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Impact and fatigue are critical loading conditions for composite aerostructures. Compression behavior after impact and fatigue is a weak point for composite fuselage panels. The purpose of this paper is to characterize experimentally the compression behavior of carbon fiber reinforced plastic (CFRP) stiffened fuselage panels after impact and fatigue.

In total, three panels were manufactured and tested. The first panel was tested quasi-statically to measure the reference compression behavior. The second panel was subjected to impact so as to create barely visible impact damage (BVID) at different locations, then to fatigue and finally to quasi-static compression. Finally, the third panel was subjected to impact so as to create visible impact damage (VID) at different locations and then to quasi-static compression. The panels were tested using ultrasound inspection just after manufacturing to check material quality and between different tests to detect impact and fatigue damage accumulation. During tests the mechanical behavior of the panel was monitored using an optical displacement measurement system.

Experimental results show that the presence of impact damage significantly degrades the compression behavior of the panels. Moreover, the combined effect of BVID and fatigue was proven more severe than VID.

The paper gives information about the compression after impact and fatigue behavior of CFRP fuselage stiffened panels, which represent the most realistic loading scenario of composite aerostructures, and describes an integrated experimental procedure for obtaining such information.

Polymer composite panels are widely used in aeronautic and aerospace structures due to the high strength-to-weight ratios of these structures. The purpose of this paper is to determine the strain fields and failure mechanisms during the failure of the impacted composite laminates when subjected to compression.

A series of compression-after-impact (CAI) tests was performed on composite plates 150×100×4 mm3 made of a carbon-fibre-reinforced epoxy resin matrix. A digital image correlation and fractographic analysis by means of optical and electron microscopy are used for this purpose.

The full-field strain measurements indicate a concentrated band of compressive strain near the impact, where buckling occurs. The results indicate that the strain concentration factor can be considered to be a failure criterion. The shear strain visualisation around the impact reveals an area of heterogeneous deformation that is comparable to the detected delamination area acquired by an ultrasonic technique. Fibre and inter-fibre fractures are described for the particular impact site regions.

These experiments could improve numerical models for the CAI analyses and help to build a new criterion for this severe failure mode.

Engineering composite laminates/structures are usually subjected to complex and variable loads, which result in interlayer delamination damage. However, damaged laminate may cause the whole structure to fail before reaching the design level. Therefore, the purpose of this paper is to develop an equivalent model to effectively evaluate compressive residual strength.

In this paper, taking carbon fiber reinforced composite T300/69 specimens as the study object, first, the compressive residual strength under different impact energy is obtained. Then, zero-thickness cohesive elements, Hashin failure criteria and Camanho nonlinear degradation scheme are used to simulate the full-process simulation for compression after edge impact (CAEI). Lastly, based on an improved Whitney–Nuismer criterion, the equation of edge hole stress distribution, characteristic length and compressive residual strength is used to verify the correctness of the equivalent model.

An equivalent relationship between the compressive residual strength of damaged laminates and laminates with edge hole is established. For T300/69 laminates with a thickness of 2.4 mm, the compressive residual strength after damage under an impact energy of 3 J is equivalent to that when the hole aperture R = 2.25 mm and the hole aperture R = 9.18 mm when impact energy is 6 J. Besides, the relationship under the same size and different thickness is obtained.

The value of this study is to provide a reference for the equivalent behavior of damaged laminates. An equivalent model proposed in this paper will contribute to the research of compressive residual strength and provide a theoretical basis for practical engineering application.

The purpose of this paper is to synthetically investigate the impact damage responses of carbon fiber reinforced polymer (CFRP) and its influence on the compression mechanical responses of CFRP laminates, including damage distribution, residual compressive strength and fracture morphology.

A progressive damage simulation model is developed to analyze the complicated damage responses of CFRP laminates that are manufactured by resin transfer method (RTM) technology. Based on the ABAQUS/explicit finite element analysis solver, a VUMAT code is proposed to descript the composite materials’ damage behaviors under both impact and compression load. Adopting this proposed model, the primary mechanical indicators of four groups’ 5284RTM/U3160 CFRP laminates with different stacking sequences are predicted. Moreover, impact and compression after impact tests are conducted to verify the accuracy of simulation results.

Both simulation and experimental results show that the impact damage with low visible detectability can significantly reduce composites’ compressive strength. For all four groups’ composite laminates, the residual strength ratio is around 35% or even lower. The kernel impact damage near the plates’ geometric center promotes the degradation process of local materials and finally leads to the early occurrence of mechanical fracture. In addition, the impact damage projection area is not sensitive to the parameters of stacking sequences, while the residual compression strength is proportional to the number of 0-degree layers within whole laminates.

This study helps to understand the effect of an impact event on CFRP laminates’ compressive bearing capacity and provides a numerical method in simulating the damage responses under both impact and compression load.

In this paper, an attempt is made to obtain buckling loads, ultimate bearing capacity and other required structural characteristics of grid structure panels. The numerical method for post-buckling behavior analysis of panels involving multiple invisible damages is also presented.

In this paper, two bidirectional stiffened composite panels are manufactured and tested. Multiple discrete invisible damages are introduced in different positions of the stringers, and the experimental and simulation investigation of buckling and post-buckling were carried out on the damaged stiffened panels.

The simulation load–displacement curves are compared with the experimental results, and it is found that the simulation model can well predict the occurrence of buckling and failure loads. The strain curve shows that the rate of strain change at the damaged site is greater than that at the undamaged site, which reflects that the debond is more likely occurred at the damaged site. The simulation verifies that the panel is usually crushed due to matrix compression and fiber–matrix shear.

