Eighth DOD/NASA/FAA Conference on Fibrous Composites in Structural Design, Part 2

Papers presented at the conference are compiled. The conference provided a forum for the scientific community to exchange composite structures design information and an opportunity to observe recent progress in composite structures design and technology. Part 2 contains papers related to the following subject areas: the application in design; methodology in design; and reliability in design

The bi-composite transition joint

The application of advanced composite materials to high performance structure frequently results in the desire to fabricate a structure from more than one composite system in order to tailor the composite material capabilities to the design requirements. The bi-composite transition provides a means of joining two different composite structural systems without the weight and complexity of mechanical attachments. The monolayer plies or combinations of plies of one composite system are interleaved with and bonded to the plies of the adjacent composite system, thereby providing a direct load transfer between the two composite structures

Simplified design procedures for fiber composite structural components/joints

Simplified step-by-step design procedures are summarized, which are suitable for the preliminary design of composite structural components such as panels (laminates) and composite built-up structures (box beams). Similar procedures are also summarized for the preliminary design of composite bolted and adhesively bonded joints. The summary is presented in terms of sample design cases complemented with typical results. Guidelines are provided which can be used in the design selection process of composite structural components/joints. Also, procedures to account for cyclic loads, hygrothermal effects and lamination residual stresses are included

Innovative design of composite structures: The use of curvilinear fiber format in composite structure design

The gains in structural efficiency are investigated that can be achieved by aligning the fibers in some or all of the layers in a laminate with the principal stress directions in those layers. The name curvilinear fiber format is given to this idea. The problem studied is a plate with a central circular hole subjected to a uniaxial tensile load. An iteration scheme is used to find the fiber directions at each point in the laminate. Two failure criteria are used to evaluate the tensile load capacity of the plates with a curvilinear format, and for comparison, counterpart plates with a conventional straightline fiber format. The curvilinear designs for improved tensile capacity are then checked for buckling resistance. It is concluded that gains in efficiency can be realized with the curvilinear format

Development of Design Data for Propulsion PMR-15 Composites

The continuing development of PMR-15 composite materials and their associated design properties is pacing the implementation of this technology on commercial aircraft. The guidelines that the FAA has issued regarding the certification of advanced composite structures are very significant with respect to future PMR-15 research and development activities. The FAA has issued an advisory circular dated 1-5-83 concerning guidelines for composite aircraft structures. Of particular significance to PMR-15 technology development is the reliance on combined environmental exposure and component testing, coupled with the stipulation that reliance on previous experience be limited to where common structures and materials have been used for a similar function. Critical environmental exposures for commercial propulsion structures include 50,000 cycle service life, exposure to skydrol, moisture and other fluids, and nacelle fire conditions

Composite structural materials

The use of filamentary composite materials in the design and construction of primary aircraft structures is considered with emphasis on efforts to develop advanced technology in the areas of physical properties, structural concepts and analysis, manufacturing, and reliability and life prediction. The redesign of a main spar/rib region on the Boeing 727 elevator near its actuator attachment point is discussed. A composite fabrication and test facility is described as well as the use of minicomputers for computer aided design. Other topics covered include (1) advanced structural analysis methids for composites; (2) ultrasonic nondestructive testing of composite structures; (3) optimum combination of hardeners in the cure of epoxy; (4) fatigue in composite materials; (5) resin matrix characterization and properties; (6) postbuckling analysis of curved laminate composite panels; and (7) acoustic emission testing of composite tensile specimens

The use of damage as a design parameter for postbuckling composite aerospace structures

Advanced fibre-reinforced polymer composites have seen a rapid increase in use in aircraft structures in recent years due their high specific strength and stiffness, amongst other properties. The use of postbuckling design, where lightweight structures are designed to operate safely at loads in excess of buckling loads, has been applied to metals for decades to design highly efficient structures. However, to date, the application of postbuckling design in composite structures has been limited, as today’s analysis tools are not capable of representing the damage mechanisms that lead to structural collapse of composites in compression. The currently running four-year European Commission Project COCOMAT [1] is addressing this issue, and aims to exploit the large strength reserves of composite aerospace structures through a more accurate prediction of collapse. A methodology has been developed to analyse the collapse of composite structures that is focused on capturing the critical damage mechanisms. One aspect of the methodology is a global-local analysis technique that uses a strength criterion to predict the initiation of interlaminar damage in intact structures. Another aspect of the approach was developed for representing the growth of a pre-existing interlaminar damage region, and is based on applying multi-point constraints in the skin-stiffener interface that are controlled using fracture mechanics calculations. A separate degradation model was also included to model the in-plane ply damage mechanisms of fibre fracture, matrix cracking and fibre-matrix shear that uses a progressive failure approach. The complete analysis methodology was implemented in MSC.Marc v2005r3 using several user subroutines, and has been validated with a range of experimental tests, including fracture mechanics coupons [2], single-stiffener specimens [3] and multi-stiffener curved panels [4]. The developed methodology was used to design and analyse fuselage-representative composite panels in various pre-damaged configurations. Two panel designs were investigated, D1 and D2, which both consisted of a curved skin adhesively bonded to blade-shaped stiffeners. For the D1 panel, the pre-damage applied was a full-width skin-stiffener debond created using a Teflon insert in the adhesive layer, whilst the D2 panel was investigated with Barely Visible Impact Damage (BVID). For both panels, parametric studies were conducted using the developed methodology in order to recommend a damaged configuration suitable for experimental testing. For the D1 panel, a 100 mm length debond was selected, and the location of the damage was investigated, whilst for the D2 panel both the location and the representation of damage was varied. Based on these parametric studies, two pre-damaged configurations of the D1 panel and one pre-damaged D2 configuration were selected for experimental testing. The selected pre-damaged configurations were manufactured by Aernnova Engineering Solutions and manufactured at the Institute of Composite Structures and Adaptive Systems at the German Aerospace Center (DLR) as part of the COCOMAT project. Following manufacture, panel quality was inspected with ultrasonic and thermographic scanning and panel imperfection data was measured using the three-dimensional (3D) optical measurement system ATOS. During the test, measurements were taken using displacement transducers, strain gauges, the 3D optical measuring system ARAMIS, and optical lock-in thermography. Under compression, the panels developed a range of buckling mode shapes, and the progression of damage was monitored leading to structural collapse. In comparison with the experimental results, the analysis methodology was shown to give accurate predictions of the load-carrying behaviour, damage development and collapse load of both panels. The results demonstrated the capability of the developed tool to capture the critical damage mechanisms leading to collapse in composite structures. The advanced analysis methodology also allowed for damage to be used as a design parameter in postbuckling structures, either in the comparative analysis context of a design procedure, to assess the damage tolerance of a design, or as pre- and post-test simulations of intact and pre-damaged structures. More broadly, the results demonstrated the potential of postbuckling composite structures, and the large strength reserve available in the postbuckling region. The success of the developed analysis methodology and the potential of postbuckling composite structures have application for the next generation of lightweight aerospace structures

