Ground vehicle blunt impact damage formation to composite aircraft structures

Composite aircraft structures such as fuselage and wings are subject to impact from many sources. Ground service equipment (GSE) vehicles are regarded as realistic sources of blunt impact damage, where the protective soft rubber is used. With the use of composite materials, blunt impact damage is of special interest, since potential significant structural damage may be barely visible or invisible on the structure's outer surface. Such impact can result in local or non-local damage, in terms of internal delamination in skin, interfacial delamination between stiffeners and skin, and fracture of internal reinforced component such as stringers and frames. The consequences of these events result in aircraft damage, delays, and financial cost to the industry. Therefore, it is necessary to understand the criticality of damage under this impact and provide reliable recommendations for safety and inspection technologies. This investigation concerns a composite-metallic 4-hat-stiffened and 5-frame panel, designed to represent a fuselage structure panel generic to the new generation of composite aircraft. The test fixtures were developed based on the correlation between finite element analyses of the panel model and the barrel model. Three static tests at certain amount of impact energy were performed, in order to improve the understanding of the influence of the variation in shear ties, and the added rotational stiffness. The results of this research demonstrated low velocity high mass impacts on composite aircraft fuselages beyond 82.1 kN of impact load, which may cause extensive internal structural damage without clear visual detectability on the external skin surface.

This paper presents a simple damage-gradient based elastoplastic model with non linear isotropic hardening in order to regularize the associated initial and boundary value problem (IBVP). Using the total energy equivalence hypothesis, fully coupled constitutive equations are used to describe the non local damage induced softening leading to a mesh independent solution. An additional partial differential equation governing the evolution of the non local isotropic damage is added to the classical equilibrium equations and associated weak forms derived. This leads to discretized IBVP governed by two algebric systems. The first one, associated with equilibrium equations, is highly non linear and can be solved by an iterative Newton Raphson method. The second one, related to the non local damage, is a linear algebric system and can be solved directly to compute the non local damage variable at each load increment. Two fields, linear interpolation triangular element with additional degree of freedom is terms of the non local damage variable is constructed. The non local damage variable is then transferred from mesh nodes to the quadrature (or Gauss) points to affect strongly the elastoplastic behavior. Two simple 2D examples are worked out in order to investigate the ability of proposed approach to deliver a mesh independent solution in the softening stage.

This paper describes the development of an instrumented tap test and imaging system for the inspection of defects and damages in composite aircraft structures, especially honeycomb sandwich constructions. The resulting tap test images based on the impact duration and displayed in a C- scan format can readily reveal the shape and extent of honeycomb core damages. The images also reveal considerable detail of internal substructures such as core splices, ply build-up, and changes of foam core density. Based on a grounded-spring mechanical model, the tap scan images can be converted into images that show the quantitative changes of the local stiffness. A manual tap scan imaging system was assembled for laboratory experiments. Using actual composites parts for airplanes, impact damages and substructures were imaged. In addition, effects of the tapper mass, impactor radius, tap velocity and operator dependence were studied. A compact fieldable system is being built and an automated tapping and imaging system is also under development. The technique was compared to other similar devices currently available.

Cost optimization of composite aircraft structures including variable laminate qualities

Primary and secondary aircraft structures are commonly made from laminated composite materials. Unfortunately, the interlaminar tensile and shear strengths are many orders of magnitude less than the in-plane strengths because they are driven by the strength of the matrix as opposed to that of the fibers. Moreover, the matrix strength is brittle in nature so that it is difficult to monitor an impending overload or fatigue failure. In order to avoid failures due to interlaminar damage without using high margins of safety, a Structural Health Monitoring (SHM) system can be used. While extensive literature on SHM systems for fiber-driven damage exists, there is no mention of an SHM system that specifically addresses these types of matrix-driven damage. The proposed SHM system spans the acquisition, location, reconstruction, and damage type decision of Acoustic Emission (AE) damage signals. An algorithm is presented to calculate the location of AE damage signals in a laminated composite structure using an over-determined approach to minimize location errors. Also, a novel procedure is presented to simultaneously reconstruct the loading source time series for a recorded damage signal and identify the source type. An example analytical study is presented to calculate the effects of uncertainty on the location and damage type predictions. Two experimental studies were undertaken to exercise the SHM algorithm presented in this study. The first experiment studies a mixed damage blunt impact case on a flat composite plate for different impact energies. The second experiment studies blunt impact of a reinforced composite plate that models a composite wing skin with bonded stringers and ribs.

