Aircraft Skin Panels Research Articles

Multifidelity Design of Structurally Embedded Waveguides

Slotted waveguides embedded in aircraft skin panels may not support a buckling load without failure in the copper waveguide due to stress concentrations created by the corners of the slots. To create designs strong enough to avoid compressive failure before buckling instability, an initial offset in the copper waveguide may be added so that the supporting skin cover and backplane take additional load to prevent failure at the waveguide slot corners. A structural model is developed to include a gap in the eigenvalue analysis through a prestress term, and subsequently to account for the gap in an optimization procedure. Additionally, a curve fit is used to account for local buckling modes that reduce the plate buckling mode of the final panel design. These factors combine to allow for a low-fidelity model to account for the gap and local buckling without the need for a high-fidelity model using plate or contact modeling. This low-fidelity model is then run through an optimizer, producing designs that are stronger and/or lighter for a fraction of the computational cost of prior high-fidelity models.

Splicing of Glass Laminate Aluminum Reinforced Epoxy (GLARE) for Large Aircraft Skin Panels

Splicing of Glass Laminate Aluminum Reinforced Epoxy (GLARE) for Large Aircraft Skin Panels

Joints of Glass Laminate Aluminium Reinforced Epoxy by splicing for the manufacture of large-sized aircraft skin panels

The schemes for joining of Glass Laminate Aluminium Reinforced Epoxy by the splicing method and the choice of the optimal scheme for the manufacture of large-sized panels are considered. The structure, mechanical and fatigue characteristics of a large-sized panel manufactured according to the optimal scheme are studied. The use of the splicing method makes it possible to reduce the weight of the structure by 10¸15 % as a result of reducing rivet joints and eliminating stress concentrators (holes), as well as increasing the reliability and damage tolerance of the aircraft. Keywords: Glass Laminate Aluminium Reinforced Epoxy, joint, splicing method, large panel, fuselage skin, fuselage, metalpolymer composite materials. org80@viam.ru, nefedovajuls@inbox.ru

Virtual shimming simulation for smart assembly of aircraft skin panels based on a physics-driven digital twin

A leading challenge in the assembly process of aircraft skin panels is the precise control of part-to-part gaps to avoid excessive pre-tensions of the fastening element which, if exceeded, impair the durability and the response under dynamics loads of the whole skin assembly. The current practice is to measure the gap in specific points of the assembly with parts already at their final location, and then be-spoke shims are machined and inserted between the mating components to fill the gap. This process involves several manual measurement-fit-adjust quality loops, such as loading parts on the assembly frame, measuring gaps, off-loading parts, adding be-spoke shims and re-positioning parts ready for the fastening operation—as a matter of fact, the aircraft is re-assembled at least twice and therefore the current practice has been proved highly cost and time ineffective. Additionally, the gap measurement relies on manual gauges which are inaccurate and unable to follow the actual 3D profile of the gap. Taking advantage of emerging tools such as in-line measurement systems and large-scale physics-based simulations, this paper proposes a novel methodology to predict the part-to-part gap and therefore minimise the need for multiple quality loops. The methodology leverages a physics-driven digital twin model of the skin assembly process, which combines a physical domain (in-line measurements) and a digital domain (physics-based simulation). Central to the methodology is the variation model of the multi-stage assembly process via a physics-based simulation which allows to capture the inherent deformation of the panels and the propagation of variations between consecutive assembly stages. The results were demonstrated during the assembly process of a vertical stabiliser for commercial aircraft, and findings showed a significant time saving of 75% by reducing costly and time-consuming measurement-fit-adjust quality loops.

Multiscale Reduced-Order Modeling of a Titanium Skin Panel Subjected to Thermomechanical Loading

This paper presents the formulation, implementation, calibration, and application of a multiscale reduced-order model to simulate a titanium panel structure subjected to thermomechanical loading associated with high-speed flight. The formulation is based on the eigenstrain-based reduced-order homogenization model and further considers thermal strain as well as temperature-dependent material properties and evolution laws. The material microstructure (i.e., at the scale of a polycrystalline representative volume element) and underlying microstructural mechanisms are directly incorporated and fully coupled with a structural analysis, concurrently probing the response at the structural scale and the material microscale. The proposed approach was fully calibrated using a series of uniaxial tension tests of Ti-6242S at a wide range of temperatures and two different strain rates. The calibrated model is then adopted to study the response of a generic aircraft skin panel subjected to thermomechanical loading associated with high-speed flight. The analysis focuses on demonstrating the capability of the model to predict not only the structural scale response, but also simultaneously the microscale response, and it further studies the effects of temperature and texture on the response.

