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NASA NTRS1991

Characterization and modeling of tensile behavior of ceramic woven fabric composites

This paper examines the tensile behavior of SiC/SiC fabric composites. In the characterization effort, the stress-strain relation and damage evolution are studied with a series of loading and unloading tensile test experiments. The stress-strain relation is linear in response to the initial loading and becomes nonlinear when loading exceeds the proportional limit. Transverse cracking has been observed to be a dominant damage mode governing the nonlinear deformation. The damage is initiated at the inter-tow pores where fiber yarns cross over each other. In the modeling work, the analysis is based upon a fiber bundle model, in which fiber undulation in the warp and fill directions and gaps among fiber yarns have been taken into account. Two limiting cases of fabric stacking arrangements are studied. Closed form solutions are obtained for the composite stiffness and Poisson's ratio. Transverse cracking in the composite is discussed by applying a constant failure strain criterion.

Kuo, Wen-Shyong↗

NASA NTRS2007

Progressive Fracture of Laminated Composite Stiffened Plate

Laminated fiber-reinforced composite stiffened plate with [0/90/plus or minus 45]s plies made of S-Glass/epoxy are evaluated via computational simulation to study damage and fracture progression. The loads are pressure and temperature which varies from 21 to 65.5 C (case I) and from 143.3 to 21 C (case II). An integrated computer code is used for the simulation of the damage progression. Results show that damage initiation begins at low load level, with matrix cracking at the 0 deg. (bottom and top) plies, fiber fracture at the bottom (0 deg.) ply and interply delamination at the top (0 deg. ) ply. Increasing the applied pressure, the damage growth is expended resulting in fracture through the thickness of the structure. At this stage, 90 percent of the plies damage at applied pressure 15.306 MPa for the case I and 15.036 MPa for the case II. After this stage the cracks propagate rapidly and the structure collapses.

Gotsis, Pascalis K.↗

NASA NTRS1993

The viscoelastic behavior of notched glassy polymers

In the bulk, glassy polymers exhibit a nonlinear viscoelastic response during deformation. Stress or strain induced damage (i.e. crazing, microshear banding) results in the production of nonrecoverable work and observed nonlinearity. Stress or strain dependent shift factors have been used to mathematically model the mechanical behavior of these polymers. Glassy polymers that have been notched, may exhibit very different load displacement response compared to the same material under bulk deformation. If a sharp notch is introduced into the body then loaded, the load displacement trace may appear to be single-valued in the absence of viscoelasticity and crack growth. This suggests the volume of damaged material is small compared to the overall dimensions of the specimen. The ability to produce a single-valued load-load-line displacement trace through the use of the Correspondence Principle may prove to be useful for fracture of viscoelastic materials.

Crook, R. A.↗

NASA NTRS1998

A New Approach to Overcoming Spatial Aliasing in Structural Damage Detection

Aircraft, reusable launch vehicles,unmanned aircraft, and other advanced structures are being built using lightweight composite materials/metals with design safety factors as low as 1.25. These advanced structures operate in uncertain and severe environments and are susceptible to damage such as delamination, fiber/matrix damage, hydrothermal strain in composite materials, and fatigue and cracking in metals. To ensure human safety and load-bearing integrity these structures must be inspected to detect and locate often invisible damage and faults before they become catastrophic. Conventional methods of non-destructive evaluation sometimes miss significant damage and are time consuming and expensive to perform. In contrast, vibrometry or vibration signature techniques are a global method of structural integrity monitoring that potentially can efficiently detect damage on large structures, including damage that is away from sensor locations, and in the interior of structures. However, a barrier problem in damage detection using vibration measurements is the need to measure the vibration response at a large number of points on the structure. Typically, model reduction or expansion procedures such as Guyan reduction or dynamic expansion are attempted to overcome the problem of insufficient measurements or spatial aliasing. These approaches depend on using information from the healthy model, and thus put error into the reduction used to represent the damaged structure. In this paper, a Frequency Response Function technique is used to detect damage to a fixed-free beam. The technique uses measured frequency response functions from the healthy structure as reference data, and then monitors vibration measurements during the life of the structure to detect damage. In an analytical simulation using a finite-element model of a beam, damage was located using only sparse measurements because the technique uses both rotation and translation measurements from the damaged structure.

