The deformation and stress have vital influences on the geometric accuracy and mechanical property of assembly structure that determine the quality of aircraft product. In order to realize real-time perceptions and controls of deformation and stress in aircraft assembly, a hybrid modelling strategy is proposed to determine the static response of the assembly structure composed of parts with high and low rigidities. The parts are meshed and separately applied by linear and nonlinear analysis models to represent explicit mappings from nodal force to nodal displacement and stress. The linear model of high-rigid part with invariant stiffness is established by the method of influence coefficient, while the nonlinear static response of low-rigid part with variable stiffness is represented by the Kriging model that is trained by the dataset of displacement and stress under multiple groups of loads. Combining individual model of each part by the application of coupled degree of displacement freedom to the nodes on connection surface, a hybrid analysis model of the whole assembly structure is established. Attributed to the decoupling of linear and nonlinear terms in the Kriging model, all the parameters in the hybrid model can be identified offline, and the mappings from force to displacement and stress of the assembly structure are rapid for real-time computation. The indicators evaluating the stress concentration and displacement magnitude are established, and a multi-objective optimization model of holding forces is further established and solved for the regulations of deformation and stress on the assembly structure. A software and a flexible fixture with servo system are developed to monitor and control the assembly stress and deformation. Simulation results show that the hybrid model can represent the static response of the assembly structure precisely and the optimization of holding force can suppress the stress and deformation efficiently. The maximum analysis errors of total displacement and Von-Mises stress on the assembly structure are separately 0.032 mm and 8.2 MPa corresponding to their real data of 0.45 mm and 186 MPa, while the displacement and stress are controlled below 0.009 mm and 1 MPa after the optimization of holding force. Experimental verification demonstrates that the assembly deformation and stress can be instantly perceived and regulated online by the sensing and controlling of holding forces.

In this study, two types of rivets, solid and blind rivets, are employed to repair delaminated composite components with varying curvatures. The forced assembly behavior of these riveted joints under drilling depth deviations and angular misalignments is systematically investigated through a combination of finite element analysis and experimental testing. The initiation and evolution of damage within the composite joints are also characterized. A three-dimensional elastoplastic damage constitutive model is developed, accounting for anisotropic material behavior, nonlinear response, and progressive damage. Based on this model, plastic deformation, multiphase damage, and residual stresses in the riveted composite joints are predicted and validated experimentally. The results demonstrate that the numerical model accurately captures the forced assembly process and microscale damage evolution around the rivet hole. In solid-riveted joints, plastic deformation propagates in a V-shaped pattern along the rivet axis, whereas in blind-riveted joints, it forms a “pine tree” distribution around the hole. The drilling depth deviation enhances the axial constraint stiffness of the rivet fasteners but also exacerbates localized damage in the countersink region. Angular misalignment produces asymmetric stress distribution across the joint, promoting unidirectional delamination propagation.

Carbon fiber reinforced polymer (CFRP) structures require efficient and reliable repair after sustaining local damage. This paper proposes an electrothermal co-curing bonding stepped repair method for CFRP by embedding a carbon nanotube film (CNTF) into the adhesive layer, endowing the repair with both Joule heating and self-sensing functionalities. The electrothermal and sensing properties of the CNTF are first discussed, followed by analyses of curing monitoring and forming quality, mechanical performance and damage sensing. The influence mechanism of the curing cycle on repair performance is then clarified. Results show that the CNTF exhibits not only rapid, stable, and uniform electrothermal performance but also a sensitive electrical resistance response to strain. Compared with conventional repair, electrothermal repair shows superior performance in temperature control and energy consumption, and both methods produce no obvious defects in forming quality. The resistance signal of the embedded CNTF can be used both to infer the cure evolution of the resin during the co-curing process and to detect damage initiation and propagation in the repaired structure under three-point bending loads. The embedding of the CNTF alters the local failure behavior of the stepped repair structure and slightly weakens its bending performance, though the effect is not significant. Furthermore, extending the holding time optimizes the distribution of curing-induced residual stress (CRS). When the holding time increases from 90 min to 150 min, the ultimate load recovery ratio rises from 75.89% to 83.34%, albeit with a diminishing beneficial effect. This study demonstrates that the CNTF holds promise as an integrated heating and sensing unit for in situ curing repair and condition monitoring of CFRP structures.

