MSc Aerospace Dissertation

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Conceptual Design of a Carbon-fibre Composite Aircraft

And

Finite Element Analysis of the Wing

Anirudh Narayan

MSc Aerospace Engineering

Student id: 0800059

Project Supervisor: Dr. Giulio Alfano

Course Director Advanced Mechanical Engineering

Brunel University, West London

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Contents Acknowledgements ....................................................................................................................... 4

Abstract ......................................................................................................................................... 4

Introduction .................................................................................................................................. 4

Literature Review .......................................................................................................................... 6

1.Carbon Fibre Composites in Aircraft: ............................................................................................... 6

2.Materials used in Carbon Fibre Composite Aircraft ........................................................................ 7

3.Aircraft Structure: ...................................................................................................................... 9

4.Adhesive Bonding Of Aircraft Structures: ...................................................................................... 14

5.Delamination/Debonding Failure: ................................................................................................. 14

6. Finite Element Analysis: ................................................................................................................ 15

7. Structural Health Monitoring System ..................................................................................... 19

Comparative Vacuum Monitoring (CVM): .......................................................................... 20

Carbon Nanotube Network: ................................................................................................ 21

Acoustic Emission Sensor: ................................................................................................... 21

8. Morphing structures ............................................................................................................... 22

9.Flutter in Composite Wings and need for Vibrational Analysis ..................................................... 24

Methodology and Results ........................................................................................................... 26

1.Aircraft Design: .............................................................................................................................. 26

2. Structural Analysis:........................................................................................................................ 45

Conclusions: ................................................................................................................................ 66

Results ............................................................................................................................................... 66

Project Management ........................................................................................................................ 67

Further Work ..................................................................................................................................... 67

Gantt chart .................................................................................................................................. 68

References: ................................................................................................................................. 69

Appendix : ................................................................................................................................... 73

Guidelines on interchanging between Abaqus and AAA .................................................................. 73

Email from Grob Aircraft ................................................................................................................... 74

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Table of Figures

Figure 1:Manufacture of CFC fuselage .......................................................................................... 7

Figure 2:Fuselage Structure .......................................................................................................... 9

Figure 3:Wing Arrangement ....................................................................................................... 11

Figure 4:Wing Structure .............................................................................................................. 11

Figure 5: Equating SHM to the Nervous system ......................................................................... 19

Figure 6:SHM flow chart ............................................................................................................. 20

Figure 7: CVM sensor .................................................................................................................. 20

Figure 8:Healing Cracks ............................................................................................................... 23

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Acknowledgements

I would like to thank my supervisor Dr. Giulio Alfano, Course Director Advanced

Mechanical Engineering, who guided me from the initial title of the thesis to its final

conceptualization and also taught me how to use Abaqus a software in which I had no

previous experience . I would also like to thank Dr. Cristnel Mares, Course Director

Aerospace Engineering who taught me about aircraft design and AAA software during

my course at Brunel University.

Abstract In this MSc thesis under the guidance of Dr. Giulio Alfano a conceptual design of a

carbon fibre composite aircraft was made and FEA was done on its wing. A procedure

to design and analyze the structural components of an aircraft in Abaqus and optimize

the design that was conceptualized in AAA was established. For this purpose the

aircraft was designed in AAA software. The shell of the aircraft was then modelled in

Aeropack and exported to Abaqus, where structural components were modelled and

assembled into the wing. Vibrational analysis was then conducted to verify the

structural integrity of the assembly and linear elastic analysis of the wing was

conducted to verify the structural integrity at steady level flight during cruise.

Introduction The current economic conditions have resulted in the transfer of billions of dollars

worth of wealth from the middle class to a few hundred elite. Thus it is necessary for

the aerospace industry to adapt to the situation and find strategies to tap into this

market of highly concentrated wealth. The practice of designing large transport jets is

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already producing losses as air travel amongst the middle class is on a decline. The

design of highly efficient business jets is one of the answers to tapping into this

wealthy market which places more emphasis on aesthetics and comfort, unlike airline

companies. Since mid-sized Carbon fibre Composite aircraft can be moulded into

desired shapes and offers better strength to weight ratio, more aesthetic and

aerodynamically efficient designs are now possible. However some components of the

aircraft can’t be moulded as whole or bolted together as this can significantly reduce

the strength and durability of the carbon fibre composite components. Therefore they

have to be adhesively bonded together. A failure mode known as debonding occurs at

these bond locations and delamination occurs inside the layers of carbon fibre. Since it

is very difficult to detect delamination/debonding, it is important to incorporate safety

features in the design phase itself. For example, the Eurofighter aircraft has a

structural health monitoring system, if it is understood where delamination has the

highest chance of occurring, piezoelectric material (material which develop a voltage

difference on being deformed) can be added. This data can be integrated into the

structural health monitoring system, if data about the rate of delamination/crack

propagation is studied for a particular aircraft. A literature review on carbon fibre

composite aircraft, delamination failure, Aircraft structures and Structural Health

Monitoring systems was done to bring all these different fields of study together in

order to have a better understanding and possible application to the carbon fibre

composite aircraft being designed in this MSc thesis.

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Literature Review The literature review was conducted on a number of topics that were relevant to the

dissertation. The reference material included published research material, books,

lecture notes and websites.

1.Carbon Fibre Composites in Aircraft: Composites are materials which compose of two or more organic and inorganic

materials. One material functions as a matrix and the other material functions as

reinforcement. The most common matrix materials are "thermosetting" materials such

as epoxy, bismaleimide, or polyimide. The reinforcing materials can be glass fibre,

boron fibre, carbon fibre, or other more exotic mixtures [19].The main driver for using

composites in aircraft is their high weight to strength ratio, which results in more fuel

efficient aircraft. Another major advantage of using carbon fibre composites in aircraft

is that they can be layered, with the fibres in each layer running in a different

direction, therefore the designer can design components which behave in a particular

way for example a component can be made to bend only in a particular direction. This

behaviour resulted in the design of forward swept wing aircraft which would not have

been possible with metals as they would bend during flight [19].Other advantages

include part reduction, complex shape manufacture, reduced scrap, improved fatigue

life, design optimisation, and generally improved corrosion resistance. The

manufacture of CFC aircraft components is generally done in three steps, first a mould

of the component is layered with composite material according to the design

specifications. Once the component is laid-up on the mould is enclosed in a flexible bag

tailored approximately to the desired shape and the assembly is enclosed usually in an

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autoclave, a pressure vessel designed to contain a gas at pressures generally up to

1.5MPa and fitted with a means of raising the internal temperature to that required to

cure the resin [20].Another method in which an expensive autoclave is not required is

the vacuum method, in this method the space between the mould and the composite

layup is evacuated of air and then it is heated to cure the resin. In the final step of

manufacturing the mould is removed and if required, cuts are made in the component

for example windows are cut into the fuselage.

Figure 1:Manufacture of CFC fuselage

[21]

The main challenges restricting the use of CFC in aircraft are material and processing

costs, damage tolerance, repair and inspection, dimensional tolerance and

conservatism associated with uncertainties about relatively new and sometimes

variable materials [20].

