Search

CANDU fuel bundles experience plastic deformations over time, and the horizontal configuration of the bundle in a crept pressure tube (PT) causes coolant to bypass the sagged lower half of the bundle. Bundle segments where the flow is limited may become more susceptible to dryout due to reactor aging. A finite element model of a 37-element fuel bundle was constructed using the commercial finite element software ANSYS to study the mechanical deformation behaviour of the bundle to maintain a coolable geometry. The main focus was on the contact between the fuel elements and between the fuel elements and PT. The complexity of the model due to all the contact pairs necessitated the use of high-powered computing hardware. Contact was demonstrated between the appendages, and sensitivity of the deformation to different boundary conditions (BC) was investigated. In particular, the radial position where the elements were welded to the endplate significantly impacted the magnitude of the element bowing. Expanding the PT up to 8% diametral creep demonstrated the proper functioning of the spacer pads (SP) and bearing pads in preventing sheath-to-sheath contact at the midplane and sheath-to-PT contact. However, the quarter plane was deemed to be the critical region due to the lack of SPs preventing excessive element bowing. This work has successfully illustrated the deformation of a CANDU fuel bundle, with contact, and its similarity with the bow profiles when compared with post-irradiation examination results and bundle heat-up tests.

A new analytical modelling method for the study of low-velocity impact response in honeycomb sandwich panels with metallic face-sheets is proposed by using Hamilton's principle. A modified Lagrange's function for this model is then proposed by extending the application of Hamilton's principle from elastic bodies to face-sheets with plasticity. The internal physical mechanism in the system is proved to be reasonable by comparing the energy converting history with published data. This study advances the understanding of the role of face-sheet plasticity in low-velocity impact response of honeycomb sandwich panels by having the ability to incorporate either the energy absorption from elastic deformation or the energy dissipation from plastic deformation of the face-sheet depending on the severity of the damage state in the dented region. The predicted maximum deflection and impact force history are compared with published experimental and finite element results, and lower and upper bounds of the maximum dent depth are obtained.

An adaptive remeshing technique using p-version finite element analysis was developed to predict the growth of cracks with highly irregular crack front geometries. In order to represent the shape of the crack front, the commonly used ellipse was replaced by a high-resolution spline. This enabled the growth of irregular crack fronts with local changes in curvature, convex and concave regions and small radii of curvature to be simulated iteratively using automatic mesh generation. The method was demonstrated for an irregular corner crack at a notch and the predicted crack front geometry was shown to be in good agreement with experimental results.

This work builds on the iodine-induced stress corrosion cracking (ISCC) model of Lewis and Kleczek by integrating the fuel performance model fuel and sheath modeling tool (FAST) to provide thermal and mechanical analysis of the fuel sheath, which was previously required as input parameters. The iodine transport methodology of the Lewis–Kleczek model has been modified to utilize the more mechanistic diffusion model in FAST, and an empirical surface multiplier term has been derived to predict iodine release rates under normal operating conditions based on measured release rates from in-reactor sweep gas tests. A fracture mechanics analysis is implemented using threshold stress intensity values and crack growth rates reported in literature. A correlation to predict crack initiation has been derived by analysis of a database of power histories with known ISCC defects. This correlation is based on the change in sheath hoop strain during a power ramp and is shown to be more accurate at discerning failure versus nonfailure than the correlations used in the previous Lewis–Kleczek model and FUELOGRAMS. Failure time prediction of the model is compared against power ramp test FFO-104 performed at the National Research Experimental (NRX) reactor. Fuel failure is predicted to occur 27% faster than experimentally measured; failure time is explored in a series of sensitivity studies to suggest areas for further development. The model improvements represent a step forward in the mechanistic modeling of stress corrosion cracking (SCC) in pressurized heavy water reactor (PHWR) nuclear fuel.

The use of eddy current (EC) arrays to detect damage in sandwich panels, such as disbonding of the carbon fiber reinforced polymer (CFRP) face-sheet to the core, is investigated. It is shown that the array is very sensitive to slight core crush and can readily find small dents and disbonds. At the same time, the eddy current array can look much deeper into the honeycomb to detect defects such as tears. The phase map of the EC signal can be used in some cases to distinguish between different types of damage. EC arrays offer the ability to rapidly scan large areas of CFRP panels.