PurposeThis paper provides a general review of automated processing methods currently being used to fabricate aircraft composite structure.Design/methodology/approachPresents a description of the Automated Tape Layer (ATL) process and the Fiber Placement (FP) process. These processes are the most “automated” of all processes being used to fabricate composite aircraft structure. Fiber Placement machines and Automated Tape Layers are composites machine tools and they are the closest comparison the composites industry has to metals machining equipment.FindingsThere is a need for more variety of composites automation and more affordable machines in the aerospace composites industry. The limited variety of automation and the cost of equipment tend to limit the spread of automation throughout the aerospace composites industry. ATL and FP are composites laminating technologies that could be adapted to a wide range of machine sizes, configurations, and price ranges.Originality/valueMore widespread use of automated processes in composites would tend to lower the cost of composite aircraft structure on a global basis.

As composite materials have been used for marine structures such as boats, ships, offshore structures, etc., the effect of FSI should be understood.In particular, FSI effect is expected to be significant for polymer composites because the water density is very comparable to the composites' densities.For example, the density of a carbon composite is approximately 50% greater than the water density.Sandwich composites consisting of very low densities of core materials have lower densities than that of water.As a result, the hydrodynamic mass associated with FSI would be very critical to composite structures under water.The growing use of composites in ship masts, superstructures, deck grates, piping, ducting, rudders, propellers, stacks, and various submarine structures requires extensive modeling and testing to help designers, builders and operators better understand composite response [Mouritz, et al., 2001].These materials are subjected to a wide spectrum of loads during manufacturing and service life.Dynamic loadings, in particular, impact type event, represent a serious design concern for use of composite.Composite structures are more susceptible to impact damage than similar metallic structures which are more ductile in nature and can absorb typically large amounts of energy without failure.Furthermore, the damage in composites from impact can go undetected even when the mechanical properties may be drastically reduced from an impact.For these reasons, numerous experimental and analytical studies have been conducted to study the dynamic response of composites subjected to impact loading [Abrate, 1994;Aslan, et al., 2003;Kwon & Wojcik, 1998].According to the review of past works, most of the research effort has been focused on low velocity impact damage, specifically, the damage predictions, and the evaluation and prediction of residual properties of damaged laminates.All of the research completed thus far has focused on damage in composites under impact loading in dry environments to support development of composites in aircraft structures.As far as dynamic response of structures under water is concerned, a great deal amount of analytical and experimental studies have been conducted on the effect of fluid force on the natural frequencies, damping ratios and mode shapes of vibrating structures in contact with fluid.This is commonly known as the Fluid Structure Interaction (FSI) problem.FSI investigations have supported many problems in submarine signaling, offshore oil structure stability, and ship structure vibrations.Through these studies, many numerical and analytical methods have been developed in order to predict the added mass and the www.intechopen.com Advances in Composites Materials -Ecodesign and Analysis 338resulting changes in natural frequency of a structure in contact with fluid.It has been determined and widely proven that the effect of fluid surrounding a structure decreases the natural frequency of a structure due to the increase in total kinetic energy of the vibrating structure and fluid from the addition of kinetic energy of the fluid.This effect can be interpreted as an added mass to the vibrating structure in the analysis of the dynamic response.Essentially as the structure vibrates, its mass is increased by the mass of the vibrating fluid with which it is in contact, consequently decreasing its natural frequency.Studies of fluid structure interaction and the added mass effect, also known as virtual mass effect, hydrodynamic mass, and hydroelastic vibration of structures, started with Lamb [Lamb, 1921] who calculated the first bending mode of a submerged circular plate.In response to a problem of submarine signaling, Lamb investigated the vibrations of a thin elastic circular plate in contact with water.In his investigation he discovered that the natural frequencies for structures in contact with fluid are lower than the frequencies in air, based on the assumption that the modes shapes are virtually the same in water as in a vacuum.The resonant frequency was determined using Rayleigh's method.Lamb's theortical results were verified experimentally [Powell & Roberts, 1923].Much later, more research was conducted for measuing and calculating natural frequencies of free vibartions of beams and plates under water [Lindholm, et al., 1965; Fu & price, 1979;Kwak, 1996].Another branch of FSI studies is underwater explosion.Both experimental and numerical studies have been conducted for metallic structures [

