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Fuselage structures are subjected to combinations of axial, bending, shear and differential pressure loads. The validation of advanced metallic and composite fuselage designs against such loads is based on the full-scale testing of the fuselage barrel, which, however, is highly demanding from a time and cost viewpoint. This paper aims to assist in scaling-down the experimentation to the stiffened panel level which presents the opportunity to validate state-of-the-art designs at higher rates than previously attainable.

Development of a methodology to successfully design tests at the stiffened panel level and realize them using advanced, complex and adaptable test-rigs that are capable of introducing independently a set of distinct load types (e.g. internal overpressure, tension, shear) while applying appropriate boundary conditions at the edges of the stiffened panel.

A baseline test-rig configuration was developed after extensive parametric modelling studies at the stiffened panel level. The realization of the loading and boundary conditions on the test-rig was facilitated through innovative supporting and loading system set-ups.

The proposed test bench is novel and compared to the conventional counterparts more viable from an economic and manufacturing point of view. It leads to panel responses, which are as close as possible to those of the fuselage barrel in-flight and can be used for the execution of static or fatigue tests on metallic and thermoplastic curved integrally stiffened full-scale panels, representative of a business jet fuselage.

This paper aims to summarize the main features of the fuselage structure description within the Common Parametric Aircraft Configuration Schema (CPACS) data format.

The CPACS fuselage structure description includes the definition of arbitrary sheets and structural profiles which can be combined with a variety of material definitions to so-called structural elements. Besides the definition of these structural elements, the definitions of structural members, such as stringers, frames, floor structures and pressure bulkheads, as well as the definitions of the complex load introduction regions that transfer loads from the wings and the empennage into the fuselage shell are introduced. Finally, exemplary models generated with different mesh generation tools developed at the DLR Institute of Structures and Design are presented. These models are suitable for subsequent static or dynamic structural analyses.

The CPACS fuselage structure description is suitable for defining standard fuselage configurations including complex load introduction regions suitable for different types of structural analysis.

The work shows exemplary fuselage models generated from the introduced CPACS fuselage description suitable for subsequent static and dynamic structural analyses. As the CPACS standard is available for download, the described definitions may be used by universities, research organizations or the industry.

The work presents the definitions of the fuselage structure within the CPACS schema that were mainly developed by the authors employed at the DLR Institute of Structures and Design. The exemplary applications show models generated completely on the basis of the definitions described in this paper.

This paper aims to develop a new method of fuselage drag optimization that can obtain results faster than the conventional methods based on full computational fluid dynamics (CFD) calculations and can be used to improve the efficiency of preliminary design.

An efficient method for helicopter fuselage shape optimization based on surrogate-based optimization is presented. Two numerical simulation methods are applied in different stages of optimization according to their relative advantages. The fast panel method is used to calculate the sample data to save calculation time for a large number of sample points. The initial solution is obtained by combining the Kriging surrogate model and the multi-island genetic algorithm. Then, the accuracy of the solution is determined by using the infill criteria based on CFD corrections. A parametric model of the fuselage is established by several characteristic sections and guiding curves.

It is demonstrated that this method can greatly reduce the calculation time while ensuring a high accuracy in the XH-59A helicopter example. The drag coefficient of the optimized fuselage is reduced by 13.3%. Because of the use of different calculation methods for samples, this novel method reduces the total calculation time by almost fourfold compared with full CFD calculations.

To the best of the authors’ knowledge, this is the first study to provide a novel method of fuselage drag optimization by combining different numerical simulation methods. Some suggestions on fuselage shape optimization are given for the XH-59A example.

In recent years, innovative aircraft designs have been investigated by researchers to address the environmental and economic issues for the purpose of green aviation. To keep air transport competitive and safe, it is necessary to maximize design efficiencies of the aircrafts in terms of weight and cost. The purpose of this paper is to focus on the research which has led to the development of a novel lattice fuselage design of a forward-swept wing aircraft in the conceptual phase by topology optimization technique.

In this paper, the fuselage structure is modelled with two different types of elements – 1D beam and 2D shell – for the validation purpose. Then, the finite element analysis coupled with topology optimization is performed to determine the structural layouts indicating the efficient distributed reinforcements. Following that, the optimal fuselage designs are obtained by comparison of the results of 1D and 2D models.

