On April 29, 1997, at 1330 eastern daylight time (EDT), a Beech D-35, N5985C, sustained substantial damage during cruise flight when it experienced tail flutter near Mt. Sterling, Kentucky. When the pilot reduced power the airplane stopped shaking, and the pilot elected to land at Lexington, Kentucky. Neither the pilot nor the passenger received injuries. The 14 CFR Part 91 flight departed Chapel Hill, North Carolina, at 1130 EDT en route to Des Moines, Iowa. A fuel stop at Seymour, Indiana, had been planned. Visual meteorological conditions prevailed and a VFR flight plan had been filed.
The pilot reported that after departing Chapel Hill he climbed to 7,500 feet. He reported that he set the power for cruise flight which produced about 160 mph. He reported that the sky was initially clear, and then turned to partly cloudy at his flight altitude. In order to keep clear of clouds he started a descent to 6,500 feet.
He reported that he experienced some light turbulence and the airplane began to vibrate and shake. After 10 to 15 seconds, the pilot pulled back on the throttle and the vibration stopped. The pilot thought the vibration was an engine related problem and suspected the engine might fail. After checking the engine gauges, he recognized the engine was operating normally and he applied power for cruise flight. He reported that before reaching cruise airspeed, the airplane started to shake again, only much more violently than the first occurrence. After 5 to 10 seconds, the pilot reduced power and the shaking stopped. He reported the shaking, "...reminded me of an old car with wheels out of alignment and balance. I thought the engine was coming out of [the] plane... ."
The pilot continued flying at 140 mph to Lexington, Kentucky. After an uneventful landing the pilot taxied to a ramp where he could do an engine run-up. The pilot applied full power and the engine and propeller operated normally with no vibrations. After the pilot deplaned, he observed that the inspection access panel door on the left side of the empennage was being held on by one screw. He then observed the substantial damage to the tail of the airplane.
Damage to Aircraft
The airplane was inspected for damage by the National Transportation Safety Board's Investigator-in-Charge, Airworthiness Inspectors of the Federal Aviation Administration (FAA), representatives from the airplane manufacturer, Raytheon Aircraft Company (RAC), and maintenance representatives from the American Bonanza Society (ABS).
The aft fuselage damage was concentrated between the fuselage station (FS) 233.5 and 256.9 bulkheads. The aft fuselage exhibited extensive skin wrinkling and tearing between the FS 256.9 (front stabilizer spar attachment) and the 233.5 bulkhead. The rest of the aft fuselage exhibited no apparent deformation or cracks.
The damage to the fuselage included:
1. Both lower longerons were bowed inward forward of FS 256.9.
2. Both sides of the fuselage were buckled with permanent wrinkles that extended across the stringers and across the doubler around the access hole on the left side. The primary wrinkle on the right side extended diagonally across the fuselage from upper forward to lower rear. On the left side, it appeared that the primary wrinkle was also extended diagonally across the fuselage from upper forward to lower rear. However, the doubler around the opening for the access hole affected the wrinkling and it was not a straight diagonal wrinkle. The wrinkles extended a short distance into the curved top skin. There was no evidence that the access door scratched the surrounding skin.
3. The lower two stringers on the right, inner side of the aft fuselage were crippled and cracked approximately 5 inches aft of the FS 256.9 frame. The lower stringer on the left, inner side of the aft fuselage was crippled inward, forward, and adjacent to the access door reinforcement.
4. The aft fuselage, lower skin, forward of the FS 256.9 frame was wrinkled and deformed down as if it pulled away from the longerons and frame. The rivets attaching the skin to the frame had pulled through the skin. The skin was separated from the longerons for approximately the downward deformation distance. The wrinkles in the skin were not symmetrical. They were approximately parallel to the longerons, but one side was buckled inward and the other side was buckled outward.
5. There was no observable damage to the top skin, except where the ends of the wrinkles in the side panels extended to the top skin.
6. Indentations, approximately 2 inches in length, running aft from the leading edge were found in the upper and lower surfaces of both the left and right stabilizers. The indentation's were between the second and third ribs from the inboard end of the stabilizer.
The examination of the empennage revealed the following:
1. Both stabilizers were securely attached to the FS 256.9 bulkhead. Both ruddervators were securely attached to the stabilizers, and exhibited no excessive hinge wear or looseness.
2. The stabilizer spar attachment holes in the bulkheads at FS 256.9 and 272.0 indicated no bushing hole elongation. Some uniform bushing wear was detected. The bulkheads showed no visible damage.
3. The stabilizer spar attachment holes appeared round and uniform.
4. The ruddervator travel stops on the ruddervator inboard hinge fitting and the contact point on the ruddervator torque fittings showed no evidence of excessive impact loading.
5. Both ruddervators were checked for balance using the force method per the Bonanza 35 Series Shop Manual. The right and left ruddervators exhibited static underbalance moments of 19.67 and 19.27 inch-pounds, respectively. The required static balance range is 16.80 to 19.80 inch-pounds.
6. The right and left ruddervator counterweight assemblies (including the attaching skin) weighed 2.86 and 2.93 pounds, respectively. Both ruddervator counterbalance weight assemblies had approved lead washers bolted to each of the factory installed counterbalance lead weights. There were four washers on the right and six washers on the left. All of the ruddervator counterbalance lead weights were attached securely. The right and left factory installed counterbalance lead weights weighed 2.41 and 2.20 pounds, respectively.
