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All sourcesNASA NTRSDOE OSTI
NASA NTRS2009

Mid-Infrared Reflectance Imaging of Thermal-Barrier Coatings

An apparatus for mid-infrared reflectance imaging has been developed as means of inspecting for subsurface damage in thermal-barrier coatings (TBCs). The apparatus is designed, more specifically, for imaging the progression of buried delamination cracks in plasma-sprayed yttria-stabilized zirconia coatings on turbine-engine components. Progression of TBC delamination occurs by the formation of buried cracks that grow and then link together to produce eventual TBC spallation. The mid-infrared reflectance imaging system described here makes it possible to see delamination progression that is invisible to the unaided eye, and therefore give sufficiently advanced warning before delamination progression adversely affects engine performance and safety. The apparatus (see figure) includes a commercial mid-infrared camera that contains a liquid-nitrogen-cooled focal plane indium antimonide photodetector array, and imaging is restricted by a narrow bandpass centered at wavelength of 4 microns. This narrow wavelength range centered at 4 microns was chosen because (1) it enables avoidance of interfering absorptions by atmospheric OH and CO2 at 3 and 4.25 microns, respectively; and (2) the coating material exhibits maximum transparency in this wavelength range. Delamination contrast is produced in the midinfrared reflectance images because the introduction of cracks into the TBC creates an internal TBC/air-gap interface with a high diffuse reflectivity of 0.81, resulting in substantially higher reflectance of mid-infrared radiation in regions that contain buried delamination cracks. The camera is positioned a short distance (.12 cm) from the specimen. The mid-infrared illumination is generated by a 50-watt silicon carbide source positioned to the side of the mid-infrared camera, and the illumination is collimated and reflected onto the specimen by a 6.35-cm-diameter off-axis paraboloidal mirror. Because the collected images are of a steady-state reflected intensity (in contrast to the transient thermal response observed in infrared thermography), collection times can be lengthened to whatever extent needed to achieve desired signal-to-noise ratios. Each image is digitized, and the resulting data are processed in several steps to obtain a true mid-infrared reflectance image. The raw image includes thermal radiation emitted by the specimen in addition to the desired reflected radiation. The thermal-radiation contribution is eliminated by subtracting the image obtained with the illumination off from the image obtained with the illumination on. A flat-field correction is then made to remove the effects of non-uniformities in the illumination level and pixel-to-pixel variations in sensitivity. This correction is performed by normalizing to an image of a standard object that has a known reflectance at a wavelength of 4 microns. After correction, each pixel value is proportional to the reflectance (at a wavelength of 4-microns) at the corresponding location on the specimen. Mid-infrared reflectance imaging of specimens that were thermally cyc led for different numbers of cycles was performed and demonstrated t hat mid-infrared reflectance imaging was able to monitor the gradual delamination progression that occurs with continued thermal cycling . Reproducible values were obtained for the reflectance associated w ith an attached and fully delaminated TBC, so that intermediate refle ctance values could be interpreted to successfully predict the numbe r of thermal cycles to failure.

Edlridge, Jeffrey I.↗
NASA NTRS2020

Challenges of Testing EB Coated CMC under SPLCF Conditions in Steam Environment at High Temperatures

Environmental barrier (EB) coated monolithic SiC (Hexoloy) and 3D woven SiC/SiC composite specimens were tested under sustained peak low cycle fatigue (SPLCF) conditions at 1204°C at 69 MPa (maximum stress) in steam environment up to 200hrs. Both types of specimens were coated with a multilayer EB coating (Mullite + rare earth disilicate (RED) bond coat and RED top coat) by slurry method. The coated Hexoloy specimen survived 200hrs testing of testing, but the coating showed through-the-thickness cracks and delamination. In contrast, the coated 3D woven SiC/SiC composite specimens failed between 48 and 160 hrs and showed extensive damage in the coating as well as within the CMC substrate.