The effect of the adhesive geometry on the impact damage of adhesively bonded aluminum sandwich panels was studied experimentally and numerically. Both the physical testing and the finite element results show that the core damage depth is dependent on the adhesive fillet height and can be linearly correlated with the dent depth. The focus of the finite element simulation is on representing the adhesive fillet height and the constraint that it provides for the cell walls at the interface with the face-sheet. A hybrid method combining the numerical study and the Bayesian method was proposed to predict the range of core damage depths from the numerical simulations based on the variability of the adhesive fillets. The variation of the adhesive fillet height between different cell walls significantly influences the core damage depth and cannot be ignored when predicting the range of the core damage depth. The non-uniform distribution of the adhesive fillets causes the core damage depth to vary within a range of up to 20% of the core thickness for the panel configurations considered in this study.

A 3D scanning method is proposed for the measurement of surface damage on aircraft structural panels. Dent depth measurements were shown to be within 0.04 ± 0.06 mm (95%) of those taken using a Starrett 643J dial depth gauge based on 54 flat panel dents, and 0.04 ± 0.05 mm (95%) based on 74 curved panel dents. Dent depths were quantified by the difference between a point cloud rendering of the damaged surface and a surface fit approximating the original, undamaged surface. Convergence studies were used to evaluate the accuracy of the surface fit, enabling this technique to be used as a stand-alone inspection method. Image processing was used to measure dent length and area, and the results showed that this method is more efficient and reliable compared to manual methods. This novel non-destructive evaluation technique thus demonstrates potential to enable the timely extraction of surface dent measurements during on-site aircraft inspections.

Aerospace sandwich panels subject to low-velocity impacts can develop damage in both the face sheet and the core that may compromise structural integrity. The standard repair manual for each aircraft specifies allowable damage limits that consider surface dents and the presence of disbonds or delaminations. A strategy for identifying the types of impact damage according to the standard repair manual using a series of non-destructive inspection techniques is presented for honeycomb sandwich panels with composite facesheets and metallic core. Eddy current is first used to identify impact sites, followed by thermography to detect disbonds or delaminations and laser scanning to measure the size of surface dents. Images of damage sites in honeycomb sandwich panels from low-velocity impacts created using a drop tower show that when used together, these inspection methods are able to detect and distinguish between three different damage configurations; core crush, surface dents, and adhesive disbonds or face sheet ply delaminations. While these three methods are commonly used non-destructive evaluation techniques, this paper focuses on their combined use in identifying different types of impact damage in aerospace sandwich panels in-situ according to the requirements of the standard repair manual. The combined methods are shown to address requirements of dent depth, dent area, and the identification of disbonds or delaminations in a more rapid, repeatable and robust manner than is currently employed.

The load carrying capability of honeycomb sandwich panels has been shown to be reduced due to core damage resulting from low-velocity impacts. Knowing the size of the core damage region would aid in predicting the reduction in residual strength of such panels. The current study determines the width and depth of the core damage region in four aluminum panels containing barely visible impact damage. Different honeycomb core configurations and impact parameters were considered experimentally using a pendulum apparatus and numerically using dynamic finite element simulations. One retired aircraft panel that contained damage from in-service impacts was also considered. Destructive sectioning showed that the core damage depth was constant within each panel for different depths of surface dents, impact velocity, impactor radius and mass. The depth of the core damage was proven to be dependent only on the height of the adhesive fillet that joins the facesheet to the core. The width of the core damage was contained to a region directly below the surface dent in the facesheet. The experimental and numerical results also showed that the configuration of the cellular core did not have a significant influence on the core damage depth.

This paper presents the interlaminar normal stress distribution across the width of L-shaped composite components in bending, with special emphasis on the free edge effects and the induced loading due to boundary conditions. As L-shaped composite components are becoming more commonplace in engineering applications, it is of vital importance to understand their failure mechanisms and stress fields. An extensive literature review reveals that interlaminar delamination is the critical failure mode for curved composite components. Despite this, most literature only provides a two-dimensional analysis of the interlaminar stresses and the free edge effects and asymmetry due to induced torsion are not considered. This paper uses three-dimensional finite element modelling to address the following with respect to L-shaped laminate components in bending: (i) radial stress due to the presence of curvature, (ii) effect of lay-up on the free edge effect, (iii) determining how induced torsion affects the interlaminar normal stress distribution. The necessity of using a three-dimensional model to analyse the interlaminar normal stress across the width of a curved composite component is discussed. It was concluded that using a two-dimensional model could underpredict the interlaminar normal stresses by up to 37% for the lay-ups analysed in this research.

Carbon-epoxy and other composite materials are widely used in the design of aircraft and other structural components, but can be susceptible to delamination at free edges due to the presence of out-of-plane stresses. Finite element (FE) modeling can be used to predict the magnitude and location of these edge stresses, but the often used macroscopic approach to representing laminates can result in predictions that vary considerably due to the numerical singularities that are created. In this study, two macroscopic FE approaches are tested for thin angle-ply laminates in tension, and recommendations are made for the most effective method for evaluating the edge-effect phenomenon early in the design process. It has been determined that perfect interface models cannot be used for predicting maximum interlaminar stresses at free edges. Perfect interface models can effectively predict the relative magnitude of the interlaminar shear stress, τxz but resin interface models must be used for the interlaminar normal stress, σz.

