Abstract This study investigates the applicability of Long Pulse Thermography (LPT) for detecting artificial defects in carbon-fiber-reinforced polymer (CFRP) composites. The work aimed to determine which types of defects can be detected in a CFRP panel using active infrared thermography and to assess the influence of defect type, size, and depth on their detectability. The test specimen was a CFRP laminate manufactured from unidirectional prepreg using the autoclave process. The panel had a variable thickness, and artificial defects were introduced at different locations to simulate typical manufacturing- and service-related damage. The investigated defects included: prepreg foil, simulating interlayer insertion caused by errors during prepreg lay-up; flat-bottom holes, simulating delamination caused, for example, by low-energy impacts; paper inserts, simulating interlayer insertion caused by the inclusion of a foreign body during manufacturing; flat-bottom holes filled with epoxy resin, simulating resin-rich areas; and graphite foil, simulating interlayer insertion defects. Thermographic measurements were performed using long-pulse thermal excitation, and the thermal response was analyzed during the cooling phase. The results showed that defect detectability depended strongly on defect type and increased with increasing defect diameter. Flat-bottom holes and resin-rich defects produced the clearest thermal indications and the highest detection repeatability, whereas prepreg foil and paper inserts yielded less distinct signals, particularly for smaller diameters. Under the applied test conditions, graphite foil inserts were not detected. The results confirm that LPT is a promising non-destructive testing method for the rapid inspection of CFRP structures; however, its effectiveness is limited for defects generating low thermal contrast.
Abstract Damage tolerance analysis provides the analytical basis for structural fatigue assessment and inspection planning in transport category aircraft and is required for certification under contemporary airworthiness regulations. In current practice, damage tolerance analysis commonly relies on generalized assumptions for initial flaw size and prescriptive inspection intervals without explicitly incorporating differences in nondestructive inspection detection capability. A baseline damage tolerance analysis was conducted on a simplified structural configuration selected to illustrate the underlying concepts without unnecessary analytical complexity, using established industry and regulatory assumptions. The analysis was repeated with the initial crack length defined by inspection capability. The results demonstrate that the assumed initial flaw size governs inspection initiation and is directly related to the maximum probable undetected flaw required by certification regulations, resulting in substantial shifts in inspection timing. A cumulative probability of detection formulation is introduced to evaluate inspection strategies in terms of total detection reliability, enabling alternative combinations of inspection method and frequency to achieve equivalent detection performance. This approach provides flexibility in inspection method selection and interval definition and may enable the use of lower-cost or more readily deployable inspection techniques while maintaining the required level of structural reliability. The proposed framework establishes a direct linkage between crack growth prediction and inspection capability, improving analytical consistency while preserving regulatory intent.
Abstract Vibration-induced fatigue cracks in turbine blades can lead to engine failure and the forced withdrawal of an aircraft or other turbine-powered machine from service. One means of preventing such cracks is to reduce blade vibration through increased damping. This paper presents a concept for enhancing the vibration damping of a turbine blade by implementing damping mechanisms within the blade structure itself. A prototype blade was manufactured using additive technology, which enabled the creation of internal pockets containing non-melted metallic powder. Lattice bars and pins of various geometries, and consequently different natural frequencies, were built into the pockets as elements fixed to the solid structure, with the function of absorbing vibration energy and transferring it into the surrounding powder volume, where the energy is dissipated. The blade was subjected to experimental testing on an electrodynamic shaker to determine its natural frequencies, mode shapes, and modal damping ratio. A significant increase in the damping ratio was observed relative to a reference Solid blade manufactured using the same technology. While the results demonstrate the effectiveness of the proposed solution, they also reveal pronounced vibration nonlinearity. Accordingly, subsequent work will focus on optimizing the blade damping structures and developing a correlated computational model.
Abstract Drilling-induced residual stresses can significantly impact the fatigue performance of aluminum alloy components. This paper presents an experimental and numerical study on the impact of rotational speed and feed rate on residual stresses experienced while drilling Al2024-T351 in a dry cutting environment. Drilling experiments were carried out using an 8 mm HSS drill at rotational speeds from 375 to 1250 rpm and feed rates between 0.12 and 0.30 mm/rev. The residual hoop stresses near the hole entrance were measured by surface-mounted strain gauges. For residual stress simulation, a coupled thermo-mechanical finite element model was developed using Abaqus/Explicit. Results showed that increasing rotational speed decreases tensile residual stress, and increasing feed rate greatly increases it. The numerical model reproduces the experimental trends but slightly overpredicts the magnitude of the stress. The findings provide practical guidance for selecting drilling parameters to minimize residual stress in aerospace aluminum alloy component.