R T Bhatt↗
NASA NTRS2004

Discrete Spring Model for Predicting Delamination Growth in Z-Fiber Reinforced DCB Specimens

Beam theory analysis was applied to predict delamination growth in DCB specimens reinforced in the thickness direction with pultruded pins, known as Z-fibers. The specimen arms were modeled as cantilever beams supported by discrete springs, which were included to represent the pins. A bi-linear, irreversible damage law was used to represent Z-fiber damage, the parameters of which were obtained from previous experiments. Closed-form solutions were developed for specimen compliance and displacements corresponding to Z-fiber row locations. A solution strategy was formulated to predict delamination growth, in which the parent laminate mode I fracture toughness was used as the criterion for delamination growth. The solution procedure was coded into FORTRAN 90, giving a dedicated software tool for performing the delamination prediction. Comparison of analysis results with previous analysis and experiment showed good agreement, yielding an initial verification for the analytical procedure.

Ratcliffe, James G.↗
NASA NTRS2020

Progressive Failure Analysis Correlation with Notched Composite Laminate Test Data

Testing of small and intermediate laminate panels with notches was performed and Progressive Failure Analyses (PFA) models were developed for aid in test planning and for correlation with the test data. Two progressive damage failure models were included in the study: the commercially available Abaqus built-in damage model and COmplete STress Reduction (COSTR) damage model developed at NASA Langley Research Center (LaRC). The finite elements models used for the analysis were developed using shell elements. The pre-test PFA results obtained from the two damage models were compared to test data. Then a post-test PFA with updated material properties based on additional available material property data was executed with the COSTR damage model and compared to the test data. The panels tested in compression all exhibited less scatter in the failure load with self-similar failure behavior. The panels tested in tension exhibited more scatter in the failure load with a failure mode that included delamination of plies. However, a self-similar crack path was also observed as the overall failure mode for all panels tested. When comparing test data to the pre-test analysis results, both damage models with shell element models were considered adequate for predicting the behavior and failure load of both the small and intermediate panels when tested in compression. However, for the tension load cases, larger differences were observed between the test and analysis results. When comparing test data to the post-test analysis performed using the COSTR damage model, better correlation was observed. However, large discrepancies were still observed for the tension load cases. Consequently, a higher fidelity finite element model including solid elements for sub-laminates and the modeling of cohesive layers between sub-laminate layers is recommended for tension loading of notched composite laminates.

Damage Model,Progressive Failure Analysis,Open Hol↗
NASA NTRS1993

Analysis of delamination in cross-ply laminates initiating from impact induced matrix cracking

Two-dimensional finite element analyses of (02/90(8)/02) glass/epoxy and graphite/epoxy composite laminates were performed to investigate some of the characteristics of damage development due to an impact load. A cross section through the thickness of the laminate with fixed ends, and carrying a transverse load in the center, was analyzed. Inclined matrix cracks, such as those produced by a low-velocity impact, were modeled in the 90 deg ply group. The introduction of the matrix cracks caused large interlaminar tensile and shear stresses in the vicinity of both crack tips in the 0/90 and 90/0 interfaces, indicating that matrix cracking may give rise to delamination. The ratio of Mode I to total strain energy release rate, G(I)/G(total), at the beginning of delamination, calculated at the two (top and bottom) matrix crack tips was 60 and 28 percent, respectively, in the glass/epoxy laminate. The corresponding ratio was 97 and 77 percent in the graphite/epoxy laminate. Thus, a significant Mode I component of strain energy release rate may be present at the delamination initiation due to an impact load. The value of strain energy release rate at either crack tip increased due to an increase in the delamination length at the other crack tip and may give rise to an unstable delamination growth under constant load.

Salpekar, S. A.↗
NASA NTRS2016

An Enriched Shell Finite Element for Progressive Damage Simulation in Composite Laminates

A formulation is presented for an enriched shell nite element capable of progressive damage simulation in composite laminates. The element uses a discrete adaptive splitting approach for damage representation that allows for a straightforward model creation procedure based on an initially low delity mesh. The enriched element is veri ed for Mode I, Mode II, and mixed Mode I/II delamination simulation using numerical benchmark data. Experimental validation is performed using test data from a delamination-migration experiment. Good correlation was found between the enriched shell element model results and the numerical and experimental data sets. The work presented in this paper is meant to serve as a rst milestone in the enriched element's development with an ultimate goal of simulating three-dimensional progressive damage processes in multidirectional laminates.

