Production and Analysis Evolution For Lattice Related Barrel Elements Under Operations With Advanced Robustness

Projektinformationen

POLARBEAR

ID Finanzhilfevereinbarung: 605448

Projekt abgeschlossen

Startdatum 1 Dezember 2013

Enddatum 30 November 2016

Finanziert unter

• Specific Programme "Cooperation": Transport (including Aeronautics)

Gesamtkosten € 3 419 215,20

EU-Beitrag € 1 195 734,00

Koordiniert durch

DEUTSCHES ZENTRUM FUR LUFT - UND RAUMFAHRT EV
Germany

Final Report Summary - POLARBEAR (Production and Analysis Evolution For Lattice Related Barrel Elements Under Operations With Advanced Robustness)

Executive Summary:
Project context and objectives
Environmental and economic issues force future aircraft designs, like the A350, to maximize efficiency with respect to weight and cost in order to keep air transport competitive and safe. As metal designs have reached a high degree of perfection today, extraordinary weight and cost savings can hardly be expected. Further potentials are seen esp. with fibre composites gaining importance in aircraft structures over the past years and still being under development.
The first generation of composite aircraft structures, such as Boeing-787 and AIRBUS-350, is based on the so-called black metal technology following the conventional semi-monocoque concept. This does not allow utilising the high strength of the fibres to the full extent because the main structural element of such a concept is a panel that consists of a highly loaded skin and thin-walled stiffeners
A thin-walled stiffened panel can be rightly called the crown of the evolution of aircraft metallic structures. It exhibits high strength equally for a single as well as for the combined loading, both for direct and for shear loads. Due to its capacity to tolerate high strains, the metal skin performs well under impact and tolerates also the so-called post buckling effect, when at high post-critical loads a part of the skin loses its stability. During stability loss it switches itself off as a load-carrying element, quasi saving itself from destruction and allowing the remaining part of the structure to continue carrying the external loads. When the external load declines the skin returns to its original state and restores its stiffness. This effect has significantly improved the structural weight efficiency without compromising on safety.
Up-to-now the potential of composites is incompletely realized in conventional airframe layout. This is due to a basic peculiarity of laminates showing high mechanical properties only in fiber direction. But for design and technological requirements a composite skin in stiffened structures generally consists of diversely oriented laminate layers. Also, composite stiffeners (frames, stringers) do not consist of unidirectional (UD) laminate layers. Because the properties of the “binder” between the fibres determine the utilisation effectiveness of the fibres’ high strength properties, the weight efficiency of composites is merely due to their low density.
For a more effective realization of high mechanical properties of up-to-date CFRP, design principles and technological concepts have to be developed which provide higher integrated structures to reduce fastening but still satisfy in-service requirements like reparability and robustness.
PoLaRBEAR project focuses on the development and analysis on a local level regarding automated manufacturing of a protection layer in winding process, automated manufacturing with prepreg technology, buckling analysis of anisogrid panel structures, damage tolerance and fatigue of the rib-structure and reparability of the elements. To increase the technology readiness level of the innovations derived in EU-ALaSCA, further analyses are necessary on local level. The main objectives of this research programme are:
• Industrial highly automated process for cost efficient barrel manufacturing
• Advanced reliability of geodesic structures under operational loads
• Design rules for robust grid structures

Project Context and Objectives:
1 Scientific and technical quality, relevant to call topics
1.1 Concept and objectives
Scope: The PoLaRBEAR project (Production and Analysis Evolution For Lattice Related Barrel Elements Under Operations With Advanced Robustness) focuses on reliable novel composite aircraft structures based on geodesic technology aiming at a significant higher Robustness and Technology Readiness Level (TRL). The project takes up issues directly from the EU-project ALaSCA.

Introduction
In the 5th call of 7th framework programme of the European Commission, the project ALaSCA (Advanced Lattice Structures for Composite Airframes) focuses on a top-down approach for maximum weight and cost reduction of airframes by developing manufacture optimised lattice fuselage structures fulfilling fundamental aspects of airworthiness. The approach consists of investigations on geodesic designs starting from aircraft configuration level, over development of potential barrel concepts to study fundamental challenges on element level.
In the ALaSCA project (see Fig. 1), the analysis of a pro-composite fuselage structure started with the investigation of most beneficial aircraft configuration achieving a possibly long undisturbed fuselage barrel section in front of the wing. After having determined fuselage barrel loads, different Pro-lattice fuselage barrel concepts were developed in addition to reference barrel concepts with conventional semi-monocoque design including a metal and composite variant. ALaSCA puts emphasis on the global analysis and understanding of the grid structure with various numerical analyses and structural optimization on global barrel level, getting suitable barrel concepts including components like window cut-outs, floor structure, attachments and barrel-barrel interface.

Fig. 1: Focus of EU-ALaSCA
It has been shown in ALaSCA that an anisogrid structure concept has beneficial effects and the potential, of getting a step further to pro-composite fuselage design compared to reference semi-monocoque structures made out of CFRP or metal.
The investigations on barrel concept level of the project ALaSCA show significant design improvements for composite geodesic fuselage structures due to:
• The resulting non-rectangular skin bays between the ribs, which show increased buckling coefficients compared to rectangular skin bays with the same weight.
• The uniaxial loading of the ribs, with which a strain allowable increase can be pursued, considering an impact protection of the highly oriented ribs.
• Aiming a possibly high axial stiffness for the stiffeners, CFRP-Metal-Hybrid shows the potential achieving a high axial stiffness, while also being damage tolerant.
Due to the focus of EU-ALaSCA on a global barrel design level, there are open questions on a local level regarding automated manufacturing of a protection layer in winding process, automated manufacturing with prepreg technology, buckling analysis of non-rectangular skin bays, damage tolerance and fatigue of the rib-structure and reparability of the elements.

Fig. 2: Focus of ALaSCA and PoLaRBEAR
EU-PoLaRBEAR will focus on the above mentioned investigations, relying on a bottom-up approach on local level to increase the technology readiness level of geodesic structures (see Fig. 2).
Investigations on manufacturing a barrel with a lattice structure are a goal out of the scope of European industry using prepreg material instead of filament winding. Without a methodology for the numerical analysis of non-rectangular skin bays, an automated optimization to exploit the full potential cannot be performed. Aircraft operations require the rib structure to be damage tolerant and the fatigue characteristics have to be analysed and tested. Consequently, repair concepts have to be investigated for in operation use. In addition to all investigations, the focus of PoLaRBEAR is put on the creation of Design Rules for geodesic structures. In addition, a set of recommendations is planned, reducing future development time and costs.

PoLaRBEAR approach and objectives: To increase the technology readiness level of the innovations derived in EU-ALaSCA, further analyses are necessary on local level. The main objectives of this research programme are:
• Industrial highly automated process for cost efficient barrel manufacturing
• Advanced reliability of geodesic structures under operational loads
• Design rules for robust grid structures

PoLaRBEAR methodology: While the global structural behavior of composite geodesic structures is investigated and understood in EU-ALaSCA, for a higher TRL the understanding of the local structural behavior, which is different from today’s aircraft structures, is necessary. The objectives will only be achieved when solutions to the following issues in terms of design concepts, damage tolerance and fatigue, manufacturing and testing are found:
• Development of advanced winding process for protection layer concepts and AFP barrel manufacturing processes under series conditions of aircraft production; manufacturing of coupons, elements and test panel;
• Numerical analysis of the skin buckling behavior dependent on the geometrical shape of the skin bays within the geodesic grid (polygonal areas) and verification by testing.
• Protection concepts and material improvement aiming at increased damage tolerance and fatigue properties. Evaluation and verification by testing.
• Development of repair concepts, derived from examination by numerical analysis and testing.
• Evaluation of individual results and conclusion of design rules for geodesic structures

PoLaRBEAR work package structure: The topics mentioned above are represented in a work package structure subsequently addressing the various design levels:
• WP 1: This work package comprises the definition of requirements of aircraft under operation as well as material allowable for laminates. The way of certification for UD-ribs gets shown and results in creation of design rules for lattice structures (considering MI 5).
• WP 2: This work package focuses on usage of production technologies, esp. addition of protection layer in a filament winding process and AFP process for barrel manufacturing. Concepts will be evaluated with test specimen on suitable element or panel level and tested.
• WP 3: In this work package a buckling methodology for non-rectangular skin bays will be investigated. Advanced FE-modeling .methods will be used for increased analysis effectiveness and validated with a panel test.
• WP 4: This work package focuses on damage tolerance and fatigue aspects for multi-orientation laminate, CFRP-Hybrid-Laminates and UD-rib-structures with protection layer. Crack propagation analysis will be performed combined with robustness analysis. The concepts will be manufactured and tested for validation.
• WP 5: In this work package repair concepts addressing skin, ribs and nodes will be worked out and sized. These repair concepts will also be manufactured and tested for validation.

