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. 2021 Sep 22;13(19):3207.
doi: 10.3390/polym13193207.

Innovation in Aircraft Cabin Interior Panels Part I: Technical Assessment on Replacing the Honeycomb with Structural Foams and Evaluation of Optimal Curing of Prepreg Fiberglass

Edgar Adrián Franco-Urquiza  1 , Annika Dollinger  2 , Mauricio Torres-Arellano  3 , Saúl Piedra  1 , Perla Itzel Alcántara Llanas  3 , Victoria Rentería-Rodríguez  3 , Cecilia Zarate Pérez  3
Affiliations

Innovation in Aircraft Cabin Interior Panels Part I: Technical Assessment on Replacing the Honeycomb with Structural Foams and Evaluation of Optimal Curing of Prepreg Fiberglass

Edgar Adrián Franco-Urquiza et al. Polymers (Basel). .

Abstract

Sandwich composites are widely used in the manufacture of aircraft cabin interior panels for commercial aircraft, mainly due to the light weight of the composites and their high strength-to-weight ratio. Panels are used for floors, ceilings, kitchen walls, cabinets, seats, and cabin dividers. The honeycomb core of the panels is a very light structure that provides high rigidity, which is considerably increased with fiberglass face sheets. The panels are manufactured using the compression molding process, where the honeycomb core is crushed up to the desired thickness. The crushed core breaks fiberglass face sheets and causes other damage, so the panel must be reworked. Some damage is associated with excessive build-up of resin in localized areas, incomplete curing of the pre-impregnated fiberglass during the manufacturing process, and excessive temperature or residence time during the compression molding. This work evaluates the feasibility of using rigid polyurethane foams as a substitute for the honeycomb core. The thermal and viscoelastic behavior of the cured prepreg fiberglass under different manufacturing conditions is studied. The first part of this work presents the influence of the manufacturing parameters and the feasibility of using rigid foams in manufacturing flat panels oriented to non-structural applications. The conclusion of the article describes the focus of future research.

Keywords: aircraft cabin interior panels; foams; non-structural composite panels; optimal curing of prepregs.

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Conflict of interest statement

The authors declare no conflict of interest.

Figures

Figure 1
Figure 1
Pictures of the individual materials used to manufacture distinct sandwich composites: (a) glass fiber prepreg, (b) honeycomb core, (c) foam core.
Figure 2
Figure 2
Pictures of representative tests of: (a) samples in the aluminum pan before hermetic close and test by differential scanning calorimetry (DSC), (b) specimen of glass fiber before testing in dynamic mechanical analysis (DMA).
Figure 3
Figure 3
Workflow for the typical ACP usage of a sandwich analysis in ANSYS.
Figure 4
Figure 4
Pictures corresponding to the GF laminates manufacturing: (a) front view of closed hot plates with temperature and pressure displays, (b) GF laminate fabricated at 200 °F (93 °C) during 1 h, (c) GF laminate fabricated at 300 °F (149 °C) during 2 h.
Figure 4
Figure 4
Pictures corresponding to the GF laminates manufacturing: (a) front view of closed hot plates with temperature and pressure displays, (b) GF laminate fabricated at 200 °F (93 °C) during 1 h, (c) GF laminate fabricated at 300 °F (149 °C) during 2 h.
Figure 5
Figure 5
DSC curves corresponding to the samples 0 to 6: (a) first heating scan, (b) cooling step, (c) second heating scan.
Figure 6
Figure 6
DSC curves corresponding to the samples 7 to 10: (a) first heating scan, (b) second heating scan.
Figure 7
Figure 7
DMA curves corresponding to the samples 1 to 5: (a) storage modulus, (b) tan δ.
Figure 8
Figure 8
Force vs. displacement with Hookean region of core material.
Figure 9
Figure 9
Stress vs. strain curves corresponding to the honeycomb core (HC) and foam tested under compression.
Figure 10
Figure 10
Side panel with fixed support and applied load.
Figure 11
Figure 11
Applicable failure criterion and related ply in brackets in the bottom left corner of the baseline panel.
Figure 12
Figure 12
Total stress of the baseline panel in ACP.
Figure 13
Figure 13
Total stress of the study case panel in ACP.
See this image and copyright information in PMC

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