Experiment and numerical investigation of the aero-thermal demise of pristine and pre-damaged GLARE™ fiber-metal laminates

Abstract

Since many decades researchers have been interrogating the re-entry physics of lunar objects in the Earth’s atmosphere. A request like ‘demise + re-entry’ in a web search engine gives hundreds if not thousands of studies concerned with the breakup and the disappearance of space flotsam in the Earth’s atmosphere. However, it is hard to comprehend a rational reason why the engineering physics involved in the thermal degradation of spacecraft structures still can be explained with inherent uncertainties. Many derelict satellites and spacecraft are orbiting the Earth. They will gradually enter the Earth’s atmosphere attributed to the gravitational force and may will not survive the aero-thermal load induced by the dense Earth’s atmosphere. This study, therefore, investigated the propensity to aero-thermal demise of a pristine and a pre-damaged spacecraft shielding structure during atmospheric re-entry. The shield was made of a GLAss fiber REinforced aluminum panel, which accommodated five quasi-isotropic S2-glass fiber reinforced FM94-epoxy laminates alternately stacked between six aluminum 2024-T3 alloy sheets. The pre-damage of the shielding system was akin to a damage can be generated by an impact of a micrometeoroid. To impart the pre-damage, a GLAss fiber REinforced aluminum plate was first impacted with a 2 mm diameter stainless steel spherical projectile at 5.49 km/s. The projectile perforated through the plate. Next, the perforated GLAss fiber REinforced aluminum plate was exposed to an air plasma-jet (oxygen dissociated) of a specific enthalpy of 5.4 × 103 kJ/kg at standard air pressure conditions. The high enthalpy plasma thermally degraded the pre-damaged plate, while debris was discharged uprange and downrange out of the plate. The dynamic pressure of the plasma flow exacerbated the debris discharge. The endothermic destruction of a pristine GLAss fiber REinforced aluminum specimen followed several steps: first, layer-after-layer oxidation and melting of the aluminum sheets; next, decomposition and sublimation of epoxy accompanied by thermal and viscous diffusion of molten aluminum and molten glass fibers; finally, the onset of a hole in the pristine specimen. By contrast, the plasma-jet that streamed through the punched hole of the pre-damaged GLAss fiber REinforced aluminum specimen, had thermally decomposed the frontal material layers of the specimen to a large extent in the initial few seconds. The punched hole, however, did not succumb to enlarge after the surface temperature of the GLAss fiber REinforced aluminum specimen reached an asymptote. To reproduce the experimental findings, a new Moving Mesh Front procedure had been developed and implemented in a Lagrangian finite element framework. The numerical analysis considered the temperature-dependent thermal properties of orthotropic S2-glass/FM94-epoxy composites and isotropic aluminum alloys. For a comparable heating condition, the agreement between the predicted and the measured temperatures of the specimen front surface was within 4.4%, while the corresponding agreement was within 6.7% for the rear surface temperature of the test specimen.