2.Materials used in Carbon Fibre Composite Aircraft

Grob Aircraft, a company which builds carbon fibre composite aircraft and has won a

contract from Bombardier for the design of the world’s first all composite business jet,

the Learjet 85 was contacted for information about the various fabrics and resins they

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use to manufacture their aircraft. The email from them is attached in the appendix.

According to them, their aircraft are produced by use of wet lay-up composite

materials, in this method the fabric or mat is saturated with liquid resin and the layup

is obtained by building layer upon layer till the desired thickness is reached. Glass fibre

is made by the following process, when quarry products (sand, kaolin, limestone,

colemanite) are blended together at 1,600 degree Celsius, liquid glass is formed. The

liquid is passed through micro-fine bushings and simultaneously cooled to produce

glass fibre filaments from 5-24m in diameter. The filaments are drawn together into a

strand (closely associated) or roving (loosely associated), and coated with a “size” to

provide filament cohesion and protect the glass from abrasion [32]. The resins that are

used in fibre-reinforced composites are sometimes referred to as ‘polymers’. All

polymers exhibit an important common property; they are composed of long chain-like

molecules consisting of many simple repeating units. Manmade polymers are generally

called ‘synthetic resins’ or simply ‘resins’. Polymers can be classified under two types,

‘thermoplastic’ and ‘thermosetting’, according to the effect of heat on their properties,

Thermoplastics, like metals, soften with heating and eventually melt, hardening again

with cooling. This process of crossing the softening or melting point on the

temperature scale can be repeated as often as desired without any appreciable effect

on the material properties in either state. Typical thermoplastics include nylon,

polypropylene and ABS, and these can be reinforced, although usually only with short,

chopped fibres such as glass. Thermosetting materials, or ‘thermosets’, are formed

from a chemical reaction in situ, where the resin and hardener or resin and catalyst are

mixed and then undergo a non-reversible chemical reaction to form a hard, infusible

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product [33]. Fibreglass is the most common composite material, and consists of glass

fibres embedded in a resin matrix. To make a composite structure, the composite

material, in tape or fabric form, is laid out and put in a mould under heat and pressure.

The resin matrix material flows and when the heat is removed, it solidifies. It can be

formed into various shapes. In some cases, the fibres are wound tightly to increase

strength [34].

3.Aircraft Structure:

Aircrafts are designed to play a variety of roles according to their mission

specifications, however all aircraft generally have certain primary components i.e.

Fuselage, Wing, Empennage, Power Plant and Landing Gear.

Fuselage:

Figure 2: Fuselage Structure

[3]

The Fuselage’s primary function is to carry the pilot and the payload or passengers.

Early fuselage designs had a box structure; the structural elements resembled those of

a bridge, with emphasis on using linked triangular elements. The aerodynamic shape

was completed by additional elements called formers and stringers and was then

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covered with fabric and painted, however this kind of a structure proved to be heavy

and modern aircraft use what is known as a semi-monocoque structure. In this type of

an arrangement the skin is the main load carrying member. A series of frames in the

shape of the fuselage cross sections are held in position on a rigid fixture, or jig. These

frames are then joined with lightweight longitudinal elements called stringers. These

are in turn covered with a skin of sheet aluminium, attached by riveting or by bonding

[1]. Since carbon fibre composites can be layered over a mould they a full monocoque

structure can be used. One important safety consideration to be taken in a carbon

fibre composite fuselage is that unlike its all metal counterpart, a carbon fibre

composite fuselage doesn’t provide shielding from lightning strikes. Some promising

developmental lightning protection methods that should be considered are aluminium

diverter strips, aluminium wire mesh, and aluminium flame spray [22].Some aircraft

designs such as the ‘flying wing’ design used in the Northrop YB-49 Flying Wing and

Northrop B-2 Spirit bomber do not have a separate fuselage ,the fuselage is a

thickened portion of the wing. Conversely some designs use the fuselage as the lifting

surface instead of the wing; Examples include NASA's experimental lifting body designs

and the Vought XF5U-1 Flying Flapjack.

Wing: The wing of an aircraft provides the lift necessary for the aircraft to fly. The

curved shape of the airfoil causes the air on the top surface to move faster than the

lower surface which causes a difference in pressure resulting in lift. Three systems are

used to determine how wings are attached to the aircraft fuselage depending on the

strength of a wing's internal structure. The strongest wing structure is the full

cantilever which is attached directly to the fuselage and does not have any type of

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external, stress-bearing structures. The semi cantilever usually has one, or perhaps

two, supporting wires or struts attached to each wing and the fuselage. The externally

braced wing is typical of the biplane (two wings placed one above the other) with its

struts and flying and landing wires [24].

Figure 3: Wing Arrangement

[24]

To maintain the aerodynamic shape of the wing, it must be designed to maintain its

shape even under extreme stress. The primary components of the wing which form its

structure are the skin, stringers, ribs and spars.

Figure 4: Wing Structure [3]

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Skin: The primary function of the wing skin is to form an impermeable surface for

supporting the aerodynamic pressure distribution from which the lifting capability of

the wing is derived. Forces are transmitted to the ribs and stringers from the skin

through plate membrane action [1].

Stringer: Although the thin skin is efficient for resisting shear and tensile loads, it

buckles under comparatively low compressive loads, increasing the thickness is not an

option because of the weight penalty. Therefore stringers are attached to skin and ribs

thereby dividing the skin into small panels and increasing the buckling and failing

stresses [1].

Spar: It is the main load carrying member of the wing. It resists shear and torsional

loads also supports the skin, the spar flanges enabling them to support large

compressive forces.

Ribs: They maintain and support the airfoil shape of the skin.

The other attachments on the wing which perform no structural function but are

important from the aerodynamic point of view are flaps, ailerons and winglets which

are used on some new wing designs.

Flaps: They provide the extra amount of lift required at low speeds during manoeuvres

like landing.

Ailerons: They are used to roll the aircraft to one side during turning manoeuvres.

When an aileron is deployed on one of the wings, more lift is generated on that wing

thus rolling the aircraft in the opposite direction.

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Winglets: These are found in new aircrafts as small vertical wings attached to the wing

tips. Their function is to prevent induced drag i.e. vortices formed at the tips which

cause drag.

Empennage: The empennage consists of the vertical and horizontal tails. The

structure of the tails is similar to the wing.

Vertical Tail: The vertical tail provides lateral stability and the attached rudder helps in

the yaw movement of the aircraft.

Horizontal Tail: The horizontal tail prevents the nose of the aircraft from pushing

downwards due to the lift provided by wings by providing negative lift i.e. a force

opposite to the direction of lift provided by the wings. In canard aircraft like the

EuroFighter the horizontal tail is near the nose but gives lift in the same direction as

the wings thus lifting the nose.

Propulsion: Aircraft primarily use propellers and jet engines for propulsion, some

aircraft use ramjet engines which can function only at supersonic speeds for added

propulsion. They are often used in conjunction with jet engines to achieve the right

velocity to function.

Propellers: The propeller blades are made in the shape of an airfoil, when the blades

are rotated they produce lift which in this case results in thrust for the aircraft. Most

propeller aircraft have propellers which pull the aircraft forward and are called tractor

propellers. Aircraft which use the propeller to push to push it forward are known as

pusher propellers. The engines used to rotate the propellers are piston engines.

