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All sourcesNASA NTRSDOE OSTI
DOE OSTI2024

Optimized Viewing Techniques to Minimize Radiation Damage From X-ray Imaging Systems

X-ray inspection of ball grid arrays (BGAs) is typically performed at one or more viewing angles to examine adhesion sites for errors such as voids, joint cracking, or head-in-pillow. During this inspection process, the cir cuit board assembly is subject to ionizing radiation exposure, which can cause trapped charge within oxide layers of semiconductor devices. Some x-ray machines allow for programmable inspection routines, which could be used to optimize radiation exposure to semiconductor components. Using Monte Carlo methods, x-ray inspection of a BGA was simulated to determine a range of acceptable viewing angles. Dose rates to circuit board components were estimated at each inspection angle to determine the view resulting in optimized radiation exposure. Results showed that for each BGA, the maximum unobstructed viewing times without exceeding a 5 Gy dose limit to a single part ranged from 82 to 94 minutes. Using a radiation cost function method, optimized viewing across all components was found. Here, it was observed that for a consistent dose limit applied to silicon-based components, performing inspection with BGAs facing the x-ray source was optimal. A third method was applied, assigning individual dose limits based on empirical data from the NASA Goddard Space Flight Center radiation database. This method showed that optimized viewing maximizes the distance between the radiation source and highly sensitive components. It was also observed that cumulative effects from viewing two BGAs will influence viewing angles, causing the optimal view of one BGA to exist nearly 180° from the other.

46 INSTRUMENTATION RELATED TO NUCLEAR SCIENCE AND ↗
DOE OSTI2025

Experimental data for damage mechanics simulation challenge

While there are many computational approaches for simulating damage in rock and other materials, few have been ground truth tested with either known experimental data or with blind data sets. Here, in this work, we present a bench-mark laboratory data set for a damage mechanics challenge to compare computational approaches on damage evolution in brittle-ductile materials. The samples were fabricated through additive manufacturing to produce repeatable specimens designed to fail in controlled ways. The failure was induced in the samples using a 3-point bending test to produce different Modes such as Mode I and mixed Modes including I-II, I-III and I-II-III Modes to generate a calibration data set and a blind challenge data set. Data collected included spatial and temporal measurements from traditional digital load–displacement sensors, 2D digital image correlation measurement to map surface deformations, 3D X-ray microscopy to ground-truth the crack-failure geometry, and laser profilometry to capture surface roughness. The data sets are available, on a data repository, to the community to advance computational models to improve our ability to predict damage in brittle-ductile materials.

3-point bending↗
NASA NTRS2000

Testing and Analysis of Composite Skin/Stringer Debonding Under Multi-Axial Loading

A consistent step-wise approach is presented to investigate the damage mechanism in composite bonded skin/stringer constructions under uniaxial and biaxial (in-plane/out-of-plane) loading conditions. The approach uses experiments to detect the failure mechanism, computational stress analysis to determine the location of first matrix cracking and computational fracture mechanics to investigate the potential for delamination growth. In a first step, tests were performed on specimens, which consisted of a tapered composite flange, representing a stringer or frame, bonded onto a composite skin. Tests were performed under monotonic loading conditions in tension, three-point bending, and combined tension/bending to evaluate the debonding mechanisms between the skin and the bonded stringer. For combined tension/bending testing, a unique servohydraulic load frame was used that was capable of applying both in-plane tension and out-of-plane bending loads simultaneously. Specimen edges were examined on the microscope to document the damage occurrence and to identify typical damage patterns. For all three load cases, observed failure initiated in the flange, near the flange tip, causing the flange to almost fully debond from skin. In a second step, a two dimensional plane-strain finite element model was developed to analyze the different test cases using a geometrically nonlinear solution. For all three loading conditions, computed principal stresses exceeded the transverse strength of the material in those areas of the flange where the matrix cracks had developed during the tests. In a third step, delaminations of various lengths were simulated in two locations where delaminations were observed during the tests. The analyses showed that at the loads corresponding to matrix ply crack initiation computed strain energy release rates exceeded the values obtained from a mixed mode failure criterion in one location, Hence. Unstable delamination propagation is likely to occur as observed in the experiments.

