This paper investigated the edge-on low-velocity impact (ELVI) behavior of unenhanced and Z-pin-enhanced carbon fiber plain-woven composite laminates under impact energies of 10 J, 30 J and 50 J through a combined experimental and finite element analysis (FEA) approach. Multi-scale characterization techniques, including visual inspection, non-destructive techniques, optical microscopy observation and FEA, were employed to systematically examine the post-impact damage states and reconstruct their three-dimensional (3D) spatial distribution characteristics. Firstly, an intralaminar continuous damage mechanics model and an interlaminar cohesive zone model applicable to plain-woven composite were established, incorporating a matrix failure criterion that was independent of empirical parameters. The 3D damage morphology and failure mechanisms of unenhanced and Z-pin-enhanced specimens during the ELVI process were interpreted based on combined experimental observations and numerical simulation. On this basis, the toughening mechanism of Z-pins was further interpreted with the assistance of FEA in combination with experimental observations. The reliability of the proposed model was assessed by comparing the mechanical response and damage characteristics between experimental and numerical results and a reasonable agreement was achieved. The research indicated that Z-pins absorb a significant amount of energy through bending and fracture under ELVI loading, thereby playing a crucial role in suppressing delamination of the laminates. After an impact energy of 30 J, compared with the unenhanced specimens, the delamination area of the Z-pin-enhanced specimens was reduced by 77.9%, and the maximum delamination crack length along the impact direction was decreased by 58.3%.

To investigate the effects of discrete source damage induced by bird strikes or engine debris impact on the load carrying capacity of structures, composite hat-stiffener panels with notch defect and debonding defect were designed. Experimental testing and numerical simulation were combined to systematically examine the mechanical behaviour of these panels under uniaxial compression. A continuum damage mechanics model, incorporating a failure initiation criterion that does not rely on empirical parameters, was developed as part of the numerical method. To quantitatively evaluate the influence of these defects on structural stability and load carrying capacity, buckling strain and failure strain were introduced as evaluation parameters, and the performance of specimens with defects was compared with that of the undamaged specimen. Results indicated that the presence of a single-bay notch increased the buckling strain by 6.9% but reduced the failure strain by 27.1%. In contrast, the double-bay notch and stiffener debonding led to significant reductions in buckling and failure strains. Consequently, all defect types significantly reduced the post-buckling load carrying capacity. In addition, experimental results revealed that for specimens with stiffener debonding, increasing the skin thickness effectively delayed the onset of initial buckling and might cause buckling to initiate in the stiffeners rather than the skin.

A theoretical model for predicting kink-band formation in unidirectional composites is developed. In contrast to previous models, the proposed model takes into account the shear deformation in the non-misaligned regions when determining the fiber misalignment angle, and introduces an analytical approach for determining the orientation of the kink-band plane based on the applied stress state, instead of the conventional approach based on maximizing a failure index. Based on this model, a three-dimensional failure criterion for composites is further developed, which considers the effect of fiber misalignment under all stress states and therefore overcomes the discontinuity in the failure envelope produced by Pinho's criterion. Using the proposed model and criterion, the failure behavior of unidirectional composites under various stress states is analyzed. A notable finding is that, under the stress state, both longitudinal tensile stress and low longitudinal compressive stress can enhance the longitudinal shear stress at failure. The proposed criterion is validated against six sets of experimental data, with prediction errors of 1.77-11% and the lowest error in four of the six cases.

PurposeTo support civil aircraft airworthiness validation, this paper aims to establish a standardized, efficient and precise technical framework for structural strength analysis and test verification, addressing the complexity of aircraft structural mechanics.Design/methodology/approachA three-level finite element analysis (FEA) methodology (Natural Grid, Refined and Detail finite element model (FEM)) is developed, integrated with parametric modeling, automated strength analysis and full-scale static test verification following the building block approach.FindingsThe three-level FEA system accurately simulates global load paths, regional stress distributions and local structural behaviors; its predictions (e.g. 103% of target failure load) are closely matched with experimental results, which ensures structural compliance with airworthiness requirements.Originality/valueThis study proposes a standardized three-level FEA framework for civil aircraft, advancing digital transformation in structural development, enhancing simulation precision and providing robust technical support for optimizing aircraft structural design and aviation safety.