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 This study investigates the influence of geometric discontinuities specifically, circular holes – on the fracture behavior of fiber-reinforced polymer composite plates containing edge cracks. Using the Finite Element Method (FEM) as a non-destructive numerical tool, we analyze how variations in hole size and positioning affect the stress distribution σ yy and the J-integral near the crack tip. The J-integral, used both as a fracture toughness parameter and a crack detection indicator, effectively captures the energy release rate and the severity of crack-tip conditions in anisotropic composite materials. Reference models without holes were first validated before introducing circular holes at varying distances from the crack front. The results reveal that strategically placed holes can reduce local stress concentrations and lower the J-integral values, mitigating the driving force behind crack propagation. These findings highlight the importance of geometric optimization and advanced simulation in improving damage tolerance and fracture resistance of composite structural components. Furthermore, the insights gained are particularly relevant to the aerospace industry, where fiber-reinforced polymer composites are increasingly used in aircraft structures due to their high strength-to-weight ratio and fatigue performance. Understanding fracture behavior under geometrical modifications enhances the structural reliability of such lightweight components under operational conditions.
This study investigates fatigue crack propagation in aluminum alloy 2024-T351 under constantamplitude loading (CAL) and variable-amplitude loading (VAL), with a focus on energy parameters. The hysteretic energy dissipated at each load level (or per block) exhibits trends comparable to those observed under CAL conditions. At high propagation rates, both the dissipated hysteretic energy and the crack growth rate vary linearly, consistent with a propagation mechanism characterized by the formation of striations in each cycle. At lower propagation rates, the relationship between crack growth rate and dissipated energy per block follows a power-law form. The analysis is extended through quantitative microfractographic observations using scanning electron microscopy, which reveal the spatial distribution of key fractographic features and the mechanisms governing crack propagation. Based on these findings, an energy-based model is proposed that allows the replacement of load spectra with an equivalent constant-amplitude loading.
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.
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.
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.
Detecting fatigue-induced progressive damage under varying environmental conditions remains a major challenge in structural health monitoring (SHM). This study investigates a baseline-free nonlinear guided wave method, which extracts nonlinear parameters to detect fatigue cracks without requiring baseline signals from the pristine state. The method demonstrates reliable detection of cracks around 3 mm in size, with the nonlinear parameter serving as a sensitive indicator of damage initiation and growth. Its independence from baseline signals enhances practicality for in-service monitoring applications. However, experimental results reveal that the method’s performance is sensitive to temperature variations, with irregular responses observed at different temperatures, which may affect detection consistency. These findings highlight both the potential and the limitations of nonlinear guided wave methods, underscoring the need for temperature compensation strategies to improve their robustness under variable environmental conditions. Overall, the proposed approach contributes to advancing baseline-free SHM techniques by offering a viable solution for progressive crack detection in realistic service environment.
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.
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.
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.
As aircraft fleets age, maintaining operational readiness at an affordable cost becomes increasingly challenging. This is largely due to the rise in Preventive Maintenance Task Requirements (PMTRs) outlined in the Aircraft Maintenance Program (AMP). While aging aircraft may require more frequent inspections, leveraging data from prior inspections enables the optimization of inspection intervals based on risk, ensuring cost efficiency by minimizing unnecessary downtime, while maintaining the required safety level.
This paper describes a study to determine whether fatigue test life scatter is best characterised by a Weibull or lognormal statistical distribution for a high strength steel used for landing gear structures. It is a response to “Face 2” of the ICAF 2017 Plantema Memorial Lecture and 2019 follow-up paper with the question; “Weibull or Lognormal Distributions to Characterize Fatigue Life Scatter?” These concluded that a Weibull distribution appears to be more suitable than a lognormal distribution for statistical modelling of fatigue life scatter to define an allowable service life at a specific probability of failure. Those studies used a homogenous dataset of 18 fatigue tests, and a non-homogenous dataset of 86 fatigue tests from a variety of sources. This paper reviewed HBK historical fatigue tests to identify a homogeneous dataset of 371 fatigue tests for a high strength steel used for landing gear structures. Weibull and lognormal statistical modelling of this dataset concluded that its fatigue life scatter is best characterized by the lognormal distribution.
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.
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.
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.
The work presented is the result of the implementation of diffraction measurements: phase composition and stresses resulting from additive manufacturing process of nickel superalloy Inconel 718 components print. With the help of diffraction methodologies, the key parameters from the point of view of the quality of prints and their strength were determined. The existence of individual phases in the material after printing was demonstrated, and the surface variation of the stress values was presented, showing its dependence on the geometry of the printed part – measurements were made at various points on the surface of samples with different geometries. In addition, the variation of the stress level was shown depending on the distance of the measurement point from the build platform on which the additive manufacturing process was carried out. Components were printed on the surface of a single build plate in order to study the effect of printing differently oriented samples with respect to the platform geometry, as well as the mutual effect of the temperature of samples printed first on the stress state of elements printed in subsequent steps of the procedure, and the effect of the temperature of elements printed later on the rate of temperature decrease, and consequently on the stress state, of elements printed first.