Materials Selection for Aircraft Skin Panels by Integrating Multiple Constraints Design with Computational Evaluations

Among the principal load-carrying parts of an aircraft, the center wing box yields the main mechanical load-carrying area and is composed of many structural components and elements such as upper and lower skin panels, internal rods, ribs, stringers, front and rear spars. For the structural analysis of such real complex systems, simplified representations of the same sections would be used alternatively to determine whether components could withstand loadings without experiencing failure, thus stay in the margin of safety. Nevertheless, these components should contribute to weight efficiency while carrying in-plane loads and distribute out-of-plane loads to in-plane members safely. Materials selection plays a crucial role for the determination of such candidate materials along with their material properties in specific structural applications. It is usually performed by considering proper objectives, constraints and free variables with respect to functions of the components of a system. Since Composite Material Handbook (CMH-17), Metallic Materials Properties Development and Standardization (MMPDS), and Preliminary Material Properties Handbook (PMP-HDBK) databases are embedded in the materials selection software (CES Edupack), best candidate materials would be easily identified by using coupling constant(s) such as in multiple-constraints designs. In this study, Ashby’s methodology was applied to determine best candidate materials for constructing skin panels by considering them as bending plates. According to materials selection approach, continuous fiber reinforced epoxy composites was stipulated as one of the best candidate materials. Computational failure analysis was then carried out by referring the proper material properties from the materials selection software.

Understanding Cold Spray for Enhanced Manufacturing Sustainability

High pressure cold spray has been showing increasing promise and application for structural repairs and coating applications where wrought like strengths are required. For example, numerous applications have been developed for repairing high cost and long lead time parts for the aerospace and defense market, such as aircraft skin panels, titanium hydraulic lines, aluminum valve actuator internal bores, hardened and chromed steel shafts, gas turbine engine parts, magnesium castings, and many more. These processes also have direct application in commercial markets like transportation and heavy industry. In particular, parts with lead times in excess of 12 months have been successfully repaired and re-introduced into service. This saves not only the direct cost of the part, but also returns the system to service much sooner. Additional benefits of field application with a hand-held nozzle assembly are also possible, particularly for power plants, refineries, and other large industrial plant operations. Cold spray consequently has a tremendous opportunity to enhance manufacturing sustainability by repairing parts that previously could only be replaced and recycled. It is environmentally friendly, as there are no toxic fumes or other harmful emissions from cold spray. Furthermore, because parts are being repaired and refurbished rather than replaced, there is tremendous cost, energy, and overall environmental benefit, making cold spray a “green” technology and an excellent technology for enhancing the long-term sustainability of high value assets.

Influence of the viscous boundary layer perturbations on single-mode panel flutter at finite Reynolds numbers

Panel flutter is an aeroelastic instability of aircraft skin panels, which can lead to a reduction in service life and panel destruction. Despite the existence of many studies related to panel flutter, the influence of the boundary layer on the panel stability has been considered in only a few of them. Up to the present day, most papers on the boundary layer effect consider only a zero-pressure-gradient boundary layer over a flat plate. The only studies of a boundary layer of arbitrary form were conducted in our previous papers (Vedeneev, J. Fluid Mech., vol. 736, 2013, pp. 216–249 and Bondarev & Vedeneev, J. Fluid Mech., vol. 802, 2016, pp. 528–552), where the boundary layer was represented as an inviscid shear layer (the Reynolds number $R=\infty$). In this paper we investigate the problem, taking viscosity into account, at large but finite Reynolds numbers. As before, we assume that the panel length is large and use Kulikovskii’s global instability criterion to analyse the panel eigenmodes and consider two different types of boundary layer profiles: a generalised convex profile and a profile with a generalised inflection point. Results show that viscous perturbations can, in general, have both stabilising and destabilising effects on the system, depending on the velocity and temperature profiles of the boundary layer and on its thickness. However, surprisingly, we prove that if the boundary layer yields a significant growth rate in the inviscid approximation, then the viscosity always produces an even larger growth rate.

