BOJCAS: bolted joints in composite aircraft structures

Bolted joints are widely used in composite aircraft structures, for their assembly. The appropriate bolted joint configuration (hole/bolt diameter, pitch, etc.) is carefully selected during the detail design phase, where high fidelity numerical models are required with substantial computational cost and time. This work presents a design criterion, which allows the selection of the bolted joint configuration during the preliminary design phase with less computational time. The developed design criterion is based on a fully parametric finite element (FE) model, built in ANSYS V19 (Canonsburg, PA, USA), of a bolted joint with progressive damage modelling (PDM) capabilities, so that the failure of the joint can be predicted. From the numerical analyses, the bearing load and the load that bypasses the hole are calculated, up to failure, for a variety of joint configurations and loading conditions. The results of each analysis are used for plotting the failure envelope for the investigated bolted-joint configuration. Consequently, a design criterion is generated for the bolted joint. The availability of these failure envelopes, as design criterion, permit the appropriate selection of the bolted-joint configuration in an earlier design phase saving valuable time and computational cost.

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.

Composite materials are used in variety of structural assemblies where joining of two or more components is imperative. Some of the joints are mechanical joints such as bolted and riveted connections. Others are fabricated by adhesive bonding, welding, brazing, and variety of other methods. In the last three decades, a number of computer codes have been developed to analyze the bolted joints in composite structures and predict the joint strength. In this paper the capabilities and limitations of these computer codes are outlined. Using some of the computer codes, the stress distribution in the hole region has been obtained and compared for a laminated plate with a circular hole in the center. This work is conducted as part of a larger effort currently underway at AdTech Systems Research, Inc. to establish a baseline to extend the joint design methodology for generalized composite systems including life prediction, thermal effects and dissimilar materials, and present the state-of-the-art approaches taken by the aerospace industry to design joints in composite structures.

Accurate representation of bolted joints in the finite element (FE) model of a built-up structure is a crucial factor in ensuring the accuracy and reliability of the dynamic behaviour prediction of the structure. Modelling of bolted joints in industrial structures has not been widely and thoroughly studied, especially in the dynamic response of the bolted joint itself. The aim of this work is to investigate several flexibility formulae for the calculation of the stiffnesses of CBUSH elements for the modelling of the bolted joints. The flexibility formulae used in this work are attributed to Huth, Tate & Rosenfeld, Boeing, Swift and Grumman. The results of the normal modes analysis of the FE models with different flexibility formulae were compared with those obtained from the experimental modal analysis (EMA). EMA was performed to measure the natural frequencies and mode shapes of the structure with bolted joints. The total errors calculated from the comparisons between FE and EMA were recorded for all flexibility formulae. The results of the investigation demonstrated that there are noticeable variations in the total errors recorded using different flexibility formulas for the CBUSH element representing the bolts in the FE model of the structure with bolted joints. It was found that the Swift’s flexibility formula is the most appropriate for the prediction of the natural frequencies of the structure with bolted joints.Accurate representation of bolted joints in the finite element (FE) model of a built-up structure is a crucial factor in ensuring the accuracy and reliability of the dynamic behaviour prediction of the structure. Modelling of bolted joints in industrial structures has not been widely and thoroughly studied, especially in the dynamic response of the bolted joint itself. The aim of this work is to investigate several flexibility formulae for the calculation of the stiffnesses of CBUSH elements for the modelling of the bolted joints. The flexibility formulae used in this work are attributed to Huth, Tate & Rosenfeld, Boeing, Swift and Grumman. The results of the normal modes analysis of the FE models with different flexibility formulae were compared with those obtained from the experimental modal analysis (EMA). EMA was performed to measure the natural frequencies and mode shapes of the structure with bolted joints. The total errors calculated from the comparisons between FE and EMA were recorded for all fle...

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.