Countersunk holes are widely used to achieve flush riveting in aircraft panels, while the geometric deviations introduced during machining can compromise the assembly quality of riveted joints. In this study, the forced-assembly behavior of composite countersunk riveted joints under various countersunk-hole geometric deviations is investigated. A three-dimensional elastoplastic progressive damage model is developed to predict the plastic deformation, residual stress distribution, and multiple damage modes in composite joints. The assembly deformation and damage evolution of flat- and curved-plate joints are comparatively investigated using experiments and numerical simulations. The results show that the countersink-depth deviation intensifies the nonuniform expansion of the rivet and consequently requires a higher riveting force than the ideal joint. The increased interference shifts the peak hoop tensile stress away from the hole edge and aggravates assembly damage within the composite laminate. For joints with hole-axis angle deviations, posture adjustments of the rivet during forming produce asymmetric radial clearances on the two sides of the hole, resulting in nonuniform hole expansion and residual hoop stress. The plastic strain field and damage zone propagate preferentially toward one side. Flat-plate joints are more sensitive to the angle deviation, whereas curved-plate joints exhibit greater sensitivity to the depth deviation.

Digital Twin (DT) technology is pushing manufacturing toward higher intelligence and adaptability. However, existing DT modeling methods still rely heavily on customization, lacking universality and scalability for assembly-oriented manufacturing systems. To address this limitation, this paper proposes a modular DT control framework that couples graphical interaction with reusable functional modules. Based on the classical five-dimensional DT model, the virtual entity is refined into geometric and physical models, and the service system is expanded into behavior and task models, enabling a clearer description and direct correspondence between system structure and operational logic. A behavior-oriented modeling workflow and a data-mapping mechanism are established to enhance scenario adaptability and reduce modeling effort. A graphical DT modeling platform is developed on top of this framework. Multiple robotic manufacturing prototypes, including robotic drilling, robotic gluing, and hybrid drilling systems, are constructed to assess the generality and reconfigurability of the proposed approach. A drilling experiment is performed on the robotic drilling system to validate the DT-based control execution mechanism. The resulting holes exhibit an average positioning error of 0.23 mm and a diameter error of 0.012 mm, both meeting aerospace drilling requirements. This confirms that virtual task commands can be accurately executed on physical system under the proposed DT framework. Overall, the DT prototype implementations and drilling experiment jointly verify the scalability of the framework and its DT-based control capability, providing a practical approach for the rapid development and deployment of DT prototypes in aircraft assembly systems.

Due to its excellent mechanical properties and strong process feasibility, CFRP is commonly used to repair damaged aerospace metal components. However, the differences in physicochemical properties among the metal, adhesive layer, and patch often induce curing residual stresses and deformation in the bonded area after forming, which can affect the load-bearing performance of the repaired structure. This study investigates the effects of residual stress and curing deformation. Tensile tests were designed for single-sided CFRP patch repairs on damaged titanium alloy components, and a multi-stage numerical analysis method was established, covering the process from co-curing bonding to quasi-static tensile loading. Through comprehensive analysis of experimental and simulation results, the strain evolution behavior of the adhesive layer during the curing stage was clarified, the stress-strain curves and damage failure modes during the tensile stage were compared, and the influence of curing processes on tensile performance was explored. The results show that the strain in the adhesive layer during curing can be divided into five stages, with a simulation error of less than 20% compared to experimental data. The multi-stage simulation method, which accounts for residual stress and curing deformation, yields stress-strain curves that align more closely with experimental results, achieving an error of only 1.39% in ultimate tensile stress. Additionally, the failure modes of different materials observed in the simulation matched those in the experiments. Regarding curing processes, reducing the heating rate and extending the dwell time improved the curing degree and modulus of the patch and adhesive layer. This also reduced tensile stress and increased compressive stress in the titanium alloy substrate along the loading direction, ultimately enhancing the overall tensile performance of the structure.

