Principles for determining material allowable and design allowable values of composite aircraft structures

While there have been significant advances in the design and construction of composite steel‐concrete building structures, their use in regions of high seismicity has been hindered by the lack of design criteria in building codes and specifications. This has prompted initiatives in the Building Seismic Safety Council and the American Institute of Steel Construction to develop seismic design provisions for composite structures. The 1997 edition of the AISC Seismic Provisions includes a new section with requirements for composite steel‐concrete structures that are cross‐referenced by the general seismic loading and design criteria in the 1997 NEHRP Provisions and the 2000 International Building Code (final draft). Intended to complement existing provisions for steel, reinforced concrete and composite structures in the AISC‐LRFD Specification and the ACI 318 Building Code , these new provisions provide an important resource for seismic design of composite structural systems, members, and connections.

Chapter 15 - Design of Composite Structure

With the increasing demands for detailed design of composite aircraft structures, the method of covering all damages with low design allowables cannot meet the current requirements for aircraft structure design. Herein, this paper proposes a novel algorithm for design allowable determination of composite laminates by combining the damage distribution with damage factor model of design allowable, so as to provide different structures with more accurate design allowables based on their initial damages. For the composite laminates with initial delaminations, a model describing the effect of delamination size and depth position on the compression design allowable is developed and the compression design allowable of different aircraft structures are individually determined by employing abundant initial delamination statistics. Compared with the design allowable offered by the single-point method, the design allowable based on the initial damage can be increased by at least 5% to 20%, greatly improving the economic benefits of the aircraft structures and providing an important support for the damage tolerance design of the composite structures.

Composite materials have the advantages of high specific strength and high specific modulus, and are widely used in aerospace, automobiles, ships and other fields. In order to carry out the lightweight design of box structure, this paper proposes a topology-size joint optimization strategy for composite structure optimization on the basis of research and analysis of topology optimization and composite material optimization related technologies, first through topology optimization to obtain the layout form of the structure, and then establish the structural model of the composite material on this basis, and optimize the size of this to obtain the optimized composite structure.

The demand for more efficient commercial aircraft with lower maintenance and operation cost has promoted more extensive use of composite materials for a significant reduction in weight and fuel consumption. Making a structure as light as possible suggests solving an optimization problem with the goal of mass minimization. To avoid failure modes related to composite structures, such as delamination and transverse matrix crack propagation, some design guidelines are commonly suggested by aircraft industries. These guidelines are: symmetry, covering ply, disorientation, percentage rule, balance, and contiguity of the layup. The manufacturability of the final design with available technologies must be guaranteed. Different regions may be subject to different loads in a large-scale structure. Laminate thickness may vary throughout the structure depending on the distributed loads in an optimized design. Additionally, for large-scale composite structures, such as an aircraft wing or fuselage, stiffeners are added to enhance structural performance in carrying compressive and tensile loads. The stiffeners divide the structure into smaller panels. To ensure manufacturability of a composite structure, it is crucial for the plies to be continuous among adjacent panels while the laminate thickness varies. Continuity of plies in adjacent panels, which is commonly referred to as blending, is a particularly difficult constraint to deal with. A feasible composite structure has to have sufficient stiffness and strength while complying with the aforementioned design and manufacturability guidelines. The feasibility of a composite structure is evaluated through the constraints added to the optimization problem. The variables in the design of a composite structure include, but are not limited to, ply stacking sequences and thickness distribution. Depending on the application, it may also be desirable to design the shape of a structure. Holes may be required in the design of parts such as ribs of aircraft wings to reduce the weight. Therefore, the shape and the topology may be additional variables in the design of a composite structure. Optimization of a composite structure subject to the design, manufacturing, and strength related guidelines is a very challenging problem. Fulfilling the manufacturability guideline in particular has been a major goal in recent studies. This confirms the interest of the experts in the field in the applicability of their proposed methods to real-life engineering problems. However, an investigation among the existing research, revealed that these methods require a significantly large number of computations and their inherent complexity makes them inadmissible for application to real structures. This motivated performing the present research with the goal of developing a design tool that can optimize complex fiber-reinforced composite structures in practice. The present research consists of a design tool for the optimization of variable stiffness composite structures (where fibers are not steered), and a method which is developed mainly for the optimization of an aircraft wing. To optimize a variable stiffness composite structure, the proposed method separates the optimization of stacking sequences from the optimization of the thickness distribution. A set of laminates with optimized stacking sequences with respect to the problem at hand is generated and stored in a reference table known as the Stacking Sequence Table (SST). The laminates in an SST must satisfy the desired laminate design guidelines. As long as the ply stacks in a design are selected from the SST, manufacturability of the final design is guaranteed. Next, a novel level-set gradient based method is introduced for the global optimization of ply drop locations. The proposed method aims at turning the discrete optimization problem associated with the integer number of plies into a continuous problem. This is done through the way the problem is parametrized; the design variables are never rounded in this approach. The level-set function gives the optimum thickness distribution over the structure for a specific SST. The developed method is verified by its application to the well-known horseshoe panel optimization problem. To investigate the performance of the method in dealing with a real problem, the proposed method is then applied to the layup optimization of a composite skin and rib structure of a wing. Local buckling and allowable strain are considered as the constraint of the problem and a standard finite element package is used to calculate buckling factors. The structural optimization of an aircraft wing is a highly complex problem. This is due to the large number of variables as well as structural and aerodynamics constraints influencing the design of skins and stiffeners. To make it computationally more efficient, a large problem can be decomposed into several smaller subproblems (subsystems) while preserving the couplings among these subproblems. A special method is subsequently introduced for the optimization of interacting skins and ribs of an aircraft wing box.

