public deliverable template - inside - centreline

Ref. Ares(2018)6036686 - 26/11/2018

CENTRELINE D2.8 Interim report on fuselage-wing junction aero-structural investigation Deliverable submission date: 26.11.2018

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Grant Agreement: 723242

Call identifier: H2020-MG-2016-Two-Stages

Project full title: CENTRELINE – ConcEpt validatioN sTudy foR fusElage wake-filLIng propulsioN

integration

Deliverable lead beneficiary: WUT

INTERIM REPORT ON FUSELAGE-WING JUNCTION AERO-

STRUCTURAL INVESTIGATION

Authors: Zdobysław Goraj, Mariusz Kowalski, Bartłomiej Goliszek

Internal Technical Auditor Name (Beneficiary short name) Date of approval

Task leader Bartłomiej Goliszek (WUT) 05.11.18

WP leader Fabian Peter (BHL) 30.10.18

Coordinator Arne Seitz (BHL) 23.11.18

Project Office Sophie Rau (ART) 26.11.2018

Abstract: The deliverable describes interim results of analyses performed as a part of Task T2.4.2 “Interim

report on fuselage-wing junction aero-structural investigation”. The following report presents the

methodology leading from basic external geometry, through choosing a structure concept and, defining

acting loads to the generation of an FEM model at the final. The report also includes a summary of future

work to be conducted in order to finish the task until month 26.

Due date: 31.10.2018

Actual submission date: 26.11.2018

Publication date: 26.11.2018

Project start date: 01.06.2017

Project duration: 36 months


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Table of Contents

Table of Contents ........................................................................................................................................ 2

List of Figures ............................................................................................................................................... 3

List of Tables ............................................................................................................................................... 4

Glossary .......................................................................................................................................................5

Executive Summary .................................................................................................................................... 6

1 Introduction .......................................................................................................................................... 7

2 Applied Methodologies ....................................................................................................................... 8

2.1 Preliminary Weight Assessment .................................................................................................. 8

2.2 Analytical Approach Methodology ............................................................................................... 8

3 Fuselage Structure Concept ................................................................................................................ 11

3.1 State of the Art for Fuselage Structure........................................................................................ 11

3.1.1 Conventional Fuselage Structure ............................................................................................ 11

3.1.2 Geodetic Fuselage Structure ................................................................................................... 11

3.2 Geodetic Structure Application in PFC ........................................................................................ 12

4 Centre-Wing Fuselage Section ........................................................................................................... 14

4.1 State-of-the Art for Centre-Wing Section ................................................................................... 14

4.1.1 Mid Fuselage Section .............................................................................................................. 14

4.1.2 Single Piece Centre Wing Box.............................................................................................. 14

4.2 Designed Structure ..................................................................................................................... 15

5 Acting Loads ....................................................................................................................................... 17

5.1 Load Cases Definition ................................................................................................................. 17

5.2 Analysed Fuselage Zones ............................................................................................................ 18

5.3 Critical Loads Selection ............................................................................................................... 19

6 Interim Structure Analysis Results ..................................................................................................... 20

6.1 Material Properties .................................................................................................................... 20

6.2 FEM Analysis Methodology ........................................................................................................ 22

6.3 Centre-Wing Mesh Definition ..................................................................................................... 23

6.4 Geodetic and Conventional Fuselage Structure Analysis ............................................................ 24

6.4.1 Analysis of Conventional Fuselage Structural Layout ......................................................... 24

6.4.2 Analysis of Geodetic Fuselage Structural Layout ................................................................. 27

7 Conclusion and Future Work ............................................................................................................... 32

8 Bibliography ....................................................................................................................................... 33


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List of Figures

Figure 2.1 Methodology flow ...................................................................................................................... 10

Figure 3.1 Classic fuselage structure [Gibson, 2016] ................................................................................... 11

Figure 3.2 Geodetic fuselage structure ....................................................................................................... 12

Figure 3.3 CAD model of a “classic” and “lattice” structure ......................................................................... 13

Figure 3.4 Geodetic structure in PFC .......................................................................................................... 13

Figure 4.1 Airbus centre wing 3D model [Airbus, 2017] ............................................................................... 15

Figure 4.2 Airbus centre-wing box technology demonstrator [Airbus, 2017] .............................................. 15

Figure 4.3 Centre fuselage exploded view .................................................................................................. 16

Figure 5.1 Analysed centre fuselage structure ............................................................................................ 17

Figure 5.2 Centre fuselage zones ................................................................................................................ 18

Figure 6.1 Step 1: Conversion of 3D Solid Body CAD model into Surface 2D model ................................... 22

Figure 6.2 Step 2: Division of the model into zones with the same laminate system ................................. 22

Figure 6.3 Step 3: Definition of finite element mesh in all zones ................................................................ 22

Figure 6.4 Step 4: Control of the material orientation in the mesh ............................................................. 23

Figure 6.5 Step 5: Laminate layup definition .............................................................................................. 23

Figure 6.6 Centre-wing mesh .................................................................................................................... 24

Figure 6.7 Loads and constraints applied to the fuselage cylinder structure .............................................. 24

Figure 6.8 Displacement of load carrying structure of the conventional fuselage (version.1) ...................... 25

Figure 6.9 External shell displacements in the conventional fuselage (version. 5) ....................................... 27

Figure 6.10 Ribs displacement in Ver. 1 ..................................................................................................... 28

Figure 6.11 Geodetic cylinder structure - skin displacement Ver.2 ............................................................. 29

Figure 6.12 Geodetic cylinder structure - Stresses in frame structure Ver. 3 ............................................... 30

Figure 6.13 Ribs displacement in a scale of 10 to 1, in Ver. 6. ...................................................................... 31


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List of Tables

Table 2.1 Weight breakdown [Seitz et al., 2018] .......................................................................................... 8

Table 6.1 UD Prepreg Mechanical Properties ............................................................................................ 20

Table 6.2 UD Prepreg Physical Properties ................................................................................................. 20

Table 6.3 Woven Prepreg Mechanical Properties ....................................................................................... 21

Table 6.4 Woven Prepreg Physical Properties ............................................................................................ 21

Table 6.5 Specification of the calculated versions of the conventional fuselage ......................................... 25

