Developing time estimating models for advanced composite manufacturing processes

Recent progress made by NASA in applications of advanced composites to aircraft structures is reviewed. Particular emphasis is given to flight service evaluation programs that are or will be accumulating experience with secondary and primary structural components on aircraft. The NASA flight service programs are conducted to obtain practical experience in design, manufacturing and operational considerations for a variety of aircraft components. The primary focus for the programs have been commercial transports. To establish confidence in advanced composites to a point where both transport builders and transport users are willing to make long-term production commitments using composites, NASA has initiated a series of programs to introduce advanced composites into airline fleet service. In implementing this program, initial applications were selective reinforcement of metal components on military aircraft. Small secondary structural components for commercial transports were added to the program; larger secondary components are being fabricated; and large primary structural components are in the design or planning stage. Results will be présented from flight service programs with composite-reinforced metal structures in the CH-54B helicopter tail cone and the C-130 transport aircraft wing boxes. Experience with advanced composite components on the L-1011, B-737 and DC-10 commercial transports will be also reported. A description of future NASA programs that are aimed at composite primary structures in commercial transport aircraft and at development of high-temperature composites technology that may be applicable to future supersonic aircraft will also be reviewed.

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.

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.

An overview of the durability design requirements and substantiation approaches, with an emphasis on aerospace applications, is provided in this chapter. The service history of both commercial and military composite aircraft structures has provided numerous examples of both good and bad designs from which current design specification and guidance are derived. Issues to be considered in the design of composites used in primary and secondary composite structures including corrosion prevention measures associated with joining composite and metallic components are described. Design details, material selection, and demonstration that designs meet performance criteria are dependent on a structure’s thermal and mechanical loading environments, service and economic life requirements, manufacturing constraints, and inspectability requirements. Based on an assessment of the durability of composite materials at the coupon level through modeling and testing, a description of the building block approach used to validate durability predictions of elements, subcomponents, components, and full-scale structural testing is given. The unique requirements for developing load spectra for accelerated full-scale durability testing for both composite and combined composite/metallic structures are discussed.

Lawrence Livermore National Laboratory and Boeing Defense and Space Group worked together to advance the understanding and aerospace application of the Resin Transfer Molding (RTM) process. Future aircraft designs will require the use of lightweight fiber composites in primary structures. RTM is the prime manufacturing process to replace existing labor and cost-intensive procedures. RTM's unique processing capabilities are that it is not labor intensive and that it provides speed and near net shape forming to complex aircraft parts. The qualities could reduce the cost of tomorrow's airline programs by several million dollars each.

: Manufacturing process models are developed for a data-driven, reconfiguarable manufacturing facility. A human assembly process model is developed to characterize assembly processes. Machine and tool resources are characterized for manufacturing processes to provide a broad characterization of manufacturing process coverage in a hypothetical final assembly factory for infantry fighting vehicles. A sample population of manufacturing process parameters and resources is collected and incorporated into a manufacturing capability and process model library. Requirements for a virtual manufacturing environment are developed, and an initial, prototype version is implemented to facilitate process model validation, configuration, and experimentation.

Interlaminar crack growth in third-generation thermoset prepreg systems

Cost optimization of composite aircraft structures including variable laminate qualities

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.

This paper describes an on going research by the authors who are seeking methods to integrate an earlier developed manufacturing process modeling method in the virtual manufacturing (VM) technology. The primary aim is to attach special manufacturability evaluation to modeling of mechanical products. A more or less generic process model is generated for a cluster of manufacturing tasks then manufacturing process entities are created by evaluation of the manufacturing process model using part, form feature, production environment, production planning and cost information. Petri net manufacturing process model representation is used. It has proved to be a versatile virtual manufacturing tool in manufacturing process planning, evaluation of manufacturability and in manufacturing environment design. A process model evaluation procedure is emphasized in the course of which feasibility and resource requirements of the manufacturing are revealed. Firstly, techniques of evaluation of manufacturability, the applied approach to integration of shape and manufacturing process modeling and the related virtual manufacturing concepts are introduced. The process model is outlined and earlier relevant publications of the authors are cited. Then problems and solutions at evaluating of process model in close connection with design and environmental changes are detailed. Following this, integration of the method into a comprehensive virtual manufacturing concept is explained. Finally, main issues and future research plans are concluded.