In this paper, post-buckling tests and numerical simulations of bidirectional stiffened composite panels with impact damage were carried out. Two panels with four longitudinal stringers and two transverse stringers were manufactured and tested. The buckling and post-buckling characteristics of the grid structure are obtained, and the failure mechanism of the structure is explained. This is helpful for the design of wall panel structure.

Compression is critical loading condition for composite airframes. Compression behaviour of structures with or without damages is a weak point for composite fuselage panels. This is one of the reasons for need of continuous in-service health monitoring of composite structures. The purpose of this paper is to characterize the compression panel behaviour on the base of a developed and implemented structural health monitoring (SHM) system.

The SHM system based on fibre optic Bragg grating (FOBG) sensors and standard resistance strain gauges (SGs) was placed onto/into (embedded or bonded) three stiffened carbon fibre reinforced polymer (CFRP) fuselage panels. The FOBG sensor system was used to monitor the structural integrity of the reference, impacted, and fatigued panels under compression loading. Both barely visible impact damage and visible impact damage were created to evaluate their influence on the panel behaviour. The functionality of the SHM system was verified through mechanical testing.

Experimental data showed the presence of impact damages significantly changes the buckling modes development and deformation behaviour of the panels. Some differences between the optical and SG sensors during buckling were observed. The buckling waves and failure development were very well indicated during loading by all sensors located on the panel surface but not by the embedded sensors. Good agreement between the data from the SGs and FOBG sensors was achieved for all sensors placed on the stringers, which did not buckle. The good reliability of FOBG sensors during the fatigue and static testing up to panel failure was verified.

The paper gives information about different buckling behaviour of CFRP fuselage stiffened panels in compression. The paper gives detailed information about measured signals from different sensors based on their location on/in the panel structure for realistic loading scenario of composite aerostructures. The paper gives an integrated overview of sensors placement considering possibilities to predicate structure behaviour.

This study aimed to determine the peak impact force and force attenuation capacity of weft-knitted spacer fabrics intended for padding that can be used for human body protection against impact.

A total of five weft-knitted spacer fabrics were fabricated with four different diameters of nylon monofilament yarns and one doubled monofilament yarns, respectively. The impact performances of the weft-knitted spacer fabrics were tested using a drop test method with a customized test rig to simulate falling. Impact tests were conducted on single- and multilayered experimental spacer fabrics to investigate the peak impact force and force attenuation capacity.

It was found that weft-knitted spacer fabric with a coarser or larger diameter of monofilament spacer yarn generated lower impact force and higher force attenuation capacity, thus resulting in better impact performance. Greater force attenuation can be achieved by utilizing a higher number of spacer fabric layers. However, the increase in thickness must be considered with the spacer fabric end use.

This study employed relatively coarse nylon monofilament yarn as spacer yarns to gain knowledge on the impact performance of weft-knitted spacer fabrics compared to warp-knitted spacer fabrics which are more common. The results showed that the diameter of spacer yarn significantly influenced the impact performance of the experimental weft-knitted spacer fabrics. These results could be useful for designing and engineering textile-based impact protectors.

Gives a bibliographical review of the finite element methods (FEMs) applied for the linear and nonlinear, static and dynamic analyses of basic structural elements from the theoretical as well as practical points of view. The range of applications of FEMs in this area is wide and cannot be presented in a single paper; therefore aims to give the reader an encyclopaedic view on the subject. The bibliography at the end of the paper contains 2,025 references to papers, conference proceedings and theses/dissertations dealing with the analysis of beams, columns, rods, bars, cables, discs, blades, shafts, membranes, plates and shells that were published in 1992‐1995.

Fibre metal laminates were developed at Delft University during the last two decades as a family of new hybrid materials consisting of bonded thin metal sheets and fibre/adhesive layers. This laminated structure provides the material with excellent fatigue, impact and damage tolerance characteristics and a low density. While the 20 per cent weight reduction was the prime driver behind the development of this new family of materials, it turns out that additional benefits like cost reduction and an improved safety level have become more and more important. The combination of these aspects in one material makes fibre metal laminates a strong candidate material for fuselage skin structures of the new generation of high capacity aircraft. The focus on this application currently leads to industrialization and qualification that makes this material available to the aircraft designer.

Stab-resistant body armor (SRBA) is used to protect the body from sharp knives. However, most SRBA materials currently have the disadvantages of large weight and thickness. This paper aims to prepare lightweight and high-performance SRBA by 3D printing truss structure and resin-filling method.

The stab resistance truss structure was prepared by the fused deposition modeling method, and the composite structure was formed after filling with resin for dynamic and quasi-static stab tests. The optimized structural plate can meet the standard GA68-2019. Digital image correlation technology was used to analyze the local strain changes during puncture. The puncture failure mode was summarized by the final failure morphologies. The explicit dynamics module in ANSYS Workbench was used to analyze the design of the overlapped structure stab resistance process in this paper.

The stab resistance performance of the 3D-printed structural plate is affected by the internal filling pattern. The stab resistance performance of 3D-printed structural parts was significantly improved after resin filling. The 50%-diamond-PLA-epoxy, with a thickness of only 5 mm was able to meet the stab resistance standard. Resins are used to increase the strength and hardness of the material but also to increase crack propagation and reduce the toughness of the material. The overlapping semicircular structure was inspired by the exoskeleton structure of the demon iron beetle, which improved the stab resistance between gaps. The truss structure can effectively disperse stress for toughening. The filled resin was reinforced by absorbing impact energy.

The 3D-printed resin-filled truss structure can be used to prepare high-performance stab resistance structural plates, which balance the toughness and strength of the overall structure and ultimately reduce the thickness and weight of the SRBA.