Composite Aircraft Fuselage Research Articles

Design of a crash energy absorber for a composite aircraft fuselage using a combined analytical–numerical approach

Design of a crash energy absorber for a composite aircraft fuselage using a combined analytical–numerical approach

Stress fields at skin-stringer junctions in composite aircraft fuselages

This paper presents a semi-analytical analysis approach for the determination of stress fields in the vicinity of skin-stringer junctions in stiffened composite panels. The situation considered in this paper is representative for a typical stiffened panel in a modern composite aircraft fuselage. The analysis method employs a two-tier approach employing a global model based on CLPT on the one hand, and a local approach on the other hand in the form of a layerwise displacement formulation. This allows for the detailed computation of the stress concentrations in the vicinity of the skin-stringer junction. The layerwise formulation utilizes a discretization of the laminate layers into mathematical layers. The principle of the minimum of the total elastic potential yields the governing equations of the given problem, and an exponential approach leads to a quadratic eigenvalue problem that can be solved numerically. The analysis method shows excellent accuracy of the stress results in comparison with comparative finite element computations at a fraction of the computational time and effort that is required for numerical analyses.

Influence of Composite Skin on the Energy Absorption Characteristics of an Aircraft Fuselage

In aviation history, accidents during landing are more frequent compared to other phases of flight and in most cases the aircraft is forced to make an emergency belly landing when landing gear malfunctions. This study aims to analyse the energy absorption of a typical composite aircraft fuselage section under such crash belly landing. A finite element model of this composite fuselage section was created and analyzed using LS-Dyna finite element software and the effects of different composite materials on the energy absorption of fuselage section were studied. Four different composite materials that are commonly used in aerospace application were selected for the purpose of the present study namely glass/epoxy, graphite/epoxy, Kevlar/epoxy, and boron/epoxy. An impact velocity of 7 m/s against rigid ground was considered. Results showed that frames and skin contribute the most in energy absorption while the passenger floor and strut contribute the least. Moreover, it was found that fuselage skin made of glass/epoxy absorbed more energy compared to other materials and it had the highest specific energy absorption.

Minimizing stress concentrations using steered fiberpaths and incorporating realistic manufacturing signatures

Robotic Automated Fiber Placement (RAFP) technology has revolutionized manufacturing large fiber reinforced composite structures. Among others, the notable advantage of RAFP is that the designer is no more restricted to the use of straight fiber paths, which means it is possible to have curvilinear fiber paths designed for optimizing structural performance. In this paper, robotic manufacturing parameters and associated constraints imposed by these parameters on the mathematical design space is explicitly identified, and these artifacts are introduced in an optimization framework. The solution to an example problem to minimize the Stress Concentration Factor (SCF) around a window cutout typical of a composite aircraft fuselage is presented. Realistic manufacturing induced deviations like gaps and overlaps during the RAFP process is explicitly accounted for in the mathematical framework, thereby producing realizable designs that can be manufactured by a RAFP machine. These unavoidable RAFP manufacturing induced effects have not been incorporated in a mathematical framework as has been presented here.

Allowables Evaluation for the Design of a Thermoplastic Fuselage

Carbon fibre reinforced polymers are commonly used in the primary and secondary structures of aircraft, mainly due to their excellent specific mechanical properties. Currently, there is a growing interest in the use of carbon fibre and thermoplastic resin composite materials in aeronautical applications, since they offer multiple advantages compared to thermosetting equivalents. This study presents the results of the mechanical tests carried out at the coupon and element levels for the characterization of the properties of a thermoplastic composite that will form a fuselage panel. In addition to standardized tests, tests are carried out to evaluate the behaviour of critical areas of the panel, such as the radius of the frames, the stringers, and the joint between the frame and the skin around the stringers (mouse-hole). In these latest tests, experimental procedures are developed for their execution, including the design of the necessary tools to reproduce the representative boundary conditions. As a result of this research, the allowables have been obtained for the design of the curved thermoplastic fuselage panel. This work is developed as part of the European project DELTA, within the framework of the Clean Sky 2 JU program. The objective of this project is to develop and execute innovative experimental test procedures that allow the validation of thermoplastic composite aircraft fuselage panels, in order to apply these solutions in medium and long-range aircraft (LPA). The ultimate goal of this project will be achieved through the execution of a curved thermoplastic fuselage panel test under representative load conditions.