Modeling and simulation of machining distortion of pre-bent aluminum alloy plate

In current commercial aviation industry sector, thin and large aircraft skin-panel is milled to create pockets on the pre-bent metal plate to reduce weight and maintain structural integrity. The distortion of the pre-bent plate is liable to occur during the machining process due to the asymmetrical distribution of the bending residual stress. In this study, the mechanism of machining distortion is analyzed by the combination of theoretical and numerical methods. The analysis model is established based on the springback analysis of the stretch bending process by considering the coupled effects of bending-induced residual stress and machining-induced stress. The numerical model is constructed by the combination of the shell element and the submodel analysis technologies. Furthermore, a series of metal plates are bent to cylindrical profile and milled sequentially to create the shallow pockets at the inner surface to verify the theoretical and simulation models. By the comparisons of the machining distortions related to the machining depth and machining-induced stress between the experimental results and predictions, it is revealed that the distribution of initial bending residual stress is emerged with unsymmetrical “N” or “Z” shape, which is different from the traditional symmetrical “M” or “∩” distribution. In addition, the distortion variation tendency is affected by the coupled effects of bending residual stress and the compressive machining-induced stress.

Experimental study on the effect of dents induced by impact on the fatigue life of 2024-T3 aluminum alloy plate

Local dent is one of the typical damages of the aircraft’s skin panels. However, to the best of our knowledge, very few studies focus on the fatigue lives of dented aluminum plates which commonly used as aircraft skin panels. In this study, drop-weight impact tests were conducted on 2024-T3 aluminum plates to obtain spherical, conical, and U-shape dents on specimens. Subsequently, fatigue tests were performed on the dented specimens to investigate the effect of the dents on specimen’s fatigue lives. It is found that all these three types of dents can obviously reduce the fatigue lives of 2024-T3 aluminum plates. Moreover, the empirical models representing relationships between residual fatigue lives and dent depths were proposed. At last, fatigue tests were carried out on specimens with both pre-cracks and dents to explore the effect of the dent on crack growth. The experimental results showed that, for the dent with medium depth relative to ultimate depth, the conical dent can aggravate the crack “macro-growth” while the spherical and U-shape dent can retard it. The observed features can provide certain references for repairing aircraft skin panels with dents.

Evolutionary-based aeroelastic tailoring of stiffened laminate composite panels in supersonic flow regime

Aircraft skin panels stiffened with attached stiffeners are commonly used in aerospace design. Recent interest on investigations of supersonic flutter of composite panels have concentrated on developments to deal with typical nonlinear aeroelastic responses. Few research on stiffened composite panels is observed in the literature, particularly on aeroelastic tailoring of laminate composites regarding supersonic flutter. This work contributes with an investigation on the aeroelastic tailoring of stiffened laminate composite panels constrained for flutter conditions, thereby allowing design approach for passive flutter avoidance. The finite element method is used to model the panel structural dynamics admitting aerodynamic coupling through the first order piston theory model. The optimization approach is based on a conventional genetic algorithm to maximize the critical flutter dynamic pressure for different arrangements of stiffened laminate composite panels. Stiffened panels were assumed to be restrained with SS1-type boundary condition or an elastic foundation of arbitrary stiffness. Optimized panels for different supersonic flow directions have shown that flutter limits can be elevated by adequate laminate orientation and by altering the flutter mechanism.

Detection and classification of structural changes using artificial immune systems and fuzzy clustering

Among all the elements that are integrated into a structural health monitoring (SHM) system, methods or strategies for damage detection and classification are nowadays playing a key role in enhancing the operational reliability of critical structures in several industrial sectors. The main contribution of this paper is the application of a new methodology to detect and classify structural changes. The methodology is based on: 1) an artificial immune system (AIS) and the notion of affinity is used for the sake of damage detection; 2) a fuzzy c-means algorithm is used for damage classification. One of the advantages of the proposed methodology is the fact that to develop and validate the strategy, a model is not needed. Additionally, and in contrast to standard Lamb waves-based methods, there is no need to directly analyse the complex time-domain traces containing overlapping, multimodal and frequency dispersive wave propagation that distorts the signals and difficult the analysis. The proposed methodology is applied to data coming from two sections of an aircraft skin panel. The results indicate that the proposed methodology is able to accurately detect damage as well as classify those damages.

