An examination of the effect of composite constituent properties on the notched-strength performance of composite materials

The effect of matrix and fiber properties on the strengths and failure processes of notched, continuous-fiber composite specimens has been investigated. Bearing, open and filled hole tension and compression tests were conducted on AS-4/3501-6, IM-8/3501-6, AS-4/APC-2 and IM-8/APC-2 specimens. Failure characteristics of tested specimens were studied by fractographic techniques. Failure progression was tracked by means of incremental loading, radiography and fractographic inspection.     

Measurement of the static compressive strength of carbon-fibre/epoxy laminates

This paper describes an experimental study of the compressive failure of T800/924C carbon-fibre/efoxy composite laminates. Undirectional laminates loaded parallel to the fibres have compressive strengths that are 70% of the tensile strength and fail by fibre-microbuckling. During microbuckling the fibre debonds from the matrix, and the fibres break in bending. Multidirectional [(±45/0₂)₃]_sm laminates were also tested in compression, and the critical failure mechanism observed was microbuckling of the 0° plies. The failure strain was almost the same as for the undirectional laminate, The failure strain was almost the same as for the unidirectional laminate, which indicated that the ±45° plies have no significant influence on the failure strength of the 0° plies.     

Modeling and testing strain rate-dependent compressive strength of carbon/epoxy composites

A technique for testing high modulus fiber-reinforced composites in compression at different strain rates is investigated. The rate-dependent compressive behavior of unidirectional AS4/3501-6 carbon/epoxy composite is characterized by using off-axis specimens. It is found that, in the compression test, a titanium coating applied at the contact ends of the off-axis specimen can greatly reduce contact frictions, allowing a fully developed extension–shear coupling so that a state of uniform stress in the specimen can be achieved. A rate-dependent nonlinear constitutive model and a dynamic compressive strength model (fiber microbuckling model) for the unidirectional AS4/3501-6 composite are established based on the low strain rate off-axis test data. Model predictions and experimental data including high strain rate data are in very good agreement indicating that the constitutive model and compressive strength model obtained with low strain rate data are valid for high strain rates as well. A technique is also developed to extract the longitudinal compressive strength of the composite from those of the off-axis specimens.     

Failure of delaminated carbon/epoxy composite beams under cyclic compression

Experimental study of the response of laminated [(45/90/ — 45/0)₆]_s carbon/ epoxy virgin and pre-cracked composite beams to compressive cyclic loading is reported. The development of failure, in the form of delaminations, recorded continuously on video tape during the tests, revealed formation of delaminations close to the external faces of the beam. Some of the delaminations developed instantly and some propagated along the beam in jumps. Total failure occurred when the stress in the central ‘undamaged’ section of the beam reached a critical value. Calculation of the buckling load of unsymmetric laminated beams supports the experimental finding that a pre-crack close to the center of the specimen has no effect on the fatigue behavior of the beam. Prediction of time to first delamination crack, based on previous results, agree with the present experimental observation.     

基于Mises屈服准则和I1断裂准则的焊接接头力学特性分析

在全面考虑焊接残余应力、加载应力和氢致应力的情况下推导了Mises等效应力σ和应力第一不变量I₁的一般表达式;理论分析了残余应力和氢致应力对屈服行为的影响,推导了单向应力状态下的表观屈服应力[σ_s]的表达式;理论分析了残余应力和氢致应力对断裂行为的影响,推导了单向应力状态下的表观断裂应力[I_<i>b]的表达式和表观屈服时的应力第一不变量特征值[Is₁]的表达式。      

The effects of geometric parameters on the failure strength for pin-loaded multi-directional fiber-glass reinforced epoxy laminate

A numerical and experimental study was carried out to determine the failure of mechanically fastened fiber-reinforced laminated composite joints. E/glass–epoxy composites were manufactured to fabricate the specimens. Mechanical properties and strengths of the composite were obtained experimentally. Tests have been carried out on single pinned joints in [0/90/0]_s and [90/0/90]_s laminated composites. A parametric study considering geometries was performed to identify the failure characteristics of the pin-loaded laminated composite. Data obtained from pin-loaded laminate tests were compared with the ones calculated from a finite element model (PDNLPIN computer code). Damage accumulations in the laminates were evaluated by using Hashin's failure criteria combined with the proposed property degradation model. Based on the results, ply orientation and geometries of composites could be crucial for pinned laminated composite joints.     

