SEA02LA116

The pilot reported that he experienced light to moderate turbulence approximately 10 minutes prior to the accident while cruising at an airspeed of about 150 mph. The pilot initiated a 1,000 foot per minute descent about 15 nautical miles northwest of the airport, and stated that his indicated airspeed was between 170 and 180 mph. During the descent, the pilot stated that the aircraft began to shake violently. The pilot applied equal pressure to the rudder pedals, pulled out the throttle, pitched up the nose to slow down, and the shaking stopped. The pilot continued the descent for landing without further incident. During the inspection of the airframe, wrinkles, creases and compression signatures were found in the skin to both the left and right sides of the empennage from fuselage stations (FS) 233.5 to FS 272 were noted. The airframe skin pulled away from five rivets located on the bottom of the empennage, along the left side, and forward of FS 256.9. The skin from left to right side pulled away from all of the empennage bottom rivets at FS 256.9. The skin pulled away from six of the rivets along the right side, and forward of FS 256.9. The two longitudinal running rivet rows on the bottom had the skin pulled away from seven rivets on the left side and six rivets on the right side. All applicable AD's had been complied with at the time. No AD speed restrictions were in effect for this model aircraft. Normal operating range for this model aircraft is between 66 - 175 mph.

History of flight

On June 21, 2002, approximately 1300 Pacific daylight time, a Beech G35, N4647D, experienced an in-flight airframe shake while descending to McAllister Field, Yakima, Washington. The aircraft is owned and operated by a private pilot. Both the pilot and his passenger were not injured; however, the aircraft received substantial damage. The 14 CFR Part 91 personal pleasure flight departed Skagit Regional Airport, Mount Vernon, Washington, approximately one hour and 30 minutes before the accident, and was being operated in visual meteorological conditions. A visual flight rules flight plan had been filed and activated.

During a telephone interview and subsequent written statement, the pilot reported that he experienced light to moderate turbulence approximately 10 minutes prior to the accident, while cruising at an airspeed of about 150 mph. He initiated a 1000 foot per minute descent about 15 nautical miles northwest of McAllister Field, and stated that his indicated airspeed was between 170 and 180 mph. During the descent, "We crossed the last ridge and had the field in sight when the aircraft began to shake violently. I applied equal pressure to the rudder pedals, pulled out the throttle, pitched up the nose to slow down, and the shaking stopped." The pilot continued the descent for landing at McAllister Field without further incident.

Aircraft information

The airplane is a 1956, model "v-tail" Beech G35, serial number D-4803. The airframe total time was approximately 5,971 hours, and the tachometer was approximately 1,419 hours at the time of the accident. The engine on the accident aircraft was a Continental E-225-8 engine, serial number 31385-D-6-8. The aircraft was equipped with a Hartzell HC-A2MV20-4A1 propeller.

A review of the aircraft logbooks and historical review of accidents/incidents revealed that the aircraft was involved in three separate gear up landings in 1963, 1973, and 1979.

Logbook entries revealed the following:

On July 3, 1984 - The aircraft was stripped and repainted. The right aileron and ruddervator were replaced with new controls and were balanced.

On June 29, 1987, Beech Service Bulletin #2188 was accomplished by inspection per Beech Instruction #35-4017 and installation of Beech Kit #35-4016-3S, S/N 125. Doublers and cuffs were installed, trim cables were replaced, the trip stop was reset, and placards were installed per drawing #35-4016. The rudder and elevator cable tensions were adjusted per the Beech Maintenance Manual.

On November 16, 1995, the aileron and ruddervator were removed, balanced and reinstalled.

On December 14, 1995, AD 94-20-04, V-tail inspection was accomplished to prevent structural failure of the V-tail, which could result in loss of control of the airplane. No cracks or structural deformation were noted.

On April 1, 1998, AD 97-06-11, Tail Control Rod Assembly was accomplished to prevent failure of the ruddervator differential tail control rod assembly, which could result in loss of control.

On May 12, 1999, during the annual inspection it was noted that AD 98-13-02 speed restrictions to prevent in-flight vibrations caused by the affected airplanes operating at excessive speed, which could result in airplane damage and possible loss of control of the airplane did not apply to the Beech G35.

