Abstract
AbstractThis review addresses a comparison, based on the literature, among nitrile rubber (NBR), ethylene-propylene-diene-monomer rubber (EPDM), and polyurethane (PU) elastomeric heat shielding materials (EHSM). Currently, these are utilized for the insulation of rocket engines to prevent catastrophic breakdown if combustion gases from propellant reaches the motor case. The objective of this review is to evaluate the performance of PU–EHSM, NBR–EHSM, and EPDM–EHSM as insulators, the latter being the current state of the art in solid rocket motor (SRM) internal insulation. From our review, PU–EHSM emerged as an alternative to EPDM–EHSM because of their easier processability and compatibility with composite propellant. With the appropriate reinforcement and concentration in the rubber, they could replace EPDM in certain applications such as rocket motors filled with composite propellant. A critical assessment and future trends are included. Rubber composites novelties as EHSM employs specialty fillers, such as carbon nanotubes, graphene, polyhedral oligosilsesquioxane (POSS), nanofibers, nanoparticles, and high-performance engineering polymers such as polyetherimide and polyphosphazenes.
Highlights
Thermal protection systems (TPS) are materials that shield the metallic or the composite casing of rocket engines by undergoing a process called ablation, which is the thermal decomposition under high temperature of the organic constituents that form the char layer
During the ablation process of a TPS, three main zones are formed (Figure 1): - The charring zone: The first zone that is affected by the high temperature in the solid propellant combustion process giving rise to the char layer. - The pyrolytic zone: Region between the charring zone and the virgin material, where the TPS decomposition and pyrolysis processes occur. - The virgin TPS zone: Region that consist on unreacted TPS [3,4]
The insulation material proposed was formulated on the basis of ethylene-propylenediene-monomer rubber (EPDM) rubber and filled with silica and magnesium oxide and was able to protect against heat, erosion, and other extreme conditions experienced by the engine or artifact during use and operation
Summary
Thermal protection systems (TPS) are materials that shield the metallic or the composite casing of rocket engines by undergoing a process called ablation, which is the thermal decomposition under high temperature (pyrolysis) of the organic constituents that form the char layer. During the ablation process of a TPS, three main zones are formed (Figure 1): - The charring zone: The first zone that is affected by the high temperature in the solid propellant combustion process giving rise to the char layer. To minimize the costs measuring the ablation properties of TPS, since the 1950s several computational codes (programs) have been developed to simulate the pyrolysis and the ablation of a TPS material under hyperthermal environment conditions, pressure, mass, and heat flux. According to the study by Natali et al [7], all indicated simulation programs are aimed at the optimization of the studies and developments of TPS, so that the test laboratories can lower the operating costs
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