The curing distortion is a critical determinant of the quality of integrally manufactured composite structures, playing a pivotal role in the design and fabrication of composite. This paper presents two simplified methods in predicting the curing distortion for large-scale composite aircraft structures manufactured through the autoclave process. Firstly, the refined finite element models of the two simplified methods were developed. Then, it was utilized to predict the curing distortion of Ω-shaped composite laminates. The comparative study between the experimental data and numerical results shows that the proposed second simplified method balanced the prediction accuracy and efficiency, which is urgently needed in practice. Finally, using the second simplified method, predictions were conducted for the curing distortion of practical large-scale composite skin structures. The results were in good agreement with the corresponding experiments. This study provides a new solution for the rapid iterative design of large-scale composite structures, as well as the efficient design of frame molds for their manufacturing.

The paper provides a brief overview of domestic and foreign guidelines (manuals) for the design of composite steel and concrete structures: steel-concrete slabs on profiled flooring, combined beams, and columns with rigid reinforcement. The necessity of creation of the actual manual corresponding to the modernlevel of development of construction science, normative documents and design practiceslinked to the new formulary SP 266.1325800.2016 is proved. It will facilitate the design, reduce labor expenditures and improve the reliability of composite steel and concrete structures. The new guidance provides general recommendations for the design of composite steel and concrete structures and the basic regulations for the calculations. The new guidance describes recommendations for modeling of composite steel and concrete structures and elements in the calculated complexes, the recommendations for calculation of combined beams fully concreting rectangular and T-section, partially concreting along with support slab on the lower flange of the beam, columns with rigid reinforcement, shear a connection of composite beams. Recommendations on the registration of creep, shrinkage and crack formation in the appointment of the modulus of elasticity are given. Recommendations on the use of diagrams of the state of concrete, reinforcement, and steel in the calculation of steel-concrete elements on a nonlinear deformation model are given. Recommendations on the use of the range of sheet flooring for steel-reinforced concrete slabs, as well as metal profiles as steel beams and rigid reinforcement in the cross sections of columns and combined beams, are presented. Recommendations on a design of units and details of composite steel and concrete structures are given, refined recommendations on buffer are presented. The examples of connection of steel beams with columns with rigid reinforcement are given. The examples of calculation of composite steel and concrete structures taking subject to the recommendations given in the Manual are presented.

