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This paper presents a structure design and optimization module, developed for the application inside multi-disciplinary optimization process chains. The loads necessary for sizing and optimization are calculated in CFD or aeroelastic calculations and applied on a Finite Element Model that represents all primary structural elements of a wing. The FE model is created automatically from a parametric geometry description. The deformations and inner loads of the wingbox are calculated via linear static FE calculations; geometry and loads are provided to an external sizing tool. Each component has a set of design variables with discrete design points which are permutated to get the component’s design candidates. A set of failure criteria is used to size the structure which can be made of composites or metal. The methodology is applied to the optimization of a forward swept composite for a short range aircraft and a design study comparing different stringer types and their influence on mass and structural deformation of the wing is performed.

Since 2011 DLR is preparing 3 orbit demonstrations of deployment and operation of a solar sail. For this mission bundle a new deployment control mechanism for DLR’s tubular CFRP shell booms has been developed that is applicable to all 3 evolution steps of the GOSSAMER—Road Map. According to the mission goal of GOSSAMER-1 and its addressed deployed sail size of 5 × 5 m2, a down scaled configuration of the booms and an adapted deployment mechanism are currently under development. The paper introduces the boom and its deployment mechanism concepts and describes and concludes tests and simulations that have been performed to proof the mechanical performance of the deployed booms.

In this paper a novel simplified structural modelling approach for stiffened panels is presented. It specifically addresses structural response of aircraft fuselage components made from carbon fibre reinforced plastics (CFRP). The proposed method considers structural modelling requirements necessary for structural optimisation with respect to system response under simultaneous static and dynamic loads. The standard explicit finite element (FE) solver LS-Dyna is utilised, providing the framework to regard statics and dynamics as well as non-linearities by material, large deformation and contact. The simplified model is fundamentally composed of shell, beam, and discrete elements. Simulating the structural collapse behaviour of thin-walled CFRP sections under bending load the detailed model provides input parameters for the discrete elements. Furthermore, skinframe- interaction attributes immanent to the present fuselage configuration are reviewed and considered by idealised deformation characteristics. Finally, the present approach is verified through analyses of the detailed model. The results illustrate the benefits of model simplification. Replacing detailed shell model features by beam and non-linear spring elements significantly reduces computational effort. At the same time the accuracy deduction is acceptable considering topological optimisation requirements.

Deorbitsail is a 3U Cubesat project that will launch, deploy, and support a large drag sail deorbiting payload. The flight payload sail system will deploy to 5 by 5 meters, increasing the frontal area of the spacecraft dramatically. A series of 4-by-4-meter sail deployments has been performed with a prototype of the proposed system. The prototype system tests included partial deployments in environmental chambers and full-scale deployment at ambient lab conditions. Design changes arising from testing of the sail system prototype are described in this paper. Engineering model construction is underway and the Deorbitsail launch will be in 2014. Deorbitsail is a European Commission 7th Framework Programme project with nine partner organizations. © 2013 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Increasing environmental, economical, and social issues force future car concepts to strive for maximum efficiency. As metal designs have reached a very high level of maturity, further potentials are seen especially with extremely lightweight (carbon) fiber reinforced composites (CFRP) exhibiting high strength and stiffness, which are advantageously integrated into multi-material designs. This section focuses on improvements in weight reduction, safety and modularisation strategies for future cars. A novel Rib and Space-Frame concept incorporating carbon fiber composites is presented. Herein, an essential component replacing the former B-pillar is the B-rib, which uses a novel mechanical principle to meet the side impact crash requirements. Furthermore, using the Rib and Space-Frame concept as an example, the entire process chain of fiber composites—from materials to design, sizing, prototype manufacture, and testing—is being presented also opening up perspectives for future mass production strategies.

Today’s bonded composite repairs rely heavily on manual grinding. The human influence on scarf tolerances and consequently on assembly and structural performance can be reduced by automating the process of scarf manufacturing. A process based on contact free surface scanning, surface reconstruction, automated repair design and automated milling of the repair scarf is presented. A machine and software design for validation purposes is described. Several repair specific design considerations relevant for the construction of a mobile scarfing machine are discussed. The redesign of a standard 3-axis milling machine to a mobile automated scarfing unit is presented and the architecture of the associated software framework is described. An outlook to future validation steps is given.

