Alan T. Nettles

Compression After Impact Testing of Sandwich Composites for Usage on Expendable Launch Vehicles

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.

Using the Climbing Drum Peel (CDP) Test to Obtain a GIC Value for Core/Face Sheet Bonds

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, GIC, 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 GIC 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 GIC value of a core/face sheet bond.

A Comparison of Quasi-Static Indentation Testing to Low Velocity Impact Testing

The need for a static test method for modeling low-velocity foreign object impact events to composites would prove to be very beneficial to researchers since much more data can be obtained from a static test than from an impact test. In order to examine if this is feasible, a series of static indentation and low velocity impact tests were carried out and compared. Square specimens of many sizes and thickness were utilized to cover the array of types of low velocity impact events. Laminates with a n/4 stacking sequence were employed since this is by the most common type of engineering laminate. Three distinct flexural rigidities under two different boundary conditions were tested in order to obtain damage due to large deflections, contact stresses and both to examine if the static indentation-impact comparisons are valid under the spectrum of damage modes that can be experienced. Comparisons between static indentation and low velocity impact tests were based on the maximum applied transverse load. The dependent parameters examined included dent depth, back surface crack length, delamination area and to a limited extent, load-deflection behavior. Results showed that no distinct differences could be seen between the static indentation tests and the low velocity impact tests, indicating that static indentation can be used to represent a low velocity impact event.

Compression after impact strength of out-of-autoclave processed laminates

Out-of-autoclave processable fiber/resin systems have been gaining much attention due to the elimination of needing a costly autoclave large enough to hold the part to be cured. For large composite structures this can pose a challenge. However, for these fiber/resin systems to replace conventional autoclave fiber/resin systems for use on aerospace structures, the damage tolerance capabilities need to meet (or exceed) those of current autoclave fiber/resin systems. In this experimental study, compression-after-impact strengths of two commercially available out-of-autoclave fiber/resin systems are compared to compression-after-impact strengths of a conventional autoclave fiber/resin system used as a baseline in this study. compression-after-impact testing was chosen since this is the most common method to assess laminates damage tolerance capabilities. Three different levels of impact severity were chosen and information on damage size and morphology are assessed along with compression-after-impact strength values. The results show the two out-of-autoclave fiber/resin systems examined in this study have similar damage tolerance characteristics to the autoclave fiber/resin system used in this study.

An Examination of the Compressive Cyclic Loading Aspects of Damage Tolerance for Polymer Matrix Launch Vehicle Hardware

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.

Compression after impact strength of thin laminates with various percentage of 0° plies

Conventional wisdom dictates that adding more 0° plies in the load-bearing direction of a laminate will increase its stiffness and strength. While this is true for undamaged laminates, the compression strength of laminates with impact damage may not be as straightforward. In this study, compression after impact strengths of relatively thin laminates with 25%, 33% or 50% of plies aligned in the 0° load-bearing direction were measured for three different damage severity levels. Results show that the increase in compression strength of the laminates with a higher percentage of plies in the 0° direction is lessened as impact damage severity increases indicating that a laminate that is stronger in compression when undamaged may not be stronger in compression when impact damage is accounted for.

Notched compression strength of 18-ply laminates with various percentages of 0° plies

The use of knockdown factors (percent reduction of undamaged compression strength) to account for flaws such as impact damage or holes have been used to infer the notched strength of laminates. It has been observed that this criterion tends to over-predict the strength of laminates with a high percentage of 0° plies. This paper examines some limited data from the literature and presents new data that compares knockdown calculated notched compression strength values with those measured experimentally for laminates with various percentages of 0° plies. Results show that the trend of over-predicting the notched compression strength of laminates as the percentage of 0°plies increases, based on a knockdown factor, is observed, but the difference can be within scatter except at very high percentages of 0° plies.

Change in damage tolerance characteristics of sandwich structure with a Thermal Protection System (TPS)

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.

The Influence of GI and GII on the compression after impact strength of carbon fiber/epoxy laminates

This study measured the compression after impact strength of IM7 carbon fiber laminates made from epoxy resins with various mode I and mode II toughness values to observe the effects of these toughness values on the resistance to damage formation and subsequent residual compression strength-carrying capabilities. A total of seven different epoxy resin systems were used ranging in approximate GI values of 245–665 J/m2 and approximate GII values of 840–2275 J/m2. The results for resistance to impact damage formation showed that there was a direct correlation between GII and the planar size of damage as measured by thermography. Subsequent residual compression strength testing suggested that GI had no influence on the measured values and most of the difference in compression strength was directly related to the size of damage. Thus, delamination growth assumed as an opening type of failure mechanism does not appear to be responsible for the loss of compression strength.

Normalizing impact energy by face sheet thickness for composite sandwich structure compression after impact testing

The amount of impact energy used to damage a composite laminate face sheet of a sandwich structure is a critical parameter when assessing residual compression strength. The compression after impact strength of impacted laminates used as face sheets on honeycomb core sandwich structure is dependent upon how thick the face sheet laminate is and this has traditionally been accounted for by normalizing (dividing) the impact energy by the laminate’s thickness. However when comparing compression after impact strength values for a given lay-up sequence and fiber/resin system, dividing the impact energy by the specimen thickness has been noted by the author to give higher compression after impact strength values for thicker face sheet laminates. A study was thus undertaken to assess the comparability of compression after impact strength data of sandwich structure by normalizing the impact energy by the face sheet thickness raised to a power to account for the higher strength of thicker laminates. Two data sets generated in this study were analyzed by dividing the impact energy by the face sheet thickness to the 1.0, 1.5, 2.0 and 2.5 powers. Results show that raising the face sheet thickness to a power of approximately 2.5 and dividing the impact energy by this quantity yields more comparable compression after impact strength data for comparing 8- and 16-ply face sheet laminates. For comparison of 24-ply face sheet laminates to 8- or 16-ply face sheet laminates, a value closer to 2 was found to give more comparable compression after impact strength data.