A hybrid method for determination of fatigue limit and non-destructive evaluation of composite structures after low-velocity impact loading

The rotating bending fatigue tests were perfomed on the specimens of nodular cast iron with unnotch, semicircular notch and 60°V-notch at room temperature, 150°C, 300°C, 400°C, 550°C, and 800°C with the frequency of 3600rpm. Their values of fatigue strength were obtained at 107 cycles.The results are as follows:(1) The fatigue strength of the unnotched specimen decreases with the rising test temperatures. At about 400°C, however, the fatigue strength takes the slightly high value due to the blue brittleness. Beyond the range of 500°C to 550°C, the fatigue strength decreases rapidly.(2) The fatigue strength of the notched specimens decreases with the rising test temperatures. At about 300°C to 400°C, however, the fatigue strength shows a near constant value. Beyond 400°C, the fatigue strength decreases rapidly.(3) The value of fatigue notch factor gradually increases with the elevating temperatures above room temperature and then increases rapidly beyond about 400°C to 500°C.

. Production efficiency as well as the reduction of energy consumption and waste are the main drivers for the improvement of large-scale production in chemical industry. In addition, the demand for new processes – these days especially related to the chemical energy conversion as basis for the Energiewende – are additional triggers for innovation. Catalysis, well recognized by the public e.g. in form of the 3-way automobile catalyst, plays an important role for about 90% of all processes in chemical industry. The development of new and improved catalysts is a challenge for chemists – as is the search for new structural materials for mechanical engineers. Due to the often-necessary noble and thus expensive metals (e.g. palladium) in many catalysts, structural and chemical function have been developed separately in the two communities. Recent progress in catalytically active materials allows replacing expensive noble metals in selected catalytic processes by significant cheaper elements [1]. This offers the new perspective to combine structural and chemical functionality by innovative materials. Within this area, intermetallic compounds with their peculiar combination of crystal and electronic structure [2], represent an interesting class of materials to address this vision as will be shown in the contribution.

The characteristics of composite materials are of high importance to engineering applications; therefore the increasing use as a substitute for conventional materials, especially in the field of aircraft and space industries. It is a known fact that researchers use finite element programs for the design and analysis of composite structures, use of symmetrical conditions especially in complicated structures, in the modeling and analysis phase of the design, to reduce processing time, memory size required, and simplifying complicated calculations, as well as considering the response of composite structures to different loading conditions to be identical to that of metallic structures. Finite element methods are a popular method used to analyze composite laminate structures. The design of laminated composite structures includes phases that do not exist in the design of traditional metallic structures, for instance, the choice of possible material combinations is huge and the mechanical properties of a composite structure, which are anisotropic by nature, are created in the design phase with the choice of the appropriate fiber orientations and stacking sequence. The use of finite element programs (conventional analysis usually applied in the case of orthotropic materials) to analysis composite structures especially those manufactured using angle ply laminate techniques or a combination of cross and angle ply techniques, as well considering the loading response of the composite structure to be identical to that of structures made of traditional materials, has made the use of, and the results obtained by using such analysis techniques and conditions questionable. Hence, the main objective of this paper is to highlight and present the results obtained when analyzing and modeling symmetrical conditions as applied to commercial materials and that applied to composite laminates. A comparison case study is carried out using cross-ply and angle-ply laminates which concluded that, if the composition of laminate structure is pure cross-ply, the FEA is well suited for predicting the mechanical response of composite structure using principle of symmetry condition. On the other hand that is not the case for angle-ply or mixed-ply laminate structure.

