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The purpose of this paper is to find the optimum configuration of the composite launch tube currently being developed in Roketsan. The winding thicknesses and winding angles of the launch tube are selected as design variables, and three different composite material alternatives are evaluated: glass/epoxy, carbon/epoxy and aramid/epoxy.

In this study, structural optimization of a composite launch tube of man portable air defense system is conducted. To achieve a cost-effective design, a cost scoring table that includes structural weight, material cost, availability and manufacturability is first introduced. Then, optimization for minimum weight is conducted, where the winding thicknesses and winding angle are taken as design variables, and the safety factor value obtained by using the Tsai–Wu damage criterion is used as constraint. A surrogate-based optimization approach is used where various options for surrogate models are evaluated. Glass/epoxy, carbon/epoxy and aramid/epoxy are considered as alternative materials for the launch tube. Finally, the selection of the most cost-effective design is performed to achieve optimum cost.

Carbon fiber-reinforced epoxy matrix material provides the optimum cost-effective design for the launch tube.

The findings of the paper can be used to design more cost-efficient composite launch tube currently being developed in Roketsan.

The existing studies are based on a design approach to achieve minimum weight of the launch tubes, whereas this study introduces a design approach to achieve optimum cost.

The purpose of this paper is to present microstructural and fractographic analysis of damage in aluminum (2024T3)/carbon-fiber reinforced laminates (AlC) after static tensile test. The influence of fiber orientation on the failure was studied and discussed.

The subject of examination was AlC. The fiber–metal laminates (FMLs) were manufactured by stacking alternating layers of 2024-T3 aluminum alloy (0.3 mm per sheets) and carbon/epoxy composites made of unidirectional prepreg tape HexPly system (Hexcel, USA) in [0], [± 45] and [0/90]S configuration. The fractographic analysis was carried out after static tensile test on the damage area of the specimens. The mechanical tests have been performed in accordance to ASTM D3039. The microstructural and fractographic analysis of FMLs were studied using optical (Nikon SMZ1500, Japan) and scanning electron microscope (Zeiss Ultra Plus, Germany).

FMLs based on aluminum and carbon/epoxy composite are characterized by high tensile properties depending on their individual components and the orientation of the reinforcing fibers, failure of hybrid laminates indicates the complexity process of degradation of these materials. The nature of damage in FML layers is similar to that typical in polymer composites with interlaminar delaminations, transverse cracks of the composite layers, degradation of fiber/matrix interface, damage process in FMLs is also associated mainly with interface between metal and fiber reinforced composite. The mixed damage – cohesive and adhesive – was observed.

One of the most important aspect in the designing and manufacturing process in the service life of composite structures is damage mechanisms. The damage processes in composite materials, particularly in FMLs, are more complex in comparison to metal materials and fiber reinforced polymers.

Composites reinforced by non-crimp fabrics (NCF) reach stiffness level close to the ideal cross-ply laminates with unidirectional plies, but damage behaviour of NCF-composites exhibit significant differences from the behaviour of the laminates with unidirectional plies (non-stitched laminates). The paper presents results of experimental studies of the initiation and development of damage in NCF-composites. Some of the observed phenomena can be explained and predicted using FE modelling of the composite deformation. Others are difficult to explain; they present a challenge for understanding the behaviour of NCF-composites. The materials are carbon/epoxy composites, reinforced by biaxial NCF 0°/90° and ±45°. Two different sets of fabrics (different producers) were used for the reinforcement, and the fibre volume fraction of the plates had two different levels: about 45% and about 55%. The samples were loaded in different directions in tension. Damage initiation and development was studied using acoustic emission and X-ray investigation.

This paper investigates the influence of moisture absorption on the mechanical properties of carbon/epoxy composites.

Three types of specimens are prepared, which are for longitudinal, transverse and shear tests. Specimens are immersed in distilled water at 70°C for 1, 3 and 9 months. These correspond to the moisture content of 2.2, 3.8 and 5.3%.