In this paper, the cohesive zone modeling (CZM) method has been employed to simulate static and fatigue debonding growth in adhesively bonded composite joints containing bolts as crack stoppers. The cracked-lap shear specimen (CLS) for static and fatigue loads and the wide single-lap shear (WSLS) structural element for static load have been modeled. In both cases, mixed-mode I+ II loading conditions are present in the bondline. Bonded and bonded/bolted configurations of the specimens have been analysed in order to evaluate the efficiency of bolts as crack stoppers. Numerical simulations have been performed using the LS-DYNA FE code. For the static debonding simulation, a standard bilinear traction-separation law has been applied while for the simulation of fatigue debonding growth a modified traction-separation law has been implemented using experimental data derived from fatigue tests. The modification has been realized by means of a user-defined subroutine (UMAT) having the capacity to account for crack divergence by making use of a crack front detection algorithm based on effective element’s length. A parametric study on the effect of distance between bolts and crack-tip has been performed. Additionally, the WSLS structural element has been subjected to impact load aiming to create a realistic initial disbond scenario. The residual strength of the element has been predicted by applying a quasi-static load after impact. The numerical results reveal that the presence of bolts under certain circumstances can cause crack growth retardation, under both static and fatigue loading conditions. This study emphasizes on the importance of fatigue crack growth simulation and aims to contribute to the numerical design of crack stopping features as a means of complying towards certification of adhesive bonding in primary composite aircraft structures.

Custom orthotics development process based on additive manufacturing

With the constant increase of energy costs and environmental impacts, improving the process efficiency is considered a priority issue for the manufacturing field. A wide knowledge about materials, energy, machinery, and auxiliary equipment is required in order to optimize the overall performance of manufacturing processes. Sustainability needs to be assessed in order to find an optimal compromise between technical quality of products and environmental compatibility of processes. In this new Industry 4.0 era, innovative manufacturing technologies, as the additive manufacturing, are taking a predominant role. The aim of this work is to give an insight into how thermodynamic laws contribute at the same time to improve energy efficiency of manufacturing resources and to provide a methodological support to move towards a smart and sustainable additive process. In this context, a fundamental step is the proper design of a sensing and real-time monitoring framework of an additive manufacturing process. This framework should be based on an accurate modelling of the physical phenomena and technological aspects of the considered process, taking into account all the sustainability requirements. To this end, a thermodynamic model for the direct laser metal deposition (DLMD) process was proposed as a test case. Finally, an exergetic analysis was conducted on a prototype DLMD system to validate the effectiveness of an ad-hoc monitoring system and highlight the limitations of this process. What emerged is that the proposed framework provided significant advantages, since it represents a valuable approach for finding suitable process management strategies to identify sustainable solutions for innovative manufacturing procedures.

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 structure has been used extensively in the wing of the AV-8B as one of the technological improvements to increase the operational effectiveness of this second generation V/STOL aircraft. In addition to reducing the structural weight the use of composite materials has permitted unconstrained optimization of the aerodynamic shape to enhance performance. The torque box consists of one-piece graphite/epoxy monolithic skins measuring 28 feet from tip to tip and a multi-spar graphite/epoxy substructure with sinewave webs. There are 84 composite substructure parts, the largest being 12 feet long, tapering from 14.5 inches to 6.5 inches in height. Metal fittings are used at all concentrated load points such as flap and aileron hinges and pylon, outrigger gear and wing-to-fuselage attachments. The innovative design and manufacturing concepts were first made concrete in test articles which were fabricated and tested prior to the final wing design to achieve a composite wing that meets all NAVAIR requirements. This paper discusses the design development of the torque box composite substructure, including considerations of sinewave size and shape, load carrying considerations for the webs and flanges; modifications required for the attachment of ribs, fuel lines, hydraulic and control rods; fuel flow consideration; corrosion protection for the aluminum parts; and the development of tooling and fabrication methods, including the assembly of parts bv mechanical fasteners.