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Jet Engines: According to Rolls-Royce the jet engines work on the principle of ‘Suck

Squeeze, Bang, and Blow’. Cold air from the atmosphere is sucked in by the fan which

is then compressed by the compressor; fuel is then mixed with the compressed air and

ignited this causes the hot high energy air to come out of the exhaust nozzle at high

velocity which gives thrust to the aircraft.

Landing Gear:

Most modern aircraft use retractable landing gear as there is a considerable increase in

drag when they are deployed. Amphibious aircraft use floatation devices instead of

wheels for landing.

4.Adhesive Bonding Of Aircraft Structures: Using adhesive bonding for joining composite parts provides many advantages such as

low cost, high strength to weight ratio, low stress concentration, fewer processing

requirements and superior fatigue resistance and environmental resistance[28].Since

welding is not possible for carbon fibre composites and riveting makes them weak,

adhesive bonding is the ideal method of joining CFC components. Adhesive bonding is

used mainly for attaching stringers to fuselage and wing skins to stiffen the structures

against buckling. It is also used to manufacture lightweight structures of metal

honeycomb cores inside metal skins for the flight control component structures [29].

5.Delamination/Debonding Failure:

Delamination is a failure mechanism in which lamina separate from each other in

laminated composites. It occurs because of poor interlaminar fracture toughness and

interlaminar stresses and results in loss in stiffness, strength, and expected life of the

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material [7]. The simulation of delamination in composites is usually divided into

delamination initiation and delamination propagation. Delamination initiation analyses

are usually based on stresses and in delamination propagation analysis, energy release

rate approach is used. The energy release rate can be evaluated using virtual crack

closure technique (VCCT) which is based on Irwin’s assumption that when a crack

extends by a small amount, the energy released in the process is equal to the work

required to close the crack to its original length [10].

A well known example in which delamination failure resulted in the loss of an aircraft

is the American Airlines Flight 587, in which the composite vertical stabilizer and

rudder separated from the fuselage of the Airbus A300-600 aircraft, rendering the

airplane uncontrollable [8]. Delamination growth can occur as a consequence of

interlaminar stresses which can arise from fuel pressure variations, stiffness mismatch

and in complex structures due to unanticipated loading such as excessive turbulence. It

is important, therefore to improve the knowledge of delamination growth both

theoretically and experimentally. It is also of interest in aircraft design to build up a

database of material toughness on advanced carbon-fibre composites, (CFC), currently

in use or considered for use in airframe structures [9].

6. Finite Element Analysis: Finite Element Analysis (FEA) was first developed in 1943 by R. Courant, who utilized

the Ritz method of numerical analysis and minimization of variation calculus to obtain

approximate solutions to vibration systems [11].Two types of analysis are generally

used in the industry 2-D and 3-D modelling. While 2-D modelling conserves simplicity

and allows the analysis to be run on a relatively normal computer, it tends to yield less

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accurate results. 3-D modelling, however, produces more accurate results while

sacrificing the ability to run on all but the fastest computers effectively [11].The finite

element method using computers took off in the 1970’s when the Boeing Company

launched a project to study stresses in the aircraft structure [1]. As the name suggests

in finite element analysis the body being studied is divided into a number of small

elements with their own physical properties such as thickness, coefficient of thermal

expansion, density, Young's modulus, shear modulus etc. The connections points

between these elements are known as nodes. The element’s geometry is dependent

on the type of problem being studied. FEA saves a lot of money and time since new

prototypes need not be made for the study, the design can be modified and studied on

the computer itself. Smaller elements or a high element density is often used to

improve the accuracy of the solution in regions where the stress gradients are high.

During a finite element analysis study a balance between computer resources available

and accuracy of results has to be achieved. The finite element analysis for this MSc

thesis was carried out in ABAQUS software, which has been adopted by major

corporations across all engineering disciplines as an integral part of their design

process. ABAQUS offers a powerful and complete solution for simple to complex linear

and nonlinear engineering problems, using the finite element method. In 2004 Abaqus

was selected by Boeing to develop and market an add-on for the software, which

incorporates the Virtual Crack Closure Technique (VCCT) proposed by Rybicki and

Kanninen [6]. Perhaps the most important function of theoretical modelling is that of

sharpening the designer's intuition; users of finite element codes should plan their

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strategy toward this end, supplementing the computer simulation with as much

closed-form and experimental analysis as possible [12].

There are five basic steps involved in developing a finite element model of a physical

system [13]:

Geometry definition

Discretization (i.e. meshing) of the geometry with a finite element mesh

Specification and assignment of material properties to finite elements

Specification of kinematic constraints

Specification of loading conditions

The different types of elements that may be used in FEA are as follows [13]:

Line Elements:

Line elements consist of 2 or more nodes that define the shape of a line. There are

three distinct types of line elements: axial line elements have only stiffness in the axial

direction, pure beam elements only have bending stiffness about one or more axes,

and combined uniaxial /beam elements have both axial and bending stiffness’s.

a) Axial line elements: Also called uniaxial or spar elements are ideal for two-

force members which are common structural components in truss-type

structures. Two-force members are mechanical components acted upon by two

equal, opposite, and collinear forces.

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b) Bending line elements: Also known as pure beam elements, these elements

are defined by 2 nodes. Each node of the beam element allows 3 degrees of

freedom. Unlike uniaxial elements, beam elements do not allow stretching or

compression stiffness. Only the component of force perpendicular to the

elements axial direction is permitted, hence these elements are used for stress

or strain calculations.

Surface or Area Elements

There are 2 types of area elements: 2 dimensional planar elements and 3

dimensional shell or plate elements:

a) Planar elements: These elements are used in 3 dimensional models which can

be simplified, since it is cost effective to follow this approach. Planar elements

permit only 2 degrees of freedom i.e. in X and Y direction. In solid mechanics

there are three different types of planar analysis problems: 2D plane strain, 2D

plane stress, and 2D axisymmetric.

b) Shell or Plate Elements: Shell elements are used to model thin 3D structures

usually acted upon by bending type loads. This element uses a different

stiffness formulation than a standard solid element allowing higher accuracy.

Solid Elements: The simplest type of solid element is the linear tetrahedral with 4

nodes and the other two sold elements are the hexahedral or brick element with 8

nodes and the quadratic hexahedral element with 20 nodes.

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7. Structural Health Monitoring System

A better understanding of how delamination occurs in aircraft structures can lead to

the development of real time structural health monitoring systems.

Figure 5: Equating SHM to the Nervous system

[14]

Structural health monitoring (SHM) can be imagined as the nervous system in the

structure of an aircraft. Different types of sensor, some embedded in the airframe,

detect cracks, corrosion, delamination and other damage and simplify their

assessment [14]. A Real time structural health monitoring system will significantly

reduce the risk of aircraft accidents due to structural failure and also reduce downtime

of aircraft in maintenance hangars thus increasing profits for airline companies.