Compared to the values at dry condition, the longitudinal modulus, shear modulus and Poisson's ratio are invariant with the moisture content. However, the transverse modulus, transverse strength and shear strength are sensitive to moisture attack. The maximum degradation is 33%, 76 and 33% for the three properties, respectively. It is also worth to note that the longitudinal tensile strength is stable at 1 and 9 months of immersion. However, at 3-months ageing period, there is only 67% of the longitudinal tensile strength retained.

The experimental results are fitted with a residual property model. Results show comparatively good fit, with a difference within 16% except the longitudinal tensile strength at 9-months immersion. This highlights that the model is not suitable to fit the experimental data with a fluctuated trend.

The purpose of this paper is to investigate the feasibility of using rapid prototyping (RP) technologies (stereolithography (SLA), fused deposition modeling (FDM), and three‐dimensional printing (3DP)) for fabrication of the core of a composite sandwich structure.

Control cores of a flat geometry were fabricated from epoxy using SLA and from acrylonitrile butadiene styrene (ABS) plastic using FDM. Corrugated geometry cores were fabricated using SLA, FDM, and 3DP. Carbon‐epoxy composite sandwich structures were fabricated from all cores using a wet‐hand layup process with vacuum cure. The performance of each core was measured using a bend test to determine bending stiffness and failure load.

Based upon bending stiffness and failure load, composite sandwich structures utilizing epoxy cores fabricated via SLA outperformed composite sandwich structures utilizing plaster powder and ABS plastic cores. Composite sandwich structures with corrugated ABS plastic cores outperformed those with flat ABS plastic cores by a margin well beyond that predicted by theory in both bending stiffness and failure load.

The marked improvement in stiffness and failure load of the composite sandwich structures with corrugated ABS plastic cores over those with flat ABS cores is not explained by the theoretical improvement due to an increased area moment of inertia and increased surface area. Additional research in the failure mechanism is warranted.

The ability to easily create complex core geometries will allow for the ability to place enhanced structural features in the regions of high stress.

This paper demonstrates that cores fabricated via RP technology and containing enhanced structural features are suitable for carbon‐epoxy composite sandwich structures.

Delamination is a common and crucial damage mode which occurs during manufacturing of layered composites or their service life. Its existence leads to degradation in mechanical properties or even structural failure of composites. Hence, the purpose of this article is to study the effect of induced delamination on flexural performance of CFRP composites.

In this article, the flexural behaviors of intact and delaminated carbon/epoxy laminates were investigated under pure bending. A circular PTFE film was introduced during fabrication to create artificial delamination. Moreover, finite element models were developed for intact and delaminated composites using ANSYS. The created models were discretized using 3D structural eight node solid elements.

The delamination influenced considerably flexural properties of composite. The composite exhibited a linear elastic nature prior to the damage of top ply on the compression side. The flexural strength and stiffness of the composite reduced to 44.5% and 18.2% respectively due to the existence of artificial delamination. The results of four point bending experiments and finite element analysis agreed for both intact and delaminated composites within acceptable error. Finally for same composites, first ply failure analysis was carried out using Tsai-Hill, Tsai-Wu and Hashin failure criteria.

In pure bending, beam section of the middle portion is free from shear. It is not so in case of three-point bending. Hence, the effect of embedded artificial defect on bending performance of CFRP composite due to pure bending has been investigated.

The purpose of this study is to investigate the effect of dimple size on the tribological performances of laser surface-textured palm kernel-activated carbon-epoxy (PKAC-E) composite.

A PKAC-E disc 74 mm in diameter was fabricated using the hot compression moulding technique. Five different types of surface contacts were prepared using a CO2 laser surface-texturing machine: a non-textured surface, and surfaces with dimples between 500 and 1,200 μm in diameter. The area density, contact ratio and depth were kept constant. A sliding test was carried out using a ball-on-disc tribometer under boundary lubricated conditions with constant sliding speed, sliding distance and applied load.

In general, the results showed that the friction coefficient decreased with an increasing dimple diameter of surface-textured PKAC-E composite. However, the appropriate dimple diameter for maintaining low friction coefficient is proposed in the range of 800 to 1,000 μm.

