R. Kim

Effect of stacking sequence on micro-cracking in a cryogenically cycled carbon/bismaleimide composite

An apparatus was developed to thermally cycle coupon-sized mechanical test specimens to −196 °C. Using this device, IM7/5250-4 carbon/bismaleimide cross-ply ([0/90]2S and [90/0/90/0/90/0/90/0/90]) and quasi-isotropic ([0/45/−45/90]S) laminates were submerged in liquid nitrogen (LN2) and returned to room temperature 400 times. Ply-by-ply micro-crack density (transverse cracks), micro-crack span, laminate modulus, and laminate strength were measured as a function of thermal cycles. The composite micro-cracked extensively in the surface plies followed by sparse micro-cracking of the inner plies. The tensile strength of the two blocked lay-ups (lay-ups with adjacent plies of the same orientation) decreased by 8.5% after 400 cycles. Sectioning of the samples revealed that the micro-cracks in the surface plies spanned the full width of the sample while many of the micro-cracks observed on the edge of the inner plies did not extend to the center of the samples, implying that a rectangular specimen with exposed free edges may result in a significantly different micro-crack density than a sample without free edges.

Fiber–Matrix Interfacial Failure Characterization Using a Cruciform-Shaped Specimen

In this study, a cruciform-shaped test specimen is utilized to characterize the fiber–matrix interface under transverse and combined (tensile and shear) loading. We first present an overview of past references of how the cruciform geometry is optimized to promote interfacial failure. We then discuss a modification of the cruciform specimen where face-sheets are adhesively bonded to reinforce the sample. These face-sheets serve a twofold purpose, namely, to prevent premature failure in the fillet region and to encourage debond initiation at the center of the gage length. Finally, an off-axis cruciform geometry, in which the wings of the cruciform sample are inclined at an angle with respect to the loading direction, is introduced to characterize the fiber–matrix interface under combined transverse and shear loading. Using the measured value of applied stress at debond initiation, and the evaluated stress concentration factor at the fiber–matrix interface, a mixed-mode failure envelope is then constructed in the normal-shear stress space, and a quadratic failure criterion is proposed.

Strength Prediction and Measurement in Model Multilayered Discontinuous Tow Reinforced Composites

The purpose of this work is to perform analytical modeling and experimentally produce idealized composite laminates reinforced with high-modulus discontinuous-fiber tows to evaluate the theoretical upper limits of tensile strength as compared to continuous-fiber composites. The idealized composite represents a staggered mosaic of prepreg tape strips of equal width and length. Three-dimensional analysis was performed to evaluate the energy release rate of the possible damage accumulation modes, such as tow delamination and splitting emanating from tape strip ends in plies with different orientations. The effect of residual stress was also evaluated on the delamination propagation energy release rates as well as tow end crack formation loads. Failure mechanisms for unidirectional and quasi-isotropic laminates were established by comparing the calculated energy release rates for tows in different locations to the critical Mode II energy release for axial cracking in the material system. Specially designed experiments were conducted to verify the failure mechanism predictions and evaluate proposed design changes for strength increase in the discontinuous-fiber tow reinforced composites.

Monitoring of Impact Damage in Composites Using Wave Propagation Methods

Composite structures in an aircraft are susceptible to impact damage, which can occur during manufacture, service or maintenance. Recent studies show that impacts with ground support equipment are the major cause of in-service damage to composite structures in an aircraft. Other sources of impact include collision with birds, runway stones or ballistic impacts. These impacts can produce various types of damage, including fiber breakage, matrix cracking, delamination, and interfacial debonding. The results of such damage can have detrimental effects on the overall structural performance and safety. A comprehensive structural health monitoring (SHM) system provides a means to significantly reduce life-cycle costs of aerospace vehicles by providing accurate diagnostics and prognostics of structural damage to reduce unnecessary inspections and support vehicle life extension. The main objective of this paper is to develop a methodology to detect and identify the damage sources and their severity in composite laminates subjected to low velocity impact using wave propagation methods. When damage occurs in a material due to mechanical load or impact, an acoustic wave emits and propagates through the material. The material chosen for this work is a 12″ long and 12″ wide, +/− 60 degree braided composite. Two edges of the plate were fixed by clamping the plate between two steel bars and secured by bolts spaced 1″ apart, while the other two edges were free, as shown in Figure 1. In order to characterize the wave propagation and damage process, two resonant type AE sensors and four accelerometers were mounted on the specimen. The specimen was then tapped lightly with a hand-held acoustic impact hammer at several different chosen locations, and stress wave signals were monitored using a commercial dynamic signal process system which contains software capable of detecting impact source location. The impact force was kept to a minimum initially such that no damage occurred in the specimen. After this initial test, the specimens were subjected to low velocity impact using drop weight impact machine with 0.5 inch spherical indenter. The impact force was increased by a number of times until substantial damage observed while monitoring signals generated from the specimen. After each incremental impact, both acoustic hammer tapping test and nondestructive inspection such as ultrasonic C-scan and/or X-ray radiography were carried out to delineate the damage source and severity. Figure 2 is an example of C-Scan of the composite plate after a series of impacts with various drop heights. Recorded signals were analyzed to determine the origin of the source and its severity. The impact hammer produced both an extensional wave and a flexural wave in these composite plate specimens. Because of dispersive characteristics of the flexural wave, the first arrival time of the extensional wave was used for source location algorithm. Besides the source location, discussion will be given on parameters such as amplitude, energy, frequency, number of events related with impact force, and damage size in detail. As an example, Figure 3 is a plot of the measured damage size as a function of the dead-weight drop height for tests conducted on various panels. As expected, the size of the damage increases with amount of drop height (or impact energy). Thus, based on C-scan measurements, critical threshold impact height of approximately 5″ is identified for “any measurable” damage to occur. The corresponding magnitude of the impact energy is ∼ 108 in-lb. On the other hand, the critical threshold for any visual damage to be detected is approximately 502 in-lb for the laminate material investigated. In summary, a methodology has been developed for estimating the damage severity from the amplitude of the signal received. The approach entails constructing design curves relating the size of the damage to impact energy, and establishing relationships between impact energy and the magnitude of the signal. These relationships can then be used to predict the estimated size of the damage based on the amplitude of the arriving signal. A critical threshold impact energy has been identified below which “no measurable” damage occurs. Three regions of damage growth, namely, a decreasing rate with magnitude of impact energy. A constant damage growth rate characterizes the steady-state region, while damage size increases almost exponentially with impact energy in the tertiary region potentially leading to catastrophic failure.© 2007 ASME

Ply-Level Behavior of Carbon/Epoxy Composites Mechanically Cycled at Cryogenic Temperatures

Polymer matrix composites are candidate materials in tanks, lines, ducts, and valves used to store and transport cryogenic liquids. The use of these materials at cryogenic temperatures necessitates an understanding of the fatigue life and microcracking extent (which could effect permeation and leakage) at these reduced temperatures. This study presents ply-level (90° and +/-45°) fatigue life test results for Cytec IM7/977-2, a carbon fiber reinforced toughened epoxy. The test results showed a power-law relationship between the applied stress level and the cycles to failure for both the 90° and +/-45° laminates, and for the three test temperatures of –321°F, room temperature, and 300°F. In addition, for both stacking sequences, the laminates tested at the cryogenic temperature showed longer fatigue lives, for a given cyclic stress level, than those tested at room temperature as well as elevated temperature.