Detecting Damage in a UAV Composite Wing Spar Using Distributed Fiber Optic Strain Sensors

Abstract

The primary structure of an unmanned air vehicle (UAV) wing is commonly composed of outer wing skins attached (bonded or fastened) to a main wing spar both composed of advanced composite materials. The main wing spar is designed to take the primary bending and transverse shear loads, whereas the wing skin is designed to take the wing torsion loads. Since these aircraft are developed for long range continuous (surveillance) flight with minimal time for ground inspection, it is desirable to have an in-flight structural health monitoring (SHM) system for insuring the structural integrity of the wing structural components and attachments. In the current study an SHM computational and experimental testbed is developed for analyzing the structural integrity of a composite wing spar, where fixturing is included to produce realistic wing bending and torsion natural frequencies and corresponding coupled bending-torsion vibration modes. The wing spar is instrumented using Bragg graded fiber optic distributed strain sensors along the spar length on three sides to determine the axial strain and bending modal curvature distributions and in a wavy configuration on the fourth side in an attempt to monitor spar twisting behavior. Initially, finite element studies of a damaged composite wing spar are performed to determine the modal curvature distributions for evaluation of numerous published SHM approaches. Finally, an experimental study is performed where the free vibration response of the undamaged and damaged wing spar is monitored based upon either an impact excitation or an initial displacement. While these results clearly show the damage using the computational models, there are difficulties locating damage using the distributed fiber optic strain sensors.