Compared to investigations on the mechanical properties of composites at room temperature, this paper explored the effects of T-joint specimen, prepared with the room temperature curing paste adhesive J-349-3, on the pullout and bending properties after exposure to both hygrothermal aging (HA) and random vibration (RV). The quasi-static pull-out and bending tests were carried out on the baseline specimen, the HA specimen, the RV specimen and the HA/RV specimen. The damage morphology, ultimate load and stiffness of the specimens were compared and analyzed, and scanning electron microscope (SEM) was used to reveal the influence mechanisms of HA and RV on the specimens. For the pull-out specimens, the ultimate load of the HA specimens increased by 14.4 %, while the ultimate load of the RV and HA/RV specimens decreased by 6.0 % and 19.9 %, respectively, compared to the baseline specimens. For the bending specimen, the environmental factors had a smaller effect (<10 %) on the ultimate load when the specimen was subjected to environmental factors. Compared to the performance of the composite laminate, the performance of the adhesive interface was more significantly influenced by the application of HA and RV. Additionally, there was a synergistic effect mechanism between the two combinations of HA treatment and RV on the effect of pull-out properties. Finally, the damage initiation and evolution of the specimen were further analyzed by the finite element model, and the validity of the model was verified by comparing the test results.

High-precision numerical simulations and physical experiments for composite structure design are costly and limited. Neural networks offer a promising solution, but conventional models rely heavily on large amount of high-fidelity data and struggle to handle the complex variable space introduced by layup sequences. This paper proposes a deep learning model-Multi-Fidelity Triple LSTM (MF-T-LSTM)-to predict composite mechanical properties using limited test samples. The model uses one LSTM to capture layup sequences and two others to process data at different fidelity levels. The MF-T-LSTM model is applied to predict the tensile load-displacement curve of composite open-hole laminates. A combination of limited experiments and extensive FEM simulations generated the training dataset. The model's components (LSTM-MLP and Dual LSTM) were validated before development. The MF-T-LSTM achieved high prediction accuracy, with an average R2 of 0.918 on the experimental test set. Quantitative analysis demonstrated its low dependence on low-fidelity (LF) data quality and strong transferability across different LF datasets. Replacing LF dataset with less accurate versions led to only a 1.4 % average drop in performance. These results underscore the MF-T-LSTM model's advantages in predicting composite mechanical responses and reducing experimental costs.

Despite the much more notorious damage caused by edge-on low-velocity impact (ELVI) when compared to outof-plane low-velocity impact (OPLVI), the research on compression after edge-on impact (CAEI), especially the techniques to improve the ELVI properties, is rather limited. This article innovatively explored the effect of macro Z-pins, CNFs (carbon nanofibers), and multi-scale CNFs/Z-pins synergistic reinforcement on the CAEI performance of carbon fibre-epoxy composite laminates after impacts at 5 J, 10 J, 30 J, and 50 J. The variation of the composite laminate strain distribution during the CAEI process was accurately visualized using 3D-Digital Image Correlation (3D-DIC). After a 50 J impact, the CAEI strength of the CNFs/Z-pins, Z-pins, and CNFsreinforced samples increased by 40.2 %, 26.2 %, and 5.6 %, respectively, compared to the unreinforced samples. Additionally, optical micrographs and scanning electron microscopy (SEM) were utilized to reveal the damage morphology, as well as to explore the synergistic toughening mechanisms of CNFs and Z-pins.

The failure mechanisms of carbon-fiber-reinforced composites are complex, and experimental tests are costly. Traditional finite element methods, limited by current theoretical models, struggle to accurately simulate the entire failure process and exhibit significant cumulative errors, complicating precise modeling and uncertainty quantification. Machine learning approaches offer a promising alternative but generally require extensive datasets to achieve satisfactory performance. We present a multi-fidelity data-driven framework that blends a small set of high-fidelity test results with an extensive collection of low-fidelity simulation data to predict the distribution of mechanical properties in composite structures. The framework is validated through tensile-failure experiments on notched laminates. To improve the statistical representativeness of the limited experimental samples, we introduce a Bayesian data-augmentation scheme and derive the theoretical distribution of the inter-group coefficient of variation to confirm its soundness. Cross-validation shows that the proposed method attains a mean absolute error of 4.99% when predicting the 10th percentile of the failure-load distribution. The study mitigates the twin challenges of scarce experimental data and weak coupling between numerical models and physical tests.