Efficient uncertainty quantification and management in the early stage design of composite applications

One of the key enablers of valued early decision-making in composite material designs is the ability to account for different aspects/scale of uncertainties within models and design processes. In order to enable well-judged decisions and to improve the trust of industrial decision-makers, measures of uncertainty, risk, and cost involved in materials are crucial. In this work, we provide i) efficient uncertainty analysis (UA) and ii) sensitivity analysis (SA) in composite structures to quantify the influence of input parameters on the output of interest accounting for the stochastic nature in multi-scale modeling with a large number of uncertain parameters. UA also provides a statistical distribution of the output of interest. The influences of different input parameters on the system responses can be estimated by conducting the global sensitivity analysis on the multi-scale models. To this end, a data-driven model approximating the relationship between the inputs and outputs is constructed by using an adaptive Sparse Polynomial Chaos Expansion (SPCE) approach. The sensitivities of the input factors on the system performances are computed analytically from the constructed data-driven model without any additional computational cost. To demonstrate the sensitivity and uncertainty management, two different test cases (composite leafspring and aircraft fuselage airframe) are considered for the stochastic structural analysis in multi-scale composite modeling. In both cases, the combined effects of multi-scale uncertainties are evaluated on the structural performances. Input parameters include the material microstructure (micro-scale uncertainties), composite layers stacking sequence (mesoscale uncertainties), and structural loading (uncertainty in macro-scale). The responses are provided in terms of their variations and probability distributions.

Composite Parts Assembly Operational Improvements

AbstractIn the era of Industry 4.0, automation plays a key role in improving manufacturing production processes in terms of working times, accuracy of operations, products’ quality, and so on. With focus on the working tasks on composite components, drilling and riveting operations require particular attention due to the possibility of damage to the final product, especially when they are executed manually. To reduce these kinds of risks, the introduction of a robot in the working area could represent an improvement to speed up the processes and to ensure a better quality result. It is worth noting that human intervention remains necessary for carrying out a part of the sub‐tasks, especially in the case of composite assembly procedures. This paper aims at presenting a case study about the drilling and riveting operations on a composite aircraft fuselage whose working cell is characterized by the coexistence of a human and a robot. Moreover, the implementation of a numerical simulation allows validating the working cycle and the adopted solutions in terms of ergonomic performance.

Robust non-destructive inspection of composite aerospace structures by extraction of ultrasonic guided-wave transfer function in single-input dual-output scanning systems

Ultrasonic guided-wave testing is widely utilized for damage detection in plate-like structural components, including composite aircraft fuselage and wing panels. Many guided-wave tests involve transducer scanning to cover finite areas, a task that is most effectively performed by non-contact wave transduction means. The most common guided-wave test implementation consists of a “single-input single-output” scheme. The single-input single-output scheme leads to a transfer function that is convolved with the particular frequency response of the transducers and that of the transducer-to-structure paths (in both excitation and detection). These responses can be unknown or generally variable, especially in non-contact scanning systems and impact excitations. This article proposes a “single-input dual-output” scheme for ultrasonic guided-wave testing in scanning systems that is based on a deconvolution operation. The single-input dual-output scheme better isolates the structural transfer function that is the only property affected by the presence of possible damage. The single-input dual-output scheme was applied to two guided-wave scanning systems under development to detect impact-type damage in stiffened skin-to-stringer panels representative of modern composite aircraft construction. The results demonstrate the dual-output technique and also shed some light on the role of the different frequency bands for the detection of damage at different locations of the skin-to-stringer assembly.

