Fuselage Joints Research Articles

The impact of passive shoulder exoskeletons during simulated aircraft manufacturing sealing tasks

Aircraft manufacturing involves an operation called sealing during which fuselage joints are sealed utilizing a sealant and smoothing tool, requiring motion of a worker's arms at heights below the elbow to overhead. The objective of this study was to assess the impact of shoulder exoskeletons on shoulder and torso muscle activation compared to no exoskeleton during simulated aircraft sealing tasks. Electromyographic signals from bilateral pairs of the anterior and medial deltoids, trapezius, latissimus dorsi, triceps, biceps, and erector spinae were collected from 16 aircraft workers as they performed simulated horizontal and vertical sealing tasks at heights ranging from chest level to above head level while wearing three different passive shoulder exoskeletons and a no-exoskeleton condition. Normalized electromyographic signals were assessed by two-way within-subjects ANOVAs and post-hoc tests. Exoskeletons consistently significantly reduced normalized EMG for the anterior deltoid and medial deltoid muscles at both horizontal sealing heights and during vertical sealing motions, with greater reduction in normalized EMG realized at higher vertical heights. While these laboratory results are encouraging, these exoskeletons must be evaluated in worksite studies to better assess their efficacy. Relevance to Industry: Passive exoskeletons are receiving increased interest as potential interventions to reduce muscular loading and work-related musculoskeletal disorders. The findings from this study on multiple passive shoulder exoskeletons highlights the impact on shoulder and torso muscles during simulated aircraft manufacturing sealing tasks. These results must be considered in context with other factors that may also impact effectiveness of passive exoskeletons including other task locations, comfort from wearing exoskeletons for extended periods, environmental conditions, hygiene, among other considerations.

The Effects of Fatigue Cracks on Fastener Loads during Cyclic Loading and on the Stresses Used for Crack Growth Analysis in Classical Linear Elastic Fracture Mechanics Approaches

High strength threaded fasteners are widely used in the aircraft industry, and service experience shows that for structures where shear loading of the joints is significant, like skin splices, fuselage joints or spar caps-web attachments, more cracks are initiated and grow from the edges of the fastener holes than from features like fillets radii and corners or from large access holes. The main causes of this cracking are the stress concentrations introduced by the fastener holes and by the threaded fasteners themselves, with the most common damage site being at the edge of the fastener holes. Intuitively, it is easy to visualize that after the crack initiation, during the growth stages, some of the load transferred initially by the fastener at the cracked hole will decrease, and it will be shed to the adjacent fasteners that will carry higher loads than in uncracked condition. Using currently available computer software, the method presented in this paper provides a relatively quick and quantitatively defined solution to account for the effects of crack length on the fastener loads transfer, and on the far field and bypass loads at each fastener adjacent to the crack. At each location, these variations are determined from the 3-dimensional distribution of stresses in the joint, and accounting for secondary bending effects and fastener tilt. Two cases of a typical skins lap splice with eight fasteners in a two rows configuration loaded in tension are presented and discussed, one representative for wing or fuselage skins configurations, and the second case representative for cost effective laboratory testing. Each case presents five cracking scenarios, with the cracks growing from approx. 0.03 inch to either the free edge, next hole or both simultaneously.

Development of structural health monitoring methods through the analysis of kinetics of local stress-strain state

The paper describes one of the most promising ways of structure health monitoring based on analysis of the kinetics of local stress-strain state. Comparison with other methods of damage detection and advantages of the stress-strain state control are considered. To implement this method of monitoring, two types of sensors are considered: strain gauges and fiber Bragg gratings (FBG). The advantages of FBG, due to which their use in operation is most preferable, are presented. The possibility of using stress-strain state control is analyzed for a promising transport aircraft using the results of tensometry under cyclic loading. The relevance of the work is confirmed by various examples of using monitoring methods in on-board systems during operation. The possibility of monitoring the airframe parts is demonstrated for longitudinal fuselage joints as an example. The change in the stress-strain state of the longitudinal joints, caused by structural damage was confirmed by tensometry analysis and by analysis of the stress-strain state of bearing sheet by the finite element method (FEM), taking into account the fatigue cracks originated during cyclic loading. The results of the calculation are compared with the tensometry data. Proceeding from the results of the calculation of the stress-strain state of the damaged and undamaged structure, the zones of the greatest deformation change are determined to optimize potential locations for the sensors. The optimal placing of the sensors for monitoring the integrity of the joints is proposed. The study confirms that the method of damage monitoring of the structure, developed on the basis of changes in the kinetics of the stress-strain state, can be successfully used in tests for maintaining the integrity of joints and other places with multi-site damages, where the methods of visual and non-destructive testing are not effective.