The aggregation of functionalities offers additional benefits to the customers such as reduced weight, reduced life cycle costs and an increased range of applications. For a compliant aggregation of functionalities according to given requirements clear instructions on how to conduct lightweight design are essential, but often not available today. High performance lightweight structures are made from carbon fiber reinforced plastics increasingly. Due to the specific composite manufacturing process four different levels of function-integration are conceivable. The pre-fabrics or components of the composite can include smart materials with enhanced functionalities. The structure design can better exploit the composite potentials of anisotropic material properties. Passive components integrated into the structure provide additional functionalities as for example de-icing and lighting protection. In adaptive systems active elements significantly improves the ability of the structure to adapt changing environmental conditions. The development of the potentials resulting from the compliant aggregation of functionalities is presented in this chapter.

In comparison to other transport systems, launch vehicles are characterized by relatively light but extremely valuable payloads. The launcher’s upper stage structures, e.g. payload adapter and fairing, offer the highest weight saving potential. An effective weight reduction can only be achieved by the combined utilization of high performance materials and adapted construction methods. To improve the structures damage tolerance a new hybrid lay-up has been developed, which combines the properties of both, steel and carbon fiber reinforced plastics (CFRP). This chapter presents a preliminary design of a payload adapter as a framework, which is based on the high performance material properties of unidirectional CFRP-steel-laminates, offering a considerable weight saving potential.

A straightforward approach to predict spring-in deformations of angled composite parts is presented. Therefore, a proposal by Radford is extended in order to calculate the spring-in contribution due to chemical shrinkage. For this, the volumetric shrinkage of neat thermoset resin, which is in the range of 2–7%, is transferred to equivalent strains on ply level assuming no shrinkage in fiber direction. As the fiber volume fraction (FVF) affects mechanical and chemical properties significantly, the spring-in angle is affected as well. Therefore, the numerical investigation accounts for the spring-in angle and its thermal and chemical contributions depending on the FVF. Classical laminate theory (CLT) is utilized to homogenize layup expansion and shrinkage properties. For validation purposes, model predictions are compared with measurement results gained from one manufactured test specimen. Good agreement between analytical and experimental results is found. Furthermore, the chemical contribution of the total spring-in angle ∆φ turned out to be significantly larger than the thermal contribution.

The manufacturing of composite structures is accompanied by fabrication induced deformations. Those deformations are undesirable and lead to transgression of geometric tolerances in the finished parts. In order to get the part within aspired dimensional tolerances, geometrical compensation of the tool is necessary. This often iterative conducted tooling-rework is commonly time consuming and costly. This paper presents an shell element based. semi-analytical simulation approach focusing on warpage deformations due to tool part interaction, in order to account for manufacturing induced deformations within the tool design process. Deviation measurements on test specimen level serve as inputs for the calculation of equivalent coefficients of thermal expansion according to the proposed analytical model. Thus, ‘warpage properties’ of different prepreg – tool–material combinations are determined. The use and the practicability of the developed approach is demonstrated by means of the calculation of a warpage compensated tool surface.

A semi-analytic probabilistic procedure is presented that enables a fast prediction of the stochastic distribution of buckling load of cylindrical shells caused by the scattering of geometry, wall-thickness, material properties and loading imperfection. Compared to Monte Carlo simulations the semi-analytic method requires significantly fewer buckling load calculations giving equally accurate results. Knowing the distribution function of buckling load, a level of reliability is chosen and a design load is defined, which is lower than all test results and less conservative than NASA SP-8007.

