Clarence C. Poe

Open hole and postimpact compressive fatigue of stitched and unstitched carbon-epoxy composites

The performance is studied of a stitched uniweave fabric composite and that of a toughened tape composite. The effects of stitching on compression fatigue life are addressed. Post impact compression fatigue and open hole fatigue tests were run on an AS4/3501-6 uniweave with stitching and a toughened IM7/8551-7 tape without stitching. Stitching was found to increase the thickness and consequently the weight of the composite material. The two materials were compared on an equal carbon content basis as well as on an equal weight basis. The excess thickness in the stitched uniweave composite was responsible for the lower fatigue life, on an equal carbon basis, compared to the toughened resin tape composite. Comparison of fatigue lives on an equal carbon content basis indicated that puncture or crimp type damage from stitching has very little effect on compression failure. Post impact fatigue test showed that although the damage area in the stitched uniweave composite was twice that of the toughened tape composite, the fatigue lives of the stitched composite were significantly longer than those of the toughened composite. Thus, it appears that the increase in thickness from stitching is much more of a penalty than crimped fibers or puncture type damage from stitching.

Strength of a thick graphite/epoxy rocket motor case after impact by a blunt object

The National Aeronautics and Space Administration is developing graphite/epoxy filament-wound cases (FWC) for the solid rocket motors of the Space Shuttle. The membrane region is about 36 mm thick. A study was made to determine the reduction in strength of the FWC due to accidental damage caused by low-velocity impacts. Two 76.2 cm diameter by 30.5 cm long cylinders were impacted every 5 cm of circumference with 1.27 cm radius impacters of various mass. The impacters represented tools and equipment dropped from various heights. One cylinder was empty and the other was filled with inert propellant. Five cm wide test specimens were cut from the cylinder. Each was centered on an impact sight. The specimens were X-rayed and loaded to failure in uniaxial tension. The strengths and depths of impact damage were analyzed in terms of maximum impact force. Rigid body mechanics and the Hertz law were used to derive an equation for impact force in terms of kinetic energy and the masses of the impacter and target. The depth of damage was predicted in terms of impact force using Loves solution of pressure applied on part of the boundary of a semi-infinite body.

Analytical and Experimental Studies of the Debonding of Stitched and Unstitched Composite Joints

The effect of stitches on the failure of a single lap joint configuration was determined in a combined experimental and finite element study. The experimental program was conducted to determine debond growth under static monotonic loading. The stitches were shown to delay the initiation of the debond and provide load transfer after total debonding of the lap joint. The experimentally determined debond length versus applied load was used as an input parameter in the finite element analysis of both stitched and unstitched configurations. The strain energy release rates at the debond front were calculated using plate elements. Discrete nonlinear spring elements were used to model the stitches and multipoint constraints were used to model the contact problem. Models of the unstitched configuration showed significant values of mode I and II strain energy release rates across the width of the joint and showed that mode III is zero at the centerline but increases near the free edge. Models of the stitched configuration showed that the stitches were effective in reducing the mode I component to zero, but had less effect on modes II and III.

Relevance of impacter shape to nonvisible damage and residual tensile strength of a thick graphite/epoxy laminate

A study was made to determine the relevance of impacter shape to nonvisible damage and tensile residual strength of a 36 mm (1.4 in.) thick graphite/epoxy motor case. The shapes of the impacters were as follows: 12.7 mm (0.5 in.) and 25.4 mm (1.0 in.) diameter hemispheres, a sharp corner, and a 6.3 mm (0.25 in.) diameter bolt-like rod. The investigation revealed that damage initiated when the contact pressure exceeded a critical level. However, the damage was not visible on the surface until an even higher pressure was exceeded. The damage on the surface consisted of a crater shaped like the impacter, and the damage below the surface consisted of broken fibers. The impact energy to initiate damage or cause visible damage on the surface increased approximately with impacter diameter to the third power. The reduction in strength for nonvisible damage increased with increasing diameter, 9 and 30 percent for the 12.7 mm (0.5 in.) and 25.4 mm (1.0 in.) diameter hemispheres, respectively. The corner impacter made visible damage on the surface for even the smallest impact energy. The rod impacter acted like a punch and sliced through the composite. Even so, the critical level of pressure to initiate damage was the same for the rod and hemispherical impacters. Factors of safety for nonvisible damage increased with increasing kinetic energy of impact. The effects of impacter shape on impact force, damage size, damage visibility, and residual tensile strength were predicted quite well assuming Hertzian contact and using maximum stress criteria and a surface crack analysis.

An evaluation of the pressure proof test concept for thin sheet 2024-T3

Abstract The concept of pressure proof testing of fuselage structures with fatigue cracks to insure structural integrity was evaluated from a fracture mechanics viewpoint. A generic analytical and experimental investigation was conducted on uniaxially loaded flat panels with crack configurations and stress levels typical of longitudinal lap-slice joints in commercial transport aircraft fuselage. The results revealed that the remaining fatigue life after a proof test was longer than that without the proof test because of crack growth retardation due to increased crack closure. However, based on a crack length that is slightly less than the critical value at the maximum proof test stress, the minimum assured life or proof test interval must be no more than 550 pressure cycles for a 1.33 proof factor and 1530 pressure cycles for a 1.5 proof factor to prevent in-flight failures.