McElroy, Mark W.↗
NASA NTRS2004

Discrete Spring Model for Predicting Delamination Growth in Z-Fiber Reinforced DCB Specimens

Beam theory analysis was applied to predict delamination growth in Double Cantilever Beam (DCB) specimens reinforced in the thickness direction with pultruded pins, known as Z-fibers. The specimen arms were modeled as cantilever beams supported by discrete springs, which were included to represent the pins. A bi-linear, irreversible damage law was used to represent Z-fiber damage, the parameters of which were obtained from previous experiments. Closed-form solutions were developed for specimen compliance and displacements corresponding to Z-fiber row locations. A solution strategy was formulated to predict delamination growth, in which the parent laminate mode I critical strain energy release rate was used as the criterion for delamination growth. The solution procedure was coded into FORTRAN 90, giving a dedicated software tool for performing the delamination prediction. Comparison of analysis results with previous analysis and experiment showed good agreement, yielding an initial verification for the analytical procedure.

Ratcliffe, James G.↗
NASA NTRS2006

Nonlinear Spring Finite Elements for Predicting Mode I-Dominated Delamination Growth in Laminated Structure with Through-Thickness reinforcement

One particular concern of polymer matrix composite laminates is the relatively low resistance to delamination cracking, in particular when the dominant type of failure is mode I opening. One method proposed for alleviating this problem involves the insertion pultruded carbon pins through the laminate thickness. The pins, known as z-pins, are inserted into the prepreg laminate using an ultrasonic hammer prior to the curing process, resulting in a field of pins embedded normal to the laminate plane as illustrated in Figure. 1. Pin diameters range between 0.28-mm to 0.5-mm and standard areal densities range from 0.5% to 4%. The z-pins are provided by the manufacturer, Aztex(Registered TradeMark) , in a low-density foam preform, which acts to stabilize orientation of the pins during the insertion process [1-3]. Typical pin materials include boron and carbon fibers embedded in a polymer matrix. A number of methods have been developed for predicting delamination growth in laminates reinforced with z-pins. During a study on the effect of z-pin reinforcement on mode I delamination resistance, finite element analyses of z-pin reinforced double cantilever beam (DCB) specimens were performed by Cartie and Partridge [4]. The z-pin bridging stresses were modeled by applying equivalent forces at the pin locations. Single z-pin pull-out tests were performed to characterize the traction law of the pins under mode I loading conditions. Analytical solutions for delamination growth in z-pin reinforced DCB specimens were independently derived by Robinson and Das [5] and Ratcliffe and O'Brien [6]. In the former case, pin bridging stresses were modeled using a distributed load and in the latter example the bridging stresses were discretely modeled by way of grounded springs. Additionally, Robinson and Das developed a data reduction strategy for calculating mode I fracture toughness, G(sub Ic), from a z-pin reinforced DCB specimen test [5]. In both cases a traction law similar to that adopted by Cartie and Partridge was used to represent z-pin failure under mode I loading conditions. In the current work spring elements available in most commercial finite element codes were used to model z-pins. The traction law used in previous analyses [4-6] was employed to represent z-pin damage. This method is intended for and is limited to simulating z-pins in composite laminate structure containing mode I-dominated delamination cracking. The current technique differs from previous analyses in that spring finite elements (available in commercial codes) are employed for simulating zpins, reducing the complexity of the analysis construction process. Furthermore, the analysis method can be applied to general structure that experiences mode I-dominated delamination cracking, in contrast to existing analytical solutions that are only applicable to coupon DCB specimens.

