Composite fuselage section Wafer Design Approach for Safety Increasing in Worst Case Situations and Joints Minimizing

Final Report Summary - WASIS (Composite fuselage section Wafer Design Approach for Safety Increasing in Worst Case Situations and Joints Minimizing)

Executive Summary:
This document collects a description of the process followed to design, manufacture, test and validate a composite fuselage section, according to a lattice structure, aiming at replacing conventional aluminium structures used in the aeronautical sector (in this case a business jet is taken as pilot case). Design is guided by a set of objectives like weight reduction, safety performance increase with respect to metallic structures and manufacturing cost feasibility. Besides, manufacturing restrictions have been observed.
Although the final aim is to achieve a fuselage section made of composite material, instead of developing a real scale component, a building block approach was implemented, starting from material samples, going through so called level 2, 3 and 4 specimens (samples featuring specific aspects like intersections between ribs, joining strength between composite and metallic interfaces, joining strength between ribs and skin…). Final validation is performed on panels featuring sections of the full scale fuselage sections, as well as downscaled fuselage sections. Three manufacturing methods are studied, automatic tape placement, semi-automatic filament winding and filament winding.
Besides obvious structural and safety requirements, further secondary aspects were assessed like reparability of the composite structure in case of small damages and prediction of the vibroacoustic performance.
The document is organised in three main chapters. Chapter one collects design aspects, chapter two tooling and manufacturing information, and chapter three testing and validation of simulation models used during the design phase.

Project Context and Objectives:
1.1 Wafer Fuselage Section with Micro Fasteners Joint Design
1.1.1 Definition of design requirements
The compliance with general standards of the structure started with the definition of the operational load cases as well as the requirements and evaluation criteria of the structure’s performance. Initially, operational load cases the fuselage structure has to comply with as well as the requirements and evaluation criteria of the structure’s performance were defined by the Consortium end users, PAI and CORVUS, based on their experience in aeronautic structures design and manufacturing in compliance with appropriate European standards and regulations. After CORVUS leave, PAI fully overtook the definition of the requirements.
1.2 Wafer Section Safety Assessment
1.2.1 Worst case scenarios for external loading
Numerical Modelling of Impact loads: Hail
The results of simulation work are summarised in the following paragraphs. UoP has developed a dynamic simulation model of the behaviour of the aluminium fuselage section.
UoP used the CAD models provided by PAI, and the software LS-Dyna to carry out the analysis. Regarding the composite wafer structure, the dimensions were established by KhAI as 1800x2000 mm. Regarding the skin of the composite structure +-/45 layup of HTS5631/MTM44-1 which was also used for the stress analysis and dimensioning. The total thickness of the skin used was 1.184 mm.

2 Manufacturing
2.1 Wafer Section Manufacturing Design
In order to sustain the micropins, the designers of the structure and the joints need to know a series of material parameters that are not standard, and thus not available in usual material databases or material supplier’s databases. In order to overcome this need, material non standard tests were identified and planned.
The first phase of this task was identifying the tests to be carried out. KhAI, as design leader, exposed that two different set of parameters were needed in order to assure that the micro pins and reels designs were correct. These two sets of parameters were:
• Fibre bending properties when bending along different radiae with a filamente winding representative tension
• Composite metal adhesion properties, to estimate the shear properties of an hybrid joint.
Two different tests were designed by Cidaut in agreement with KhAI and with NET.
2.2 Scaled down prototypes manufacturing
2.2.1 Wafer test panels
a. Panels manufactured at IVW
b. Short overview about the activities from INEGI
c. Short overview about the activities from KHAI
3 Validation