This is the first study, to the authors’ knowledge, to investigate the effects of dimple size, which is larger than 500 μm, on the tribological performances of laser surface-textured PKAC-E composite.

The indentation behaviour of sandwich panels is significant to incipient damage and is known to be affected by a number of dominant parameters. However, it is challenging not only to demonstrate how those few dominant parameters influence the indentation behaviour but also to ascertain that such influence was coupled to the variation of the other dominant parameters. The paper aims to discuss these issues.

In this work, the authors adopted a controllable quasi-static testing to carry out a diagnostic interrogation on the nature of incipient damage in laminate-skinned sandwich panels using hemispherical indenter and used photographs taken from the cross-sections of all the cut-up tested specimens, which were stopped both just before and after the initial critical loads, respectively, to confirm the mechanism of the incipient damage. Sandwich panels with aluminium honeycomb core had carbon/epoxy skins of two different thicknesses and lay-ups and hemispherical nosed indenter had three different diameters.

The authors found that: the incipient damage mechanism in all the panels was combined delamination in the skin and core crushing without debonding; doubling the skin thickness had the significant enhancement on critical load and indentation and this enhancement became greater for the larger indenter diameters; the indenter diameter had the moderate effect on critical load in the thick panels from 8 to 14 mm but had the negligible effect on thin panels and no effect on the thick panels from 14 to 20 mm; varying the skin lay-up or support had little effect on the indentation behaviour.

These findings were limited to the constant core density and core thickness. Varying the former significantly could alter the findings accordingly.

The results of this work should be tremendously useful to design and analysis in industrial applications of sandwich structures in aircraft, vehicles, marine vessels and transport carriages for situations involving localised loading and deformation.

The results of this research work is one of the very few that demonstrated a systematic understanding of the indentation behaviour characteristics of sandwich construction, which is vital to the establishment of indentation law for sandwich structures in future.

When a thick structure is, on the contrary, subjected to moisture absorption, a fairly long time may be needed to reach full saturation. It is, therefore, important to understand and predict the areas of complex composite structures that are more prone to saturation. The material knock-down factors (proportional to the moisture content) may be applied only to these zones, in order to obtain a less pessimistic structural response prediction. The purpose of this paper is to investigate an FE diffusion model that was used to validate the absorption testing results of thick carbon epoxy laminates.

The experimental results were validated by using a diffusion model in Abaqus FE code.

The absorption results of three 15 mm thick carbon epoxy laminates are presented and reproduced via a mass diffusion model. The laminates were conditioned at 70°C and 85 per cent relative humidity in a moisture chamber. Areas more prone to saturation have been predicted by the FE model and the moisture content in the non-saturated areas has been calculated.

The practical implications of the absorption model are discussed on an example of an aero-engine fan blade-like structure.

Validation of thick panels’ absorption data is an important point of novelty of this paper, given the lack of experimental and modelling validation in the open literature.

The purpose of this paper is to present the process of design and prototyping of a two-seat, electric-powered, self-launching motorglider AOS-71 closely connected with the teaching process conducted by the academic staff of Warsaw University of Technology (WUT) within a unique educational ULS – Ultra Light Sailplanes programme.

The selected design methods and tools used during the development of the motorglider have been described. The computer aided design/computer aided manufacturing modules of the Siemens NX software were used to work on the structural design, tools and technical documentation. The core of the ULS educational programme is to educate aerospace engineering students by providing an opportunity for them to participate in each phase of the aircraft life cycle – from conceptual drawings through structural design and prototyping to manufacturing, testing and maintenance.

The main innovations of the AOS-71 design are: retractable ecological electric propulsion, spacious cockpit where seats are located side by side and the all-composite airframe made of 90 per cent advanced carbon epoxy composites.

The electric motorglider can be used as a multifunctional flying laboratory for flight research and student education.

The AOS-71 project and its continuation are a valuable example of involving aerospace students in each phase of the aircraft life cycle. It also contributes to the research in the field of using innovative electrical propulsion systems in aircraft designs.