Compared with the common numerical analysis of out-of-plane low-velocity impact (LVI) of composite laminates, this paper established a finite element model (FEM) for out-of-plane LVI of unreinforced laminates and Z-pin-reinforced laminates. The mechanical response and damage morphology of the model after impact at energies of 5, 10, 30, and 50 J were verified with experiments. Through mechanical response research, it was found that the numerical analysis models of LVI of unreinforced and Z-pin-reinforced laminates have higher prediction accuracy for maximum impact force, maximum displacement, and energy absorption of the laminate. The simulation results showed that the Z-pin reinforced laminates had higher impact force and lower maximum displacement compared to the unreinforced laminates. From the LVI simulation, it was observed that the damage morphology of the laminate's surface and interior indicated minor fiber damage, which correlated well with the fiber damage observed in the experiments. In contrast, matrix damage played a dominant role during the impact, leading to extensive damage. This extensive damage occurred due to the loss of the stress transfer medium caused by matrix damage, which subsequently resulted in fiber damage.Highlights Comparison of simulated and experimental mechanical response Criteria for determining the initiation of matrix/fiber damage Simulation analysis of Z-pin reinforced laminates Intra-laminar fiber/matrix damage and inter-laminar delamination

The composite thin-walled stiffened panel exhibits considerable potential for post-buckling bearing capacity. However, due to the limitations in post-buckling analysis and experimental evaluation capabilities, local buckling under limited loads is not permissible for current aircraft composite panels, significantly diminishing the weight reduction effect of composite panels. In this paper, experimental research on composite fuselage curved panels under combined compression and shear loads was conducted with an experimental system. The buckling analysis and failure prediction of composite fuselage curved panels were performed by the finite element method, using the interlaminar failure criterion and an improved Tsai-Wu failure criterion. Meanwhile, interaction formulaes were utilized to predict buckling correlation curves and failure envelopes. The results show that the finite element model with improved criteria and the rapid analysis method based on interaction formulaes exhibit high accuracy, which are available for properties analysis of composite panels. The experimental techniques and analytical methods proposed in this paper can provide a way to evaluate the buckling and post-buckling performance of composite stiffened panels.

A new longitudinal failure criterion for unidirectional fibre reinforced (UD) composites is proposed, incorporating the contribution of local fibre misalignment to fibre failure. The failure function was established based on the quadratic polynomial expansion of stress components in the local coordinate system parallel to the misaligned fibres, with coefficients rationally determined based on the basic strength parameters of UD composites. The local stress components were used to determine the failure status associated with fibre compression, while the tensile load along fibres was assumed to have negligible influence on the development of the fibre misalignment. The predicted failure envelopes agreed well with the experimental data for different stress combinations, particularly in the regimes where the fibre failure was influenced by transverse stresses. This validates the predictive ability of the proposed criteria and highlights the importance of the contribution of matrix-failure-related stresses to the fibre-dominated failure behaviour of fibre reinforced composites.

Based on Hashin's fracture plane assumption, a matrix failure criterion for brittle fiber-reinforced composites is proposed. The failure function is expressed as a quadratic polynomial of the stress components on the fracture plane. The unknown coefficients in the failure criterion are only calibrated by the three basic strengths of unidirectional composites, i.e., the transverse tensile strength, transverse compressive strength, and longitudinal shear strength, thus overcoming the limitation of requiring empirical parameters in most previous matrix failure criteria. Especially, under plane stress states (622, 1-21), an analytical solution for the fracture angle of unidirectional composites can be provided. The prediction results of the proposed criterion are consistent with a large number of experimental data, confirming its applicability. In addition, the study establishes the relationship among the three transverse basic strengths (Yt, Yc and S23). It can be used to predict S23 which is difficult to measure experimentally.

This paper investigated the effects of carbon nanofibers (CNF), Z-pin, and multi-scale CNF/Z-pin on the compression-after-impact (CAI) properties and damage morphology of laminates after 10 J, 30 J, and 50 J impact. The deformation and strain field variations of the laminate during the CAI process were clearly characterized by 3D-Digital Image Correlation (DIC). The process from local buckling to the global failure of the laminate was clearly revealed. The mechanical response results showed that the CAI strength of unreinforced, CNF, Z-pin and CNF/Z-pin-reinforced laminates increased in order after impact with the same energy. After 50 J impact, the CAI strength of CNF, Z-pin and CNF/Z-pin-reinforced laminates was increased by 18.6%, 28.7% and 36.4%, respectively, compared to unreinforced laminates. By constructing a simulation model to analyze damage and mechanical response, the model's effectiveness was verified. Additionally, the damage morphology, damage mechanism, and synergistic toughening mechanism between CNF and Z-pin in CAI specimens were revealed from optical photographs and scanning electron microscopy (SEM).

This paper investigated the effect of CNF (carbon nanofibre)/Z-pin multi-scale synergistic reinforcement on improving the impact damage resistance of carbon fiber-reinforced composite laminates. The out-of-plane low-velocity impact response of unreinforced, CNF -reinforced, Z-pin-reinforced and CNF/Z-pin multi-scale synergistic reinforced laminates were investigated. The use of CNF concurrently with Z-pin can not only eliminate the incapability of Z-pin in improving the damage resistance at low energy (damage initiation) but can also synergistically maximize the improvement of the impact resistance at high energy (damage propagation). At 50 J impact energy, the damaged area of CNF, Z-pin and CNF/Z-pin reinforced laminates were reduced by 10.3%, 29.5% and 34.8%.