Deployable structures are required to either enable orbit transfer of very large structure or to make the orbit transfer of medium and small size structures more affordable. Hereby, deployable booms are basic building blocks of such deployable structures. DLR is providing a concept for deployable booms that utilize very thin CFRP material and which can be stowed by coiling. The given paper introduces the concept of the CFRP booms and discusses the problems of their self deployment tendency. Furthermore, different mechanisms are presented that are able to control the deployment. Tests under artificial zero-g environment have been conducted to verify the applicability of the control concepts. Hence, the paper also gives insight in objectives, setup and the results of the experiment as well as a final evaluation of the concepts. Finally, an outlook on current and future projects that use the introduced booms or equivalent systems is given.

The load carrying capability of axially compressed cylindrical shells is dependent on imperfections like geometric deviations from the perfect shell or loading imperfections. The scattering of imperfections induces that the buckling load is randomly distributed. Knowledge about the distribution of buckling load allows an efficient and save design of cylindrical shells. The stochastic distribution can be predicted with purely numerical methods like Monte Carlo simulation, which require a multitude of buckling load calculations. In the present paper a fast semi-analytic procedure is presented, that predicts the distribution of buckling load with the same accuracy as a Monte Carlo simulation, but requires much less computational effort.

The complex three-dimensional structure of textile composites makes the experimental determination of the material parameters very difficult. Not only the number of constants increases, but especially through-thickness parameters are hardly quantifiable. Therefore an information-passing multiscale approach for computation of textile composites is presented as an enhancement of tests, but also as an alternative to tests. The multiscale approach consists of three scales and includes unit cells on micro- and mesoscale. With the micromechanical unit cell stiffnesses and strengths of unidirectional fiber bundle material can be determined. The mesomechanical unit cell describes the fiber architecture of the textile composite and provides stiffnesses and strengths for computations on macroscale. By comparison of test data and results of numerical analysis the numerical models are validated.

Providing the possibility to develop energy-efficient, lightweight adaptive components, Pressure-Actuated Cellular Structures (PACS) are primarily conceived for aeronautics applications. The realization of shape-variable flaps and even airfoils provides the potential to safe weight, increase aerodynamic efficiency and enhance agility. The herein presented holistic design process points out and describes the necessary steps for designing a real-life PACS structure, from the computation of truss geometry to the manufacturing and assembly. The already published methods for the form finding of PACS are adjusted and extended for the exemplary application of a variable-camber wing. The transfer of the form-finding truss model to a cross-sectional design is discussed. The end cap and sealing concept is described together with the implementation of the integral fluid flow. Conceptual limitations due to the manufacturing and assembly processes are discussed. The method’s efficiency is evaluated by Finite Element Method (FEM). In order to verify the underlying methods and summarize the presented work a modular real-life demonstrator is experimentally characterized and validates the numerical investigations.

Sealing concepts allowing a large change of cross-sectional area are investigated. Shape variable seals are indispensable for the biologically inspired pressure actuated cellular structures (PACS), which can be utilized to develop energy efficient, lightweight and adaptive structures for diverse applications. The requirements regarding extensibility, stiffness and load capacity exceed the characteristics of state of the art solutions. This work focuses on the design of seals suitable for extensional deformations greater than 25 %. In a first step, a number of concepts are generated. Then the most suitable concept is chosen based on numerical characterization and experimental examination. The deformation supportive end cap (DSEC) yields satisfying results as it displays a stress optimized shape under maximum load, an energetically inexpensive bending-based deformation mechanism and utilizes the applied forces to support distortion. In the first real-life implementation of a double row PACS demonstrator, which contains DSEC, the proof of concept is demonstrated.

A remarkable property of nastic, shape changing plants is their complete fusion between actuators and structure. This is achieved by combining a large number of cells whose geometry, internal pressures and material properties are optimized for a given set of target shapes and stiffness requirements. An advantage of such a fusion is that cell walls are prestressed by cell pressures which increases, decreases the overall structural stiffness, weight. Inspired by the nastic movement of plants, Pagitz et al. 2012 Bioinspir. Biomim. 7 published a novel concept for pressure actuated cellular structures. This article extends previous work by introducing a modular approach to adaptive structures. An algorithm that breaks down any continuous target shapes into a small number of standardized modules is presented. Furthermore it is shown how cytoskeletons within each cell enhance the properties of adaptive modules. An adaptive passenger seat and an aircrafts leading, trailing edge is used to demonstrate the potential of a modular approach., Comment: 15 pages, 9 figures

