Abstract Carbon-fiber-reinforced (CFR) composites in aircraft structures are subjected to complex, multiaxial loading conditions that may induce fatigue damage prior to final failure. To ensure structural safety, reliable failure criteria must be established for both undamaged and fatigue-affected materials. This study presents experimental investigations of tubular CFR composite specimens subjected to combined axial force and internal pressure, generating complex stress states in the thin-walled gage section. The specimens were loaded to failure along various stress paths, enabling construction of a failure surface in principal stress space. Three distinct failure modes were observed: resin matrix puncture, longitudinal cracking, and circumferential cracking with specimen separation. A probabilistic approach was introduced to account for the large scatter in experimental data, replacing deterministic failure stresses with stress values corresponding to specified survival probabilities. The results indicate that the maximum principal stress criterion, formulated in three-dimensional principal stress space with axes aligned to fiber directions, provides a suitable framework for the investigated composite. Incorporating probabilistic assessment improves reliability in predicting composite failure under complex loading.

Abstract Thermoplastic composites enable weldable, recyclable aircraft structures, but thermal mismatch between metals and polymers can introduce detrimental residual stresses. This study develops a finite element method (FEM) framework to predict residual stress fields in resistance-welded joints between aluminum 7075 and carbon-fiber-reinforced polyamide 6 (PA6). Transient thermal analyses with multilinear, temperature-dependent properties were coupled to mechanical analyses; contact conditions transitioned from frictional to bonded at PA6 melting. Three thermal cycles (20°C→220°C→20°C, 20°C→240°C→20°C, 20°C→260°C→20°C) were examined to assess peak-temperature effects. The simulations show stress contours that decay with distance from the bond and reveal pronounced peaks in both normal and shear components at weld edges, consistent with shear-lag theory. Within the bonded interior, average stresses are relatively low, whereas edge concentrations identify likely sites for debonding or delamination initiation. The magnitude of residual stresses increases with thermal gradient, underscoring the need for parameter control during welding. The FEM outputs will be validated against uniaxial tension and three-point bending tests on welded specimens, with future work quantifying fatigue-life reduction under combined thermal and mechanical cycling. The results highlight mitigation priorities for bonded repairs and hybrid aerospace structures, including process-curve tuning (current/pressure/cooling) and edge-region design measures.

Abstract Bolt connections are widely used in aircraft engines due to their advantages of high stiffness, low weight, and ease of assembly and disassembly. However, they are subjected to complex stress states in service, including high preloads, combustion-induced forces, and random vibrations, which may lead to fatigue failure. Since the fatigue performance of bolts directly affects the reliability of engines and their fuel accessories, accurate fatigue life estimation is essential for safe design. This study proposes a frequency-domain method for evaluating the random vibration fatigue life of bolts in aeroengine fuel accessories. A detailed finite element model of the bolted connection was established, with excitation boundary conditions defined by the assembly configuration. Modal, frequency response, and random response analyses were performed to obtain the stress power spectral density (PSD) of the bolts. The Dirlik method, combined with the material’s S–N curve, was then applied to estimate fatigue life under broadband and narrowband vibration excitation. Results show that bolts in the Y-direction experience the highest RMS stresses, leading to the shortest fatigue life – approximately 4.44 hours under critical loading conditions – which does not meet design requirements. The proposed method enables rapid evaluation of bolt fatigue life under random vibration environments, providing a practical tool to support bolt selection and design optimization in aeroengine applications.

Abstract The lower wing section of an aircraft is considered particularly vulnerable to fatigue failure due to the presence of inspection holes, which create stress concentrations and increase local stress in the surrounding material. This study estimates the fatigue life of the lower wing structure, including rivet holes around the inspection openings, in a new-generation Indonesian short takeoff and landing (STOL) aircraft under cyclic flight loads. Fatigue assessment was conducted in five stages: (1) development of a 3D design model of the lower wing skin; (2) stress analysis of the skin without rivet holes, using finite element analysis (FEA), to identify critical areas around the inspection hole; (3) stress analysis of the skin with rivet holes in these critical areas; (4) compilation of a stress spectrum from flight test data; and (5) fatigue life estimation using the cumulative damage method with the application of a scatter factor. The analysis results indicate a maximum fatigue life of 67,750 flight cycles for rivet holes in the lower wing skin, exceeding the industry target of 30,000 cycles. However, when a scatter factor is applied, the maximum fatigue life is reduced to 13,550 flight cycles, establishing the required inspection threshold for the STOL aircraft.

