1

DESIGN AND FABRICATION OF

THORP T-211 WING

A PROJECT REPORT

Submitted by

ARUN CELESTIN.P (611311101002)

ASWIN SHANKAR.P.S (611311101003)

SUDARSAN.N (611311101020)

VIGNESHWARAN.S (611311101022)

in partial fulfillment for the award of the degree

of

BACHELOR OF ENGINEERING

IN

AERONAUTICAL ENGINEERING

MAHENDRA ENGINEERING COLLEGE

NAMAKKAL-637 503

ANNA UNIVERSITY :CHENNAI – 600 025

APRIL 2015

ANNA UNIVERSITY :CHENNAI- 600 025

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BONAFIDE CERTIFICATE

Certified that this project report “FABRICATION AND INSTALLATION

OF THORP-T211 WING” is the bonafide work of,

ARUN CELESTIN.P (611311101002)

ASWIN SHANKAR.P.S (611311101003)

SUDARSAN.N (611311101020)

VIGNESHWARAN.S (611311101022)

Who carried out the project work under my supervision.

SIGNATURE SIGNATURE

Dr.C.DHAVAMANI,M.E., Ph.D., Mr.K.BALAKRISHNAN, M.Tech

M.I.S.T.E., M.I.E. SUPERVISOR,

HEAD OF THE DEPARTMENT, ASSISTANT PROFESSOR,

Department of Aeronautical Engineering, Department of aeronautical

Engineering

Mahendra Engineering College, Mahendra Engineering College,

Namakkal-637 503. Namakkal-637 503.

Submitted for the University examinations held on………………. at ………….

INTERNAL EXAMINER EXTERNAL EXAMINER

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ACKNOWLEDGEMENT

Behind every achievement lies an unfathomable sea of gratitude to those

who actuated it, without them it would never have into existence .To them we

lay the word of gratitude imprinted within us.

We express our sincere thanks and gratitude to our honorable Chairman,

Shri.M.G.BHARATKUMAR, M.A., B.Ed., M.I.S.T.E., who has provided

excellent facilities for us to complete our project as successful one.

We wish to express our sincere thanks to our respected Principal

Dr.M.MADHESWARAN, B.E., M.E., Ph.D., M.B.A., (Ph.D).,

M.I.S.T.E.,for all the blessing and help provided during the period of project

work..

We wish to express our sincere thanks to Dr.V.SHANMUGAM, M.E.,

Ph.D., M.B.A.,Dean School of Mechanical Science for the continuous help

over the period of project work

We wish to express our sincere thanks to Dr.C.DHAVAMANI, M.E.,

Ph.D., M.I.S.T.E., M.I.E.,Head of the Department of Aeronautical

Engineering for the continuous help over the period of project work.

With sincere gratitude respect and pride, we express our thanks to

MR.U.V.RAO, SENIOR ENGINEER, AVIONICS DEPARTMENT,

TAAL, HOSUR for his excellent guidance and encouragement throughout the

successful completion of the project.

We are indebted to our guide Mr.K.BALAKRISHNAN, M.Tech.,

Assistant Professor, Department of Aeronautical Engineering for his constant

help and creative ideas over the period of Project work.

We would like to extend our warmest thanks to all our Staff members and

Lab Technicians for helping us in this venture.

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ABSTRACT

Our main aim is to implement the composite materials to the

thorp T-211 wing by fabrication of the carbon fiber and aramid fiber

by the process of lapping of the sandwich panels.

In the initial stage of manufacturing of the thorp T-211 wing

was done with the metals like aluminum. Aluminum has more

strength, corrosion resistant and also less weight. So, aluminum has

used in all aircraft parts.

But, now the technology has been increased in the material

science. So, there is a new material has introduced in the field of

materials. That is composite material these materials, Light weight,

Resistance to corrosion, High resistance to fatigue damage, reduced

machining Tapered sections and compound contours easily

accomplished, Can orientate fibers in direction of strength/stiffness

needed.

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LIST OF TABLES

6.1 Wing 1: with propeller & landing gear 84

6.2 Wing 1: without landing gear 85

6.3 Wing 2: without landing gear 86

6.4 Carbon fabric construction data-Hexcel fibers 88

6.5 PAN Carbon fibers data 89

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LIST OF FIGURES

1.1 Structural wing design 3

1.2 Aerofoil shape 5

1.3 Aerofoil 9

1.4 Aerodynamic forces 10

1.5 Lift direction 12

1.6 Lift induced drag 14

1.7 Various wing tip devices 20

1.8 Types of aircraft wings 23

1.9 Parts of wings 24

1.10 Types of fibers 32

1.11 Cross sectional view of composite blade 35

1.12 Explore view of T-211 38

1.13 Thorp T-211 aircraft 39

3.1 Tools for fabrication 54

3.2 Fabrication Process 55

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3.3 Composite layers 56

4.1 Open circuit wind tunnel 73

5.1 Rear beam assemby 699L(699R) 78

5.2 AFT RIB ASSY 698-7L(698-7R) 79

5.3 Drilling holes 81

5.4 Rear spar attachment 82

6.1 Cross sectional view of wing 90

6.2 Ribs and Spar 91

]

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LIST OF ABBREVATIONS

DOC Direct Operating Cost

ATL Automated Tape Laying

FRP Fiber Reinforced Plastic

AFP Automated Fiber Placement

UHMPE Ultra High modules Polyethylene

LC Laminate Composite

HLC Hybrid Laminate Composite

VARTM Vacuum Assisted resin Transfer Molding

SRPP Self Reinforced Poly Propylene

HLU Hand Lay Up

RTM Resin Transfer Molding

RIP Resin infusion Process

TSHB Tensional Split Hopkinson Bar

HAL Hindustan Aeronautics Limited

ISRO Indian space Research Organization

NAL National Aerospace Limited

ADE Aeronautical Development Establishment

E-beam Electron beam

RIM Reaction Injection Molding

CFRP Carbon Fiber Reinforced Plastic

NACA National Advisory committee for

Aeronautics

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CHAPTER-1

INTRODUCTION

1.1 WING AND WINGDESIGN

Aircraft preliminary design – the second step in design process – was

introduced. Three parameters were determined during preliminary design,

namely: aircraft maximum takeoff weight (WTO); engine power (P), or engine

thrust (T); and wing reference area (Sref). The third step in the design process is

the detail design. During detail design, major aircraft components such as wing,

fuselage, horizontal tail, vertical tail, propulsion system, landing gear and

control surfaces are designed one-by-one. Each aircraft component is designed

as an individual entity at this step, but in later design steps, they are integrated

as one system – aircraft- and their interactions are considered.

This chapter focuses on the detail design of the wing. The wing may be

considered as the most important component of an aircraft, since a fixed-wing

aircraft is not able to fly without it. Since the wing geometry and its features are

influencing all other aircraft components, we begin the detail design process by

wing design. The primary function of the wing is to generate sufficient lift force

or simply lift (L). However, the wing has two other productions, namely drag

force or drag (D) and nose-down pitching moment (M). While a wing designer

is looking to maximize the lift, the other two (drag and pitching moment) must

be minimized. In fact, a wing is considered as a lifting surface that lift is

produced due to the pressure difference between lower and upper surfaces.

Aerodynamics textbooks are a good source to consult for information about

10

mathematical techniques for calculating the pressure distribution over the wing

and for determining the flow variables.

Basically, the principles and methodologies of ―systems engineering‖ are

followed in the wing design process.

Limiting factors in the wing design approach originate from design

requirements such as performance requirements, stability and control

requirements, producibility requirements, operational requirements, cost, and

flight safety. Major performance requirements include stall speed, maximum

speed, takeoff run, range and endurance. Primary stability and control

requirements include lateral-directional static stability, lateral-directional

dynamic stability, and aircraft controllability during probable wing stall.

During the wing design process, eighteen parameters must be determined. They

are as follows:

1. Wing reference (or plan form) area (SW or Serf or S)

2. Number of the wings

3. Vertical position relative to the fuselage (high, mid, or low wing)

4. Horizontal position relative to the fuselage

5. Cross section (or aerofoil)

6. Aspect ratio (AR)

7. Taper ratio

8. Tip chord

9. Root chord

10. Mean Aerodynamic Chord (MAC or C)

11. Span

12. Twist angle (or washout)

13. Sweep angle

14. Dihedral angle

15. Incidence (iw) (or setting angle)

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16. High lifting devices such as flap

17. Aileron

Fig1.1. Structural wing design

Thus, the wing design begins with one known variable (S), and

considering all design requirements, the other 17 wing parameters are obtained.

The wing must produce sufficient lift while generate minimum drag, and

minimum pitching moment. These design goals must be collectively satisfied

throughout all flight operations and missions. There are other wing parameters

that could be added to this list such as wing tip, winglet, engine installation,

fairing, vortex generator, and wing structural considerations.