Strength of an unidirectional lamina is computed with a representative volume element. An approximative “voxel” meshing method is used in conjunction with continuum damage mechanics to simulate crack growth in the RVE. Two material models for the nonlinear material behaviour of the epoxy resin are compared. (© 2006 WILEY-VCH Verlag GmbH & Co. KGaA, Weinheim)

Thin-walled shell structures like circular cylindrical shells are prone to buckling. Imperfections, which are defined as deviations from perfect shape and perfect loading distributions, can reduce the buckling load drastically compared to that of the perfect shell. Design criteria monographs like NASA-SP 8007 recommend that the buckling load of the perfect shell shall be reduced by using a knock-down factor. The existing knock-down factors are very conservative and do not account for the structural behaviour of composite shells. To determine an improved knock-down factor, several authors consider realistic shapes of shells in numerical simulations using probabilistic methods. Each manufacturing process causes a specific imperfection pattern; hence for this probabilistic approach a large number of test data is needed, which is often not available. Motivated by this lack of data, a new deterministic approach is presented for determining the lower bound of the buckling load of thin-walled cylindrical composite shells, which is derived from phenomenological test data. For the present test series, a single pre-buckle is induced by a radial perturbation load, before the axial displacement controlled loading starts. The deformations are measured using the prototype of a high-speed optical measurement system with a frequency up to 3680 Hz. The observed structural behaviour leads to a new reasonable lower bound of the buckling load. Based on test results, the numerical model is validated and the shell design is optimized by virtual testing. The results of test and numerical analysis indicate that this new approach has the potential to provide an improved and less conservative shell design in order to reduce weight and cost of thin-walled shell structures made from composite material.

The full lightweight potential of carbon fiber reinforced plastics in complex structures can only be exploited using an adequate combination of material and design. Potential combinations can be found, using an optimization routine integrated into the design process. This paper presents a detailed comparison of the optimization results for distinctive rocket interstage design concepts. The complex cylindrical grid structure has to withstand the mission loads while being as light as possible. For the efficient use of the superior material properties, carbon fiber reinforced plastics are combined with a foam core to form a sandwich configuration. To minimize the total mass, parametric finite element models are used within a constrained, mixed-integer optimization. The structure is modeled using a bottom up strategy. An optimization framework, incorporating the CAE software Ansys into the multi-purpose optimization package pyOpt, is presented. The augmented lagrangian particle swarm algorithm is used to find the configurations having minimal weight. As manufacturability is an important design requirement, a winding process is represented by constraints in the parameter set of the geometry. The stiffnesses of the structure, as well as strength criteria of the laminate are used as inequality constraints for the optimization. Further on, a lower bound of the first natural eigenfrequency using additional constraints, is investigated. Finally, the impact of the governing geometry, material parameters and hence their influences on the optimization problem are carefully reviewed. It is shown that appropriate design – material combinations can significantly reduce the structural mass.

The Mobile Asteroid surface SCOuT (MASCOT) is a 11 kg small body lander package developed under DLR lead for the Japanese space probe Hayabusa2. Based on MASCOT a modified MASCOT-2 lander package was studied for the AIM (Asteroid Impact Mission) mission study. The objective of the study was to develop a lander package, which maximizes heritage and reuse from the MASCOT predecessor. Thus the landing module’s framework structure retained the MASCOT design variant, being up-scaled by approx. 20% (MASCOT-2: 330mm x 300mm x 210mm, 13 kg). In contrast, the interface structure called Mechanical and Electrical Support Structure (MESS) experienced a more significant re-design, due to the need of interface simplification and the peculiarities of the lander deployment for AIM mission and the Didymos system. This paper introduces three possible design variants for the MESS, which are later narrowed down to two and presented in greater detail: a CFRP-honeycomb sandwich plate with additional unidirectional stiffening plies and an X-shaped solid CFRP hat beam structure. Both variants are much simplified compared to the MASCOT support structure, but retain the overall mounting variant. This conceptual discussion is followed by a detailed structural analysis of both mechanical support structures. Provided by a set of mechanical loads and stiffness requirements, the sandwich and beam interface structures are separately simulated in a finite element model, consisting of shell and beam elements, respectively. The attached landing module is modelled with both, shell and beam elements, allowing a coupled structural analysis of the system. By varying geometrical and material parameters in the structural MESS models, a trade-off between the resulting minimal masses, the stiffness and the strength requirements is performed. Specifically considering also development risks it is concluded that the sandwich design variant shows an overall better performance.