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Figure 6:SHM flow chart

[15]

Some structural health monitoring technologies which can be embedded into the

aircraft structure are listed below:

Comparative Vacuum Monitoring (CVM):

Figure 7: CVM sensor

[16] CVM Sensor [17]

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CVM technology is based on the principle that a small volume maintained at a low

vacuum is extremely sensitive to any ingress of air. When a crack develops it forms a

leakage path between the atmospheric and vacuum galleries which can easily be

detected. Self adhesive elastomeric sensors have been developed for this purpose.

When a crack propagates from outside or from an atmospheric gallery, the seal

between the atmospheric and vacuum galleries are broken which is detected by the

CVM system [18].

Carbon Nanotube Network:

With advancements in nano technology highly efficient sensors without any significant

weight penalty can be developed, Carbon Nanotube are ideal sensors for incorporation

into the structural health monitoring system of an aircraft. Carbon Nanotubes exhibit a

behaviour called piezoresistivity i.e. change in resistance with strain. Such a sensor

could measure large strain and form a grid over a large area of a structure for

structural health monitoring (SHM) applications. Also, unlike other smart materials,

CNTs are potentially simultaneously structural, functional and smart materials because

of their load carrying capability, high thermal and electrical conductivity and sensing

properties [30].

Acoustic Emission Sensor:

Acoustic Emission (AE) is a phenomenon of all materials that when forces are applied

stress waves are propagated through the material structure, which are measurable

with suitable sensors. AE sensors are piezo-electric elements in most cases. They

transform the stress waves into a voltage, which can be analysed with a suitable

system. The frequency response of the sensors must be suitable for the frequency

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range to be detected. AE stress wave sources are associated with breaks in molecular

structure, i.e. in polymers between main-chain linkages or weak secondary linkages.

The waves have a high frequency content (100 kHz – 2 MHz) which makes this

technique insensitive to mechanical vibrations usually generated by the engines and

other aircraft parts. As a crack propagates AE is generated and so, particularly for

composite materials, the growth of flaws like delamination or cracks in the matrix or

fibres can be detected before they become dangerous [31].

8. Morphing structures

Advances in composite material research and further study of failure modes in

composite structures will lead to a new breed of aircraft which can heal themselves

and also perform multiple roles by altering the shape of their components thus

changing their aerodynamic properties and mission specifications. Proof of concept

was demonstrated by Duenas et al that a low volume-fraction (5-10%) of magnetic

particles is sufficient for enabling self healing of an approximate 150 micron x 5000

micron crack in a mendomer polymer using inductive heating. It was also

demonstrated that carbon-fibre-composites can be fabricated to morph using an

apparent shape memory effect of the same mendomer that was used to demonstrate

the self-healing [25].In their paper Duenas et al describe a self healing system which

can automatically heal its cracks without the requirement of an external sensing

system or actuator. According to them, the autonomous crack healing is accomplished

by dispersing microspheres containing a healing chemical called dicyclopentadiene

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(DCPD) and a polymerizing agent known as Grubbs’ catalyst. When a crack is initiated

in the material, the high stresses associated with it cause the nearest microspheres to

break, releasing the chemical, which after interacting with the catalyst, initiates a

chemical polymerization reaction of the DCPD that heals the crack. Similarly fibres

storing healing resin have also demonstrated by Pang et al, where when fractured, the

resin flows into the damage sites within the structure. However the research into

these carbon compounds is at a very early stage and some drawbacks still exist such as

the catalyst and the healing agent degrade at high temperatures, at low temperatures

their response time becomes slow and once the microspheres burst they can’t be

reused thus the crack can be healed only once at a particular location.

Figure 8:Healing Cracks

[27]

Many engineering ideas came from observing nature; aircraft themselves were

envisioned by observing nature. When Animal tissue is damaged blood flows out which

clots and is also sensed by the brain which sends signals to increase the body

temperature. Precisely this can be accomplished if the research done by Zako & Taka is

combined with, one of the structural health monitoring systems described above.

According to Zako & Taka a polymeric material which hosts a second solid-state

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polymer phase can migrate to the damage site under the action of heat thus healing

the crack [26].Biomemetic self healing i.e. healing mimicking nature can be

encapsulated by a table prepared by Trask et al.

[27]

9.Flutter in Composite Wings and need for Vibrational

Analysis Aeroelasticity is a phenomenon which causes great instability in aircraft, vibrations in

the wing causes flutter. Emergence of flutter compromises not only the long-term

durability of the wing structure, but also the operational safety, flight performance and

energy efficiency of the aircraft, Flutter in a wing causes its tips to rise and fall which

will change the angle of incidence , thus resulting in instability [35]. The aeroelastic

analysis of laminated composite wings is also vital to the prevention of failures induced

by oscillatory motion. The aeroelastic instabilities will change, however, when a crack

has initiated in a wing structure and must be accounted for by adjustment to the

structural and dynamic model [36].Therefore Flutter not only results in aerodynamics

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instability it also causes crack initiation and propagation in carbon fibre composite

wings. Taking all this into account it becomes apparent that vibrational analysis of

composite wings is necessary for a safe aircraft design.

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Methodology and Results The aircraft design was made using Advanced Aircraft Analysis (AAA) software and the

analysis of the wing structure was done using Abaqus version 6.9.

1.Aircraft Design:

Aircraft design has now become an iterative process; therefore no new aircraft is built

from scratch. A base aircraft is taken and improvements are made on its design

depending on the mission specifications. Therefore for the purpose of designing a

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completely carbon fibre composite aircraft the Learjet 85 was taken as the base

aircraft which is the world’s first completely CFC aircraft due to enter production in

late 2012.Data of other similar aircraft designs for the iterations to be carried out in

AAA (Advanced Aircraft Analysis Software) was found from a number of sources. The

more the number of similar aircraft, the more accurate the iterations would be

especially in the weight sizing module of AAA. Therefore a number of similar designs

were researched and the aircraft solution that came up in this MSc thesis was based on

the Learjet 85 but is a new design since all the data was not be available and was

calculated in AAA from data that was available.

An initial sketch of the Learjet design drawn in AutoCAD gave a rough idea of the

design parameters such as wing span, fuselage length etc.

(1)

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Most of the parameters changed as the design process progressed, however this rough

sketch was extremely useful to keep the final design as close as possible to the original

idea.

Weight Estimation: The weight of the aircraft determines all other aspects of the

design such as the wing span, because the lift that the wings are required to produce

will depend directly on the weight it has to lift, this in turn will affect the geometry of

the control surfaces and other components. Therefore it is very important to estimate

the weight of the aircraft depending on its mission specification. The aircraft designed

in this MSc thesis is a midsized business jet and its mission specifications are given

below:

No of passengers: 8 Max Cruising Speed: 470 Knots

Crew: 2 Specific fuel consumption (sfc) = 0.5 lb/h/lb

No of Engines: 2 (Turbo Fan Jet Engines) Range: 2500 nautical miles

After the mission specifications were finalized the flight segments were defined and

then created in AAA.

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(2)

The software contains in-built typical values of “Mff” i.e. fuel fractions required to

calculate the weight of the aircraft in each segment. However for the cruise and climb

segment the software requires manual input based on the mission specifications.