Table 6.6 Comparison of displacements of the conventional fuselage in each calculated version ............. 26

Table 6.7 Angles between ribs, ribs numbers, frame weights and increments/decrements in percent ........ 27

Table 6.8 Results FEM simulation .............................................................................................................. 28


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Glossary

Abbreviation / acronym Description

CAD Computer-Aided Design

CENTRELINE ConcEpt validatioN sTudy foR fusElage wake-filLIng propulsioN

intEgration

CPACS Common Parametric Aircraft Configuration Schema

CFRP Carbon Fibre Reinforced Polymer

CS Certification Specifications

CPACS Common Parametric Aircraft Configuration Schema

EASA European Aviation Safety Agency

FEM Finite Element Method

IGES Initial Graphics Exchange Specification

MLG Main Landing Gear

PFC Propulsive Fuselage Concept

STEP Standard for the Exchange of Product Model Data

TiGL TiGL Geometry Library

TRL Technology Readiness Level

UD Unidirectional


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Executive Summary

This report documents the actions taken so far to accomplish the ConcEpt validatioN sTudy foR fusElage

wake-filLIng propulsioN integration (CENTRELINE) Task T2.4.2 "Fuselage-wing junction aero-structural

investigation". The document contains information on the methods used, a description of 3D Computer-

Aided Design (CAD) models and the presentation of the preliminary results obtained. The approach to the

preparation of a parametric CAD model for the PFC (Propulsive Fuselage Concept) aircraft is shown. The

prepared model facilitates modifications and optimization of the structure, and enables graphical

presentation of various solutions and layouts. The report contains the description of classic and geodetic

fuselage structure, and presents the proposition of usage of the geodetic fuselage in the CENTRELINE

initial PFC design. The methodology of preparing the model for the Finite Element Method (FEM) analysis

is described, and the finite element mesh prepared for the centre-wing is presented. The studies include

the comparative analysis results of the fuselage stiffness in geodetic and conventional layout. Finally,

important aspects and further steps are presented to complete the task until month 26.


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1 Introduction

According to “Flightpath 2050” [EC, 2011] and SRIA issued by ACARE [ACARE, 2012], great challenges are

posed for aviation in order to sustainably protect the natural environment. The guidelines regulating the

permitted emission of harmful gases to the atmosphere are more and more stringent every year. The

CENTRELINE project [Seitz et al., 2018] aims at the prove-of concept for a wake-filling propulsion system

integrated with the aft fuselage section – the so-called Propulsive Fuselage Concept. This very promising

approach can bring a significant reduction in both fuel consumption and the amount of harmful gases

emitted into the atmosphere. However, to achieve the intended effect, the mass penalty, associated with

the installation of additional equipment, should be minimized. The interim report concerns the task T2.4.2

"Fuselage-wing junction aero-structural investigation". The task is dedicated to the pre-design of the wing

and the fuselage junction and their integration. This report describes previous activities, documents the

applied methodology and presents the preliminary results obtained.

Section 2 provides information about the methodology of building the 3D CAD model. It presents how the

parametric model is built step by step and what are the individual stages leading to obtaining useful results.

The workflow is presented on the graphic diagram. This section also presents a mass breakdown

estimated using the semi-empirical methods. Section 3 deals with the concept of the fuselage and wing

structure. Two versions of the fuselage are described: geodetic and conventional. Section 3 shows also the

proposed concept of the fuselage made by combining sections with different layouts. In Section 4, the

centre-wing structure is analysed. At the beginning, the state of the art for centre-wing design is presented.

The preliminary design of the centre-wing and fuselage load-carrying structures are also described.

Section 5 of the report outlines the most important loads acting on the designed structure. The analysis of

the fuselage zones subjected to the various loads are presented and the most significant loads were

selected for later analysis. Section 6 contains information on numerical analysis. Physical and mechanical

properties of selected composite materials that will be used in the FEM analysis are listed in the tables. The

methodology of the model preparation for the FEM and prepared finite element mesh made on the

structure of the centre-wing are presented. The next subsection contains the results obtained in the

comparative analysis of the fuselage stiffness in the geodetic and conventional version. The summary and

definition of future activities to complete the task before the month 26 are shown in Section 7.


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2 Applied Methodologies

In this section, the key facts about applied methodologies are outlined. The first paragraph shows the

preliminary weight assessment generated using semi empirical methods, while the second paragraph

presents the analytical approach methodology. Each stage of work has been defined, aiming at the

transition from the supplied external geometry to the solid model and then to the surface model that will

be used for the Finite Element Analysis. This section presents also the parametric CAD model that has

been prepared to simplify and speed up future activities connected with optimization and modification of

the load carrying structure concept.

2.1 Preliminary Weight Assessment

Table 2.1 presents basic weight breakdown of the aircraft structure, obtained using semi-empirical

methodology [Seitz, 2011].In the table mass changes compared to R2035 reference aircraft are presented.

The analytical approach presented in this report, using numerical methods, is aimed at estimating more

specifically the mass penalty caused by mounting additional equipment and adapting the load-carrying

structure to the additional loads.

Table 2.1 Weight breakdown [Seitz et al., 2018]

Parameter Delta vs. R2035 [%]

Wing +0.9

Fuselage +4.5

Horizontal Tail -8.9

Vertical Tail +26.7

Pylons -17.5

Landing Gear +6.6

Structures Total +2.0

2.2 Analytical Approach Methodology

The CENTRELINE project utilizes the Common Parametric Aircraft Configuration Schema (CPACS)

language to define geometry and performance data of both reference R2035 and the initial Propulsive

Fuselage Concept (PFC) design. To visualize the geometry defined in CPACS, the TiGL Geometry Library

(TiGL) viewer software is used. The TiGL viewer allows to generate Standard for the Exchange of Product

Model Data (STEP) or Initial Graphics Exchange Specification (IGES) format files, that can be read by

almost any CAD software available. The STEP file of the geometry has been generated and implemented

into the Siemens NX® software. Using this software, the work on the concept of the internal structure of

the aircraft equipped with a propulsor has begun. The designed load carrying structure model has been

parametrized to the greatest extent possible, to facilitate later modifications and optimization. The

parametric CAD models of fuselage, wing, empennage and propulsor were prepared [Goraj et al., 2018].