The topological results reveal the need for horizontal stiffeners to be concentrated near the upper and lower extremities of the fuselage cross section and a lattice pattern of criss-cross stiffeners should be well-placed along the sides of the fuselage and near the regions of window locations. The slight influence of windows on the optimal reinforcement layout is observed. To form clear criss-cross stiffeners, modelling the fuselage with 1D beam elements is suggested, whereas the less computational time is required for the optimization of the fuselage modelled using 2D shell elements.

The authors propose a novel lattice fuselage design in use of topology optimization technique as a powerful design tool. Two types of structural elements are examined to obtain the clear reinforcement detailing, which is also in agreement with the design of the DLR (German Aerospace Center) demonstrator. The optimal lattice layout of the stiffeners is distinctive to the conventional semi-monocoque fuselage design and this definitely provides valuable insights into the more efficient utilization of composite materials for novel aircraft designs.

This paper aims to investigate the rebound process and the secondary-impact process of the fuselage section that occurs in the actual crash events.

A full-scale three-dimensional finite element model of the fuselage section was developed to carry out the dynamic simulations. The rebound process was simulated by removing the impact surface at a certain point, while the secondary-impact process was simulated by striking the impact surface against the fuselage bottom after the first impact.

For the rebound process, the fuselage structure restores deformation due to the springback of the fuselage bottom, and it results in structural vibration of the fuselage section. For the secondary-impact process, the fuselage deformation is similar with that of the single impact process, indicating that the intermittent impact loading has little influence on the overall deformation of the fuselage section. The strut failure is the determining factor to the acceleration responses for both the rebound process and the secondary-impact process.

The rebound process and the secondary-impact process, which is difficult to study by experiments, was investigated by finite element simulations. The structure deformations and acceleration responses were obtained, and they can provide guidance for the crashworthy design of fuselage structures.

This research first investigated the rebound process and the secondary-impact process of the fuselage section. The absence of the ground load and the secondary-impact was simulated by controlling the impact surface, which is a new simulating method and has not been used in the previous research.

This paper aims to present the results of calculations that checked how the longerons and frames arrangement affects the stiffness of a conventional structure. The paper focuses only on first stage of research – analysis of small displacement. Main goal was to compare different structures under static loads. These results are also compared with the results obtained for a geodetic structure fuselage model of the same dimensions subjected to the same internal and external loads.

The finite element method analysis was carried out for a section of the fuselage with a diameter of 6.3 m and a length equal to 10 m. A conventional and lattice structure – known as geodetic – was used.

Finite element analyses of the fuselage model with conventional and geodetic structures showed that with comparable stiffness, the weight of the geodetic fuselage is almost 20 per cent lower than that of the conventional one.

This analysis is limited to small displacements, as the linear version of finite element method was used. Research and articles planned for the future will focus on nonlinear finite element method (FEM) analysis such as buckling, structure stability and limit cycles.

The increasing maturity of composite structures manufacturing technology offers great opportunities for aircraft designers. The use of carbon fibers with advanced resin systems and application of the geodetic fuselage concept gives the opportunity to obtain advanced structures with excellent mechanical properties and low weight.

This paper presents very efficient method of assessing and comparison of the stiffness and weight of geodetic and conventional fuselage structure. Geodetic fuselage design in combination with advanced composite materials yields an additional fuselage weight reduction of approximately 10 per cent. The additional weight reduction is achieved by reducing the number of rivets needed for joining the elements. A fuselage with a geodetic structure compared to the classic fuselage with the same outer diameter has a larger inner diameter, which gives a larger usable space in the cabin. The approach applied in this paper consisting in analyzing of main parameters of geodetic structure (hoop ribs, helical ribs and angle between the helical ribs) on fuselage stiffness and weight is original.

The purpose of this paper is to present some of the key achievements. At DLR, a sophisticated interdisciplinary aircraft design process is being developed, using the CPACS data format (Nagel et al., 2012; Scherer and Kohlgrüber, 2016) as a means of exchanging results. Within this process, TRAFUMO (Scherer et al., 2013) (transport aircraft fuselage model), built on ANSYS and the Python programming language, is the current tool for automatic generation and subsequent sizing of global finite element fuselage models. Recently, much effort has gone into improving the tool performance and opening up the modeling chain to further finite element solvers.