Aluminum hexagonal honeycomb panels are commonly used in the aerospace industry to reduce weight due to their high stiffness to mass ratio. The panels are commonly involved in incidents where they are dented in the out-of-plane direction which causes plastic deformation in the face-sheet and buckling collapse of the thin repeating cell-walls in the core. This paper investigates the responses to barely-visible-impact-damage (BVID) in aluminum honeycomb sandwich panels in the out-of-plane direction with attention to the structural adhesive. The structural adhesive forms a fillet shape between the face-sheet and the aluminum core during the curing process and in some cases can encompass over 50% of the honeycomb core thickness. The adhesive fillets become stiff after curing and are able to brace the thin metallic cell-walls and prevent buckling in sections of the core enclosed in adhesive. It was shown that larger fillets cause the damage to occur deeper in the core. Force-displacement data collected from quasi-static experiments showed that as the amount of adhesive used in honeycomb panels was increased, the peak force required to produce a specified maximum dent depth increased as well. Absorbed energy positively correlated with an increasing quantity of adhesive; showing improvements of up to 50% when comparing panels with the largest amount of adhesive and no adhesive. This paper provides relationships between the quantity of adhesive used to fabricate metallic honeycomb sandwich panels and the damage resistance and energy absorption under BVID conditions.

Sandwich panels consisting of aluminum face-sheets and honeycomb core are widely used in transportation systems. The composite structure has a high stiffness and strength, but it is susceptible to impacts in service. An experimental investigation of surface deformation and core damage in a honeycomb sandwich panel subjected to three different low-velocity impact energies was undertaken. Surface damage evaluation using 3D laser scanning technology was conducted to assess the surface damage and a comparison was made with two typical indentation profiles which were proposed mathematically in the past. The experimental dent profile shows a good agreement with one of the two analytical dent profiles. The impacted sandwich panel was then cut transversely to study the damage inside the honeycomb core. The number of buckled or collapsed folds under the damaged top face-sheet and the depth of the core damage were utilized as two parameters to quantify the damage of the honeycomb core. It is concluded that the core damage depth and the number of folds is independent of impact energy and is constant within each dent.

Horizontally oriented fuel bundles, such as those in CANada Deuterium Uranium (CANDU) reactors present unique modeling challenges. After long irradiation times or during severe transients the fuel elements can laterally deform out of plane due to processes known as bow and sag. Bowing is a thermally driven process that causes the fuel elements to laterally deform when a temperature gradient develops across the diameter of the element. Sagging is a coupled mechanical and thermal process caused by deformation of the fuel pin due to creep mechanisms of the sheathing after long irradiation times and or high temperatures. These out-of-plane deformations can lead to reduced coolant flow and a reduction in coolability of the fuel bundle. In extreme cases element-to-element or element-to-pressure tube contact could occur leading to reduced coolant flow in the subchannels or pressure tube rupture leading to a loss of coolant accident. This paper evaluates the capability of the Multiphysics Object-Oriented Simulation Environment (MOOSE) framework developed at the Idaho National Laboratory to model these deformation mechanisms. The material model capabilities of MOOSE and its ability to simulate contact are also investigated.

This paper documents the extension of the Fuel And Sheath modeling Tool (FAST) for modeling transient conditions and presents a proof of concept validation exercise. This validation compares the predictions of FAST and ELESTRES/ELOCA fuel performance codes against experimental measurements from a simulated loss of coolant test conducted at Chalk River Laboratories. The comparison includes in reactor measurements of fuel temperature and internal gas pressure along with post irradiation (hot-cell) measurements of cladding deformation and oxidation. The results of this comparison show that the single pellet version of the FAST code was insufficient to fully account for strong axial dependence along an element. This was addressed by modifying the FAST code geometry to allow an arbitrary number of pellets in the fuel-stack. This full element version demonstrated improved agreement with measured cladding deformation, including circumferential ridging effects.