Results – PoLaRBEAR main innovations (MI): Applying the above-mentioned methodology and WP structure, the objectives will be achieved by main innovations in the following areas:
• MI1: Automated Fiber Placement (AFP) process for barrel production, i.e. investigating and evaluating the potentials of this promising manufacturing process vs. the time and cost-efficient available filament winding process.
• MI2: Buckling methodology for non-rectangular skin bays considering pure loads, e.g. shear or compression, or combination of these loads: procedure for analysis and optimization of polygonal skin bays in terms of buckling using effective numerical analysis methods.
• MI3: Damage tolerant ribs with high axial stiffness, i.e. increasing damage and impact tolerance of axially high-stiff ribs by improved protection philosophies e.g. protective covering, multi-orientation laminates and hybrid materials.
• MI4: Pro-repair design concepts addressing the structural members skin, ribs and nodes and their interdependencies, i.e. integrated and certifiable repair methods with a high degree of automation and reproducibility.
• MI5: Design Rules for geodesic rib structures summarizing and derived from the results of MI1 to MI4 based on local level investigations for increase of Technology Readiness Level (TRL).

Project Results:
PoLaRBEAR – Production and Analysis Evolution for Lattice Related Barrel Elements under Operations with Advanced Reliability

Project – Consortium - Partner Abbreviations
DLR German Aerospace Centre - EU
TsAGI Central Aerohydrodynamic Institute - RU
CRISM Central Research Institute for Special Machinery - RU
TU BS Technical University of Braunschweig - EU
NIK LLC NIK-Samara - Research Engineering Centre - RU
QMUL Queen Mary University of London - EU
SMR S.A. Engineering and Development - Switzerland
AIRBUS Airbus Operations GmbH - EU
MIPT Moscow Institute of Physics and Technology - RU
MSC MSC.Software GmbH (MSC) - EU
VIAM All-Russian Scientific Research Institute of Aviation Materials (RU)
MUCTR Mendeleev University of Chemical Technology of Russia - RU
ULeeds University of Leeds - EU
Abstract
In the level 1 projects ALaSCA (Advanced Lattice Structures for Composite Airframes) and PoLaRBEAR (Production and Analysis Evolution for Lattice Related Barrel Elements under Operations with Advanced Reliability) the potential of a novel lattice airframe architecture is demonstrated. Derived from a well proven spacecraft structure design, all classical design levels from coupon to full-size structural test level are investigated. During the projects different barrel design concepts are developed including design solutions on component level like window cut-outs, barrel-barrel interface or floor-barrel attachment. Accordingly, the project partners developed new analysis methods and sizing algorithms as well as manufacturing processes and test methods for these unconventional stiffened fuselage structures.
Keywords: Composite, Anisogrid, Lattice Structure, Analysis, Test, Airframe,
INTRODUCTION
High demands regarding safety, ecology and comfort for state-of-the-art civil aircrafts, lead to the necessity of using new materials, such as CFRP, to reduce weight and cost a the new generation of airframes.
Implementation of new materials to high-loaded structures inevitably leads to the necessity of fundamental re-assessment of structure concepts and layouts. This is reasonable as physical properties of the new CFRP materials are fundamentally different from formerly used metallic alloys.
Following this philosophy, lattice composite structures offer the possibility of a new structure concept which purpose it is to realize the potential of current CFRP materials while at the same time minimize their critical shortcomings. A load-bearing lattice structure which consists of ribs with unidirectional stacking sequence increases the maximum levels of loading for high-strength fibres (CFRP) hence a higher weight efficiency of the structure results. A high level of integration allows manufacturing of large lattice composite sections in one manufacturing cycle which reduces costs and the assembly time.
For about 10 years the lattice composite technology is successfully used for the manufacturing of Russian launch-vehicle primary structures, high benefits regarding weight and manufacturing cost as compared to metallic prototypes [1] have been achieved.
The main purpose of ALaSCA [2] and PoLaRBEAR projects was the adaptation of the lattice composite technology to a fuselage section of a civil aircraft and the assessment of weight and cost saving potentials for this type of structures in comparison to conventional metallic and composite semi-monocoque analogues.
In the present report a description of the main activities and results of PoLaRBEAR project is given. In addition main results of ALaSCA project are also included to give a better context. In Chapter 2 the main lattice design concepts considered in the projects are described. In Chapter 3 an overview of methods for numerical strength analysis of the new concepts is given. Chapters 4 and 5 describe the two main lattice structure concepts and the research activities on these concepts performed during the projects. Chapter 6 summarizes the main results of the projects and gives an outlook regarding further research topics.
Anisogrid Fuselage design concepts
On airplane level, a predestined aircraft configuration is identified with overall lower fuel consumption. For this configuration a beneficial fuselage section is chosen in order to apply different Anisogrid concepts. A wide variety of concepts is investigated including component design solutions like window cut-outs, barrel-barrel interface and floor-barrel interface.
Airframe Configuration
The selected aircraft configuration has a low wing, rear mounted engines and a T-tail. This engine arrangement allows a relatively short landing gear, but leads to higher loads for the wing structure under manoeuvre flight conditions.

Figure 1: fuselage barrel section in focus
The configuration combines the aerodynamic benefits of the NLF wing (natural laminar flow) with the potential for future engine concepts (ultra-high bypass ratio turbofan, open rotor) with larger dimensions [3]. In addition, the long undisturbed fuselage section is suited for the usage of carbon fibre reinforced plastic materials. Therefore this configuration was selected for the detailed investigations on the lattice structure fuselage design.
Variety of Fuselage Design Concepts
Monolithic highly integral Lattice Design Concept
According to this concept, the fuselage barrel includes layered skin and lattice stiffeners as main primary structure elements. This structure concept is very close to lattice composite structures of rocket adapters, which are well-proven for space applications. A monolithic structure, including skin and lattice stiffeners is manufactured in one shot by means of wet winding technology which was a basic lattice design in ALaSCA [4]. In the PoLaRBEAR project a developed AFP-manufactured Prepreg concept was focused on [5].

Figure 2: Monolithic highly integral Lattice/Anisogrid Design Concepts
Multifunctional Lattice Concept with pressurized inner Skin
The Multifunctional Lattice Structure Concept is substantially different to the conventional aircraft structures with load-bearing stiffened skin. In frames of the concept, the main primary structure element responsible for bearing moments and forces is a stiff lattice grid on unidirectional (UD) composite ribs, while skin located inside the grid is used for pressurizing only (see Figure 3) [6]. Besides the lattice grid and the pressurized inner skin, the structure has the outer (“aerodynamic”) skin and a foam filler, located between two skins, which can serve both for insulation and for protection of composite ribs from the environmental factors.

Figure 3: UD-rib structure concept
The proposed concept has the following main advantages (in comparison to conventional composite structures with laminated skin and stiffeners):
force flows are mainly directed along the direction of carbon fibres that allows to realize higher levels of loading in the composite primary structure elements;
due to its topology, the grid of ribs (i.e. the primary structure) can be reliably protected from external and internal factors (e.g. impacts, environment, hot/wet etc.) with a slight weight penalty;

The Multifunctional Lattice Concept has higher damage tolerance due to high robustness of lattice structures. Furthermore, impacts of skin cells between ribs do not damage the primary structure (lattice grid), so they have almost no influence on its load-bearing capability. Also, some available scenarios of repair for this type of lattice structures have been developed during the projects.
Multifunctional Differential Anisogrid Concept with inner load bearing Skin
This design concept is based on a load carrying skin, stiffeners in grid architecture and frames as circumferential stiffeners. The stiffeners are placed as geodesic lines on the cylindrical skin (called helix-stiffeners), non-rectangular skin bays are generated consequently, which increase the stability properties of the skin. To avoid intersections of the stiffeners, they are placed on two sides of the skin. Due to the load bearing skin on the inside and the need of an aerodynamic skin, this concept is called the Skin-Inside concept. In order to close the space between the two skins, a foam core with low density is considered as not load carrying in dimension process. See Fig. 8 for the concept illustration. This concept shows in contrast to a Skin-Outside concept a high differential approach, which means that most components are manufactured separately and assembled afterwards. [7,8].

Figure 4: Pro-Lattice Skin-Inside Concept Overview
The primary structure consists of frames with C-shape and stiffeners with Omega-shape (hat shape). Due to the position of the stiffeners with an angle and the loss of axial panel stiffness, CFRP-Steel-Hybrid material is assumed for the stiffeners because of the increased axial stiffness properties and the fulfilment of repair, joining and damage tolerance demands at the same time [9]. The foam and the aerodynamic skin are multifunctional elements for the concept. The foam is filling the space between the skins, but is also acting as thermic insulation for the cabin as well as protection of the primary structure elements from outside impacts. The aerodynamic skin layer is assumed to be an aluminium layer to provide lightning strike protection in addition to the aerodynamic shape.
METHOD DEVELOPMENT
Within the projects the partners developed numerical and analytical methods to solve the sizing problems for the proposed concepts. On global aircraft level a multilevel optimization cycle is implemented for the sizing of anisogrid fuselage barrel sections. A numerical heuristic optimization loop is set up for assessment of buckling and material strength criteria of Anisogrid panels with the finite element method. A new semi-analytical method has been developed for calculation of the buckling value of grid-stiffened skin shells on element level. A new shell element formulation for lattice structures was developed which allows faster and more robust numerical simulations compared to conventional used FE shell elements while having the same accuracy. On the global barrel sizing level a methodology to investigate the structural robustness against local imperfections has been applied for the first time for the stiffened cylindrical shell.