On August 2, 2001, during the annual inspection, AD 89-05-02, to prevent the failure of the magnesium elevator control fitting, and AD 94-20-04, to prevent structural failure of the V-tail, which could result in loss of control of the airplane were complied with. A visual inspection of the bulkheads at fuselage station (FS) 256.9 and FS 272 found no defects at this time.

On December 21, 2001, during a pre-purchase inspection, the panels from the wings and tail were removed for an internal inspection for corrosion and defects. None were noted.

Meteorological information

The surface observation for Yakima at 1256 reported clear sky conditions with winds variable at three knots. The temperature was 30 degrees Celsius. The pilot reported light to moderate turbulence over the mountains.

Damage to aircraft

On July 17, 2002, investigators from the National Transportation Safety Board, Federal Aviation Administration and Raytheon Aircraft met in Yakima to inspect the aircraft. The inspection noted wrinkles in the skin to both the left and right sides of the rear fuselage from fuselage stations (FS) 233.5 to station 272.

The lower left side of the rear fuselage at FS 233.5 noted a shallow recessed dimple in an area about three inches in diameter. A diagonal approximate four inch crease traveled aft and upward from this dimple. At the upper area between FS 233.5 and FS 246.3, two areas with longitudinal creases were noted. At FS 246.3, just below the inspection door, a compression dimple was noted in the skin that traveled from about three inches forward of FS 246.3 and traveled aft to FS 256.9.

The skin pulled away from five rivets located on the bottom of the rear fuselage, along the left side, and forward of FS 256.9. The skin from left to right side pulled away from all of the rear fuselage bottom rivets at FS 256.9. The skin pulled away from six of the rivets along the right side, and forward of FS 256.9. The two longitudinal running rivet rows on the bottom had the skin pulled away from seven rivets on the left side and six rivets on the right side. Diagonal crossing creases were noted in the bottom section panel between FS 246.3 and FS 256.9.

The lower right side of the rear fuselage at FS 233.5 just above the circular inspection panel, two about three inch longitudinal traveling creases were noted. At the upper section between FS 233.5 and FS 246.3, just forward of the right side stabilator, buckling in the shape of a chevron was noted. The area between FS 256.9 and FS 246.3 exhibited diagonal traveling creases from the bottom of FS 256.9 and traveled upward to and forward of FS 246.3. Three other diagonal traveling creases were noted from the upper ends of FS 256.9 and traveled forward and downward to FS 246.3. The lower section of the rear fuselage between FS 256.9 and 246.3 was buckled upward which created an "S" shaped compression signature.

There was no visible damage observed to either the left or right stabilizer and ruddervator assemblies. There were chipped paint marks on the top and bottom sides of both elevator control torque fittings. The marks corresponded to the torque fittings contacting the elevator stops mounted on the inboard elevator support hinge assembly. The elevator stops were not deformed. A flexible case static wick was attached to each elevator trailing edge to the outboard edge of the trim tab. The weight of each static wick was determined to be .4 oz. Installation of static wicks are not approved.

Tests and research

On July 17, 2002, a ruddervator balance check was performed using the Model 35 Maintenance Manual, Balancing Procedure Counterbalancing Method. The test revealed that the right elevator moment was calculated to be 19.012 in-lbs. static underbalanced. It took 4 lbs., 4.9 oz. weight hanging 4.415-inches distance forward of the hinge line to achieve a level balance of the elevator.

The left elevator moment was calculated to be 17.665 in-lbs. static underbalanced. It took 3 lbs., 15.3 oz. weight hanging 4.465-inches distance forward of the hinge line to achieve a level balance of the elevator.

The static underbalance weight range is listed as 16.8 to 19.8 inch-pounds on airplanes prior to serial number D-8118.

Additional information

The FAA's National Track Analysis Program (NTAP) of recorded radar data that contained position and altitude information, in 10 and 12 second intervals, showed that the aircraft was traveling southbound through the Seattle, Washington, area and turned easterly through Snoqualmie Pass, along the