Krueger, Ronald↗
NASA NTRS2000

Testing and Analysis of Composite Skin/Stringer Debonding under Multi-Axial Loading

A consistent step-wise approach is presented to investigate the damage mechanism in composite bonded skin/stringer constructions under uniaxial and biaxial (in-plane/out-of-plane) loading conditions. The approach uses experiments to detect the failure mechanism, computational stress analysis to determine the location of first matrix cracking and computational fracture mechanics to investigate the potential for delamination growth. In a first step, tests were performed on specimens, which consisted of a tapered composite flange, representing a stringer or frame, bonded onto a composite skin. Tests were performed under monotonic loading conditions in tension, three-point bending, and combined tension/bending to evaluate the debonding mechanisms between the skin and the bonded stringer. For combined tension/bending testing, a unique servohydraulic load frame was used that was capable of applying both in-plane tension and out-of-plane bending loads simultaneously. Specimen edges were examined on the microscope to document the damage occurrence and to identify typical damage patterns. For all three load cases, observed failure initiated in the flange, near the flange tip, causing the flange to almost fully debond from the skin. In a second step, a two-dimensional plane-strain finite element model was developed to analyze the different test cases using a geometrically nonlinear solution. For all three loading conditions, computed principal stresses exceeded the transverse strength of the material in those areas of the flange where the matrix cracks had developed during the tests. In a third step, delaminations of various lengths were simulated in two locations where delaminations were observed during the tests. The analyses showed that at the loads corresponding to matrix ply crack initiation computed strain energy release rates exceeded the values obtained from a mixed mode failure criterion in one location. Hence, unstable delamination propagation is likely to occur as observed in the experiments.

Krueger, Ronald↗
NASA NTRS2000

Formation of Leading-Edge Pinholes in the Space Shuttle Wings Investigated

The space shuttle wing leading edge and nose cap are composed of a carbon/carbon composite that is protected by silicon carbide. The coefficient of thermal expansion mismatch leads to cracks in the silicon carbide. The outer coating of the silicon carbide is a sodium-silicate-based glass that becomes fluid at the shuttles high reentry temperatures and fills these cracks. Small pinholes roughly 0.1 mm in diameter have been observed on these materials after 12 or more flights. These pinholes have been investigated by researchers at the NASA Johnson Space Center, Rockwell International, the Boeing Company, Lockheed Martin Corporation, and the NASA Glenn Research Center at Lewis Field to determine the possible sources and the extent of damage. A typical pinhole is illustrated in the photomicrographs. These pinholes are found primarily on the wing leading edges and not on the nose cap, which is covered when the orbiter is on the launch pad. The pinholes are generally associated with a bead of zincrich glass. Examination of the orbiter and launch structure indicates that weathering paint on the launch structure leads to deposits of zinc-containing paint flakes on the wing leading edge. These may become embedded in the crevices of the wing leading edge and form the observed zinc-rich glass. Laboratory experiments indicate that zinc oxide reacts vigorously with the glass coating on the silicon carbide. Thus, it is likely that this is the reaction that leads to pinhole formation (Christensen, S.V.: Reinforced Carbon/Carbon Pin Hole Formation Through Zinc Oxide Attack. Rockwell International Internal Letter, RDW 96 057, May 1996). Cross-sectional examination of pinholes suggests that they are enlarged thermal expansion mismatch cracks. This is illustrated in the photomicrographs. A careful microstructural analysis indicates that the pinhole walls consist of layers of zinc-containing glass. Thus, pinholes are likely formed by zinc oxide particles lodging in crevices and forming a corrosive zinc-rich glass that enlarges existing cracks. Having established the likely source of the pinholes, we next needed to model the damage. Our concern was that if a pinhole went through the silicon carbide to the carbon/carbon substrate, oxygen would have a clear path to oxidize the carbon at high temperatures. This possibility was examined with studies in a laboratory furnace. An ultrasonic drill was used to make artificial pinholes in a sample of protected carbon/carbon. After exposure, the specimens were weighed and cross-sectioned to quantify the extent of oxidation below the pinhole. The results at higher temperatures showed good agreement with a simple diffusion-control model. This model is based on the two-step oxidation of carbon to carbon monoxide and carbon dioxide. The fluxes are illustrated in the final figure. The model indicates a strong dependence on pinhole diameter. For smaller diameters and short times, the oxidation of carbon is very limited.