Multi-Level Sizing Algorithm for Airframe Barrel Section
An important aspect for the fuselage section is the sizing process of the sections with lattice and alternative layouts. The main purpose of sizing is the weight estimation of different section structure variants and use of the results in a comparative analysis. Thus, for all alternative concepts parametrical FE-models for varying basic parameters of models are built up. The main requirement for the models of the different fuselage structure designs is the same accuracy for strength analysis in order to assure correct comparison of the results.
The 4-level algorithm [10–12], which satisfies these conditions, was adapted for the determination of rational parameters of the composite lattice structures within the fuselage section (Figure 5). Rational parameters of the structure are determined during the iterative process, which include solving different strength tasks at 4 detail levels: level of aircraft (level 1), level of section (level 2), level of specific elements (level 3), level of element/micro level (level 4).
The 4-level algorithm is fully automated and all FE and analytical models are built automatically on the basis of the specialized database, which contains all structure parameters.
The main results of each iteration are weight parameters of the structure and reserve factors for strength and buckling. The techniques and algorithms for these analyses were validated using comparison of results obtained via various modelling techniques and experimental data.

Figure 5: Scheme of 4-level designing algorithm (for lattice structures)
At the 1st level (geometrical model), the loads on fuselage section are specified. At the 2nd level, the mass and stiffness parameters are analyzed. In the case of the lattice structure, the ribs are made with winding technology, the barrel section is an integrated structure, and the stiffness parameters concern to groups of ribs (helical, circular, and longitudinal). At the 3rd level, the strength analysis of specific components of the structure: framing of cut-outs, attachments "floor-to-barrel" and interfaces between different fuselage units is performed. At the 4th level the detailed FE-model for analysis of local stress-strain state is generated and the local strength analysis is performed.
In addition to FE-models, a number of analytical models are automatically built up on the basis of the universal database, thus providing fast and complex analysis of the structure with low labour input and low time expenses.
The specialized database and special architecture of the 4-level algorithm allow to provide automatic communications between FE- and analytical models and to exclude the accidental errors connected with data transfer from one model to another.
Figure 6 shows modelling of the lattice barrel for different kinds of the strength analysis: global, local, and on micro level (level of fibre).

Figure 6: FE modelling of lattice structure for different strength analyses [13]
Numerical panel optimization
The design process of a stiffened anisogrid composite panel is a multi-parameter optimisation problem, which is performed with the MSC Software SimXpert. Using SimXperts capabilities (Pyhton API) and SciPy methods; a numerical optimization loop (model setup/parametrization and optimization setup) is set up for anisogrid panels using the Finite Element method (see Figure 7).
The advantage of this high-fidelity numerical optimization is that besides global and local panel stiffness criteria also material strength criteria can be assessed. Analytical approaches regarding material strength delivers in general no or inaccurate prediction regarding local stresses and strains of the individual panel elements. This is especially important for composite since they are more sensitive to damages compared to metals. Therefore it is possible to further reduce weight of Anisogrid composite panels because both the buckling and strength criteria can be fulfilled.
The high-fidelity numerical optimization loop consists of a global and a subsequent local optimization in order to ensure that the optimization is not trapped in local optima. Two modern heuristic optimization methods, ant colony optimization and differential evolution were consequently implemented.

Figure 7: Optimization Temple Flow Chart
Skin-Bay Buckling Factor Determination Method
A procedure to analyse the buckling behaviour of curved skin fields in grid-stiffened shells is developed by Weber and Middendorf [14,10]. As load cases, combinations of biaxial compression and in-plane shear are considered. The laminate of the skin is assumed to be symmetric, balanced and orthotropic. Furthermore the material law of the Classical Laminate Theory is applied and the curvature is taken into account by using kinematic relations of Kirchhoff/Love for thin singly-curved shells.
The buckling load is obtained by minimizing the total potential energy of the system according to Ritz and solving the resulting eigenvalue problem. Since the stiffness matrices only depend on the shape functions they only need to be calculated once. This is a major advantage compared to finite element analyses where a change in geometry demands a change of the model. Compared to referencing finite element and analytical results, the method shows a high level of accuracy even in cases where the finite element method encounters numerical problems (Weber and Middendorf [14]).

Figure 8: Stiffening patterns and dimensions of skin segments for a) orthogrid, b) isogrid, c) diamond grid and d) kagome grid [14].
A comparison of buckling factors by Weber and Middendorf [14] for different grid patterns, see Figure 8, and varying half vertex angle α is performed. As load cases axial compression and in-plane shear are considered. In case of the skin field model, interaction of adjacent skin segments leads to an increase of the buckling factor compared to the case of a single panel. For orthogrid structures this effect is limited to situations including shear load. Iso- and kagome-grids show considerable higher buckling factors compared to orthogrids for same α.
Ribs Buckling Methodology
As a rule, FEM analysis of buckling of a structure consists of two parts: computation and analysis of results. The results include the data both margins on global buckling (section) and on local buckling (elements) (Figure 9). Determination of necessary value of critical buckling load consists in the visual analysis of the results with choosing required buckling mode with minimal eigenvalue. This conventional "manual" operation is labor- and time-consuming and thus is not suitable for use in the automated procedure of sizing (optimization) of the lattice structure of fuselage sections.
To solve this problem, new approaches for the global and local buckling analysis of the lattice structure of fuselage sections have been proposed [13]. To avoid confusion regarding different data in the results output at the global buckling analysis, the results corresponding to the local buckling modes are excluded by means of simplification of the section structure model, namely each rib is modelled by one "beam" finite element between neighbor crossings.
The required level of accuracy (which decreases by 4-8 % because of the simplified modelling) is reached due to use of correction factors. These factors are derived and written into the specialized database of the 4-level design algorithm at a stage of preliminary parametrical investigation (Figure 10).
The "fast" algorithm of local buckling analysis is developed for use in the sizing procedures. The algorithm includes the following operations:
Parametrical calculations of dependence of correction factors on structural and geometric parameters of lattice section and ribs; adaptation the results to the database format.
FE-analysis of linear strain state of lattice section; selection of critical areas (compressed ribs).
Calculation of the correction factors k_l in accordance with current parameters of ribs and information from the database.
Determination of critical loads of ribs local buckling basing on the analytical solutions corrected by the factors k_l.
The factor k_l takes into consideration conditions of the rib supporting, the rib’s curvature and twisting.

Figure 9: Output of results at buckling analysis of lattice section structure by means of «Patran»

Figure 10: Scheme of section buckling analysis with "adjustment" of result for simplified model
Use of the new methods and algorithms for global and local buckling allowed significant reduction of labor and time expenditures at the sizing of lattice section structures.
Finite Elemente Formulation – SMR

Skin-stiffened grids, and in anisogrids in particular, are more challenging to numerical analysis than conventional stringer-frame semi-monocoque structures. Triangular shell elements from several FEM codes exhibited an overly soft behaviour on anisogrids, resulting in bad mesh convergence. Hence the motivation to develop fast and yet accurate and reliable numerical analysis methods. This FEM formulation realizes this goal; it has been applied successfully to anisogrids as well as unstiffened shells [15].
The developed formulation does not contain any removal of locking. Transverse-shear locking is inhibited by adopting a rotation-free discretization of the mid-surface and exactly calculated shell surface normals (and derivatives) at the Gauss points. Hence, only membrane locking may occur, for example, in shells and in nonlinear analysis of plates.
Such rotation-free Kirchhoff-Love element formulations require discretization’s where the direction of the surface normal must not change when passing from one element to the other. This condition is called G1 continuity. In PoLaRBEAR, it is implemented by using a C0 discretization to which a set of nonlinear constraints is added. The C0 discretization consists of Bernstein-Bézier triangles of order 5 or higher(see Figure 11). The high order leads to better mesh convergence rates and largely removes membrane locking. Triangles give for more modeling flexibility than quadrilaterals and untrimmed NURBS. The constraints are formulated as integrals along the sides of adjacent elements and are enforced by means of penalties.

Figure 11: Control points of a quintic Bernstein-Bezier element
Additionally, C1 constraints have been developed where not only the direction of the surface normal must not change when going from one element to the next, but also the length of the surface normal. These constraints are linear which means that the constrained degrees-of-freedom can be eliminated from the set of equations, resulting in the reduction of the degrees-of-freedom by roughly 66%. Using these constraints where possible greatly improves the effectiveness of the formulation. But because C1 constraints are stronger than necessary, they must be used only where the solution field has sufficient regularity, such as in skin fields. At the model boundary, at material boundaries, thickness jumps, and junctions, etc., the weaker G1 constraints must be used instead. Otherwise, stress oscillations will occur and mesh convergence will be poor.
The formulation is implemented in the nonlinear B2000++ FEM code. Different polynomial orders (5 to 9) are supported and quadrilateral elements in addition to triangles. Extensions such as MITC (for the quadrilateral Bernstein-Bézier elements), Hellinger-Reissner mixed elements, and transverse-shear have been implemented as well, mainly for comparison with literature results and to find out how the performance could be further improved.
The selection of G1/C1 constraints is carried out automatically by the pre-processor. Boundary conditions such as clamping or symmetry can be specified by means of boundary elements. The creation of FEM models with this method is nearly as simple as with conventional Reissner-Mindlin low-order shell elements. This includes the quite complex grid models.
The new formulation is an effective complement to the existing low-order Reissner-Mindlin shell elements. It can be applied to a large variety of thin-shell problems. Because it is a Ritz procedure (which means that it is a fully conforming method), mesh convergence is monotonic. It is simple to use and has excellent performance. It does not fully replace the low-order shell elements, however. Although less effective if high accuracy is desired, low-order elements are more practical when dealing with very complex geometries.