A hybrid structural health monitoring (SHM) system, consisting of a piezoelectric transducer and fiber optic sensors (FOS) for generating and monitoring Lamb waves, was investigated to determine their potential for damage detection and localization in composite aerospace structures. As part of this study, the proposed hybrid SHM system, together with an in-house developed algorithm, was evaluated to detect and localize two types of damage: a through thickness damage (hole of 2 mm in diameter) and a surface damage (2 mm diameter bore hole with a depth of 0.65 mm) located on the backside of the plate. The experiments were performed using an aircraft representative composite plate skin, manufactured from carbon fiber reinforced polymer (CFRP).

Composite aircraft structures commonly use bonded joints to adhere the wing and fuselage skins to the wing spars, stringers, ribs, and frames. Unfortunately the bond out-ofplane normal (peel) strength is many orders of magnitude less than the fiber strength of the adjacent laminates. Moreover, the adhesive failure is brittle in nature so that it is difficult to monitor an impending overload or fatigue failure. In order to avoid failures due to interlaminar damage without using high margins of safety, a Structural Health Monitoring (SHM) system can be used. While extensive literature on SHM systems for fiber-driven damage exists, there is no mention of an SHM system that specifically addresses these types of matrix-driven damage. The proposed SHM system spans the acquisition, location, reconstruction, and damage type decision of Acoustic Emission (AE) damage signals. An algorithm is presented to calculate the location of AE damage signals in a laminated composite structure using an over-determined approach to minimize location errors. Also, a novel procedure is presented to simultaneously reconstruct the loading source time series for a recorded damage signal and identify the source type. Parametric studies are presented for both algorithms to gauge their effectiveness under varying levels of noise, number of sensors, material properties and different types of uncertainty. Two experimental studies were undertaken to exercise the SHM algorithm presented in this study. The first experiment models a predominantly single type quasi-static damage case, and the second experiment studies a mixed damage impact case.

To expand further the composites in aircraft structures a consortium of leading aerospace organisations, led by British Aerospace (BAe), is currently engaged in the second phase of the Affordable Manufacture of Composite Primary Structures (AMCAPS II) research and development programme. The aim of this programme is to investigate innovative design, material and manufacturing concepts and technologies with the potential of providing cost‐effective composite primary wing structures. In order to support AMCAPS II tasks within BAe, process time estimating models are being developed for each of the main manufacturing processes being considered for part manufacture and assembly. The outputs from these models form the basic data from which process costs are established. The objective of this paper is to describe, using examples from the automated tape laying process, the approach to model development being applied within the AMCAPS II programme.