Monitoring of the elements of aviation structures using strain-gauge measurement

Analysis of the kinetics of the stress-strain state was carried out using data of tensometry as one of the methods for monitoring the integrity of the aircraft structure and a promising approach to ensure safety of aviation equipment during long-term operation. The advantages of condition monitoring using on-board systems prior to routine procedures aimed at maintaining the airworthiness (which consist mainly in periodic inspections) are presented. The results of using monitoring when studying the fatigue characteristics of the aircraft structure elements at the Central Aerohydrodynamic Institute (Russia) are shown. The study was carried out on samples of longitudinal fuselage joints. Several types of fatigue damages typical for operation of longitudinal joints are considered. The possibility of detecting defects by monitoring of the local stress-strain state with the help of strain gauges under various modes of destruction is assessed. Data on the of fracture surfaces of the damaged elements, fronts and crack propagation rates corresponding to different number of applied cycles were used. The assessment of the controllability of several typical structural and technological variants of longitudinal joints is presented. The results of determining zones of changes in stress-strain state in a bearing sheet with a single main crack are presented, which is the most typical fatigue damage to the longitudinal fuselage joints. The optimal location of the sensors for monitoring the integrity of the joint is proposed. Studies proved that the method of structural health monitoring using tensometry data can be successfully applied to tests for durability and has a prospect of being used in maintaining the integrity of joints and other places of limited access by instrumental non-destructive test methods.

Development of fatigue cracks from mechanically machined scratches on 2024‐T351 aluminium alloy—part I: experimentation and fractographic analysis

AbstractClad and unclad 2024‐T351 aluminium alloy sheets, weakened by mechanically machined scratches, were fatigued to investigate the effect of small surface damage, like scribe marks, on aircraft fuselage joints. The role of scratch cross section geometry on fatigue life of scribed components was analysed. Scratches between 25 and 185 µm deep, with 5, 25 and 50 µm root radii, were cut on sample surface by using diamond‐tipped tools. After testing, fracture surfaces were examined using a scanning electron microscope, and crack growth rates were measured by striation counting. Scratches reduced aluminium fatigue life under tensile and bending load up to 97.8% due to multiple crack nucleation at their roots. Short cracks nucleated from sharp scratches coalesced to form unique elongated cracks growing through sample thickness. Cracks initiated from scratches were typical short cracks, growing faster than conventional long cracks. Despite the different scribing process, fatigue data of regular diamond tool cut scribes can be used to conservatively predict life reduction owing to ploughed in‐service scribe marks on fuselage joints. Finite element analyses on scribed samples and the fatigue life prediction models are described in Part II of this paper.

Fatigue Life Estimation of Aeronautical Joints Based on Stress Severity Factor

The goal of the present work is to investigate the validity limits and safe application range of Stress Severity Factor methodology in estimating fatigue life of aircraft fuselage joints. Fatigue tests were conducted and recorded data from aluminum alloys joints was subjected to analytical evaluation. FE models were created to obtain fasteners load distribution and determine normal stress due to secondary bending. Severity Factor method conservatively estimated fatigue life of 74% for the analyzed joints. Its robustness was verified for lap joints fatigue life estimative, but for almost all single strap joints, secondary bending effect were significant. Thus for this kind of joints, a term accounting for bending stress was added to the original severity factor formulation to increase level of safety in fatigue life estimates.

Finite element analysis of catastrophic failure of dynamically-loaded countersunk composite fuselage joints

Accurate models of dynamic structural failure are important for crashworthiness studies. To date, catastrophic failure of dynamically-loaded composite bolted joints has been studied using global or stacked shell element models. In this paper, high-fidelity (three-dimensional solid) explicit FE models are used to simulate catastrophic failure of countersunk composite fuselage joints. While current state-of-the-art 3D modelling approaches focus almost exclusively on the prediction of composite damage, this study also investigates the treatment of fastener damage. Fastener fracture is a common catastrophic joint failure mode, particularly in joints designed to initially fail in bearing. A Johnson–Cook material model and cohesive elements were used to predict plasticity, damage and fracture of the titanium (Ti–6Al–4V) fastener. Although a model calibration was required, due to the complex interaction of model parameters, numerical results demonstrate key trends of experiments and provide a starting point for the development of more predictive approaches for simulating fastener failure.