3.1 Validation tests
3.1.1 Building Block approach and tests planning
In the case that composite materials are intended for use in structural components, a design development programme is generally initiated during which the performance of the structure is assessed prior to use. The process of validating the structural performance and durability of composite structural components consists of a complex mix of testing and analysis. Testing alone can be excessively expensive because of the vast number of specimens needed to verify every geometry, loading condition, environment, and failure mode. Analysis techniques alone are not usually sophisticated enough to adequately predict results under every set of conditions. By combining testing and analysis, analytical predictions are verified by tests, test plans are guided by analysis, and the cost of the overall effort is reduced while reliability is increased.
3.1.2 Static performance of tested probes
Grid stiffened structures or advanced grid stiffened (AGS) structures are shells supported by a grid lattice of stiffeners. They were the subject of many researches for the past many decades as a possible replacement of monocoque, skin-stringer, and honeycomb sandwich structures. In the past decade these arrangements have finally become an option for using in several different structures with cylindrical shapes such as airplane fuselage.
3.1.3 Fuselage panels performance
Compression tests were performed on the full scale panel to define its response to the loading and study the associated failures. Deformation in different parts of the specimen was quantified by the use of strain gauges.
b. Loading
c. Deformation
Several strain gauges were positioned on different locations of the panel to monitor and record the local deformations caused by the compressive load.
d. Failure
The load on the fuselage panel increased until failure occurred at around 150 kN.
3.1.4 Fatigue and Damage Tolerance results
Grid stiffened structures or advanced grid stiffened (AGS) structures are shells supported by a grid lattice of stiffeners. They were the subject of many researches for the past many decades as a possible replacement of monocoque, skin-stringer, and honeycomb sandwich structures. In the past decade these arrangements have finally become an option for using in several different structures with cylindrical shapes such as airplane fuselage.
3.1.5 Static test results on larger scale prototype
The fuselage prototype section manufactured by CirComp was tested under bending in order to extract load-strain data for validation of FE models of the structure and the performance under load. Prior to test the part was strain gauged to enable measurement of the strain of different sections under the loading. In addition to strain gauges, the Digital Image Correlation (DIC) technique was used to collect more deformation data from the top side of the prototype during the test. The bending load was applied on the part using two hydraulic actuators and via a specially designed fixture. The strain data was recorded using a custom build LabView program.
3.2 Simulation models validation, correlation FEA vs. experiment
A typical geometrically repetitive element of the proposed fuselage wafer structure is the intersection of spiral and hoop ribs.
3.3 Secondary requirements
3.3.1 Design Implications from the Reparability point of view
This report describes the a methodology developed for repairing wafer panels for different damage modes. Two damage modes were repaired during this procedure, skin/rib debonding and skin damage as a result from impact. The process of repairing is described and the results from compression tests after the repair are shown.
3.3.2 Vibroacoustic prediction of composite wafer structures

Project Results:
1.1 Wafer Fuselage Section with Micro Fasteners Joint Design
1.1.1 Definition of design requirements
g. Conclusions
According to project plan and together with project partners influence of structural arrangement of fuselage section including reels quantity, location and all possible load-carrying schemes of regular wafer structure including hoop-spiral, longitudinal-spiral and spiral-spiral ones were analyzed.

Taking into account manufacturing and structural restrictions the most rational hoop-spiral system of ribs was selected for further analysis.

Geometrical and weight parameters of hoop, spiral ribs and skin were determined analytically and comparative and checking analysis of suggested load-carrying schemes was conducted using FE method.

Several contradictive approaches to developing similarity criterion for design of 1m and 0.5m prototypes were suggested.

Preliminary design of prototypes was conducted and recommended geometrical parameters of prototype’s ribs and external loads for further testing at bending were calculated.

Developed structural parameters of prototypes, some manufacturing and testing recommendation were worked out for following prototypes manufacturing and testing.
A number of optimized designs were shown in ¡Error! No se encuentra el origen de la referencia.. As can be seen, multiple of those comply with the requirements. In reality, the compliant design pool is much bigger hence a lot more configuration could potentially be chosen. Considering the fact that the optimization process was not finalized due to time related issues, upon further optimizations a bigger design pool of compliant designs could be obtained potentially with a better design fitness (lower weight). This design pool also offers the possibility to include the reel restriction/equality requirement, it could also be introduced as a constraint in the optimization process.
Using the current findings, the lowest fuselage barrel weight obtained so far is 40.48 kg, which is significantly lower than the weight of the aluminium reference structure corresponding to the current design of the Piaggio P180 aircraft (62.55kg). This shows the weight saving potential of the lattice composite application for aircraft fuselage applications. Thus using OOA composite materials and a lattice architecture a 54.5% weight saving is possible with respect to the aluminium reference design.
1.2 Wafer Section Safety Assessment
1.2.1 Worst case scenarios for external loading
m. Conclusions
The material properties of carbon fibre composite specimens produced by the ATP and FW processes were evaluated. For the ATP there was no issue and the exact same process that would be followed for the fuselage panel fabrication could be used to produce flat panels from which standard specimens were extracted. The resulting material was of high quality and yielded repeatable material property data. The difficulty with FW arises from the fact that tubular specimens (i.e. replicating the exact manufacturing method) can only generate material property information transverse to the fibre direction. For properties along the fibre direction it was attempted to wound on a flat mandrel and extract standard size and type specimens. This proved unsuccessful since the resulting quality of the material was very low. It can be concluded that since the two material systems use the same type of fibres it would be safe to use the ATP material properties to derive the necessary material design allowables.