Cost optimization of composite aircraft structures including variable laminate qualities

One of the key problems of designing composite primary aircraft structures is the problem of development of lightweight and reliable joints for such structures. For conventional bolted joints used for connection of laminated composite primary structure elements, the loads are transferred via the contact of metallic and composite parts having considerably different strength and stiffness parameters. As the local strength properties of composites are defined by polymer resins, having several times less strength as compared to the metallic alloys, the maximal load that can be transferred via the joint is limited by the properties of resins. As a consequence, the metallic parts of such joints are loaded to a very low extent, that causes low weight efficiency. In the presented paper, the approach to development of metal–composite joints based on the principle of harmonization is proposed. The harmonization principle is to minimize the disbalance of strength properties of metallic and composite parts in the local zones of contact. This principle can be realized by two ways: either by removing the resin from the zone of contact to increase local stiffness of composite part (“stiff” joint concept) or by creation of metallic part with porous structure to decrease local stiffness of the metallic part (“soft” joint concept). In this paper, the evaluation of both concepts is given, based on numerical and experimental investigations carried out in this work.

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.

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 [

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.

The weight and fuel savings offered by composite materials make them attractive not only to the military, but also to the civilian aircraft, space, and automobile industries. In these industries, bolted and riveted joints are extensively used as a primary method for structural joining. Bolted joints in composite materials have complex failure modes, and hence the demand for improving their performance exists. The main objective of this work is to improve the performance of bolted joints in composite structures by introducing nanoparticles/fibers around the expected failure zone. The literature on this issue showed shortcomings in the investigations of such materials. Most of the investigations in this field aimed to enhance the mechanical properties of epoxy materials, which cannot be used alone for high performance structural applications due to their low mechanical properties. In the present work, epoxy resin was modified with different types of nanofillers including multi walled carbon nanotubes (MWCNT). Nano-phased epoxy was used to fabricate different types of nanocomposites as well as nano-hybridized glass fiber reinforced composite laminates. Therefore, six different advanced materials were fabricated including a nanocomposite material (MWCNT/E), a quasi-isotropic nano-hybridized composite laminate (QI-GFR/MWCNT/E), a unidirectional nano-hybridized composite laminate (UD-GFR/MWCNT/E) and a control panel manufactured without nano-fillers (neat epoxy, QI-GFR/E, UD-GFR/E). The materials were characterized by tension and compression tests. The obtained properties are essential for the validation of respective finite element analysis. The results showed improvements in the tensile and compressive properties (strength and modulus) of the fabricated nanocomposites (MWCNT/E) compared with neat epoxy. The hybridized composite laminate with MWCNT showed high improvements in their mechanical properties compared to the composite laminates without nanofillers.

Fatigue characteristics of the bolted joints for unidirectional composite laminates

Aircraft structures are exposed to impact damage caused by debris and hail during their service life. One of the design concerns in composite structures is the resistance of layered surfaces to damage, which occurs from impacts with various foreign objects. Therefore, the impact localization and damage quantification of impacts should be studied and considered to address flight safety and to reduce costs associated with a regularly scheduled visual inspection. Since the structural components of the aircraft are large scale, visual inspection and monitoring are challenging and subject to human error. This paper presents a promising solution that can automatically detect and localize an impact that may occur during flight. To achieve this goal, acoustic emission (AE) is employed as an impact monitoring approach. Random forest and deep learning were adopted for training the source location models. An AE dataset was collected by conducting an impact experiment on a full-size thermoplastic aircraft elevator in a laboratory environment. A dataset consisting of AE parametric features and a dataset consisting of AE waveforms were assigned to a random forest classifier and deep learning network for the investigation of their applicability of impact source localization. The results obtained were compared using the source localization approach in previous research using a conventional artificial neural network. The analysis of results shows the random forest and deep learning leads to better event localization performance. In addition, the random forest model can provide the importance of features. By deleting the least important features, the storage required to save the input and the computing time for the random forest is greatly reduced, and an acceptable localization performance can still be obtained.

: This report contains a summary of the workshop, a list of attendees, a prepaper entitled 'Some Areas for Discussion Suggested by RAE,' and the presentations on Composite Aircraft Structures, Civil Aviation Concerns, Impact Damage, RAE Composite Certification, and the Effect of Observed Climatic Conditions on the Moisture Equilibrium Level of Fibre-Reinforced Plastics.

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.