Use of nondestructive inspection and fiber optic sensing for damage characterization in carbon fiber fuselage structure

Stephen Neidigk; Jacqui Le; Dennis Roach; Randy Duvall; Tom Rice

doi:10.1117/12.2048895

10 April 2014 Use of nondestructive inspection and fiber optic sensing for damage characterization in carbon fiber fuselage structure

Stephen Neidigk, Jacqui Le, Dennis Roach, Randy Duvall, Tom Rice

Author Affiliations +

Proceedings Volume 9063, Nondestructive Characterization for Composite Materials, Aerospace Engineering, Civil Infrastructure, and Homeland Security 2014; 90631S (2014) https://doi.org/10.1117/12.2048895
Event: SPIE Smart Structures and Materials + Nondestructive Evaluation and Health Monitoring, 2014, San Diego, California, United States

Abstract

To investigate a variety of nondestructive inspection technologies and assess impact damage characteristics in carbon fiber aircraft structure, the FAA Airworthiness Assurance Center, operated by Sandia National Labs, fabricated and impact tested two full-scale composite fuselage sections. The panels are representative of structure seen on advanced composite transport category aircraft and measured approximately 56”x76”. The structural components consisted of a 16 ply skin, co-cured hat-section stringers, fastened shear ties and frames. The material used to fabricate the panels was T800 unidirectional pre-preg (BMS 8-276) and was processed in an autoclave. Simulated hail impact testing was conducted on the panels using a high velocity gas gun with 2.4” diameter ice balls in collaboration with the University of California San Diego (UCSD). Damage was mapped onto the surface of the panels using conventional, hand deployed ultrasonic inspection techniques, as well as more advanced ultrasonic and resonance scanning techniques. In addition to the simulated hail impact testing performed on the panels, 2” diameter steel tip impacts were used to produce representative impact damage which can occur during ground maintenance operations. The extent of impact damage ranges from less than 1 in2 to 55 in2 of interply delamination in the 16 ply skin. Substructure damage on the panels includes shear tie cracking and stringer flange disbonding. It was demonstrated that the fiber optic distributed strain sensing system is capable of detecting impact damage when bonded to the backside of the fuselage.

Citation Download Citation

Stephen Neidigk, Jacqui Le, Dennis Roach, Randy Duvall, and Tom Rice "Use of nondestructive inspection and fiber optic sensing for damage characterization in carbon fiber fuselage structure", Proc. SPIE 9063, Nondestructive Characterization for Composite Materials, Aerospace Engineering, Civil Infrastructure, and Homeland Security 2014, 90631S (10 April 2014); https://doi.org/10.1117/12.2048895

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KEYWORDS

Inspection

Skin

Fiber optics

Ultrasonics

Composites

Carbon

Nondestructive evaluation

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