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(3)

In the cruise segment the range, cruising velocity and fuel consumption were based on

the base aircraft i.e. the Learjet 85.The lift to drag ratio was estimated from coefficient

of lift (Cl) value of the wing from the equation Lift (L)= ½ p V^2 S Cl and from the drag

coefficient (Cd),then a typical value of l/d was chosen from the Roskam Tables in AAA.

This value however changed when the aerodynamics module was completed and the

values had to be adjusted till the required range, cruising velocity and fuel

consumption was achieved along with the design point.

In order to get the second curve for the design point a regression curve was plotted by

finding similar aircraft in the same weight category as the required design.

Input Parameters

R 2500.0 nm V 470.00 kts cj 0.500lb/hr

lbL/D 12.31

Output Parameter

Mf f 0.8057

Fuel-Fraction in Cruise Segment: Flight Condition 1

Advanced Aircraft Analysis 3.12 Project 09/14/09 2:16 PM

Page 31

(4)

After the regression curve was defined the number of passengers and crew was

entered into the program along with their estimated weights.

Page 32

(5)

The design point was finally achieved after some further adjustments in the

aerodynamics module.

The program then gives the useful load as an output.

(6)

Page 33

(7)

Geometry:

Wing: The wing airfoil chosen for this aircraft design was the eppler 423 and the airfoil

coordinates were obtained from the UIUC Airfoil coordinate database [29].This

database contains coordinates for all known airfoils which can be converted to

‘afl’ format from ‘txt’ by simply renaming the file for use in AAA. The values for

the wing geometry were balanced according to the results obtained in the

aerodynamics module. Initial values were estimated then later adjusted after

the performance module required changes in the aerodynamics which in turn

affected the geometry.

Page 34

(8)

(9)

Fuselage: The fuselage is constructed by defining a series of concentric circles bounded

by a square which determines the circularity. The more the number of circles the

straighter the fuselage section will be. For ρ= A/B as shown in the figure below, A was

calculated by using the Pythagoras theorem while B was chosen depending on how

circular the section being created needed to be.

Wing Horizontal Tail Vertical Tail Canard V-Tail Angles

Ventral Fin Fuselage Landing Gear Airplane AeroPack Scale

Wing Type Selection

Straight Tapered Cranked Wing Fuel Volume Flap/Aileron/Tab Chord Length

Select Input Parameters Combination

b, cr, c

t,

c/4AR, S, ,

c/4 AR, S, cr,

c/4AR, S, ,

LE

Airplane Geometry

Advanced Aircraft Analys is 3.12 Project 09/14/09 3:43 PM

crw

= 13.40 ft

ctw

= 5.80 ft

cw = 10.10 ft

xmgcw

= 3.91 ft

ymgcw

= 13.35 ft

bw /2 = 30.75 ft

Page 35

(10)

(11)

Input Parameters

Xapexf 0.00 ft

Yapexf 0.00 ft

Zapexf 0.00 ft

if 0.00 deg

(X,Z)apexf Apex is included

(X,Y,Z)f us Fuselage Coordinate System

Nfstations 16

Output Parameter

Coordinates Defined

Fuselage Table: double click for Cross-Section Dialog

Station xfus

1

ft yfus

1

ft zfus

1

ft yfus

2

ft zfus

2

ft yfus

3

ft zfus

3

ft yfus

12

ft zfus

12

ftfus

12y

fus23

ft zfus

23

ftfus

23

1 2.6444 0.0000 0.6111 0.6111 0.0000 0.0000 -0.6111 0.6111 0.6111 0.6621 0.6111 -0.6111 0.6966

2 5.2888 0.0000 1.1000 1.1000 0.0000 0.0000 -1.1000 1.1000 1.1000 0.6552 1.1000 -1.1000 0.7103

3 8.5943 0.0000 1.6528 1.6528 0.0000 0.0000 -1.6528 1.6528 1.6528 0.6138 1.6528 -1.6528 0.7103

4 16.5275 0.0000 3.0000 3.0000 0.0000 0.0000 -3.0000 3.0000 3.0000 0.7103 3.0000 -3.0000 0.7103

5 27.7662 0.0000 3.0000 3.0000 0.0000 0.0000 -3.0000 3.0000 3.0000 0.7103 3.0000 -3.0000 0.7103

6 31.0717 0.0000 3.0000 3.0000 0.0000 0.0000 -3.0000 3.0000 3.0000 0.7103 3.0000 -3.0000 0.7103

7 39.6660 0.0000 3.0000 3.0000 0.0000 0.0000 -3.0000 3.0000 3.0000 0.7103 3.0000 -3.0000 0.7103

8 46.2770 0.0000 3.0000 3.0000 0.0000 0.0000 -3.0000 3.0000 3.0000 0.7103 3.0000 -3.0000 0.7103

9 55.5324 0.0000 2.3139 2.3139 0.0000 0.0000 -2.3139 2.3139 2.3139 0.6345 2.3139 -2.3139 0.6345

10 60.2300 0.0000 2.3139 2.3139 0.0000 0.0000 -2.3139 2.3139 2.3139 0.7103 2.3139 -2.3139 0.6345

11 64.1267 0.0000 2.3139 2.3139 0.0000 0.0000 -2.3139 2.3139 2.3139 0.7103 2.3139 -2.3139 0.6345

12 66.1267 0.0000 2.3139 2.3139 0.0000 0.0000 -2.3139 2.3139 2.3139 0.7103 2.3139 -2.3139 0.6345

13 66.8123 0.0000 2.3139 2.3139 0.0000 0.0000 -2.3139 2.3139 2.3139 0.7103 2.3139 -2.3139 0.6345

14 67.1235 0.0000 1.8000 1.8000 0.0000 0.0000 -1.8000 1.8000 1.8000 0.7103 1.8000 -1.8000 0.6345

15 67.5000 0.0000 1.2000 1.2000 0.0000 0.0000 -1.2000 1.2000 1.2000 0.7103 1.2000 -1.2000 0.6345

16 68.1000 0.0000 0.2000 0.2000 0.0000 0.0000 -0.2000 0.2000 0.2000 0.7103 0.2000 -0.2000 0.6345

Fuselage Geometry: Flight Condition 1

Advanced Aircraft Analysis 3.12 Project 09/14/09 4:12 PM

Page 36

(12)

Horizontal and Vertical Tail: The horizontal and vertical tail construction is the same as

the wing construction. However care must be taken while choosing the vertical tail

airfoil since unlike the wing, it is mandatory for the vertical tail to have a symmetrical

airfoil otherwise there will be a lift force generated only on one side causing the

aircraft to become unstable.