7. The empennage control systems exhibited no excessive looseness or improper installations.
8. The aluminum skin thicknesses for the top, side, and bottom skins were the required thicknesses, 0.032, 0.020, and 0.016, respectively.
The entire airplane was inspected for any discrepancies that might have contributed to the accident. The inspection revealed the following information:
1. The propeller was checked for dynamic balance. The frequencies of the forward and aft accelerometers were about 2.0 and 0.7 inches per second (ips), respectively. The normal acceptable frequency range was 0.2 ips or less.
2. The airspeed indicator was accurate within 3%.
3. The left outboard wing near the tip had been damaged in the past. Neither wing contained any buckling, tearing, shearing, or other signs of overload.
4. The number 5 rivet on the right hand, aft engine mount was sheared.
5. There was a crack in the skin of the nose gear "dog house."
6. A large mouse nest was in the left hand outboard wing just forward of the rear spar.
7. The number 6 cylinder had been hitting the cowl door and had worn a hole into the door structure.
The airplane manufacturer's structural integrity specialist reported that, "The symmetry of the failure about the vertical centerline (except for the buckling of the bottom skin) indicates failure due to a vertical load down on the tail. The side panels probably buckled first and caused the longerons to fail by forced crippling. The damage to the bottom skin and stringers then occurred."
The airplane was inspected for items installed that were not original to the aircraft as manufactured. The installed items included:
1. Wing tips.
2. Mirrors installed on wing tips.
3. Windshield, cabin door, and pilot's side windows.
4. Cleveland brakes and wheels.
5. Continental E-225-8 engine.
6. 10 gallon fuselage fuel tank.
The aircraft manufacturer's technical service representative reported that the aircraft, "...had been somewhat modified. However, in my opinion these modifications would not be a contributing factor to the inflight vibration encountered by the pilot/owner."
The pilot held a private certificate with a single engine land rating. He had 2,000 hours total flight time with 25 hours in make and model.
The pilot purchased the Beech D35 Bonanza, serial number D-3410, in December 1996. An annual inspection was done on June 26, 1996. The airframe had a total of 3,508 hours. The engine was a 225 horsepower Continental E-225-8. The engine had about 253 hours since a major overhaul. The propeller was an original Beech propeller with electric pitch control.
The aircraft logbooks indicated that in June 1996 the nose landing gear and nose gear doors had received extensive repair. In September 1996, the right ruddervator and left flap had been reskinned due to corrosion. The aircraft logbooks indicated that all repairs were performed in accordance with applicable maintenance manuals.
The Pilot's Operating Handbook listed the Vne, Never Exceed speed, as 203 mph. The Maximum Structural Cruising speed, Vno, was 160 mph.
Under the FLUTTER section of the Beechcraft Single Engine (Piston) Safety Information guide, the following information is provided in case excessive vibration in the controls is encountered:
"If an excessive vibration, particularly in the control column and rudder pedals, is encountered in flight, this may be the onset of flutter and the procedure to follow is:
1. IMMEDIATELY REDUCE AIRSPEED (lower the landing gear if necessary).
2. RESTRAIN THE CONTROLS OF THE AIRPLANE UNTIL THE VIBRATION CEASES.
3. FLY AT THE REDUCED AIRSPEED AND LAND AT THE NEAREST SUITABLE AIRPORT.
4. HAVE THE AIRPLANE INSPECTED FOR AIRFRAME DAMAGE, CONTROL SURFACE ATTACHING HARDWARE CONDITION/SECURITY, TRIM TAB FREE PLAY, PROPER CONTROL CABLE TENSION, AND CONTROL SURFACE BALANCE BY ANOTHER MECHANIC WHO IS FULLY QUALIFIED."
Tests and Research
A section of the airplane's empennage was sent to the National Transportation Safety Board's Materials Laboratory for examination. The longerons were examined for pre-existing damage. The metallurgist reported that, "All four fracture surfaces were composed almost entirely of ductile dimples, consistent with an overstress separation. Each also contained a very small area of fatigue striations near the terminal crack front, as if the crack had endured a small amount of cyclic stress after the main loading event."
The pilot reported that he flew the airplane to Colorado in December 1996. He reported that during the flight he experienced a vibration, but thought it was turbulence related since he was flying over the foothills of the Rocky Mountains. He had been flying at about 8,000 feet between 160 to 170 mph. He described the vibration as a "mild shudder," and did not think it was threatening.
The pilot reported that he had been flying to Colorado in late January or February 1997, when he experienced another vibration incident. He was flying at 3,000 feet and he thought it was another case of turbulence. He did not experience any negative G's and he described it as a "mild shudder." He reduced power and the vibrations stopped. He did not have the tail of the airplane inspected after the two vibration occurrences.
He reported that he had waxed the airplane twice between the time when he had purchased the airplane and the day of the accident. He reported that he had not noticed any wrinkles or dents to the tail surfaces, or any creases in the longerons.
Parties to the investigation included the Federal Aviation Administration, Raytheon Aircraft Company, and the American Bonanza Society.
The airplane has been undergoing repair, and it is intended to be back in service in the spring of 1998.