Ratcliffe, James G.↗
NASA NTRS2010

Thermography Inspection for Detection and Tracking of Composite Cylinder Damage During Load Testing

Two thermography techniques, passive and active, are used to detect damage initiation and progression in a cyclically loaded composite cylinder. The passive thermography tracks damage progression in real time during cyclic loading. Active flash thermography, using a flash tube enclosed within the cylinder, images delaminations in a cylinder under different loads. A differential thermography processing technique eliminates normal material variations and improves sensitivity to and sizing of delaminations. The thermography results were compared to nonimmersion ultrasonic results.

Zalameda, J. N.↗
NASA NTRS2013

LSP Composite Susbtrate Manufacturing Processing Guide

This document is intended to define Carbon Fiber Reinforced Plastic (CFRP) test panel configurations that can be employed for the purposes of evaluating the protection capabilities of Lightning Strike Protection (LSP) materials developed by the Aerospace Industry. The configurations are intended to provide consistent behavior in their response to simulated lightning strikes at pre-defined levels when tested by a capable vendor according to a test procedure written to enable consistent results (ref section 2.1.2). In response to an attachment of a simulated lightning strike on a CFRP panel, one can expect to see various levels of ablation and delamination, both through the thickness of the panel and with respect to the amount of panel surface area that exhibits damage. Panel configurations defined in this document include: An "unprotected" configuration 128694-1 (ref section 4.1), consisting of a cured CFRP laminate stackup of tape and fabric prepregs, coated with a typical aerospace primer and paint finishing scheme, attached to aluminum grounding bars intended to draw electrical current from the lightning attachment point to the panel edges and thus to ground. A "protected" configuration 128694-2 (ref section 4.1), wherein a layer of an LSP material form often used in the Aerospace Industry is included in the laminate stackup prior to cure. The CFRP materials, finishes and grounding arrangement for ths configuration are the same as for the "unprotected" configuration.

Kovach, Daniel J.↗
NASA NTRS2005

Nondestructive Evaluation of Thermal Barrier Coatings by Mid-infrared Reflectance

The application of mid-infrared reflectance (MIR) imaging to monitor damage in thermal barrier coatings (TBCs) has been extended from a previously demonstrated area-averaged spectroscopic analysis tool to become a practical imaging tool that provides the spatial resolution needed to quickly identify localized regions of TBC damage by visual inspection, Illumination optics and image collection procedures were developed to produce illumination-normalized flatfield reflectance images after subtraction of the background thermal emission. MIR reflectance images were collected with a bandpass filter centered at a wavelength of 4 microns, which provided the optimum balance between good sensitivity to buried cracks and coating erosion, but with a desirable sensitivity to TBC sintering and absorption from ambient gases. Examples are presented of the application of MIR reflectance imaging to monitor damage progression in plasma-sprayed 8wt% yttria-stabilized zirconia (8YSZ) TBCs subjected to either furnace cycling or alumina particle jet erosion. These results show that MIR reflectance imaging can reliably track the progression of buried delamination cracks produced by thermal cycling and can also be used to determine when any local section of the TBC has eroded beyond an acceptable limit. Modeling of the effects of buried cracks and erosion on reflectance will be presented to show the dependence of damage sensitivity to TBC thickness.

Eldridge, Jeffrey I↗
NASA NTRS1982

Effects of moisture, residual thermal curing stresses, and mechanical load on the damage development in quasi-isotropic laminates

This investigation demonstrates how moisture absorbed in (0/+ or - 45/90)s and (0/90/+ or - 45)s graphite epoxy laminates significantly alters the stress state and chronology of damage development along the laminate edge during static tension and tension-tension cyclic loading. Emphasis is placed on using reasonable approximations for wet and dry elastic properties, including out-of-plane properties (nu sub 23 and G sub 23), since these properties are required by finite element and shear lag models to predict the stress state at the laminate edge. Moisture was observed to alter the dry edge stress state in the 90-deg plies of the (0/+ or - 45/90)s laminate such that delaminations occurred at a lower load and transverse cracks occurred at a higher load. A model was developed which predicted the differences in loads required to initiate damage in the 90-deg plies of the two laminates in the wet and dry conditions. Although moisture can alter the chronology of damage development, the damage state in each laminate observed prior to fracture appeared to be independent of moisture content.