The fuselage of the PFC aircraft is divided into sections, see section 3.2. For the cylindrical sections (except

fuselage-wing junction area) the geodetic layout was used, remaining sections make use of the


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conventional design with stringers and frames. For basic optimization of the geodetic fuselage structure,

the following variables were defined:

• Number of hoop ribs,

• Number of helical ribs,

• Angle between helical ribs.

For the conventional sections the variables were:

• Number of stringers,

• Number of frames,

• Thickness of frames.

Having two versions of structure (classical and geodetic) already parametrized, a combination of both

was used to create an aircraft fuselage (see section 3.2).

The same parametrization procedures were applied to all wing elements. For the wings the classic

composite structure was chosen because it is well known, widely proven by experience and reliable enough.

The wing structure consists of skins, spars and ribs. The first action for each wing element type is to define

the division for fixed and movable parts (main parts, slats, flaps, ailerons for main wing and main parts,

elevators and rudders for tail plane and vertical stabiliser). When starting the structure of a fixed wing part

parametrisation one has to define how the thickness and number of plies made of unidirectional (UD)

Carbon Fibre Reinforced Polymer (CFRP) are changing versus wingspan. For the description of the wing

skin the following main variables could be used:

• Thickness of a single carbon composite ply,

• Number of plies in the wing root section,

• Number of plies in the wing tip section,

• Change of skin thickness along wing span.

The second step in the wing structure modelling is a spar design. For the PFC aircraft, a two spar structure

has been chosen. As the structure is fully composite, the C-shape spar cross sections was selected, because

such a shape is easier to manufacture by automated fibre placement machines. Basing on previously

defined wing division, the spars placement has been parametrized.

For the spars, the main parametrized variables were selected as follows:

• Spar placement,

• Thickness of a single carbon composite ply,

• Spar cap width and its change versus wingspan (both for front and rear spar),

• Number of carbon composite plies in upper spar cap and its change versus wingspan (both for

front and rear spar),

• Number of carbon composite plies in lower spar cap and its change versus wingspan (both for

front and rear spar),

• Number of carbon composite plies in web spar and its change along wingspan (both for front

and rear spar),

The third group of important element of wing design are stringers. They are designed as flats made of an

UD carbon composite and are attached to the skins in the co-curing process.

For stringers the following independent variables are assumed:

• Number of stringers,


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• Angle between stringers and wing leading edge,

• Flat thickness,

• Height of each stringer (the closest to the fuselage) and its change versus wingspan.

Finally, a model for the wing ribs needs to be built. The geometry of each rib results from the geometry of

spars, skins and stringers.

For the ribs the following independent variables were selected:

• Number of ribs,

• Distance between adjacent ribs,

• Thickness of each rib.

The prepared solid body model, after defining appropriate parameters, can be used to estimate the mass

of the structure, but for more accurate determination, it is necessary to perform time and labour-

consuming calculations for various load cases. Such calculations require a huge amount of work and

computing power. Based on the solid model, a surface model needed for the FEM analysis, was generated.

In the scope of the CENTRELINE project, the FEM analysis of the selected load cases for selected

structures will be conducted. In particular, the analysis and optimization of the load-carrying structure of

the centre-wing fuselage section will be carried out. The Figure 2.1 presents the design and analysis

process.

Described methodology was used also for the structural pre-design of the aft-fuselage and Fuselage Fan

nacelle and their integration [Goraj, Kowalski et al. 2018]

Figure 2.1 Methodology flow


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3 Fuselage Structure Concept

This section concerns the load-carrying structure concept of the fuselage. The concept of geodetic

fuselage and the conventional fuselage structure are described. Attached figures present the construction

details in both versions of the fuselage structure. The concept utilizing a geodetic layout in the fuselage

sections of the PFC aircraft is also presented.

3.1 State of the Art for Fuselage Structure

3.1.1 Conventional Fuselage Structure

Conventional (or classic) fuselage structure is an example of a semi-monocoque structure, which has been

widely used in commercial aviation for a very long time. This technology was an evolution of the original

monocoque structure, which had some disadvantages. However, in monocoque structures of significant

size, the challenge of ensuring strength to mass ratio occurred. A low strength to mass ratio is usually

caused by the high weight of the skin, which has to be sufficiently stiff and durable to carry all bending

loads. Conventional fuselage structure, like the monocoque structure, consists of frames and skin, but has

also additional strengthening longerons that extend along the entire length of the fuselage and help to

maintain the highest bending loads. Moreover, the longerons prevent the skin from buckling. Frames are

used to maintain the shape of the fuselage cross-sections as well as increasing the longerons buckling

strength. They also allow the application of concentrated loads to the airframe. All above mentioned

elements are connected with rivets, bolts, screws and nuts. Originally, conventional structures were made

mainly of aluminium alloys. Nowadays, metal has been replaced by composite materials in many

applications, but the conventional structure concept remained unchanged. An example of a conventional

fuselage structure is presented in Figure 3.1.

Figure 3.1 Classic fuselage structure [Gibson, 2016]

3.1.2 Geodetic Fuselage Structure

A geodetic structure made of aluminium was used for the first time in the construction of the Wellington

Mk. X HE239 aircraft. The biggest problem with those structures at that time was the complicated

strength calculations [Michalski, 2016].

Geodetic fuselage structures consist of an external shell and ribs grid. The external shell in this structure is

responsible mainly for carrying the internal pressure loads. Geodetic grid made of UD carbon fibre


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composite is responsible mainly for carrying external loads. The load carrying grid consists of hoop and

helical ribs that entwine the fuselage. Helical ribs act like diagonals of a rectangular element lying on a

cylindrical shell subjected to twisting – one diagonal is extended and the other shortened. Thanks to this

principle, the loads acting on the opposite ribs are eliminated. According to research carried out so far, a

geodetic fuselage design in combination with advanced composite materials yields an additional fuselage

weight reduction of approximately 10% [Hühne, 2013]. 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. A fuselage sample with geodetic structure is presented in Figure 3.2.