Innovative NDT Technique Based on Ferrofluids for Detection of Surface Cracks

An innovative NDT technique is proposed for surface inspection of materials not necessarily magnetic or conductive, based on local magnetic field variations due to ferrofluid deposited in the cracks. The feasibility of the technique is assessed preliminarily, based on signal detectability without applied external magnetic field, and under applied DC fields. The signals (local magnetic flux density variations) are quantified analytically, experimentally and numerically. The model agrees well with the tests, showing that the signal increases with the applied field strength, up to the saturation magnetization of the ferrofluid, and decreases with the distance to the crack longitudinal axis, and thus it can provide useful estimations of the signal. The proposed technique, requiring application of external fields to magnetize the ferrofluid to enhance the signal, seems promising: the model suggests that signals associated to cracks significantly smaller than surface cracks in a target application like aircraft skin panel inspection NASA STD-5009 are easily detectable with commercial magnetometers.

Investigation of an expert health monitoring system for aeronautical structures based on pattern recognition and acousto-ultrasonics

Aeronautical structures are subjected to damage during their service raising the necessity for periodic inspection and maintenance of their components so that structural integrity and safe operation can be guaranteed. Cost reduction related to minimizing the out-of-service time of the aircraft, together with the advantages offered by real-time and safe-life service monitoring, have led to a boom in the design of inexpensive and structurally integrated transducer networks comprising actuators, sensors, signal processing units and controllers. These kinds of automated systems are normally referred to as smart structures and offer a multitude of new solutions to engineering problems and multi-functional capabilities. It is thus expected that structural health monitoring (SHM) systems will become one of the leading technologies for assessing and assuring the structural integrity of future aircraft. This study is devoted to the development and experimental investigation of an SHM methodology for the detection of damage in real scale complex aeronautical structures. The work focuses on each aspect of the SHM system and highlights the potentialities of the health monitoring technique based on acousto-ultrasonics and data-driven modelling within the concepts of sensor data fusion, feature extraction and pattern recognition. The methodology is experimentally demonstrated on an aircraft skin panel and fuselage panel for which several damage scenarios are analysed. The detection performance in both structures is quantified and presented.

Structural damage detection using principal component analysis and damage indices

One of the most important tasks in structural health monitoring corresponds to damage detection. In this task, the existence of damage should be determined. In the literature, several potentially useful techniques for damage detection can be found, and their applicability to a particular situation depends on the size of the critical damages that are admissible in the structure. Almost all of these techniques follow the same general procedure: the structure is excited using actuators, and the dynamical response is sensed at different locations throughout the structure. Any damage will change this vibrational response. The state of the structure is diagnosed by means of the processing of these data. Several studies have shown that the detection of changes in a structure depends on the distance from the damage to the actuator as well as the configuration of the sensor network. In this article, the authors considered the advantage of using an active piezoelectric system, where the lead zirconate titanate transducers are used as actuator and sensors in different actuation phases. In each actuation phase of the diagnosis procedure, one lead zirconate titanate transducer is used as actuator (a known electrical signal is applied), and the others are used as sensors (collecting the wave propagated through the structure at different points). An initial baseline model for undamaged structure is built applying principal component analysis to the data collected by several experiments and after the current structure (damaged or not) is subjected to the same experiments, and the collected data are projected into the principal component analysis models. Two of these projections and four damage indices ( T2-statistic, Q-statistic, combined index, and I2 index) by each actuation phase are used to determine the presence of damages and to distinguish between them. These indices are calculated based on the analysis of the residual data matrix to represent the variability of the data projected within the residual subspace and the new space of the principal components. To validate the approach, data from two aeronautical structures—an aircraft skin panel and an aircraft turbine blade—are used.