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1.2 Wing design - selection of wing parameters

1.2.1 Introduction

In the context of wing design the following aspects need consideration.

1. Wing area (S) : This is calculated from the wing loading and gross weight

which have been already decided i.e. S= W / (W / S)

2. Location of the wing on fuselage : High-, low- or mid-wing

3. Aerofoil : Thickness ratio, camber and shape

4. Sweep (Λ) : Whether swept forward, swept backward, angle of sweep,

5. cranked wing, variable sweep.

6. Aspect ratio (A) : High or low, winglets

7. Taper ratio (λ) : Straight taper or variable taper.

8. Twist (ε) : Amount and distribution

9. Wing incidence or setting (iw)

10. High lift devices : Type of flaps and slats; values of CLmax, Sflap/S

11. Ailerons and spoilers : Values of Saileron/S; Sspoiler/S

12. Leading edge strakes if any.

13. Dihedral angle ( Γ ).

14. Other aspects : Variable camber, planform tailoring, area ruling, braced

15. wing, aerodynamic coupling (intentionally adding a coupling lifting

surface like canard).

The above parameters are dealt with in the following order.

i) Airfoil selection

ii) Aspect ratio

iii) Sweep

iv) Taper ratio

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v) Twist

vi) Incidence

vii) Dihedral

viii) Vertical location

ix) Wing tips

x) Other aspects

Fig1.2. Aerofoil shape

1.3 Airfoil selection

Large airplane companies like Boeing and Airbus may design their own

airfoils However, during the preliminary design stage, the usual practical is to

choose the airfoil from the large number of airfoils whose geometric and

aerodynamic characteristics are available in the aeronautical literature. To

enable such a selection it is helpful to know the aerodynamic and geometrical

characteristics of airfoils and their nomenclature. These topics are covered in

the next three subsections.

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Following six terms are essential in determining the shape of a typical

airfoil:

(1) The leading edge

(2) The trailing edge

(3) The chord line

(4) The camber line (or mean line)

(5) The upper surface

(6) The lower surface

For Thorpedo T211 aircraft,

Wing Span, b = 7.62m

Wing Area, S = 9.75m2

S = b x Croot

Solving,

Croot = 1.28m

15

Aspect Ratio, A.R =

= 5.953m

Wing loading,

= 576.607N/m

2

Stall Velocity, Vstall = 39 knots

=20.063m/s

CLmax =

Density at sea level = 1.225kg/m3

Hence,

CLmax = 2.338746

Reynold’s number:

Reynolds's number, Re =

µ0 = 1.667x10-5

Ns/m2

ρ0 = 1.225kg/m3

Re = 3.8318023 x 106

Hence it is transient flow.

When a retractable landing gear is installed it needs provisions to be stored

within airplane body. In Thorpedo T211 aircraft fuel is stored within the

fuselage. Hence the wings are hollow. This space can be utilized for storing the

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under carriage once it’s retracted. But, the existing airfoil NACA 1410 is a thin

airfoil and cannot accommodate it.

So a new airfoil which is thicker and has more CLmax, in order to counter the

extra weight of landing gear mechanism, is selected.

NACA 4415 airfoil meets all these requirements.

1.4 National Advisory Committee for Aeronautics (NACA)

Mathematical theory has not, as yet, been applied to the discontinuous

motion past a cambered surface. For this reason, we are able to design aerofoil

only by consideration of those forms which have been successful, by applying

general rules learned by experience, and by then testing the airfoils in a reliable

wind tunnel.

NACA 4415 is defined as a shape that has a maximum camber of 4 percent of

the chord (first digit); the maximum camber occurs at a position of 0.4 chord

from the leading edge (the second digit), and the maximum thickness is

15percent (the last two digits).

NACA 1410 is defined as a shape that has a maximum camber of 1percent of

the chord (first digit); the maximum camber occurs at a position of 0.4 chord

-5

0

5

10

15

0 20 40 60 80 100 120

NACA 4415

17

from the leading edge (the second digit), and the maximum thickness is 10

percent (the last two digits).

1.5 Geometrical characteristics of airfoils

In this procedure, the camber line or the mean line is the basic line for

definition of the aerofoil shape . The line joining the extremities of the camber

line is the chord. The leading and trailing edges are defined as the forward and

rearward extremities, respectively, of the mean line. Various camber line shapes

have been suggested and they characterize various families of airfoils.. Various

thickness distributions have been suggested and they characterize different

families of airfoils. The maximum ordinate of the thickness distribution as

fraction of chord (ytmax/c) and its location as fraction of chord (xytmax/c) are

the important parameters of the thickness distribution.

1.6 Airfoil shape and ordinates

The aerofoil shape (Fig.1.3) is obtained by combining the camber line

and the thickness distribution in the following manner.

a) Draw the camber line shape and draw lines perpendicular to it at

various locations along the chord.

b) Lay off the thickness distribution along the lines drawn perpendicular

to the mean line.

c) The coordinates of the upper surface (xu, yu) and lower surface (xl, yl)

of the airfoil are given by the four equations presented in fig1.3.

Fig. 1.3 Aero foil

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1.7 AERODYNAMIC FORCES

Aerodynamic force is exerted on a body by the air (or some other gas) in which

the body is immersed, and is due to the relative motion between the body and

the gas. Aerodynamic force arises from two causes the normal force due to the

pressure on the surface of the body the shear force due to the viscosity of the

gas, also known as friction. Pressure acts locally, normal to the surface, and

shear force acts locally, parallel to the surface. The net aerodynamic force over

the body is due to the pressure and shear forces integrated over the total exposed

area of the body. When an aerofoil (or a wing) is moving relative to the air it

generates an aerodynamic force, in a rearward direction at an angle with the

direction of relative motion. This aerodynamic force is commonly resolved into

two components. Drag is the force component parallel to the direction of

relative motion, lift is the force component perpendicular to the direction of

relative motion. In addition to these two forces, the body may experience an

aerodynamic moment also, the value of which depends on the point chosen for

calculation. The force created by a propeller or a jet engine is called thrust and it

is also an aerodynamic force (since it also acts on the surrounding air). The

aerodynamic force on a powered airplane is commonly represented by three

vectors: thrust, lift and drag. The other force acting on an aircraft during flight is

its weight. Weight is a body force and is not an aerodynamic force.

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1.7.1 Lift (Force)

A fluid flowing past the surface of a body exerts a force on it. Lift is the

component of this force that is perpendicular to the oncoming flow direction. It

contrasts with the drag force, which is the component of the surface force

parallel to the flow direction. If the fluid is air, the force is called an

aerodynamic force. In water, it is called a hydrodynamic force. Lift is most

commonly associated with the wing of a fixed-wing aircraft, although lift is also

generated by propellers, kites, helicopter rotors, rudders, sails and keels on

sailboats, hydrofoils, wings on auto racing cars, wind turbines, and other

streamlined objects. Lift is also exploited in the animal world, and even in the

plant world by the seeds of certain trees. While the common meaning of the

word "lift" assumes that lift opposes weight, lift in the technical sense used in

this article can be in any direction with respect to gravity, since it is defined

with respect to the direction of flow rather than to the direction of gravity.

When an aircraft is flying straight and level (cruise) most of the lift opposes

gravity. However, when an aircraft is climbing, descending, or banking in a turn

the lift is tilted with respect to the vertical. Lift may also be entirely downwards

in some aerobatic manoeuvres, or on the wing on a racing car. In this last case,

the term downforce is often used. Lift may also be largely horizontal, for

instance on a sail on a sailboat.

Fig. 1.4.Aerodynamic force

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Aerodynamic lift is distinguished from other kinds of lift in fluids.

Aerodynamic lift requires relative motion of the fluid which distinguishes it

from aerostatic lift or buoyancy lift as used by balloons, blimps, and dirigibles.

Aerodynamic lift usually refers to situations in which the body is completely

immersed in the fluid, and is thus distinguished from planing lift as used by

motorboats, surfboards, and water-skis, in which only a lower portion of the

body is immersed in the lifting fluid flow.

An airfoil generates lift by exerting a downward force on the air as it flows past.

According to Newton's third law, the air must exert an equal and opposite

(upward) force on the airfoi , which is the lift. The air flow changes direction as

it passes the airfoil following a path that is curved downward, and the overall

result is that a reaction force is generated opposite to the directional

change.[17][18]

In the case of an airplane wing, the wing exerts a downward force

on the air and the air exerts an upward force on the wing. Some of the air

passing the airfoil has downward momentum imparted to it at a rate equal to the

lift (see "Momentum balance in lifting flows" for details). This is consistent

with Newton's second law of motion which states that the rate of change of

momentum is equal to the resultant force.

