Meng_Aero_Vergara_Camilo

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DEPARTMENT OF MECHANICAL ENGINEERING

SCHOOL OF ENGINEERING AND DESIGN

BRUNEL UNIVERSITY

Conceptual Design and Structural Analysis of an 80 PAX

Aircraft

By Camilo Vergara

MARCH 2013

Supervisor: Dr Narcis Ursache

ABSTRACT

This report describes and follows the steps of conceptual design of an original 80 PAX

regional aircraft from its sketch, initial sizing, and composition with the use of computer

aided methods. It also suggests ways of improving existing structures and novel

configurations.

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Contents |Page

1| Introduction………………………………………………………………………………………………………………….8

- 1.1 Comparison of existing techniques………………………..……...………….……...……9

- 1.2 Project aim………………………………………………..……………….……………….…….….14

2|Literature Review……………………………………………………………………………….………………………..15

3| Project Plan……………………………………………………………………………………………………………..….19

4|Conceptual Sketch.……………………………….…………..………..……………………………………………….20

5|Aircraft Sizing…………….…………………………………………………………………………………………………24

6|Final Geometry Selection & Configuration…………………………………………….……………………..32

- 6.1 Wing…………………………………….………………………………………………………………32

- 6.2 Tail……………………………………….………………………………………………………………34

- 6.3 Fuselage……………………………….………………………………………………………………34

- 6.4 Canard………………………………….………………………………………………………………35

7|Modal Analysis……………………………………………………………………………………………………..………38

8|Aero-elastic Analysis………………………………………………………………..……………………………………44

- 8.1 Static Analysis…..………….…….………………………………………………..……………..44

- 8.2 Structural Considerations…………………………………..………………………………..52

- 8.3 Rigid VLM/DLM………………….……………………………………………...…..…………...55

9|Comparison of XATA against a standard configuration………………………………………………….61

10|Conclusions…………………………………………………………………………………………………………………64

- 10.1 Report Findings….……………………………………………………………………………..64

- 10.2 Validation of results………………………………………………………………….……….65

- 10.3 Critical Analysis………………………………………………………………………….………66

11|References………………………………………………………………………………………………………………….67

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12|Appendix…………………………………………………………………………………………………………………….70

Table of Figures |Page

1.1 CEASIOM chart [1] 10

1.2 NeoCASS chart [5] 10

1.3 RDS Software chart [3] 12

3.1 Gant Chart 19

4.2 Conceptual Design Wheel [8] 21

4.3 Initial Design Ideas 70

4.4 LamAiR Project [17] 71

4.5 Cabin Parameters; Passenger, Crew and Baggage Configuration 22

4.6 Fuel Parameters; Fuel Tank Location throughout Fuselage and Wing 22

5.1 Typical Mission Profile of a Commercial Jet 24

5.2 Sizing from a Conceptual Sketch [8] 25

5.3 NACA 64A010 Aerofoil Properties [15] 26

5.4 Payload Trade Study for the XATA [8] 28

5.5 Range Trade for the XATA [8] 29

5.6 Range-Payload Trade Study [24] 29

5.7 Flow Diagram of Modules from NeoCASS used in Sizing Stage 30

5.8 GUESS Output plot Example 79

6.1 Side view of XATA 36

6.2 Front view of XATA 37

6.3 Top view of XATA 37

7.1 Beam mesh of XATA 38

7.2 Set-up for Modal Analysis 39

7.3 Settings Display for Modal Analysis 40

7.4.1 Mode 7 Frequency 5.0068Hz 40








8.1.1 Deformation plot of XATA at 1.0g [M 0.8; Zacc 1.0g; 9000m] 45

8.1.2 XATA at Half Capacity at 1.0g [M 0.8; Zacc 1.0g; 9000m] 45

8.2A The Relation of Sweep to the Divergence, Alieron Reversal & bending

46 Torsion Flutter Speeds [11]

8.1.3 XATA at 1.42g [M 0.5; Zacc 1.42g; 5000m] 46

8.1.4 XATA at 2.0g [M 0.5; Zacc 2.0g; 5000m] 47

8.1.5 XATA at -1.0g [M 0.5; Zacc -1.0g; 5000m] 48

8.1.6A XATA Composite Wings at Fig8.1.1 conditions [M 0.8; Zacc 1.0g; 9000m] 48

8.1.6B XATA Composite at Fig 8.1.3 conditions [M 0.5; 1.42g; 5000m] 49

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8.1.6C XATA Composite at fig 8.1.4 conditions [M 0.5; Zacc 2.0g; 5000m] 49

8.1.6D Composite Deformation [M 0.5; Zacc 2.5g; 5000m] 49

8.1.6E Composite XATA at Fig 8.1.5 conditions [M 0.5; Zacc -1.0g; 5000m] 50

8.1.8 V-n Diagram at 5000m 51

8.2.1 Location of Spars along Wingspan 52

8.2.2 Quasi-Unbalanced Smart Spar Fibre Layout & Structure Composition [20] 53

8.2.3 Composite Wing Fibre Orientation 54

8.3.1 XATA Normal Force Distribution plot on Aerodynamic Surfaces

55 [M 0.8; Zacc 1.0g; 9000m]

8.3.2 XATA [M 0.5; Zacc 2.0g; 5000m] 56

8.3.3 Rigid VLM/DLM [M 0.5; Zacc -1.0g; 5000m] 88

8.3.4A Location of Aero-Panel Chosen for Analysis at[M 0.8; Altittude 9000m] 87

8.3.4B Location of Aero-Panel Chosen for Analysis at[M 0.5; Altittude 5000m] 88

8.3.6 Set-up of Static Aero-elastic and Rigid DLM/VLM Analysis 60

9.1 Deformation plot on a Standard Configuration of Similar Size

61 [m 0.8; Zacc 1.0g; 9000m]

9.2 Deformation Comparison between Winglet & Non Winglet Design 61

9.3 Deformation of Standard Configuration [M 0.5; Zacc 2.0g; 5000m] 62

9.4 Rigid VLM/DLM Standard Configuration [M 0.5; 5000m] 63

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List of Tables |Page

4.1 Current Regional Aircraft Specifications 20

8.1.6 Table of Material Properties 47

8.1.7 Table of Material Properties of Carbon Composite AS-4 [13] 50

8.3.5 Table of Deformation Angles 58

10.2.1 GUESS Validation of Fuselage Weights Estimation Comparison with real 65

World Values [22]

10.2.2 GUESS Validation of Wing Weights Estimation Comparison with real 66

World Values [22]

Nomenclature

L= Lift

D= Drag

= Coefficient of Lift

= Coefficient of Drag

= Thrust Specific Fuel Consumption

= Empty Weight

= Gross Weight

= Density of Air with respect to altitude

s= Wingspan

M= Mach number

= Never Exceed Velocity (structural limit)

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= Stall Velocity

= Dive Velocity

E= Endurance

R= Range

= Zero Fuel Weight

Units

The units featured and used throughout this report are all consistent with the International

System of Units, (SI Units). And this all calculations were made using these prefixes.

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Acknowledgements

It was really quite a task completing the final report, sometimes time just seemed to

evaporate. All that was done during the week was eat, sleep, and project work.

I would like to give special mention to my family, who regardless of all the stress and work,

never failed to encourage for me to take a break every now and then. It is imperative to be

able to relax even whilst having such a big piece of work on your shoulders.

This leads me onto the next set of people, to my friends, my colleagues, my fellow Microsoft

word warriors. We all shared our Dissertation-syndrome together in the engineering towers,

slowly turning hysterical as the deadline seemed to appear out of nowhere.

Another needing a special mention is my girlfriend, who put up many times with my high

stress levels and me getting back in the early hours of the morning on several occasions.

Of course, where would I be without the academic support provided in the form of my

Supervisor Dr Narcis Ursache, who took the time out to make sure I was going in the right

direction and who readily available whenever asked for (literally at a moment’s notice).

Lastly a special mention goes to Professor Sergio Ricci, of Politecnico di Milano. It is due to

his patience with me that I managed to ultimately understand the software, and use it to

the degree that I did in this report. I wasn't always the easiest person to help out, but he

made sure to promptly reply whenever was possible.

If there’s anything that can be taken from this experience, it’s the fact that one should never

underestimate the power of a helping hand. To the people who support you, who

experience this alongside you day in, day out. You all made the madness bearable, and to

you, I say thank you.

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1| Introduction

Conceptual design is one if not, the most important stage in the process tree of

aircraft design and construction. It is in this stage in which concept and theory must be

proven in order to push forward onto the more detailed analysis. Any errors and issues not

fixed or discovered in this early stage can lead to disastrous delays in the later stages of

production.

In the early days of aircraft design, engineers used to have to manually work through

mountains of equations and work in order to validate a conceptual design, and even then

there was a significant margin of error due to inconsistencies and random errors. The drive

to automate parts of the process began in the 1940’s, with the Second World War making a

demand for major advances and pushing forward various technologies, including in

aerospace. From here began the effort to simplify conceptual design phases, in an attempt

to minimize design faults, and get rid of overlooked errors before expensive tests and

prototyping started. With the use of computers, software was developed and still is to this

day being improved and refined specifically with conceptual design in mind, to be able to

quickly and easily draw up ideas and prove them before moving onto more expensive stages

of development. In order to successfully complete the conceptual problem within the set

objectives and aims of the exercise two software modules called NeoCASS, and AcBuilder

were employed. As well as this research was done into various different software packages

and suites in order to get a fuller understanding of the processes, and steps universally

involved with designing, and optimising a conceptual aircraft.

There is a justifiable need for a project involving conceptual design, as a valuable insight will

be gained into current Industry techniques and trends, but in addition to this, the outcome

should be able to present viable design improvements and alternatives to current aircraft

configurations by the end of the process.

The main aim of this project is to understand the processes of conceptual design. This

requires knowledge of the current tools available to help with the optimisation procedures.

In this case it is the use of software specifically designed with aircraft design in mind.

Through this means of developing a conceptual construct, it is possible to analyse structural

design, aero elasticity, as well as model all forces acting on the fuselage and any

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deformations that may occur, these are all vital parameters that need to be accounted for

before the next step in the design phase can commence.

1.1 Comparison of Existing Numerical Methods

In order to successfully design and optimize a passenger aircraft, computer-aided methods

of aircraft design were implemented coupled with methods covered in Daniel P. Raymer’s

Aircraft Design: A Conceptual Approach. An initial construct was put together through

AcBuilder (Aircraft Builder) which was influenced heavily on current similarly sized aircraft

(for example Embraer’s E-170, British Aerospace’s 146-200, and the Comac ARJ21-700).

Once complete the aircraft is saved as an .xml file, which is then used in NeoCASS (Next

generation Conceptual Aero-structural Sizing Suite) for structural & Aero-elastic modelling

for proof of concept.

Both of the above modules come from the same umbrella software package that is called

CEASIOM (Computerized Environment for Aircraft Synthesis and Integrated Optimisation

Methods). The Objective of CEASIOM is to provide a wholesome package where all

predictions, computations, and optimizations of the early conceptual design phase can be

run through to give an accurate picture of how the conceptual aircraft will fair through

various conditions. This unlike most other counterpart programs is free of charge. Its

purpose is to provide a medium where the aerospace community can exchange ideas of

concepts and knowledge, freely and easily. All of CEASIOM’s modules run on Windows and

Linux, through the use of Matlab (version 2007 and later). CEASIOM is still in development,

but has support and links with multiple organisations such as Dassault Aviation, and CFS

Engineering [1]. Below is a diagram showing all the different branches of the CEASIOM

modules and how they interact with one another

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Figure 1.1[1]

For the purpose of narrowing a broad design problem, out of all the programs CEASIOM

encompasses, NeoCASS and AcBuilder are the relevant modules that seemed to deal with a

very specific part of the aircraft’s development phase. They are particularly suited for the

purpose of conceptual and preliminary design. In actuality the modules are explicitly tuned

for transport and passenger aircraft configurations. Below is a figure of a flow chart which

covers the function of the two software programs in relation to each other.

Figure1.2 [5]

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In addition to CEASIOM, there are alternate software solutions to the design problem of

conceptual development. A few of them require paid licensing for use such OAD and RDS

and are fully released and developed Packages used by engineers in industry.