Kriz, R. D.↗
NASA NTRS2009

Improved Fabrication of Ceramic Matrix Composite/Foam Core Integrated Structures

The use of hybridized carbon/silicon carbide (C/SiC) fabric to reinforce ceramic matrix composite face sheets and the integration of such face sheets with a foam core creates a sandwich structure capable of withstanding high-heatflux environments (150 W/cm2) in which the core provides a temperature drop of 1,000 C between the surface and the back face without cracking or delamination of the structure. The composite face sheet exhibits a bilinear response, which results from the SiC matrix not being cracked on fabrication. In addition, the structure exhibits damage tolerance under impact with projectiles, showing no penetration to the back face sheet. These attributes make the composite ideal for leading edge structures and control surfaces in aerospace vehicles, as well as for acreage thermal protection systems and in high-temperature, lightweight stiffened structures. By tailoring the coefficient of thermal expansion (CTE) of a carbon fiber containing ceramic matrix composite (CMC) face sheet to match that of a ceramic foam core, the face sheet and the core can be integrally fabricated without any delamination. Carbon and SiC are woven together in the reinforcing fabric. Integral densification of the CMC and the foam core is accomplished with chemical vapor deposition, eliminating the need for bond-line adhesive. This means there is no need to separately fabricate the core and the face sheet, or to bond the two elements together, risking edge delamination during use. Fibers of two or more types are woven together on a loom. The carbon and ceramic fibers are pulled into the same pick location during the weaving process. Tow spacing may be varied to accommodate the increased volume of the combined fiber tows while maintaining a target fiber volume fraction in the composite. Foam pore size, strut thickness, and ratio of face sheet to core thickness can be used to tailor thermal and mechanical properties. The anticipated CTE for the hybridized composite is managed by the choice of constituents, varying fiber tow sizes and constituent part ratios. This structural concept provides high strength and stiffness at low density 1.06 g/cm3 in panels tested. Varieties of face sheet constructions are possible, including variations in fiber type and weave geometry. The integrated structures possible with this composite could eliminate the need for non-load-bearing thermal protection systems on top of a structural component. The back sheet can readily be integrated to substructures through the incorporation of ribs. This would eliminate weight and cost for aerospace missions.

Hurwitz, Frances I.↗
NASA NTRS2019

Potential Weight Benefits of IM7/8552 Hybrid Thin-Ply Composites for Aircraft Structures

Composite materials have increasingly been used for aerospace applications due to improved performance and reduced weight compared to their metallic counterparts. Inclusion of thin-ply material, plies with cured thickness half or less than standard-ply composites, have potential to improve performance and reduce structural weight further. The effect of thin-ply material on the weight of aircraft structure was investigated by examining wing cover weight reduction. To minimize the effects on manufacturing due to using thin plies, hybrid laminates were examined that used thin ±45-degree plies to replace their standard-ply counterparts in laminates. Compression after impact (CAI) tests were conducted to examine the possible weight savings that could be gained by increasing the design allowables that were used to size the wing upper cover of a semi-span test article. A large increase in CAI strength was observed for quasi-isotropic hybrid laminates, whereas less improvement was seen for hard hybrid laminates such as found in the wing cover. For laminates design by CAI strength, weight savings of about 13% were found using the hybrid hard laminates compared to the standard-ply laminates. Whether similar weight savings could be expected for structure sized using tension after impact allowables will have to be investigated further. Notched specimens were tested to examine possible weight savings using hybrid laminates in regions that are sized using discrete source damage requirements. As expected, the hybrid laminate had marginal improvements over the standard-ply laminate for compression with a notch present. The hybrid laminate, however, exhibited about 20% lower strength than the standard-ply laminate counterpart for tension with a notch. The failure mode of the hybrid specimens was a brittle, self-similar crack, which differs from the standard-ply specimens that failed by significant amounts of delamination and fiber splitting. In light of the apparent reduction in notched tensile strength, additional investigation is required to assess the use of hybrid laminates for areas containing discrete source damage, and their effect on weight of such regions.

