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During process development testing of the ARES I composite interstage, large regions of node disbonds were observed in the aluminum honeycomb core. The sandwich panels were composed of carbon fiber/epoxy face sheets and one-eighth inch cell size aluminum core co-cured in an autoclave. Node disbonds were initially observed after cure with ultrasonic inspection and subsequently confirmed with X-ray inspection and cross sectioning. A stress analysis was performed on the residual thermal stresses resulting from post-cure cool down. Analysis indicated that the thermal stresses in the aluminum core are a function of foil thickness of the aluminum core. Thermal stresses are relieved through disbonding of the adhesive node. The node disbonds were shown to significantly reduce the shear capability of the sandwich structure. Funding was provided by the ARES I Upper Stage Project. [ABSTRACT FROM AUTHOR]

Most composite damage tolerance assessments are made on bare laminates where the impactor comes into direct contact with the outermost ply. However, structures such as those used on launch vehicles are often covered with a thermal protection system (TPS) during the majority of the life of the part. This TPS covering may change the impact characteristics of the laminate rendering damage tolerance testing on bare laminates irrelevant to the part. This study examines the composite interstage structure of the ARES I launch vehicle which is scheduled to be covered with a sprayable foam TPS after manufacture. Damage tolerance testing is performed on bare sandwich structure and sandwich structure covered with the TPS selected for use on the ARES I composite interstage. Instrumented impact, infrared thermography, visual, cross-sectional and compression after impact CAI data are compared. Results show that the TPS covering does change most of the impact characteristics of the sandwich structure. It was found that the TPS created a larger damage zone as detected by IRT, however the TPS covered specimens possessed a higher residual compression strength for a given impact energy and damage size. These results are attempted to be explained by the different damage morphology that occurs between the bare and TPS covered specimens. [ABSTRACT FROM AUTHOR]

The issue of fatigue loading of structures composed of composite materials is considered in a requirements document that is currently in place for manned launch vehicles. By taking into account the short lives of these parts, coupled with design considerations, it is demonstrated that the necessary coupon level fatigue data collapse to a static case. Data from a literature review of past studies that examined compressive fatigue loading after impact and data generated from this experimental study are presented to support this finding. In other studies from the literature, a stress amplitude of about 60% of the static compression after impact (CAI) strength was found to exist, below which fatigue had no deleterious effects up to one million cycles. In this study, a stress amplitude of about 80% of the static (CAI) strength was found to exist, below which fatigue had no deleterious effects up to 10,000 cycles. A launch vehicle structure should never experience one cycle above 61.4% of static CAI strength, much less 10,000 at 80%. Despite utilizing severe fatigue amplitude loading in impact damaged coupons, residual strength after fatigue was consistently higher than expected. Unrealistically high fatigue stress amplitudes were needed to fail 5 of 15 specimens, before 10,000 cycles was reached. Since a typical launch vehicle structure, such as the ARES I interstage, only experiences a few cycles near limit load, it is concluded that static CAI strength data will suffice for most launch vehicle structures. [ABSTRACT FROM AUTHOR]