OAD (whose software is known as ADS –Aircraft Design Software) like CEASIOM, is an in

depth Conceptual design tool. A significant difference is whereas CEASIOM is free, OAD

software incurs licensing costs. These range from £41.77 for the basic ‘ADS maker’ package

(which is intended for fans of the Flight simulator X game to design and fly a custom

aircraft), up to £2,043.81 for the ‘ADS professional’ suite which includes an optimisation

module to streamline preliminary designs, and is intended for industry use; annual

maintenance fee is 20% of original price [2]. Professional packages include similar modules

to CEASIOM; this includes load analysis, weight & balance, wing optimisation, integrated

CAD, and more. In addition there are extra features not found on the free software such as

the ability to export files onto other CAD programs, and the ability to directly export files

onto flight simulators. Much like CEASIOM, ADS has been in development for several years,

and has backing from various institutes and organisations including heavy weights such as

EADS defence & security, right round to the LAA (Light Aircraft Association) [2].

Another Conceptual design program that was looked into was RDS Raymer’s Aircraft Design.

Immediate similarities with the previous two were its ability to analyse loading, run initial

sizing optimisation, and weights analysis. Coming in at £12,269.20 per copy, plus 25% yearly

maintenance costs [3], it is realistically an option only for industry projects.

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Figure 1.3[3]

Ultimately RDS has a differently structured layout which begins with the design layout,

much like AcBuilder, but goes one step further in its latter stages than its counterparts by

including Cost estimation, and incorporating Range & sizing, as well as performance before

running it through its own dedicated optimiser. Figure 1.3 is a flow chart from the RDS

website to visualise the process which the designer goes through within the actual software.

It is worth mentioning that RDS can be used with many different types of aircraft and

spacecraft, ranging from passenger aircraft, to advanced fighter aircraft, reusable launch

vehicles, UAVs, and dynamic lift airships [6]. Users of RDS include Honeywell Engines &

Systems engineers [3].

Aside from CEASIOM there very few collections of individual standalone modules by third

party developers aimed at aiding the initial concept design stage. These are less complicated

programs designed with a specific task of tackling a single design aspect. Among these is

included:

XFOIL [4]: Written by a professor from MIT, this focuses on the analyses of subsonic airflow,

pressure distribution, effects of blended aerofoil designs, and the impact of flaps around a

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specified aerofoil design. It would be useful if the weight of my objectives were purely

focused on the aerofoil shape. However this is not the case, so a more complete package

capable of analysing in 3D is needed.

CalculiX [26]: Another free software solution aimed at conceptual design. This is a three

dimensional tool used for finite element analysis in a structure. However this in itself is

purely a structural analysis tool, nothing more. In stark contrast CEASIOM is a more

complete package & although still in its later stages of development and refinement, offers

as much as other suites which cost several thousands of pounds to operate.

The last package found worth a mention is open source software which belongs to NASA.

This is called VSP (Vehicle Sketch pad). It is essentially an aircraft geometry tool (like

acbuilder), but with a lot more options, and is much more detailed in its final geometry

compositions. In it you can model virtually any type of fuselage, and according to their

website [7], you can input aerodynamic preferences as well as mass properties. From the

demo video online, it states you can also model different engines, landing gears and their

movement patterns, for example it shows a model of a UAV with VTOL capabilities, and

shows how the engines spin 90 degrees from horizontal to vertical, as well as the landing

gears deploying underneath. From the research done on the background and functionality

of the software it was apparent that although it had superior abilities to sketch aircraft, it

deals purely with the geometric composition aspect, and features no type of in depth

analysis, and iterative processes, that define the design process at the conceptual stage.

Of the available methods that were researched, It can be confidently said that the objectives

can be achieved with the use of the CEASIOM modules NeoCASS & AcBuilder. Given the

circumstances, it is also the most economically viable option. Not only does it have the most

user friendly format (easily operable GUI), it encompasses the most important aspect of a

conceptual design program, the ability to run iterations on vital design parameters in either

pre-set, or custom flight manoeuvres, and optimise accordingly. In addition both NeoCASS

and AcBuilder run on Matlab, which is a vital tool and widely used worldwide by engineers.

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1.2 Project Aim

1] Design & optimise an original 80PAX aircraft

For this purpose the research proposal is the study and exercise in conceptual design of an

80PAX aircraft. Objectives will be to optimise and analyse the structure, aero elasticity and

forces acting on the theoretical aircraft. This should shed light on improvements that can be

made to current conventional configurations in order to get overall better performance and

aircraft characteristics at a conceptual level.

Objectives-

1) Improve original design

2) Analyse structure & modify mesh density

3) Analyse Modes of Vibration

4) Analyse Force distribution along aerofoil

5) Analyse static Aero-elastic (Trim properties)

6) Successfully complete Conceptual design phase (Prove concept)

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2| Literature Review

The materials required for background research and study into the subject of

conceptual aircraft design are readily available in the form of academic books written as

introductions to their respective fields, as well as declassified NASA flight program technical

reports for more in depth knowledge that is only referred to in the books. These were

invaluable in presenting much needed data, and information on the whole field of

conceptual design.

The first item with regards to progress on the chosen field of study was to get a full grasp

and thorough introduction to the general aspect of a design project. In order to do this

research was done into what other people had to say on the subject. The first book acquired

was ‘Introduction to Aircraft Design’ by John P Fielding. This served as a good starting point

for aircraft design. It provided a broad insight into all aspects of a design process. The book

touches upon many aspects that go beyond just the conceptual design; it covers specific

topics from a broad angle including civil and military aircraft from an analytical standpoint. It

included structural diagrams and information about different types of configurations within

the two classifications on aircraft. What it does is introduce the basics of structure design

considerations with respect to the loads experienced on the aircraft during normal

operation, the basics of the aircraft systems, as well as various sub sections which touched

upon topics such as fatigue and the causes of failure of certain components. It included

information and diagrams used to display the purpose of each design feature. The

information in the later stages however was tailored towards more military designs, and

served little use in the type of civil design that is done in this report.

Even though a general introduction had been made it was felt that there needed to be more

information from the perspective of aircraft handling characteristics. ‘Design for flying’ by

David Thurston covers this aspect as the descriptive form of the book aids with more

information about certain existing configurations, as well as a brief history of their

development. It serves as a brief introduction to the factors affecting performance, flight

operation and reasons for which the design process often is a continuous loop of

optimization for key flight parameters. It finishes with information about aircraft

certification, and the phases associated with this process.

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With a better understanding of what it means to produce an aircraft, the amount of work

that is associated with such an undertaking to be done and the number of factors affecting a

design, a closer inspection of the design phase at just the conceptual step was required.

‘Aircraft Design: A Conceptual Approach’ by Daniel P Raymer was chosen due to the link it

had of the author designing some of the software previously mentioned in the introduction

(RDS). Although the general phases of design had been shown in the previous books, A

Conceptual Approach deals purely with the concept stage of development. It is widely

considered as the key literature to which to base a primary conceptual design, it holds a

complete guide to developing a full concept to the preliminary stage, with the ability to

optimize it to such a degree that little or no alterations may be made on the detailed design.

Several key aspects which aid in the initial stages of conceptual design were applied

throughout the report. Range and trade sizings to estimate aircraft performance, basic

aircraft geometry arrangement from the shape of the fuselage, tail arrangement, engine

locations to any additional geometric components, aerodynamic & structural considerations

on the chosen design and weights comparison to general industry standard classes of

aircraft were all derived from this book. It was used as a guide to generate the initial

conceptual sketches, and from there start upon the sizing and estimating of aircraft

performance to provide a solid design base to analyse structurally.

With background research covering the steps involved in a conceptual design, the chosen

analytical methodology needed to be researched as-well. ‘Aero-elasticity’ by Raymond L

Bisplinghoff, Holt Ashley, Robert L Halfman provided an insight into the field of structural

analysis. Where the previous books have focused on the whole concept design phases, this

particular book explains in-depth the mathematical theory and steps for analysing and

calculating the bending structural divergence experienced by the wings during flight, as-well

as provide an insight into analysis of natural modes and frequencies of complicated aircraft

structures. This is something that is core to the static aero-elastic and modal analysis in

NeoCASS. The program uses the matrices to define the structure at each point drawn up

from AcBuilder and GUESS as well as the structural material properties, and assigns forces at

each section of the mesh, under the set flight condition. For Modal analysis, it sets

vibrations at each node and calculates accordingly to find natural frequency of the

structure. For static aero-elastic analysis, the wings along with any other control surfaces

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experiencing loads are modelled like Cantilever beams all connected to the fuselage. Force,

moment and element displacement is then calculated by NeoCASS through the use of the

aforementioned matrices, and plotted on a 3D chart in Matlab. Any error due to the effects

of sweep on the coupling of bending and torsional actions along the wing is negligent and

minimal up until angles of 45°.

Further information with respect to the structure was thought to be crucial in giving a wider

picture of the whole subject. ‘Fundamentals of Aircraft Structural Analysis’ by Howard D

Curtis provided key information on modelling the geometry of a vehicle to analyse it

statically. What the book was further used for was the understanding behind the choice of

beam mesh composition by the program NeoCASS. It helped understand the reasons as to

why it is commonly used in order to analyse a developing vehicle structure, and allow an

accurate portrayal of forces in virtual 3D space. In conceptual design the structure is

simplified to the minimum to calculate initial properties in operation. Beams are ideal for

this purpose as the manner in which the structure is put together means they experience

shear , displacement and moments about their axis in conjunction to the other beams next

to them. In essence creating the mesh out of beams means that the end result acts like a

differential beam segment, where there are so many that the subtle changes between each

of them mould together and the structure acts as one integral component. NeoCASS is not

the only design program to incorporate this; others such as ANSYS [16] use a beam mesh to

model components for analysis.

All the information already reviewed appeared to omit any solid form of information for

forward swept wings, the topic was touched upon but very briefly in all the books reviewed.

Upon further study there seemed to be a constant mention of a forward swept test aircraft

X-29. ‘X-29 Forward Swept Wing Flight Program Status’ by Gary A Trippensee, David P Lux

helped with this respect. Although theoretical and general information was presented in the

books used for the purpose of developing a valid design, the mention of the X-29 testing

aircraft used by NASA to develop data on forward swept wings and this report was

invaluable as it provided solid official information on the characteristics of such a

configuration. Using actual original reports helped to validate and clear up any uncertain

information regarding the prototype aircraft.

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Further information on the way forward sweep affects wings in terms of divergence during

flight was given by ‘Wind Tunnel Experiments on Divergence of Forward Swept Wings’ by

Rodney H Ricketts, Robert V Doggett Jr. When the aircraft geometry was in the process of

selection from the initial ideas more specific information was required on the behaviour and

relationship of forward sweep in wings, flutter frequencies and the possible effect on static

divergence. This paper aided in a much fuller understanding on the properties of forward

sweep not only on the divergence, but on the subsequent modal values of the wing.

Selection of an appropriate aerofoil from those provided in AcBuilder required information

and characteristics of the ones available. Of these, the one chosen was the NACA 64A010.

‘Tests of the NACA 64-010 and 64A010 Airfoil Sections At High Subsonic Mach Numbers’ by

Albert D Hemenover was used to as the base on which to get the appropriate data used in

the aerofoil of the conceptual design, it was necessary to acquire more information of the

aforementioned for use in and calculations which were further used in approximate

initial range and endurance. In addition, diagram presented in the paper helped in the

optimization of the wing incidence angle for cruise conditions through the iterations of the

Lift equation to achieve the condition Lift = Weight.

The aforementioned equations for lift and Drag at cruising speed combined with

assumptions stated in Raymer’s ‘Aircraft Design A Conceptual Approach’, as well as basic

range and endurance for a jet aircraft provided initial estimation of basic performance of

the aircraft were acquired from the Lecture slides of the aircraft design module of the

course at Brunel University. As well as these, some from the airworthiness module were

also incorporated in to help plot diagrams used in the sizing section of this report.

These sources were chosen to be due to the nature of the main topic of this report. Aircraft

design is a very complicated and large field of study, it is not something that is to be taken

lightly as chances are lives will depend upon the outcome of whether an aircraft was well

designed or not. Usually a design would involve many different teams of engineers combing

through each and every parameter of the aeroplane. For this reason the aim has been

narrowed down to a much more realistic target for one person to achieve. The emphasis has

been put on structure as well as general performance values to back up any type of

configuration that is developed.

