilot sold the airplane to a corporation in Washington, District of Columbia. Per a pre-sale agreement, the pilot maintained possession of the airplane, and was responsible for maintaining it.

The airplane's maintenance records were not recovered at the accident site, and the owner's representative could not locate them. In addition, the maintenance records for two other airplanes under the accident pilot's control could not be located.


The pitot tube from the airplane's right wing was located at the start of the debris path. The path was orientated on a magnetic heading of 288 degrees, and started approximately 25 feet to the right of runway 26, and continued to the main wreckage. At its origin, the path measured approximately 2 feet wide, and progressively got wider. At its terminus, it was approximately 60 feet wide.

The main wreckage was located approximately two-thirds the way down, and 300 feet to the right of runway 26. Oriented northeast was the fuselage laying on its left side. The first one-third of the fuselage was destroyed by impact, and difficult to recognize. The left wing was separated from the fuselage, and approximately 15 feet to the northwest. The carry-though structure with the right wing attached had also separated from the fuselage. This section was northwest of the fuselage with the carry-though partially under the fuselage. The nose gear assembly had separated from the fuselage, and was located approximately 60 feet to the southeast. The right horizontal stabilator was attached. The left horizontal stabilator had separated about 3 feet outboard of the fuselage, and was located approximately 30 feet to the southeast. The right horizontal stabilator was attached. The vertical stabilizer was attached. All flight control surfaces were accounted for, and the alternate flight control hydraulic accumulator indicated 2,500 psi.

On June 20, 1999, the wreckage was recovered to a hangar located at the airport for further examination. During the recovery, approximately 20 gallons of fuel was observed leaking from the wreckage.

The right wing's entire leading edge was fragmented, and compressed in approximately 14 inches. The right main landing gear was attached, and in the down position. When a rotational force was applied to the wheel, it rotated freely. The right wing drop-tank was attached, and had ruptured. The right flap was in the down position, and approximately 1 foot of the outboard end was torn aft. The right aileron was attached, and in the neutral position. Approximately 16 inches of the inboard section, and approximately 12 inches of the outboard section of the aileron was damaged. The right wing-tip was torn and fragmented. Damage on the wing-tip covered approximately 120 square inches, and was more severe towards the leading edge.

The left wing's entire leading edge was fragmented and compressed in approximately 11 inches. The left main landing gear was attached, and in the down position. When a rotational force was applied to the wheel, it rotated freely. The left wing drop-tank had separated, and ruptured. The left flap was in the down position, and approximately 14 inches of the outboard section was bent down 90 degrees. The left aileron was attached, and in the neutral position. The outboard-half of the aileron displayed more damage then the inboard-half, and the trailing edge displayed more damage then the leading edge. The skin of the aileron was deformed, but not fragmented. The left wing-tip was intact, and displayed random damage from its leading edge to its trailing edge.

On the empennage, the right speed-brake had separated. The left speed-brake was partially attached. Fracture surfaces on the left horizontal stabilator displayed a 45 degree shear lip, and were consistent with overload. The top 14 inches of the vertical stabilator was crushed, and bent to the right 45 degrees.


An autopsy was preformed on the pilot, on June 20, 1999, at the Medical Examiner's Office for Burlington County, New Jersey.

A toxicological test was performed on the pilot by the Federal Aviation Administrations Toxicology and Accident Research Laboratory, Oklahoma City, Oklahoma.


On June 24, 1999, the engine was transported to a maintenance facility in Millville, New Jersey, for further examination. The engine's compressor section was examined via a borescope. Damaged was isolated to the first two stages. Approximately one-third of the blades in this area had nicks measuring approximately 0.020 of an inch deep on their leading edge. Approximately one-third of the stator vanes in this area had damage to their trailing edge. The damage consisted of nicks, and gouges ranging from 0.020 to 0.200 of an inch deep. No compressor blades or stator vanes were identified as missing.

