Impact Damage Tolerance of Composite Laminates with Through-The-Thickness Stitches

The ability of through-the-thickness stitches to contain damage during a low-velocity impact event and increase the residual strength of stitched panels was investigated in this research. The impact response, spread of interlaminar delaminations, dent depth, surface damage, and static residual strength after impact were studied for carbon-epoxy specimens fabricated from HTS40 TohoTenax standard modulus fibers, stitched together with Vectran 1200 denier thread and infused with API 1078 resin after through-the-thickness stitching. Three different stitch patterns were used to explore the ability to arrest impact damage during and after impact. Simply supported boundary conditions were maintained during the impact testing. Non-destructive evaluations were performed using ultrasonic C-scans and X-ray computed tomography (CT) imaging to determine the shape, size, and location of delaminations. Results indicate that while the dynamic response during the impact event was almost the same for the unstitched specimens and all stitch patterns considered, the extent of delamination and the compression strength after impact varied greatly. For both the 60 J and 80 J impact energies, the delamination area was significantly less for the stitched specimens than for the unstitched specimens, but the range of delamination areas among the stitch patterns was much larger for the lower impact energy than for the greater impact energy. Similarly, while the presence of stitching influenced the compression after impact strength, the strength values for all stitch patterns were very similar. These results are a step toward quantifying the influence of through-the-thickness stitching

Modeling Progressive Damage Using Local Displacement Discontinuities Within the FEAMAC Multiscale Modeling Framework

A method for performing progressive damage modeling in composite materials and structures based on continuum level interfacial displacement discontinuities is presented. The proposed method enables the exponential evolution of the interfacial compliance, resulting in unloading of the tractions at the interface after delamination or failure occurs. In this paper, the proposed continuum displacement discontinuity model has been used to simulate failure within both isotropic and orthotropic materials efficiently and to explore the possibility of predicting the crack path, therein. Simulation results obtained from Mode-I and Mode-II fracture compare the proposed approach with the cohesive element approach and Virtual Crack Closure Techniques (VCCT) available within the ABAQUS (ABAQUS, Inc.) finite element software. Furthermore, an eccentrically loaded 3-point bend test has been simulated with the displacement discontinuity model, and the resulting crack path prediction has been compared with a prediction based on the extended finite element model (XFEM) approach

IMECE2002-33784 PROCESS MODELING OF SHAPE ROLLING FOR AEROSPACE INDUSTRY

ABSTRACT Shape rolling of seamless rings constitutes an efficient manufacturing process offering excellent material yield, energy conservation, and component production, which require a minimum of subsequent machining operations. An increasing number of rings are being produced from high temperature Titanium and Nickel based super alloy materials for gas turbine engine parts such as vane and fan casings, exhaust casings, turbine shrouds, and combustion liners. With the increasing cost of super alloy raw materials and growing demand for costcompetitive parts, the importance of ring rolling to contoured shape becomes an increasingly important factor. This paper describes a new process modeling technique based on Upper Bound Elemental Technique (UBET) for shape rolling of super alloys. This tool provides a new design paradigm for an industry relying to heavily on designer experience and cut-and-try methods. As a rapid software tool to aid designers in developing ring-rolling process schedules thereby helping in reducing the design and analysis cycle time, the potential to capture the unique 3-D flow situation experienced in shape rolling of seamless rings is being explored. Numerical results have been compared with data available for high temperature alloys such as IN718 and Ti6Al-4V

3D Progressive Damage Modeling for Laminated Composite Based on Crack Band Theory and Continuum Damage Mechanics

A simple continuum damage mechanics (CDM) based 3D progressive damage analysis (PDA) tool for laminated composites was developed and implemented as a user defined material subroutine to link with a commercially available explicit finite element code. This PDA tool uses linear lamina properties from standard tests, predicts damage initiation with an easy-to-implement Hashin-Rotem failure criteria, and in the damage evolution phase, evaluates the degradation of material properties based on the crack band theory and traction-separation cohesive laws. It follows Matzenmiller et al.'s formulation to incorporate the degrading material properties into the damaged stiffness matrix. Since nonlinear shear and matrix stress-strain relations are not implemented, correction factors are used for slowing the reduction of the damaged shear stiffness terms to reflect the effect of these nonlinearities on the laminate strength predictions. This CDM based PDA tool is implemented as a user defined material (VUMAT) to link with the Abaqus/Explicit code. Strength predictions obtained, using this VUMAT, are correlated with test data for a set of notched specimens under tension and compression loads