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3| Project Plan

Figure 3.1

Gant chart

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4| Conceptual sketch

From the first initial sketch, various design implements have been considered and

scrapped. The location and placement of each component externally, as well as their

internal loads of the XATA (Abbreviation for the conceptual aircraft developed in this report)

series conceptual designs has been considered extensively. The design started off based

from data of similar sized passenger Aircraft of varying configurations, this included

performance data from aircraft like the E170 80 PAX (Embraer) through to the Chinese built

ARJ21-700 which is 70-95 PAX. This is a kind of ‘wish list’ and guidelines as to the current

standard aircraft in use. Below on figure 4.1 is the table with data researched about each

aircraft.

80 PAX Aircraft preliminary design variables list

Flight Parameters E170 LR Bae 146-200 CRJ-700LR ARJ21-700 X-

ATA

Wingspan 26.0 m 26.34 m 23.2 m 27.28 m TBA

TBA

Max take-off weight 37,200 kg 42,184 kg 34,930 kg 40,500 kg TBA

Max Landing weight 32,800 kg 36,741 kg 30 660 kg

TBA

Take-off Distance 1,690 m 1,550 m 1,850 m 1,700 m TBA

Landing Distance 1,160 m 1,200 m 1,560 m

TBA

height 9.8 m 8.61 m 7.2 m 8.44 m TBA

length 29.9 m 28.55 m 32.5 m 33.46 m TBA

range 3,889 km 2,963 km 3,700 km 2,200 km TBA

Maximum Altitude 11,900 m 9,509.76 m 12,500 m 11,900 m TBA

fuel capacity 9,470 kg 10,300 kg 8,820 kg 10,100 kg TBA

Cruise speed 890 km/h 800 km/h 830 km/h 870 km/h TBA

passenger capacity 70-78 85-115 70-79 78-90 80

Table 4.1 Current Regional Aircraft Specifications

Based upon this information, in the early stages of development, several designs were

drawn up, with the basic wing length, fuselage length and weight values from its real life

counterpart.

P a g e | 21

Requirements

Sizing Trade

Studies

Design Concept

Design

Analysis

Figure 4.2[8]

AcBuilder, a module from NeoCASS was employed to help shape and put together the

designs from the very basic initial criteria.

Figure 4.2 shows: the design wheel. This depicts the whole conceptual phase in a compact

and visible manner. As can be seen, the design is often changed non-stop throughout the

complete process.

Figure 4.3 in the appendix shows the various initial different designs initially created,

alongside alterations of the main idea focused on in later stages of the proposed aircraft.

Where-as the others were interesting concepts, the XATAV5 in particular I felt was a

promising design. The subject of forward sweep in wings is often overlooked and not widely

used by commercial aircraft, However it does have benefits that are referred to in a later

stage of this report. As well as this it appeared as the most feasible design allowing for

considerations of future trends in the industry. In addition an on-going study was found that

mentioned the positive effects of negative sweep on the website of NASA partner, DLR

(German aerospace centre). This mentions briefly that the aspect of a forward swept wing

may help meet future milestones of aircraft efficiency: ‘By 2020, aircraft are supposed to be

50 percent more economical’ [17]. It mentions that the flow is largely laminar due to the

forward sweep which helps to decrease aerodynamic drag. A diagram of the aircraft being

developed by them is included in the appendix as figure 4.4. This further indicated that the

best initial aircraft design to develop was the XATAv5.

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Figure 4.5 Cabin parameters; Passenger, Crew and Baggage Configuration

Using AcBuilder as the means to generate a detailed conceptual sketch presented many

advantages, as it was possible to input an initial computational value for all the required

components of the aircraft. As-well as this it was possible to modify vital design aspects and

have them presented visually. This can be seen in figures 4.5 and 4.6.

This is invaluable as a tool as it gives the opportunity to compact all the conceptual design

features, not just the ability to sketch it in virtual space into the same process. This means as

the geometry is sketched out; values for weight and various other internal components can

be assigned instantaneously in parallel with the sketch giving a much more thorough picture

of the aircraft at the beginning.

Figure 4.6 Fuel parameters; Fuel Tank Location throughout Fuselage and Wing

P a g e | 23

In addition to these Acbuilder allows for the modification of the mesh used in analysis at a

later stage, and several material and weight inputs that define the structure.

These advanced options to modify the way the structure is defined in 3D space comes under

the ‘technology’ tab of AcBuilder, where it is possible to alter the Beam and Aerodynamic

panel meshes, the material composition, the loading limits, the ability to not define

components, and it is possible to alter the number of nodes in the model analysed through

GUESS.

From this initial geometry composition the design has been further altered and fine-tuned

as the stages become more detailed, and more emphasis is placed on predicted

performance values, stability, and aircraft balance.

Throughout this project alongside the software and during the design process of the

conceptual passenger aircraft, manual calculations will be done not only to corroborate

results, but to take a step further past the limitations of the software and extend the design

processes to other areas not covered by NeoCASS. This will further aid in the successful

proof of concept ready for the next phase of development.

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5| Aircraft Sizing

Every Conceptual design starts with an idea or a sketch. In this report this is done in

CEASIOM’s geometry module Acbuilder. At this stage it only has to be a rough

representation of the aircraft. Following the concept design sketch, comes the next stage,

the first-guess sizing. It is here where all the basic parameters of the aircraft are calculated.

It is widely regarded as the most important calculation in the aircraft design, and if they are

unsatisfactory go back to the sketch, modify it, and see the results of the improvements. In

essence sizing is defined as the design of the aircraft with regards to limitations and

expressed requirements by calculation, and how that affects how big the vehicle and its

various components are drawn up. One of the first steps to help clarify the purpose of the

new design is to draw up a mission profile. Below is Diagram drawn up for the aircraft being

developed in this project.

Mission Profile:

Figure 5.1 Typical Mission Profile of a Commercial Jet

Typically an aircraft must have enough fuel to loiter for between 20-30minutes at 10,000 ft

(3048m). As-well as this, they must have 30 minutes of extra cruise fuel for daytime, and 45

minutes for night or any other instrument condition flights, this can be seen reflected on the

diagram.

From this the geometry generated was run through the Sizing module of NeoCASS, what this

does is optimize any weights and the balance of the aircraft itself. From this came essential

30,000ft

10,000ft

P a g e | 25

data such as Maximum Take-off weight, Operational empty weight, zero fuel weight,

weights of individual components and their centres of gravity and the aircraft centre of

gravity with respect to percentage mean aerodynamic chord at each loading condition. This

then forms the basis for all future analysis calculations done by the software. After GUESS is

run, if the results tab on the NeoCASS GUI and ‘plot GUESS results’ are selected, a plot is

generated of Fuselage length vs Shear force, which shows the distribution of stress on the

airframe before any type of analysis is begins, an example of this is visible in the appendix

on figure 5.8.

At this point, to check that the design thus far is on track with current trends of similar

aircraft, the Empty weight fraction was taken from the values deduced by GUESS and

compared it to the sized max take-off weight. When superimposed on the diagram below

(Figure 5.2), it comes close to the trend line of twin turboprop aircraft, coming in just slightly

above it, touching the trend ranges of jet fighters. This is due to the increased structural

weight of the wing section due to the forward sweep which hasn’t been compensated for by

the use of lighter composites and updated technologies.

Figure 5.2 Sizing from a Conceptual Sketch [8]

0.60

~40,000Kg

0.586

P a g e | 26

In order to grasp basic flight capability as a passenger aircraft, the weights values from

GUESS, the assumption of a cruising speed of 0.6 Mach at an altitude 9144 m (30,000ft), the

supposition that span-wise efficiency e= 0.8 due to defining the aircraft in a state of cruise

(α=0), will help develop a basis for a very early and rough prediction of maximum range and

the endurance of the aircraft in a constant state of cruise. To calculate initial values of

range and endurance, it is necessary to derive the coefficients of Lift ( ) and Drag ( )

under the conditions of flight.

L=W for un-accelerated straight and level flight.

Using the vs mach number chart below,

Figure 5.3NACA 64A010 aerofoil properties [15]

P a g e | 27

was derived from this and assuming L= W at full fuel +10,000 kg of payload. The diagram

also gave the opportunity to size and calculate the angle of incidence of the wings to 1.29°

onto the oncoming flow to tailor cruising at Mach 0.8, at an altitude of 9000 m (~30,000 ft).

From this the overall value of across the wing came out as 0.2 at steady level un-

accelerated flight.

was derived from the ratio of Lift to Drag and the assumption that L/D for a jet aircraft at

cruise is assumed at 13.9 [8]

This produced a value of 0.014 for

Using data from GE Aviation’s website of a thrust specific fuel consumption of the CF34-8

Turbofan ( ) and values calculated for the coefficients of lift and Drag, and density

of air at 0.46 (~30,000 ft) endurance and range were estimated by the following

equations[9];

√

From this initial endurance was estimated at a flight time of 5 hours. This seems appropriate

for a regional aircraft.

Range was estimated as 4,241 Km, at constant cruise speed, so that would be less if it

included taxiing, take-off, and landing/ landing reattempt. This comes very close to the

ranges of similarly sized aircraft, such as the E-170 which has a range of 3,889 km, which is a

good indicator as although the result is very rough, the aircraft does seem to be able to

perform similar to its current day counterparts. These equations were used to the same

purpose as figure 5.2, to make sure that the design is feasible and realistic at an early stage.

P a g e | 28

22000

27000

32000

37000

42000

47000

0 5000 10000 15000 20000

Take

Off

We

igh

t (

Kg)

Payload (Kg)

Payload Trade Study

StandardAluminium

CompositeWings

Trade studies can be drawn from the data of weights as-well as the iterative use of the

range equation found on the previous page. This is useful for potential customers in order to

help visualise the basic capabilities and characteristics of the aircraft at an early stage. It

provides a means to check how heavy the aircraft will have to be under certain situations,

for example figures 5.4, 5.5 and 5.6 below show graphs of Trade studies, which display

Payload weight vs take-off weight (Payload Trade), Range vs Take-off weight (Range Trade)

and Payload weight vs Range (Payload-Range Trade).

Figure 5.4 Payload Trade Study for the XATA [8]

Some important points to note on the trade studies are that;

- Payload weight refers to the passengers and anything that is associated with them,

- Any values referring to range do not take into consideration the additional 30 minute

cruise fuel, taxiing fuel, and a 400Km diversion prior to the displayed ranges,

- In the case of the Payload Trade Study (Figure 5.4), the aircraft is assumed to have full fuel

capacity, with varying payload,

P a g e | 29

20000

25000

30000

35000

40000

45000

0 1000 2000 3000 4000 5000

Take

Off

We

igh

t (K

g)

Range (Km)

Take Off Weight- Range Trade study

StandardAluminium

CompositeWings

0

2000

4000

6000

8000

10000

12000

14000

16000

18000

20000

0 1000 2000 3000 4000 5000 6000 7000

Pay

load

(K

g)

Range (Km)

Payload-Range Trade Study

StandardAluminium

Composite

- Where-as in the Take Off Weight- Range Trade Study (Figure 5.5) the aircraft is assumed to

be fully loaded, with varying fuel capacity from a full tank, down to a reserve of 542.6 Kg of

fuel.

- The second set of results visible in both graphs corresponds to a material modification

which is referred to in the later stages of the report.

Figure 5.5 Range Trade for the XATA [8]

The ‘Take Off Weight-Range Trade’ is used as an initial indicator to show the correlation

between fuel capacity and range on a flight which is fully loaded with max payload. Figure

5.6 below uses results from the previous graphs and depicts the effects on range with

varying payload on a full fuel load.

P a g e | 30

Figure 5.6 Range-Payload Trade Study [24]

The aircraft underwent many improvements and was sized accordingly, with each version

more refined that the previous counterpart. Some GUESS module results from the different

versions of the XATA conceptual aircraft are located in the appendix along with a set from

an aircraft with a standard configuration with extra weight (the purpose of the additional

weight is referred to later in the report).

Figure 5.7 Flow Diagram of Modules from NeoCASS used in Sizing Stage

Conceptual

Sketch

Are extra

considerations

implemented?