Structural Robustness Assessment
The structural robustness, the sensitivity of a structure to local failure, of the anisogrid shells is assessed with so called lower-bound methods which are used to determine a knock-down factor for the buckling load of cylindrical shells. These knock-down factors are defined as the ratio of the imperfect to the perfect buckling load (Figure 12). Different perturbation approaches were developed to assess the sensitivity of shell structures to local failure by means of knock-down factors.

Figure 12: Load-displacement curve and corresponding buckling loads of a cylindrical shell
The single boundary perturbation approach was introduced in [18] and delivers knock-down factors whose consider the combined influence of geometric and loading imperfections(compare with illustration shown in Figure 13). Further details regarding numerical model and settings are described in [18].

Figure 13: Illustration of the SBPA method in the numerical analysis
The rigid plane moves in direction of the cylinder axis within the geometrically non-linear simulation which results at first in a local contact between boundary perturbation height and plane. The rigid plane moves further until there is a full contact between plane and cylinder edge. A localized single dimple appears under the boundary perturbation height which moves gradually to the cylinder middle during axial compression (Figure 14). The buckling failure of the shell is initiated as soon as the single dimple reaches a critical amplitude. The parameter which determines the amplitude of the single dimple within the SBPA is the boundary perturbation height h. This procedure is repeated multiple times for different boundary perturbation heights h which results in the characteristic buckling load vs. boundary perturbation height diagram of the SBPA (Figure 15).

Figure 14: Pre-buckling deformation pattern for the proposed SBPM numerical model
The corresponding lower-bound buckling load is defined as NSBPA whereas the knock-down factor is defined as ρSBPA (ratio of NSBPA to NPer). The knock-down factor can then be used to quantify the structural robustness of the structure.

Figure 15: Illustration of characteristic SBPA diagram for shells with ring deformation pattern in the pre-buckling range [1]
However, this approach requires non-linear finite element simulations which are unfortunately very time consuming and costly. Therefore it is currently investigated how to simplify this method in order to enable an implementation within an optimization loop.
MONOLITHIC HIGHLY INTEGRAL ANISOGRID PREPREG DESIGN CONCEPT
In the PoLaRBEAR project two main design concepts are analysed and the following chapter shows the highly integral monolithic design concept.
This design concept is using the high stiffness properties of a Prepreg material with constant fibre volume content of 60% in the whole rib structure and a load bearing skin. The concept enables a fully automated manufacturing of the integral rib structure and optionally also the skin in one process. The patent process is in progress for the structural as well as for the manufacturing concept.
Primary Airframe Design Concept
The Anisogrid Prepreg Design Concept consists of helical and circumferential ribs and a load bearing skin see Figure 16. The integral rib structure is built of Rib Layers as main load-bearing layers in rib direction and Interface Layers for an improved connection of the stiffener structure with the load bearing skin. The interface layers are also intended to be automatically placed like the rib layers.

Figure 16: Anisogrid Prepreg Design Concept
Considering a high production respectively fibre placement rate, the advantage of constant fibre volume content in the rib structure results in the absence of over-pressing the rib layers in the knot. Consequently in the current concept, the rib layers of two rib directions have to be cut and one layer is running endless through the knot. This procedure alternates over the knot height for all rib directions, see Figure 17. Under consideration of three rib directions, equal height of all stiffeners and an equal distribution of cut and uncut layers in all directions, the concept results in 33% of endless layers for each rib direction.

Figure 17: Sequence of cut and uncut layers in Anisogrid-Prepreg-Concept
In order to further increase productivity of the fibre placement process; pre-stacked package layers are used for the ribs. This means that the package layers are prepared in advance to place several single layers in one step. An additional advantage of such package layers is the possibility to implement also other fibre orientations into this layer. As result this approach is enabling the fulfilment of current stacking rules in civil aviation.

Figure 18: Details of rib stacking with Interface and Pre-stacked layers
The stacking of a rib with alternating pre-stacked layers (package layers) and additional interface layers is shown in Figure 17. The interface layers covers the rib stacking and connects the rib structure with the skin.
In summary following design details are considered for the Anisogrid Prepreg Design Concept:
Integral Anisogrid rib structure concept with load-bearing outside skin
High stiffness properties due to constant 60% fibre volume content
Possibility of using post-buckling skin due to interface layers
Fully automated manufacturing process for rib structure including the interface layers and the skin
AFP process enables adaptable rib cross-section along the rib length
Manufacturing process
As mentioned before an Automated Fibre Placement (AFP) process is intended for manufacturing the whole structure. For this process a positive tooling is used whereon the ribs are placed into grooves (negative tooling), see Figure 19.

Figure 19: Visualization of AFP-process for Anisogrid manufacturing
The placement of the package layers is state-of-art for current fibre placement machines. An angle-adaptive knife is seen as simple to implement. The placement of the interface layers is in contrast not state-of-art. A placement head is under development (patent pending) which has the ability to drape the interface layers directly into the grooves.
It is intended to use a fully metal tooling to ensure high dimensional accuracy, reproducibility and tool life for reduced recurring costs. One drawback of a metal tool in comparison to e.g. a partly elastic tool is that the occurring undercuts of the positive curved grid structure leads to a multi core design. Due to the height and the radial orientation of the ribs core elements needs draft angles on all sides. Resultant a probably complex multi-core tool is needed.
The application of the aerodynamic, load-bearing skin laminate can be done optionally wet on the wet grid structure in a co-curing process or in a second co-bonding step wet on the cured grid structure for example.
Element specimen for property determination - Prepreg specimens
For the specific Anisogrid-Prepreg Design Concept, static tension and compression tests are performed and strength and stiffness results are compared [16]. The goal is to determine the reduction of grid knot specimens compared to undisturbed Grid rib specimens. Especially the cut layers of the Prepreg concept in the rib knot are substantially different to the wet winded lattice structure.
Therefore a high amount of undisturbed rib and knot specimens are manufactured as well as additional kinds of reference specimens for a better understanding of effects. Therefore typical laminate specimens are produced with the same rib and knot fibre orientation ratio. Second undisturbed and disturbed rib specimens are investigated, which are cut out of a thick plate. Overall 65 specimens are produced and tested.

Figure 20: Testing of Anisogrid Prepreg specimen
According to the test results the stiffness of the undisturbed rib (rib between knots) can be used for sizing the whole rib structure including knot areas. The resulting different stacking in the knot area showed no remarkable influence to the stiffness for tension as well as for compression. In consequence the advantage of achievable high stiffness can be fully used for the Anisogrid Prepreg concept. This can be exploited especially for stiffened structures, where stability is a more critical design driver than strength.
In terms of strength, the Anisogrid Prepreg concept has to operate with much lower maximum allowable stresses due to the knot area. Compared to the undisturbed rib, a reduction to circa 34% in tension and even 50% in compression is massive. The matrix interface of the cut plies in the knot area are failing very early and because of the very narrow ply widths, cracks seem to separate the layers directly from side to side. In consequence the load bearing cross-section gets reduced very quickly.
Test panel for post-buckling capabilities of Anisogrid structures with skin
The post-buckling behaviour is investigated for anisogrid and orthogrid structures. The design parameters for the examined panels are defined in a way that the only difference between the two panels is the stiffener alignment. A large series of different panels (orthogrid and anisogrid panels) are analytically pre-designed for this investigation. In order to analyse the post-buckling behaviour of the panels, geometrically non-linear FEM simulations are used.

Figure 21: Buckling pattern of the anisogrid between limit and ultimate load [17]
The results showed that the investigated anisogrid panels in contrast to the orthogrid panels have a benevolent post-buckling behaviour between limit load and ultimate load leading to an on average 18 % higher collapse load and correspondingly higher strain energy. Additionally a constant buckle deformation pattern for the anisogrid panels differs significantly from the orthogrid panels, see Figure 21.
For validation of this effect a compression post-buckling test panel is sized in an Anisogrid Prepreg design with load bearing skin. Even that the design concept is intended as a primary airframe structural design, the test panel is designed as generic test panel to validate observed effects of this special architecture.
The panel is manufactured in a manual placement process for the whole structure, whether the methods used are similar to the automated manufacturing process aimed to.
The panel is tested in a static pure compression test until global failure of the structure. After several load steps, the panel cracked in the middle of the panel due to compression strength failure of the knots. The post-buckling loading state is validated for the load bearing skin wherein the skin is eluding from the compression load due to local buckling.

Figure 22: Anisogrid Post-Buckling Compression Test [18]
Three main test goals could be validated successfully. It could be shown that the buckling pattern in the numerical model and the test panel equals in a remarkably way. The load level is simulated correctly as well as e.g. buckle deformation direction and position of the skin bay buckle. Even the shape of the skin bay buckles is correctly simulated. The difference of the load-displacement curve between test and simulation is the missing load drop after global panel failure, see Figure 23. Reason is a missing implementation of strength criteria with successive degradation of the simulation model.