Fuselage Station, xf/lf %

Area

Afusi

ft2

0.00 10.00 20.00 30.00 40.00 50.00 60.00 70.00 80.00 90.00 100.00

50.00

0.00

-50.00

-100.00

-150.00

-200.00

-250.00

Z-location

zclf

ft

30.00

20.00

10.00

0.00

-10.00

Nose

Center

Tail

Afusi

zclf

Page 37

(13)

(14)

Nacelles

The nacelles which cover the jet engines were designed using the nacelle coordinate

system without defining the apex so that front end can be open. Defining the nacelles

helped in calculating the CG in Class 1 weights.

crh = 8.59 ft

cth = 5.29 ft

ch = 7.07 ft

xmgch = 0.08 ft

ymgch = 4.87 ft

bh/2 = 10.58 ft

crv = 13.22 ft

ctv = 7.93 ft

cv = 10.80 ft

xmgcv = 0.61 ft

zmgcv = 2.73 ft

bv = 5.95 ft

Page 38

(15)

Input Parameters

Xnosen 41.65 ft

Ynosen 4.60 ft

Znosen 1.0000 ft

in 0.00 deg

n 0.0 deg

n 0.0 deg

(X,Z)apexn Apex is not included

(X,Y,Z)n Nacelle Coordinate System

Nnstations 7

Output Parameter

Coordinates Defined

Nacelle 1 Table: double click for Cross-Section Dialog

Station xn

1

ft yn

1

ft zn

1

ft yn

2

ft zn

2

ft yn

3

ft zn

3

ft yn

12

ft zn

12

ftn

12y

n23

ft zn

23

ftn

23

1 0.5000 0.0000 1.6500 1.6500 0.0000 0.0000 -1.6500 1.6500 1.6500 0.7103 1.6500 -1.6500 0.7103

2 1.0000 0.0000 1.6500 1.6500 0.0000 0.0000 -1.6500 1.6500 1.6500 0.7103 1.6500 -1.6500 0.7103

3 1.5000 0.0000 1.6500 1.6500 0.0000 0.0000 -1.6500 1.6500 1.6500 0.7103 1.6500 -1.6500 0.7103

4 4.0000 0.0000 1.6500 1.6500 0.0000 0.0000 -1.6500 1.6500 1.6500 0.7103 1.6500 -1.6500 0.7103

5 5.0000 0.0000 1.6500 1.6500 0.0000 0.0000 -1.6500 1.6500 1.6500 0.7103 1.6500 -1.6500 0.7103

6 7.0000 0.0000 1.3220 1.3220 0.0000 0.0000 -1.3220 1.3200 1.3200 0.7103 1.3200 -1.3200 0.7103

7 11.2387 0.0000 1.3220 1.3220 0.0000 0.0000 -1.3220 1.3200 1.3200 0.7103 1.3200 -1.3200 0.7103

Nacelle Geometry: Flight Condition 1

Advanced Aircraft Analysis 3.12 Project 09/22/09 4:18 PM

Page 39

Loads

(15)

The VN diagram obtained shows the manoeuvrable region of the aircraft in the green

curve. Starting from the left of the green curve the top and bottom end points indicate

the value of the load factor ‘n’ at the 2 stall speeds. To the right of the curve the top

end point indicates the load factor at cruise and the bottom end point indicates the

value of the load factor at dive speed.

An aircraft experiences aerodynamic loads induced by the pilot and loads induced by

atmospheric turbulence. Pilot induced load limits are quantified in a manoeuvring V-n

diagram. Gust loads that result from sudden wind gusts are calculated by forming a

gust V-n diagram. An aircraft must be designed for both limit and ultimate loading. FAR

§25.301 defines a limit load to be the maximum load an aircraft is expected to see in

service. Ultimate loads are limit loads multiplied by a factor of safety. The factor of

safety applied to a commercial aircraft is defined to be 1.5 by FAR §25.303. The

Speed, Veas keas

Load Factor

n

g

0.00 100.00 200.00 300.00 400.00 500.00 600.00 700.00 800.00 900.00

3.00

2.00

1.00

0.00

-1.00

-2.00

VS VAeas

VBeas

VCeas

VDeas

Maneuver Diagram

Gust Diagram

Page 40

following excerpts from FAR §25.305 explain the structural requirements for the two

load categories:

§25.305 Strength and deformation.

(a) The structure must be able to support limit loads without detrimental permanent

deformation. At any load up to limit loads, the deformation may not interfere with safe

operation.

(b) The structure must be able to support ultimate loads without failure for at least 3

seconds.

The Velocity to load factor plot was plotted with values calculated from other modules

and the Veas was calculated from the formula Veas= ρ √V where V= true air speed and

ρ is density of air at that altitude.

Aerodynamics

Lift: The lift for the wing and empennage group was calculated using typical values

found in the Roskam tables which are in built in the AAA software. The values in the

aerodynamics module are adjusted according to their effect on other modules. Since

some of the values such as the range and estimated aircraft weight are constant, these

can be used as a reference for adjusting the aerodynamic module.

Page 41

(16)

Drag:

The drag segment in AAA is similar to the lift segment, however since this aircraft is a

carbon fibre composite aircraft typical values of skin friction could not be used and had

to be researched.

Input Parameters

Altitude 34000 ft

T 0.0 deg F

U1 470.00 kts

CLw cln p.of f 3.0000

Sw 590.40 ft2

ARw 6.41

w 0.43

c/4w 13.0 deg

cl

rw@M=0 6.3025 rad

-1

cl

tw@M=0 6.3025 rad

-1

orw M=0 2.0 deg

otw M=0 2.0 deg

(t/c)rw 9.03 %

(t/c)tw 9.00 %

clmax

rw1.189

clmax

tw1.134

gw 1.0 deg

Output Parameters

M1 0.812

cl

rw 10.7857 rad-1

cl

tw 10.7857 rad-1

orw 1.7 deg

otw 1.7 deg

aw 1.0 deg

Wing Lift Distribution: Flight Condition 1

Advanced Aircraft Analysis 3.12 Project 09/18/09 3:42 PM

Page 42

(17)

The above plot is an output after the drag segment is completed; it gives coefficient of

lift vs. coefficient of drag for various aircraft conditions such as take off gear up or

landing gear up etc.

(18)

Drag Coefficient, CD

Lift Coefficient

CL

0.0000 0.2500 0.5000 0.7500 1.0000 1.2500

4.0000

3.5000

3.0000

2.5000

2.0000

1.5000

1.0000

Take-off Gear Down

Take-off Gear Up

Clean

Landing Gear Down

S = 590.40 ft2

Input Parameters

WTO 22571.5 lb Sw 590.40 ft2 CD

0clean,M

0.0161BDP

clean 0.0621

Mission Profile Table

Mission Profile Wbegin

lb WF

begin

lb WF

usedlb h ft V kts C

LC

DL/D

Segment Input Input Input Input Input Output Output Output

1 Warmup 22571.5 5905.2 225.7 0 0

2 Taxi 22345.7 5679.5 111.7 0 0

3 Take-off 22234.0 5567.7 111.2 0 300

4 Climb 22122.8 5456.6 345.3 43000.0 410 0.3052 0.0219 13.92

5 Cruise 21777.5 5111.3 4230.8 43000.0 470.00 0.2081 0.0188 11.05

6 Loiter 17546.7 880.5 290.0 40000 430 0.1906 0.0184 10.36

7 Descent 17256.7 590.4 172.6 15000 300

8 Land/Taxi 17084.2 417.9 136.7 0 0

L/D from Weights: Flight Condition 1

Advanced Aircraft Analysis 3.12 Project 09/18/09 4:05 PM

Page 43

The lift to drag ratio from weights is also given as an output in the Class 1 drag

segment, these values affect the original values in the weight segment and might

change the design point completely, and therefore they were adjusted accordingly.