2 Manufacturing
2.1 Wafer Section Manufacturing Design
l. Results and Discussion
A summary of the results obtained within this test programme are presented in Figure 2.104 while all experimentally obtained data can be found in the Appendix. Some solid conclusions can be derived from the data in Figure 2.104:
• Producing flat panels by wet winding, to replicate the manufacturing process and at the same time keep the specimen preparation and testing simple, yielded inconsistent material with large variation in the fibre volume fraction and void content. The measured material properties from these panels where characteristically low even along the fibre direction (i.e. tensile strength below 2000 MPa), which could only be justified by the poor consolidation achieved during manufacture. Overall, modulus values measured were acceptable, however the produced strength data cannot be utilised.
• Producing tubular specimens by wet filament winding resulted in very consistent high quality material with exceptionally low void content and high fibre volume fraction (i.e. 58%). Because of the shape of the mandrel the fibre compaction during winding was high and hence the quality of the resulting product and specimens exceptionally good. The material properties obtained from these specimens would possibly form the upper (i.e. higher) boundary of the performance of the material as in the actual application the curvature of the component would be significantly larger and therefore the fibre compaction fairly lower.
• The ATP process could produce flat panels replicating the exact procedure as it would be followed in the fuselage panel manufacturing. The material was of high quality with non-existent void content and significantly high fibre volume fraction. The material properties presented a good degree of consistency. It should be noted that a direct explicit comparison between the ATP and FW materials couldn’t be made as the resin of the two materials differs. Fibre dominated properties should be very comparable.
m. Conclusions
The material properties of carbon fibre composite specimens produced by the ATP and FW processes were evaluated. For the ATP there was no issue and the exact same process that would be followed for the fuselage panel fabrication could be used to produce flat panels from which standard specimens were extracted. The resulting material was of high quality and yielded repeatable material property data. The difficulty with FW arises from the fact that tubular specimens (i.e. replicating the exact manufacturing method) can only generate material property information transverse to the fibre direction. For properties along the fibre direction it was attempted to wound on a flat mandrel and extract standard size and type specimens. This proved unsuccessful since the resulting quality of the material was very low. It can be concluded that since the two material systems use the same type of fibres it would be safe to use the ATP material properties to derive the necessary material design allowables.

3.1 Validation tests
3.1.1 Building Block approach and tests planning
e. Conclusions
A building block programme to support the development of the integrated lattice structure within WASIS project has been formulated. The particular limitations regarding materials and processes as well as timing of the development are considered. It is anticipated that the testing programme will yield information of considerable importance for the validation of the proposed design and the preferred manufacturing method.
3.1.2 Static performance of tested probes
h. Conclusions
A large number of FW and ATP produced specimens was tested in order to investigate the performance of the produced lattice structural elements and quantify their strength for design considerations. In general the quality of the FW structural components was vastly better that those produced by the ATP method. This was somehow expected as the ATP approach for producing lattice structural components is currently at very low technology readiness level.
Specimens that didn’t contain any skin failed by local compression of the ribs which was a follow up of variable degree of instability (buckling). This supports the necessity of using a skin as a means of reducing the local buckling of the ribs.
When skin was present, and excluding any local buckling at the skin edges, the main failure modes were (a) compressive failure of the ribs for the FW scaled down specimens and (b) skin-to-rib debonding for the full scale ATP specimens. This finding highlights the limitations in using scaled down specimens for investigating the mechanical performance of the structure, because the scaling was only applied to the ribs and not the skin. This resulted in lower stiffness mismatch between the two (i.e. ribs and skin) and diversion of the failure to the ribs themselves. In the case of the full scale specimens, where the stiffness mismatch is much higher (i.e. same skin thickness but larger ribs), the failure was clearly driven by the skin-to-rib debonding, exposing the main disadvantage of the current design.
For the attachment frame details which didn’t contain any skin, failure was clearly driven by pull-out of the reel and strength was dependant on the metallic reel surface quality and resin quantity available to form a good bond between the composite and the metal part. In the presence of the skin and full metallic attachment frame, failure is again dominated by skin-to- rib debonding.
Finally, the use of micro-pin joints for floor attachments was investigated by applying the worst case scenario loading i.e. direct tensile load. This resulted in very poor performance with pull-out of the co-cured micro-pins from the composite. The clean surfaces of the metallic parts (incl. the micro-pins) suggest the that laminating resin is obviously not enough to provide the necessary adhesion. This can be addressed by incorporating another structural adhesive in the manufacturing process and optimise the surface treatment of the metallic parts. As a comparison adhesively bonded, using an aerospace qualified adhesive, floor attachments where tested. These presented significantly better performance. In this case the performance was limited by the fracture toughness of the composite, since failure was always related to top ply composite delamination.
3.1.3 Fuselage panels performance
e. Conclusion