Microdebonding investigation of the effects of thermal residual stress on the bond strength of a graphite/polymide composite

The microdebonding test was used to investigate the effects of thermal residual stresses resulting from different lay-ups in fabrication on the fiber/matrix bond strength of a graphite-fiber-reinforced polyimide composite. This was accomplished by comparing the results of a cross-plied laminate with those of a unidirectional laminate. The results indicated that the measured interfacial bond strength of the unidirectional composites was greater than that of the cross-plied laminate. The thermal radial stress distribution around the fiber for the unidirectional and the [0₂, 90₂]_s laminates were estimated, to explain this reduction of the interfacial bond strength.     

Prediction for delamination initiation around holes in symmetric laminates

The purpose of this paper is to present an analytical technique for predicting the delamination initiation load and delamination location for balanced symmetric laminates containing a hole. Basically, this approach includes two phases. First, the stress distribution around the hole region is calculated by the authors' recently developed efficient method. Second, the delamination initiation load and delamination location are predicted by using an average stress failure hypothesis together with the Hashin-Rotem delamination criterion. A computer program was developed to calculate the delamination initiation load and location for notched laminates with different ply orientations, different material properties and different stacking sequences. Numerical results for (θ/ – θ/0°)_s and (0°/90°)_s composite laminates generated by the current approach are compared with existing solutions. The results show that the present method is very efficient and useful for the purpose of engineering design.     

Experimental investigation of compression failure of sandwich specimens with face/core debond

An experimental study of the in-plane compressive failure mechanism of foam cored sandwich specimens with an implanted through-width face/core debond is presented. Tests were conducted on sandwich specimens with glass/vinylester and carbon/epoxy face sheets over various PVC foam cores. Observation of the response of the specimens during testing showed that failure occurred by buckling of the debonded face sheet, followed by rapid debond growth towards the ends of the specimen. The compression strength of the sandwich specimens containing a debond decreased quite substantially with increasing debond size. A high-density core resulted in less strength decrease at any given debond size. Examination of the failure surfaces after separation of the face sheet and core revealed traces of core material deposited on the face sheet evidencing cohesive core failure. The amount of core material adhered to the face sheet decreased with increasing foam density indicating increasing tendency for core/resin interfacial failure.     

Effects of gage length and stress concentration on the compressive strength of a unidirectional CFRP

Compression tests for a unidirectional CFRP were carried out for a wide range of gage lengths, three different configurations of end tab edges and two different end tab materials under the conditions of constant specimen thickness. The relation of σ_c-L is to be divided into two parts, namely, a part where the compressive strength, σ_c, is nearly constant independent of gage length, L, and beyond that, a part where σ_c decreases with increasing L. The apparent compressive strength measured by Celanese test method was lower than the true compressive strength because of the stress concentration near the end tab edges of the specimen. The true compressive strength was obtained by using the specimen where the gage lengths of 3.2 mm and 6.4 mm, and the material of end tab is stainless steel.     

A seawater tank approach to evaluate the dynamic failure and durability of E-glass/vinyl ester marine composites

Durability and dynamic failure properties are critical parameters for naval composite ships in seawater. However, previous measurements have significantly underestimated the intrinsic durability of composite structures. In this study, a simple composite seawater tank approach is developed to accurately simulate the material/mechanics conditions of composite structures in seawater. Only one surface of each composite specimen is exposed to seawater and this exposed surface is subjected to drop-weight impact testing followed by characterization of residual compressive strength. A series of experiments on dry and wet composite specimens were conducted to characterize the impact damage, and the residual compressive strength. Results from seawater exposure indicate that the compression-after-impact strengths of the wet specimens reduced by around 10% compared to the baseline dry specimens over 29-month seawater exposure. Therefore, durability of current marine composite is better than the previously measured properties. Our simple seawater tank durability experiment is expected to provide accurate justification for the required maintenance period of new composite ship structures.     

Life prediction for fatigue of T800/5245 carbon-fibre composites: I. Constant-amplitude loading

This paper presents a study of the fatigue behaviour of a [(±45,0₂)₂]_s laminate of the T800/5245 composite system. This modern aerospace material consists of high-failure-strain, intermediate-modulus carbon fibres in a toughened bismaleimide resin system. The fatigue response in repeated tension, repeated compression, and mixed tension-compression was determined at constant stress levels over a wide range of R values, and the results are analysed in terms of constant-life diagrams of the kind that can be used to provide design data. The use of such information for life prediction purposes is discussed.     