Widely used in aerospace structures, composite honeycomb sandwich materials are costly and time-consuming to replace once damaged. Effective repair has become an essential approach for restoring function and extending service life. This study explored an efficient scarf repair strategy using adhesively bonded carbon fiber-reinforced prepreg patches. Facesheet-notched sandwich specimens were repaired with scarf patches, varying in scarf type, scarf angle, ply orientation, step pitch, and thickness of additional patches. Intrinsic modal frequency, compressive loading capacity, and strain distribution were employed to evaluate repair performance. A parametric analysis model based on analysis of variance was established to quantify the influence of each factor and guide optimization. Results show that the stepped scarf method outperformed the tapered one, especially when the scarf angle was lower than 1/20, due to its larger and more accurate bonding interface. Additional patch plies improved the load transfer by reinforcing the repair zone and providing overlapping paths with parent facesheet to redistribute stress. Among all variables, scarf angle had the greatest effect, followed by thickness, step pitch, and ply orientation. Specimens repaired with optimized parameters validated the accuracy of the quadratic regression model, and their load-bearing capacity was fully restored to the level of intact specimens.

To investigate the mechanical properties of woven composite hybrid bonded/bolted joints,a mechanical failure model for hybrid bonded/bolted joints based on 3D progressive damage model and cohesive force model was developed to simulate the mechanical behavior and damage evolution of the joints.Based on the Abaqus finite element software,a finite element simulation model for hybrid bonded/bolted joints woven composites was established.The damage initiation and propagation of the composite material were judged using the three-dimensional Linde criterion.The cohesive force model was used to simulate the damage failure process of the adhesive layer.On the basis of test verification of the model accuracy,the strength and damage failure process of the joints were analyzed under different tightening torques.The test and simulation results indicate that as the tightening torque increases,the extension of adhesive layer damage can be effectively suppressed.However,the shear strength of the adhesive layer in the hybrid bonded/bolted joints firstly increases and then decreases,because increasing the tightening torque can reduce the peeling stress of the adhesive layer in the joints.However,the excessive tightening torque will strengthen the stress around the adhesive layer hole,leading to a decrease in the shear srength of the adhesive layer and a decrease in the strength of the connection structure.The load-displacement curve of the numerical simulation is consistent with the test results,and the predicted adhesive layer fracture load is equivalent to the test results.At the same time,the fiber damage,matrix damage,and delamination damage on the laminated plate can also be well reflected in the numerical model,which is similar to the damage form after the connection test,verifying the effectiveness of the damage prediction model.

Aircraft structures consist of numerous complex components that require a high level of precision to assemble. Tooling plays a crucial role in the assembly of aircraft components, providing the functions of positioning, shape maintenance, and support to guarantee the accuracy of the product. Aiming to obtain reusable assembly tooling that can be rapidly reconfigured, this study focuses on the modular design and layout of tooling structures. The concept of functional elements for the characterization of tooling parts is proposed, and the relationship between each pair of elements is established to clarify the similarities and dependencies among various tooling structures. Based on the analysis of functional elements and their relationships, the tooling structures are divided and recombined into several modules. The detailed module designs are demonstrated by using typical structures such as platforms, columns, and locators as examples. A parametric representation of the multi-source information of tooling modules is proposed, and optimization methods for the layout and configuration of locators and platforms are developed using their parametric information. A reconfigurable tooling process integrated with a monitoring system is designed, realized, and successfully applied to the assembly of a practical type of fuselage. The results from verifying these methods’ efficiencies show that the modular design and reconfiguration planning of tooling only takes about 10 min and a few seconds, respectively, which is far less than the time consumed during traditional tooling design (from several days to weeks). The work in this study provides an engineering paradigm for the serialization and reconfiguration of assembly tooling in aviation manufacturing.