Abstract The refill friction stir spot welding (refill FSSW) process is an innovative solid-state spot-welding method, which has evolved from the concept of friction stir welding. Compared to riveting, the process has the advantage of avoiding stress concentration by eliminating holes. In addition, weight can be saved compared to riveting as no additional material is needed. However, the fatigue strength of refill FSSW joints under cyclic loading is still not satisfactory. To address this challenge, laser shock peening (LSP) is investigated as an innovative residual stress engineering technique to improve the fatigue performance of refill FSSW AA2024-T3 joints. Two application scenarios are investigated, one investigating the LSP technique as a complementary manufacturing process to the refill FSSW technology, and the other investigating the LSP technique as a repair process for damaged joints. The fatigue test results showed that the application of the LSP treatment can significantly improve the fatigue behaviour of the refill FSSW overlap joints. In terms of Basquin fatigue strength, the LSP treatment resulted in an improvement by a factor of 1.51 and 2.82 for the one- and two-sided LSP-treated specimens, respectively. The life of specimens with refill FSSW joints that had been specifically pre-damaged by stopping the fatigue test at approximately 51%, 75% and 83% of the number of cycles to the Basquin fatigue strength, applying LSP treatment and continuing the fatigue test was also significantly extended. The results of this study show that LSP is a very effective technique for significantly extending the fatigue life of refill FSSW joints. Therefore, the combination of these two manufacturing processes, refill FSSW and LSP, represents a promising technology for industrial companies that require high fatigue performance for their structural components.

Abstract Digital twins (DTs) can connect inspection data with product models to support safer, more efficient lifecycle decisions. This paper proposes a CAD-native workflow for implementing a digital twin that visualizes and manages non-destructive testing (NDT) results directly on a 3D model. The method supports over-the-surface data (ultrasonic C-scans, UT) via UV mapping and projected images (thermography, TT) via planar projection, both executed in Siemens NX with custom macros for point localization and on-surface measurement. We validate the approach on a bottom nacelle panel from a Honeywell HTF7000 turbofan engine, acquired via 3D scanning and reverse engineering. The resulting digital twin preserves a persistent spatial link between inspection images and geometry, enables remote sizing and review, and centralizes result management in the CAD environment for PLM use cases (e.g., defect history, trend analysis). Timelines indicate higher initial effort but reduced on-site workload and travel for qualified inspectors thereafter. Limitations include large file sizes when storing geometry and multiple images in a single model; we outline a lightweight distribution strategy and future automation/VR enhancements. The findings demonstrate the feasibility and practical value of CAD-resident digital twins for NDT visualization, remote evaluation, and product lifecycle management.

Abstract The increasing demand for greener aviation technology has driven the adoption of advanced composite materials in aircraft structures, offering significant weight saving and fuel efficiency improvements. Wing structures made of torsion boxes composed of stiffened panels have shown over the decades the benefits provided by composite technology, which is able to adapt the material properties to the structural constraints to which the structure is subjected. A convenient manufacturing technology for stiffened panels consists of co-bonding stiffeners on pre-cured skin. On these structures, the certification authorities require the demonstration of the residual strength at limit load of the panel with a disbonded stiffener. This is typically a post-buckling problem where the complete failure of the panel is due to a secondary buckling mode or the failure of adjacent stiffeners due to the combination of compressive and tearing loads, the larger buckled panel bay generating pull out loads on the adjacent stiffeners. This demonstration is classically performed by tests on large stiffened panels. In a composite wing box where geometrical parameters and loading modes can vary significantly from one zone to another, the numerical simulation can bring significant benefits to reduce the number of tests. This article presents a numerical damage model able to predict the damage and disbond of a co-bonded stiffener and applicable to a large aircraft structural model that can be integrated into a post-buckling simulation. As a validation case, this model has been applied to a multi stiffened composite panel where a central stiffener has been disbonded. The simulation results gave accurate predictions for the buckling loads and modes as well as for the appearance of damages on stiffeners until the panel failure.

