The Preload Relaxation Mechanism of Composite Bolted Joints Under Bending Vibration Load: Experimental and Numerical Study

Persistent bending vibrations in aircraft wing composite bolted joints induce progressive preload relaxation and fastener loosening, threatening structural integrity and flight safety. This study integrates vibration testing and numerical simulations to investigate the mechanisms of preload relaxation. Controlled bending vibrations were applied using a precision exciter, while an ultrasonic system monitored real‐time preload. Post‐test analysis with microscopy and 3D profilometry characterized damage in fasteners and laminates, complemented by finite element simulations. Results reveal a two‐stage preload relaxation: an initial rapid attenuation (18% loss) followed by gradual reduction (5%). Prolonged vibration caused severe fretting damage, including matrix cracking and fiber fracture in laminates, and complex mixed wear on thread surfaces involving fatigue, abrasion, adhesion, and oxidation. Wear at thread interfaces was more pronounced than in laminates, decreasing at higher preload but intensifying under larger load amplitudes. A finite element model incorporating thread wear effects was developed, employing a cyclic jump technique for efficiency and predictive accuracy. Numerical results quantified the evolution of wear depth and contact parameters, confirming that preload relaxation stems from contact pressure redistribution and fretting wear accumulation. High preload suppressed relaxation by increasing contact pressure, while high load amplitude accelerated it by enhancing relative slip.