Figure 3.2 Geodetic fuselage structure

3.2 Geodetic Structure Application in PFC

The CENTRELINE project aims at the market entry at 2035 [Seitz, 2018] and therefore the considered

technologies must correspond that time. Currently an old type of design for the airframes is reconsidered

in connection with progress of maturity of composite technology. Aeronautical experts believe that the

lattice composite structure would be a good design selection for such a structure [Vasiliev, 2006], [Vasiliev,

2012], [Khairi, 2010], [Araujo, 2010]. A lattice structure consists of skin and helical and hoop ribs made

from UD CFRP.

Based on the cylindrical part of R2035 fuselage, two sets of the composite structures samples (Figure 3.3)

have been designed: the “classic” concept with longerons, frames and skin, and the “lattice” concept with

helical ribs, hoop ribs and skin.


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Figure 3.3 CAD model of a “classic” and “lattice” structure

Both samples have been analysed using FEM within the Nastran module and the results are presented in

the section 6.4.

A combination of classic and geodetic structure was used to create an aircraft fuselage. There are two

reasons that combination both the classical and geodetic structure are used [Botelho, 2006]. First, in the

wing-fuselage section the classical structure must be used instead of geodetic one because of

complication in joining wing-centre box just to geodetic structure. Second, it must be underlined that the

lattice structure cannot be used on non-developable surfaces. Figure 3.4 shows where the lattice structures

(see the crosshatched area) are used. In Figure 3.4 doors and windows cut-outs are not presented.

However, this aspect will be considered in the final structure concept.

Figure 3.4 Geodetic structure in PFC


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4 Centre-Wing Fuselage Section

Section 4 contains description of the centre-wing and the fuselage wing junction area. First paragraph

presents the current state-of-the-art for mid fuselage section, it also outlines ongoing research on

composite centre wing box. The second part of the chapter presents and describes preliminary design of

the centre fuselage section for the PFC aircraft.

4.1 State-of-the Art for Centre-Wing Section

4.1.1 Mid Fuselage Section

This region of the fuselage is located around wing-fuselage intersection. It contains part of the cabin and,

the air conditioning, hydraulic and main landing gear bays beneath the cabin floor, together with the

integration structure for the wing centre box. The zone beneath the cabin floor is unpressurized, the actual

pressure boundary is formed by the upper skin panels of the centre wing box and a pressure diaphragm

extending from the wing box to the end of the main landing gear bay. The forward pressure boundary is

formed by the frame ahead of the centre wing box. The aft boundary is formed by the pressure bulkhead

installed at the end of the MLG (Main Landing Gear) bay. The centre section of the wing extends across the

width of the fuselage and forms an integral fuel tank. It is bounded by the front and rear spars, the upper

and lower skins and the root ribs. Internally there are spanwise truss type ribs extending from the front

spar to the rear spar. The centre wing box provides external attachments and internal supports at the root

ribs for the fuselage frames.

4.1.2 Single Piece Centre Wing Box

In 2017, Airbus has created the first-ever single-piece composite centre wing box – representing an

important evolution of this vital structural component that provides support and rigidity for an aircraft’s

wings. It was designed as an upgrade to conventional multi-part centre wing boxes, with advantages of the

new version including a 20% reduction in manufacturing costs. [Airbus, 2017]

The single-piece composite centre wing box fully leverages advances in composite technologies, including

the moulding of complex parts combined with continuous fibre. This makes it easier to assemble and

provides improved load-bearing properties. Due to a reduced parts count, the weight of the component as

a whole is smaller [Airbus, 2017]. An example of the Single Piece Centre Wing Box is shown in Figure 4.1.


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Figure 4.1 Airbus centre wing 3D model [Airbus, 2017]

Figure 4.2 Airbus centre-wing box technology demonstrator [Airbus, 2017]

4.2 Designed Structure

The designed structure of the mid fuselage section combines the conventional structure described in

chapter 4.1.1 along with advanced composite technologies (see chapter 4.1.2). The classic layout has been

chosen due to the high complexity of connecting the geodetic structure with the remaining parts. The

length of the centre section is minimized, so the remaining parts of fuselage can be made as a geodetic

structures to decrease weight as much as it is possible. The centre wing box is designed as a CFRP

monobloc with strut type frames inspired by the latest Airbus technology demonstrator (see section 4.1.2).

Front and rear pressure bulkheads are designed in a similar way as the CFRP monobloc. In the lower part of

this section there is a keel beam, that is responsible for carrying the tension and compressing loads in this

part of the fuselage. The upper part of the mid fuselage section has a conventional design with frames,

stringers and skin. Moreover, two frames attached to the centre wing box are strengthened to ensure a

sufficient load transfer around the assembly. Figure 4.3 presents the 3D CAD model of the mid fuselage

structure with skin and stringers not visible.


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Figure 4.3 Centre fuselage exploded view


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5 Acting Loads

In this section, forces and loads are presented. Outlined load cases will be used in later stages of the

project to calculate the state of stress and deformation of designed structure. The structure will be

designed to withstand the described loads that may occur separately or combined.

5.1 Load Cases Definition

For the PFC aircraft the0 Certification Specifications (CS) - 25 are applicable. Those regulations are issued

by European Aviation Safety Agency [EASA, 2017] and apply to large civil transport aircrafts. Certification

specifications are guidelines for the manufacturer and engineers, how structure should be designed to

ensure not only aircraft structural integrity but also to ensure passengers safety and to minimize risk of

fatal accidents.

Task T2.4.2 focuses on investigating the fuselage-wing junction, therefore the design and analysis of

centre part of aircraft fuselage is the most important challenge. It is a complicated problem as this part of

fuselage concentrates all acting loads from all of the aircraft components. For this reason, this part of

structure is the most strenuous section of the aeroplane. The analysed structure is shown on the Figure 5.1

Figure 5.1 Analysed centre fuselage structure

Concerning the fuselage and the CS-25 regulations that pertain to its structure, four categories of acting

loads are addressed: flight loads, flight loads with pressurization, ground loads and pressure loads. The

structure must withstand the described cases in aircraft operational configuration. The described cases

consist of the following main limit loads:

Flight cases (with or without pressurization)

• Manoeuvre and gust loads,

• Lateral gust,

• Elevator deflection,


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• Flight with side slipping.

Ground Cases

• Level landing,

• Tail down landing.

Pressure loads

• Cabin over pressurization.

5.2 Analysed Fuselage Zones

These sources of the loads are not all described in the CS-25 regulations, however they have the most

severe influence on the dimensioning of the aircraft fuselage structure. In the analysed centre part of the

fuselage, six zones can be distinguished that are typically sized by different sets of loads.

