This study investigates the damage initiation and failure process of six honeycomb sandwich composite fuselage panels subjected to quasi-static pressurization and axial loading using the Full-Scale Aircraft Structural Test Evaluation and Research (FASTER) facility located at Federal Aviation Administration William J. Hughes Technical Center in Atlantic City, NJ. The FASTER fixture is capable of testing full-scale fuselage panel specimens under conditions representative of those seen by an aircraft in actual operation. The lay-up of the test articles represents a typical general aviation sandwich composite aircraft fuselage structure, though the dimensions reflect that of a widebody aircraft for use in the FASTER fixture. The panels have an internal radius of 74 inch and are 125 inch long by 73 inch wide. The test articles contain 0.0085 inch-thick Toray Composites T700SC-12K-50C/2510 plain weave carbon fabric prepreg facesheets and a 0.75 inch -thick Plascore PN2-3/16-3.0 Nomex honeycomb core. The test section lay-up is [45/0/45/Core/45/0/45]. The artificial damage configurations evaluated in this program include 10 inch through-thickness notches oriented in the longitudinal, hoop, and 45 inclined directions, and a 10 inch diameter hole through the exterior facesheet. All six panels were loaded quasi-statically to failure, studying damage growth and strain redistribution behavior with increasing load and recording the residual strength. The test articles were instrumented with strain gages near the damage and in the farfield regions. A digital image correlation method was used to obtain full-field displacement and strain measurements at equal load intervals and after any visible surface damage was first observed. Several nondestructive inspections methods were used to scan for non-visual damage, including flash thermography, ultrasonics, radiography, and computer aided tap testing. The experimental results are being used to develop a material failure model for a fuselage of this composite layup under in-flight loading conditions with multiple internal damage modes.
Full-scale testing, inspection, and modeling of damaged hoeeyeomb sandwich composite fuselage panels
Dimensionsgerechte Prüfung, Inspektion und Modellierung von beschädigten Sandwich-Wabenstrukturlaminaten für Flugzeugrumpfpaneele
2009
16 Seiten, 9 Bilder, 3 Tabellen, 11 Quellen
Conference paper
English
Prüfung , Inspektion , Modellierung , Beschädigung , Wabenstrukturlaminat , Sandwich-Bauweise , Flugzeugrumpf , Kaltpressen , Pressen (Durchdrücken) , statische Belastung , Axialkraft , carbonfaserverstärkter Kunststoff , Prepreg , Zwischenschicht , Aramidfaser , Kerbung , axiale Verschiebung , Verstärkungsgewebe
Compression strength prediction of impact-damaged composite sandwich panels
Tema Archive | 2004
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