Fixture is widely used for locating the aircraft parts and ensuring the skin shape of aviation product in assembly. This research proposes an approach that determines the deformations of both fixture structure and skin using strain data gauged by fiber Bragg gratings to monitor their assembly status. The strain data of skin are transferred to curvatures and multiple curves are reconstructed. The skin surface is formed by the curves and its posture is registered to the fixture by location reference. An improved curvature interpolation method that considers the spatial lengths of gratings is proposed to enhance the shape accuracy of skin. On the purpose of adaptation to the bidirectional bending surface with variance of convexity, curve reconstruction and coordinate transformation algorithms are proposed. A mapping model from strain to displacement is established for the fixture structure by modal superposition. The orientations of multiple sensors on the structure are optimized to minimize the conversion error. The algorithms for deformation sensing are integrated into a developed monitoring software and undergo experiment tests. The results show that the maximum relative errors in skin displacement are 8.28 % and 1.94 % for traditional and improved curvature interpolations, respectively. The maximum error of board deformation is 6.13 %, which validates the efficiencies of the proposed methods. Verification of the software demonstrates that it can provide accurate and instant sensing of the deformations in aircraft assembly application.

As the lightest structural metal material, magnesium alloy is increasingly widely used in aerospace, rail transportation and other fields. However, during its service, the action of alternating loads often induces fatigue fracture damage, which seriously threatens its service safety and stability. This study investigates the fatigue behaviors of Mg-3Al-Zn (AZ31) magnesium alloy along its rolling direction (RD) and normal direction (ND) under strain-controlled tension-tension cyclic loading. The strain-life curve reveals that ND specimens in which the deformation was dominated by twinning-detwinning exhibits higher fatigue life than RD specimens dominated by slip. Morphological analysis of crack patterns indicates many grain boundary fractures and long surface cracks in RD specimens, contrasting with the ND specimens' uniform distribution of short, linear cracks throughout the texture. Scanning Electron Microscopy (SEM) and Electron Backscatter Diffraction (EBSD) analyses suggest that crack initiation in RD specimens is likely at grain boundaries due to deformation incompatibility between the grains with soft and hard directions. Conversely, the cracks of ND specimen predominantly formed at the {10 1 ¯ $$ \overline{1} $$ 2} twinning boundaries, owing to the strain incompatibility between the matrix and twin due to the discrepancy of basal dislocation activations.

Digital twin (DT) models with high-fidelity could map physical entity states precisely, raise the credibility of simulation, enhance the accuracy of processing decisions and improve feedback control precision in intelligent manufacturing, while the modeling process is frequently constrained by the complexity of the physical entity structure. This paper aims to propose a high-precision DT modeling method for aircraft assembly equipment and a drilling robot system with complex structures is taken as the research object. A physical model detailing the structure of the hybrid drilling robot is developed via the combination of Denavit-Hartenberg (D-H) and the virtual mechanism methods. A logical model is established based on the kinematic model of the hybrid drilling robot to express its behavior for drilling. The Levenberg-Marquardt (L-M) least-squares method is applied for calibration of DT model, which reduces the influence of geometric errors by identifying structural parameters in the physical model. The average position and normal errors have decreased to 1/10 and 1/8 respectively compared to before calibration, leading to enhanced accuracy in DT modeling. DT control software is developed to integrate physical model and logical model and is combined with hybrid drilling robot to construct DT system. The drilling quality experiment of DT system for flat and single curvature plates is designed and the results showed that the average positioning errors after the DT model calibrated are reduced by 39.29 % and 49.25 %, respectively. In addition, these drilling quality meets the drilling requirements of large aircraft body fastener assemblies.