Integrated analysis and design of thick composite structures for optimal passive damping characteristics

The development of novel composite mechanics for the analysis of damping in composite laminates and structures and the more significant results of this effort are summarized. Laminate mechanics based on piecewise continuous in-plane displacement fields are described that can represent both intralaminar stresses and interlaminar shear stresses and the associated effects on the stiffness and damping characteristics of a composite laminate. Among other features, the mechanics can accurately model the static and damped dynamic response of either thin or thick composite laminates, as well as, specialty laminates with embedded compliant damping layers. The discrete laminate damping theory is further incorporated into structural analysis methods. In this context, an exact semi-analytical method for the simulation of the damped dynamic response of composite plates was developed. A finite element based method and a specialty four-node plate element were also developed for the analysis of composite structures of variable shape and boundary conditions. Numerous evaluations and applications demonstrate the quality and superiority of the mechanics in predicting the damped dynamic characteristics of composite structures. Finally, additional development was focused on the development of optimal tailoring methods for the design of thick composite structures based on the developed analytical capability. Applications on composite plates illustrated the influence of composite mechanics in the optimal design of composites and the potential for significant deviations in the resultant designs when more simplified (classical) laminate theories are used

Design and evaluation of three-phase fibrous composite structures

Study reveals composite combination evaluations for boron binder reinforcements for unidirectly reinforced boron/epoxy, glass binder reinforcements for unidirectionally reinforced boron/epoxy, and glass binder reinforcements for unidrectionally reinforced glass/epoxy

Lessons learned for composite structures

Lessons learned for composite structures are presented in three technology areas: materials, manufacturing, and design. In addition, future challenges for composite structures are presented. Composite materials have long gestation periods from the developmental stage to fully matured production status. Many examples exist of unsuccessful attempts to accelerate this gestation period. Experience has shown that technology transition of a new material system to fully matured production status is time consuming, involves risk, is expensive and should not be undertaken lightly. The future challenges for composite materials require an intensification of the science based approach to material development, extension of the vendor/customer interaction process to include all engineering disciplines of the end user, reduced material costs because they are a significant factor in overall part cost, and improved batch-to-batch pre-preg physical property control. Historical manufacturing lessons learned are presented using current in-service production structure as examples. Most producibility problems for these structures can be traced to their sequential engineering design. This caused an excessive emphasis on design-to-weight and schedule at the expense of design-to-cost. This resulted in expensive performance originated designs, which required costly tooling and led to non-producible parts. Historically these problems have been allowed to persist throughout the production run. The current/future approach for the production of affordable composite structures mandates concurrent engineering design where equal emphasis is placed on product and process design. Design for simplified assembly is also emphasized, since assembly costs account for a major portion of total airframe costs. The future challenge for composite manufacturing is, therefore, to utilize concurrent engineering in conjunction with automated manufacturing techniques to build affordable composite structures. Composite design experience has shown that significant weight savings have been achieved, outstanding fatigue and corrosion resistance have been demonstrated, and in-service performance has been very successful. Currently no structural design show stoppers exist for composite structures. A major lesson learned is that the full scale static test is the key test for composites, since it is the primary structural 'hot spot' indicator. The major durability issue is supportability of thin skinned structure. Impact damage has been identified as the most significant issue for the damage tolerance control of composite structures. However, delaminations induced during assembly operations have demonstrated a significant nuisance value. The future challenges for composite structures are threefold. Firstly, composite airframe weight fraction should increase to 60 percent. At the same time, the cost of composite structures must be reduced by 50 percent to attain the goal of affordability. To support these challenges it is essential to develop lower cost materials and processes