With the implementation of major aircraft structures fabricated from carbon fiber reinforced plastic materials, lightning protection has become a more complicated issue for designers and engineers to solve.

Cracking of coatings and surface layers is evident on a variety of structures including buildings, automobiles, and aircraft. In some situations, the appearance of the coated or painted surface is degraded and the aesthetic appeal is lessened. However, in others, such as composite aircraft structures, paint cracking is both aesthetically undesirable and potentially deleterious from the electromagnetic effects aspect.

Figure 1. Representative surface protection scheme modeled using COMSOL Multiphysics. The composite was modeled as two layers indicated by the red and light blue regions.

In the latter case, cracks can propagate into the structure or around fasteners, providing a path for moisture and other environmental species to enter, resulting in corrosion and degradation of the protection measures including expanded metal foils (EMF) required for lightning abatement and safe operation. Consequently, over several decades there have been numerous efforts and investigations concerning the degradation of surface layer protection schemes.

Figure 2. Examples of SWD/LWD ratios from 0.25 to 1.00 for a 1 in2 EMF.
Cracking typically develops over extended periods of time due to environmental factors and thermal cycling of the surface layers and substructure. The thermal cycling of aircraft is a direct result of the typical ground-to-air-toground, often repeated, flight cycle. Subsequently, stresses accumulate in the coatings, eventually leading to failure of their protective functionality.

There are several contributors to the stress buildup, including the paint, primer, corrosion isolation layer, surfacer, EMF, and the underlying composite substructure. Boeing recently did a study that focused primarily on the EMF contribution to the cracking mechanism.

A representative surface layer protection scheme was addressed that was composed of the layers mentioned above. The approach taken was to simulate the temperature cycle of the layers using a coefficient of thermal expansion (CTE) model developed with the commercially available COMSOL Multiphysics software.

The simulation allowed determination of the thermal stress and displacements that result from repeated duty cycles. Though the full complexity of crack genesis was not included, some insight could be gained regarding what the sensitive parameters of the EMF may be and the variations that can be employed to mitigate the resulting stress and displacements that lead to cracking. Of particular interest are the EMF width, height, aspect ratio, composition (aluminum (Al) or copper (Cu)), and surface layup structure.

In the case of Al used for EMF, there is a need for fiberglass between the aluminum and the structure to prevent galvanic corrosion. Though not the major thrust of this research, the potential effect on stress and displacement that results from the glass transition temperature of the paint layer must be considered.

Model Solving

Boeing developed a CTE model to simulate the effects of EMFs incorporated in a composite surface protection scheme undergoing thermal excursions using COMSOL Multiphysics software, a finite element solver that contains a variety of physics and engineering applications with an emphasis on coupled or multiphysics analysis.

In particular, Boeing used the Thermal Stress Multiphysics Interface that combines solid mechanics with heat transfer. Coupling occurs where the temperature from heat transfer acts as a thermal load for the solid mechanics, causing thermal expansion. The interface has the equations and features for stress analysis and general linear solid mechanics, solving for the displacements.

A representative surface protection scheme was created using the COMSOL model builder. The layup consisted of paint, primer, fiberglass, surfacer, EMF (Cu or Al), and fiberglass on a composite substrate. The height and size of all the layers could be varied using input parameters. In addition to geometrical parameters, the initial and final temperatures could be varied and were chosen to be typical of altitude and ground, respectively. The strain reference temperature was the initial temperature and a representative heat transfer coefficient of 5 W/m2K was used.