Following the structural test of the full scale fuselage panel it can be concluded that the skin-to-ribs debonding is the first type of failure caused by the compression loading applied. It was also evident that no damage was recorded in the ribs as a result of compression loading. Hence it is advisable to improve the bonding between the skin and the ribs to directly improve the mechanical performance of the structure under compression.
3.1.4 Fatigue and Damage Tolerance results
g. Conclusion
Following the fatigue tests undertaken in this work programme the following conclusions can be derived:
• The failure of the Attachment Frame without Skin specimens is dominated by the strength of the bond between the metallic reel and the composite and not by the strength of the composite itself. The large concentration of voids and the insufficient amount of resin close to the centre of the reel contribute significantly to early damage initiation and subsequent failure of the part. It is recommended that another structural gap filling adhesive should be used prior to the application of the composite and possibly an increase in the reel spacing that would allow easier compaction of the composite.
• For the Attachment Frame with Skin specimens the debonding of the skin from the metallic attachment frame (i.e. damage initiation) occurred at a very low cycle count and well before the final failure of the specimen, which was debonding of the skin from ribs. The type of failure under fatigue highlights the limitations of the current design with the very low skin-to-rib crack growth resistance and fairly high stiffness mismatch.
• The Ribs with Skin specimens failed in the same way as when tested under quasi-static loading. Failure initiated from the rib intersections and propagated along the ribs. Secondary failures of skin-to-rib debonding at the edges were artefacts of the structural elements and the free edges rather than driven by the design. Again the scaled down specimen could not capture the skin-to-rib debonding as the stiffness mismatch was much less than the full scale component. However, a very interesting finding is the damage tolerant nature of the lattice structural design with very limited influence on the fatigue performance even after the introduction of substantial impact damage on the rib intersection or the skin. Even for these specimens damaged initiated and propagated from the same locations and after the same number of fatigue cycles as with the undamaged specimens.
3.1.5 Static test results on larger scale prototype
i. Conclusions
The fuselage section was successfully tested without producing any damage in the structure as no residual strain seen in the part after the loading-unloading sequence
The strain measurement from strain gauges and the DIC are very similar which verifies the consistency of the techniques employed.
3.2 Simulation models validation, correlation FEA vs. experiment
3.2.5 Conclusions
A correlation of finite element analysis with experimental reuslts as performed on three levels of the building block approach. Several options for modelling composite materials by the FE method were investigated and the results were compared with physical test data.
On Level 2, two types of structural elements were tested in compression and modelled by FEA. The structural element was a typical intersection of the ribs extracted from the scaled down version of the lattice stiffened fuselage structure in two variants; with and without the skin. A good correlation was obtained for the “skin and ribs” component, modelled either by conventional elements or by 3D solid elements and user subroutine. The “Ribs only” component has proven to be more sensitive to initial geometric imperfections and better prediction was achieved by 3D solid elements.
On Level 3, a full-scale fuselage section panel was tested in compression. The models comprised two different approaches previously used at Level 2; the skin was modelled by conventional shell elements and orthotropic material properties; ribs were modelled by 3D solid tetrahedrons and a user defined material subroutine. The correlation of load-strain achieved by the models was good, even though some level of uncertainty arises from manufacturing imperfections. The model was not capable to predict ultimate failure of the structure.
On Level 4, a scaled down fuselage section was tested in bending. Conventional shell elements were used in the model and the load-strain results are in a very good agreement with the experimentally obtained data. Also the strain map obtained by DIC correlates well with the simulation results. Whether the model would be able to predict the ultimate failure cannot be elaborated, as the fuselage section was not tested up to this point.
Finally, a full scale fuselage section under realistic load conditions was further analysed at several levels of detail. Global models were used to obtain stresses and failure indices of the whole structure and to drive boundary conditions of the local model. The failure indices shown here suggest that the structure would be able to withstand the imposed loading, however, no experimental data are available for validating such prediction.
3.3 Secondary requirements