Zone 1: pressurized gust or maneuver or crash

Zone 2: pressurized gust or maneuver

Zone 3: dynamic landing tail down landing

Zone 4: maneuver or gust - dynamic landing

Zone 5: lateral gust

Zone 6: cabin pressure

Figure 5.2 Centre fuselage zones

Zone 1

For the centre wing section, the sizing loads are for the wing bending with pressurization condition. These

loads are considered in gust or manoeuvre cases. Cabin pressure is taken into account for these loads. The

forward parts of centre wing are generally sized by crash.

FWD


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Zone 2

The upper centre fuselage is sized by gust or manoeuvre loads associated with cabin pressure.

Zone 3

This section is generally sized by dynamic level landing. Especially the keel beam needs to be checked for

buckling under compression loads.

Zone 4

This zone consists of lateral panels and slanted frames. It is sized with the shear force occurring in the

cases of manoeuvre, gust or dynamic landing.

Zone 5

Parts presented as a Zone 5 in Figure 5.2 are generally sized by lateral gusts, that induce tension and

compression loads in lateral parts of the fuselage.

Zone 6

Forward and aft pressure bulkheads are usually sized by cabin over pressure.

5.3 Critical Loads Selection

Investigating the above described loads and zones, the following combination of limit loads has been

chosen as critical to be analysed:

• Maximum pressurized manoeuvre or gust,

• Lateral gust,

• Dynamic level landing,

• Fuselage over pressurizing.

Mentioned above load cases were selected from the long list of cases considered during the aircraft design,

however they are sufficient to a sufficient preliminary dimensioning of aircraft structure and to achieve

Technology Readiness Level (TRL) 3. These combined load cases will be implemented in the FEM analysis.


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6 Interim Structure Analysis Results

This chapter presents the methodology for the preparation of the FEM analysis, as well as preliminary

results of the stiffness analysis for the geodetic and conventional fuselage. Section 6.2 presents the

process of preparing a model for the FEM analysis step by step. Next, the prepared Finite Element Mesh

for the centre-wing is described, and the results of the comparative analysis of the conventional and

geodetic structure are presented.

6.1 Material Properties

The CENTRELINE project envisages the use of highest quality materials with the best strength to mass

ratio available. Carbon composite provides strength several times better than steel with a weight lower

than aluminium. In order to perform FEM calculations, the material had to be pre-selected. In the case of

the load-carrying structure, the most important are UD carbon tapes and carbon woven fabrics. Many

carbon, glass or aramid fabrics of various weights and weaves are used in the final structure of the aircraft,

however, at the conceptual design stage, only carbon-based woven fabrics and UD tapes, which are mostly

responsible for transferring the loads, will be considered. After analysing the materials available on the

market, the most durable systems used in aircraft constructions were selected. The HEXCEL prepregs are

used, among others, in the Airbus A350 XWB which has the highest percentage of the composites in the

load carrying structure [Hexcel 1, 2018]. The physical and mechanical properties of the materials selected

for the analysis were collected in the tables below.

• Unidirectional prepreg, M21 Resin, IMA-12k Fibres [Hexcel, 2018]

Table 6.1 UD Prepreg Mechanical Properties

Mechanical Properties Units Value

Tension Strength MPa 3050

Tension Modulus GPa 178

Compression Strength MPa 1500

Compression Modulus GPa 146

In-plane Shear Strength MPa 94

In-plane Shear Modulus GPa 5.2

Table 6.2 UD Prepreg Physical Properties

Physical Properties Units Value

Nominal Prepreg Mass g/m2 294

Theoretical Cured Ply Thickness Mm 0.184

Theoretical Fibre Volume % 59.2

Theoretical Laminate Density g/cm3 1.58


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• Woven prepreg, M21 Resin, AS4C-6K Fibres [Hexcel, 2018]

Table 6.3 Woven Prepreg Mechanical Properties

Mechanical Properties Units Value

Tension Strength MPa 885

Tension Modulus GPa 67.6

Compression Strength MPa 835

Compression Modulus GPa 59.7

In-plane Shear Strength MPa 23

In-plane Shear Modulus GPa 23

Table 6.4 Woven Prepreg Physical Properties

Physical Properties Units Value

Nominal Prepreg Mass g/m2 475

Theoretical Cured Ply Thickness Mm 0.285

Theoretical Fibre Volume % 55.8

Theoretical Laminate Density g/cm3 1.56


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6.2 FEM Analysis Methodology

Figure 6.1 Step 1: Conversion of 3D

Solid Body CAD model into Surface

2D model

• Step 1: Conversion of 3D Solid Body CAD model into Surface 2D

model

The prepared parametric CAD solid body model, which can be

used for weight estimation or components arrangement

planning, has to be transformed into a surface model to perform

strength calculations. The needed 2D model is created by

isolating the surface of selected elements and sections. The

external surfaces of elements or contact surfaces are the most

frequently selected for export.

Figure 6.2 Step 2: Division of the

model into zones with the same

laminate system

• Step 2: Division of the model into zones with the same laminate

system

After generating the surface model, it should be divided into

appropriate zones depending on the fabric layout. Each

composite element has a laminate layout adapted to the load

pattern. To prepare the FEM model, the area of the element

should be divided into zones in which the composite system is

different.

Figure 6.3 Step 3: Definition of

finite element mesh in all zones

• Step 3: Definition of finite element mesh in all zones

After dividing the surface element into zones, it is necessary to

define a finite element mesh in each zone. Depending on the size

of the model, the desired accuracy, and the computing

capabilities of the computer used, the appropriate size and type

of finite elements should be selected. In most cases, large

elements are used for the first calculations, then the mesh is

refined in zones where the stress concentration occurs.


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Figure 6.4 Step 4: Control of the

material orientation in the mesh

• Step 4: Control of the material orientation in the mesh

Due to the fact that the carbon fibre reinforced composite used in

the CENTRELINE project is not an isotropic material, the main axis

of the material in finite elements should be checked and corrected

if necessary. The surface normal vector should also be under

control to build a laminate thickness in the right direction.