Performance:

The objective in the performance sizing is to get a matching plot between various

performance factors such as landing distance, maximum cruise speed and stall speed.

The values had to be adjusted in the various modules till all the curves passed through

the same point as shown below.

(19)

Aeropack: After the design was completed in AAA the model was then exported to

Aeropack software for a 3-D model of the design.

(W/S)TOlb

ft2

(T/W)TO

0.00 25.00 50.00 75.00 100.00 125.00 150.00

1.00

0.90

0.80

0.70

0.60

0.50

0.40

0.30

0.20

0.10

0.00

CLmax

L

= 0.70

CLmax

L

= 1.70

CLmax

TO

= 3.00

CLmax

TO

= 4.00

CLmax

TO

= 5.00

Stall Speed Clean

Stall Speed

Take-off Distance

TTO = 0 deg F

Maximum Cruise Speed

Landing Distance

TL = 0 deg F

WTO = 22571.46 lb

Page 44

(20)

Page 45

2. Structural Analysis:

The second phase of the project i.e. structural analysis and modelling was carried out

in Abaqus 6.9.The primary focus of analysis was on the wing of the aircraft. The wing

was cut in half for ease of calculation and modelling as the behaviour of the left and

right wings would be similar.

Modelling Structural Elements

Wing/Skin: The wing was imported from Aeropack into Abaqus as a part; it was then

cut along its mid span using the geometry repair function by removing the shell faces.

This was done as only half of the wing was required for analysis and for easy insertion

of structural components.

Page 46

I-Beam Flanges: To create I-beam flanges which followed the curvature of the wing,

datum lines were made at locations where the spars needed to be inserted.

Partitions in the wing skin were made at these datum line locations and then a copy of

the partitioned wing was made, in order to cut the unrequired faces of the skin to form

the I-Beam flanges

Page 47

I-Beam/Spars

Once the flanges were obtained, the beams had to be inserted into the flanges to

obtain the I-beams which would make up the spars for the wing skin. In order to this

the coordinates were obtained from the flanges and then sketched using the Abaqus

Sketcher.

2-D side-view sketch of the beams were made and then extruded width wise by 0.025

meters. The beams were then assembled and merged with the flanges to form the I-

Beams.

Page 48

Ribs: The ribs were then sketched and extruded the same way as the spars. More co-

ordinates were needed for the ribs as they followed the airfoil shape which was more

complex than the spars.

Page 49

Page 50

Material Assignment

Material properties now had to be assigned to all the parts, for this purpose materials

were created in Abaqus by giving the materials mechanical properties such as Density,

Young’s modulus and Poisson’s ratio .

Sections were then assigned to each face of the irregular parts separately, since in

Abaqus only sections having the same geometry can be assigned material properties

together.

Page 51

Mesh:

After the section assignments were completed, the parts were meshed individually.

The skin was assigned hexagonal elements and the solid parts were given quadratic

elements. The spacing between elements was reduced at locations where it was

believed that stresses would be higher in order to get an accurate picture while

analyzing the assembly. A number of iterations had to be performed till the right mesh

for analysis was obtained.

Page 52

Assembly:

After each part was meshed, the ribs were assembled into the spars and then the spar-

rib assembly was assembled into the wing. This was done by a series of rotations and

translations which took some time to master as these manoeuvres had to be very

accurate. The model was then ready for analysis.

Page 53

Analysis:

Vibrational Analysis:

In order to check and verify the behaviour of the model, a vibrational analysis was

conducted for the first 10 natural frequencies.

Page 54

Steps

In the ‘steps’ module of Abaqus, the nature of analysis was defined. In this case the

frequency step was chosen and the number of Eigen values was entered as 10 for the

first 10 natural frequencies.

Constraints:

Since the model was made of different parts, constrains had to be assigned so that the

model did not break apart during analysis. The flanges were constrained to the skin of

the aircraft and the beams. The ribs were also constrained to the wing skin and spars.

In Abaqus the inner or outer surface selections are determined by 2 colours i.e.

Brown= outer surface and Purple= inner surface.

Page 55

Page 56

Boundary Conditions:

To simulate the wing being attached to fuselage, the wing skin root and the ends of the

spars were encastered preventing rotation and translation in all directions at this

location.

Interaction Properties:

In order to study debonding interaction properties could be assigned instead of

boundary conditions and constraints. However due to the lack of time the interaction

properties were not defined.

Page 57

Analysis:

The initial results were very disappointing with large perturbations in the wing skin and

in one trial the I-beams broke apart and came out of the skin.

Page 58

The constraints were then adjusted and it was also found that the skin thickness had to

be increased. The material properties were also researched again and adjusted as

some errors had crept in during conversion from imperial to S.I. units.

The behaviour of the model for the first 10 natural frequencies was then successfully

obtained with no perturbations in the skin:

Mode 1:

Page 59

Mode 2:

Mode 3:

Page 60

Mode 4:

Mode 5:

Page 61

Mode 6:

Mode 7:

Page 62

Mode 8:

Mode 9:

Page 63

Mode 10:

Page 64

Linear Elastic Analysis:

A linear elastic analysis of the wing was then carried out at steady level flight during

cruise. Since at this condition lift is equal to weight and only half of the wing was being

analyzed, the spars were assigned a load of half the weight of the aircraft and the skin

was assigned an evenly distributed lift which was equal to half the aircraft weight, in

the opposite direction. The weight of the aircraft at cruise was obtained from AAA.

Page 65

The linear elastic analysis job was then submitted yielding the following result. The

undeformed shape is shown as shadow under the deformed shape.

Page 66

Conclusions:

Results

A conceptual design of an 8 seater business jet was completed. Vibrational and linear

elastic analysis on its carbon fibre composited wing was also done. A procedure has

now been established to design an aircraft in AAA and then design and analyze its

structural components in Abaqus. Any changes that are required after structural

analysis for example change in wing span; root/tip thickness etc. can easily be

reinserted into AAA in order to analyze the effect of these changes on performance

and if needed the design can be changed and revaluated in Abaqus till an optimum

design is achieved.

Page 67

Further Work This project has great scope for further work. Due to time constraints only vibrational

analysis and Linear Elastic analysis in steady level flight of the wing could be

completed. If time permitted analysis of delamination could be performed and linear

elastic analysis in other flight conditions could be performed on the wings. The

fuselage and other components of the aircraft can also be given structural attributes

and analyzed. Different materials could be assigned and changes produced could be

studied. Debonding between skin and spars can be now studied, and delamination

within the sandwich panes making the skin can be also analysed. Doing this project

helps the student understand both the structural analysis and design process of

aircraft design.

Project Management The literature review and learning how to use the Abaqus Software were done

simultaneously. The aircraft design software AAA was taught as part of the course and

therefore there was no need to learn it again for the dissertation. The aircraft design

was done in the Howell building and structural analysis was done in the Michael

Sterling building of Brunel University. There were some minor delays caused to the

project due to upgrades done in the lab, however this was accounted for as a number

of copies of the data were made.