Fig. 1.5. Lift direction

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1.7.2 Drag (Force)

In fluid dynamics, drag (sometimes called air resistance or fluid resistance)

refers to forces that oppose the relative motion of an object through a fluid (a

liquid or gas). Drag forces act in a direction opposite to the oncoming flow

velocity. Unlike other resistive forces such as dry friction, which is nearly

independent of velocity, drag forces depend on velocity.

For a solid object moving through a fluid, the drag is the component of the net

aerodynamic or hydrodynamic force acting opposite to the direction of the

movement. The component perpendicular to this direction is considered lift.

Therefore drag opposes the motion of the object, and in a powered vehicle it is

overcome by thrust. In aerodynamics, and depending on the situation,

atmospheric drag can be regarded as an inefficiency requiring expense of

additional energy during launch of the space object or as a bonus simplifying

return from orbit.

VARIOUS TYPES OF DRAG:

1) PARASITE DRAG:

i) FORM DRAG

ii) SKIN FRICTION DRAG

iii) INTERFERENCE DRAG

2) LIFT-INDUCED DRAG

3) WAVE DRAG

1.7.2.1 PARASITE DRAG:

Parasitic drag (also called parasite drag) is drag caused by moving a solid

object through a fluid. Parasitic drag is made up of multiple components

including viscous pressure drag (form drag), and drag due to surface roughness

(skin friction drag). Additionally, the presence of multiple bodies in relative

22

proximity may incur so called interference drag, which is sometimes described

as a component of parasitic drag. In aviation, induced drag tends to be greater at

lower speeds because a high angle of attack is required to maintain lift, creating

more drag. However, as speed increases the induced drag becomes much less,

but parasitic drag increases because the fluid is flowing faster around protruding

objects increasing friction or drag. At even higher speeds in the transonic, wave

drag enters the picture.

Each of these forms of drag changes in proportion to the others based on speed.

The combined overall drag curve therefore shows a minimum at some airspeed -

an aircraft flying at this speed will be at or close to its optimal efficiency. Pilots

will use this speed to maximize endurance (minimum fuel consumption), or

maximize gliding range in the event of an engine failure.

1.7.2.2 LIFT-INDUCED DRAG:

Lift-induced drag (also called induced drag) is drag which occurs as the result

of the creation of lift on a three-dimensional lifting body, such as the wing or

fuselage of an airplane. Induced drag consists of two primary components,

including drag due to the creation of vortices (vortex drag) and the presence of

additional viscous drag (lift-induced viscous drag). The vortices in the flow-

field, present in the wake of a lifting body, derive from the turbulent mixing of

air of varying pressure on the upper and lower surfaces of the body, which is a

Fig. 1.6. Lift in induced drag

23

necessary condition for the creation of lift. For an aircraft in flight, this means

that as the angle of attack, and therefore the lift coefficient, increases to the

point of stall, so does the lift-induced drag. At the onset of stall, lift is abruptly

decreased, as is lift-induced drag, but viscous pressure drag, a component of

parasite drag, and increases due to the formation of turbulent unattached flow

on the surface of the body.

1.7.2.3 WAVE DRAG:

Wave drag (also called compressibility drag) is drag which is created by the

presence of a body moving at high speed through a compressible fluid. In

aerodynamics, Wave drag consists of multiple components depending on the

speed regime of the flight. In transonic flight (Mach numbers greater than 0.5

and less than 1.0), wave drag is the result of the formation of shockwaves on the

body, formed when areas of local supersonic (Mach number greater than 1.0)

flow are created. In practice, supersonic flow occurs on bodies traveling well

below the speed of sound, as the local speed of air on a body increases when it

accelerates over the body, in this case above Mach 1.0. Therefore, aircraft flying

at transonic speed often incur wave drag through the normal course of

operation. In transonic flight, wave drag is commonly referred to as transonic

compressibility drag. Transonic compressibility drag increases significantly as

the speed of flight increases towards Mach 1.0, dominating other forms of drag

at these speeds.

In supersonic flight (Mach numbers greater than 1.0), wave drag is the result of

shockwaves present on the body, typically oblique shockwaves formed at the

leading and trailing edges of the body. In highly supersonic flows, or in bodies

with turning angles sufficiently large, unattached shockwaves, or bow waves

will instead form.

24

Additionally, local areas of transonic flow behind the initial shockwave may

occur at lower supersonic speeds, and can lead to the development of additional,

smaller shockwaves present on the surfaces of other lifting bodies, similar to

those found in transonic flows.

1.8 DRAG REDUTION TECHNIQUES

Drag reduction is one of the main objectives of the transport aircraft

manufacturers. The drag breakdown of a transport aircraft at cruise shows that

the skin friction drag and the lift-induced drag constitute the two main sources

of drag, approximately one half and one third of the total drag. Hybrid laminar

flow technology and innovative wing tip devices offer the greatest potential for

drag reduction. Aircraft performance improvement in off-design conditions can

also be obtained through trailing edge optimization, control of the shock

boundary layer interaction and of the boundary layer separation. The paper will

give an overview of the results obtained for the different mentioned topics and

will try to evaluate the potential gains offered by the different technologies.

Drag reduction of civil transport aircraft directly concerns performance, but also

indirectly, of course, cost, and environment. Fuel consumption represents about

22% of the Direct Operating Cost (DOC) which is of utmost importance for the

airlines, for a typical long range transport aircraft.

Drag reduction directly impacts on the DOC: a drag reduction of 1% can

lead to a DOC decrease of about 0.2% for a large transport aircraft. Other trade-

offs corresponding to a 1% drag reduction are 1.6 tons on the operating empty

weight or 10 passengers.

25

The environmental factors, such as noise, air pollution around airports and

impact on climate change, which are well underlined in [1], will also play an

important role for future growth of the civil aviation.

The impact of air travel on the environment will then become an increasing

powerful factor on aircraft design. It is also important to recall the main goals of

the vision 2020 launched by the European commission: a 50% cut in CO2

emissions per passenger kilometer and an 80% cut in nitrogen oxide emissions.

These objectives cannot be reached without breakthrough in today technologies.

Drag reduction is a great challenge but there is certainly room for

improvements. The drag breakdown of a civil transport aircraft shows that the

skin friction drag and the lift-induced drag constitute the two main sources of

drag, approximately one half and one third of the total drag for a typical long

range aircraft at cruise conditions. This is why specific research on these topics

has been initiated in European Research centers and it seems that Hybrid

Laminar Flow technology and innovative wing tip devices offer the greatest

potential. Aircraft performance improvement can also be obtained through

trailing edge optimization, control of the shock boundary layer interaction and

of boundary layer separation. In the following sections, the different

technologies which were investigated at ONERA will be presented and

illustrated by experimental results.

1.8.1 SKIN FRICTION DRAG REDUCTION

Two methods are generally considered for skin friction drag reduction.

The first one aims at reducing the turbulent skin friction while the second one

aims at delaying transition to maintain large extent of laminar flow.

1.8.1.1Turbulent skin friction reduction

A skin friction drag reduction can be obtained with the use of passive

boundary layer manipulators. Among the various devices, V-groove rib-lets

26

have demonstrated substantial reductions (up to 8%) of the local skin friction.

An experimental verification in a large wind tunnel was carried out in 1988 on a

1/11 scale complete model of the Airbus A320.

For the test, 2/3 of the wetted model surface was covered with the rib-lets for

which the previously mentioned V-groove cross-section has been chosen.

Viscous flow computations on the wing and on the fuselage have shown that a

rib-let depth of 0.023 mm can allow a average value of h + w=8 to be obtained.

Wind tunnel test was successful and total drag reductions up to 1.6% have been

demonstrated at corresponding cruise Mach number conditions.

With the guidelines of the previous wind tunnel investigations and the

recommendations coming from the structure, material and system teams, a flight

test was prepared with the Airbus A320 No 1. Overall performance and local

data were measured with and without the rib-lets, and drag reduction predictions

based on the wind tunnel tests were confirmed.

1.8.2 Hybrid laminar flow technology

A substantial reduction in fuel consumption and in CO2 emissions will

certainly require the adoption of laminar flow control in order to reduce the skin

friction. For small aircraft with low swept wing, laminar flow can be maintained

by shaping the airfoil and this concept is currently considered for new small jet

aircraft. However for high Reynolds number and high sweep encountered on a

large transport aircraft, suction has to be applied.

In the Hybrid Laminar Flow concept, the laminar flow can be maintained

by the application of suction in the region of the leading edge to control the

development of cross flow instabilities combined with favorable pressure

gradients in the spar box region. It is first necessary to ensure that the

attachment line remains laminar and to avoid contamination phenomenon. Anti

contamination devices have to be used to avoid the contamination of the

attachment line by the turbulent structures coming from the fuselage.