Much functionality has been shifted from specific routines in ANSYS to Python, including the automatic creation of global finite element models based on geometric and structural data from CPACS and the conversion of models between different finite element codes. Furthermore, a new method for modeling and interrogating geometries from CPACS using B-spline surfaces has been introduced.

Several new modules have been implemented independently with a well-defined central data format in place for storing and exchanging information, resulting in a highly extensible framework for working with finite element data. The new geometry description proves to be highly efficient while also improving the geometric accuracy.

The newly implemented modules provide the groundwork for a new all-Python model generation chain, which is more flexible at significantly improved runtimes. With the analysis being part of a larger multidisciplinary design optimization process, this enables exploration of much larger design spaces within a given timeframe.

In the presented paper, key features of the newly developed model generation chain are introduced. They enable the quick generation of global finite element models from CPACS for arbitrary solvers for the first time.

The aim of this paper is to present a new variation modeling method for fuselage structures in digital large aircraft assembly. The variation accumulated in a large aircraft assembly process will influence the dimensional accuracy and fatigue life of airframes. However, in digital large aircraft assembly, variation analysis and modeling are still unresolved issues.

An elastic structure model based on beam elements is developed, which is an equivalent idealization of the actual complex structure. The stiffness matrix of the structure model is obtained by summing the stiffness matrices of the beam elements. For each typical stage of the aircraft digital assembly process, including positioning, coordinating, joining and releasing, variation models are built using the simplified structure model with respective loads and boundary conditions.

Using position errors and manufacturing errors as inputs, the variations for every stage of the assembly process can be calculated using the proposed model.

This method has been used in a large fuselage section assembly project, and the calculated results were shown to be a good prediction of variation in the actual assembly.

Although certain assumptions have been imposed, the proposed method provides a better understanding of the assembly process and creates an analytical foundation for further work on variation control and tolerance optimization.

The purpose of this paper is to present the result of calculations that were performed to estimate the structural weight of the passenger aircraft using novel technological solution. Mass penalty resulting from the installation of the fuselage boundary layer ingestion device was needed in the CENTRELINE project to be able to estimate the real benefits of the applied technology.

This paper focusses on the finite element analysis (FEA) of the fuselage and wing primary load-carrying structures. Masses obtained in these analyses were used as an input for the total structural mass calculation based on semi-empirical equations.

Combining FEA with semi-empirical equations makes it possible to estimate the mass of structures at an early technology readiness level and gives the possibility of obtaining more accurate results than those obtained using only empirical formulas. The applied methodology allows estimating the mass in case of using unusual structural solutions, which are not covered by formulas available in the literature.

Accurate structural mass estimation is possible at an earlier design stage of the project based on the presented methodology, which allows for easier and less costly changes in designed aircrafts.

The presented methodology is an original method of mass estimation based on a two-track approach. The analytical formulas available in the literature have worked well for aeroplanes of conventional design, but thanks to the connection with FEA presented in this paper, it is possible to estimate the structure mass of aeroplanes using unconventional technological solutions.

The purpose of this paper is to introduce a novel method to study the flutter coupling mechanism of the twin-fuselage aircraft, which is becoming a popular transportation vehicle recently.

A new method of flutter mode indicator is proposed based on the principle of work and power, which is realized through energy accumulation of generalized force work on generalized coordinates, based on which flutter coupling mechanism of the twin-fuselage aircraft is studied using ground vibration test and computational fluid dynamics/computational solid dynamics method.

Verification of the proposed flutter mode indicator is provided, by which the flutter mechanism of the twin fuselage is found as the horizontal tail’s torsion coupled with its bending effect and the “frequency drifting” phenomenon of twin-fuselage aircraft is explained logically, highlighting the proposed method in this paper.

This paper proposed a new method of flutter mode indicator, which has advantages in flutter modes indexes reliability, clear physical meaning and results normalization. This study found the flutter coupling mechanism of twin-fuselage aircraft, which has important guiding significance to the development of twin-fuselage aircraft.