Schulz, Mark J.↗

NASA NTRS2017

Damage Simulation in Composite Materials: Why It Matters and What Is Happening Currently at NASA in This Area

Use of lightweight composite materials in space and aircraft structure designs is often challenging due to high costs associated with structural certification. Of primary concern in the use of composite structures is durability and damage tolerance. This concern is due to the inherent susceptibility of composite materials to both fabrication and service induced flaws. Due to a lack of general industry accepted analysis tools applicable to composites damage simulation, a certification procedure relies almost entirely on testing. It is this reliance on testing, especially compared to structures comprised of legacy metallic materials where damage simulation tools are available, that can drive costs for using composite materials in aerospace structures. The observation that use of composites can be expensive due to testing requirements is not new and as such, research on analysis tools for simulating damage in composite structures has been occurring for several decades. A convenient approach many researchers/model-developers in this area have taken is to select a specific problem relevant to aerospace structural certification and develop a model that is accurate within that scope. Some examples are open hole tension tests, compression after impact tests, low-velocity impact, damage tolerance of an embedded flaw, and fatigue crack growth to name a few. Based on the premise that running analyses is cheaper than running tests, one motivation that many researchers in this area have is that if generally applicable and reliable damage simulation tools were available the dependence on certification testing could be lessened thereby reducing overall design cost. It is generally accepted that simulation tools if applied in this manner would still need to be thoroughly validated and that composite testing will never be completely replaced by analysis. Research and development is currently occurring at NASA to create numerical damage simulation tools applicable to damage in composites. The Advanced Composites Project (ACP) at NASA Langley has supported the development of composites damage simulation tools in a consortium of aerospace companies with a goal of reducing the certification time of a commercial aircraft by 30%. And while the scope of ACP does not include spacecraft, much of the methodology and simulation capabilities can apply to spacecraft certification in the Space Launch System and Orion programs as well. Some specific applications of composite damage simulation models in a certification program are (1) evaluation of damage during service when maintenance may be difficult or impossible, (2) a tool for early design iterations, (3) gaining insight into a particular damage process and applying this insight towards a test coupon or structural design, and (4) analysis of damage scenarios that are difficult or impossible to recreate in a test. As analysis capabilities improve, these applications and more will become realized resulting in a reduction in cost for use of composites in aerospace vehicles. NASA is engaged in this process from both research and application perspectives. In addition to the background information discussed previously, this presentation covers a look at recent research at NASA in this area and some current/potential applications in the Orion program.

McElroy, Mack↗

DOE OSTI2023

Multiscale numerical investigation of ratchet growth damage effects in PBX 9502

This paper presents results of numerical experiments conducted on the high explosive PBX 9502 to investigate how recently observed grain-scale damage mechanisms of ratchet growth affect uniaxial compression measurements. Simulations are multiscale in the sense of directly resolving grains, pores, cracks, and grain-interfaces based upon scanning electron microscope (SEM) images of damaged and undamaged samples. The combined finite-discrete element method (FDEM) is utilized to resolve both grain-scale microfracture and elastoplastic deformation of solid grains. Pristine (undamaged) and damaged microstructures are compared in simulation of unconfined compression tests of the same material from the literature. Here, the simulation results show the observed microscale mechanisms of damage, specifically microfracture predominantly around and sometimes through grains and crack-associated pore growth, can well-explain the effective degradation of strength and stiffness observed in the laboratory measurements.

36 MATERIALS SCIENCE↗

NASA NTRS1994

Residual strength of thin panels with cracks

The previous design philosophies involving safe life, fail-safe and damage tolerance concepts become inadequate for assuring the safety of aging aircraft structures. For example, the failure mechanism for the Aloha Airline accident involved the coalescence of undetected small cracks at the rivet holes causing a section of the fuselage to peel open during flight. Therefore, the fuselage structure should be designed to have sufficient residual strength under worst case crack configurations and in-flight load conditions. Residual strength is interpreted as the maximum load carrying capacity prior to unstable crack growth. Internal pressure and bending moment constitute the two major components of the external loads on the fuselage section during flight. Although the stiffeners in the form of stringers, frames and tear straps sustain part of the external loads, the significant portion of the load is taken up by the skin. In the presence of a large crack in the skin, the crack lips bulge out with considerable yielding; thus, the geometric and material nonlinearities must be included in the analysis for predicting residual strength. Also, these nonlinearities do not permit the decoupling of in-plane and out-of-plane bending deformations. The failure criterion combining the concepts of absorbed specific energy and strain energy density addresses the aforementioned concerns. The critical absorbed specific energy (local toughness) for the material is determined from the global specimen response and deformation geometry based on the uniaxial tensile test data and detailed finite element modeling of the specimen response. The use of the local toughness and stress-strain response at the continuum level eliminates the size effect. With this critical parameter and stress-strain response, the finite element analysis of the component by using STAGS along with the application of this failure criterion provides the stable crack growth calculations for residual strength predictions.

