Flutter analysis of anisotropic panels with patched cracks

Abstract

A finite element approach is developed to analyze the panel flutter problems of thin plate-like composite panels with patched cracks. The panel studied is a compound structure that comprises three layers including a thin cracked panel, an adhesive, and a thin patch. A 48-degree-of-f reedom (DOF) triangular crack patching element is derived for numerically analyzing the compound structure. The aerodynamic pressure acting on the panel is estimated using linearized piston theory with aerodynamic damping effect neglected. By a proper tailoring of the materials, very high flutter and/or divergence boundary could usually be obtained for the composite panels. The existence of a crack usually reduces the aeroelastic stability boundary; however, some exceptions were found for the composite panel within certain range of specific filament orientation. The deterioration in flutter/divergence performance due to a crack can, in general, be cured by means of patching, and anisotropic patching is more effective as compared to the isotropic patching provided that the tailoring of the structural parameters is done correctly. I. Introduction A RELIABLE and accurate flutter analysis is of primary importance in the flight vehicle design. The onset of flutter would quickly develop into either catastrophic structural failure or undesirable limit cycle type oscillation. The local panel flutter of the vehicle could be essential for a global flutter to occur. For instance, the initiation of cracks in wing structure is not unusual as they may be produced by either molding, fatigue, or impact damage.1'2 These defects could induce extremely high stresses around the cracktip regions and, thus, precipitate structural failure. When cracks are present, this local structural problem may have significant influence on the local flutter/divergence of the panel structure. The onset of the local aeroelastic instability may cause the failure of the cracked panel and cumulatively affect the global structure as well as the airload distribution and finally result in