Examination of the turbine section by borescope revealed no damage, and no turbine blades or turbine nozzles were identifies as missing. Using two fingers, a rotational force was applied to the aft turbine wheel, and it rotated freely for 350 degrees. Pass the 350 degree mark, resistance was felt consistent with rubbing, and additional force was required. Also, while rotating the turbine section, the compressor section rotated and the female drive splines to the engine driven fuel pumps rotated.

The aft electric driven fuel boost pump was stamped February, 1952, and had a rubber cure date of November, 1956. With the fuel pump submerged, 27 volts D.C. was applied. The electric motor spun, and water was expelled via the outlet. The forward electric driven fuel boost pump was stamped August 1953, and had a rubber cure date of November, 1956. With the fuel pump submerged, 27 volts D.C. was also applied. The electric motor spun, and water was expelled via the outlet. Both the left and right engine driven fuel pumps were rotated by hand, and residue fuel was expelled via their outlets. When 27 volts D.C. was applied to the backup electric hydraulic pump, the motor spun, and residual hydraulic fluid was expelled via the outlet. The aft fuselage pump, and fuel level transmitter were not examined because impact forces prevent the engine from being separated from the fuselage. In addition, extensive use of a cutting torch would have been required in close proximity to the aft fuel cell.


According to a pilot who flew the accident airplane on the airshow circuit, in April of 1999, he experienced a flame-out at altitude do to fuel starvation. he executed a successful air-start, and landed the airplane without incident. At first, it was thought the fuel control may have caused the loss of power, but the idea was dismissed after conducting extensive reach, and not finding any documented failures of the type of fuel control unit being used. Ultimately, the right electric boost pump, and the left electric boost pump were identified as inoperative and changed. In addition, both engine driven fuel pumps were replaced. After the incident, a procedure was developed to test the operation of the boost pumps, and the engine drive pumps to insure proper operation prior to each flight.

The airshow pilot added that during a 12 minute routine the airplane would consume between 1,000 and 1,200 pounds of fuel with a power setting of 90 to 95 percent N1, or 83.3 to 100 pounds per minute. He estimated that the accident pilot used a similar power setting. He also stated that the engine was not susceptible to compressor stalls at low altitudes, but a rapid increasing in throttle could cause the compressor section to "hang".

Examination of the fuel quantity indicator by the Safety Board's Material Laboratory revealed a pair of parallel lines on the painted surface of the dial. These lines were consistent with needle impact, and had a radial direction from the center of the indicator towards the 1,500 pound line.

According to a witness, the airplane had approximately 1,360 pounds of fuel onboard prior to engine start. In addition, a lineman stated that the pilot requested 340 pounds of additional fuel, but then changed his mind.

According to an F-86 mechanic, a pilot would have no indication that the aft fuselage pump or fuel level transmitter had failed. He added that the failure of the pump is usually detected when the airplane is weighed for weight and balance calculations.

A warning in the Aircraft Operating Instructions stated that, "A possibility of fuel starvation and subsequent engine flameout will exist when failure of the aft fuselage pump or fuel level transmitter occurs with the aircraft in a climbing attitude at low fuel state i.e., below 500 lbs." In addition, "A considerable amount of fuel can be trapped in the aft fuselage cell under these conditions. Where flight is necessary with low fuel state pilot should avoid nose high attitudes. At level or slightly nose down attitudes all the fuel from the aft cell will be available."

In addition, the Aircraft Operating Instructions stated that "When the airflow through the compressor becomes less than required for a given rpm, the resultant relative airflow to the compressor blades is then above the angle of attack for stall, and the individual blades stall in the same manner...a wing does. This is compressor stall. Compressor stall can be initiated by a rapid throttle movements during unstabilized compressor inlet conditions such as...during slipping or skidding maneuvers or low airspeed high angle of attack conditions, particularly at high altitudes. If throttle movement is required during these conditions, a steady progressive movement of 3 to 4 seconds will produce the same rate of acceleration with less danger of initiating a stall."

On July 5, 1999 the wreckage was released to the insurance company minus the fuel quantity gauge. That item was then released on October 5, 1999.

(c) 2009-2011 Lee C. Baker. For informational purposes only.