Test and Analysis of a Composite Multi-bay Fuselage Panel under Uni-axial Compression

A composite panel containing three stringers and two frames cut from a vacuum-assisted resin transfer molded (VaRTM) stitched fuselage article was tested under uni-axial compression loading. The stringers and frames divided the panel into six bays with two columns of three bays each along the compressive loading direction. The two frames were supported at the ends with pins to restrict the out-of-plane translation. The free edges of the panel were constrained by knifeedges. The panel was modeled with shell finite elements and analyzed with ABAQUS nonlinear solver. The nonlinear predictions were compared with the test results in out-of-plane displacements, back-toback surface strains on stringer flanges and back-toback surface strains at the centers of the skin-bays. The analysis predictions were in good agreement with the test data up to post-buckling. INTRODUCTION Efforts have been made to demonstrate composite structural technology for rotorcraft primary structures to save weight and cost [1]. However, the cost and weight savings cannot be realized without a better understanding of the structural integrity issues associated with unitized composite structures. A high fidelity failure analysis methodology was previously proposed to analyze unitized composite structures [2]. The objective of this proposed method was to understand the global nonlinear behavior of the entire structure due to the interactions among its components and define local failure modes at the joining locations of the structural components. The high fidelity failure analysis methodology was able to address the limitations of strength based analysis used in the current design methods and identify, before full-scale static tests, potential failure modes missed by traditional strength based analysis [2-3]. The effectiveness of this high fidelity failure analysis methodology needs to be validated by carefully conducted experiments. Therefore, the objective of this paper is to test, under uni-axial compression a panel cut from the Rotary Wing Structures Technology Demonstrator (RWSTD) [1] composite fuselage tool proof article and to model this test article with finite elements. The paper will provide comparison of the model results and the test results in strains and displacements. This paper reports the experimental methods and numerical simulation. SPECIMEN DESIGN AND EXPERIMENTAL PROCEDURES A three-stringer multi-bay panel (Fig. 1) was cut from a RWSTD [1] composite fuselage tool proof article (Fig. 2) and made into a compression specimen. The fuselage tool proof article was manufactured from stitched, warp knit, and plain weave, AS4 carbon fiber preforms infused with SI-ZG-5A resin system using vacuumassisted resin transfer molding (VaRTM) process. The skin was stitched together and the frame and stringer flanges were stitched to the skin. The stringers and frames divided the specimen into six bays with two columns of three bays each along the compressive loading direction. For the remainder of this paper the bays will be identified by their relative location on the test specimen (i.e. upper left, middle right, etc) as shown in Fig. 3 when viewing from the stringer side. The specimen was potted at both ends to facilitate load introduction. The two middle frames were supported at the ends. Fig. 1. Three-stringer multi-bay compression specimen. Frames Stringer 45th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics & Materials Conference 19 22 April 2004, Palm Springs, California AIAA 2004-2056 Copyright © 2004 by The Boeing Company. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. Fig. 2. RWSTD composite tool proof article. MPCs for pin Slider MPCs for knife edge