Lovejoy, Andrew E.↗
NASA NTRS2009

Inert Welding/Brazing Gas Filters and Dryers

The use of hybridized carbon/silicon carbide (C/SiC) fabric to reinforce ceramic matrix composite face sheets and the integration of such face sheets with a foam core creates a sandwich structure capable of withstanding high-heat-flux environments (150 W/sq cm) in which the core provides a temperature drop of 1,000 C between the surface and the back face without cracking or delamination of the structure. The composite face sheet exhibits a bilinear response, which results from the SiC matrix not being cracked on fabrication. In addition, the structure exhibits damage tolerance under impact with projectiles, showing no penetration to the back face sheet. These attributes make the composite ideal for leading-edge structures and control surfaces in aerospace vehicles, as well as for acreage thermal protection systems and in high-temperature, lightweight stiffened structures. By tailoring the coefficient of thermal expansion (CTE) of a carbon fiber containing ceramic matrix composite (CMC) face sheet to match that of a ceramic foam core, the face sheet and the core can be integrally fabricated without any delamination. Carbon and SiC are woven together in the reinforcing fabric. Integral densification of the CMC and the foam core is accomplished with chemical vapor deposition, eliminating the need for bond-line adhesive. This means there is no need to separately fabricate the core and the face sheet, or to bond the two elements together, risking edge delamination during use. Fibers of two or more types are woven together on a loom. The carbon and ceramic fibers are pulled into the same "pick" location during the weaving process. Tow spacing may be varied to accommodate the increased volume of the combined fiber tows while maintaining a target fiber volume fraction in the composite. Foam pore size, strut thickness, and ratio of face sheet to core thickness can be used to tailor thermal and mechanical properties. The anticipated CTE for the hybridized composite is managed by the choice of constituents, varying fiber tow sizes and constituent part ratios. This structural concept provides high strength and stiffness at low density 1.06 g/cu cm in panels tested. Varieties of face sheet constructions are possible, including variations in fiber type and weave geometry. The integrated structures possible with this composite could eliminate the need for non-load-bearing thermal protection systems on top of a structural component. The back sheet can readily be integrated to substructures through the incorporation of ribs. This would eliminate weight and cost for aerospace missions.

Goudy, Jerry↗
NASA NTRS1984

Compression failure characteristics of + or - 45-deg-dominated laminates with a circular hole or impact damage

An investigation of the compression failure characteristics of + or - 45-deg-dominated laminates with a circular hole or impact damage was conducted. Graphite-epoxy laminates consisting of all + or - 45-deg plies and of + or - 45-deg and 90-deg plies were evaluated. Inplane shearing between the fibers and matrix (i.e., matrix shearing) was found to be the primary compression failure mechanism for an all + or - 45-deg laminate with a hole. The failure mechanism for the + or 45-deg-dominated laminates with a hole appeared to be a combination of delamination and matrix shearing. The compression failure stresses and strains for both the all + or - 45-deg laminate and the + or - 45-deg-dominated laminates were unaffected by impact damage at an energy level of 5.5 J when compared to results for undamaged control specimens. The failure stresses and strains for all laminates with impact damage at an energy level of 17.0 J or of 34.3 J were significantly lower than the corresponding data for the undamaged control specimens. The failure stresses and strains for these impact-damaged specimens were independent of the percentage of 90-deg plies in the laminate.