Figure 23: Anisogrid post-buckling test panel
The functionality of the interface layers could also be shown in a very successful way. The observed first and second separation of the skin and grid at the panel sides are due to the open edges and resulting very high deformation levels. In a realistic structure without open edges respectively stiffened edges this behaviour doesn’t occur in this way. Remarkable is that the separation is not running through the interface in a wide extent. The failures stayed locally and the panel could be additionally loaded in a great extent. The same behaviour can be seen after global failure of the panel. The knots where failing and the skin didn’t separated in wide areas from the failed middle part of the panel. In the top and bottom area the skin-rib interface is still intact. Also the great skin deformation in the deep post-buckling state didn’t generate high enough peel stresses to the interface layers to separate both.
The third test goal is also validated due to the strains of circa 4000 µstrains which has been reached in the helical ribs when the global failure of the panel occurred. These maximum strains were also the result for static compression tests of knot element specimen.
Resultant it can be shown successfully that the Anisogrid architecture with triangular skin bays has a different post-buckling behaviour compared to classical orthogonal stiffened structures. The missing mode switches, which are in fact dynamic processes in the structure, the authors are interpreting as more stable post-buckling behaviour beneficial for application. Comparing currently known test results of grid structures with load bearing skin, it is also concluded that interface layers are needed if a post-buckling panel state wants to be used.
Multifunctional Lattice Concept with pressurized inner Skin
The main purpose of Multifunctional Lattice Concept with Pressurized skin is to minimize the main shortcomings of conventional laminated composite structures.
Multifunctional Lattice structures have unidirectional layout of load-bearing elements, allowing to bear higher level of loading. The stress concentrations at the crossings are low due to extra-high volume ratio of fibres in these zones;
Multifunctional Lattice structures due to their topology have a capability to protect the primary structure against impacts and environmental factors with a slight weight penalty.
Fuselage Design Concept
Multifunctional Lattice composite fuselage structures have the structure layout that is substantially different from the conventional aircraft structures having load-bearing stiffened skin. In frames of Multifunctional Lattice concept the main primary structure element responsible for bearing moments and forces is a stiff grid on unidirectional composite ribs, while the skin located inside the grid is used for pressurizing only (Figure 24). The structure concept also has additional secondary structure elements, the main of which are: layer of protective foam and smooth outer skin (to form the aerodynamic shape).

Figure 24: Main components of Multifunctional Lattice structure concept
The proposed Multifunctional Lattice structure concept has the following main advantages (in comparison with conventional composite structures with laminated skin and stiffeners):
force flows in Multifunctional Lattice structure are directed mainly along the direction of carbon fibres that allow to realize higher levels of loading in the composite primary structure elements;
due to its topology the grid of unidirectional (UD) ribs (i.e. the primary structure) can be reliably protected from external and internal factors (e.g. impacts, environment, hot/wet etc.) with a slight weight penalty;
for some feasible combinations of design parameters, Multifunctional Lattice structure allows to realize windows’ cut-outs without “cutting” primary structure elements and thus avoid stress concentrations.

Manufacturing process
The currently available lattice composite technology was developed by CRISM (Khotkovo, Russia). This manufacturing technology, based on wet winding methods, has proven its high performance for rocket applications. But it does not allow realizing the Multifunctional UD-rib structure concept described above with the frames of one technological cycle, i.e. the main structure elements cannot be manufactured simultaneously so far. Thus, the proposed production process consists of the following stages:
manufacturing of lattice grid;
manufacturing of protection layers and inner skin;
installing the inner skin into the grid;
fastening of protection layers;
winding of the outer skin on the structure.
During these stages the manufacturing of the load-bearing lattice grid and the secondary elements are performed separately.
The lattice grid is manufactured using the production process developed by CRISM. This manufacturing technology is based on wet filament winding method. The lattice composite structure is winded into special preforms (substrates), which are put on the rotating mandrel (Figure 25).

Figure 25: Manufacturing of a lattice grid
The lattice grid here is a monolithic structure, while the inner skin and the foam plates are assembled from the fragments (panels). Foam plates are attached to the lattice grid by a glue attachment. The inner skin is fastened to the structure in the following way (Figure 26):
the fragments of the inner skin are attached to each other by the means of polymerization (vulcanization) to form the entire skin;
fuselage section is pressurized from the inside to force the proper installation of the inner skin.

Figure 26: Installation of inner skin and foam plates
After that, the energy-absorbing foam is placed on the outer surface of the structure, followed by forming the outer impact-resistance layer. The outer skin is made of composite material having elastic thermoplastic matrix and can be manufactured using out-of-autoclave technologies.
Generally, the inner skin could be also manufactured as a whole. In this case it should be used as a preform for winding, as the installation of such large-size structure element at this stage of manufacturing is very difficult to perform. In the case when the inner skin is assembled from fragments (as it is considered here) the adjacent fragments of the inner skin should be accurately bonded to provide pressurizing. This leads to a requirement to the structure materials for the inner skin. Such materials should allow the polymerization/curing of the fragments that can be provided locally and without influence on the other structure elements. One of the possible solutions can be using thermoplastic composites for the inner skin. These materials also have high reparability, which is crucially important for the pressurized inner skin.
It is worth to mention that in order to provide robust attachment of the outer skin and foam plates to the lattice grid the special fasteners can be installed into the structure at the stage of assembling. These fasteners can be realized e.g. by the special clamps in the zones of rib knots (intersections). As the outer skin is not included in the primary structure, the role of such clamps would be only to prevent the outer skin debonding caused by aerodynamic forces. Thus, such clamps can be lightweight and their amount can be considerably low.
Element specimen for property determination - Winded specimens
The experimental investigations of elementary specimens, simulating regular ribs of a lattice structure, have been performed by CRISM. The main objective of these investigations is to define the strength and stiffness characteristics of unidirectional ribs manufactured using current winding technology. This technology is usually used by CRISM for serial production of real lattice composite space structures. For manufacturing of specimens the real cylindrical section structure has geometrical parameters close to the real fuselage structure which was manufactured. The lattice elements (ribs) for the experimental study are cut from this section, upgraded and prepared for the testing (some attachments and local reinforcements were made for the testing bench). Therefore, the manufacturing process which is used for lattice ribs is very close to the manufacturing of full-scale UD-rib cylindrical fuselage structures.
Following specimens are cut off the lattice UD-rib section (see Figure 27): specimen of rib crossing zone (compression testing), regular rib zone (compression testing), long (elongated) rib with cut crossing ribs (tension testing).

Figure 27: Specimens of UD-rib structure
It is important to notice that the ultimate compressive and tensile strength of composite material in UD-ribs is significantly lower than the corresponding characteristics of unidirectional extruded CFRP. According to the obtained experimental results for UD-ribs σ_ult^-= 650.72 MPa and σ_ult^+= 1389.6 MPa, thus for the corresponding extruded CFRP these values are σ^- = 1300 MPa and σ^+ = 1700 MPa. The difference between these values (on approx.50% for compression and approx.20% for tension) is connected with different volume ratios of fibers V in the UD-ribs and the extruded CFRP. For UD-ribs the volume ratio is V = 40% for regular zones and V = 75% for crossings. So the average volume ratio of UD-ribs is lower than for extruded CFRP and hence the ultimate strength is lower.
Based on experimental results for ultimate stresses and elastic modules the values of ultimate strains for compression and tension can be easily found: ε- = 0.6 – 0.75% (compression), ε+ = 1.2 – 1.6% (tension).
The results show that the tensile strains of UD-ribs at failure are very close to the ultimate strains of carbon fibers, i.e. the failure of a UD-rib at tension is caused mainly by the failure of fibers. As UD-ribs have no orthogonal layers at regular zones, the authors assume that the failure of resin and fibers in these zones occurs simultaneously. Of course, airframe structures have complex loading and the values of maximum allowable strains for composite structures should be formulated taking into account influence of environmental factors on the primary structure.
One of the main advantages of UD-rib structures is the possibility to assure robust protection for the main primary structure element (UD-grid of ribs). The values of ultimate strains highly depend on the extent of their protection, i.e. how the typical impact decreases the strength characteristics of a rib. For an “undisturbed” UD-rib structure (i.e. the protection system fully absorb any impacts of typical energies of 35 J (from inside) and 50 J (from outside) the values of ultimate strains, proven by the experiments are the following:
εultimate UD-rib =
0.6% - compression;
0.7-0.8% - tension.
But these levels can be used only for fully protected Multifunctional UD-rib structures. The main question here is the minimal weight expenses needed to build such protection system. In the next chapter the main activities on development of such protection system for UD-rib structure with advanced robustness are described.
It is important to notice that if the developed protection system would presume that the rib can be damaged to a certain extent in the case of impact of typical energy, these values should be corrected and decreased.
Impact Results on Ribs with Protection
One of the shortcomings of lattice structures based on unidirectional composite ribs is their high sensitivity to impacts. The influence of impacts is even more critical than for laminated composites. The experiments show that a typical composite UD-rib loses almost 100% of its strength after an impact with energy of 15 J.

