No abstract

Composite material usage is necessary on NASA’s future launch vehicles in order to obtain a low mass vehicle. While aircraft and launch vehicles that utilize load-bearing composite components have many similar damage tolerance requirements, the distinct differences between a part that has a lifetime of ∼500 s (one launch) and can be inspected in detail before use and one that has a lifetime of many tens of thousands of flight hours and can only undergo a ‘walk around’ inspection before each flight (commercial transport) needs to be taken into account. This article presents these differences and uses data from the ARES I composite interstage as an example of how to arrive at preliminary compression after impact strength values for the sandwich structure in the acerage of this part using residual strength curves. Results show that if severity of damage can be quantified by a nondestructive method (other than dent depth), the mass of the structure can be reduced due to better characterization of the damage.

A method of measuring the mode I (peeling) fracture toughness of core/face sheet bonds in sandwich structures is desired, particularly with the widespread use of models that need this data as input. This study examined if a mode I critical strain energy release rate, G IC , can be obtained from the climbing drum peel (CDP) test. The CDP test is relatively simple to perform and does not rely on measuring small crack lengths such as required by the more commonly used double cantilever beam (DCB) test. Simple energy methods were used to calculate G IC from CDP test data on composite face sheets bonded to a honeycomb core. Face sheet thicknesses from 2 to 5 plies (0.51-1.27 mm) were tested to examine the upper and lower bounds on face sheet thickness requirements. Suggestions on conducting the test and on modifying the CDP apparatus to test composite face sheets (as opposed to metallic) are also presented. Results from the study suggest that the CDP test, with certain provisions, can be used to find the G IC value of a core/face sheet bond.

No abstract

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.