4 Conclusions
Along the whole project duration the consortium has accomplished the entire implementation of developing a fuselage section made of CFRP using affordable manufacturing methods. Step by step the original idea has gone through design, development and validation.
Early in the project a set of materials were identified as best appropriate for the fuselage section application. Depending on the manufacturing method, hot cure resin “Araldite® LY 556 / Aradur® 917 / Accelerator DY 070“ was used for wet winding, together with HTS40 (12K)-F13 (800 Tex) as the reinforcement. For prepeg manufacturing, MTM44-1 /HTS40 (12K)-F13 /134gsm was chosen. Ideal material for micro fasteners and joints is Titanium alloy, Ti-6Al-4V, however for practical reasons it was stainless steel AISI 301which was eventually used in prototypes, the same as the one used for the attachment frames.

Design of the wafer fuselage section was subjected to boundary conditions related to the component structural arrangement, i.e. reels quantity, location and all possible load-carrying schemes of regular wafer structure including hoop-spiral, longitudinal-spiral and spiral-spiral ones were analyzed. Besides also manufacturing limitations were taken into account. With all this in mind, geometrical and weight parameters of hoop, spiral ribs and skin were determined analytically. Load-carrying schemes were studied using FE method.
An important step in the design process was the selection of a criterion to extrapolate design solutions to downscale prototypes, i.e. 1m and 0.5m in diameter. Based on the agreed criterion, geometrical parameters of scaled fuselage section prototypes were identified, like ribs shape, size and distribution. The same criterion was also applied to determine bending loading conditions for later testing.

It was found that multiple designs complied with manufacturing and loading conditions according to the initial requirements. Designed was optimized aiming to reduce total weight while meeting stiffness requirements. It was found that fuselage barrel weight could be reduced up to 40.48 kg, which is significantly lower than the weight of the aluminium reference structure corresponding to the current design of the Piaggio P180 aircraft (62.55kg). This showed the weight saving potential of lattice composite structures for aircraft fuselage applications. Thus using out of autoclave composite materials and implementing the lattice architecture a 54.5% weight saving is possible with respect to the aluminium reference design.
In parallel with the composite structure, the metallic attachment frame was designed. The level of loads applied to each element of attachment frame was estimated analytically and by FEA. Tearing, bearing and shear loading cases were considered to check structural integrity. Several possible options regarding arrangement of reels within the attachment frame were analyzed. Most crucial considerations of fuselage section step-by-step manufacturing (including restrictions developed by partners-manufacturers) and assembling were taken into account. Finally for the full scale fuselage section, the attachment frame structural solution consisted of an angle section and twenty sets of joining fittings and reels with correspondent fasteners. Equivalent design solutions were developed for downscale components accounting for 12 (1m) and 8 (0,5m) reels.
Once the wafer section design was released, a safety study using FEM was accomplished taking into account worst loading cases. With respect to hail impact, it was found that the component integrity was out of risk. It was also found that even in case the skin might be damaged by hail impact, it would practically doesn’t influence in any way on loss of load-carrying capacity.