27

The suction system has to be designed according to various aerodynamic

and structure requirements. Main features of suction systems are laser drilled

titanium panel and suction chambers controlled by independent ducts.

The geometrical characteristics of perforated panel such as hole diameter,

porosity as well as chamber sizes are determined taking into account the suction

velocity range, computed by stability approach, and pressure distributions for

various aerodynamic conditions.

With suction systems, premature transition can be caused by outflow and

by roughness effects due to high velocities in the suction holes. Pressure drop

methods and suction criteria have to be used to avoid these premature

transitions.

Surface imperfections such as isolated roughness, gaps, steps and

waviness can provoke premature transition. It is then necessary to study their

effects on transition and to develop calculation methods and criteria in order to

estimate these effects. Recent studies have shown that modern manufacturing

techniques can provide smooth surfaces, compatible with laminar flow.

Recent progress carried out towards the understanding of transition

characteristics of swept-wing flows would allow to control the transition by

passive means. Some experiments presented in have shown that transition

governed by cross flow instabilities can be delayed using artificial roughness. In

this concept, the artificial vortices interact nonlinearly with the natural vortices

in such a way that the natural vortices are strongly reduced.

In this approach, the drag reduction could be lower than the one expected with

the HLF concept, but the drawbacks are also very limited. It is worthwhile to

investigate these passive means through basic experiments and non-linear PSE

computations, because they can contribute to the system simplification needed

for a future laminar aircraft.

28

1.8.3 LIFT-INDUCED DRAG REDUCTION

The second major drag component is the lift-induced drag. The classical

way to decrease the lift-induced drag is to increase the aspect ratio of the wing.

Wing aspect ratio is a compromise between aerodynamic and structure

characteristics and it is clear that for a given technology there is not a great

possibility to increase aspect ratios. The alternative is to develop wing tip

devices acting on the tip vortex which is at the origin of the lift-induced drag.

Basic studies have shown that drag reduction can be obtained with

variations in plan form geometry along a small fraction of the wing-span and

with aft-swept configurations.

Furthermore, the presents, as examples among the investigated shapes, the

wing tip turbine, the wing tip sails, the wing grid, the blended winglet and the

spiroid tip.

Fig.1.7. various wingtip devices

The concept of the blended winglet is to modify a large part of the wing

tip together with the winglet itself in order to obtain a very smooth blended

shape. The blended winglet is expected to be more efficient than a narrow one

to reduce the flow acceleration that occurs in the cross flow curvature and to

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decrease the vortex intensity as important chord variation is avoided. The

spiroid tip is a spiral loop obtained when joining by their tip a vertical winglet

and a horizontal one. This unconventional device seems promising to reduce the

tip vortex intensity but has a complex geometry difficult to optimize.

Total drag reduction of about 2% can be expected with such wing tip

devices. However, for industrial applications, wingtip devices have a strong

influence on the wing structure and aero- elastic effects have to be taken into

account through a multidisciplinary optimization approach.

1.8.4 WAVE DRAG REDUCTION

Even if the wave drag contribution to the total drag of a modern transport

aircraft is not high, there is room for some significant improvements through

adaptation of the aircraft to the variation of the flight conditions : an increase of

the cruise Mach number for example. This aerodynamic adaptation can be

realized with shock control or trailing edge devices.

1.8.4.1 Shock control devices

Among the different passive shock boundary layer control concepts

investigated, the bump concept seems promising. This concept is based on the

local modification of the airfoil surface in the shock region. The straight shock

is transformed into a lambda shock configuration and its strength is reduced by

the presence of the compression waves.

1.8.4.2 Trailing edge devices

For wave drag reduction, the concept of the thick cambered trailing edge

which increases the rear loading and reduces the upper surface pressure

recovery seems also very promising. This concept has then been investigated on

a wing body configuration under a co-operation with Airbus France. Tests were

carried out on a half-model in the wind tunnel and the results have been

carefully analyzed through far-field drag extraction techniques. The computed

and measured drag reduction obtained when the thick cambered trailing edge is

installed in the outer part of the wing. It is clear that the thick cambered trailing

30

edge concept can be used by the designer as an additional degree of freedom. Its

effects can also be obtained through a trailing edge deflector.

These results show that characteristics of the flow can be strongly modified with

the use of a trailing edge device which allows drag reduction and greater buffet

margin to be obtained. Important investigations are currently carried out to

adapt the wing geometry to the different flight conditions: cruise, take-off and

landing.

1.9 TYPES OF WINGS

Fixed-wing aircraft, popularly called aeroplanes, airplanes or just planes may be

built with many wing configurations. This page provides a breakdown of types,

allowing a full description of any aircraft’s wing configuration. For example the

Spitfire wing may be classified as a conventional low wing cantilever

monoplane with straight elliptical wings of moderate aspect ratio and slight

dihedral.Sometimes the distinction between types is blurred, for example the

wings of many modern combat aircraft may be described either as cropped

compound deltas with (forwards or backwards) swept trailing edge, or as

sharply tapered swept wings with large ―Leading Edge Root Extension‖ (or

LERX). All the configurations described have flown (if only very briefly) on

full-size aircraft, except as noted.

Some variants may be duplicated under more than one heading, due to their

complex nature. This is particularly so for variable geometry and combined

(closed) wing types.

1.9.1MONO PLANE

One wing plane. The most aeroplanes have been monoplanes. The wing

may be mounted at various positions relative to the fuselage.

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1. Low wing: mounted near or below the bottom of the fuselage.

2. Mid wing: mounted approximately halfway up the fuselage.

3. Shoulder wing: mounted on the upper part or ―shoulder‖ of the fuselage,

slightly below the top of the fuselage. A shoulder wing is sometimes

considered a subtype of high wing.

4. High wing: mounted on the upper fuselage. When contrasted to the

shoulder wing, applies to a wing mounted on a projection (such as the

cabin roof) above the top of the main fuselage.

Low wing

Mid wing

Shoulder wing

Fig.1.8. Types of aircraft wings

1.10 WING PARTS

Aileron

Flap

Rib

Spar

Spoiler

Stall strips

High wing

Parasol wing

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Wing tip

1.10.1 AILERON

An aileron is a hinged flight control surface usually attached to

the trailing edge of each wing of a fixed-wing aircraft. Ailerons are used in pairs

to control the aircraft in roll (or movement around the aircraft's longitudinal

axis), which normally results in a change in flight path due to the tilting of

the lift vector. Movement around this axis is called 'rolling' or 'banking'.

Fig 1.9 Parts of wing

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1.10.2 Flaps

Flaps are devices used to alter the lift characteristics of a wing and are

mounted on the trailing edges of the wings of a fixed-wing aircraftto reduce the

speed at which the aircraft can be safely flown and to increase the angle of

descent for landing. They shorten takeoff and landing distances. Flaps do this by

lowering the stall speed and increasing the drag.

Extending flaps increases the camber or curvature of the wing, raising the

maximum lift coefficient — the lift a wing can generate. This allows the aircraft

to generate as much lift, but at a lower speed, reducing the stalling speed of the

aircraft, or the minimum speed at which the aircraft will maintain flight.

Extending flaps increases drag, which can be beneficial during approach and

landing, because it slows the aircraft. On some aircraft, a useful side effect of

flap deployment is a decrease in aircraft pitch angle which lowers the nose

thereby improving the pilot's view of the runway over the nose of the aircraft

during landing. However the flaps may also cause pitch-up depending on the

type of flap and the location of the wing.

1.10.3 Spoiler

In aeronautics, a spoiler (sometimes called a lift dumper) is a device

intended to reduce lift in an aircraft. Spoilers are plates on the top surface of a

wing that can be extended upward into the airflow to spoil it. By so doing, the

spoiler creates a controlled stall over the portion of the wing behind it, greatly

reducing the lift of that wing section. Spoilers differ from airbrakes in that

airbrakes are designed to increase drag without affecting lift, while spoilers

reduce lift as well as increasing drag.

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Spoilers fall into two categories: those that are deployed at controlled angles

during flight to increase descent rate or control roll, and those that are fully

deployed immediately on landing to greatly reduce lift ("lift dumpers") and

increase drag. In modern fly-by-wire aircraft, the same set of control surfaces

serve both functions. Spoilers are used by nearly every glider (sailplane) to

control their rate of descent and thus achieve a controlled landing. An increased

rate of descent can also be achieved by lowering the nose of an aircraft, but this

would result in increased speed. Spoilers enable the approach to be made at a

safe speed for landing.