Figure 28: Specimens of lattice ribs with and without protection after impact of 15J
For conventional laminated composite fuselage structures the negative influence of impacts is usually compensated by enlarging the thickness of the skin. This cannot be called an effective way to prevent dangerous consequences of damages on thin laminated skin. But the other way to solve this problem - protection of the load-bearing structure from impacts, would mean covering the whole skin by a protective layer, and thus considerably increasing the weight of the structure.
For a Multifunctional Lattice structure enlarging of geometrical parameters of ribs makes no sense and the only way to provide its long-term operation is the full and robust protection of composite UD-ribs.
Fortunately, the lattice topology of the grid allows to provide such protection with a slight weight expenses. Figure 28 shows the results of experimental investigation of impact strength of two lattice elements (with and without protective layer) after impact of 15J. The rib without protection was totally destroyed after the impact. The rib itself absorbs only 3.1 J (20.7% of total impact energy). The protected rib almost totally saved its integrity after impact and absorbed about 12.4 J (82.7% of total impact energy). For this test the weight of a protective layer was about 20% of weight of the rib.
The results of a number of tests described above have shown that the UD-rib structures can be reliably protected against impacts with slight weight expenses.
Based on these promising assumptions, a number of investigations dedicated to development of structure concept selection of rational materials for the protection system of UD-rib structure have been carried out in TsAGI and VIAM within the frames of FP7 PoLaRBEAR project.
The developed concept of the protective system (including two variants) for lattice rib is shown on Figure 29.

Figure 29: Concept of protective system for ribs in Multifunctional Lattice structure
The protection system has the following main elements:
outer skin with smooth surface made of thin CRFP-composite;
impact-resistant layer made of fabric of high-strength fibres (e.g. Aramid fibres);
energy absorbing layer;
primary protection layer;
lightweight filler between two skin;
Impact-resistant elastic inner skin.
The materials for different elements of the protection system are selected on the basis of the current (existing) materials developed by VIAM for other applications. The investigations of the rational structure concept for the protective system of UD-rib fuselage structure are started in TsAGI.
According to the weight comparison made within the frames of ALaSCA, the weight of the protection system together with inner/outer skins should not exceed 40% of total weight of the barrel in order to obtain weight decrease not less than 10% in comparison with metallic and composite “Black metal” analogues. The preliminary results show that this task can be effectively solved using current (existing) materials.
Repair Concept for Multifunctional Lattice Structure
Providing reparability of lattice fuselage structure is necessary for preservation of structure strength characteristics during long-term operation. Reparability, as a rule, is reached by increase of structure technological effectiveness, including:
Increase of controllability,
Increase of accessibility to damaged details,
Providing interchangeability of damaged details.
For damages of categories 1-2 (damages of outer and internal skins) for which repair is carried out on the basis of bonded connection, providing reparability consists only in providing controllability, first of all for internal elastic skins. Taking into account that elastic skins of lattice fuselage structure have quite small thickness, controllability is provided by occurrence of visible damages on skin surface even at small impact energy. Such damages represent dents, delamination or through holes.
Damages of categories 3-5 (damages of UD-ribs or fragments of load-bearing UD-rib structure) will demand repair of UD-ribs, structure fragments and skins. Controllability in this case is provided at the expense of a skin - it plays a role of the indicator of damage. However, the structure health monitoring system is necessary for obtaining the information on damage category (whether the UD-ribs are damaged) as a visual control in this case is inapplicable. Accessibility of the damaged structure in case of damages of categories 3-5 is provided by skin disassembly.
For reparability of UD-rib structure at damages of categories 3-5 it is necessary to provide interchangeability of structure elements, or possibility of restoration of initial properties of its fragments. For this purpose, it is necessary to provide possibility of bolted connections, as the basic repair method.
The main difficulty in providing reparability of the UD-rib structure is the impossibility of direct bolted connections of ribs as any holes in load-bearing composite elements with unidirectional package will lead to considerable loss of their load-bearing capability, and stress/strain concentration will extend on the most structure part.
For repair of UD-ribs a bolted and bonded connections can be used. Unfortunately, strength characteristics of glues are quite low and as a result there is a constraint on the minimum length of bonded patch for bonded repair. The numerical investigations were performed in order to determine the dependence of minimum length of bonded patch on the UD-rib parameters (Figure 30). The results of calculations show that for the rational ribs parameters (rib cross-section area is ~200-250 mm2 and average rib length ~350 mm) the glues with strength characteristics ϭbond<5 kg/mm2 will not allow to restore the initial strength parameters of UD-ribs.

Figure 30 – Constraint for bonded repair of UD-ribs
Bolted repair is currently considered as the basic repair method of the UD-rib structure. Embedded fixing elements are necessary in zones of rib crossings for carrying out UD-ribs repair based on bolted connections. In order to restore initial structure characteristics, titanic patches on the damaged ribs from the outer and internal sides of lattice structure are used. Patches are pulled together with bolts through embedded elements (bushes) in zones of ribs crossings (Figure 31).

Figure 31 – Details for bolted repair of damager UD-rib
Two variants of repair connection have been considered:
Patches are glued to ribs,
Patches are not glued to ribs.
Results of calculations have shown that glue connection of bushes to ribs allows to lower stress/strain concentrations of in a zone of contact of bushes with ribs. Thus values of stresses/strains do not exceed ultimate values. The bolted repair method allows to restore completely initial strength and stiffness properties of the structure.
At the most severe damage of a category 5 replacement of the damaged structure fragment by a new one can be required. For this reason, the fuselage structure should consist of some separate panels. Light and reliable longitudinal and cross-section joints are necessary for providing connection of lattice composite fuselage panels.
In the presence of light and reliable longitudinal and cross-section joints of lattice panels it could be possible to carry out not only replacement of the damaged panels, but also to assemble the fuselage section from lattice panels with different design parameters that would allow to raise even more weight efficiency of the lattice fuselage structure.
Bending test of Full-Size Lattice Panels
One of the main advantages of Multifunctional Lattice structure Concept is that the composite elements are loaded almost only by compression/tension forces directed along high-strength carbon fibers. This assumption should be proved by tests of real full-scale UD-rib structures under real boundary conditions in order to assure that the results obtained on ideally straight specimens are correct also for real structures.
Such experimental investigations have been carried out in TsAGI on full-scale lattice composite panels (Figure 37). For these panels the specialized testing facility was built, which allow to provide real boundary conditions for lattice composite fuselage panels.

Figure 37: Experimental facility for testing of full-scale lattice fuselage panels under real boundary conditions
It is important to notice that the current lattice structures are integral, i.e. they are manufactured as a whole, unlike conventional fuselage structure assembled from panels. That means that the term “panel” for lattice structure doesn’t mean any technological unit and is applied only to describe some selected/considered part of the lattice structure. The experimental investigation of lattice panels is very difficult from the viewpoint of providing correct boundary conditions.
Because of that, the investigated experimental panel is considered within the frames of the lattice composite shell with outer composite skin fixed into the experimental facility. The tests are accompanied by the numerical FE-analysis. The obtained experimental data is used for the validation of the FE-modeling methods. The values measured by the strain gauges are compared with the values of strains at the corresponding elements of the FE model. Figure 38 shows the comparative analysis of experimental and calculation data for the considered experimental panel.

Figure 38: Results of model validation on experimental panel
The validation has shown that the error of the numerical analysis does not exceed 10% for all of the control points of the panel. In some points the errors are bigger and can reach up to 20%, but most of these zones are slightly loaded, so the errors are mainly caused by the low accuracy of the strain gauges at low levels of strains.
The test results of the real full-scale panel put into real boundary conditions have shown that the UD-rib panel with rectangular grid can be modeled good enough (error less than 10%). The FE model is based on 2D shell elements, where the local strain distribution in ribs is averaged. Thus the local behavior of ribs in the regular UD-structure with real boundary conditions is very close to the local behavior of the straight specimens. That means that the results of tests on specimens of the UD-structure are correct for the real Multifunctional Lattice UD-rib structures.
CONCLUSION AND OUTLOOK
In the level 1 project PoLaRBEAR, follower of EU-ALaSCA, the potential of lattice airframe architecture has been demonstrated. Different design concepts are developed, showing a wide variety of possible approaches. These concepts consist of the primary stiffening airframe structure as well as design solutions on component level like window cut-outs, barrel-barrel-interface and barrel-floor-interface.
In order to design these concepts, automated sizing algorithms are developed like a 4-level iterative sizing process for analysing from fuselage section level to the stress-strain state on element level. On section level a parameterized optimisation procedure has been set up with automated finite element pre-, solve and post-processing. On element level a new semi-analytical method has been developed to determine the buckling factor of rectangular and non-rectangular skin-bays (triangular, diamond, kagome).
The investigation of the impact of the weight reduction on aircraft level shows that a mass reduction of 10% of primary fuselage structure, results in 1% overall fuel consumption reduction for the flight mission of a short and middle range aircraft.
In the PoLaRBEAR project more specific investigations, including questions of damage tolerance and repair, were performed for Anisogrid structures. In different investigations the Technological Readiness Level could be increased as intended.
On element level static and also impact tests were performed for winded and for Prepreg design specimens. The winded specimens achieve remarkably high tension and compression properties corresponding to the fibre volume ratio of the regular rib. To overcome the impact sensitivity of the unidirectional material a protection system was proposed and investigated in first impact tests. The Prepreg rib and knot specimens have shown the expected high stiffness properties for tension as well as for compression corresponding to the high fibre volume ratio. As drawback, a remarkable reduction of the strength properties has to be considered due to the amount of the cut plies in the knot area.
For anisogrid structures with load bearing skin and triangular skin-bays a different post-buckling behaviour was observed in numerical analysis in terms of missing buckle mode switches. This post-buckling behaviour could be successfully validated with a panel compression test. It could also be shown that the proposed design with interface layers is providing a strong connection of the skin and the rib structure.
In a full-scale barrel test with 4 m diameter and 6 m length the winded lattice barrel concept could be successfully validated in terms of demonstration and very important for the numerical models generated. The difference of the numerical model and the strains measured is for the whole barrel structure within 10% and satisfying in full extent.
A new shell element formulation for lattice structures is developed which allows faster and more robust numerical simulations compared to conventional used FE shell elements while having the same accuracy. A new method to assess the structural robustness of stiffened shells has been successfully applied to Anisogrid stiffened barrel structures and can be implemented now within a numerical optimization loop.

