Event WPR24LA038

Factual narrative

On November 17, 2023, about 2030 mountain daylight time, a Beech 95, N2024C, was substantially damaged when it was involved in an accident near Yuma, Arizona. The flight instructor and pilot receiving instruction were not injured. The airplane was operated as a Title 14 Code of Federal Regulations Part 91 instructional flight. The flight instructor reported that, after departing from the airport, the pilot receiving instruction noticed the rudder pedals moving back and forth. The instructor assumed control of the airplane, also felt the rudder pedals moving back and forth, and elected to return to the airport. =The oscillation of the rudder pedals intensified, and the airplane began to shake violently, which caused the front windscreen to separate from the airplane. The pilot receiving instruction declared an emergency while the instructor landed the airplane without incident. Postaccident examination of the airplane by a Federal Aviation Administration inspector revealed that the rudder trim control rod end was fracture-separated near the trim tab attachment point and the forward windscreen had separated, which resulted in substantial damage. The rudder trim control rod end and jam nut were sent to the National Transportation Safety Board Materials Laboratory in Washington, DC, for examination, which revealed that the rod end had fractured through the threads approximately 3 to 4 threads away from the clevis end. The upper ear of the clevis was bent and twisted, with contact damage noted on the left side of the clevis and on the threads between the clevis and the fracture surface. The forward side of the clevis showed a fractured surface with a curving arrest line, features consistent with fatigue. A scanning electron microscope (SEM) was used for further examination of the fatigue region. The SEM images revealed that the fatigue features emanated from a broad origin area at the root of the thread and propagated preferentially along the thread root. Course striations and secondary cracking were identified at higher magnifications and showed features consistent with a relatively high-stress, low-cycle fatigue fracture mechanism. It was also noted that the fracture surface appeared to be free of any substantial oxidation from a preexisting crack. The overall fracture appearance indicated that the fracture likely occurred due to repeated high compressive loads. Review of maintenance records revealed that about a month prior to the accident flight, maintenance and been performed on the rudder trim tab hardware. Flutter is an aeroelastic phenomenon that can occur when an airplane’s natural mode of structural vibration couples with the aerodynamic forces to produce a rapid periodic motion, oscillation, or vibration. Flutter can be somewhat stable if the natural damping of the structure prevents an increase in the forces and motions. Flutter can become dynamically unstable if the damping is not adequate or speed is increased, resulting in increasing self-excited destructive forces being applied to the structure. Flutter can range from an annoying buzz of a flight control or aerodynamic surface to a violent destructive failure of the structure in a very short period of time. Due to the high frequency of oscillation, even when flutter is on the verge of becoming catastrophic, it can still be very hard to detect. Aircraft speed, structural stiffness, and mass distribution are three inputs that govern flutter. An increase in airspeed, a reduction in structural stiffness, or a change in mass distribution can increase the susceptibility to flutter. Shortly after departing on the instructional flight, the pilot receiving instruction noted the rudder pedals moving back and forth without his input. The flight instructor assumed control of the airplane, also felt the rudder pedals moving, and chose to return to the airport. The oscillation of the rudder pedals intensified, the airplane began to shake violently, and the windscreen separated from the airplane. The pilots declared an emergency and subsequently landed uneventfully at the departure airport. Postaccident examination of the airplane revealed that the rudder trim control rod end was fracture separated near the trim tab attachment point. Metallurgical examination revealed fatigue cracking parallel to the clevis attachment bolt. The overall fracture was consistent with fatigue crack growth from bending loads on the fitting where the bending moment’s axis was parallel to the clevis attachment bolt axis. It is likely that buckling of the control rod end produced the fatigue crack initiation and propagation. The fracture surface appeared free of any substantial long-term oxidation from a preexisting crack, suggesting that the fracture likely occurred due to repeated, high-compressive loads associated with the reported rudder trim flutter condition, and was secondary to the flutter condition, not the cause. Flutter is governed by aircraft speed, structural stiffness, and mass distribution. An increase in airspeed, a reduction in structural stiffness, or a change in mass distribution can increase the susceptibility to flutter. Review of maintenance records revealed that maintenance had been performed on the rudder trim tab hardware about one month before the accident flight. While it is likely that the flutter event developed first and resulted in the fracture of the rudder trim control rod end, the initiating event of the rudder flutter could not be determined based on the available evidence. Source: NTSB Aviation Accident Database Retrieved: 2026-02-12

NTSB Findings

Hierarchical cause / factor breakdown from the FAA bulk avdata database. Each finding tagged C (Cause) or F (Factor).

• — Aircraft-Aircraft structures-Empennage structure-Rudder-Unknown/Not determined