1.10.4 Spar

In a fixed-wing aircraft, the spar is often the main structural member of the

wing, running spanwise at right angles (or thereabouts depending on wing

sweep) to the fuselage. The spar carries flight loads and the weight of the wings

while on the ground. Other structural and forming members such as ribs may be

attached to the spar or spars, with stressed skin construction also sharing the

loads where it is used. There may be more than one spar in a wing or none at all.

However, where a single spar carries the majority of the forces on it, it is known

as the main spar. Spars are also used in other aircraft aerofoil surfaces such as

the tailplane and fin and serve a similar function, although the loads transmitted

may be different from those of a wing spar.

1.10.5 Ribs

In an aircraft, ribs are forming elements of the structure of a wing, especially in

traditional construction. By analogy with the anatomical definition of "rib", the

ribs attach to the main spar, and by being repeated at frequent intervals, form a

skeletal shape for the wing. Usually ribs incorporate the airfoil shape of the

wing, and the skin adopts this shape when stretched over the ribs.

1.10.6 Wing tip

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A wing tip is the part of the wing that is most distant from the fuselage of

a fixed-wing aircraft. Wing tips are also an expression of aircraft design style,

so their shape may be influenced by marketing considerations as well as

byaerodynamic requirements.Wing tips are often used by aircraft designers to

mount navigation lights, anti-collision strobe lights, landing lights, handholds,

and identification markings.Wing tip tanks can act as a winglet and distribute

weight more evenly across the wing spar.Aerobatic aircraft use wingtip

mounted crosses for visual attitude reference.

1.11 DESCRIPTION OF COMPOSITE STRUCTURES

1.11.1 Introduction

Composite materials are becoming more important in the construction of

aerospace structures. Aircraft parts made from composite materials, such as

fairings, spoilers, and flight controls, were developed during the 1960s for their

weight savings over aluminum parts. New generation large aircraft are designed

with all composite fuselage and wing structures, and the repair of these

advanced composite materials requires an in-depth knowledge of composite

structures, materials, and tooling. The primary advantages of composite

materials are their high strength, relatively low weight, and corrosion resistance.

1.11.2 Laminated Structures

Composite materials consist of a combination of materials that are mixed

together to achieve specific structural properties. The individual materials do

not dissolve or merge completely in the composite, but they act together as one.

Normally, the components can be physically identified as they interface with

one another. The properties of the composite material are superior to the

properties of the individual materials from which it is constructed. An advanced

composite material is made of a fibrous material embedded in a resin matrix,

36

generally laminated with fibers oriented in alternating directions to give the

material strength and stiffness. Fibrous materials are not new; wood is the most

common fibrous structural material known to man.

Applications of composites on aircraft include:

• Fairings

• Flight control surfaces

• Landing gear doors

• Leading and trailing edge panels on the wing and stabilizer

• Interior components

• Floor beams and floor boards

• Vertical and horizontal stabilizer primary structure on

large aircraft

• Primary wing and fuselage structure on new generation

large aircraft

• Turbine engine fan blades

• Propellers

1.11.2.1 Major Components of a Laminate

An isotropic material has uniform properties in all directions. The

measured properties of an isotropic material are independent of the axis of

testing. Metals such as aluminium and titanium are examples of isotropic

materials.

A fiber is the primary load carrying element of the composite material. The

composite material is only strong and stiff in the direction of the fibers.

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Unidirectional composites have predominant mechanical properties in one

direction and are said to be anisotropic, having mechanical and/or physical

properties that vary with direction relative to natural reference axes inherent in

the material. Components made from fiber reinforced composites can be

designed so that the fiber orientation produces optimum mechanical properties,

but they can only approach the true isotropic nature of metals, such as

aluminum and titanium.

A matrix supports the fibers and bonds them together in the composite

material. The matrix transfers any applied loads to the fibers, keeps the fibers in

their position and chosen orientation, gives the composite environmental

resistance, and determines the maximum service temperature of a composite.

1.11.2.2 Strength Characteristics

Structural properties, such as stiffness, dimensional stability, and strength

of a composite laminate, depend on the stacking sequence of the plies. The

stacking sequence describes the distribution of ply orientations through the

laminate thickness. As the number of plies with chosen orientations increases,

more stacking sequences are possible. For example, a symmetric eight-ply

laminate with four different ply orientations has 24 different stacking

sequences.

1.11.2.3 Fiber Orientation

The strength and stiffness of a composite buildup depends on the

orientation sequence of the plies. The practical range of strength and stiffness of

carbon fiber extends from values as low as those provided by fiberglass to as

high as those provided by titanium. This range of values is determined by the

orientation of the plies to the applied load. Proper selection of ply orientation in

advanced composite materials is necessary to provide a structurally efficient

design. The part might require 0° plies to react to axial loads, ±45° plies to react

38

to shear loads, and 90° plies to react to side loads. Because the strength design

requirements are a function of the applied load direction, ply orientation and ply

sequence have to be correct. It is critical during a repair to replace each

damaged ply with a ply of the same material and ply orientation. The fibers in a

unidirectional material run in one direction and the strength and stiffness is only

in the direction of the fiber. Pre-impregnated tape is an example of a

unidirectional ply orientation. The fibers in a bidirectional material run in two

directions, typically 90° apart.

A plain weave fabric is an example of a bidirectional ply orientation.

These ply orientations have strength in both directions but not necessarily the

same strength. The plies of a quasi-isotropic layup are stacked in a 0°, –45°,

45°, and 90° sequence or in a 0°, –60°, and 60° sequence. These types of ply

orientation simulate the properties of an isotropic material. Many aerospace

composite structures are made of quasi-isotropic materials.

1.11.2.4 Warp Clock

Warp indicates the longitudinal fibers of a fabric. The warp is the high

strength direction due to the straightness of the fibers. A warp clock is used to

describe direction of fibers on a diagram, spec sheet, or manufacturer’s sheets.

If thewarp clock is not available on the fabric, the orientation is defaulted to

zero as the fabric comes off the roll. Therefore, 90° to zero is the width of the

fabric across.

1.11.2.5 Fiber Forms

All product forms generally begin with spooled unidirectional raw fibers

packaged as continuous strands. An individual fiber is called a filament. The

word strand is also used to identify an individual glass fiber. Bundles of

filaments are identified as tows, yarns, or rovings. Fiberglass yarns are twisted,

while Kevlar yarns are not. Tows and rovings do not have any twist. Most

fibers are available as dry fiber that needs to be impregnated (impreg) with a

39

resin before use or prepreg materials where the resin is already applied to the

fiber.

1.11.2.6 Roving

A roving is a single grouping of filament or fiber ends, such as 20-end or

60-end glass rovings. All filaments are in the same direction and they are not

twisted. Carbon rovings are usually identified as 3K, 6K, or 12K rovings, K

meaning 1,000 filaments. Most applications for roving products utilize mandrels

for filament winding and then resin cure to final configuration.

1.11.2.7 Unidirectional (Tape)

Unidirectional prepreg tapes have been the standard within the aerospace

industry for many years, and the fiber is typically impregnated with

thermosetting resins. The most common method of manufacture is to draw

collimated raw (dry) strands into the impregnation machine where hot melted

resins are combined with the strands using heat and pressure. Tape products

have high strength in the fiber direction and virtually no strength across the

fibers. The fibers are held in place by the resin. Tapes have a higher strength

than woven fabrics.

1.11.2.8 Bidirectional (Fabric)

Most fabric constructions offer more flexibility for layup of complex

shapes than straight unidirectional tapes offer. Fabrics offer the option for resin

impregnation either by solution or the hot melt process. Generally, fabrics used

for structural applications use like fibers or strands of the same weight or yield

in both the warp (longitudinal) and fill (transverse) directions. For aerospace

structures, tightly woven fabrics are usually the choice to save weight,

minimizing resin void size, and maintaining fiber orientation during the

fabrication process.

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1.12 TYPES OF FIBERS

Fig1.10 Types of Fibers

1.12.1 Fiberglass

Fiberglass is often used for secondary structure on aircraft, such as

fairings, radomes, and wing tips. Fiberglass is also used for helicopter rotor

blades. There are several types of fiberglass used in the aviation industry.

Electrical glass, or E-glass, is identified as such for electrical applications. It has

high resistance to current flow. E-glass is made from borosilicate glass. S-glass

and S2-glass identify structural

fiberglass that have a higher strength than E-glass. S-glass is produced from

magnesia-alumina-silicate. Advantages of fiberglass are lower cost than other

composite materials, chemical or galvanic corrosion resistance, and electrical

properties (fiberglass does not conduct electricity). Fiberglass has a white color

and is available as a dry fiber fabric or prepreg material.