Inter-laminar shear stress in hybrid CFRP/austenitic steel

Bolted joints are the most common solution for joining composite components in aerospace structures. Critical structures such as wing to fuselage joints, or flight control surface fittings use bolted joining techniques. Recent research concluded that higher bearing strengths in composite bolted joints can be achieved by a CFRP/ Titanium hybrid lay-up in the vicinity of the bolted joint. The high costs of titanium motivate a similar research with the more cost competitive austenitic steel. An experimental program was performed in order to compare the apparent inter-laminar shear stress (ILSS) of a CFRP reference beam with the ILSS of hybrid CFRP/Steel beams utilizing different surface treatments in the metallic ply. The apparent ILSS was determined by short beam test, a three-point bending test. Finite element models using cohesive elements in the CFRP/Steel interface were built to simulate the short beam test in the reference beam and in the highest interlaminar shear stress hybrid beam. The main parameters for a FEM simulation of inter laminar shear are the cohesive elements damage model and appropriate value for the critical energy release rate. The results show that hybrid CFRP/Steel have a maximum ILSS very similar to the ILSS of the reference beam. Hybrid CFRP/Steel is a competitive solution when compared with the reference beam ILSS. FEM models were able to predict the maximum ILSS in each type of beam.

Static and high-rate loading of single and multi-bolt carbon–epoxy aircraft fuselage joints

Single-lap shear behaviour of carbon–epoxy composite bolted aircraft fuselage joints at quasi-static and dynamic (5m/s and 10m/s) loading speeds is studied experimentally. Single and multi-bolt joints with countersunk fasteners were tested. The initial joint failure mode was bearing, while final failure was either due to fastener pull-through or fastener fracture at a thread. Much less hole bearing damage, and hence energy absorption, occurred when the fastener(s) fractured at a thread, which occurred most frequently in thick joints and in quasi-static tests. Fastener failure thus requires special consideration in designing crashworthy fastened composite structures; if it can be delayed, energy absorption is greater. A correlation between energy absorption in multi-bolt and single-bolt joint tests indicates potential to downsize future test programmes. Tapering a thin fuselage panel layup to a thicker layup at the countersunk hole proved highly effective in achieving satisfactory joint strength and energy absorption.

Effect of cold driving process on fatigue life of helicopter fuselage joints

Cold driving is a common procedure in aeronautical structures in presence of riveted holes; this process has influence on the fatigue life of the whole riveted joint. Using detailed FE models comprehensive of the rivet forming operation, the effect of the squeeze force on fatigue life of the joint has been investigated. Different multiaxial fatigue criteria (Sines, Crossland, KakunoKavada and Papadopoulos) have been used in the fatigue evaluations in order to consider the complete stress state of the riveted hole (comprehensive of the residual stress field). The numerical analyses have been supported by an experimental test carried out on a panel specimen from rotorcraft frame.

Automated rivet detection in the EOL image for aircraft lap joints inspection

An enhanced visual inspection technique named Edge of Light TM (EOL) was invented and developed at the Institute for Aerospace Research, National Research Council Canada. One application is the detection and measurement of “pillowing” deformation caused by corrosion in the faying surfaces of riveted aircraft fuselage joints. To quantify the hidden corrosion products, one solution is to relate the average corrosion in a region to the peak-to-peak amplitude of the EOL image brightness profile between two diagonally adjacent rivet centers. This raises the requirement to automatically detect and locate the rivets in an EOL image. The variation in illumination over individual rivets degrades the performance of most of the edge detection algorithms. In this study, a phase congruency algorithm is employed to provide an absolute measurement of EOL image features. The detected feature together with the knowledge of rivet size and the number of rows are used to locate rivets in the EOL image. The performance of the approach is compared with commonly used circle detection algorithms and an adaptive threshold approach. The efficiency of the proposed algorithm is assessed experimentally and the results are used in the quantitative evaluation of aircraft hidden corrosion.

Detection and monitoring of hidden fatigue crack growth using a built-in piezoelectricsensor/actuator network: II. Validation using riveted joints and repair patches

A built-in diagnostic technique for monitoring hidden fatigue crack growth in aircraftstructures has been developed in part I of the study. The technique uses diagnostics signals,generated from nearby piezoelectric actuators built into the structures, to detect crackgrowth. In this second part of the study, the proposed diagnostic technique was applied tomonitor fatigue crack growth in riveted fuselage joints and a cracked metallic plate repairedwith a bonded composite patch. A complete built-in diagnostic system for the tests wasdeveloped, including a sensor network, hardware and the diagnostic software. Predictionswere correlated quite well with measurements from the eddy current test and theultrasonic scan methods as well as visual inspection. The damage index successfullydetected both crack growth and debond damage for the structures considered.