A simplified model for structural stiffness and crashworthiness optimisation of composite fuselages

In this paper, a multi-objective optimisation for the initial design of crashworthy composite structures is proposed. The focus is on the optimisation model with its minimisation of inhomogeneous deformation capability. It promotes homogeneous contribution of all crash elements, as a representative of structural crashworthiness. For the first time, homogeneous contribution has been identified as a metric for crashworthiness of composite structures and been transferred into a mathematical expression. The structural model uses discrete elements for very efficient computation in combination with the genetic algorithm NSGA-II. Clustering as a machine learning technique is applied to the Pareto set of solutions in order to identify representative structural solutions. The approach uses positioning of elements and the shape of the spring characteristic of these elements as variables. The method enables the simultaneous consideration of static and crash loads, which is demonstrated by a case study featuring a composite aircraft fuselage substructure. So far, static and crash loads have been only considered separately, but never at the same time. The novelty of this approach is in the combination of an appropriate simplified modelling technique and a new formulation of the optimisation model. The proposed optimisation is beneficial in improving the crashworthiness of composites, as optimisation of the geometry and material behaviour enables a non-linear response to be obtained in an otherwise brittle material.

Evaluation of a hybrid TMM-SEA method for prediction of sound transmission loss through mechanically coupled aircraft double-walls

This paper presents a detailed experimental validation of a hybrid TMM (transfer matrix method)/SEA (statistical energy analysis) method to predict the sound transmission through mechanically coupled aircraft double-walls. The accuracy of the model is investigated using lightweight structures representative of a composite aircraft fuselage cross-section over a large frequency band. The studied structure is composed of a thick composite sandwich panel representative of primary structure, lined with an acoustic insulation layer (glass wool), and structurally connected via vibration isolators to a thin composite sandwich lining panel representative of an interior trim panel. Three types of vibration isolators are considered: a soft, stiff and rigid mechanical links. Results show that the transfer matrix model is able to predict changes in the overall transmission caused by different structural couplings. Modelling limitation is mainly traced to the uncertainties in the measurement of the dynamic stiffness of vibration isolators which are directly propagated to the coupling loss factor and, therefore, to the transmission loss.

On the applicability of acoustic emission to identify modes of damage in full-scale composite fuselage structures

The acoustic emission method was applied during the testing of six full-scale sandwich composite aircraft fuselage panels containing through-the-thickness notches. The panels were subjected to different combinations of quasi-static internal pressure, the corresponding hoop loads, and longitudinal loads. The applicability of conventional acoustic emission signal feature analysis to identify the dominant modes of failure and extraneous emission in large composite structures was investigated. It was concluded that no clear distinction could be made among the different failure mechanisms based on the conventional acoustic emission signal features alone. Further, emission generated by fretting, either among fracture surfaces or of loading fixtures, has acoustic emission signal waveform features that are similar to those of damage-generated emission signals.

Analysis of the deformation energy of straight composite z-frames subjected to critical buckling loads

In this paper the correlation of deformation energy and critical buckling loads is examined. An advanced analysis technique is derived for the investigation of z-section frames typically applied in composite aircraft fuselage. This technique is verified through closed form solutions of z-section columns under quasi-static loading conditions. The method is applied to study the effects of flange width and laminate thickness on the critical failure modes of composite frames. Therein, the focus is on composite structural elements favourable for deformation kinematics in aerospace applications subjected to crash loads.

AFP mandrel development for composite aircraft fuselage skin

Automatic fiber placement (AFP) has become a popular processing technique for composites in the aerospace industry, due to its ability to place prepregs or tapes precisely in the exact position when complex parts are being manufactured. This paper presents the design, analysis, and manufacture of an AFP mandrel for composite aircraft fuselage skin fabrication. According to the design requirements, an AFP mandrel was developed and a numerical study was performed through the finite element method. Linear static load analyses were performed considering the mandrel structure self-weight and a 2940 N load from the AFP machine head. Modal analysis was also performed to determine the mandrel’s natural frequencies. These analyses confirmed that the proposed mandrel meets the design requirements. A prototype mandrel was then manufactured and used to fabricate a composite fuselage skin. Material load tests were conducted on the AFP fuselage skin curved laminates, equivalent flat AFP, and hand layup laminates. The flat AFP and hand layup laminates showed almost identical strength results in tension and compression. Compared to hand layup, the flat AFP laminate modulus was 5.2% higher in tension and 12.6% lower in compression. The AFP curved laminates had an ultimate compressive strength of 1.6% to 8.7% higher than flat laminates. The FEM simulation predicted strengths were 4% higher in tension and 11% higher in compression than the flat laminate test results.

