Multi-layered Self-healing Material System for Impact Mitigation

Puncture healing melt blends were developed by melt blending self-healing polymers with non self-healing polymeric materials. The self-healing polymeric materials consisted of Surlyn&#174 8940, Affinity&#8482 EG 8200 G, and poly(butadiene)-graft-poly(methyl acrylate-co-acrylonitrile) or Barex&#174 210 IN. The non-self-healing polymeric materials consisted of poly(ether ether ketone) (PEEK), LaRC phenyl ethynyl terminated imide 330 (LaRC PETI 330), and Raptor Resins Bismaleimide-1 (BMI-1). Puncture healing blends were also prepared with chopped glass and chopped carbon fibers. The overall goal was to develop a product with superior properties relative to either of the starting materials. The melt blends were prepared in varying compositions to optimize desired properties of the resulting matrix. Ballistic testing was conducted to determine the self-healing characteristics of several developmental polymers subjected to micrometeoroid type damage.

A composite fabrication process cycle was developed from composite precursor materials developed at LaRC to fabricate composite laminates. The precursor material is a pre-impregnated unidirectional carbon fiber preform, or prepreg. In the pre-pregging process, the high strength, structural reinforcing carbon fiber is wetted by a solution containing a self-healing polymer. The resulting material is of aerospace quality and exhibits a significant decrease of internal damage following impacts tests (using ASTM D 7137 standard).

This technology is a flexible, lighter weight radiation shield made from hybrid carbon/metal fabric and based on the Z-grading method of layering metal materials of differing atomic numbers to provide radiation protection for protons, electrons, and x-rays. To create this material, a high density metal is plasma spray-coated to carbon fiber. Another metal with less density is then plasma spray-coated, followed by another, and so on, until the material with the appropriate shielding properties is formed. Resins can be added to the material to provide structural adhesion, reducing the need for mechanical bonding. This material is amenable to molding and could be used to build custom radiation shielding to protect cabling and electronics in situations where traditional metal shielding is difficult to place.

NASA used three strategies to develop a family of modified epoxy resins for use in an improved composite layup configuration. By tailoring molecular structures, incorporating secondary additives, and adjusting composite architecture, NASA has produced fiber-reinforced polymer composites that suppress viscoelastic creep and relaxation. Specifically the three strategies were: (1) Controlling the cross-linking density using reactive functional groups and selecting appropriate monomers with stoichiometry adjustments. A higher stoichiometric ratio is responsible for reduced relaxation compared to a lower ratio (resulting in 26% relaxation after 1 year vs. 49%). (2) Increase the steric hindrance by reducing free volume and enhancing intermolecular interaction. This involved the preparation of two different reactive low molecular weight oligomers The best epoxy resin created was from the VCD/HAA additives with 15-17% relaxation; DEGBA/HAA had 22-25% relaxation. (3) Optimizing different carbon fiber layup configurations for woven structures made into a highly stiff boom, highly flexible boom, a small size boom, and a large size boom application. The best configuration was that of a highly stiff boom with 4-5% relaxation. Other carbon fiber layup configurations had 16-30% relaxation. As shown in the figure below, NASA tested the relaxed composite using an accelerated test with variable temperatures.

The invention includes an exposed surface cap with a specially formulated coating, an insulator base adjacent to the cap with another specially formulated coating, and one or more pins that extend from the cap through the insulator base to tie the cap and base together through ceramic bonding and mechanical attachment. The cap and insulator base have corresponding depressions and projections that mate and allow for differences in thermal expansion of the cap and base. The cap includes a high-temperature, low density, carbonaceous, fibrous material whose surface is optionally treated with a High Efficiency Tantalum-based Ceramic Composite (HETC) formulation, the fibrous material being drawn from the group consisting of silicon carbide foam and similar porous, high temperature materials. The insulator base and pin(s) contain similar material. The mechanical design is arranged so that thermal expansion differences in the component materials (e.g., cap and insulator base) are easily tolerated. It is applicable to both sharp and blunt leading edge vehicles. This extends the possible application of fibrous insulation to the wing leading edge and/or nose cap on a hypersonic vehicle. The lightweight system comprises a treated carbonaceous cap composed of Refractory Oxidation-resistant Ceramic Carbon Insulation (ROCCI), which provides dimensional stability to the outer mold line, while the fibrous base material provides maximum thermal insulation for the vehicle structure. The composite has graded surface treatments applied by impregnation to both the cap and base. These treatments enable it to survive in an aero-convectively heated environment of high-speed planetary entry. The exact cap and base materials are chosen in combination with modified surface treatments and a specially formulated surface coating, taking into account the duration of exposure and expected surface temperatures for the particular application.