AcBuilder

Internal Loads

Component

Properties

Geometry

Aircraft technical

weights

NeoCASS GUESS

module

Mesh model

of geometry

Aircraft Balance

No

Is it stable? Yes

Analysis stages

No

Yes

P a g e | 31

Figure 5.7 shows the flow of the first stage of the conceptual design. The section referring to

extra considerations includes criteria such as tail sizing, sweep, angle of incidence, dihedral,

fuel tank capacity, as well as choice of power plant. Some of these required iterative

methods for calculation and are mentioned in the next section.

NeoCASS allows the use of AcBuilder in conjunction with the GUESS module which

essentially encompasses the optimization of the aircraft through individual component

weights, mean aerodynamic chord, and the centres of gravity at certain aircraft loading

conditions, this in turn will help the designer optimize the aircraft in a structural manner in

terms of balance, so as to create a stable aircraft. For example from the results presented in

GUESS, the aircrafts geometry was shifted and modified several times so as to give it the

required balance. This will give a value for the whole aircraft which is then used in the rest

of the various solvers, and analyses conducted through NeoCASS.

P a g e | 32

6| Final Geometry Selection & Configuration

The only civil jet ever produced to feature forward swept wings was the Hamburger

Flugzeugbau HFB-320 Hansa Jet, produced between 1964 and 1973. Despite this, the

original study of the X-29A stated that forward swept wings had the capability of significant

reduction of drag, among other things. The conclusion of the study [14] went on to confirm

the aircraft had reached and exceeded its performance predictions. The XATA aircraft was

designed with specific considerations in mind. It has to be an original design, and be able to

potentially present a new style of flight. In addition to this it has to be a viable concept with

future trends in industry development being accounted for. For these reasons the

configuration has been justified in the following sets of paragraphs, giving reason and

description of each section of the geometry.

6.1 Wing

The XATA at present consists of NACA0012 and N64A410 aerofoils with an angle of

incidence of around 1.29°, however since this is conceptual design there is not a major in

depth analysis on this aspect yet in this development phase. It is being designed with a

forward swept wing, with a span of 30metres, and an area of 125 m2. This gives for an

aspect ratio of 7.2. The reason for the use of such a configuration is due to the lack of

commercial use in the industry today. The forward swept wing actually has many benefits to

the overall performance even though it is not commonly seen on commercial aircraft, for

example, it has a better span wise distribution than backwards swept wing, and a slightly

forward swept wing has been known to have some spin resistance. One side effect however

is the fact that for every 10° of forward sweep, the overall effect on the wing is as if there

was 1° of geometric Anhedral. Some of these include the fact that unlike conventional wing

design, it stalls at the root instead of the tip, this allows advanced warning to a pilot and can

aid in avoiding a stall in a crucial moment of flight. The design also improves max lift with

aileron control, and it reduces frontal area and drag. From a structural standpoint, the

forward sweep allows the whole wing construct along with its internal supports to be placed

further aft of the nose, clearing space at the centre of the aircraft for cargo, or any other

payload.

P a g e | 33

With the XATA, the forward swept wing would also present another issue entirely. Usually

these designs are scrapped as in order to move the wing further aft, and sweep it forward,

the wing itself must be stronger than a conventional one to avoid any excessive bending

near the tip regions (tip divergence) and avoid any tip stall during flight. In the past this

meant a weight penalty that reduced the efficiency of the overall aircraft below designs

specifications, however in the modern day world, composites are proving to be effective at

reducing the weight of all the important components in an air-vehicle. An example would be

the fuselage of the new Boeing 787 Dreamliner; this is a novel leap in design, entirely made

up of composite materials. As such this reduces a significant amount of weight from the

overall structure. If the wing and its supports were to be constructed of similar materials,

the weight penalty of the chosen configuration would essentially be reduced to a point

where it is not an issue. In a way this is also looking to the future of aircraft design.

Composite structures which are more rigid, stronger and lighter than conventional aircraft

materials are very likely to replace them in the industry in the near future. An example of

current advances in the materials industry is the recent resurfacing of the use of Carbon-

Titanium composites which have superior qualities of strength and reduced weight to

standard Carbon Fibres [27]. The wing incorporates dihedral to add stability during flight

and prevent a Dutch roll. It has a 5° geometric dihedral on the outboard and mid-sections

which is reduced to 4° due to the effects of the forward sweep, whilst there is a sharper 10°

at the inboard section. This is enough for the aircraft to experience the stability benefits of

the geometric composition, and low enough to avoid adding excess dihedral which causes

the aircraft to go into a sideslip. A point worth mentioning is the characteristics of forward

swept wings under high load factors, although the aircraft being developed is not expected

to experience such forces, the wing has to be constructed with particular emphasis on

eliminating the chances of structural divergence in flight. This is a proven concept if

constructed correctly, as demonstrated in the Grumman X-29A [10]. This experimental

aircraft employed construction techniques that help strengthen the wing in the axis where

the structure would shift by adding graphite epoxy load carrying covers on to the entire

wing, this prevented any type of structural divergence, and essentially eliminated any

chance of any tip stall.

P a g e | 34

6.2 Tail

Multiple tail designs were considered for the subject aircraft, of which the V tail was chosen

initially in contrast to a standard Tail seen on the typical configuration XATAV2. This is due

to the position of the engines, as well as the chance to reduce interference drag. The tail

was sized using the tail volume coefficient method from the length of the fuselage, this was

to compensate for the lack of control surfaces usually presented by a conventional tail to

make sure that the V tail had enough volume and made the aircraft stable and capable of

returning to straight flight after a gust or any other disturbances in the air. The only major

downsides to V tails are the complex control systems required to provide the necessary

movement on the “Ruddervators”, and the effect of the rolling movement created by the

control surfaces, creating adverse roll- yaw coupling in the opposite direction of movement.

This issue can be avoided by inverting the V tail thus creating a proverse roll-yaw moment.

However there are two problems with doing this, since the aircraft already has ventral fins

on the fuselage, the V tail remains above so as to avoid disturbing the air due to the surfaces

being too close together. The other is that part of the tail would still be behind the exhaust

from the engines, potentially damaging or melting the structure. The tail structure was sized

according to appropriate equations bearing in mind that the canard counts as part of the

surface area required for stability [23]

6.3 Fuselage

The fuselage has a slight angle towards the front, a technique drawn from aircraft with a

lifting fuselage type configuration. The principle can be applied to standard fuselages as-

well. The idea is that a little bit of ‘free lift’ and a reduction in separation drag are generated

by making the fuselage resemble an aerofoil. This is achieved by angling the whole front

section of the hull such that at cruise the fuselage is at an angle even though the aircraft

itself is flying horizontal. Most aircraft currently employ this; a particular exaggerated

example was on the Lockheed L-1011. On the XATA this is incorporated aft of the downward

nose right up till the front section of the wing, whereupon the fuselage levels out. In

addition the fuselage has ventral fins which are in place to further reduce parasitic drag due

to the angle at which the fuselage ends, as well as provide extra lateral stability at cruise.

These are particularly helpful as the fuselage angle at the aft of the aircraft to the end point

P a g e | 35

is 15°, standard practice is for an angle between 10° - 12° to avoid excess drag effects due to

separation of flow.

The Turbofans which will initially be chosen are GE aviation’s CF34-8 series engines [19].

These are widely used for small regional aircraft, such as the E170, &175 series, as-well as

the bombardier CRJ900. From the data on the official manufacturer’s website, the thrust to

weight ratio was calculated as 0.17 from the dry weight of the power plant, with a

maximum thrust of 64.5 KN per engine. This information was added into the initial AcBuilder

model which was used in GUESS. The engine configuration itself is such that the power-

plants are podded. This is so that the inlet is placed away from the fuselage with shorter

inlet ducts, as well as creating a reduction of the noise generated. These are located above

the wing near the rear of the fuselage, which gives the aircraft a lower ground clearance,

and a smooth underbelly in the event of an emergency crash landing or water ditch landing,

it is a similar configuration to the military Learjet (USAF C21-A) [8]. The landing gear height

is then reduced as a result, reducing the weight of the undercarriage. The exhaust can also

be directed above the flaps if a form of vectored thrust capability is installed, which through

the Coanda effect generate extra lift due to the downward flow so long as the nacelles

conform to the shape of the wing.

6.4 Canard

When an aircraft is being designed with a forward swept wing, aft of the centre of the

fuselage such as in one of the designs being developed in this report, there are ways to

balance out the forces so as to allow the centre of gravity to be in a suitable location for

flight. One of those ways is to incorporate a canard into the front of the aircraft. With this

design modification there are two variations. Either a lifting-type canard is installed, or a

control-type canard. The difference between the two comes in the main purpose of their

use. A lifting canard is a fixed wing which is design to just produce lift at the forward section

of the fuselage changing the distance between the Centre of gravity and the mean

aerodynamic chord. By rule of thumb, initial crude calculations of this distance show it has

to be 15% of the mean aerodynamic chord. The Control canard on the other hand is either a

fixed wing with control surfaces or a completely moveable wing at the front which can

essentially do the job of the horizontal stabilizer. This distance is more flexible, ranging

P a g e | 36

between 15-20% of the mean aerodynamic chord. Even though a control canard carries

none of the aircrafts weight, it has the added benefit of being able to control the aircraft

through a region of undisturbed flow, where-as under certain conditions a traditional tail

section may lose its effectiveness if it is caught behind the wake of the main wing, a control

canard will always be in a position to control the pitch in whatever situation.

With respects to flight safety, the canard itself can be designed to stall prior to the main

wing allowing for early warning to allow pilots to act to prevent the aircraft from

experiencing the stall. As-well as this, it is well known, a forward swept wing has dangerous

pitch-up stall qualities, however this can be combated by a control canard which is capable

of downward deflections of up to 45°, which can be used to restore the aircraft pitch to a

stable horizontal level in almost any situation. This is key to preventing any accidents due to

loss of control of the aircraft. An extra bonus to this design also factors into safety, since all

the controls of the control canard will be located nearby or underneath the flight deck, In

the event of a tear in the fuselage, or an emergency situation where vital control lines to the

rear of the aircraft are severed, the pilots will always have pitch control of the aircraft,

which is critical in saving it.

Figure 6.1, 6.2 and 6.3 below displays an image of the full exterior of the aircraft; its

geometry which has been justified and explained can be seen. From this, AcBuilder creates a

mesh which is used throughout the rest of the program, and of which, the results are

ultimately based from. This particular version is XATAv6 which only differs from the v5-5

from having altered winglets, and having all the geometric components aside from the

fuselage and the engine pods made from composite.

Figure 6.1 Side view of XATA

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Figure 6.2 Front view of XATA

Figure 6.3 Top view of XATA

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7|Modal Analysis

Modal analysis has become the standard practice to find the modes of vibration of

any machine that may be subjected to such loads and forces in its day to day use. It is

essential to assess whether or not the structure has been made so that any excessive

oscillation can be swiftly minimized and eliminated in order to prevent flutter and its

disastrous effects within the operational limits expected by both the client and designers of

the mechanical component.

NeoCASS encompasses the ability to analyse a conceptual construct by a modal means.

During the initial construct stage in AcBuilder, the input is developed as the program assigns

a beam mesh to the structure of the whole aircraft; the density of this mesh is up to the

designer’s choice. This simple beam mesh fits well with the whole conceptual design

modelling aspect of an aircraft, as in this phase, the structure tends to be represented by

very simple plates or beams, these cover the wings, fuselage, tail and canard sections

thoroughly to build up an accurate mathematical model, which represents the aircraft

through the structural aspects of design. In addition to this, there is a separate aerodynamic

panel mesh which acts as the virtual skin for other analyses types involving forces due to

flight conditions. There is also an option to change the location of the aircraft wing spars

altering the structural weight of the wing. Below is diagram displaying the structural beam

mesh created from the inputs from AcBuilder through the GUESS module of the XATA.

Figure 7.1 Beam mesh of XATA

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Figure 7.2 Set-up for Modal Analysis

Figure 7.2 above displays how to run modal analysis, you have to first set the solver input

for ‘Modal’ where upon you have the choice to run the computational steps(normalization)

as a ‘mass’, ‘max’ or a ‘3 point’ calculation. After this the SmartCAD file is assembled by

joining the solver file and the results from both the main GUESS and the GUESS COMN mass

configuration files, to enable the ‘Modal’ function. The great thing about NeoCASS is that it

is capable of analyses of vibrations with respect to up to six Degrees of freedom, in order to

best simulate loads and conditions in flight. What this means is that the mass representation

of the aircraft can move fully in a virtual space, not only along the three axis of movement

instead of only one or two directions which would be one or two degrees of freedom, but in

addition it is able to take into account the three additional movements of motion on an

aircraft (Pitch, Roll, Yaw). Figure 7.3 below shows the inputs to the solver of Modal Analysis.