Manufacturability with respect to materials and joints was an important issue to take into account before releasing the final design. In this respect non standardised test was conducted to evaluate fibers strength wounded around a set of reel heads, depending on the wounding angle and reel head diameter. It was found that using the configuration in which the relative angle between composite and reel head is 30º only the smaller diameter was critical, while 8mm and 10mm of rod diameter presented satisfactory results. Adhesive strength between metallic and composite parts was also evaluated. A set of alternative adhesives and adhesion methods were studied. It was found there are adhesive solutions able to perform higher than 20 MPa in adhesive shear strength (22.4 MPa) Besides the identified failure mode of tested joints was mainly cohesive thus selected sub-layer ensured enough strength, indicating that in case there is a need to achieve better results, additional attention has to be paid to resin strength properties.
Influence of manufacturing methods on material properties was evaluated. The material properties of carbon fibre composite specimens produced by the ATP and FW processes were evaluated. For the ATP there was no issue and the exact same process that would be followed for the fuselage panel fabrication could be used to produce flat panels from which standard specimens were extracted. The resulting material was of high quality and yielded repeatable material property data. The difficulty with FW arises from the fact that tubular specimens (i.e. replicating the exact manufacturing method) can only generate material property information transverse to the fibre direction. For properties along the fibre direction it was attempted to wound on a flat mandrel and extract standard size and type specimens. This proved unsuccessful since the resulting quality of the material was very low. It can be concluded that since the two material systems use the same type of fibres it would be safe to use the ATP material properties to derive the necessary material design allowables.
During most of the project duration a continuous validation process was implemented, based on a building block programme, from level 1 samples up to the larger prototype, which was the latest validation activity. Such building block programme supported the development of the integrated lattice structure. Aim of this validation approach is to provide not only validation of the proposed design, but also collect evidences about the preferred manufacturing method.
Test panels were tested under static loading conditions. Based on the extensive set of collected test results it can be concluded that the skin-to-ribs debonding is the first type of failure caused by the compression loading applied. It was also evident that no damage was recorded in the ribs as a result of compression loading. Hence it is advisable to improve the bonding between the skin and the ribs to directly improve the mechanical performance of the structure under compression.
The static study was extended up to the dynamic loading range, conducting fatigue testing on level 2 samples. The failure of the Attachment Frame without Skin specimens is dominated by the strength of the bond between the metallic reel and the composite and not by the strength of the composite itself. The large concentration of voids and the insufficient amount of resin close to the centre of the reel contribute significantly to early damage initiation and subsequent failure of the part. It is recommended that another structural gap filling adhesive should be used prior to the application of the composite and possibly an increase in the reel spacing that would allow easier compaction of the composite. For the Attachment Frame with Skin specimens the debonding of the skin from the metallic attachment frame (i.e. damage initiation) occurred at a very low cycle count and well before the final failure of the specimen, which was debonding of the skin from ribs. The type of failure under fatigue highlights the limitations of the current design with the very low skin-to-rib crack growth resistance and fairly high stiffness mismatch. The Ribs with Skin specimens failed in the same way as when tested under quasi-static loading. Failure initiated from the rib intersections and propagated along the ribs. Secondary failures of skin-to-rib debonding at the edges were artefacts of the structural elements and the free edges rather than driven by the design. Again the scaled down specimen could not capture the skin-to-rib debonding as the stiffness mismatch was much less than the full scale component. However, a very interesting finding is the damage tolerant nature of the lattice structural design with very limited influence on the fatigue performance even after the introduction of substantial impact damage on the rib intersection or the skin. Even for these specimens damaged initiated and propagated from the same locations and after the same number of fatigue cycles as with the undamaged specimens.
Eventually the large prototype was tested under static load. This prototype was submitted to a low load level in order to ensure its structural integrity for further dissemination purposes. The fuselage section was successfully tested without producing any damage in the structure as no residual strain seen in the part after the loading-unloading sequence. The strain measurement from strain gauges and the DIC are very similar which verifies the consistency of the techniques employed.