Digital twin (DT) technology is changing the current pattern of intelligent manufacturing, it makes up for the shortcomings of process parameter optimization methods to improve real-time and predictability. This paper developed DT models for the robotic gluing system to predict the quality (width and thickness) of glue lines and optimize gluing parameters (trajectory and extrusion speeds). The DT framework based on the geometric, physical, behavioral, and rule models is constructed to monitor and optimize the gluing parameters in real-time. An improved backpropagation neural network (BPNN) prediction model based on whale optimization algorithm (WOA) is established to predict the width and thickness of glue lines from historical and real-time data, while simultaneously enabling real-time calculation of the cross-sectional area of glue lines. A multi-objective optimization model constructed using non-dominated sorting genetic algorithm (NSGA-II) is used to optimize the gluing parameters. The DT prototype of the robotic gluing system has been developed and verified experimentally. The position calibration of the geometric model is used to correct the gluing trajectory before gluing, and the position errors of the gluing points are within +/- 0.5 mm. The gluing trajectory is designed to test the effectiveness of the adaptive optimization of gluing parameters. The prediction errors of the width and thickness of the glue line are controlled between +/- 0.5 mm and +/- 0.3 mm, individually. After parameter optimization, the width and thickness of the glue line at the corner are reduced by 4.53 % and 7.54 %, respectively, thus avoiding glue accumulation. This reduction solves the problem of poor consistency in the quality of glue lines and verifies the feasibility of integrated monitoring, prediction, and optimization based on the DT model.

Riveting is a reliable joining process that is widely used in aircraft assembly, where axial loads can be transferred from one sheet to another through the load transfer action of rivets. In this article, the effects of hole diameter and squeeze force on the fatigue behavior of riveted joints were systematically researched. The results demonstrated that the hole diameter and squeeze force had great effects on the interference fit size and the interface contact state between the rivet and the hole, thereby affecting the fatigue properties. The fatigue life shortens as the initial hole diameter increases. The greater the squeeze force, the higher the fatigue life. With the increase of the initial hole diameter, the decreased degree of fatigue life is larger with the lower squeeze force. In addition, the fretting wear behavior was generated during the fatigue loading process. Two failure modes of outer sheet fracture and mixed fracture occurred in riveted joints.

Digital Twin (DT) technology is one of the key approaches to enhancing the intelligence of aircraft assembly equipment. However, the diversity of such equipment types and significant structural differences present substantial challenges to the development of DT models. This article proposes a unified V-shaped DT modeling paradigm to support high-accuracy and structured modeling. The robotic drilling system is used as an example to validate this paradigm. The modeling requirements of this system are established based on a comprehensive analysis of its structural characteristics and operational tasks. A corresponding virtual entity is constructed through parametric modeling based on kinematic analysis. The behavior model represents the interaction protocols and decision logic of the physical system, with basic modules for communication and behavioral analysis. These modules are then systematically integrated to form a complete task model for drilling. The structural validation of the virtual entity is performed, accompanied by the formulation of behavioral matching degree and task execution consistency to evaluate the effectiveness of the proposed modeling paradigm. Meanwhile, kinematic parameter identification is integrated to calibrate the virtual entity, thereby further enhancing the DT modeling accuracy. The experimental results show that the behavior matching degree for positioning after calibration is 0.204 +/- 0.228 mm, with an increase of 78.71 %. The average errors of hole position and diameter are reduced by 78.43 % and 14.27 %, respectively, after calibration. The corresponding task execution consistency is improved to 1.465 and 1.462. This indicates that the high-accuracy DT model constructed by the proposed paradigm effectively enhances the intelligence and assembly quality of the equipment.

The composite interlamination is the key part of load transfer and plays a decisive role in the overall mechanical performance of composite structures. The interlaminar property is predominantly influenced by the matrix resin which is easily affected by the hygrothermal environment. This paper delved into the effect of hygrothermal aging duration on the interlayer debonding behavior of unidirectional (UD) laminates. The pre-cracked specimens were treated by hygrothermal aging with a period of 0, 15, 30, 60, 90, and 120 days. The double cantilever beam (DCB) tests were employed to evaluate the interlaminar property. The effect of hygrothermal aging on load-displacement responses, R-curves, crack growth speed, and fracture surface topography was analyzed. The results revealed that the load-bearing capacity of aged specimens presented a reduction of 18.6%-29.0% compared with unaged ones. The interlaminar fracture toughness of specimens subjected to 15 days of treatment decreased significantly compared with that of unaged specimens due to the increase of water absorption caused by hygrothermal aging. With the increase in hygrothermal aging time, the interlaminar fracture toughness peaked at 60 days because of the post-curing effect, followed by a declining trend influenced by the dominant hygrothermal aging effect. The fiber bridging of aged samples was increasingly obvious, became the most prominent at 60 days, and then gradually reduced. The fiber bridging diminished the crack propagation speed, hindered interlaminar debonding, and contributed to the interlaminar toughening.Highlights The samples underwent hygrothermal aging for 0, 15, 30, 60, 90, and 120 days. Load ability of aged samples was at least 18.6% lower than that of unaged ones. R-curves rose before the crack length of 79 mm and then tended to stabilize. Fracture toughness first decreased, then increased and reduced with aging time. Fiber bridging could hinder crack growth and improve the interlayer property. Mechanisms of interlaminar toughening (post-curing) and property degradation (aging) of unidirectional composite laminates under different hygrothermal aging durations. image