Abstract This study introduces a Bayesian-informed framework for fatigue life prediction in shallow shell structures. The methodology focuses on inferring the Equivalent Initial Flaw Size Distribution (EIFSD), a critical parameter for structural durability. Bayesian inference, combined with a Co-Kriging surrogate model, enables statistically robust predictions while accounting for uncertainties in material properties, geometry, and loading. The Dual Boundary Element Method (DBEM) is employed for crack propagation due to its efficiency and re-meshing-free modelling. To improve inference efficiency, an iterative parameter space narrowing strategy is proposed. Instead of exhaustively sampling the entire space, the method begins with coarse discretisation to locate high-probability EIFSD regions, then refines them adaptively. A numerical example involving a fuselage window under cabin pressure demonstrates the method. Surrogate models trained on DBEM-generated data significantly reduce computational cost. The proposed strategy achieves high-precision inference, with only 0.059% error in the inferred mean and 5.2% in standard deviation, while reducing CPU time by 52% compared to dense sampling.

Abstract In a context of growing importance of mass reduction and reliability of structures towards greener aircrafts, fatigue of metallic materials is a key issue in the structural optimization. The process used by aeronautic industrials to compute the fatigue life is often based on a large empirical experience and meets a need for efficiency in their application, requiring a compromise between accuracy and ease of use. According to legacy crack initiation methodologies, lifetime computation is based on the analysis of elastic stress fields, calculated analytically or by Finite Element Method. Evaluation of lifetime is calibrated on elementary tests, mainly uniaxial, with geometric specificity (bone, hole, notch…). One of the limits of this approach appears when parts are subjected to multi-axial loads. Nowadays, these particular stress states are justified by conservative approaches to ensure flight safety and by tests on full-scale aircrafts. Whether for the operational maintenance or the structural optimization of new aircrafts, it is intended to enhance crack initiation methodologies, taking into account multiaxiality of loads, stress gradient effects, and complex material behaviours. Dassault-Aviation implements a crack initiation lifetime computation based on a local approach. These developments go hand in hand with a PhD (Nutte, 2023) on a multiaxial fatigue criterion in order to predict crack initiation in metallic assemblies. This work was supported by an innovative dedicated test campaign. The identification of material’s parameters is based on uniaxial and multi-axial mechanical tests, specifically designed to calibrate these models. Then, novel geometry of specimen for bolted assemblies, facilitating various biaxial non-proportional loadings, is used to evaluate the methodology. Also, multiaxial fatigue models require a precise assessment of the local mechanical fields to which the structure is subjected. For this, a finite element analysis must be conducted with a level of complexity associated with the level of accuracy targeted. The material constitutive equations used in the finite element analysis are therefore at the heart of these fatigue substantiation approaches. Applications to complex aeronautical structures such as massive 3D parts or assembly by fastener will highlight the benefits and perspectives for this local fatigue approach. It will require the use of multi-scale data science.

Abstract JAXA has been conducted the research to evaluate the fatigue life up to form a certain size of fatigue crack in a CFRP/Aluminum hybrid joint. Thermal stress occurs in the hybrid joints during operation due to the difference of coefficient of thermal expansion between Aluminum and CFRP. In ICAF2023, we presented the experimental and numerical results for the hybrid joint under thermal cycles. In this study, the mechanically fastened hybrid joint specimens composed of two Aluminum plates and two CFRP plates are prepared. Most of the dimensions of the hybrid joint such as thickness, width of the plates and types and pitch of the fasteners and etc. are same as those evaluated in the previous research. The cyclic thermal and external loads are simultaneously applied to the hybrid joint and stress and strain on the Aluminum plate are evaluated experimentally and numerically. From the behavior of the elastic strain, which is the total strain minus the thermal strain, it is shown that when thermal load and external load are coupled, the hysteresis loop becomes larger than when only the external load is repeatedly applied. In addition, it is shown that the strain and stress around the fastener holes in the top row through the load direction are high, and this could be a critical area for fatigue failure.