Jacobson, Nathan S.↗
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 NTRS1985

Tough composite materials: Recent developments

The present volume broadly considers topics in composite fracture toughness and impact behavior characterization, composite system constituent properties and their interrelationships, and matrix systems' synthesis and characterization. Attention is given to the characterization of interlaminar crack growth in composites by means of the double cantilever beam specimen, the characterization of delamination resistance in toughened resin composites, the effect of impact damage and open holes on the compressive strength of tough resin/high strain fiber laminates, the effect of matrix and fiber properties on compression failure mechanisms and impact resistance, the relation of toughened neat resin properties to advanced composite mechanical properties, and constituent and composite properties' relationships in thermosetting matrices. Also treated are the effect of cross-link density on the toughening mechanism of elastomer-modified epoxies, the chemistry of fiber/resin interfaces, novel carbon fibers and their properties, the development of a heterogeneous laminating resin, solvent-resistant thermoplastics, NASA Lewis research in advanced composites, and opportunities for the application of composites in commercial aircraft transport structures.

Vosteen, L. F.↗
NASA NTRS1987

Bithermal Low-Cycle Fatigue Behavior of a NiCoCrAlY-Coated Single Crystal Superalloy

Specimens of a single crystal superalloy, PWA 1480, both bare and coated with a NiCoCrAlY alloy, PWA 276, were tested in low-cycle fatigue at 650 and 1050 C, and in bithermal thermomechanical fatigue tests. In the two bithermal test types, tensile strain was imposed at one of the two temperatures and reversed in compression at the other. In the high-strain regime, lives for both bithermal test types approached that for the 650 C isothermal test on an inelastic strain basis, all being controlled by the low ductility of the superalloy at 650 C. In the low-strain regime, coating cracking reduced life in the 650 C isothermal test. The bithermal test imposing tension at 650 C, termed out-of-phase, also produced rapid surface cracking, but in both coated and bare specimens. Increased crack growth rates also occurred for the out-of-phase test. Increased lives in vacuum suggested that there is a large environmental contribution to damage in the out-of-phase test due to the 1050 C exposure followed by tensile straining at the low temperature.

Gayda, J.↗
NASA NTRS1987

Bithermal low-cycle fatigue behavior of a NiCoCrAlY-coated single crystal superalloy

Specimens of a single crystal superalloy, PWA 1480, both bare and coated with a NiCoCrAlY alloy, PWA 276, were tested in low-cycle fatigue at 650 and 1050 C, and in bithermal thermomechanical fatigue tests. In the two bithermal test types, tensile strain was imposed at one of the two temperatures and reversed in compression at the other. In the high-strain regime, lives for both bithermal test types approached that for the 650 C isothermal test on an inelastic strain basis, all being controlled by the low ductility of the superalloy at 650 C. In the low-strain regime, coating cracking reduced life in the 650 C isothermal test. The bithermal test imposing tension at 650 C, termed out-of-phase, also produced rapid surface cracking, but in both coated and bare specimens. Increased crack growth rates also occurred for the out-of-phase test. Increased lives in vacuum suggested that there is a large environmental contribution to damage in the out-of-phase test due to the 1050 C exposure followed by tensile straining at the low temperature.

Gayda, J.↗
NASA NTRS2006

Thermal Cycle Reliability of PBGA/CCGA 717 I/Os

Status of thermal cycle test results for a nonfunctional daisy-chained peripheral ceramic column grid array (CCGA) and its plastic ball grid array (PBGA) version. both having 560 I/Os. were presented in last year's conference. Test results included environmental data for three different thermal cycle regimes (-55 C/125 C, -55 C/100 C, and -50 C /75 C). Update information on these - especially failure type for assemblies with high and low solder volumes-were presented. The thermal cycle test procedure followed those recommended IPC-9701 for tin-lead solder joint assemblies. Revision A of this specification covers guideline thermal cycle requirements for Pb-free solder joints. Some background information discussed during release of this specification with its current guideline recommendations were also presented. In a recent reliability investigation a fully populated CCGA with 717 I/Os was also considered for assembly reliability. evaluation. The functional package is a field-programmable gate array that has much higher processing power than its previous version. This new package is smaller in dimension, has no interposer, and has a thinner column wrapped with copper for reliability improvement. This paper will also present thermal cycle test results for this package assembly and its plastic version with 728 I/Os. both of which were exposed to three different cycle regimes. Two cycle profiles were those specified by IPC- 9701A for tin-lead, i.e. -55 to 100 C and -55 to 125 C and one was a cycle profile specified by Mil-Std-883, i.e.. -65 C/150 C which is generally used for ceramic hybrid packages. Per IPC-9701 A, test vehicles were built using daisy chain packages and were continuously monitored. The effects of many process and assembly variables-including corner staking commonly used for improving resistance to mechanical loading such as drop and vibration loads--were also considered as part of the test matrix. Optical photomicrographs were taken at various thermal cycle intervals to document damage progress and behavior. Representative samples of these along with cross-sectional photomicrographs at higher magnification taken by scanning electron microscopy (SEM) to determine crack propagation and failure analyses for packages are also presented.