Madenci, Erdogan↗

NASA NTRS2004

Lamb Wave Assessment of Fatigue and Thermal Damage in Composites

Among the various techniques available, ultrasonic Lamb waves offer a convenient method of evaluating composite materials. Since the Lamb wave velocity depends on the elastic properties of a structure, an effective tool exists to monitor damage in composites by measuring the velocity of these waves. Lamb wave measurements can propagate over long distances and are sensitive to the desired in-plane elastic properties of the material. This paper describes two studies which monitor fatigue damage and two studies which monitor thermal damage in composites using Lamb waves. In the fatigue studies, the Lamb wave velocity is compared to modulus measurements obtained using strain gage measurements in the first experiment and the velocity is monitored along with the crack density in the second. In the thermal damage studies, one examines samples which were exposed to varying temperatures for a three minute duration and the second includes rapid thermal damage in composites by intense laser beams. In all studies, the Lamb wave velocity is demonstrated to be an excellent method to monitor damage in composites.

Seale, Michael D.↗

NASA NTRS2018

Simulating the Clamped Tapered Beam Specimen Under Quasi-Static and Fatigue Loading Using Floating Node Method

As part of the NASA Advanced Composites Project (ACP), a sub-element has been designed to provide validation data for progressive damage analysis models. The clamped tapered beam is a cross-ply laminated composite specimen designed to validate the simulation of the onset of matrix cracks and their interaction with delaminations, including delamination migration. A tapered geometry was used to localize the first damage occurrence in the tapered region, without prescribing an initial crack. The boundary and loading conditions were chosen to favor delamination growth and subsequent migration after the first damage occurrence. The typical sequence of events consists of a matrix crack located at the tapered region, leading to delamination onset, followed by delamination growth and subsequent delamination migration to a different interface via a dominant matrix crack. The Clamped Tapered Beam (CTB) was tested in both quasi-static and fatigue regimes. The results obtained are used in this study to assess and validate a methodology based on the Floating Node Method (FNM) implemented as an Extended Interface Element. In this methodology, quasi-static and fatigue damage formation and development are modeled by combining FNM to represent crack networks, with Directional Cohesive Zone Elements (DCZE) and Virtual Crack Closure Technique (VCCT), respectively. Qualitatively, the methodology is capable of predicting the sequence of events and overall failure morphology. Quantitatively, the simulation results generally bound the experimental data, based on the range of the characterization data used. In this paper, the results from quasi static and fatigue simulations are compared and correlated with experimental data.

Seshadri, B. R.↗

NASA NTRS2016

Model-Based Fatigue Prognosis of Fiber-Reinforced Laminates Exhibiting Concurrent Damage Mechanisms

Prognostics of large composite structures is a topic of increasing interest in the field of structural health monitoring for aerospace, civil, and mechanical systems. Along with recent advancements in real-time structural health data acquisition and processing for damage detection and characterization, model-based stochastic methods for life prediction are showing promising results in the literature. Among various model-based approaches, particle-filtering algorithms are particularly capable in coping with uncertainties associated with the process. These include uncertainties about information on the damage extent and the inherent uncertainties of the damage propagation process. Some efforts have shown successful applications of particle filtering-based frameworks for predicting the matrix crack evolution and structural stiffness degradation caused by repetitive fatigue loads. Effects of other damage modes such as delamination, however, are not incorporated in these works. It is well established that delamination and matrix cracks not only co-exist in most laminate structures during the fatigue degradation process but also affect each other's progression. Furthermore, delamination significantly alters the stress-state in the laminates and accelerates the material degradation leading to catastrophic failure. Therefore, the work presented herein proposes a particle filtering-based framework for predicting a structure's remaining useful life with consideration of multiple co-existing damage-mechanisms. The framework uses an energy-based model from the composite modeling literature. The multiple damage-mode model has been shown to suitably estimate the energy release rate of cross-ply laminates as affected by matrix cracks and delamination modes. The model is also able to estimate the reduction in stiffness of the damaged laminate. This information is then used in the algorithms for life prediction capabilities. First, a brief summary of the energy-based damage model is provided. Then, the paper describes how the model is embedded within the prognostic framework and how the prognostics performance is assessed using observations from run-to-failure experiments

