N129B

History of Flight

On October 6, 1997, at 1330 central daylight time (CDT), a Beech G35, N129B, sustained substantial damage during descent when it experienced tail flutter near Poplar Grove, Illinois. When the pilot reduced power the airplane stopped oscillating. The 14 CFR Part 91 flight had departed Janesville, Wisconsin, at 1300 CDT and landed at Poplar Grove Airport, Poplar Grove, Illinois, the intended destination. The pilot and one passenger were not injured. Visual meteorological conditions prevailed and no flight plan had been filed.

The pilot reported that after departing Janesville he climbed to 4,000 feet. He reported that the sky was clear with smooth air. He initiated a descent to Poplar Grove. The power was set at 23 inches manifold pressure and 2,300 RPM. He reported that the airplane was trimmed for descent, and he was not holding any pressure to keep the nose down. The airplane was descending at 600 to 700 fpm and at 185 mph. The airplane was in a 15 degree left turn with a 5 degree nose down attitude.

He reported that at about 3,200 feet the yoke started to move fore and aft about 1/2 inch. He described the motion to be, "...like [an] autopilot malfunction in Altitude Hold." The fore and aft travel on the yoke then increased to about 1 inch. He reported that the movement was not violent. During the 1 inch fore and aft movement of the yoke, the airplane experienced 3 to 4 solid jolts, and the entire airframe shook or banged 3 to 4 times. The pilot described the jolts as being:

1. Like severe chop. 2. Like going over a wake in a speedboat. 3. Like crossing wake turbulence. Not tossing. 4. Like bottoming out. Not rapid.

He reported that he did not feel any vibration in the rudder pedals. After about 5 seconds of oscillation, the pilot pulled back on the throttle and the banging stopped. He continued to his destination and landed uneventfully. When he was pushing the airplane back into the hangar, he discovered the damage to the empennage.

Damage to Aircraft

The airplane was inspected for damage by the National Transportation Safety Board's Investigator-in-Charge, an Airworthiness Inspector of the Federal Aviation Administration (FAA), a representative 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. The longerons were removed from the airplane. Permanent downward bends were set in the longerons. The aft end of the right longeron had been displaced about 3 inches, and the aft end of the left longeron had been displaced about 2.5 inches.

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, the primary wrinkle 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 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.

4. 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.

5. The right stabilizer had three creases on the top of the leading edge and two creases on the bottom. They were located between the second and third rib. There were no creases on the left 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 balance moments of 18.76 and 18.2 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 3.20 and 3.09 pounds, respectively. Both ruddervator counterbalance weight assemblies had approved lead washers bolted to each of the factory installed counterbalance lead weights. There were eight 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 2.48 and 2.49 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.

9. With the trim wheel set a zero, the right ruddervator trim tab was up about 5/16 inch. The left ruddervator trim tab was centered on the ruddervator. (The pilot reported that the airplane flew very slightly right wing low, ball centered, with hands off the controls.)

10. The ruddervator control cables were slightly loose, but the pilot indicated that prior to the accident the cables were tight.

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 0.25 to 0.27 and 0.7 inches per second (ips), respectively. The normal acceptable frequency range was 0.2 ips or less.

2. The pitot static system was checked about a month prior to the accident.

The airplane was inspected for items installed that were not original to the aircraft as manufactured. The installed items included:

1. Wing tips.

2. One piece windshield.

3. Hartzell hydraulic propeller

4. Battery box forward of the firewall.

5. Air/oil separator.

6. Gage oil filter.

7. Modified instrument panel and seats.

Personnel Information

The pilot held a airline transport certificate with a single and multi-engine land rating. He had about 18,000 hours total flight time with 3,500 hours in single engine airplanes. He had 75 hours in make and model. He had flown the airplane 40 hours within the last 90 days.

Aircraft Information

The airplane was a Beech G35 Bonanza, serial number D-4433. The last annual inspection was done on July 8, 1997. The airframe had a total of 5,439 hours. The engine was a 225 horsepower Continental E-225-8. The engine had about 390 hours since a major overhaul. The propeller was a Hartzell two bladed hydraulic propeller.

The aircraft logbooks indicated that the engine was overhauled in March 1993. The fuel primer hose separated while in flight and the pilot made an off field landing. This occurred 8.5 hours after the engine overhaul. An engine teardown inspection was completed. The propeller blades and clamps were replaced and the propeller was overhauled. The nose gear and nose gear doors were repaired. There was no indication of damage to the empennage. 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 202 mph. The Maximum Structural Cruising speed, Vno, was 175 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

Metallurgists from the National Transportation Safety Board's Materials Laboratory examined the airplane. (See Materials Laboratory Factual Report)

Radar data was analyzed and it indicated that the airplane was travelling at an average ground speed of 184 mph when...