Surface crack analysis applied to impact damage in a thick graphite/epoxy composite

The residual tensile strength of a thick graphite/epoxy composite with impact damage was predicted using surface crack analysis. The damage was localized to a region directly beneath the impact site and extended only part way through the laminate. The damaged region contained broken fibers, and the locus of breaks in each layer resembled a crack perpendicular to the direction of the fibers. In some cases, the impacts broke fibers without making a visible crater. The impact damage was represented as a semi-elliptical surface crack with length and depth equal to that of the impact damage. The maximum length and depth of the damage were predicted with a stress analysis and a maximum shear stress criterion. The predictions and measurements of strength were in good agreement.

Estimating Residual Strength in Filament Wound Casings from Nondestructive Evaluation of Impact Damage

The purpose of this study is to improve the ability to detect hidden impact damage in thick composites caused by low velocity impact and to predict the remaining strength of those materials. An impact study has been undertaken on filament wound graphite/epoxy casings, such as those proposed for NASA’s space shuttle solid fuel rocket boosters. In thick composite materials, low-velocity impact damage may not be visually evident, depending on the impacter shape; yet the damage may compromise the composite’s ultimate strength. A model of a filament wound casing was fabricated with one fifth of the diameter (30 inches) but with the full thickness (1.4inches) of the full rocket motor (12 feet and 1.4 inches, respectively). It was impacted with various masses and energy levels using a one inch diameter ball as the indenter. This casing was subsequently cut into coupons of 2 in. width by 12 in. length. These samples were nondestructively examined for the degree of damage. Next, these samples were loaded in tension until failure. Efforts to accurately detect the damage with dye penetrants and x-ray methods have proven unsatisfactory in the samples that displayed no visible damage. In spite of the high attenuation of this material, ultrasonic phase velocity and attenuation images show promise in predicting the residual strength of the coupons. Predictions of the damage profile, and therefore the cross-section of the damage in the direction of loading, were obtained by assuming an “effective” value for the attenuation of the damaged part of the filament wound casing material (15 dB/MHz-cm) and an “effective” value for the velocity of the damaged part of the filament wound casing material (2250 m/s). These estimates were based partially on measurements made on impact damaged thin composite material. The remaining strength predictions from these ultrasonic data showed a significant improvement over the x-ray predictions of remaining strength and the method may be usable for predictions of remaining strength of full scale rocket motors that may have suffered impact damage.

Strain Intensity Factor Approach for Predicting the Strength of Continuously Reinforced Metal Matrix Composites

A method was previously developed to predict the fracture toughness (stress intensity factor at failure) of composites in terms of the elastic constants and the tensile failing strain of the fibers. The method was applied to boron/aluminum composites made with various proportions of 0 deg and +/- 45 deg plies. Predicted values of fracture toughness were in gross error because widespread yielding of the aluminum matrix made the compliance very nonlinear. An alternate method was develolped to predict the strain intensity factor at failure rather than the stress intensity factor because the singular strain field was not affected by yielding as much as the stress field. Far-field strains at failure were calculated from the strain intensity factor, and then strengths were calculated from the far-field strains using uniaxial stress-strain curves. The predicted strengths were in good agreement with experimental values, even for the very nonlinear laminates that contained only +/- 45 deg plies. This approach should be valid for other metal matrix composites that have continuous fibers.

A Plate Element-Based Model for Mixed-Mode Debonding of Composite Stitched Stiffened Panels

An analysis based on plate finite elements and the virtual crack closure technique is used to study the effect of stitching on Mode I and Mode II strain energy release rates for a stitched warp-knit composite debond configuration. The stitches were modeled as discrete nonlinear fastener elements with a compliance determined by experiment. The axial and shear behavior of the stitches was considered with both the compliances and failure loads assumed to be independent. The effect of model slenderness ratio on the accuracy of the strain energy release rates determined with the plate element models for configurations without stitching was determined by comparison with similar plane strain models. The analysis showed that stitches are very effective in reducing Mode I strain energy release rate, G I , by closing the debond faces near the debond front; however, they are less effective in reducing the Mode II strain energy release rate, G I I .

An Evaluation of the Pressure Proof Test Concept for 2024-T3 Aluminium Alloy Sheet

The concept of pressure proof testing of fuselage structures with fatigue cracks to insure structural integrity was evaluated from a fracture mechanics viewpoint. A generic analytical and experimental investigation was conducted on uniaxially loaded flat panels with crack configurations and stress levels typical of longitudinal lap splice joints in commercial transport aircraft fuselages. The results revealed that the remaining fatigue life after a proof cycle was longer than that without the proof cycle because of crack growth retardation due to increased crack closure. However, based on a crack length that is slightly less than the critical value at the maximum proof stress, the minimum assured life or proof test interval must be no more than 550 pressure cycles for a 1.33 proof factor and 1530 pressure cycles for a 1.5 proof factor to prevent in-flight failures.