PurposeThe purpose of this paper is to investigate the mechanical properties’ behaviors of woven composite cut-out structures with specific parameters. Because of the complexity of micro-scale and meso-scale structure, it is difficult to accurately predict the mechanical material behavior of woven composites. Numerical simulations are increasingly necessary for the design and optimization of test procedures for composite structures made by the woven composite. The results of the proposed method are well satisfied with the results obtained from the experiment and other studies. Moreover, parametric studies on different plates based on the stacking sequences are investigated.Design/methodology/approachA multi-scale modeling approach is suggested. Back-propagation neural networks (BPNN), radial basis function (RBF) and least square support vector regression are integrated with efficient global optimization (EGO) to reduce the weight of assigned structure. Optimization results are verified by finite element analysis.FindingsCompared with other similar studies, the advantage of the suggested strategy uses homogenized properties behaviors with more complex analysis of woven composite structures. According to investigation results, it can be found that 450/−450 ply-orientation is the best buckling load value for all the cut-out shape requirements. According to the optimal results, the BPNN-EGO is the best candidate for the EGO to optimize the woven composite structures.Originality/valueA multi-scale approach is used to investigate the mechanical properties of a complex woven composite material architecture. Buckling of different cut-out shapes with the same area is surveyed. According to investigation, 45°/−45° ply-orientation is the best for all cut-out shapes. Different surrogate models are integrated in EGO for optimization. The BPNN surrogate model is the best choice for EGO to optimization difficult problems of woven composite materials.

The damage tolerance of composite structures is intimately linked to the morphology and extent of flaws or damage. Impacts are a commonly occurring source of threat to composites that can produce a “seed” damage state which inherently controls the subsequent durability and damage tolerance of the affected structure. Transverse impact to composites is of particular concern due to the possibility of exciting damage modes that are difficult, or even not possible, to visually detect from the exterior (impact-side) surface. Some examples of such damage are delamination, backside-only fiber failure, debonding of internal substructure (e.g., stringers and stiffeners, doublers, joints), and crushing and separation of sandwich panel core. Impact damage is highly dependent upon the nature of the threat and conditions associated with the impact event. This chapter will provide an overview of impact damage threats that are common to composite aircraft structures and describe the relationship of these threat sources to the damage that is seeded. An historical overview of the impact damage tolerance methodologies developed by military aircraft programs, and subsequently widely adopted across the composite structures community, is also provided.

Wing stiffness center should be determined firstly for structure detail design. The present study focused on the impact analysis of beam and stringer for wing cross-section stiffness center based on thin wall structure mechanics theory. In order to discuss the impact of beam and stringer on stiffness center, the sensitivity formulas for stiffness center of wing cross-section were derived and expressed in terms of the dimension and layout of beams and stringers. The results indicated that the structural layouts of beam and stringer were important influencing factors in stiffness detail design of full composite wing structure. The research results can provide an important reference for the stiffness design and aeroelastic design of the full composite wing.

Composite structures require a rigorous program of nondestructive inspection and maintenance to detect and characterize hidden defects at an early stage of their occurrence so that preventive measures can be taken before the structure loses its load carrying capacity or suffers from catastrophic failure. Current methods for defects detection in large aircraft and aerospace structures are slow, labor intensive and costly. This is especially true for composite structures where conventional techniques are often ineffective. Ultrasonic guided waves offer an attractive complementary tool for improving inspection techniques in relatively large plate-like structural components due to their large propagation range and sensitivity to defects in their propagation path. Since the waves are affected by the geometrical structural features (e.g. stringers) as well as harmful defects (e.g. delaminations), the application of guided waves in the NDE or SHM of real structures requires a good understanding of these interaction effects. This will help identify the defects from their distinguishing features in the signal in structural components with complex geometry. In this paper a detailed study of the interaction of guided waves with defects in an aluminum plate and a honeycomb composite sandwich structure is carried out using numerical simulations and laboratory experiments. The simpler aluminum plate is used for model validation and understanding the basic characteristics of the interaction phenomena. The agreement between the simulated waveforms and those measured from the experiments are found to be excellent in both cases indicating the possibility of applying guided wave based techniques to more realistic structures.