An experimental investigation of the biaxial strength of IM6/3501-6 carbon/epoxy cross-ply laminates using cruciform specimens

Several variations of a thickness-tapered cruciform specimen have previously been used to experimentally determine the biaxial strength of an AS4/3501-6 carbon/epoxy cross-ply laminate. The present work represents a follow-up study of the original specimen design, and incorporates numerous specimen improvements made in an attempt to generate more accurate biaxial results. A total of 52 tests were performed at numerous biaxial stress ratios, utilizing six different specimen configurations. The experimental data generated in the present study for all specimen geometries, as well as a complete biaxial failure envelope in σ₁–σ₂ stress space for this laminate configuration, are presented. A desirable failure mode in the gage section of the specimen was achieved for all specimens tested in the present study, indicating that accurate biaxial stress states were being generated at ultimate specimen failure. The ability of the thickness-tapered cruciform specimen to determine the biaxial strength of composite materials at any stress ratio has been demonstrated.     

Delamination threshold loads for low velocity impact on composite laminates

An investigation of the low velocity impact load level at which a composite laminate will delaminate is presented. The delamination threshold load is described as the load level, obtained from the load–time history or load–displacement plot, at which a sudden load drop occurs due to specimen stiffness loss as a result of laminate level damage. Approximately 500 low velocity impact load–time histories from the Air Force Research Laboratory (AFRL) Low Velocity Impact Database are used to investigate the delamination threshold load. The database contains laminate impact test data for graphite/epoxy (AS4/3501-6), graphite/PEEK (AS4/APC-2), and graphite/BMI (IM7/5260) material systems. The delamination threshold load observations are compared to C-scan damage measurements of impacted specimens to determine if the sudden load drop corresponds to delamination development.     

基于应变能耗散的复合材料层合板面内缺口强度分析CDM模型

基于伴随能量释放的渐进损伤演化思想,建立了复合材料层合板面内失效分析的连续介质损伤力学（CDM）分析模型,该模型包含损伤表征、损伤起始判定和损伤演化法则3个方面。基于CDM模型,通过引入损伤状态变量表征损伤,建立了平面应力状态下的材料损伤本构模型。采用损伤参量 f_E改写Hashin准则,以判定损伤的起始。损伤演化由特征长度内的应变能释放密度控制,建立了损伤状态变量关于等效应变的渐进损伤演化法则。模型中还同时考虑了面内剪切非线性和网格敏感性,并进行了对比分析。对含缺口的[90/0/±45]_3s和[（±θ）₄]_s 2类典型复合材料层合板的面内拉伸失效进行了分析,结果表明,本文中的模型能有效预测复合材料层合板的面内拉伸强度。     

Nonlinear analysis of interlaminar stress in graphite/epoxy laminates with and without delamination cracks

A quasi three-dimensional yield function, which is quadratic in stresses except for σ₁₁, is proposed for graphite/epoxy composites. The elastic-plastic interlaminar stress response near a free edge in the [90/0]_s, [0/90]_s, and [45/−45]_s laminates with and without delamination cracks was investigated using the pseudo three-dimensional finite element technique. The plasticity model was evaluated by comparison with off-axis experimental data. Since shear response is the key element for nonlinear stress-strain behavior of graphite/epoxy composites, the plasticity theory predicts interlaminar stresses in the [45/−45]_s laminate significantly different from linear elasticity. Moreover, the existence of a delamination crack caused more plasticity effects on interlaminar stresses.     

2-D biaxial testing and failure predictions of IM7/977-2 carbon/epoxy quasi-isotropic laminates