Surface damage on honeycomb aircraft panels is often measured manually, and is therefore subject to variation based on inspection personnel. Eddy current testing (ECT) is sensitive to variations in probe-to-specimen spacing, or lift-off, and is thus promising for high resolution profiling of surface damage on aluminum panels. Lower frequency testing also allows inspection through the face sheet, an advantage over optical 3D scanning methods. This paper presents results from the ECT inspection of surface damage on an approximately flat aluminum honeycomb aircraft panel, and compares the measurements to those taken using optical 3D scanning technology. An ECT C-Scan of the dented panel surface was obtained by attaching the probe to a robotic scanning apparatus. Data was taken simultaneously at four frequencies of 25, 100, 400 and 1,600 kHz. A reference surface was then defined that approximated the original, undamaged panel surface, which also compensated for the effects of specimen tilt and thermal drift within the ECT instrument. Data was converted to lift-off using height calibration curves developed for each probe frequency. A damage region of 22,550 mm2 area with dents ranging in depth from 0.13-1.01 mm was analyzed. The method was accurate at 1,600 kHz to within 0.05 mm (2s) when compared with 231 measurements taken via optical 3D scanning. Testing at 25 kHz revealed a 3.2 mm cell size within the honeycomb core, which was confirmed via destructive evaluation. As a result, ECT demonstrates potential for implementation as a method for rapid in-field aircraft panel surface damage assessment.

During off-normal conditions, the temperature transient experienced by a CANDU fuel bundle may lead to deformation of the bundle via fuel element bowing between the spacer pads due to sub-channel temperature gradients, or fuel element sagging under gravity. The resulting deformation could impact the coolant flow distribution through the bundle, causing further degradation of the fuel cooling. Moreover, the deformation of the fuel elements may also lead to contact with the pressure tube, leading to localized heating that may compromise the integrity of the pressure tube. The objective of this work is to create a 3D finite element model to analyze bundle deformation behaviour under transient conditions. The methodology and preliminary results are presented herein.

The effect of various p-method finite element model parameters on the prediction of planar crack front geometry in three dimensional structures is evaluated. An automatic crack growth method was developed using the commercial software StressCheck coupled with Microsoft Excel. The geometry of the evolving crack front was predicted using iterations of three dimensional finite element models to determine the stress intensity values at discrete points along the crack front and incrementing the crack length locally using the Paris growth law. Modeling parameters such as the global mesh size, p-level, crack tip mesh size, and the extraction radius used for determining the stress intensity factors were all considered. The effect of these parameters on the predicted shape of a growing elliptical crack front was evaluated through a series of simulations of a corner crack emanating from a centered hole in a plate. A single model of a through crack with a straight crack front was also used to compare the effects of the model parameters on a different geometric configuration. This study has shown that the corner crack geometry and the growth series are more sensitive to the number of elements along the crack front and the extraction radius than the through thickness crack. In addition, convergence of the predicted crack front geometry does not guarantee convergence of the distribution of stress intensity factors along the crack front. It was also determined that for the models considered, using a smaller overall mesh size is a more effective way of decreasing discretization error than increasing the p-level.

This study focuses on the effects of prestrain magnitude on 3 mm AA5754 sheet in bending at nominal applied strain rates of 0.001/s and 0.1/s. The necessity of incorporating prestrain and strain rate effects into numerical simulations of bending is also evaluated. A series of experimental bend tests using axial compression in the longitudinal material direction were performed following plane strain prestrain in the transverse material direction. An inelastic buckling mode of deformation was produced with the peak buckling load increasing and the minimum load decreasing with larger magnitudes of prestrain. A semi-empirical material model, referred to as the Voce-MA model, was developed which incorporates strain rate-sensitivity of the flow stress, the prestrain magnitude and their interaction. Simulations of the bend tests using this material model were then performed in LS-DYNA at nominal applied strain rates of 0.001/s and 0.1/s for samples with 0, 3, 6 and 12% plane strain prestrain. It was shown that for AA5754 sheet in bending, prestrain effects must be considered in terms of current sheet thickness and material hardening. While the peak and minimum loads are not strain rate sensitive at the low rates used in this study, a rate-dependant material model is still necessary in order to account for the deviations in local strain rate from the applied strain rate. The Voce-MA material model was capable of representing prestrain and strain rate effects for all cases of AA5754 sheet in bending considered in this study.

The effect of prestrain with a path change on the strain rate sensitivity of AA5754 sheet was investigated. Prestrain magnitudes between 0% and 12% were applied in plane strain in either the transverse or longitudinal (rolling) material direction. Samples were then loaded in uniaxial tension in the longitudinal direction at strain rates of 0.001/s and 0.1/s. Results show that when a path change is involved between prestrain and subsequent uniaxial loading, the strain rate sensitivity of the hardening rate at 0.1/s compared to 0.001/s is reduced. The rate sensitivity of the yield stress remains constant with increasing magnitudes of prestrain, while the rate sensitivity of the elongation to failure decreases with increasing prestrain. A permanent softening of the flow stress is also observed, which is greater when the path change is combined with a change in orientation.