Damage Tolerance of Pre-Stressed Composite Panels Under Impact Loads

An experimental test campaign studied the structural integrity of carbon fibre/epoxy panels preloaded in tension or compression then subjected to gas gun impact tests causing significant damage. The test programme used representative composite aircraft fuselage panels composed of aerospace carbon fibre toughened epoxy prepreg laminates. Preload levels in tension were representative of design limit loads for fuselage panels of this size, and maximum compression preloads were in the post-buckle region. Two main impact scenarios were considered: notch damage from a 12 mm steel cube projectile, at velocities in the range 93–136 m/s; blunt impact damage from 25 mm diameter glass balls, at velocities 64–86 m/s. The combined influence of preload and impact damage on panel residual strengths was measured and results analysed in the context of damage tolerance requirements for composite aircraft panels. The tests showed structural integrity well above design limit loads for composite panels preloaded in tension and compression with visible notch impact damage from hard body impact tests. However, blunt impact tests on buckled compression loaded panels caused large delamination damage regions which lowered plate bending stiffness and reduced significantly compression strengths in buckling.

The use of topology optimisation in the conceptual design of next generation lattice composite aircraft fuselage structures

AbstractConventional commercial aircraft fuselages use all-aluminium semi-monocoque structures where the skin carries the external loads, the internal fuselage pressurisation and is strengthen using frames and stringers. Environmental and economic issues force aircraft designers to minimise weight and costs to keep air transport competitive and safe. But as metal designs have reached a high degree of perfection, extraordinary weight and cost savings are unlikely in the future. Carbon composite materials combined with lattice structures and the use of topology optimisation have the potential to offer such weight reductions. The EU FP7 project Advanced Lattice Structures for Composite Airframes (ALaSCA) was started to investigate this. This article present some of this research which has now led to the development of a new airframe concept which consists of: a load carrying inner skin; transverse frames; CFRP-metal hybrid stiffeners helically arranged in a grid configuration; insulating foam and an additional aerodynamic outer skin.

Dynamic testing and modelling of composite fuselage frames and fasteners for aircraft crash simulations

In crash simulations of composite aircraft fuselage sections, frame breaking, skin bending and failure of mechanically fastened joints can typically be identified as major contributors to crash energy absorption. In order to generate a database for model validations, corresponding static and dynamic tests have been performed on coupon and structural element level to characterise the rate-dependent failure behaviour and energy absorption. Skin-bending, frame-bending and joint-failure tests under pull-out, bearing and peeling loads were performed on 977-2/HTS carbon fibre/epoxy specimens. On the one hand, effects of loading rate on the frame-bending behaviour could be observed. On the other hand, fastener failure did not appear to be depending on loading rate for the test speeds up to 10 m/s involved in this study. Adequate modelling methods in Abaqus/Explicit were derived and validated, and finally applied to a global aircraft crash simulation model.