Prognostics↗

NASA NTRS2020

Ultrasonic Lamb Wave Inspection of Composite Defects

Composites have been extensively used in aerospace engineering due to their advantages of light weight, high strength, and engineering design flexibility. Manufacturing defects such as wrinkle and porosity can affect the performance of the composites and may lead to failure in the end, while damage such as delamination, fiber fraction, and matrix cracking can directly cause failure of the composites. In this paper, a Lamb wave based nonintrusive nondestructive evaluation system, which employs piezoelectric transducer for actuation and scanning laser Doppler vibrometer for wavefield sensing, is presented for typical composite defect and damage inspection and evaluation. Two composite panels with different geometry (flat or curved) and with various embedded defects (wrinkle and delamination) are inspected using the nonintrusive Lamb wave system. Both the wrinkles and delamination are detected from the wavefield and approximately quantified through wavefield imaging methods.

Wrinkles↗

NASA NTRS2000

Crack Turning in Integrally Stiffened Aircraft Structures

Current emphasis in the aircraft industry toward reducing manufacturing cost has created a renewed interest in integrally stiffened structures. Crack turning has been identified as an approach to improve the damage tolerance and fail-safety of this class of structures. A desired behavior is for skin cracks to turn before reaching a stiffener, instead of growing straight through. A crack in a pressurized fuselage encounters high T-stress as it nears the stiffener--a condition favorable to crack turning. Also, the tear resistance of aluminum alloys typically varies with crack orientation, a form of anisotropy that can influence the crack path. The present work addresses these issues with a study of crack turning in two-dimensions, including the effects of both T-stress and fracture anisotropy. Both effects are shown to have relation to the process zone size, an interaction that is central to this study. Following an introduction to the problem, the T-stress effect is studied for a slightly curved semi-infinite crack with a cohesive process zone, yielding a closed form expression for the future crack path in an infinite medium. For a given initial crack tip curvature and tensile T-stress, the crack path instability is found to increase with process zone size. Fracture orthotropy is treated using a simple function to interpolate between the two principal fracture resistance values in two-dimensions. An extension to three-dimensions interpolates between the six principal values of fracture resistance. Also discussed is the transition between mode I and mode II fracture in metals. For isotropic materials, there is evidence that the crack seeks out a direction of either local symmetry (pure mode I) or local asymmetry (pure mode II) growth. For orthotropic materials the favored states are not pure modal, and have mode mixity that is a function of crack orientation.

Pettit, Richard Glen↗

DOE OSTI2021

A micromechanical hydro-mechanical-damage coupled model for layered rocks considering multi-scale structures

An anisotropic Hydro-Mechanical-Damage coupled model was established in the framework of thermodynamics for saturated layered rocks, which contain arbitrarily-distributed cracks of smaller scales and orientated bedding planes of much larger scales. Here, the proposed model considers the anisotropic damage growth and frictional sliding of cracks and shear sliding and dilatancy of bedding planes, as well as the interaction between cracks and bedding planes. Numerical implementation of the nonlinear coupled model was developed based on TOUGHREACT, which is a well-established simulator for multiphase fluid flow and reactive transport analysis. The proposed model was first validated with laboratory water injection test results on rock blocks, and then used to study the effects of injection rates, in situ stresses and bedding planes on fluid injection processes at the field scale, demonstrating the importance to consider the variation of multi-scale structures for better understanding the coupled Hydro-Mechanical responses in layered rocks.