Special session honoring the contributions of Dr. T. Kevin O'Brien, NASA Langley Research Center

Aeronautical composite stiffened structures have the capability to carry loads deep into postbuckling, yet they are typically designed to operate below the buckling load to avoid potential issues with durability and structural integrity. Large out-of-plane postbuckling deformation of the skin can result in the opening of the skin-stringer interfaces, especially in the presence of defects, such as impact damage. To ensure that skin-stringer separation does not propagate in an unstable mode that can cause a complete collapse of the structure, a deeper understanding of the interaction between the postbuckling deformation and the development of damage is required. The present study represents a first step towards a methodology based on analysis and experiments to assess and improve the strength, life, and damage tolerance of stiffened composite structures subjected to postbuckling deformations. Two regions were identified in a four-stringer panel in which skin-stringer separation can occur, namely the region of maximum deflection and the region of maximum twisting. Both regions have been studied using a finite element model of a representative single-stringer specimen. For the region of maximum deflection, a seven-point bending configuration was used, in which five supports and two loading points induce buckling waves to the specimen. The region of maximum twisting was studied using an edge crack torsion configuration, with two supports and two loading points. These two configurations were studied by changing the positions of the supports and the loading points. An optimization procedure was carried out to minimize the error between the out-of-plane deformation of the representative single-stringer specimen and the corresponding region of the four-stringer panel

Study of Skin-Stringer Separation in Postbuckled Composite Aeronautical Structures

Aeronautical composite stiffened structures have the capability to carry loads deep into postbuckling, yet they are typically designed to operate below the buckling load to avoid potential issues with durability and structural integrity. Large out-of-plane postbuckling deformation of the skin can result in the opening of the skin-stringer interfaces, especially in the presence of defects, such as impact damage. To ensure that skin-stringer separation does not propagate in an unstable mode that can cause a complete collapse of the structure, a deeper understanding of the interaction between the postbuckling deformation and the development of damage is required. The present study represents a first step towards a methodology based on analysis and experiments to assess and improve the strength, life, and damage tolerance of stiffened composite structures subjected to postbuckling deformations. Two regions were identified in a four-stringer panel in which skin-stringer separation can occur, namely the region of maximum deflection and the region of maximum twisting. Both regions have been studied using a finite element model of a representative single-stringer specimen. For the region of maximum deflection, a seven-point bending configuration was used, in which five supports and two loading points induce buckling waves to the specimen. The region of maximum twisting was studied using an edge crack torsion configuration, with two supports and two loading points. These two configurations were studied by changing the positions of the supports and the loading points. An optimization procedure was carried out to minimize the error between the out-of-plane deformation of the representative single-stringer specimen and the corresponding region of the four-stringer panel

Study of Skin-Stringer Separation in Postbuckled Composite Aeronautical Structures

Aeronautical composite stiffened structures have the capability to carry loads deep into postbuckling, yet they are typically designed to operate below the buckling load to avoid potential issues with durability and structural integrity. Large out-of-plane postbuckling deformation of the skin can result in the opening of the skin-stringer interfaces, especially in the presence of defects, such as impact damage. To ensure that skin-stringer separation does not propagate in an unstable mode that can cause a complete collapse of the structure, a deeper understanding of the interaction between the postbuckling deformation and the development of damage is required. The present study represents a first step towards a methodology based on analysis and experiments to assess and improve the strength, life, and damage tolerance of stiffened composite structures subjected to postbuckling deformations. Two regions were identified in a four-stringer panel in which skin-stringer separation can occur, namely the region of maximum deflection and the region of maximum twisting. Both regions have been studied using a finite element model of a representative single-stringer specimen. For the region of maximum deflection, a seven-point bending configuration was used, in which five supports and two loading points induce buckling waves to the specimen. The region of maximum twisting was studied using an edge crack torsion configuration, with two supports and two loading points. These two configurations were studied by changing the positions of the supports and the loading points. An optimization procedure was carried out to minimize the error between the out-of-plane deformation of the representative single-stringer specimen and the corresponding region of the fourstringer panel.Aerospace Structures & Computational Mechanic