staking↗
NASA NTRSDate not supplied

Experimental Observations of Damage States in Unnotched and Notched 3D Orthogonal Woven Coupons Loaded in Tension

This paper describes a series of unnotched and notched tensile tests on two 3D orthogonal woven architectures. Dog bone specimens were used for the unnotched tests, and open-hole specimens were used for the notched tests. For each architecture and test type, two specimens were loaded under monotonic displacement control to failure and two specimens were incrementally loaded to failure in five steps. X-ray computed tomography (CT) was used to characterize the woven architecture and defects on each specimen prior to testing. Digital image correlation (DIC) images were obtained from both sides of the specimen throughout the loading and unloading. Acoustic emission (AE) data were also continuously acquired throughout the loading. For the incremental tests, X-ray CT was used to document the damage progression after each load increment. The damage progression was analyzed based on the results from these three techniques. The results for both architectures and test specimens followed a pattern where matrix cracks formed at the surface perpendicular to the tensile load direction. New surface cracks formed continuously throughout the loading and cracks extended in depth and width. In general, fiber failure was not detected until specimen failure.

Wade C Jackson↗
NASA NTRS1969

Investigation of the development of cracks in solder joints

Study consisted of an analytical approach, in which a mathematical model of existing printed circuit board component mounting techniques was analyzed, and an empirical investigation was performed to determine the extent of damage caused by temperature cycling of the printed circuit boards.

Moore, R. L.↗
NASA NTRS2012

In-Situ Nondestructive Evaluation of Kevlar(Registered Trademark)and Carbon Fiber Reinforced Composite Micromechanics for Improved Composite Overwrapped Pressure Vessel Health Monitoring

NASA has been faced with recertification and life extension issues for epoxy-impregnated Kevlar 49 (K/Ep) and carbon (C/Ep) composite overwrapped pressure vessels (COPVs) used in various systems on the Space Shuttle and International Space Station, respectively. Each COPV has varying criticality, damage and repair histories, time at pressure, and pressure cycles. COPVs are of particular concern due to the insidious and catastrophic burst-before-leak failure mode caused by stress rupture (SR) of the composite overwrap. SR life has been defined [1] as the minimum time during which the composite maintains structural integrity considering the combined effects of stress level(s), time at stress level(s), and associated environment. SR has none of the features of predictability associated with metal pressure vessels, such as crack geometry, growth rate and size, or other features that lend themselves to nondestructive evaluation (NDE). In essence, the variability or surprise factor associated with SR cannot be eliminated. C/Ep COPVs are also susceptible to impact damage that can lead to reduced burst pressure even when the amount of damage to the COPV is below the visual detection threshold [2], thus necessitating implementation of a mechanical damage control plan [1]. Last, COPVs can also fail prematurely due to material or design noncompliance. In each case (SR, impact or noncompliance), out-of-family behavior is expected leading to a higher probability of failure at a given stress, hence, greater uncertainty in performance. For these reasons, NASA has been actively engaged in research to develop NDE methods that can be used during post-manufacture qualification, in-service inspection, and in-situ structural health monitoring. Acoustic emission (AE) is one of the more promising NDE techniques for detecting and monitoring, in real-time, the strain energy release and corresponding stress-wave propagation produced by actively growing flaws and defects in composite materials [3,4,5,6,7,8]. To gain further insight into the mechanisms responsible for composite rupture, broadband modal acoustic emission analysis was used. Also, since AE data reduction proved to be very time consuming, specialized data reduction software was written to automate the process.