58 GEOSCIENCES↗

NASA NTRS1995

Characterization of Two-Dimensionally Braided Composites Subject to Static and Fatigue Loading

This paper presents a characterization of damage processes in four different two-dimensionally braided architectures consisting of graphite fibers in an epoxy matrix, subjected to both static and fatigue loading. Both notched and unnotched specimens were tested from each architecture. Damage mechanisms in static compression testing were found to range from matrix cracking in an architecture controlled by shearing characteristics of the matrix to axial fiber bundle kinking and splitting in architectures controlled by unidirectional fiber bundles. In both static tension and tension-tension fatigue testing, it was found that all architectures sustained the same basic damage types and progression sequences. The first types of damage to occur were splits in the braider bundles and cracks in the resin rich areas. Once a sufficient density of this type of damage had occurred, the axial bundles began to disbond from the surrounding constituents, and continued to do so until ultimate failure. One architecture sustained an additional mode of failure, which involved splitting of the axial bundles along their length. Varied degrees of notch sensitivity were found among the different architectures. In one architecture, this was attributed to differences found in the stacking of the individual plys of the braided architecture. As the density of axial bundles adjacent to the notch increased, the rate of stiffness loss increased and the life decreased.

Burr, Scott T.↗

NASA NTRS1982

Improved damage tolerance of titanium by adhesive lamination

Basic damage tolerance properties of Ti-6A1-4V titanium plate can be improved by laminating thin sheets of titanium with adhesives. Compact tension and center cracked tension specimens made from thick plate, thin sheet, and laminated plate (six plies of thin sheet) were tested. The fracture toughness of the laminated plate was 39 percent higher than the monolithic plate. The laminated plate's through the thickness crack growth rate was about 20 percent less than that of the monolithic plate. The damage tolerance life of the surface cracked laminate was 6 to over 15 times the life of a monolithic specimen. A simple method of predicting crack growth in a crack ply of a laminate is presented.

Johnson, W. S.↗

NASA NTRS1983

Damage tolerance evaluation of adhesively laminated titanium

Johnson, W. S.↗

NASA NTRS2007

Triboluminescent Materials for Smart Optical Damage Sensors for Space Applications

Triboluminescence is light that is produced by pressure, friction or mechanical shock. New composite materials are constantly being reengi neered in an effort to make lightweight spacecrafts for various NASA missions. For these materials there is interest in monitoring the con dition of the composite in real time to detect any delamination or cr acking due to damage, fatigue or external forces. Methods of periodic inspection of composite structures for mechanical damage such as ult rasonic testing are rather mature. However, there is a need to develop a new technique of damage detection for composites, which could dete ct cracking or delamination from any desired location within a materi al structure in real time. This could provide a valuable tool in the confident use of composite materials for various space applications. Recently, triboluminnescent materials have been proposed as smart sen sors of structural damage. To sense the damage, these materials can b e epoxy bonded or coated in a polymer matrix or embedded in a composi te host structure. When the damage or fracture takes place in the hos t structure, it will lead to the fracture of triboluminescent crystal s resulting in a light emission. This will warn, in real time, that a structural damage has occurred. The triboluminescent emission of the candidate phosphor has to be sufficiently bright, so that the light signal reaching from the point of fracture to the detector through a fiber optic cable is sufficiently strong to be detected. There are a large number of triboluminescent materials, but few satisfy the above criterion. Authors have synthesized a Eu based organic material know n as Europium tetrakis (dibenzoylmethide) triethylammonium .(EuD(sub 4)TEA), one of the bright triboluminescent materials, which is a pote ntial candidate for application as a damage sensor and could be made into a wireless sensor with the addition of microchip, antenna and el ectronics. Preliminary results on the synthesis and characterization of this material shall be presented.

Aggarwal, Mohan D.↗

NASA NTRS1981

Experimental and analytical study of fatigue damage in notched graphite/epoxy laminates

Fatigue damage development in notched graphite/epoxy laminates is investigated in both compression and tension, with tested specimens being examined for damage type and location using light microscopy, scanning electron microscopy, ultrasonic C-scans, and X-radiography. Delamination and ply cracking were found to be the dominant types of fatigue damage. Comparison of observed delamination locations with finite element calculations indicates that both interlaminar shear and peel stresses must be considered when predicting delamination. Fatigue loading generally caused only small stiffness losses, and residual strengths were in all cases greater than, or equal to, the virgin strengths.

Whitcomb, J. D.↗