Due to the non-isotropic nature of carbon fibre reinforced plastics (CFRP) unavoidable deformations occur during part production. These deformations often lead to dissatisfaction of tolerances or result in cost and time intensive rework of the tooling. In a cost driven production environment, similar to the automotive industry, it is essential to predict the deformations beforehand in order to compensate the used tooling. In future applications Fast Curing Epoxy Resins (FCER), with curing times of less than 20 minutes, will play a key role in order to enable high rate CFRP-production at lower cost. Until now, minimal experimental data exists for FCER systems. The present paper reports on a comprehensive experimental study on different FCER systems. It includes the characterization of neat resin samples as well as the quantification of the spring-in deformation of different L-profiles. Additionally, a study of parameters such as lay-up, temperature, cooling rate and fibre volume fraction concerning their influence on the process induced deformations is performed. Finally, the results are compared to slower curing systems. The findings of the study are relevant for the future development of robust prediction methods for process induced deformations of FCER-based composites.

A new deployable mast for realization of large, deployable space structures such as antennas of high arperture, long instrument booms, large solar arrays or solar sails has been developed at the DLR. The triangular truss is based on the use of longerons of composite tapes and shows a signicant increase in mass eciency compared to existing deployable mast concepts. The longerons of thin carbon bre tapes enable realization of large cross-sectional dimensions which leads to a strong insensitivity of critical compression load towards element length. Thereby, in comparison to trusses with longerons of solid rods - the most commonly used type of longerons - higher truss bay length and radius can be achieved for a lower amount of structural mass. This directly contributes to bending stiness and strength. The folding mechanism makes use of the high deformation capability of longerons of tapes in attened state. It is based on a two path folding pattern where the truss is attened rst in cross-direction and reeled up afterwards. This is enabled by the combined use of exible longerons and hinges added to one batten row and ensures high volume eciency. This is especially true for very long masts as the stowage volume mainly depends on the intial reeling radius.

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

The paper describes the results of the deployment testing of the De-Orbit Sail flight hardware, a drag sail for de-orbiting applications, performed by DLR. It addresses in particular the deployment tests of the fullscale sail subsystem and deployment force tests performed on the boom deployment module. For the fullscale sail testing a gravity compensation device is used which is described in detail. It allows observations of the in-plane interaction of the booms with the sail membrane and the central satellite body to which the booms and sails are attached. During the tests issues with the sail deployment tests were encountered which lead to design changes regarding the sail and type of sail attachment.

Fibre reinforced plastics with tough epoxy and thermoplastic matrices are spreading increasingly in many lightweight applications. For an efficient and reliable design the mechanical behaviour, considering non-linear plasticity, various failure modes under complex loading and damage progression, has to be estimated with numerical simulations. Most state-of-the-art continuum damage mechanics models do not consider the non-linear behaviour of the matrix material or are not suited for 3D solid elements. This work proposes a combined 3D continuum damage/plasticity model. It uses a single parameter flow criterion in combination with Cuntze’s Failure Mode Concept (FMC) for intralaminar failure. The FMC requires no iterative fracture angle search as the Action Plane Strength Criterion by Puck (APSC). This work describes details of the developed model like the coupling of the FMC with a degradation model as well as the implementation into Abaqus/Standard. A validation against open-hole tension tests made out of AS4/PEEK from literature is performed. It can be shown that the prediction of experimental failure loads with the FMC as well as with the APSC provides comparable results. The maximum deviations are between - 7.85 % and + 12.85 % . However, the computation times for predictions with the FMC are significantly less than with the APSC.