NeoCASS draws up complicated matrices of motion with respect to vibrations between 0

and 999999 Hz, in the case of the current analysis being looked through in this project, the

lowest mode of vibration was 7, producing a frequency of 5.0068Hz on the deformed model

of the XATA, at this state, the wings show the preferred behaviour of transferring the

vibrations towards the fuselage. This can be seen in figure 7.4.1 at the end of this section

along with figures of vibration effects at modes 10, 15,20,25,30,35,40, all plotted with a

scale factor of 10 (as are the exported animations on the electronic version of this report).

GUESS

Module

GUESS.inc

GUESSCONM_CONF1

Modal Analysis

Modal Inputs:

Normalization, Degrees of

Freedom, Number of Modes

& Vibration range

Solver.Inc

Modal SmartCAD

NeoCASS

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Figure 7.3 Settings Display for Modal Analysis

Upon running the modal analysis the program, NeoCASS sorts the nodes on the aircraft

mesh setting the coordinate system for the problem, reads through the data created in

earlier steps through GUESS and AcBuilder, setting degrees of freedom, Eigen values and

key information such as material property database. From this it assembles the stiffness

matrix, mass matrix, and goes on to solve for the allocated frequency range, from the

minimum all the way to the maximum. This is done so each node gets a value with regards

to vibration on each section of the aircraft. In doing so, NeoCASS can be made to export a

short animation with a user defined number of frames, under more than one set of results,

of the aircraft under fluctuating amplitudes of vibration. This also displays the effect of the

vibration on the structure and geometry of each area in flight. It is possible to see the

effects of vibration on different frequencies as on the results section of the NeoCASS GUI,

the ‘selected set’ can be changed from 1 to any of the set mode vibrations calculated by the

program in the analysis stage. From here it is possible to export the animation of the

vibration in question, or select it to continue and analyse flutter at that particular condition.

Two of the animations exported by the program (at mode 7 and mode 10) are presented on

the CD alongside the electronic copy of this report, which shows shifts in the geometry

represented by a yellow version of the structure with its outline shown by red markers at

different points in the amplitude at that frequency. This is due to the vibrations acting on

the exterior at the selected mode of vibration.

Figure 7.4.1 Mode 7 Frequency 5.0068Hz

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The different vibrations effect different areas of the aircraft, this is a good feature so as to

contain any resonance through the airframe. Modes 7 to 40 show how each section vibrates

at different natural frequencies, it cycles through from the wing with the lowest value, to

the V-tail with the highest, and it starts the cycle again, at figure 7.4.6. The difference with

higher values of vibration however is exhibited on figure 7.4.5, where effects start

overlapping. At this point the V-tail isn’t the only component experiencing vibration; the

Canard at the front vibrates a small amount as-well.

It is important to mention that through the computational method of modal analysis of

natural vibration mode frequencies, the first solution presented is the convergence, which is

the lowest frequency that it occurs at. In other-words iterations and calculations always

converge on the first mode solution [11]; in this case it was mode seven. This set of modal

results is then used in further flutter analysis through NeoCASS at either a particular set

flight condition, or it can also be used to draw up a diagram of the flutter envelope of the

aircraft [5].

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8|Aero-elastic Analysis

NeoCASS also includes a means to analyse how much physical bending occurs at

different geometric locations during specific in-flight manoeuvres, it has the ability to

analyse the aero-elasticity of the structure. This shows the extent to which the components

will deform with respect to a change in surroundings or extreme flight condition. The more

the geometry deviates from its original location, the larger the resultant induced forces that

act on the aircraft will be, and this can be a major problem if the aforementioned forces act

in such a way as to inhibit control, destabilize the aircraft, or prematurely cause stall at the

wing tips. In the program this is triggered by inputting data for the solver for aero-elastic

analysis under a user specified number of flight conditions, this is then coupled like in modal

analysis with the results from GUESS, combining both the GUESS beam stick model, and the

GUESS**COMN file which contains the masses and forces on board the aircraft, which is to

say the percentage of passengers, baggage, fuel. All this is saved as a whole SmartCAD file,

which is then used to run both the ‘trim’ and the ‘Rigid VLM/DLM’ functions on the

program, these together make up the aero elastic analysis. ‘Trim’ provides the means to

statically analyse the aircraft under aero-elastic bending due to speeds and any G-load

manoeuvres. ‘Rigid VLM/DLM’ provides the means to assess normal forces acting on the

surfaces of the aerodynamic components of the aircraft, excluding the fuselage. Figure 8.3.6

at the end of this section visually shows the way the different parts of NeoCASS are set up

for either of the two analyses completed on the conceptual design. All of the results on this

section have drawn up with a scale factor of 1.0.

8.1 Static Analysis (Trim)

NeoCASS’s beam mesh of the aircraft serves it particularly well for the purpose of aero-

elastic analysis, it allows for accurate results of bending due to the distribution of the beams

through each structural component of the aircraft. It enables the bending to be calculated at

small increments along the wing, also replicating the effect on the ribs along the aerofoil.

P a g e | 45

Figure 8.1.1 Deformation plot of XATA at 1.0g [M 0.8; Zacc 1.0g; 9000m]

The inputs NeoCASS uses to base its results off, is the velocity in terms of z-acceleration

(load factor), Mach number, and altitude at which the aircraft is flying. It takes these values,

getting the forces along the wing and surfaces, combines it with the material data from

AcBuilder and then exports the results in the form of a 3D plot of the whole aircraft with the

deformed airframe superimposed on the same image. Any structural divergence is displayed

as a yellow and blue line superimposed onto the original structure. This is seen in figure

8.1.1, as a wing bends under the specified loading of 100% passengers and 100% baggage,

at a velocity of Mach 0.8, an altitude of 9000 metres and experiencing a Load factor (n) of

1.0.

Figure 8.1.2 XATA at Half Capacity at 1.0g [M 0.8; Zacc 1.0g; 9000m]

The figure above is a deformation plot of the same aircraft as in figure 8.1, flying under the

same conditions; however the structural divergence is lessened due to the aircraft only

being filled to half capacity. From this it is reasonable to presume that the percentage

capacity of the aircraft in terms of passengers and baggage is a significant factor in wing

bending, however the structural materials and flight conditions have a much heavier weight

in determining how much the wing deforms.

P a g e | 46

The aero-elastic characteristics of a forward swept wing must not be overlooked. It is critical

to use materials that maximise the divergence speed, to minimize the amount of structural

bending during the expected flight operating envelope. The most noteworthy aspect of the

wing design that differs to a normal and rear swept wing, is such that the wing divergence

speed is lower, due to the movement of centre of lift outwards across the wings. However

the possibility of aileron reversal is decreased as the speed at which this occurs increases.

This is depicted as a comparison between forward and rear swept wings in figure 8.2A

below.

Figure 8.2A The Relation of Sweep to the Divergence, Aileron Reversal & Bending Torsion

Flutter Speeds [11]

A good example to help display extreme structural divergence can be seen upon

investigation of multiple flight manoeuvres. The following plots are of the same aircraft

experiencing a number of G loads. Figure 8.1.3 below shows how much of a difference there

is between the structural divergence at a steady level cruise speed, to a pull up manoeuvre.

This shows an aircraft travelling at Mach 0.5 at an altitude of 5000 metres with its wings

experiencing load factor of 1.42.

Figure 8.1.3 XATA at 1.42g [M 0.5; Zacc 1.42g; 5000m]

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Figure 8.1.4 below displays the deformation plot of the XATA at fully loaded capacity, flying

at Mach 0.5 at an altitude of 5000m, but most importantly experiencing a load factor of 2.0,

twice the normal gravitational acceleration (19.62 m/s). The wing experiences extreme

bending due to the forces along it , however as in the previous figures, the canard seems

unaffected, this is due to the canard being of a control nature, so it bears very little of the

force and loads of the aircraft, leaving the main wings to support it all. This simulates an

extreme condition in normal flight, or a condition on landing or take-off. The plots so far

have been of the aircraft with the default material settings.

Figure 8.1.4 XATA at 2.0g [M 0.5; Zacc 2.0g; 5000m]

Figure 8.1.5 below shows a further example of wing bending under forces, this time inverted

at the same velocities as figures 8.1.3 & 8.1.4. As previously stated this material is the

default set structural material set by NeoCASS which is standard Aluminium. The material

property values are present in the table below in figure 8.1.6

Table 8.1.6 Table of Material Properties

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Figure 8.1.5 XATA at -1.0g [M 0.5; Zacc -1.0g; 5000m]

With the structure deflecting upwards or downwards at such angles, Aluminium as a

material is not appropriate as it is not rigid enough to prevent structural divergence, and the

tip stall that occurs with it on forward swept wings. For this reason another alternative

material was investigated, and used as a replacement through AcBuilder. This modified

version of the XATA aircraft was then placed through the same computational analysis with

the same forces and flight conditions as the previous figures. Below are the results of the

alterations made to the structure of the wing, and their impact on structural divergence.

Figure 8.1.6A XATA Composite Wings at Fig 8.1.1 Conditions [M 0.8; Zacc 1.0G; 9000m]

At first comparison the difference at cruise is minimal; there is only a minor difference

between the deflections of the composite wings, to the aluminium ones. This swiftly

changes however when the aircraft starts experiencing harsher flight conditions and forces.

For example the following three diagrams, figures 8.1.6B, C,D & E. These show a dramatic

reduction in structural divergence in comparison to their aluminium counterparts. The

drastic changes to the behaviour of the wings under acceleration forces are due to the

material that has replaced the aluminium.

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Figure 8.1.6B XATA Composite at Fig 8.1.3 conditions [M 0.5; Zacc 1.42g; 5000m]

Figure 8.1.6C XATA Composite at Fig 8.1.4 conditions [M 0.5; Zacc 2.0g; 5000m]

The amount of structural divergence may be significantly lower, however running rigid

VLM/DLM analysis on the normal forces along the wings displays plots which show the true

magnitude of the effects that this phenomena have on the lift.

Figure 8.1.6D Composite Deformation [M 0.5; Zacc 2.5g; 5000m]

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Figure 8.1.6E Composite XATA at Fig 8.1.5 conditions [M 0.5; Zacc -1.0g; 5000m]

Earlier in the report a reference was made to a carbon-titanium composite, however due to

lack of information the composite present is current technology that was used to model the

structure of the wing section. This is done assuming that composites will continue to get

lighter and stronger, so other slightly heavier tougher metals may be incorporated into the

wing at vital sections to further reduce the bending of the wing than already displayed, as

well as improve upon the fatigue properties of the structure. The hope is that as demand

increases for carbon composites, new more efficient and cost effective ways of machining

and manufacturing the material will be developed, reducing the cost, and making it

accessible to even small aircraft manufacturers. This is all done in a bid to increase the

divergence speed, and structural strength of the airframe, as well as keep the costs

relatively low.

This material is a type of carbon composite known as Hex Tow AS-4, which is produced by

Hexcel, a U.S. company that is one of the world’s leading material experts. Below is a table

of the material properties used through AcBuilder of the composite.

Table 8.1.7 Table of Material Properties of Carbon Composite AS-4 [13]

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-2

-1.5

-1

-0.5

0

0.5

1

1.5

2

2.5

3

3.5

4

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

Load

Fac

tor

(n)

Mach Number

Figure 8.1.8 V-n Diagram at 5000m

Pictured above is the flight envelope constructed for the XATA Composite, at an altitude of

5000m. This was done by combining the stall velocity, cruise velocity, and the Dive velocity.

This first two were either given already or easily calculated using the appropriate equations

[9]

The Dive velocity was derived from the ‘never exceed velocity’ of the aircraft at 5000m.