A Bioinspired Methodology Based on an Artificial Immune System for Damage Detection in Structural Health Monitoring

Among all the aspects that are linked to a structural health monitoring (SHM) system, algorithms, strategies, or methods for damage detection are currently playing an important role in improving the operational reliability of critical structures in several industrial sectors. This paper introduces a bioinspired strategy for the detection of structural changes using an artificial immune system (AIS) and a statistical data-driven modeling approach by means of a distributed piezoelectric active sensor network at different actuation phases. Damage detection and classification of structural changes using ultrasonic signals are traditionally performed using methods based on the time of flight. The approach followed in this paper is a data-based approach based on AIS, where sensor data fusion, feature extraction, and pattern recognition are evaluated. One of the key advantages of the proposed methodology is that the need to develop and validate a mathematical model is eliminated. The proposed methodology is applied, tested, and validated with data collected from two sections of an aircraft skin panel. The results show that the presented methodology is able to accurately detect damage.

Survey of numerical approaches to analyse the behavior of a composite skin panel during a water impact

Water impacts may have tragic consequences for the passengers of helicopters. Most of the passive safety devices developed for helicopter crashworthiness is designed for ground impact. The loading that characterizes the impact with hard and soft surfaces is different and therefore energy absorption devices developed for ground impact are not effective during a water impact.Various researches focus on the use of composite materials for aircraft and helicopter fuselage. In this paper, in particular, it is investigated the behavior of a composite panel during the impact with water and the approaches to study the event by means of finite element codes.In order to collect reliable data for numerical model validation, water impact drop tests were carried out. A sample panel, made with a Carbon Fiber Reinforced Plastic material similar to those used for modern aircraft skin panels, was manufactured. A specific test device was created and used in the tests. Impact decelerations and deformation of the panel were measured.Numerical models of the tests were created. Meshless approaches were used, in addition to Lagrangian and Eulerian Finite elements, to model the water region. Eventually, a close experimental–numerical correlation was obtained for each model in terms of impact dynamics, decelerations and composite panel deformation. The main features of the event and the differences between the four numerical approaches were discussed. Guidelines for further investigations were also drawn.

Numerical Analysis for the Multi-Point Stretch Forming Process of Aircraft Skin Panel

Multi-point stretch forming (MPSF) is a new technique to form aircraft outer skin panel. Since multi-point die is composed by the discrete punches, the result of the MPSF aircraft outer skin panel need to study in depth. The thickness of elastic cushion and free length are two important factors to affect the accuracy, and they must be chosen reasonably. A series of numerical simulations on typical MPSF processes were carried out to an aircraft outer skin panel part. The results show that the thicker the elastic cushion is, the more valid the dimple will be suppressed .The longer free length is, the smaller the equivalent strain and thinning and more uniform the distribution of thickness will be. When the free length is shorter, the degree of effect is relatively obvious on the equivalent strain and thinning and the distribution of thickness; when the free length is longer than a certain value, the degree of effect is small.

Dynamic Product Modeling with Inter-Features Associations: Comparing Customization and Automation

AbstractThe customization of CAD tools and the development of trade-oriented applications make it possible to eliminate tedious tasks of modeling that would otherwise monopolize the time of designers. This article reviews some concepts, such as entity typifying, associations, imprints and pseudo-imprints, used to associate features to form a dynamic product model. Concepts are tested in an aerospace case-study where CAD tools are used for modeling pockets in aircraft skin panels. Specific companies’ know-how can be modeled by CAD tools through either Customization or Automation. Customization is achieved by implementing the concepts through User-Defined Features (UDF). A UDF is seen as a specific aggregation of constraints established between lowlevel geometrical objects, so as to transform an imprint into a pseudo-imprint. Once defined, this aggregation of constraints is manipulated as a single entity to accelerate design work. Automation goes beyond customization by performing UDF instantiation with lim...

Finite element-based three-dimensional stress analysis of composite bonded repairs to metallic aircraft structure

The use of high-modulus composite bonded patches has been shown to be a very useful life extension strategy when used to reinforce damaged or damage-prone aircraft structural components. When applied to aircraft skin panels such as fuselage and wing skins, the patch attracts load up through the thickness of the panel as well as from the wider regions creating a stress concentration on the surface of the panel at the ends of the patch. The main objective of this work is the numerical evaluation of this stress concentration for a single layer patch, a multi-layer equivalent-stiffness patch with uniform taper stepping, and a multi-layer patch with non-uniform taper stepping based on shape optimisation principles. Modelling of the residual thermal effects caused by heat curing of dissimilar materials has been included, as well as comparison of the through-thickness and in-plane 2D results with the 3D results.