Search

Advanced structural analysis methods, known as progressive damage and failure analysis tools, are being developed to predict initiation and propagation of damage under repeated loading based on capturing individual and interacting damage modes. This work shows the ability of the progressive damage and failure analysis method implemented in CDMat software developed at the University of Texas Arlington Advanced Materials and Structures Lab to predict strength and fatigue failure of an advanced mechanically fastened aerospace structural joint, the common feature test component (CFTC)— representative of flight-critical structural attributes and failure modes—without a priori knowledge of the test result. The CFTC structural features include a composite tape skin, a composite fabric stiffener, and an aluminum rib applying pullthrough load through multiple countersunk bolts combined with the axial compression of skin and stiffener. Failure and damage predictions under static and constant-amplitude cyclic loading are compared with tests. Developed by Boeing under the Air Force Research Laboratory Composite Airframe Life Extension Program, the CFTC has been the most complex progressive damage and failure analysis validation article to date.

Aircraft composite parts are commonly manufactured from resin-saturated thermoset pre-impregnated plies laid-up over a rigid tool and consolidated and cured in an autoclave. In many applications, autoclave consolidation alone is not sufficient to remove the air, or bulk, that has been entrapped within the laminate during layup and may lead to the formation of voids in the final part. Entrapped air at ply interface can also participate in the formation of ply wrinkles, especially in curved laminates. High bulk content in curved sections is associated with an excess of ply length material that builds up during the layup. During compaction, this excess ply length may be squeezed into a tighter radius, which might lead to ply buckling and formation of wrinkles. Wrinkles and voids may significantly affect structural integrity of composite structures and increase rejection rates in the production cycle. Debulking, or vacuum consolidation, has been a common practice extensively used by aircraft manufacturers to reduce the amount of air entrapped during the layup of composite parts prior to curing. Yet, the underlying physical principles governing formation of defects during debulking are not well understood. This work is part of the Office of Naval Research (ONR) project “Physics-Based Composite Process Simulation” that seeks to fill the gaps in understanding the physical principles governing the formation and evolution of manufacturing defects. Recently, the authors of this work have introduced an approach for modeling debulking in resinsaturated prepregs where cohesive elements enriched with pore-pressure degree of freedom are inserted at ply interfaces, using simulation concepts originally developed for the analysis of hydraulic fracturing in geomechanics problems. Further improvement of the methodology combined the pore-pressure cohesive zone model with cohesive contact to capture the tacky behavior of the interface. This work presents a verification of the ability of the method to capture the physics involved during debulking at the coupon scale using a two-ply carbon/epoxy IM7/8552 debulk test specimen with a seeded wrinkle. The methodology is then further extended and verified at the element scale for the first time for simulation of debulking in a rotor blade grip test element.

В данном докладе рассматривается брачная дипломатия династии Комнинов как средство укрепления власти. Император опирался на «родственников и свойственников» — свой род и кланы, породнившиеся с правящим домом. Выявлено, что такая брачная политика была выгодна лишь в краткосрочной перспективе. This scientific report reviews the marriage diplomacy of Komnenoi as a method of strengthening power. Emperor has relied on his relatives and other families connected with the house of Komnenoi by marriage. It was found that such a policy was efficient only in a short term.

Progressive damage and failure analysis (PDFA) methods have been proposed as an essential tool for reducing the certification costs of primary and secondary composite structures. Successful application of PDFA methods have been demonstrated on test coupons and relatively small composite components; however, to achieve the goal of cost reduction for certification programs, performance of PDFA methods must be verified at larger length scales than typical aerospace allowable tests. The United States Air Force Research Laboratory (AFRL) has been actively investigating the use of these methods for aerospace durability and damage tolerant applications through the Composite Airframe Life Extension (CALE) program. Prior programs have examined applications such as unnotched tension/compression, open hole tension/compression, bearing, filled hole compression and bearing bypass. The Boeing-led CALE Project 4 seeks to build on the previous efforts and assess the tools for durability and damage tolerance of advanced composite structural features. In particular, this program focused on a Common Feature Test Component (CFTC) representative of key aircraft design considerations and failure modes. CFTC advanced structural features include a composite skin made of unidirectional tape, a fabric hat stiffener, and a mechanically fastened aluminum rib. In this work, the authors present an evaluation of fatigue simulations by CDMat – a PDFA software developed at the University of Texas at Arlington, used for the progressive damage and failure analysis of CFTC under static and cyclic loads. Both constant amplitude and block fatigue loading are considered, and residual strength capability of the element is predicted. This work will present comparison of CFTC simulation predictions with tests that have been conducted at AFRL including blind predictions of critical failure loads and fatigue damage growth for the CFTC specimens.

Composite helicopter rotor components are typically thick and often have areas with a tight radius of curvature, which make them especially prone to process-induced defects, including wrinkles and voids at ply interfaces. Such flaws cause high rejection rates in production of flight-critical components and structures. This work seeks to fill the gaps in understanding generation of the noted defects in contoured polymer–matrix composite laminates. In particular, understanding and modeling defect formation at the early stages of the manufacturing process might be the missing link to enable the development of practical engineering solutions allowing for better control of the manufacturing process of contoured composite parts. In this work, an approach based on a continuum description of the uncured prepreg material, including the initial bulk or void content, and finite element modeling (FEM) is used to simulate the consolidation process at the early stages of manufacturing of contoured laminates. The simulation predicts instabilities leading to formation of both wrinkles and voids at ply interfaces during laminate debulking or vacuum consolidation. Applicability of the method to consolidation in both closed-cavity and open-face tooling is also demonstrated. FEM results show good correlation with X-ray computed tomography data. This work also introduces a new simulation concept based on finite element and discrete modeling of voids at ply interfaces to improve the accuracy of predicting their evolution during the debulking operations.

Advanced polymeric composites are playing a major role in designing high-performance and lightweight vertical lift structures. However, uncertain residual strength and remaining useful life of the composite rotor and airframe structures due to complexity of failure mechanisms and susceptibility to manufacturing irregularities, which may be precursors to structural damage, impose risks that cannot be mitigated exclusively by time-consuming and costly experimental iterations. Validated analysis techniques accelerating design, certification, and qualification of composite structures are needed. Our team has been taking essential steps toward improving confidence in material qualification for laminated composites. The first step started with our reduced lamina test methods, short-beam shear, and small-plate twist based on digital image correlation measuring as a subset the standard material properties and, in addition, key properties that cannot be currently measured using any standard test methods. The lamina properties provide essential material input data for laminate analysis. The laminate analysis was the second step increasing confidence in material qualification. A known weakness of the existing progressive damage analysis methods is the lack of effective techniques to predict ultimate failure. The newly developed methodology relies on explicit finite element modeling and eliminates convergence issues in the ply-level progressive damage analysis methods due to severe nonlinear discontinuities after propagation of damage beyond detectable size. This work shows results of applying this methodology to nanosilica-toughened IM7/PMT-F3GHT open-hole tension strength/fatigue, open-hole compression strength/fatigue, and bearing strength multidirectional laminate configurations. The ability to predict progression of damage from initiation to ultimate strength and fatigue for advanced material systems including IM7/PMT-F3GHT carbon/epoxy reinforced by nanosilica has been demonstrated for the first time.