Upon close inspection of the deformation plots at a load factor of 1.0, all displayed the same

bending. This is simply the position of the wings during normal flight; they lift a little as they

support the weight of the fuselage. From there the deformation of the XATA at 5000m was

then plotted from Mach numbers 0.3 to 1.0. These were all analysed to see at what speed

the structure diverges. At this point it is travelling at its maximum structural velocity. This

was at 0.8 M. the wings bent downwards, which also correlates to the first mode of

vibration, (mode 7) where the wings experience downward displacement. It is at this point

where the isotropic structure will begin to experience flutter instabilities.

[25]

These critical values formed the basis of the V-n diagram and the flight envelope at this

altitude. The diagram also shows the locations at which the static aero-elastic properties of

the wings have been analysed and presented in figures 8.1.6A through to E.

Locations at

which the static

aero-elastic

deformation of

the wing was

analysed

P a g e | 52

8.2 Structural Considerations

Figure 8.2.1 Location of Spars along Wingspan

Until recent times it has been unheard of to build a wing primarily from composite

materials. The A400M from airbus is a pioneering design for this very reason [18]. It is still in

the process of getting fully certified having had it’s first production unit’s maiden flight

recently on the 3rd of March 2013. The majority of its wing components are composite,

including the spars and fuel pipes. The only parts not constructed of composites were the 24

ribs, reasons were it was not cost effective. The manner in which the wing skin was woven

from the fibre is very familiar. It resembles the wing covers originally found on NASA’s X-29

from around two decades ago. Like the airbus, the XATA spars shown in figure 8.2.1 will be

comprised of Composite. The structure of the wings of the X-29 is a guide as to how the

wings on the XATA would be made aswell. The majority of the skin will be alligned so that

the fibres run at 90° to the direction of the forces acting along the wing causing divergence.

This will be reinforced by fibres offset by +45° and -45° to the main fibre direction spaced

along the wing span. Like in the other two aircraft, it increases the materials resistance to

structural divergence. It is a testament to the Composite materials industry if a tactical

airlifter can successfully be designed to incorporate a wing made up of such materials which

supports its weight of the aircraft with a payload of more than twice that of the XATA (37

tonnes). This is with present day technology, and if it is applicable to an aircraft that can

P a g e | 53

take so much loading on the wings, then it is feasible for a lighter, smaller passenger

aircraft.

The deformation plots thus far have been made with the assumption that the material and

fibre used, has been as such that it imitates an isotropic construction, which is to say that

the material strength has been made so the fibres orientation causes it to have the same

properties if it were to experience a force from any direction.

The structural divergence, and divergence velocity can be further improved by the use of

modifying the fibre orientation of the composite on aircraft wing skin, as well as the

implementation of specialized structural supports.

The spars on the composite XATA need to be extra resistant to the forces which cause wing

deflection and twisting, as well as being sufficiently light to retain the benefits over its

aluminium counterpart. To this extent quasi-unbalanced smart spars (QUSBs) could be

implemented [20]. These are key in also reducing the risk of delamination experienced by

both the spar component and wing skin as well as excessive deformation and possible

buckling failure in the face of excessive forces under extreme conditions.

Figure 8.2.2 Quasi-Unbalanced Smart Spar Fibre Layout & Structure Composition [20]

The combined structure of the layers as well as the additional support surface internally is

shown by figure 8.2.3.

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Figure 8.2.3 Composite Wing Fibre Orientation

Figure 8.2.3 above shows the main fibre direction to counteract the divergence tendencies

of the forward sweepof the wings. Fibres alligned to 45°, 0° & 90° along the aircraft spanline

as the axis (for example visualise the spar at the rear of the wing as being the axis) as well as

the predominant layer at β° [21](between 20-60° depending on sweep angle, and aspect

ratio) in combination with the QUSBs running through the wings counteract the divergence

and the flutter instabilities that are generated from it by generating a convergent twist-bend

coupling along the whole span focusing on the wing tips and outboard leading edges. This is

only worth briefly mentioning as this type of in-depth analysis is part of the pre-liminary

stages of design, not conceptual.

The XATA relies on improvements like this to make sure its structure allows the toughest

and lightest materials of today, at tomorrows prices, which is to say that with projects like

the A400M, Tooling and manufacturing costs of creating appropriate material for such uses

will go down as lessons learnt spread from the military, through to the civil sector. With this

in mind it may not be currently cost effective, however sometimes manufacturers must take

a gamble, as with the A400M, if this were to go into development, it is probable that the

company which does it would benefit greatly from lessons which could be learnt and

implemented, even on standard swept back configurations, with respects to the materials

and technology being employed.

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8.3 Rigid DLM/VLM

This analysis function in NeoCASS provides the ability to see directly how the change of

altitude and velocity of the aircraft affects the lift generated, as-well as any effects on the

other control surfaces on the aircraft. with the trim analysis depicted in the section prior,

this form of analysis will run under the same conditions, to give a visual reference of

the difference each flight condition makes to the overall force distribution.

Figure 8.3.1 XATA Normal Force Distribution plot on Aerodynamic Surfaces

[M 0.8; Zacc 1.0g; 9000m]

Above in figure 8.3.1 is the predicted force distribution at cruise of the XATA, the

same conditions as in figure 8.1.6A. The lift force is distributed along most of the

span of the wing, with the centre making up the highest values. The scale of the

force is shown on the right of the diagram. These set of values along the colour scale

change between plot as they alter in order to best show the distribution of the normal force

across the surfaces at that altitude and depict force in Newtons. The canard is under a high

amount of force towards the front on both images due to it being a control canard, as it is

responsible for pitch of the aircraft which takes the force of the incoming airflow at an

angle.

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Below figure 8.3.2 is a plot of the aircraft at Mach 0.5, and an altitude of 5000 metres

experiencing a load factor of 2.0

Figure 8.3.2 XATA [M 0.5; Zacc 2.0g; 5000m]

The fuselage is not incorporated into the output plots done by NeoCASS as the program

assumes very little to no lift is generated, which has a negligible effect on flight operation at

this primary design phase.

A close up of the wing sections below reveals a drop in the lift along the wing. As velocity of

the aircraft decreases, the lift decreases along the outer edges of the wings. The distribution

is centred closer to the fuselage, and as expected, less lift force is generated.

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A point worth noting is that this type of analysis on NeoCASS only takes into account

velocity and altitude to present a normal lift distribution along the aerofoil at that given

flight condition, it does not factor in any manoeuvres, an example of the indifference

between results at the same altitude and speed, at a different load factor is listed in the

appendix as figure 8.3.3, which is at the same altitude and velocity as 8.3.2, however it is

experiencing a load factor of -1.0. It is plain to see that this has no effect on the results from

Rigid VLM/DLM.

Regardless of the structural divergence being omitted as a factor in the plots, the normal

force distribution allows the comparison of rough calculation of loss of lift at a certain wing

bending condition between the aluminium and composite wings. Both will generate the

same lift along their identical wings, but the magnitude of force lost due to divergence

depends upon the angle at which the wings bend upwards. This is used underneath to give

further indications as to why it is important to eliminate as much structural divergence as

possible.

Figures 8.3.4A&B in the appendix show the patches on the aero-panel mesh that the

following results are based off. This will be used to figure out the loss at each state of

structural divergence, at the same velocity and altitude.

The following represents the losses on the standard aluminium wing configuration, the wing

deformed 10° at [M 0.8; Altitude 9000 m] effective lift force is reduced to 1969N from the

total normal force of 2000N generated by the aerofoil. Where-as on the composite wing

structure, the wing deflected at 8° under the same circumstances resulting in the effective

lift force being reduced to 1980N in that patch of the aerodynamic panel mesh.

Close up of Figure 8.7A wing

Close up of Figure 8.3.2 wing Close up of Figure 8.3.1 wing

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Below is a table of corresponding deformations and lift reduction in the section with respect

to set conditions in figures 8.1.1, 8.1.3 and 8.1.4.

Wing Deformation

Total Normal Force (N)

Effective Lift (N)

Manoeuvre Force (g)

Aluminium Wing

M 0.8 Altitude 9000 10° 2000 1969 1

M 0.5 Altitude 5000 12° 1000 978 1.42

17° 1000 956 2

Composite Wing

M 0.8 Altitude 9000 8° 2000 1980 1

M 0.5 Altitude 5000 9° 1000 987 1.42

11° 1000 981 2

Table 8.3.5 Table of deformation Angles

Table 8.3.5 shows how much lift is lost by the bending of the structure on the wings. This is

only accounting for a section of the wing a little less than ¾ outboard from the fuselage. It is

not even at the outer edges, where in the case of the aluminium wing the angle of

deformation increases considerably. It aids to further highlight the importance of preventing

any divergence if possible.

As well as the increase in structural strength and resistance to wing divergence, the weight

reduction due to the use of AS4 is most visible with the output of the GUESS sizing module

of NeoCASS, Below are the two sets of summary weights of the XATA with standard

aluminium, then composite structured wings.

XATA Aluminium wings:

Operative Empty Weight (OEW) 26250.26 Kg

Max Zero Fuel Weight (MZFW) 34233.38 Kg

Maximum Take-Off Weight (MTOW) 43775.94 Kg

XATA composite wings:

Operative Empty Weight (OEW) 24850.78 Kg

Max Zero Fuel Weight (MZFW) 32833.90 Kg

Maximum Take-Off Weight (MTOW) 42376.46 Kg

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This also reduces the

ratio from figure 4.1 to 0.586 at a lower take-off weight

The resultant decrease is just with replacing only the wing structural material with

composite. This reduction also has an effect on the Trade studies, and it is shown on figures

5.4, 5.5 and 5.6. It is capable of carrying the same payload, at a lower overall take off

weight.

If the canard and V tail also have their materials changed from aluminium to composite, the

overall weight further decreases to: 9895.64 Kg structural weight producing a value of

weight saved as 668Kg. Although the full composite result was not used in the main analysis

even though it is lighter than the wing composite counterpart, its results will be kept in

mind, as the XATA would be made with a composite canard and V-tail. The extra weight

saving however will be kept to one side, in case any part of the structure needs to be

reinforced, or any extras need to be added on. The 668Kg should help to decrease how

much extra weight these changes will incur on the design values. Otherwise if no other

changes are made, the fuel tanks will be increased to allow for a larger capacity, and add

extra range capability to the regional jet.

The full set of outputs acquired from the GUESS stage of NeoCASS are listed in the appendix,

displaying individual component weights, and the distance of their centres of gravity from

the nose.

The results of this aero-elastic tailoring [14] reflect the original report of the experimental X-

29A. Conventional metals do not have sufficient strength to keep the wings from deflecting

at fatal angles, however with the advent of composites including the ever improving fatigue

resistance of current materials as well as the reduction in manufacturing costs, it seems that

composites are ideal as structural materials due to its superior strength and lightweight

properties, as opposed to in this case, standard aluminium. It is critical in this aircraft to

delay structural divergence as much as possible to minimize the chances of tip stall and

maintain a stable, flyable, economical aircraft.

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Figure 8.3.6 Set-up of Static Aero-elastic and Rigid DLM/VLM Analysis

The diagram above displays the steps taken in this section to acquire the set of results

displayed, on both the deformation plots and the normal force distribution on the wing at

specified conditions.

NeoCASS Solver.inc

Rigid DLM/VLM

SmartCAD

Static Aero-elastic

Inputs: Velocity,

Altitude & Load

Factor

GUESS

Module

GUESS.inc

GUESSCONM_CONF1

Static Aero-elastic

SmartCAD

Static Aero-

elastic Analysis

Rigid DLM/VLM

Analysis

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9|Comparison of XATA against a standard configuration

The forward swept design in this section will be compared to a conventional design

made in parallel alongside itself.

Figure 9.1 Deformation plot on a Standard Configuration of Similar Size

[M 0.8; Zacc 1.0g; 9000m]

Aero-elastic analysis is an easy comparison. Since the XATA has been tailored with stronger

structural materials than its counterpart it is lighter and stiffer than the conventional design.

Even though the results displayed were under the assumption that the material was

isotropic, The carbon fibre is still a step up from materials used on current ageing small

regional aircraft, which is to say that even though with forward swept wings the divergence

speed is lower, the structural material plays a key part in delaying this as much as possible

and avoiding any loss of lift from the bending, as well as reducing the weight penalty due to

the stiffer supports required for displacing the wing so far backwards.