Page 68

Gantt chart

Page 69

References: 1) Aircraft Structures for engineering students by T.H.G. Megson

2) Handbook of Adhesives and Sealants by Edward M. Petrie

3) Free online private pilot ground school (http://www.free-online-private-pilot-

ground-school.com/aircraft-structure.html )

4) Aircraft design a conceptual approach by Daniel P. Raymer

5) About.com(Composites/Plastics)

(http://composite.about.com/cs/miscellaneousnews/a/bpr_abaqus.htm)

6) Delamination in Composite Materials Dr. Richard Chung San Jose State

University

7) Materials Examination of the Vertical Stabilizer from American Airlines Flight

587 1National Transportation Safety Board, NASA Langley Research Centre,

8) A numerical and experimental investigation of delamination behaviour in the

DCB specimen,Joakim Schöna, Tonny Nyman, 2002

9) Mixed-Mode Decohesion Finite Elements for the simulation of delamination in

composite Materials P.P. Camanho ,C.G. Davila

10) Peter Widas ,Virginia Tech Material Science and Engineering

12) Finite Element Analysis, David Roylance , Department of Materials Science

and Engineering, M.I.T.

13) Biomesh (www.biomesh.org)

14) European Aeronautic Defence and Space Company (EADS)

(http://www.eads.com/800/en/madebyeads/endurance/shm.html)

Page 70

15) Structural Health Monitoring for Life Management of Aircraft, Sridhar

Krishnaswamy , North Western University ,U.S.A

16) CVM, Holger Speckman Airbus, Bremen, Germany

17) Sandia National Laboratories team leader Dennis Roach

18) Structural Health Monitoring ,Fu Ku Chang,2003

19) Composites and Advanced Materials ,U.S. Centennial of Flight

Commission(http://www.centennialofflight.gov/essay/Evolution_of_Technolog

y/composites/Tech40.htm)

20) Carbon fiber reinforced plastics in aircraft construction, C. Soutis, 2005

21) Star Telegram (www.star-telegram.com)

22) “Design considerations for composite fuselage structure of commercial

transport aircraft”, G.W. Davis ,I.F. Sakata, NASA Contractor Report CR-159296

23) Online Video Lecture Series on Computational Methods in Design and

Manufacturing by Dr. R. Krishnakumar, Department of Mechanical Engineering,

IIT Madras. (http://nptel.iitm.ac.in/)

24) Aeronautics Learning Laboratory for science technology and

research(http://www.allstar.fiu.edu/aero/flight12.htm)

25) “Multifunctional self healing and morphing composites”, T. Duenas*1, E.

Bolanos2, E. Murphy2, A. Mal3, F. Wudl2, C. Schaffner2, Y. Wang3, H. T. Hahn3,

T. K. Ooi4, A. Jha1,2007, US Army Aviation and Missile Research, Development,

and Engineering Centre

Page 71

26) “Intelligent Material Systems Using Epoxy Particles to Repair Micro cracks and

Delamination Damage in GFRP”, M. Zako and N. Takano,1999, Department of

Manufacturing Science, Osaka University

27) “Self-healing polymer composites”, R S Trask, H R Williams and I P Bond,

Department of Aerospace Engineering, University of Bristol

28) “Use of epoxy/multiwalled carbon nanotubes as adhesives to join graphite fibre

reinforced polymer composites”, Kuang-Ting Hsiao, Justin Alms,Suresh G

Advan,2000

29) UIUC Airfoil Coordinate database(http://www.ae.uiuc.edu/m-

selig/ads/coord_database.html)

30) “A carbon strain sensor for structural health monitoring”, Inpil Kang, Mark J

Schulz University of Cincinnati ,2006

31) “Intelligent Structural Health Monitoring (SHM) of Composite Aircraft

structures using Acoustic Emission sensors”,Dirk Aljets ,2005

32) Net composites (www.netcomposites.com)

33) Azom materials (www.azom.com)

34) Composites and advanced materials

(http://www.centennialofflight.gov/essay/Evolution_of_Technology/composites/Tech

40.htm)

35) “Wing instability of composite wing aircraft” Mahmood ,Fatholla,University of

Iran

Page 72

36) “Flutter prediction, suppression and control in aircraft composite

wings”,Nagarjuna, Cranfield university

Page 73

Appendix :

Guidelines on interchanging between Abaqus and AAA

1) The design should be exported to Aeropack after it is completed in AAA

2) The file format used to export the model from Aeropack to Abaqus should be

IGES

3) While exporting the model from Aerpack care must be taken about the units

used as this can effect the whole project. Abaqus uses the same units the user

has provided from the beginning and has no predefined units.

4) The whole aircraft can’t be meshed as a whole but can be meshed separately,

however this provides little benefit to structural analysis as the model imported

from Aeropack is a shell and structural components have to be added to it.

5) Ideally a single component from Aeropack should be imported and given

structural attributes.

6) To get coordinates from the wing skin in order to model structures like spars

and ribs, the skin or flanges should be given an arbitrary mesh and then the

coordinates of the nodes can easily be found using the query option in the tools

menu.

Page 74

Email from Grob Aircraft

Dear Mr. Narayan,

All Grob aircraft are produced by use of wet lay-up composite materials.

The resin system for motorplanes is L20/SL (today called ERP L20 / EPH 960).

Some gliders are produced from the Scheuffler resin system L285 / H285, H286,

H287. The very old gliders from Epicote / Laromin.

Fibre Fabrics are: Interglas 92110, 92125, 92140, 92145, 92146 and comparable

fabrics. Carbon fabrics: Mainly 98141 or ECC 452, also ECC459.

Glas Fibre Rovings: Vetrotex EC9, Carbon HTA Rovings.

We hope that helps.

Best regards

Jörg Unbehend

Joerg Unbehend

Head of Design

*** Bitte beachten Sie meine neue E-Mail Adresse ***

Phone: +49 (0) 8268 998 424

Fax: +49 (0) 8268 998 221

GROB AIRCRAFT AG

Lettenbachstrasse 9

Page 75

86874 Tussenhausen-Mattsies

Germany

www.grob-aircraft.com

Sitz Tussenhausen-Mattsies, Amtsgericht Memmingen, HRB 13686

Vertretungsberechtigter Vorstand:

Johann Heitzmann, Andre Hiebeler, Andreas Konle

Aufsichtsratsvorsitzende: Antoinette Hiebeler-Hasner

-----Ursprüngliche Nachricht-----

Von: Gehling Ulrich

Gesendet: Montag, 29. Juni 2009 09:15

An: Unbehend Joerg

Betreff: WG: Grob-Contactform Message (Product Support)

Bitte um KURZE Antwort direkt an Sender: Mr Anirudh Narayan

Danke

-----Ursprüngliche Nachricht-----

Von: Vodermeier Rudolf

Gesendet: Montag, 29. Juni 2009 07:59

An: .GF

Betreff: WG: Grob-Contactform Message (Product Support)

Page 76

-----Ursprüngliche Nachricht-----

Von: [email protected] [mailto:[email protected]]

Gesendet: Sonntag, 28. Juni 2009 17:01

An: -EVL-Productsupport

Betreff: Grob-Contactform Message (Product Support)