In previous research, a series of a thickness-tapered cruciform specimen configurations have been used to determine the biaxial (two-dimensional, in-plane) and triaxial (three-dimensional) strength of several carbon/epoxy and glass/vinyl-ester laminate configurations. Refinements to the cruciform geometry have been shown capable of producing acceptable results for cross-ply laminate configurations. However, the presence of a biaxial strengthening effect in quasi-isotropic, [(0_N/90_N/ ± 45_N)_M]_S, laminates have brought into question whether the cruciform geometry could be used to successfully generate two-dimensional strength envelopes. In the present study, a two-dimensional failure envelope for a IM7/977-2 carbon/epoxy laminate was developed at the Air Force Research Laboratory, Space Vehicles Directorate, using a triaxial test facility. The electromechanical test frame is capable of generating any combination of tensile or compressive stresses in σ₁:σ₂:σ₃ stress space and can evaluate the uniaxial (one-dimensional, in-plane), biaxial or triaxial response of composite materials. Results are promising as they indicated that failure in the majority of the IM7/977-2 specimens occurred in the gage section. This leads the authors to believe that maximum biaxial stress states were correctly generated within the test specimen. In addition to the experimental data presented, multi-continuum theory (MCT) was used to predict and analyze the onset of damage and ultimate failure of a biaxially loaded IM7/977-2 laminate. Multi-continuum theory is a micromechanics based theory and associated numerical algorithm for extracting, virtually without a time penalty, the stress and strain fields for a composites’ constituents during a routine structural finite element analysis. Damage in a composite material typically begins at the constituent level and may, in fact, be limited to only one constituent in some situations. An accurate prediction of constituent failure at sampling points throughout the laminate provides a genesis for progressively analyzing damage propagation in a composite specimen allowing identification of intermediate damage modes. A constituent-based, quadratic, stress-interactive, failure criterion was used to take advantage of the micro-scale information provided by MCT. There was reasonable correlation between analytically and experimentally developed IM7/977-2 2D failure envelope which leads us to believe that the thickness-tapered cruciform specimen can be used to determine the biaxial strength of quasi-isotropic laminates.     

Effect of different loading conditions on the mechanical behavior of [0/±45/90]s woven composites

The main objective of this paper is to investigate the behavior of [0/±45/90]_s woven FRP composites under tension, bending, and combined bending/tension loading conditions. First, the mechanical properties of the composite were determined experimentally using the ASTM testing standards. Bending properties were determined using 3-point and 4-point bending tests. The results showed that the woven composites performed better under bending loading than under tension loading. Finally, special test fixtures were designed to facilitate the study of the effect of the combined bending/tension loading. The bending moments were applied using offset shims of various thicknesses placed between the plane of the specimen and the loading axis. At the beginning, the load–strain diagrams at the specimen center showed the domination of bending strains, tension on one surface and compression on the other. With the advance of the loading process, the tension strain dominated and the strain on both sides were almost equal. The failure under combined bending/tension loading was due to the high stresses near the fixture. However, in pure bending, the material failed at the center because of the excessive delamination on the compressive side.     

Life prediction for fatigue of T800/5245 carbob-fibre composites: II. Variable-amplitude loading

This is the second of two papers describing the fatigue response of a [(±45,0₂)₂]_s laminate of the T800/5245 composite system, a modern aerospace material consisting of high-failure-strain, intermediate-modulus carbon fibres in a toughened bismaleimide resin system. In the first paper, the fatigue response in repeated tension, repeated compression, and mixed tension-compression was determined at constant stress levels over a wide range of R values. The results of those constant-amplitude experiments were then used to define a programme of four- and two-block variable-loading experiments in an attempt to derive predictive methods for such loading conditions. Formulae have been developed to predict life under non-linear cumulative damage conditions and empirical data derived to validate the procedures.     

Behavior of notched and unnotched [0/±30/±60/90]s GFR/EPOXY composites under static and fatigue loads

The main objective of the present paper is to study the bending behavior of notched and unnotched angle-ply, [0/±30/±60/90]_s, glass fiber reinforced epoxy (GFRE) composites under static and fatigue loads. Static and fatigue bending properties have been determined for notched and unnotched angle-ply specimens. For this purpose different circular notch sizes (2, 4.5, 7, 9 mm) were drilled at the specimen center. Constant-deflection bending fatigue tests were performed at zero mean stress and 25 Hz. A 15% reduction of the initial applied moment was taken as a failure criterion. S–N diagrams for notched GFRE specimens have been constructed based on gross and net cross-section area. The results show that the ultimate bending strength of notched GFRE specimens decreased linearly with increasing notch diameter. Based on gross-section the fatigue life increases with decreasing notch size and the longer fatigue life was for the unnotched specimens. On the other hand, the S–N diagrams based on net-section indicate the insensitivity of angle-ply composites to the notch size. This is considered to be a peculiar phenomenon to composite materials. The results also show that the S–N diagrams have not any fatigue limit rigorous within 10⁷ cycles.