Figure 9.2 Deformation Comparison between Winglet & Non Winglet Design

[M 0.8; Zacc 1.0G; 9000m]

Close up of Figure 7.7A wing

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The Geometric differences of the two aircraft are a major design factor. The purpose is to

design an aircraft that is an improvement on current day standard designs. This is reflected

in the program. NeoCASS has the ability to discern between forward and aft swept wings

and incorporate it into the final plotted results. The deformation plot figure 9.1 shows that

even with an increased structural load the deformation is very close to the XATA with

composite wings.

XATA structure weight: 10,564 Kg

Normal configuration weight: 26,991.7 Kg

The reason for the increased weight was a bigger fairing section than in the XATA, but even

with the substantial increase in weight, the deformation is minimal at cruise, like the XATA.

Additionally figure 9.2 shows a comparison of the same aircraft at same conditions, only one

without a winglet attached. From this it can be seen that the configuration with the winglet

experiences more structural divergence by a fractional amount than the wing with no

winglet.

Figure 9.3 Deformation of Standard Configuration [M 0.5; Zacc 2.0g; 5000m]

In the above figure, is the deformation at the same conditions of the composite on fig

8.1.6C. Immediate differences notable are the similarities in the very small amount of

divergence when compared to the aluminium XATA, as-well as the horizontal stabiliser

acting as a load bearing component of the aircraft. The diagram however does not show the

additional bend that more dihedral on the outboard section of the wings would incur on the

design. On the next page is the resultant normal force distribution across the wing and tail

sections.

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Figure 9.4 Rigid VLM/DLM Standard Configuration [M 0.5; 5000m]

There is a larger concentration of force on the wings compared to the XATA; however this is

due to the increase in weight. Meaning the wing needs to support a bigger load in flight.

Ultimately what the normal configuration shows in comparison to the XATA, even in its

composite form, is the drawback that the forward sweep has on the structural divergence.

However as explained before the composite has been modelled as such that the material

properties are isotropic. Although the report briefly went into a more in depth look at the

structure, for ways to avert divergence in flight, ideas were presented on how to further

minimize the bending, even though this aspect of design falls under the preliminary design

phase rather than the conceptual one.

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10|Conclusion

The purpose of this report was to generate an understanding as well as prove an

original conceptual design for an aircraft which is required to take the role of a regional

Passenger jet aircraft.

10.1 Report Findings

To generate an original 80PAX aircraft, at a conceptual stage, and improve it with the help of

computer software, was the task set at the beginning. NeoCASS has been invaluable in

providing information to analyse the structure itself in order to achieve a solid conceptual

aircraft design.

The main aim was to design and optimize the XATA aircraft. The design at this conceptual

level can be concluded as viable, although there are many things that can still be further

improved. The concept underwent several analyses and the iterative geometry sizing stages

put forward reliable computational results that suggested the aircraft proposed in this

report is feasible. It has been found that the uses of high strength- low weight composites

are an improvement over standard construction materials, in combating aero-elastic issues

as-well as reducing the overall weight, even if the material is woven to give isotropic

qualities. The unorthodox design does present some challenges, however it is possible to

overcome the structural flutter instabilities inherent with forward sweep even with

technology that exists presently, which has its roots in the experimental aircraft of the past.

Alongside the main aim there were various objectives, and milestones that were set in order

accomplish as feasible a design as possible:

At the conceptual sketch and sizing sections of the project was redone various times with

each modification increasing the airworthiness of the conceptual design. In this sense the

original aircraft did improve significantly. In addition to the improvement of the main

geometry both the Aero-panel, and beam mesh were modified to be denser and provide a

more in depth and accurate result.

The modes of vibration were successfully analysed with related results appearing in the

static aero-elastic section when creating the V-n diagram. This was achieved through the

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modal analysis in NeoCASS. The modes of vibration were helpful in identifying at which

frequencies have an effect on which parts of the geometry.

The normal Force distribution along aerofoil was analysed appropriately and used in

computing the loss of lift due to the normal force becoming a component of the upwards

vector of effective lift at each load factor.

The static aero-elastic properties of the aircraft under set flight conditions was successfully

analysed with the deformation plots of the aircraft with two sets of wing materials under

particular load factors being compared in the report, along with a third set from an

overburdened standard configuration.

From the beginning proving a conceptual design would be difficult to do with just one

person doing the task that would normally fall under a dedicated team of experts. Because

of this it was chosen that certain aspects of the conceptual stage should take priority and it

would be enough just to prove the feasibility of the original Concept idea.

10.2 Validation of results

The authenticity of the results used throughout this report to base the properties of the

XATA aircraft is a factor that needs special consideration. If the GUESS module is highly

inaccurate, then it will lead to inconsistencies, and incorrect results. Below on tables 10.2.1

& 10.2.2, are depicted the computational values of several real world aircraft weights

through GUESS. Next to them are the real world counterparts. It is apparent that NeoCASS

has the ability to accurately predict the structure weights not just at a conceptual stage, but

at the later advanced stages further along the development cycle.

Table 10.2.1 GUESS Validation of Fuselage Weights Estimation Comparison with Real World Values [22]

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Table 10.2.2 GUESS Validation of Wing Weights Estimation Comparison with Real World Values [22]

From these results it is safe to assume that those presented as a result through modal and

aero-elastic analyses are valid due to being based on authentic virtual structural

representation.

10.3 Critical Analysis

Over the course of the report, a lot of emphasis was placed on the structural aspect of the

aircraft. In retrospect it would have been better to focus more on the overall performance

values of the conceptual aircraft.

It should be stated that conceptual design is no small task. It is usually undertaken by a team

of professionals all with experience in the field of design. For this reason the main aim and

specifics of the XATA conceptual aircraft project was narrowed down from the general

overall conceptual design phase, to more specific workable project title, where the stability

along the fuselage axis was reason enough to assume satisfactory flight handling.

If this report were to be done again, an attempt would be made to run through the flutter

analysis function of NeoCASS, as well as acquire more v-n diagrams at varying flight

conditions to get a fuller picture of the flight envelope of the aircraft. It would have helped

visualise the limits of operational flight further than just at 5000m for the XATA. If there had

been enough time, an alternate means of representing orthotropic materials through similar

results and seeing the difference between outcomes would have been attempted, as even

though composites were used for the XATAv5-5, it didn’t fully encompass the improvements

that could have been achieved had it been modelled correctly.

If at all possible it would have been extremely beneficial to have produced a scale model of

the aircraft and tested its aerodynamic features in one of the wind tunnels to have

experimental values of its behaviour in fast flowing air. However that would have required a

much larger amount of time, and allocation of resources.

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11|References

___________________________________________________________________________

[1] CEASIOM module tree 2013

http://www.ceasiom.com/ceasiom-modules.html

[2] OAD aircraft design software 2012

http://www.oad.aero/

[3] RDS windows Aircraft Design 2012

http://www.aircraftdesign.com/rds.shtml

[4] XFOIL Subsonic Airfoil Development System by Mark Drela 2008

http://web.mit.edu/drela/Public/web/xfoil/

[5] NeoCASS V2.0 Tutorial R1 2011

[6] RDS software overview by D P Raymer 2012

http://www.aircraftdesign.com/RDSwin_Overview_2012.pdf RDS information pdf

[7] Vehicle Sketch Pad: open source NASA geometry Aircraft design tool http://www.openvsp.org/ Open source VSP

[8] Aircraft Design: A Conceptual Approach Fourth Edition by Daniel P. Raymer 2006

[9] Aircraft Design Lecture notes by Dr Alvin Gatto, Dr Cristinel Mares & Dr Mark Jabbal 2012

[10] Fundamentals of Aircraft Structural Analysis by Howard D Curtis 1997

[11] Aero-elasticity by Raymond L Bisplinghoff, Holt Ashley & Robert L Halfman 1996

[12] Design For Flying 2nd edition by David Thurston 1994

[13] Hexcel Website 2013

http://www.hexcel.com/

[14] NASA technical Memorandum 100413: X-29A Forward-Swept-Wing Flight Research

Program Status by Gary A Trippensee & David P Lux 1987

http://www.ceasiom.com/ceasiom-modules.html

http://www.oad.aero/

http://www.aircraftdesign.com/rds.shtml

http://web.mit.edu/drela/Public/web/xfoil/

http://www.aircraftdesign.com/RDSwin_Overview_2012.pdf

http://www.openvsp.org/

http://www.hexcel.com/

P a g e | 68

[15] NACA Research Memorandum Tests of the NACA 64-010 & 64A010 Airfoil Sections at

High Subsonic Mach Numbers by Albert D Hemenover 1949

[16] ANSYS Student/Customer support website 2013

http.//support.ANSYS.com/

[17] Lam AIR Project 2012

http://www.dlr.de/dlr/en/desktopdefault.aspx/tabid-10660/1147_read-4498/

[18] A400M wing assembly: Challenge of integrating composites by Jeff Sloan 2012

http://www.compositesworld.com/articles/a400m-wing-assembly-challenge-of-integrating-

composites

[19] GE Aviation CF34-8 Engine Data 2013

http://www.geaviation.com/engines/commercial/cf34/cf34-8.html

[20] Smart Spars: Intrinsically Smart Composite Structures by Moishe Garfinkle &

Christopher Pastore 1999

http://www.underwater.pg.gda.pl/didactics/ISPG/W%B3%F3kna/Fiber%20Architects%20Ae

rospace.htm

[21] Design and Analysis of a Composite Forward Swept Wing by Konstantin V Jensen 2009

[22] NeoCASS Next Generation Conceptual Aero Structural Sizing by L Cavagna S Ricci 2012

[23] How Big The Tail by Stan Hall 2002

http://www.eaa62.org/technotes/tail.htm

[24] Introduction to Aircraft Design by John P Fielding 1999

[25] 14 C.F.R. PART 23—AIRWORTHINESS STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES Title 14 - Aeronautics and Space; Sub part G operating limitations and information. 23.1505 Airspeed limitations.

http://law.justia.com/cfr/title14/14-1.0.1.3.10.7.html

[26] CALCULIX A Free Software Three-Dimensional Structural Finite Element Program by Guido Dhondt & Klaus Wittig 2012

http://www.calculix.de/

http://www.compositesworld.com/articles/a400m-wing-assembly-challenge-of-integrating-composites

http://www.compositesworld.com/articles/a400m-wing-assembly-challenge-of-integrating-composites

http://www.geaviation.com/engines/commercial/cf34/cf34-8.html

http://www.underwater.pg.gda.pl/didactics/ISPG/W%B3%F3kna/Fiber%20Architects%20Aerospace.htm

http://www.underwater.pg.gda.pl/didactics/ISPG/W%B3%F3kna/Fiber%20Architects%20Aerospace.htm

http://www.calculix.de/

P a g e | 69

[27] Carbon- Titanium Composite Patent application by William R Kingston 1995

http://www.patentstorm.us/patents/5733390/description.html

http://www.patentstorm.us/patents/5733390/description.html

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12|Appendix

Figure 4.3 initial Design Ideas

P a g e | 71

Figure 4.4 LamAiR Project [17]

---------------------------------------XATAv5-3 GUESS Results -----------------------------------------------

-------------------------------------------- SUMMARY -----------------------------------------------

------------- Fuselage [Kg] -------------------------

Ideal structural mass 3558.51

Total structure mass 6715.63

------------- Semi-wingbox [Kg] ---------------------

Bending material mass 196.22

Shear material mass 68.96

Predicted wingbox mass 265.17

Actual wingbox mass 815.44

------------- Wing Carrythrough [Kg] ----------------



P a g e | 72

Torsion material mass 82.56

CarryThrough mass 454.34

Final CarryThrough mass 610.72

------------- Wing [Kg] -----------------------------


Structural mass 1605.28

Primary structure mass 2192.24


Total structure including CT 3443.89

------------- Canard [Kg] ---------------------------





Total structure mass including CT 860.29

------------- Vertical tail [Kg] --------------------






Weight referred to one fin.