Effect of thickness and laminate taper on the stiffness, strength and secondary bending of single-lap, single-bolt countersunk composite joints

This paper presents an experimental investigation of single-lap, single-bolt countersunk composite joints. Issues related to design of mechanically fastened joints in composite aircraft fuselage structures are studied, such as the effects of joint thickness, laminate taper (i.e. the positioning of the plies in the taper region) and secondary bending on joint stiffness and strength. It is demonstrated that bearing strength measurement is subject to a scaling effect due to joint thickness and that this needs to be considered in design. Tapered joints display only moderate differences in joint stiffness and bearing strength when compared to similar constant thickness specimens, suggesting that significant weight savings may be achieved. Little or no difference in bearing strength occurs with change in taper type. Three-dimensional digital image correlation is used to measure out-of-plane displacement in all joints tested. It is shown that thin joints experience higher degrees of out-of-plane displacement. This has a direct effect on the final mode of failure in the joint, and explains why thin laminates have a higher tendency to fail in fastener pull-through. It is also demonstrated that taper type has a significant effect on secondary bending and hence the final mode of failure in the joint.

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.

Contribution to an improved crash design for a composite transport aircraft fuselage—development of a kinematics model and an experimental component test setup

Crashworthiness research for transport aircraft fuselage structures is important particularly with regard to the increasing ratio of carbon fibre reinforced plastic (CFRP) in primary structures. Although today’s aluminium fuselage structures offer sufficient crashworthiness purely due to the ductile behaviour of metal, the brittle behaviour of a CFRP fuselage structure implicates the need for special crash devices to avoid uncontrolled failure with little energy absorption. Hence, a specific crash design has to be developed for a CFRP fuselage structure. A numerical methodology was developed to investigate potential crash concepts on fuselage section level and to design appropriate crash devices. The essential of this so-called kinematics model is the potential to define the behaviour of structural crash devices by characteristic input curves using macro elements in the frames, vertical struts and the sub-cargo structure. By varying the load–deformation characteristics of the crash devices, different crash scenarios, in the sense of alternative crash sequences, can be defined, assessed and compared to each other. Different potential crash scenarios for narrow-body fuselage structures were analysed using this modelling approach. The final output characteristic of the macro elements in the frame structure represents the basis for an experimental investigation of energy absorbing frame bending mechanisms. A test setup for four point flexural test configurations was developed to investigate such concepts on the generic frame level. The setup was used to analyse the failure behaviour of frame components made of pure CFRP as well as hybrid CFRP/titanium laminates. Four quasi-static as well as four dynamic frame bending tests were conducted to investigate the energy absorbing failure mechanisms of such hybrid frame structures.

Preliminary design of composite fuselage structures using analytical rapid sizing methods

The increasing use of composite materials in aircraft structures aims in reducing the structural weight significantly. In order to exploit the advantages of composite materials especially within a large-scale optimization calculation, a model for a computationally efficient structural analysis needs to be developed. In this regard, algorithms need to be developed to rapidly compute stress distribution and critical loads for both strength and stability of the composite aircraft fuselage. Therefore, sizing methods for metallic and composite orthotropically stiffened fuselage structures have been reviewed. For critical load computation, fibre fracture and inter-fibre fracture need to be taken into consideration with respect to strength. Regarding stability, the critical buckling loads of skins and stringers as well as the critical crippling load need to be taken into account. The buckling of stringers often occurs after the skin buckling load is exceeded. Hence, the postbuckling behaviour needs to be analyzed and load redistributions in the postbuckling range have to be taken into consideration. These load redistributions can generally be calculated numerically using either the finite element method or the finite strip method (Mocker and Reimerdes in Compos Struct 73:237–243, 2006) as well as analytically. In order to minimize the computational time, the postbuckling behaviour of the skins regarded as composite plates is computed analytically within this work by means of a computation of effective stiffnesses for global analysis and local failure load computation. Even though the postbuckling behaviour of metallic and composite plates has been widely studied in literature, only few work has been spent on the analytical or semi-analytical derivation of methods for the common load case of combined compression and shear loading. As the preliminary design of an aircraft fuselage requires a rapid and sufficiently accurate description of the postbuckling behaviour, the postbuckling behaviour of an orthotropic composite plate under combined compression and shear loading is analytically analyzed within the present work. The derived rapid sizing method for postbuckling significantly reduces the computational time when compared to the computational time needed for a nonlinear finite element computation. In this regard, it even allows for the consideration of postbuckling behaviour within a large-scale optimization computation of complete fuselage structures.