Currently,experimental research on variable stiffness design mainly focuses on lami-nates.To ensure adaptability in practical application,it is imperative to conduct a systematic study on stiffened variable stiffness structures,including design,manufacture,experiment,and simula-tion.Based on the minimum curvature radius and process schemes,two types of T-stiffened panels were designed and manufactured.Uniaxial compression tests have been carried out and the results indicate that the buckling load of variable stiffness specimens is increased by 26.0％,while the fail-ure load is decreased by 19.6％.The influence mechanism of variable stiffness design on the buck-ling and failure behavior of T-stiffened panels was explicated by numerical analysis.The primary reason for the reduced strength is the significantly increased load bearing ratio of stiffeners.As experimental investigations of stiffened variable stiffness structures are very rare,this study can be considered a reference for future work.

This paper proposes a three-dimensional failure criterion model for unidirectional fiber-reinforced composites. In contrast to previous models that only accounted for the effect of fiber misalignment in the presence of longitudinal compressive stress, the proposed failure criterion model comprehensively considers the effect of localized misaligned regions on the failure under any stress state. Another key contribution of this study is the introduction of the effective misalignment angle. Considering effective misalignment angles, the proposed failure criterion model can reasonably reveal the effect of localized misaligned regions on the failure behavior under different stress states. The agreement between the predicted results and the experimental data proves that the proposed model has good applicability. Furthermore, the influence of initial misalignment angles on failure is analyzed under varying stress conditions. The results indicate that even under longitudinal tensile stress, the initial misalignment angle still plays an important role in the failure behavior of materials.

A proposed strategy for managing airspace and preventing illegal drones from compromising security involves the use of autonomous drones equipped with three key functionalities. Firstly, the implementation of YOLO-v5 technology allows for the identification of illegal drones and the establishment of a visual-servo system to determine their relative position to the autonomous drone. Secondly, an extended Kalman filter algorithm predicts the flight trajectory of illegal drones, enabling the autonomous drone to compensate in advance and significantly enhance the capture success rate. Lastly, to ensure system robustness and suppress interference from illegal drones, an adaptive fast nonsingular terminal sliding mode technique is employed. This technique achieves finite time convergence of the system state and utilizes delay estimation technology for the real-time compensation of unknown disturbances. The stability of the closed-loop system is confirmed through Lyapunov theory, and a model-based hardware-in-the-loop simulation strategy is adopted to streamline system development and improve efficiency. Experimental results demonstrate that the designed autonomous drone accurately predicts the trajectory of illegal drones, effectively captures them using a robotic arm, and maintains stable flight throughout the process.

Composite structures can achieve greater performance through variable stiffness design. In consideration of inevitable manufacturing defects introduced by the Automatic Fiber Placement machine, modified models are established to calculate equivalent properties for materials with gaps or overlaps. A modeling method for variable stiffness structures with defects is proposed based on the modified models. This approach significantly reduces the dependence on mesh size for analysis accuracy, thereby improving modeling and calculation efficiency. Upon validation of the proposed modeling method, a Hierarchical Kriging surrogate modeling is employed to optimize the buckling performance of a variable stiffness wing box with defects. The impact of different manufacturing strategies on optimization results is also investigated. The findings demonstrate that the variable stiffness design improves the wing box buckling performance under a combined torsion-bending condition. Finally, the potential of utilizing overlaps without cut-restart is analyzed for weight reduction in the design of variable stiffness wing boxes.

This paper proposes a cascaded dual closed-loop control strategy that incorporates time delay estimation and sliding mode control (SMC) to address the issue of uncertain disturbances in logistic unmanned aerial vehicles (UAVs) caused by ground effects, crosswind disturbances, and payloads. The control strategy comprises a position loop and an attitude loop. The position loop, which functions as the outer loop, employs a proportional–integral–derivative (PID) sliding mode surface to eliminate steady-state error through an integral component. Conversely, the attitude loop, serving as the inner loop, utilizes a fast nonsingular terminal sliding mode approach to achieve finite-time convergence and ensure a quick system response. The time-delay estimation technique is employed for the online estimation and real-time compensation of unknown disturbances, while SMC is used to enhance the robustness of the control system. The combination of time-delay estimation and SMC offers complementary advantages. The stability of the system is proven using Lyapunov theory. Hardware-in-the-loop simulation and flight tests demonstrate that the control law can achieve a smooth and continuous output. The proposed control strategy can be effectively applied in complex scenarios, such as hovering, crash recovery, and high maneuverability flying, with significant practicality in engineering applications.