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1.12.2 Kevlar®

Kevlar® is DuPont’s name for aramid fibers. Aramid fibers are light

weight, strong, and tough. Two types of Aramid fiber are used in the aviation

industry. Kevlar® 49 has a high stiffness and Kevlar® 29 has a low stiffness.

An advantage of aramid fibers is their high resistance to impact damage, so they

are often used in areas prone to impact damage. The main disadvantage of

aramid fibers is their general weakness in compression and hygroscopy.

Service reports have indicated that some parts made from Kevlar®

absorb up to 8 percent of their weight in water. Therefore, parts made from

aramid fibers need to be protected from the environment. Another disadvantage

is that Kevlar® is difficult to drill and cut. The fibers fuzz easily and special

scissors are needed to cut the material. Kevlar® is often used for military

ballistic and body armor applications.

It has a natural yellow color and is available as dry fabric and prepreg

material. Bundles of aramid fibers are not sized by the number of fibers like

carbon or fiberglass but by the weight.

1.12.3 Carbon/Graphite

One of the first distinctions to be made among fibers is the difference

between carbon and graphite fibers, although the terms are frequently used

interchangeably. Carbon and graphite fibers are based on graphene (hexagonal)

layer networks present in carbon. If the graphene layers, or planes, are stacked

with three dimensional order, the material is defined as graphite. Usually

extended time and temperature processing is required to form this order, making

graphite fibers more expensive. Bonding between planes is weak. Disorder

frequently occurs such that only two-dimensional ordering within the layers is

present. This material is defined as carbon. Carbon fibers are very stiff and

strong, 3 to 10 times stiffer than glass fibers. Carbon fiber is used for structural

aircraft applications, such as floor beams, stabilizers, flight controls, and

42

primary fuselage and wing structure. Advantages include its high strength and

corrosion resistance. Disadvantages include lower conductivity than aluminum;

therefore, a lightning protection mesh or coating is necessary for aircraft parts

that are prone to lightning strikes. Another disadvantage of carbon fiber is its

high cost. Carbon fiber is gray or black in color and is available as dry fabric

and prepreg material. Carbon fibers have a high potential for causing galvanic

corrosion when used with metallic fasteners and structures.

1.12.4 Lightning Protection Fibers

An aluminum airplane is quite conductive and is able to dissipate the high

currents resulting from a lightning strike. Carbon fibers are 1,000 times more

resistive than aluminium to current flow, and epoxy resin is 1,000,000 times

more resistive (i.e., perpendicular to the skin).

The surface of an external composite component often consists of a ply or

layer of conductive material for lightning strike protection because composite

materials are less conductive than aluminum. Many different types of

conductive materials are used ranging from nickel-coated graphite cloth to

metal meshes to aluminized fiberglass to conductive paints. The materials are

available for wet layup and as prepreg. In addition to a normal structural repair,

the technician must also recreate the electrical conductivity designed into the

part. These types of repair generally require a conductivity test to be performed

with an ohmmeter to verify minimum electrical resistance across the structure.

When repairing these types of structures, it is extremely important to use only

the approved materials from authorized vendors, including such items as potting

compounds, sealants, adhesives, and so forth.

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1.13 MATRIX MATERIAL

1.13.1 Thermosetting Resins

Resin is a generic term used to designate the polymer. The resin, its

chemical composition, and physical properties fundamentally affect the

processing, fabrication, and ultimate properties of a composite material.

Thermosetting resins are the most diverse and widely used of all man-made

materials. They are easily poured or formed into any shape, are compatible with

most other materials, and cure readily (by heat or catalyst) into an insoluble

solid. Thermosetting resins are also excellent adhesives and bonding agents.

1.13.2 Polyester Resins

Polyester resins are relatively inexpensive, fast processing resins used

generally for low cost applications. Low smoke producing polyester resins are

used for interior parts of the aircraft. Fiber-reinforced polyesters can be

processed by many methods. Common processing methods include matched

metal molding, wet layup, press (vacuum bag) molding, injection molding,

filament winding, pultrusion, and autoclaving.

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Fig 1.11 cross sectional view of composite blade

1.13.3 Vinyl Ester Resin

The appearance, handling properties, and curing characteristics of vinyl

ester resins are the same as those of conventional polyester resins. However, the

corrosion resistance and mechanical properties of vinyl ester composites are

much improved over standard polyester resin composites.

1.13.4 Phenolic Resin

Phenol-formaldehyde resins were first produced commercially in the

early 1900s for use in the commercial market. Ureaformaldehyde and

melamine-formaldehyde appeared in the 1920–1930s as a less expensive

alternative for lower temperature use. Phenolic resins are used for interior

components because of their low smoke and flammability characteristics.

1.13.5 Epoxy

Epoxies are polymerizable thermosetting resins and are available in a

variety of viscosities from liquid to solid. There are many different types of

epoxy, and the technician should use the maintenance manual to select the

correct type for a specific repair. Epoxies are used widely in resins for prepreg

materials and structural adhesives.

45

The advantages of epoxies are high strength and modulus, low levels of

volatiles, excellent adhesion, low shrinkage, good chemical resistance, and ease

of processing. Their major disadvantages are brittleness and the reduction of

properties in the presence of moisture. The processing or curing of epoxies is

slower than polyester resins. Processing techniques include autoclave molding,

filament winding, press molding, vacuum bag molding, resin transfer molding,

and pultrusion. Curing temperatures vary from room temperature to

approximately 350 °F (180 °C). The most common cure temperatures range

between 250° and 350 °F (120–180 °C).

1.14 INTRODUCTION TO THORP T-211

THORP T211– Even the name screams power and performance. Affectionately

named after its designer, John Thorp, the six cylinders Jabiru 3300 equipped

T211 is not an ordinary aircraft. The combination of a light, yet strong airframe

with 120 horsepower provides a tremendous power to weight ratio which

creates short take off runs, strong climbs and impressive cruise speeds. The

Thorpedo is the first U.S. manufactured aircraft to earn the Special

Airworthiness certificate under the Light Sport Aircraft ruling. The FAA type

certified heritage ensures a proven design that has been tested to a higher

standard. With all its power, this nimble aircraft outperforms many in its class.

The available digital panel, luxurious interior and other options make this an

efficient or spirited recreational aircraft, suitable for both the seasoned pilot and

the new sport pilot alike.

Almost all the trainer and light sport aircraft have fixed landing gear

system. The landing gear system itself produces about 20 – 40% of the total

drag produced in an airplane. We know that the resultant power needed to

overcome this drag will vary as the cube of velocity, hence if the drag produced

46

in the aircraft is reduced, the total power consumed by the aircraft will be

reduced by a great extent. In order to do so, the perfect alternative would be the

retractable landing gear system, which will not only increase the performance of

the aircraft but will also enhance the maneuverability of the aircraft. We will

also be observing the various changes which will occur with respect to

aerodynamics and performance of the aircraft. The present wing of the aircraft

does not have the thickness to incorporate the landing gear of the aircraft, thus

we will have to change the wing of the aircraft keeping in mind the lift co-

efficient and the Reynolds no. at which the aircraft flies. Hence to check the

results we have made a prototype of the aircraft and tested the same in the wind

tunnel.

47

Fig 1.12 Explore view of T-211

1.14.1 Development

Thorp constructed eight prototypes, and had the design certified by the

FAA, but was unable to find a foothold in the Cessna-dominated post-war US

market. The T-211 was developed with a 90 horsepower continental upgrade in

1953. The project was therefore shelved until the homebuilding boom saw the

rights to the aircraft acquired first by Adams Industries and then by Thorp Aero

in the 1970s, the latter firm building five examples as the Thorp Arrow or T-

211 Aero Sport built in Sturgis Kentucky, but only sold overseas or part 141

operations due to current liability laws. The kits were then manufactured by AD

Aerospace in the United Kingdom and Venture Light Aircraft in the United

48

States.

Fig 1.3 Thorp T-211 aircraft

Indus Aviation began production of the T-211 to the guidelines of Light

Sport Aircraft in the mid-2000s. The Thorp T-211 was the first US-designed

Special Light Sport Aircraft to receive certification from the Federal Aviation

Administration. The light-sport version uses the 120 hp (89 kW) Jabiru 3300

engine, while the type certified version uses a 100 hp (75 kW) Continental O-

200 engine and is equipped for both VFR and IFR flying.