The present research aims to investigate efficient repair techniques of cracked Ti-alloy aircraft structures with adhesively bonded carbon fiber-reinforced polymer prepreg patches. The repaired specimens in the configuration of a Ti-alloy butt joint with one-side bonded composite patch were prepared under multiple repair factors including patch thickness, patch length, adhesive thickness, cure pressure, patch layup and surface treatment. The repair efficiency was evaluated by loading behavior, bonded interface microstructure and failure mode. The results reveals that the geometric factors affect the loading performance and alter failure modes by adjusting stress distribution in the repair system, whereas the cure pressure and surface treatment act on the bondline and change interfacial properties. A sensitivity-optimization model based on analysis of variance was established for parametrical study to quantify the contribution of repair factors and obtain optimal values. The optimum parameters were validated by repaired central-cracked specimens via static and fatigue tests, which proved that the repaired structure could restore 90.7% loading capacity of intact ones and endure more than 106 fatigue cycles of 25% ultimate failure load level of center-cracked ones. The proposed experimental and parametrical study possessed good efficacy in refurbishing strength and stiffness of cracked metallic structures.

The paper investigates the validity and reliability of the blind-bolt repair method for repairing delaminated composite aircraft panels. The delaminated specimens are prepared by inserting Teflon films during the manufacturing process to simulate interlayer damage. Subsequently, these specimens are repaired using the blind-bolt method. Modal and uniaxial compression tests are conducted to quantitatively evaluate the natural frequency, mode shape and load-bearing strength of both delaminated and bolt-repaired specimens. Digital image correlation and ultrasonic phased array techniques are employed to characterize buckling instability and damage evolution of specimens. The results reveal that the natural frequency and compressive buckling strength of delaminated specimens significantly decrease. The mode shape also changes nonlinearly with the stiffness reduction. This variation is proportional to the size and the quantity of delaminations. The blind-bolt repair method effectively restores the vibration and mechanical properties of the delaminated composite structure by reconnecting separated sub-laminates. A repair tolerance of 20 mm-60 mm is recommended for a single blindbolt. When the delamination length is 35 mm, the repair efficiency for the critical buckling load and the ultimate load is the highest, at 58.3% and 64.4%, respectively.

The quality of aircraft assembly is mainly guaranteed by toolings which are vital to the geometrical accuracy and service performance of aviation products. In this research, a real-time monitoring system that determines the structural deformation and clamping force of reconfigurable toolings using strain data is developed to perceive the service state of the toolings. By laying fiber Bragg gratings on positioning beam and baseplate of the reconfigurable tooling, strain data of the tooling structures are gauged and transferred to curvatures. The beam and baseplate are modeled as one and two-dimensional objects respectively and shape reconstruction algorithms are established to obtain their deflection curve and surface using curvature information. Distribution of fibers is optimized to minimize the conversion error from strain to curvature. An estimation that reveals the mathematical relationship between the shape reconstruction error and measurement interval is implemented, and a mapping model from strains to clamping force of the beam is established. These algorithms are integrated into the self-developed monitoring software and undergo simulating and experimental tests. The maximum relative errors of deformation and force are 4.53% and 4.12% respectively in simulation, and 9.21% and 7.29% individually in experiment, which validates the efficiencies of the method. Tests of the monitoring system suggest that it can provide a timely and accurate sensing of the deformation and force of the tooling.