Shuart, M. J.↗
NASA NTRSDate not supplied

Least Squares Reverse Time Migration (LSRTM) for Damage Imaging in Composite Laminates

A method for adapting least squares reverse time migration (LSRTM) for ultrasonic guided wave imaging of composite laminates is proposed in this paper. As composites become more widely used in fields such as the aerospace industry, the need for high-resolution imaging in structural health monitoring (SHM) and nondestructive evaluation (NDE) is also growing. For instance, delamination is a common problem in composite laminates, which has led to a certain degree of apprehension in the use of composite materials for load-bearing structures. Although the solver-based imaging techniques using conventional reverse time migration (RTM) methods illuminate damage with a wide range of damage-scattering effects, the resulting images do not fully define the damage regions due to the limited data acquisition aperture, sensor density, frequencies/wavelengths, and incompleteness of adjoint reconstruction. Previously, the LSRTM theory was derived by incorporating a least squares migration (LSM) algorithm into RTM, and benchmarked its high-resolution damage imaging performance for isotropic plates. To improve damage imaging in composite laminates, a process to create an ultrasonic guided wave-based LSRTM method for anisotropic materials is proposed in this paper. The derivation of the forward modeling operator and the adjoint operator is presented. Numerical case studies were conducted to show the improvement of LSRTM over RTM in mapping damage in composite plates. Multiple damage sites or damage with a complex shape were created in the numerical studies based on 1) Born approximation-based modeling or 2) reduction of the plate stiffness-based modeling. LSRTM begins by using a forward modeling process. The forward modeling accuracy for the flexural wave solver used in this study for LSRTM was also compared with the results from a three-dimensional rotated staggered grid (RSG) elastodynamic finite difference solver, previously open-sourced by NASA Langley Research Center. These benchmarking studies have shown that the LSRTM method is more effective in reducing artifacts, improving resolution, and enhancing damage model reconstruction compared to RTM in composite laminates.

Jiaze He↗
NASA NTRS1994

Assessment of impact damage of composite rocket motor cases

This contract reviewed the available literature on mechanisms of low velocity impact damage in filament wound rocket motor cases, MDE methods to quantify damage, critical coupon level test methods, manufacturing and material process variables and empirical and analytical modeling off impact damage. The critical design properties for rocket motor cases are biaxial hoop and axial tensile strength. Low velocity impact damage is insidious because it can create serious nonvisible damage at very low impact velocities. In thick rocket motor cases the prevalent low velocity impact damage is fiber fracture and matrix cracking adjacent to the front face. In contrast, low velocity loading of thin wall cylinders induces flexure, depending on span length and the flexure induces delamination and tensile cracking on the back face wall opposed to impact occurs due to flexural stresses imposed by impact loading. Important NDE methods for rocket motor cases are non-contacting methods that allow inspection from one side. Among these are vibrothermography, and pulse-echo methods based on acoustic-ultrasonic methods. High resolution techniques such as x-ray computed tomography appear to have merit for accurate geometrical characterization of local damage to support development of analytical models of micromechanics. The challenge of coupon level testing is to reproduce the biaxial stress state that the full scale article experiences, and to determine how to scale the composite structure to model full sized behavior. Biaxial tensile testing has been performed by uniaxially tensile loading internally pressurized cylinders. This is experimentally difficult due to gripping problems and pressure containment. Much prior work focused on uniaxial tensile testing of model filament wound cylinders. Interpretation of the results of some studies is complicated by the fact that the fabrication process did not duplicate full scale manufacturing. It is difficult to scale results from testing subscale cylinders since there are significant differences in out time of the resins relative to full scale cylinder fabrication, differences in hoop fiber tensioning and unsatisfactory coupon configurations. It appears that development of a new test method for subscale cylinders is merited. Damage tolerance may be improved by material optimization that uses fiber treatments and matrix modifications to control the fiber matrix interface bonding. It is difficult to develop process optimization in subscale cylinders without also modeling the longer out times resins experience in full scale testing. A major breakthrough in characterizing the effect of impact damage on residual strength, and understanding how to scale results of subscale evaluations, will be a sound micromechanical model that described progressive failure of the composite. Such models will utilize a three dimensional stress analysis due to the complex nature of low velocity impact stresses in thick composites. When these models are coupled with non-contact NDE methods that geometrically characterize the damage and acoustic methods that characterize the effective local elastic properties, accurate assessment of residual strength from impact damage may be possible. Directions for further development are suggested.

Paris, Henry G.↗