1.14.2 Specifications (T-211)

General characteristics

Crew: 1 pilot

49

Capacity: 1 passenger

Length: 18 ft 2 in (5.49 m)

Wingspan: 25 ft 0 in (7.62 m)

Height: 6 ft 1 in (1.92 m)

Wing area: 105 ft² (9.67 m²)

Empty weight: 750 lb (339 kg)

Max. takeoff weight: 1,270 lb (575 kg)

Power plant: 1 × Continental O-200-A, 100 hp (75 kW)

Fuel capacity: 21 gal usable (78 l)

Baggage capacity: 40 lb (18 kg)

Performance

Maximum speed: 120 mph (193 km/h)

Range: 375 miles (764 km)

Service ceiling: 12,500 ft (3,810 m)

Rate of climb: 750 ft/min (229 m/min)

Wing loading: 12.1 lb/ft² (60 kg/m²)

Power/mass: 0.08 hp/lb (0.13 kW/kg)

1.15 DESCRIPTION OF COMPANY

The project is an industrial project sponsored by Taneja Aerospace and

Aviation Ltd., Hosur. Part of the Pune based Indian Seamless group, TAAL was

established in 1994 as the first private sector company in the country to

manufacture general aviation i.e. non-military aircraft. The company’s vision at

the time was to create a nucleus facility for the development of an aeronautical

industry in India, TAAL entered into collaboration with Partenavia of Italy to

manufacture the six-seat twin piston engine P68C aircraft and the eleven-seat

twin turbo-prop Viator aircraft. While TAAL continues to manufacture Light

50

Transport and Trainer Aircraft, the company has since diversified its activities

and has established a significant presence in many segments of the aviation and

aeronautical industries in India.

Part of the Pune based Indian Seamless group, TAAL was established in 1994

as the first private sector company in the country to manufacture general

aviation i.e. non-military aircraft. The company's vision at the time was to

create a nucleus facility for the development of an aeronautical industry in India

and in particular to promote affordable general aviation in the country. To kick-

off this process, TAAL entered into a collaboration with Partenavia of Italy to

manufacture the six-seat twin piston-engine P68C aircraft and the eleven-seat

twin turbo-prop Viator aircraft.

While manufacture of Light Transport and Trainer Aircraft continues to be in

TAAL’s capability, the company has since diversified its activities and has

established a significant presence in many segments of the aviation and

aeronautical industries in India.

TAAL is into all Aviation related business activities namely, Aircraft

Manufacturing & Maintenance Centre and Aviation Infrastructure - Airfield &

MRO.

1.15.1 Aircraft Manufacturing & Maintenance Centre

This business has evolved from the initial business of the company, which was

to manufacture the Partenavia P68C, six seat, twin-engine aircraft in India (We

are the first and only private sector company in India to have built and certified

an aircraft).

We currently manufacture aero structures for Hindustan Aeronautics Limited

(HAL), Indian Space Research Organization (ISRO), National Aerospace

Laboratories (NAL) Aeronautical Development Establishment (ADE), and

51

Number of modifications on Indian Navy and Air force Helicopters and

Aircraft. Of these, the largest structures that we manufacture are for ISRO

where we build most of the structural assemblies for the Booster rockets of the

GSLV program. We have also built major structures of 14 seat Saras aircraft

developed by NAL. Once again, we would regard ourselves as the largest

dedicated private sector aero structure manufacturer in India.

1.15.2 Aviation Infrastructure - Airfield & MRO

TAAL has entered into an Aviation Infrastructure - Airfield & MRO facility

agreement with Air Works India (Engg) for establishment of commercial

Aircraft Maintenance and Operating Aviation Infrastructure - Airfield & MRO

Division services at TAAL's private airfield (Licensed) at Hosur, near

Bangalore. The runway at this airfield is capable of accepting Airbus A 320 and

Boeing 737 Series class of aircraft and the hanger is capable of accommodating

narrow body aircraft.

TAAL has DGCA Maintenance approval under CAR 145 for Maintenance of

Cessna Jet and other Light Transport Aircraft.

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CHAPTER-2

LITERATURE SURVEY

The evolution of composite material has replaced most of the

conventional material of construction in automobile, aviation industry etc. Fibre

reinforced composites have been widely successful in hundreds of applications

where there was a need for high strength materials. There are thousands of

custom formulations which offer FRPs a wide variety of tensile and flexural

strengths. When compared with traditional materials such as metals, the

combination of high strength and lower weight has made FRP an extremely

popular choice for improving a product’s design and performance.

2.1 Literature Survey Related to Present Work

Polymer matrix composites are predominantly used for the aerospace

industry, but the decreasing price of carbon Fibres is widening the applications

ofthese composites to include the automobile, marine, sports, biomedical,

construction, and other industries .Carbon Fibre polymer-matrix composites

have started to be used in automobiles mainly for saving weight for fuel

economy. The so-called graphite car employs carbon Fibre epoxy-matrix

composites for body panels, structural members, bumpers, wheels, drive shaft,

engine components, and suspension systems. This car is 570 kg lighter than an

equivalent vehicle made of steel. It weighs only 1250 kg instead of the

conventional 1800 kg for the average American car. Thermoplastic composites

with PEEK and polycarbonate (PC) matrices are finding use as spring elements

for car suspension systems. An investigation was conducted by Issac M Daniel

et.al on failure modes and criteria for their occurrence in composite columns

and beams. They found that the initiation of the various failure modes depends

on the material properties, geometric dimensions and type of loading.

53

They 18 reported that the loading type or condition determines the state of stress

throughout the composite structure, which controls the location and mode of

failure. The appropriate failure criteria at any point of the structure account for

the biaxiality or triaxiality of the state of stress. Jeam Marc et. investigates the

modeling of the flexural behavior of all-thermoplastic composite structures with

improved aesthetic properties, manufactured by isothermal compression

moulding. A four noded plate element based on a refined higher order shear

deformation theory is developed by Topdar et. for the analysis of composite

plates. This plate theory satisfies the conditions of inter-laminar shear stress

continuity and stress free top and bottom surfaces of the plate. Moreover, the

number of independent unknowns is the same as that in the first order shear

deformation theory. Banerji and Nirmal reported an increase in flexural strength

of unidirectional carbon Fibre/ Poly(methyl methacrylate), composite laminates

having polyethylene Fibres plies at the lower face Li and Xian showed that the

incorporation of a moderate amount of carbon Fibres into ultra-high-modulus

polyethylene (UHMPE) Fibres reinforced composites greatly improved the

compressive strength, flexural modulus while the addition of a small amount of

UHMPE Fibres into a carbon Fibre reinforced composite remarkably enhanced

the ductility with only a small decrease in compressive strength. Rohchoon and

Jang studied the effect of stacking sequence on the flexural properties and

flexural failure modes of aramid-UHMPE hybrid composites. The flexural

strength depends upon the type of Fibres at the compressive face and dispersion

extent of the Fibres. Matteson and Crane reported increase in flexural strength

by using unidirectional steel wire tapes in glass Fibre composites and carbon

Fibres composites.

54

They showed that the increase in flexural strength was due to a change in

failure mode from compressive buckling to nearly ductile tensile failure.

Bradley and Harris used unidirectional high carbon steel wires to improve the

impact properties of epoxy resin reinforced with unidirectional carbon Fibre

reinforced.

Unfortunately, flexural design methodologies rely on their experimental

boundary conditions and the particular laminate setup, since a scaling of the

results is very difficult. The occurrence of usual failure modes under flexural

loading conditions, like delamination, matrix tensile fracture, localized

compressive failure and Fibre shear failure is strongly dependent of the material

configuration (Fibre type, resin type, lay-up, and thickness), the loading type. In

this respect, three point bend test equipment along with specimen indicated in

figure 1 was used as a fast and cost efficient comparison tool.

Jawad Kadhim Uleiwi :

Studies Investigated the effect of fibre volume fraction on the flexural

properties of the laminated composite constructed of different layers, one of

them having reinforced glass fibre and the other layer reinforced with Kevlar

fibre has been investigated experimentally and the results illustrate that tension

stress decreases with the increase in fibre volume fraction of glass fibre of the

lower layer while it increases with the increase of Kevlar volume fraction of the

upper layer.

Wen-Pin Lin :

Studies analysed the Failure of Fibre-Reinforced Composite Laminates

under Biaxial Tensile Loading. With the onset of failure for individual lamina is

determined by a mixed failure criterion composed of the maximum stress

criteria. The lamina was described and observed to be brittle or degrading

modes with the collapse of the entire laminate.

55

Amjad J. Aref :

Examined the structural behaviour of the fibre reinforced polymer-

concrete hybrid bridge superstructure system subjected to negative moment

flexural loads through experimental procedures.

The experimental results showed that the design of the hybrid FRP-

concrete bridge superstructure under a negative flexural moment is found to be

stiffness- driven instead of strength-driven.

Slimane Metiche and Radhouane Masmoudi :

Studied the flexural behaviour of light weight fibre reinforced polymer

(FRP) poles. Experimental results show that the use of low linear density glass-

Fibres could provide an increase of the ultimate load carrying capacity up to 38

% for some fibre reinforced polymer poles. the positioning of the hole in the

compression side compared to the tension side leads to an increase of the

ultimate load carrying capacity up to 22 % for the 5.4m (18 feet) fibrereinforced

polymer poles and it was learnt that there was no significant effect (3,5%) for

the 12m (40 feet) fibre reinforced polymer poles. This is mainly due to the

stacking sequence and the stress states generated around the hole.