----------- Item Weights [Kg] ---------------------

Fuselage 6715.63

Wing 3443.89

Vertical tail 554.40

Canard 860.29

Interior 3237.22

P a g e | 73

Systems 5786.78

Nose landing gear 0.00

Main landing gear 1159.59

Engines1 3504.93

Engines2 0.00

Pilots 170.00

Crew 150.00

Passengers 7257.52

Baggage 725.60

Central tank 7842.56

Wing tank 1700.00

Fuel wing span fraction from 13.7562 to 40 %

Aux. tank 0.00

------------- Item CG [m] from nose -----------------

Fuselage 15.45

Wing 17.04

Vertical tail 29.45

Canard 6.78

Interior 14.95

Systems 14.95



Engines1 20.41

Engines2 20.41

Pilots 3.55

Crew 16.17

P a g e | 74

Wing tank 18.09

Central tank 18.30

Aux. tank 0.00

Passengers 17.44

Baggage 14.28

------------- Aircraft Weights [Kg] -----------------

Operative Empty Weight (OEW) 25613.10

Max Zero Fuel Weight (MZFW) 33596.22

Maximum Take Off Weight (MTOW) 43138.78

------------- Aircraft Balance [m] from nose --------

Longitudinal Operative Empty Weight CG 16.17

Longitudinal Max Zero Fuel Weight CG 16.40

Longitudinal Maximum Take Off Weight CG 16.82

------------- Aircraft MAC [m] ----------------------

Wing mean aerodynamic chord MAC 4.99

Wing mean aerodynamic chord apex 15.03

------------- Aircraft Balance wrt MAC --------------

Longitudinal Operative Empty Weight CG at MAC 22.99%

Longitudinal Max Zero Fuel Weight CG at MAC 27.66%

Longitudinal Maximum Take Off Weight CG at MAC 35.92%

-------------------------------------------- CONVERGENCE -------------------------------------------

- Refinement loop history:

Iter 1: Total structural mass: 11567.8 Kg. Tolerance: 8.904e-03.





P a g e | 75







- GUESS model saved in D:\NeoCASS 11-02-2013\Project F\GUESS_guess.mat file.

- GUESS summary saved in D:\NeoCASS 11-02-2013\Project F\GUESS_guess.txt file.

- SMARTCAD main file with OEW configuration saved in D:\NeoCASS 11-02-2013\Project F\GUESS.inc.

- SMARTCAD configuration file saved in D:\NeoCASS 11-02-2013\Project F\GUESSCONM_CONF1.inc file

--------------------------------XATAV5-4 GUESS results--------------------------------------------------

-------------------------------------------- SUMMARY -----------------------------------------------

------------- Fuselage [Kg] -------------------------



------------- Semi-wingbox [Kg] ---------------------








P a g e | 76




------------- Wing [Kg] -----------------------------






------------- Canard [Kg] ---------------------------






------------- Vertical tail [Kg] --------------------







------------- Item Weights [Kg] ---------------------

Fuselage 6710.41

Wing 3423.62


Canard 1413.19

Interior 3237.22

P a g e | 77

Systems 5786.78



Engines1 3504.93

Engines2 0.00

Pilots 170.00

Crew 150.00

Passengers 7257.52

Baggage 725.60


Wing tank 1700.00


Aux. tank 0.00


Fuselage 15.45

Wing 17.44

Vertical tail 29.44

Canard 4.56

Interior 14.98

Systems 14.98



Engines1 20.41

Engines2 20.41

Pilots 3.53

Crew 16.19

Wing tank 18.11

P a g e | 78

Central tank 18.37

Aux. tank 0.00

Passengers 17.44

Baggage 14.28









------------- Aircraft MAC [m] ----------------------







-------------------------------------------- CONVERGENCE -------------------------------------------


Iter 1: Total structural mass: 12362 Kg. Tolerance: 6.881e-02.




- GUESS model saved in D:\NeoCASS 11-02-2013\GUESSTEST\GUESS_guess.mat file.

P a g e | 79

- GUESS summary saved in D:\NeoCASS 11-02-2013\GUESSTEST\GUESS_guess.txt file.

- SMARTCAD main file with OEW configuration saved in D:\NeoCASS 11-02-2013\GUESSTEST\GUESS.inc.

- SMARTCAD configuration file saved in D:\NeoCASS 11-02-2013\GUESSTEST\GUESSCONM_CONF1.inc file.

Figure 5.8 GUESS Output Plot Example

_________________ XATA5-5 Composite wing GUESS Results_________________

-------------------------------------------- SUMMARY -----------------------------------------------

------------- Fuselage [Kg] -------------------------



------------- Semi-wingbox [Kg] ---------------------


P a g e | 80










------------- Wing [Kg] -----------------------------






------------- Canard [Kg] ---------------------------






------------- Vertical tail [Kg] --------------------







P a g e | 81

------------- Item Weights [Kg] ---------------------

Fuselage 6705.73

Wing 1971.05


Canard 1381.59

Interior 3237.22

Systems 5786.78



Engines1 3495.99

Engines2 0.00

Pilots 170.00

Crew 150.00

Passengers 7257.52

Baggage 725.60


Wing tank 1700.00


Aux. tank 0.00


Fuselage 15.45

Wing 17.36

Vertical tail 29.41

Canard 4.56

Interior 14.98

Systems 14.98


P a g e | 82


Engines1 20.41

Engines2 20.41

Pilots 3.53

Crew 16.19

Wing tank 18.11

Central tank 18.37

Aux. tank 0.00

Passengers 17.44

Baggage 14.28









------------- Aircraft MAC [m] ----------------------







-------------------------------------------- CONVERGENCE -------------------------------------------

P a g e | 83





- GUESS model saved in D:\NeoCASS Carbon composite\Project composite\GUESS_guess.mat file.

- GUESS summary saved in D:\NeoCASS Carbon composite\Project composite\GUESS_guess.txt file.

- SMARTCAD main file with OEW configuration saved in D:\NeoCASS Carbon composite\Project composite\GUESS.inc.

- SMARTCAD configuration file saved in D:\NeoCASS Carbon composite\Project composite\GUESSCONM_CONF1.inc file.

-----------------------XATAv2-1 Standard Configuration GUESS results-----------------

-------------------------------------------- SUMMARY -----------------------------------------------

------------- Fuselage [Kg] -------------------------



------------- Semi-wingbox [Kg] ---------------------






P a g e | 84






------------- Wing [Kg] -----------------------------






------------- Vertical tail [Kg] --------------------






------------- Horizontal tail [Kg] ------------------






------------- Item Weights [Kg] ---------------------

Fuselage 6726.59

Wing 4685.55

Horizontal tail 1382.68

P a g e | 85


Interior 3237.22

Systems 5786.78



Engines1 3743.82

Engines2 0.00

Pilots 170.00

Crew 300.00

Passengers 7257.52

Baggage 725.60


Wing tank 1700.00


Aux. tank 868.00


Fuselage 15.14

Wing 17.42

Horizontal tail 32.42

Vertical tail 31.83

Interior 14.49

Systems 14.49



Engines1 13.75

Engines2 0.00

Pilots 2.96

P a g e | 86

Crew 15.45

Wing tank 16.93

Central tank 16.57

Aux. tank 0.00

Passengers 17.04

Baggage 16.18









------------- Aircraft MAC [m] ----------------------







-------------------------------------------- CONVERGENCE -------------------------------------------




Iter 3: Total structural mass: 26544 Kg. Tolerance: 3.284e-02.

P a g e | 87



- GUESS model saved in D:\NeoCASS LOCAL\XATAv2trail\GUESS_guess.mat file.

- GUESS summary saved in D:\NeoCASS LOCAL\XATAv2trail\GUESS_guess.txt file.

- SMARTCAD main file with OEW configuration saved in D:\NeoCASS LOCAL\XATAv2trail\GUESS.inc.

- SMARTCAD configuration file saved in D:\NeoCASS LOCAL\XATAv2trail\GUESSCONM_CONF1.inc file.

Figure 8.3.4A Location of Aero-panel Chosen for Analysis at [M 0.8; Altitude 9000m]

-2000N

P a g e | 88

Figure 8.3.4B Location of Aero-panel Chosen for Analysis at [M 0.5; Altitude 5000m]

Figure 8.3.3 Rigid VLM/DLM [M 0.5; Zacc -1.0g; 5000m]

P a g e | 89

----------------------XATAV2-2 Standard Configuration GUESS results Heavier Structure--------- -------------------------------------------- SUMMARY ----------------------------------------------- ------------- Fuselage [Kg] ------------------------- Ideal structural mass 3554.10 Total structure mass 6707.29 ------------- Semi-wingbox [Kg] --------------------- Bending material mass 151.31 Shear material mass 33.06 Predicted wingbox mass 184.37 Actual wingbox mass 989.33 ------------- Wing Carrythrough [Kg] ---------------- Bending material mass 240.34 Shear material mass 19.43 Torsion material mass 33.97 CarryThrough mass 293.75 Final CarryThrough mass 394.85 ------------- Wing [Kg] ----------------------------- Ideal structural mass 1978.66 Structural mass 1947.59 Primary structure mass 2659.71 Total structure mass 3437.32 Total structure including CT 3832.17 ------------- Vertical tail [Kg] -------------------- Ideal structural mass 8652.42 Structural mass 8516.58 Primary structure mass 11630.58 Total structure mass 15030.98 Total structure mass including CT 15030.98 ------------- Horizontal tail [Kg] ------------------ Ideal structural mass 751.69 Structural mass 739.89 Primary structure mass 1010.42 Total structure mass 1305.83 Total structure mass including CT 1421.21 ------------- Item Weights [Kg] ---------------------

P a g e | 90

Fuselage 6707.29 Wing 3832.17 Horizontal tail 1421.21 Vertical tail 15030.98 Interior 3237.22 Systems 5786.78 Nose landing gear 0.00 Main landing gear 1852.53 Engines1 9635.22 Engines2 0.00 Pilots 170.00 Crew 300.00 Passengers 7257.52 Baggage 725.60 Central tank 8000.00 Wing tank 1546.41 Fuel wing span fraction from 13.7105 to 40 % Aux. tank 800.00 ------------- Item CG [m] from nose ----------------- Fuselage 15.13 Wing 13.26 Horizontal tail 30.65 Vertical tail 31.85 Interior 14.37 Systems 14.37 Nose landing gear 0.00 Main landing gear 17.50 Engines1 12.10 Engines2 0.00 Pilots 2.78 Crew 15.14 Wing tank 13.72 Central tank 14.05 Aux. tank 0.00 Passengers 16.99 Baggage 16.23 ------------- Aircraft Weights [Kg] ----------------- Operative Empty Weight (OEW) 48204.48 Max Zero Fuel Weight (MZFW) 56187.60 Maximum Take Off Weight (MTOW) 66487.60

P a g e | 91

------------- Aircraft Balance [m] from nose -------- Longitudinal Operative Empty Weight CG 20.03 Longitudinal Max Zero Fuel Weight CG 19.59 Longitudinal Maximum Take Off Weight CG 18.55 ------------- Aircraft MAC [m] ---------------------- Wing mean aerodynamic chord MAC 6.13 Wing mean aerodynamic chord apex 10.45 ------------- Aircraft Balance wrt MAC -------------- Longitudinal Operative Empty Weight CG at MAC 156.37% Longitudinal Max Zero Fuel Weight CG at MAC 149.16% Longitudinal Maximum Take Off Weight CG at MAC 132.28% -------------------------------------------- CONVERGENCE ------------------------------------------- - Refinement loop history: Iter 1: Total structural mass: 25882.5 Kg. Tolerance: 2.251e-01. Iter 2: Total structural mass: 27241 Kg. Tolerance: 4.067e-02. Iter 3: Total structural mass: 26939.2 Kg. Tolerance: 9.035e-03. Iter 4: Total structural mass: 27010 Kg. Tolerance: 2.120e-03. Iter 5: Total structural mass: 26991.7 Kg. Tolerance: 5.505e-04. - GUESS model saved in D:\NeoCASS 11-3-2013\Project Composite 1\GUESS_guess.mat file. - GUESS summary saved in D:\NeoCASS 11-3-2013\Project Composite 1\GUESS_guess.txt file. - SMARTCAD main file with OEW configuration saved in D:\NeoCASS 11-3-2013\Project Composite 1\GUESS.inc. - SMARTCAD configuration file saved in D:\NeoCASS 11-3-2013\Project Composite 1\GUESSCONM_CONF1.inc file.

--------------Lvl3XATAV6FullComposite.xml GUESS Structural Weight results---------------- ----------------------------------- CONVERGENCE------------------------------------------- - Refinement loop history: Iter 1: Total structural mass: 9899.64 Kg. Tolerance: 7.995e-003. Iter 2: Total structural mass: 9895.64 Kg. Tolerance: 4.079e-004.

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