Figure 6.5 Step 5: Laminate layup definition

• Step 5: Laminate layup definition

The final stage of building the FEM model is to

define the laminate layout in each of the

predefined zones. At this stage, the angle of

the fibre positioning, the thickness of the

individual layers and the type of fabric or

sandwich material are determined. The angle

of the fibre arrangement should be defined

taking into account the predefined main axis

of the material defined in the previous step. In

the Laminate Modeler, the reference plane

relative to which the laminate will be built

should also be determined.

The same methodology was used for preparing the FEM analysis of Fuselage Fan Nacelle and Initial PFC

design aft-fuselage section [Goraj, Kowalski et al. 2018].

6.3 Centre-Wing Mesh Definition

In order to carry out the strength analysis of the centre-wing, a finite element mesh was generated for the

surface model. The outer shell was modelled using CQUAD4 four-node elements with a maximum size of

40 mm. The Stringers and stiffening struts were modelled with one-dimensional BEAM elements. The

stringer elements have a rectangular cross section with a size of 20 mm, while the struts elements have a

circular cross section, and the maximum element size is equal to 20 mm. The frames were modelled using

four-node two-dimensional CQUAD4 elements with a maximum size of 20 mm. The obtained mesh is

presented in the Figure 6.6.


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Figure 6.6 Centre-wing mesh

6.4 Geodetic and Conventional Fuselage Structure Analysis

6.4.1 Analysis of Conventional Fuselage Structural Layout

The FEM analysis was carried out for a section of the fuselage with a diameter of 6.4 m and a length equal

to 10 m. A conventional structure was used. The purpose of the calculations was to check, how the number

of longerons and frames affect the rigidity of the structure and to compare the obtained results with those

obtained for the geodetic structure. The fuselage model was fixed on the one side, with a bending

moment applied to the other side of the sample. An internal pressure of 0.0763 MPa (differential pressure

between cabin altitude FL70 and cruise altitude – FL410 multiplied by safety factor) was applied to the

inner surface of the cylindrical section, and the bending moment of 1500 Nm was applied to the ending

side of the fuselage model Loads and constraints are presented in Figure 6.7.

Figure 6.7 Loads and constraints applied to the fuselage cylinder structure

The influence of three parameters including number of longerons, number of frames and thickness of the

frame on the state of stress, strain and displacement was investigated. The initial case – a baseline –

corresponds to a fuselage with 80 longerons 1mm thick and 20 frames with a thickness of 2 mm. For all

cases the outer shell consists of four layers of carbon fibre fabric and a 1mm thick polyurethane core. The

total thickness of the external shell is therefore equal to 2.32 mm. Version 1 was treated as a reference


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case (baseline). In Version 2 the number of longerons has been changed to 60. In Version 3 both the

number of longerons and frames were reduced to 60 and 12, respectively. Version 4 was equipped with 80

longerons and 12 frames, the fifth version is analogous to version 4 but the thickness of the frames was

changed to 4mm, while Version 6 is a case similar to 3 but with frames of 4mm thickness. The differences

between individual versions are illustrated in Table 6.5.

Table 6.5 Specification of the calculated versions of the conventional fuselage

Table 6.5 contains also the information about the total weight for each version. For the baseline, the total

weight of the analysed structure is equal to 786.8 kg. The weight of the external shell is 433 kg and it is the

same for all versions being tested. For the baseline the skin constitutes 55% of the total weight. In the FEM

modelling of the fuselage section, two dimensional four-node elements were used. The maximum size of

the finite elements used in the case of the external shell is 50 mm, while for the longerons and frames it is

30 mm. Conducted analyses of the conventional structure have shown that for the baseline, the

displacement of the “grid” created by the longerons and frames, yielded 36.04 mm. A contour map of the

displacement in the load carrying structure is shown in Figure 6.8.

Figure 6.8 Displacement of load carrying structure of the conventional fuselage (version.1)

The modification introduced in the second version consisted in the reduction of the number of longerons

to 60. Due to this modification, it was possible to achieve both an increase in stiffness and a reduction in

Version Number of

stringers

Number of

Frames

Frame Thickness

[mm]

Weight [kg] +/- % Weight

1 80 20 2 786.79 0

2 60 20 2 742.79 -5.6

3 60 12 2 671.30 -14.7

4 80 12 2 715.30 -9

5 80 12 4 822.53 +4.5

6 60 12 4 778.54 -1


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the weight of the structure. In Version 2, displacement of frames and longerons decreased to 29.97 mm,

which means that the result is improved by 17% compared to the baseline. The weight of the structure has

also been reduced by 5.6% compared to the baseline. In this case, the reduction of the number of

longerons caused a stiffening of the structure. This effect was achieved probably by increasing the rigidity

of frames by limiting the number of cut-outs for longerons. In the applied load scenario, the frames are

under higher loads than longerons and therefore more responsible for transferring the loads. In the third

version, in addition to reducing the number of longerons (to 60 from 80), the number of frames has also

been reduced – to 12 from 20. In this case, the displacement of the grid is 32.03 mm, which is 11.1% less

than for the baseline. This result is slightly worse than in Version 2, but the weight of the structure has

decreased significantly. Compared to the baseline, Version 3 is lighter by almost 15%. In the fourth version,

the number of longerons was kept equal to that of the baseline, and the number of frames was reduced to

12. In this case, the rigidity of the structure was decreased. Displacement equal to 38.25 mm was recorded

which means it is higher by 6.1% than in the case of the baseline. Analysis of this case confirms the

conclusions drawn after calculating version 2. Because the frames are responsible for carrying the loads, so

increasing the number of cut-outs in frames in Version 4 resulted in a reduction of stiffness. To obtain a

better rigidity, the thickness of the frames for cases 5 and 6 was set to 4 mm, twice as high as in cases 1-4.

In version 5, in which 80 longerons and 12 frames were used, it was possible to obtain a stiffness

improvement of 36.4%, however with an increase of the total weight of the structure by 4.5%. In the case

of Version 6, in which the number of longerons was decreased by 20 (from 80 to 60), the maximum

displacement of the load carrying structure was equal to 18.6 mm, which is almost 50% less than in the

baseline. This effect was achieved with a simultaneous, slight weight profit. The weight of the Version 6 is

about 1% lower than the weight of the baseline. The comparison of results obtained for all versions is

presented in Table 6.6.