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Potential Impact:
3 Impact
3.1 Expected impacts listed in the work programme
The proposal will further enhance the cooperation in research and in innovation between the EU and the Russian Federation in the field of civil transport aircraft.
Based on an existing network and on existing knowledge developed over the last two decades, especially in the EU-Russia-FP7-project ALaSCA, the international cooperation between the EU and Russian partners gives the opportunity:
• To increase the market appeal through innovations and increased competitiveness
• To acquire and access science and technology that are complementary to current European knowledge and of mutual benefit
• To develop technologies which fulfil global needs, e.g. climate change

More details about the topics, which are investigated in international co-operation between European and Russian partners, are given in the Work Program. The current proposal relates to topic AAT.2013.8-1 – Domain 1: “Reliable novel composite aircraft structures based on geodesic technology”. The aim is to promote an effective cooperation in the development of light, low-cost airframe fuselage structures made of new generation of composite materials based on geodesic / isogrid technologies addressing:
• An increase of Technology Readiness Level (TRL) of geodesic technology
• An understanding of the local structural behaviour
• The investigation of a robust design, its manufacturing and the local structural behaviour of grid intersections and rib-skin interface, as well as repair methods

In agreement with the strategic research agenda of ACARE and the “vision of 2020” report, the PoLaRBEAR project thoroughly focuses on top level impacts, which target society’s needs for:
• The greening of air transport
• Safer and environment friendly air transport
• Improving cost efficiency
• Pioneering the air transport of the future

The PoLaRBEAR proposal is based on the results of the ALaSCA project. The impact of the ALaSCA results, which is related to the European and Russian needs in aeronautical development, is:
• Pro-composite design leads to an adequate structural exploitation of composite materials and therefore to a weight saving in the composite fuselage structure of 25 percent (greening).
• Increasing safety by damage tolerant design. (safety)
• Winding technology is a highly automated manufacturing process yielding high output and strongly reduced manufacturing costs. (cost efficiency)
• Pro-composite design, using lattice technology, design of aircraft concepts and manufacturing, leads to efficient pro-composite driven novel aircraft configurations based on lattice fuselage structures (greening, cost efficiency, pioneering aircraft structures)

PoLaRBEAR’s expected impact, which is related to the European and Russian needs in aeronautical development, is:
• Maximum weight & cost reduction of fuselage structures through geodesic designs.
• Development of manufacture-optimized geodesic designs satisfying airworthiness requirements.
• Verification of airworthiness by manufacture and testing of representative geodesic components.
• Strongly improving avenues of industrial exploitation of geodesic structures.

In the Work Programme 2013 “Cooperation Theme 7 TRANSPORT” PoLaRBEAR also addresses challenge 3: The strengthening of the competitiveness of European transport industry through innovation, as competition from developed and emerging economies, is intensifying in a global economy. The game changing geodesic structure is fully in line with this challenge.

In this challenge, the expected impacts for ACTIVITY 7.1.4: “Improving Cost Efficiency” are addressing the following objectives for the technology readiness by 2020:
• To reduce aircraft development costs by 50%
• To create a competitive supply chain able to halve time to market
• To reduce travel charges

The aim stated in ACTIVITY 7.1.4 is to ensure cost efficiency in air transport by focussing on the reduction of aircraft acquisition costs. Concepts, innovative solutions and technologies must result in lower lead time, and lower aircraft costs and more efficient systems from design to production, including certification, resulting in a more competitive supply chain.

3.1.1 Strategic impact of PoLaRBEAR
Numbers of flights offered on passenger routes will more than double. This is a more rapid rise than in previous years and will, given current levels of congestion and delays, present a continued challenge to the world’s airports and air traffic management systems. World jet aircraft size, including regional jets, will increase by 20% over the next 20 years, as a result of increased congestion, diminishing returns of traffic stimulation from increased frequencies and the overall growth of the fleet. Accordingly to aircraft manufacturer, over 22.000 new passenger aircraft and freighters will be required over the next 20 years (Fig. 26).
The environmental (noise, place requirement) and ecological burden is increasing synchronously but is already eminent today, as transportation and traffic account for ca. 20% of the total CO2 emissions. Further, the rising cost of fuel, whose consumption is directly linked to the CO2 emissions, increases operating costs significantly. All these factors contribute to an ever-increasing demand for more fuel-efficient aircraft technology.
One of the most promising ways to reduce fuel consumption (and in consequence operating costs) and environmental impact of aircraft is to reduce weight through the increased use of carbon composite structures (chapter 1.2.1.1 and Fig. 34 respectively).

Fig. 34: Development of carbon composite share of structural weight in aircraft [21]

Both the increasing number of required new aircraft during the next decade and the increasing percentage of carbon composite parts demands for pro-composite design to exploit the potential of composite material and for a dedicated production process to drastically reduce production costs. As an example AIRBUS’ A320 replacement called NSR is baseline against the A320-200 for overall performance and cost terms and is aimed at a provisional service-entry date of 2012-13. AIRBUS NSR, Phase 1 results are: the specific fuel consumption reduction would be 4%, the best operating cost reduction 3% and the best emissions reduction would be 5%. Game changers in design and manufacturing processes are needed.
Lattice structures are very close to a pro-composite design (see section 1.1) having the potential to be the game changer. The development of manufacture-optimized lattice designs satisfying airworthiness requirements will have a crucial impact on weight and cost of future fuselage structures.

PoLaRBEAR is a cooperative project that is clearly aligned with these objectives. Its primary impacts the strengthening of competitiveness of European and Russian aerospace industries and associated.

PoLaRBEAR’s direct impact concerning the European and Russian needs in aeronautical development by the end of 2017 is:

• Greening
o Lattice lay-outs satisfy the composite design requirements in a nearly perfect way and will therefore lead to an eminent weight saving ratio of the composite fuselage structure by 25 percent.
o Investigation of new, environmentally more acceptable manufacturing techniques and processes.
o Further decrease of weight through synergy effects.
o Special effort in developing structures with long-run durability causes low energy consumption seen over the whole manufacturing and life cycle.

• Increasing safety
o Increased safety by damage tolerant design providing redundant load paths around area of destruction.
o Damage tolerance and fatigue tests to ensure an improved operational safety of civil aircraft.

• Cost efficiency
o Highly automated manufacture yielding high output and strongly reduced manufacturing costs.
o Integral construction, strongly reduced number of joining elements leading to drastically reduced assembly cost.
o Reduction of operating cost by decrease of fuel consumption or increase of payload through weight-saving potentials of lattice fuselage.

PoLaRBEAR’s indirect impact responding to societal needs by the year 2017 is:
• improvement of working conditions by introduction of more automated processes while minimizing contact with and exposure to potential irritating agents such as resin (in case of preform-infusion technology).
• Support of green aircraft in terms of chromate free structures is given, the structural weight reduction ideas to reduce fuel consumption and substantially decreased emissions.
• Implementation of technological progresses in materials and manufacturing techniques increasing the skillfulness of the European workforce and securing the employment level in Europe.

PoLaRBEAR’s impact on global topics by the year 2017:
• offering an excellent technological basis for lighter, more efficient and safer aircrafts
• Promoting the European – Russian cooperation of research and industry,
• Ensuring and increasing of European employment
• Improving the knowledge of researchers and engineers
• Enhancing the competitiveness of the European and Russian aircraft industries by strong exploitation.

PoLaRBEAR’s future direct impact until 2035 is expected in:
• Provision of further adapted and optimized manufacturing methods regarding lattice elements and structures under airworthiness aspects (short-term)
• Certification of lattice elements and structures for aircraft applications (mid-term)
• Manufacturing and testing of complete lattice fuselage barrels (long-term)
• First flight of lattice-relevant aircraft (fuselage barrel and beyond)

Furthermore PoLaRBEAR is expected to significantly influence light weight design in other fields of application and will have a future indirect impact on e.g. architecture, wind energy, transportation, i.e. especially railway or automotive structures etc.

PoLaRBEAR’s expected indirect impact responding to societal needs will result in much more eco-friendly, sustainable and efficient air and ground transport fulfilling the needs of future societies in a globalized world.