ABSTRACTFollowing the failure of the Fundão mine tailings dam in Brazil, approximately 32 million cubic meters of Fe ore tailings were released into the downstream riverine system. The postevent monitoring surveys implemented the use of noninvasive acoustic methods to improve the understanding of the fish biomass distribution patterns and aquatic habitat condition of the impacted reaches of the Rio Gualaxo do Norte, Rio do Carmo, and Rio Doce. The primary focus of the program was to collect hydroacoustic measurements of fish biomass, which was accompanied by sonar imaging of aquatic habitats using high‐resolution side scan and downward imaging sonar at each site visited. The biannual surveys began in April 2017 and were fundamentally a multiple control–impact design because before data (prior to the dam failure event) were not available. Up to 22 sites were visited for each survey, including reservoir and river sites. Results indicate similar levels of instream habitat between control and impact river and reservoir sites. Average volumetric biomass was not significantly different between impact sites and their corresponding controls in the August 2018 survey (latest to date). Preliminary temporal analysis of the biomass data set collected indicates that either stable or increasing levels of biomass are detected at tailings impacted sites within the Rio Doce. Integr Environ Assess Manag 2020;16:615–621. © 2020 Hydrobiology QLD Pty Ltd. Integrated Environmental Assessment and Management published by Wiley Periodicals LLC on behalf of Society of Environmental Toxicology & Chemistry (SETAC)

Compressive behaviour of CFRP laminates repaired with adhesively bonded external patches

The aircraft industry continues to increase its use of composite materials, most noteworthy in the arena of principle structural elements. This expanded use, coupled with difficulties associated with damage tolerance analysis of composites, has placed greater emphasis on the application of accurate nondestructive inspection (NDI) methods. Traditionally, a few ultrasonic-based inspection methods have been used to inspect solid laminate structures. Recent developments in more advanced NDI techniques have produced a number of new inspection options. Many of these methods can be categorized as wide area techniques that produce two-dimensional flaw maps of the structure. An experiment has been developed to assess the ability of both conventional and advanced NDI techniques to detect voids, disbonds, delaminations, and impact damage in adhesively bonded composite aircraft structures. A series of solid laminate, carbon composite specimens with statistically relevant flaw profiles are being inspected using conventional, hand-held pulse echo UT and resonance, as well as, new NDI methods that have recently been introduced to improve sensitivity and repeatability of inspections. The primary factors affecting flaw detection in laminates are included in this study: material type, flaw profiles, presence of complex geometries like taper and substructure elements, presence of fasteners, secondarily bonded joints, and environmental conditions. One phase of this effort utilized airline personnel to study Probability of Detection (POD) in the field and to formulate improvements to existing inspection techniques. In addition, advanced NDI methods for laminate inspections — such as thermography, shearography, laser ultrasonics, microwave, and phased/linear array UT — were applied to quantify the improvements achievable through the use of more sophisticated NDI. This report presents the composite laminate experiment design and the POD results for advanced NDI with comparisons to results achieved by airline inspectors using conventional UT methods. A companion report provides the full set of results from the conventional NDI testing.

Non-destructive test analysis and life and residual strength prediction of composite aircraft structures

Real time ultrasonic imaging of aerospace composites Brad Regez1, Sridhar Krishnaswamy1 Summary Ultrasonic inspection techniques have a proven record of applications during manufacturing and maintenance of metallic and composite aircraft structures. Traditional ultrasonic inspection techniques use a point-by-point scanning method to acquire a full-field image. This scanning method is slow and inadequate in satisfying modern manufacturing and maintenance time line requirements. The next generation of ultrasonic inspection eliminates the need for point-by-point scanning using an imaging array that is capable of displaying thousands of points to create a full-field image in real-time. In this study, a next generation dual-channel ultrasonic imaging system is used for real-time three-dimensional evaluation of small composite aerospace structures with programmed damage present. The system’s pulser uses a piezoelectric transducer to insonify the structure and the system’s receiver consists of two integrated data acquisition channels that are utilized for concurrent in-plane and in-depth realtime evaluation of structures. The first channel uses an ultrasonic imaging array to create a large real-time planar (X-Y plane) image of the structure while another acquisition channel analyzes in-depth (Z-axis) ultrasonic scans displaying data in an A-scan format. The images acquired with the ultrasonic system are then compared to images obtained using flash thermography. Flash thermography is capable of fast, fullfield images by measuring the material heat transfer response to thermal excitation. Flash thermography has gained acceptance as an inspection technique in the manufacturing and maintenance of aerospace composites the last couple of years and provides a good benchmark to compare the next generation ultrasonic imaging system results.