Table 6.6 Comparison of displacements of the conventional fuselage in each calculated version

Version Shell displacement [mm] „Grid” displacement [mm] +/- % (grid displacement)

1 18.95 36.04 0

2 25.09 29.97 -16.8

3 24.59 32.03 -11.1

4 19.69 38.25 +6.1

5 18.71 22.91 -36.4

6 24.54 18.60 -48.4

Table 6.6 also shows the displacements of the outer shell for all six versions. It can be observed, that the

displacements depend mostly on the number of longerons. For cases 2, 3 and 6, where the number of

longerons is 60, the displacements of the outer shell are between 24.5 and 25.1 mm, while for the

remaining cases, where the number of longerons is equal to 80, the displacement values are between 18.7

and 19.7 mm. These displacements are dependent on the space between the longerons and frames,

through which the shell is subject to the internal pressure. The smallest value of the displacement in the

outer shell is obtained in Version 5. The contour map of the outer shell displacement in Version 5 is shown

in Figure 6.9.


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Figure 6.9 External shell displacements in the conventional fuselage (version. 5)

6.4.2 Analysis of Geodetic Fuselage Structural Layout

The prepared parametric CAD 3D model was used to investigate the impact of a geodetic grid

configuration on the stiffness of the fuselage section. The model of the fuselage section had length of 10 m

and a diameter equal to 6.4 m. The parameters that were being changed in the FEM simulations included

the number of hoop ribs, number of helical ribs and the angle between hoop ribs. Ribs were modelled with

a square cross sections 10 by 10 mm for each version. The laminate layup was the same in every case as

well. The external shell consists of 4 carbon fibre woven fabric plies and 1 mm polyurethane core layer in

the following scheme: 0/90 deg, +-45 deg, PU, +-45 deg, 0/90 deg. In the geodetic fuselage concept, the

external shell is responsible mainly for transferring the internal pressure to the load carrying composite rib

grid.

Table 6.7 Angles between ribs, ribs numbers, frame weights and increments/decrements in percent

Version Angle between helical

ribs

Helical ribs

number

Hoop ribs

number

Frame weight

[kg]

Total

weight

[kg]

+/-

[%]

1 60 80 24 224.76 658.54 0

2 60 100 24 261.75 695.52 +5.6

3 60 80 41 279.16 712.93 +8.2

4 90 80 24 257.06 690.83 +4.9

5 50 80 24 218.21 651.98 -1

6 50 80 31 240.61 684.38 +4

Table 6.7 shows the differences between different computational versions. The first version (Ver.1) was

treated as a reference case and it consists of 80 helical ribs and 24 hoop ribs, while the angle between

helical ribs is set to 60 degree. Versions 1-5 differ from each other only by one parameter, to check the

sensitivity of the model versus the angles between ribs, helical and hoop ribs numbers and frame weights.

In Ver. 2, the number of helical ribs was increased to 100. Ver. 3 consists of 41 hoop ribs. In Ver. 4 the angle


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between helical ribs was increased to 90 deg. and in the Ver. 5 the angle was decreased to 50 deg. Table 3

also shows how the weight of the model has changed due to the performed modifications. External shell

weight was the same for all variants, the weight differences were caused only by the geodetic grid

modifications. The lattice structure was modelled using one dimensional finite elements with the size of 20

mm, while for modelling of the external shell 2D the 4 nodes elements were used. For the reference case,

the number of 1D elements was equal to 72760, whereas the number of 2D elements is 599689. Those

numbers change insignificantly depending on the computing version. The fuselage part under

consideration was loaded with internal pressure of 0.0763 MPa, and with a bending moment of 1500 Nm (it

is a synthetic load scenario) applied to the ending side of the geodetic cylinder structure. The degrees of

freedom were fixed at the unloaded end of the fuselage model. Loads and constraints are presented at

Figure 6.7.

Table 6.8 Results FEM simulation

Version Max Disp. Skin [mm] Max Disp. Frame [mm]

Max Stress

In frame

[MPa]

+/-Frame Displacement [%]

1 26.35 21.13 770 0

2 26.01 19.07 856 -9.7

3 22.43 14.89 555.16 -29.5

4 29.55 22.57 721.71 +6.8

5 27.41 18.17 723.28 -14

6 24.73 15.95 706.16 -24.7

Figure 6.10 Ribs displacement in Ver. 1

In the reference version the maximum displacement of the geodetic grid was 21.13 mm (Figure 6.10), while

the maximum displacement of the external shell was 26.35 mm. This difference is due to internal pressure,

which pushes the skin outwards. The highest recorded value of stress in ribs reached 770 MPa.

For the Ver.2, where the number of helical ribs was increased to 100, the maximum lattice ribs

displacement value of decreased to 19 mm, which is almost 10% lower compared to the reference variant


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displacement. Unfortunately, the better stiffness was obtained at the expense of a weight increased by

5.6 % compared to the Ver. 1. Another revealed undesirable effect was connected to the increase in

stresses observed in the load carrying grid. In this case the stress level exceeded 850 MPa, which is almost

100 MPa more than in the reference version. The displacement of the external skin was slightly decreased,

which is probably caused by the smaller distance between the ribs. For Ver.2 (Table 6.7 shows the

differences between different computational versions. The first version (Ver.1) was treated as a reference

case and it consists of 80 helical ribs and 24 hoop ribs, while the angle between helical ribs is set to 60

degree. Versions 1-5 differ from each other only by one parameter, to check the sensitivity of the model

versus the angles between ribs, helical and hoop ribs numbers and frame weights. In Ver. 2, the number of

helical ribs was increased to 100. Ver. 3 consists of 41 hoop ribs. In Ver. 4 the angle between helical ribs was

increased to 90 deg. and in the Ver. 5 the angle was decreased to 50 deg. Table 3 also shows how the

weight of the model has changed due to the performed modifications. External shell weight was the same

for all variants, the weight differences were caused only by the geodetic grid modifications. The lattice

structure was modelled using one dimensional finite elements with the size of 20 mm, while for modelling

of the external shell 2D the 4 nodes elements were used. For the reference case, the number of 1D

elements was equal to 72760, whereas the number of 2D elements is 599689. Those numbers change

insignificantly depending on the computing version. The fuselage part under consideration was loaded

with internal pressure of 0.0763 MPa, and with a bending moment of 1500 Nm (it is a synthetic load

scenario) applied to the ending side of the geodetic cylinder structure. The degrees of freedom were fixed

at the unloaded end of the fuselage model. Loads and constraints are presented at Figure 6.7.