3.1.2 Added value due to European and Russian level work
The project could not be carried out as quickly and efficiently without a strong interdisciplinary European and Russian consortium. Indeed, it requires capabilities and expertise that are spread over several countries, and all of these are necessary to gain the maximal benefit of the project. It is remarkable that the leading European aeronautical industry combine their effort to make a major step forward in a cost efficient production of carbon composite aero structures.
SME’s are regarded as potential suppliers of composites and tools. Generation of new jobs through expansion of SMEs is possible. All types of entities are represented: SMEs, universities, research centres and industrial partners, and certainly the differences in approach will enrich the project. However, the main drivers of the project are the industrial partners who urgently need an answer to the questions posed by the PoLaRBEAR proposal.
The proposed PoLaRBEAR framework is based on the development of cooperative research topics to further improve, in the medium term, the technology base and develop innovative aspects. Therefore, the present consortium of the project clearly indicates the level of the expected strategic impact.

3.1.3 Related national and international research activities
PoLaRBEAR is embedded in a long standing and continuous research on carbon composite aircraft components. Up to now this research mainly has been focussing on the performance and design principles of the structures, e.g. the black fuselage in CFK-Rumpf NG, the carbon composite wing in the IP ALCAS, or singularities in MAAXIMUS. Thus all these projects are mainly targeting the structural weight reduction. In the following related activities are listed:

Acronym Funding PoLaRBEAR-related key features Content
European projects
ALCAS European 6th FP Fuselage design (WP2), interfaces (WP3) Advanced low cost Airframe Structures
BOJCAS FP5 Interfaces (WP3) Design of composites joints / damage criteria for composites
COCOMAT European 6th FP Numerical analysis of fuselage structures Improved MATerial Exploitation at Safe Design of COmposite Airframe Structures by Accurate Simulation of COllapse
DCF German/Dutch project under contract with AIRBUS Hamburg Experiences for stability analysis Design Components Fast Buckling Handbook (ESA, ECSS-E-30-24)
MOJO European 6th FP Interfaces (WP3) Modular Joints for Aircraft Composite Structures
MAAXIMUS European 7th FP Fuselage design, requirements, virtual testing (WP1-4) More Affordable Aircraft Structure Lifecycle through extended, Integrated and Mature Numerical Sizing
ALaSCA European 7th FP Novel aircraft and fuselage structures Lattice structures from aircraft to element level
National projects
CFK-Rumpf NG DLR Fuselage design concepts (WP2,4) Future fuselage design and manufacturing aspects, etc.
TopDesign LuFo, BMWi Design and manufacturing of PAX surround door structure Design and manufacturing of PAX surround door structure
Tab. 4: Other related national and international research activities

3.2 Dissemination and exploitation of Project results
The dissemination and exploitation of the results of the PoLaRBEAR project has three levels:
• Education and lecturing
• Scientific publications
• Industrial application
The composition of partners coming from industry (suppliers and end-users), research establishments and universities is a very good basis for a well balanced consideration of all three fields.
The contribution of the major European aeronautic industries guarantees the rapid exploitation of the project results. Nevertheless, it is planned that the consortium will implement internal structures to exploit the carbon composite lattice technology further developed under the FP7.
3.2.1 Exploitation manager
This action will be headed and supervised by the Project exploitation manager. The exploitation manager will be designated by the project Coordinators at the start of the project. He will be responsible for the following activities:
• To summarize the original and updated plans to the overall projects summary exploitation and dissemination report.
• To make sure that all results will find their way directly to the lecturing content of the universities and to the development departments of the companies in order to allow a quick and broad utilisation.

3.2.2 Exploiting mechanism
Formal mechanism for exploiting all technologies and capabilities is outlined below:
• Market specific aeronautic technologies are exploited through a panel comprising the industrial partners and headed by the exploitation manager.
• Enable a smooth interaction between the consortium members providing the innovations through the exploitation panel.
• Extension of the methods developed in PoLaRBEAR to other industrial segments like automotive industry or spacecraft manufacturers. This will be guaranteed by the consortium coordinator who is cross linked to the related European industry.
Project results will be disseminated at two levels: detailed and general.
At the detailed level technical reports will be provided to the exploitation manager and, after content check and permission of the creator, distributed between all the project partners. Those reports will describe the technical details and the conclusions of the research.
At the global level DLR will organize a scientific workshop at the end of the PoLaRBEAR project. That will ensure that the increased technological capability of European and Russian site and its partners in taking full advantage of lattice design superior characteristics becomes evident to the world-wide aerospace community. Any dissemination will be made in strictly accordance with the agreement of every single participant. Any communication (paper, data sheets, etc.) will be submitted to all consortium members for approval.
Followed are listed the different possible ways how the dissemination will be raised by the PoLaRBEAR consortium:
• Scientific publications in international journals and presentations at international conferences
• Presentation at appropriate international conferences
• Invention disclosure, patent applications considering the European and Russian law in particular
• At the end of the project the project coordinator will organize a scientific international workshop on carbon composite lattice design.
• Producing flyers, posters and videos about the PoLaRBEAR project.

3.2.3 Dissemination and exploitation of results
The results will comprise novel composite fuselage concepts, reliable methods of designing and design guidelines and experimental database on characteristics of specially obtained composite materials oriented to high-loaded fuselage structures. Publications based on PoLaRBEAR work, close interaction between industry and research, PoLaRBEAR partners, will ensure wide dissemination and exploitation. In particular, the following activities will be performed for exploitation and dissemination of results:
Exploitation:
• Direct exploitation by the end users in future aircraft designing.
• Development of know-how and capability by PoLaRBEAR’s research organizations for future research and business.

Dissemination:
• Passing on knowledge to the public-at-large through:
• Training at the research organizations;
• Publication of the main achievements on the PoLaRBEAR web page;
• Using the instruments of European Aeronautics Science Network (EASN);
• Hosting workshops;
• Publication of papers in international journals;
• Presentation of papers to international conferences and meetings is planned (e.g. in a choice from the following: AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference; AIAA Multidisciplinary Analysis and Optimization symposium; AIAA SDM Conference; COMPOSITE-EXPO; DGLR; JEC; ILA; SAMPE; World Congress of Structural and Multidisciplinary Optimization)

The following table shows how the partners will exploit the project results:
Partner General exploitation & specific needs Timetable for use and expected market size
Research Centres
(DLR, TsAGI, VIAM) Mission: identify, develops and applies high-tech knowledge in the aerospace sector. Bridging the gap between Research & Industry
Through the research on lattice structures, these Research Centres will increase & their strengthen capability to support their (National and European) aerospace industry. This ranges from the identification of possible applications, through the design and analysis up to and including manufacturing techniques and industrialisation. For use to strengthen position with respect to aircraft materials / composites: as of now.
For commercial use by the aerospace industry: 2020 - 2030
Aero Industry (including SME):
(AIRBUS, CRISM, NIK, SMR, MSC) New or more competitive products
The use of lattice fuselage structures and the ability to design, size, manufacture, test and certify these (fuselage) design concepts open the way to new CFRP applications, currently still ruled by metallic materials. Targeted applications are in fuselage structures optimized in terms of cost, weight and safety. The potential market is depicted as medium to highly loaded safety critical parts on the future replacements of the A320 and B737 single aisles.
Universities:
ULeeds, TUBS, MIPT, MUCTR,
QMUL Expanding expertise
Increase level of scientific competences and dissemination of knowledge From 2013

3.2.4 Management of intellectual property
The rules for industrial property rights and exploitation of results will be defined in a consortium agreement signed by all partners. The principles of the intellectual properties will be the following:
• Any invention generated by a partner is the property of the partner. Every participant is obligated to communicate made inventions and declared patens and give a abstract of content.
• If any joint invention results from the co-operation under the consortium agreement, i.e. joint inventions made by employees of more than one partner, and if the features of that joint invention cannot patented separately, the parties concerned can jointly apply for patent protection.
• A continuous and substantial interaction between the personnel of the partners will take place during the project. The partners will agree that the partner generating foreground information shall own any this information but all partners shall be entitled to use such information without any financial compensation to or the consent of the owing partner.
• A considerable amount of background information is being brought into the project, which will remain the property of the specific partner and may, in some cases, be commercially confidential. This specific know-how is owned by the specific partner and will not be shared within this project. The exception is the transfer of know-how in order to enable a cooperating partner to perform tasks vital to achieve the project goals. The legal modalities are usually sorted out on bilateral level. In order to gain maximum synergy all results and developments obtained by the partners within the project will be made available among the partners in the consortium. This includes all manufacturing, test and simulation results. Some results, which are obtained after having applied new methods on company specific elements or subcomponents, might be confidential due to competitiveness reasons. However, the dissemination level of the associated deliverables and its confidentiality ranking are regulated in the list of deliverables. It is worthwhile to note that such technological development as it will be performed in the PoLaRBEAR project must involve a significant collaborative effort, with each partner engaged on a particular element of the task, which may be eventually assembled to address the overall objectives. But please keep in mind:

A project of the PoLaRBEAR nature can only succeed if data are exchanged openly between the partners but limiting the disclosure of the information to the required dissemination to the public.

Summary of the dissemination and exploitation
• PoLaRBEAR will establish an exploitation panel comprising the industrial partners headed by the DLR research entity.
• PoLaRBEAR will continuously develop and update the exploitation plan.
• PoLaRBEAR will transfer the results to other industrial segments like automotive industry
• PoLaRBEAR will disseminate the project results through scientific publications, an international workshop, via an internet homepage, flyers, posters and videos about the PoLaRBEAR project.

List of Websites:
Prof. Dr.-Ing. Christian Hühne - [e-mail]
MSc. Steffen Niemann - [e-mail]