Test results were used to check consistency of simulation methods used in design. Thus a correlation of finite element analysis with experimental results as performed on three levels of the building block approach was attempted. Several options for modelling composite materials by the FE method were investigated and the results were compared with physical test data.
On Level 2, two types of structural elements were tested in compression and modelled by FEA. The structural element was a typical intersection of the ribs extracted from the scaled down version of the lattice stiffened fuselage structure in two variants; with and without the skin. A good correlation was obtained for the “skin and ribs” component, modelled either by conventional elements or by 3D solid elements and user subroutine. The “Ribs only” component has proven to be more sensitive to initial geometric imperfections and better prediction was achieved by 3D solid elements.
On Level 3, a full-scale fuselage section panel was tested in compression. The models comprised two different approaches previously used at Level 2; the skin was modelled by conventional shell elements and orthotropic material properties; ribs were modelled by 3D solid tetrahedrons and a user defined material subroutine. The correlation of load-strain achieved by the models was good, even though some level of uncertainty arises from manufacturing imperfections. The model was not capable to predict ultimate failure of the structure.
On Level 4, a scaled down fuselage section was tested in bending. Conventional shell elements were used in the model and the load-strain results are in a very good agreement with the experimentally obtained data. Also the strain map obtained by DIC correlates well with the simulation results. Whether the model would be able to predict the ultimate failure cannot be elaborated, as the fuselage section was not tested up to this point.
Finally, a full scale fuselage section under realistic load conditions was further analysed at several levels of detail. Global models were used to obtain stresses and failure indices of the whole structure and to drive boundary conditions of the local model. The failure indices obtained suggest that the structure would be able to withstand the imposed loading, however, no experimental data are available for validating such prediction.
Further aspects not directly related to manufacturability and strength performance were assessed, like reparability. In this case level 2 samples were used, as repaired panels, which were resin potted and subjected to compression testing. It was found that repaired panel exhibits comparable stiffness and equal strength with the pristine ones. This leads to the conclusion that the process elaborated to repair detached ribs from the skin is feasible and promising.
The overall project conclusion is that design and lower cost manufacturing methods have been developed for a composite lattice structure, which is potentially suitable for aeronautical application as a fuselage section. Particular issues have been identified related to joining performance between composite and metallic parts, as well as debonding of ribs and skin. Apart from these cases, structurally speaking the lattice design implemented with composite material showed to offer enough load carrying capacity, and the combination of materials and manufacturing methods are promising in terms of cost saving, which is a key factor to further promote the implementation of composite structures in fuselage applications. Weight reduction is clearly feasible, even taking into account that metallic parts are still necessary in the link between composite and conventional metallic structures.

Potential Impact:
Contribution to growth and competitiveness by enhancing European aeronautics industry: The structural concept design that will result from the WASIS project is essential to achieve these goals. While lattice structures have been known in Russia for years, and their advantages have been proved by design engineers, the opportunity of joining structural benefits with modern composite materials advantages is now unique. Combining existing material knowledge with highly automated complex manufacturing methods and an advantageous structural concept, whose know how is not known to the European manufacturing industry, will result in a highly efficient weight reduced primary structure.

Meeting demands on fuel consumption and emissions reduction: Among the mitigation technologies and strategies, a major area of research and development is the improvement of fuel efficiency. This can be obtained through different alternatives, one of which being weight reduction, which this project will address.

Reducing material wastes with opening design: The use of specially designed joints during the section manufacturing will allow having openings in the structure without having to cut fibres. This relates directly to the material wastes associated to a new product.

Promoting the European aeronautics manufacturing industry: This project manufacturing processes will focus on two highly technologically complex process, robotic filament winding and fibre placement. The high level of expertise required to operate these technologies to generate complex geometries, as the fuselage section, will promote the know how of the European manufacturing industry.

Addressing safety on composite damaged structures: The developed fuselage structure will be analyzed to fulfil the same safety levels as any conventional structure would. Regarding repair, lattice structure’s loading capability is supplied mostly by the reinforcement ribs, so skin can be made thinner, which in turn reduces weight and allows for easier repairing techniques.

Societal benefits: In view of the current crisis it is essential to maintain the employment across the European aeronautic industry. Technologies that will result from the WASIS project are expected to have essential societal impact. WASIS project results exploitation will offer opportunities for the employment of highly skilled professionals. In addition, other societal benefits will be addressed by the current proposal, as are public health, with the emissions reduction and the induced noise reduction. WASIS will aim for a cleaner environment.

List of Websites:

http://www.wasis.eu/

Final Report Summary - WASIS (Composite fuselage section Wafer Design Approach for Safety Increasing in Worst Case Situations and Joints Minimizing)

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