Table 6.8) this displacement is equal to 26.01 mm, which is about 1.5 % lower than in the reference version

(i.e. Ver. 1). Figure 6.11 shows the increased displacement arrows of geodetic skin in a scale of 10 to 1.

Figure 6.11 Geodetic cylinder structure - skin displacement Ver.2

In the Ver.3, the number of hoop ribs was increased, which resulted with an increase of load carrying grid

weight by 8.2 %. Weight increase, however, is compensated by noticeable improvement in stiffness of the

load carrying structure. The maximum value of displacement of the geodetic grid for Ver.3 is 14.89 mm,

which is almost 30% lower than in the case of reference scenario. A significant reduction in maximum


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stress can be also observed – the maximum stress value in ribs was equal to 770 MPa in the case of the

reference scenario, whereas it was decreased to 555 MPa in the case of Ver.3. A map presenting the von

Mises stresses in the geodetic fuselage load carrying grid is presented at Figure 6.12. The maximum

displacement in the composite skin was also decreased. In the case of Ver. 3 it is under 23 mm, which is 10%

lower than in the reference case with the geodetic cylinder structure.

Figure 6.12 Geodetic cylinder structure - Stresses in frame structure Ver. 3

The variant with more hoop ribs proved to be more beneficial, despite the increase in weight. In the next

lattice structure version (Ver. 4), the angle between helical ribs was changed to 90 deg. In this case, a

deterioration of stiffness and increase in weight by almost 5% with respect to the reference case, was

observed. The maximum calculated displacement in the grid was 22.57 mm, which is 6.8% higher than in

the reference case. Stresses in the geodetic grid decreased slightly compared to the reference Ver.1, but

the maximum displacement in skin was raised to almost 30 mm. In consequence the conclusion can be

made, that the increase of the angle between helical ribs negatively affects both weight and stiffness of

the structure.

In Ver. 5, the spiral ribs are crossing at angles of 50 degrees. Due to that modification, it is possible to

achieve an increase in overall stiffness and at the same time to obtain a slight weight loss. The maximum

displacement in the ribs grid was 18.17 mm, which is 14 % less than in the base variant. The weight of the

design model has decreased by 1 %. The only undesirable result of this change is related to the increase in

stresses in the external shell.

The carried out analysis proved that increasing the number of hoop ribs and reducing an angle between

helical ribs has the most beneficial impact on the fuselage stiffness.

Based on that conclusion, an additional version was analysed, in which both above mentioned parameters

(i.e. number of hoop ribs and angle between helical ribs) were changed. In Ver. 6 the design model was

equipped with 31 hoop ribs, while the angle between helical ribs was decreased to 50 deg. The purpose of

these changes was to achieve stiffness similar to that of the case in Ver. 3, but with a smaller weight

penalty. This goal was successfully achieved.


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Figure 6.13 Ribs displacement in a scale of 10 to 1, in Ver. 6.

The maximum displacement of the lattice grid was reduced to 15.95 mm, which is almost 25% less than for

the reference case. The obtained stiffness is about 7 % worse than that of the Ver. 3, but the weight of the

frame only was increased by less than 16 kg, while for Ver. 3 it was equal to 55 kg. Unfortunately, the

decrease in the maximum stress values is not so big compared to that of the Ver. 3, which was 28 % stiffer

than in the reference version. Comparing results for geodetic and conventional versions with similar

displacement values (geodetic Ver. 4 with conventional Ver. 5, and geodetic Ver. 5 with conventional

Ver.6), it can be observed, that the weight of the fuselage sample in the conventional version is about 19%

higher than in case of the geodetic fuselage sample. The analysis of several geodetic configurations have

shown, that the most promising results can be obtained for the Ver. 5 and Ver. 6, and these are the

versions selected for further analysis.


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7 Conclusion and Future Work

This report summarises all actions taken so far as a part of task T2.4.2 aiming for predesign of fuselage-

wing junction integration. Work methodology was presented: starting from aircraft external geometry in

CPACS language to internal structure designed in NX software. Moreover, 3D CAD model of inner

structure parametrization was described so as to enable in upcoming part of task quicker and reliable

optimization of fuselage structure.

Two kinds of fuselage structure concepts were presented: the first one – currently used so called

conventional one and the second one: a geodetic structure. These two structures have been taken under

discussion and analysis. As a result, the best promising fuselage concept have been chosen as a

combination of these two ideas.

The central parts of the designed fuselage that concentrate all aircraft loads have been described.

Moreover, the analysis of the CS-25 regulation load cases has been presented. On its basis and study of

components of the central fuselage, critical loads cases acting on PFC aircraft were chosen, which will be

implemented in the FEM calculation in second stage of this task.

The reports concluding section about the interim results presents the material selection and its properties,

describes the structure mesh preparation for the FEM analysis and results of the comparison of the classic

versus the geodetic fuselage structure.

The following main tasks need to be performed by the end of the task T2.4.2 month 26:

• Calculation of all of the defined in this report load cases

• Load implementation into FEM analysis. Performing FEM analysis and critical evaluation of the

results, will be performed repeatedly to obtain the best structure concept and therefore

redesigning the structure if needed and weight estimation update.

• Perform a functional hazard analysis to evaluate concept intrinsic redundancy and ensure

robustness against relevant failure modes.

• Harmonization of technical assessment of fuselage propulsor maintainability, reliability and safety.

Preparation of aerodynamic analysis that will allow to examine the influence of selected structural

decisions on flow field in the fuselage fan plane.


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8 Bibliography

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Increasing in Worst Case Situations and Joints Minimizing.” Final Progress Report 2010,

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[Airbus, 2018] Airbus, „Airbus’ new centre wing box design holds great promise for future aircraft”,

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[Khairi, 2010] Khairi, Y., Nukman Y., Dawal, S., Z., Devi C., Sofia, N., “Conceptual Design of Fuselage

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[Michalski, 2016] Michalski, K.: „Concept and preliminary strength analysis of composite geodetic

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