Waller, Jess↗
NASA NTRS2008

Fatigue Crack Growth Threshold Testing of Metallic Rotorcraft Materials

Results are presented for a program to determine the near-threshold fatigue crack growth behavior appropriate for metallic rotorcraft alloys. Four alloys, all commonly used in the manufacture of rotorcraft, were selected for study: Aluminum alloy 7050, 4340 steel, AZ91E Magnesium, and Titanium alloy Ti-6Al-4V (beta-STOA). The Federal Aviation Administration (FAA) sponsored this research to advance efforts to incorporate damage tolerance design and analysis as requirements for rotorcraft certification. Rotorcraft components are subjected to high cycle fatigue and are typically subjected to higher stresses and more stress cycles per flight hour than fixed-wing aircraft components. Fatigue lives of rotorcraft components are generally spent initiating small fatigue cracks that propagate slowly under near-threshold cracktip loading conditions. For these components, the fatigue life is very sensitive to the near-threshold characteristics of the material.

Newman, John A.↗
DOE OSTI2025

An FFT-based micromechanical model for gradient enhanced brittle fracture

Damage models incorporated within FFT-based micromechanical methods have received much attention recently because of the need to better understand and predict brittle and ductile fracture. An important aspect of a damage model is non-local regularization, which removes the mesh dependence of the predictions that otherwise become physically unacceptable upon grid refinement. In this work, the Helmholtz-type equation for non-local gradient regularization of a damage model on a distorted grid is solved using an FFT-based approach. Further, the resulting system of equations is solved using the Jacobi iterative method. The model is applied to simulate brittle fracture of an intermetallic. The influence of the time and space discretization, the length-scale parameter, and intermetallic crystallographic orientation on crack evolution is studied.

36 MATERIALS SCIENCE↗
NASA NTRS1988

Fatigue damage modeling for coated single crystal superalloys

A high temperature, low-cycle fatigue life prediction method for coated single crystal nickel-base superalloys is being developed. The method is being developed for use in predicting crack initiation life of coated single crystal turbine airfoils. Although the models are being developed using coated single crystal PWA 1480, they should be readily adaptable to other coated nickel-base single crystal materials. The coatings choosen for this effort were of two generic types: a low pressure plasma sprayed NiCoCrAlY overlay, designated PWA 286, and an aluminide diffusion, designated PWA 273. In order to predict the useful crack initiation life of airfoils, the constitutive and failure behavior of the coating/substrate combination must be taken into account. Coatings alter the airfoil surface microstructure and are a primary source from which cracks originate. The adopted life prediction approach addresses this complexity by separating the coating and single crystal crack initiation regimes. This provides a flexible means for using different life model formulations for the coating and single crystal materials. At the completion of this program, all constitutive and life model formulations will be available in equation form and as software. The software will use the MARC general purpose finite element code to drive the constitutive models and calculate life parameters.

Nissley, David M.↗
NASA NTRS1996

Durability and Damage Tolerance of High Temperature Polymeric Composites

Modern durability and damage tolerance predictions for composite material systems rely on accurate estimates of the local stress and material states for each of the constituents, as well as the manner in which the constituents interact. In this work, an number of approaches to estimating the stress states and interactions are developed. First, an elasticity solution is presented for the problem of a penny-shaped crack in an N-phase composite material system opened by a prescribed normal pressure. The stress state around such a crack is then used to estimate the stress concentrations due to adjacent fiber fractures in composite materials. The resulting stress concentrations are then used to estimate the tensile strength of the composite. The predicted results are compared with experimental values. In addition, a cumulative damage model for fatigue is presented. Modifications to the model are made to include the effects of variable amplitude loading. These modifications are based upon the use of remaining strength as a damage metric and the definition of an equivalent generalized time. The model is initially validated using results from the literature. Also, experimental data from APC-2 laminates and IM7/K3B laminates are used in the model. The use of such data for notched laminates requires the use of an effective hole size, which is calculated based upon strain distribution measurements. Measured remaining strengths after fatigue loading are compared with the predicted values for specimens fatigued at room temperature and 350 F (177 C).

Case, Scott W.↗
NASA NTRS1982

The effects of microcracking on the thermal expansion of graphite-epoxy composites

A study of the effects of thermal environment and microcracking in graphite epoxy composites was made. Research indicates that microcracking does affect the thermal expansion of composite laminates. The amount of reduction in thermal expansion was a function of the crack density. A maximum reduction of approximately 25% occurred in a quasi-isotropic specimen with a crack density of 2.05 mm 1 in the 90 deg plies. Laminate analysis with appropriate reductions in E sub 2 and Alpha sub 2 of the damaged plies appears to be capable of modeling the observed

Bowles, D. E.↗