H. A. Rijsdijk :

Investigated the influence of maleic-anhydride-modified polypropylene

(m-PP) on monotonic mechanical properties of continuous-glass-fibrereinforced

polypropylene (PP) composites. This study showed an increase in composite

strength as a result of the addition of maleic-anhydride- modified PP to

continuous-glass-fibre-reinforced PP composites. An optimum in both

longitudinal and transverse flexural strength was reached for composites based

on a PP matrix with 10wt% m-PP.

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P.N.B. Reis :

Studied the flexural behaviour of hand manufactured hybrid laminated

composites with a hemp natural fibre/polypropylene core and two glass

fibres/polypropylene surface layers at each side of the specimen.

Laminate composites (LC) present an ultimate strength about 4% higher

than the hybrid laminated composites (HLC) associated to changes in failure

mechanisms, while the stiffness modulus was also about 3.8% higher. Fatigue

strength of hybrid laminated composites 21 is also about 20% lower than the

laminated composites as consequence of the change of the failure mechanisms

and of the different static strengths.

M. Davallo :

Investigated the Mechanical behaviour of unidirectional glasspolyester

composites to identify performance differences of composites with different

glass lay-ups and laminate thicknesses during flexure and tensile testing formed

by hand lay-up moulding (HLU). es. The damage generated in the composites

exhibited matrix cracking on the lower face followed by the coalescence of

delaminations formed within the reinforcing plies.

Michel Espinosa Klymus :

Evaluated the fracture pattern of four composites for indirect dental

restoration relating to three-point flexural strength. Further the compressive

strength and modulus of elasticity were also addressed. Composites

polymerized under high temperatures (belle Glass and Targis) had higher

flexural strength and elastic modulus values than composites polymerized by

light temperatures (Artglass and Solidex). It was found that they failed earlier

under compression because they were more rigid and showed partial fracture in

the material bulk.

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S. Benjamin Lazarus :

Investigated the mechanical properties of natural Fibre developed using a

plant fibre which is used for green manuring called Sunhemp. Polyester is used

as the matrix to prepare the composite.

From the results the applications of the composite for some specific

purposes can be decided upon since the maximum value of strength is achieved

for a particular Fibre length and Fibre weight ratio.

M. Wesolowski :

Studied the elastic properties of laminated composites by different Non-

Destructive techniques. Two carbon fibre XP45 Turane Resin laminated

composite plates and four beams cut from the plates along their principal

directions 1, 2 (two beams from each plate), are chosen for the study. Among all

proposed methods for the elastic properties characterization, the approach based

on the inverse technique is most suited for the convenient, fast, and accurate

identification of elastic properties.

J. Davies and H. Hamada:

Investigated the flexural properties of hybrid unidirectional fibre

reinforced polymer (FRP) composites containing a mixture of carbon (C) and

silicon carbide (SiC) fibres were evaluated at span to depth (S/d) ratios of 16,

32, and 64. The hybrid composite flexural strength was generally higher than

either the pure CFRP or SiC fibre composites. The work of fracture was a factor

of 2.6 larger for the S 4 /C 4 specimen compared to the S specimen and

suggests that these hybrid FRP composites may have a role as energy absorption

materials. The compressive stress, compressive strain and modulus to failure of

the SiC fibre were estimated to be 3.46 GPa, 157 GPa, and 0.018, respectively.

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IH Tacir :

Studied the reinforcing effect of glass fibres on the fracture

resistance and flexural strength of acrylic resins. In this study, statistically

significant differences were found in the flexural strength of the specimens. The

injectionmoulded, fibre reinforced polymers had significantly lower flexural

strength than the injection-moulded composites, and the microwave-moulded,

fibre reinforced composites had lower flexural strength than the microwave-

moulded composites.

The fracture resistance was significantly higher in the injection moulded,

fibre-reinforced composites than in the injection-moulded composites, and the

fracture resistance was significantly higher in the microwave moulded, fibre-

reinforced composites, than in the microwave-moulded composites.

Hoo Tien Kuan :

Evaluated the mechanical properties of composite

materials based on two types of self-reinforced polypropylene (SRPP) and a

glass fibre reinforced polypropylene are investigated under quasi-static and

dynamic loading conditions. Hybrid laminates based on glass fibre reinforced

polypropylene skins and a self-reinforced polypropylene core was manufactured

using a compression moulding technique. Hybridising the glass and

polypropylene fibre composites in this manner combines the strength and

stiffness of the glass fibres system with the excellent impact resistance and low

density of the self-reinforced polypropylene composite. Tests have shown that

increasing the volume fraction of self-reinforced polypropylene can enhance the

energy-absorbing characteristics of the hybrid composites.

Further with the increase in the amount of glass fibre in the reinforced

composite there was an increase in the flexural modulus of the hybrid

composites.

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M. Davallo :

Flexural properties of continuous random glass-polyester composites

formed by resin transfer moulding (RTM) and hand-lay up (HLU) moulding

have been studied to determine the effects of glass content, composite thickness,

reinforcement geometry and type of fabrication on damage developed during

flexure tests. Strain values both at maximum-load and failure were determined.

The failure strains of the two sets of composite series were relatively constant.

Hence, both types of composite series appeared to fail at a critical strain value.

The damage developed during the test was monitored on the side of each

polished beam using an optical microscope.

S. Tolson and N. Zasara:

Investigated the computational model for determining the ultimate

strength of an arbitrary laminated composite plate. A new higher order shear

deformation plate theory was developed. The theory utilizes seven degrees of 24

freedom at each node. An improvement in the accuracy of the transverse shear

stress was obtained by calculating these stresses using three-dimensional

elasticity equilibrium equations. The composite failure analysis is used to

determine the first and last ply failure of a laminated composite plate. The

computer programming was developed based on the seven degree of freedom

higher order shear deformation plate theory.

Geon-Woong Lee :

Studied the mechanical properties and failure mechanisms of through-the-

thickness stitched plain weave glass fabric/polyurethane foam/epoxy

composites. Hybrid composites were fabricated using resin infusion process

(RIP). Stitched sandwich composite increased drastically the flexural properties

as compared with the unstitched fabrics.

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Breaking of stitching in yarns was observed during the flexural test and

thus the failure mode yielded relatively high flexural properties. Polymer

composites with stitched sandwich structure improved the mechanical

properties with increasing the number of stitching yarns. It was concluded from

the study that proper combination of stitching density and types of stitching

fibre is important factor for through-the-thickness stitched composite panels.

N.K. Naik:

Investigated the inter laminar shear behaviour of typical polymer matrix

composites under high strain rate shear loading. Tensional split Hopkinson bar

(TSHB) apparatus is used for the studies in the shear strain rate range of 496–

1000/ s. It is observed that the interlaminar shear strength at high strain rate is

enhanced compared with that at quasistatic loading. Further, it is observed that

the inter laminar shear strength increases with increasing shear strain rate within

the range of shear strain rate considered.

Slavisa Putic:

This paper outlines the experimental investigation of inter laminar shear

strength as the critical mechanical property of composite constructions of

structure elements Placed between two thin glass mat layers where a layer is

placed on the glass fabric of the same structure but of different density, with

different polyester resin matrices. The significance of the shear strength lies in

the fact that for all types of composites it is strongly influenced by factors

weakening the interface binds.

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W.Richards Thissels:

The IM6 Fibres 3051/6 epoxy resin showed a 40% increased in stress

strain slope under compression loading at strain rate of 2000 L/S than 1×10-3

L/S when the applied load was parallel,45 0 and normal to the Fibre axis.The

compression test showed that delamination significant failure component.

The applicability of current hole in plate analytical methods to highly

anisotropic material is there questionable. Both hole in a plate analytical

methods indicates that GI is about 50% higher than G1.

Jane Maria Faulstich de Paiva:

This paper shows a study involving mechanical (flexural, shear, tensile

and compressive tests) and morphological characterizations of four different

laminates based on 2 epoxy resin systems (8552TM and F584TM). The results

show that the F584-epoxy matrix laminates present better mechanical properties

in the tensile and compressive tests than 8552 composites. Further it is observed

that PW laminates for both matrices show better flexural and inter laminar shear

properties.

Roberto J. Cano and Marvin B. Dow:

In this study, the unidirectional laminate strengths and moduli, of

notched (open-hole) and un notched specimens in tension and Compression

tests were performed and the properties of quasi isotropic polymer composite

laminates, and compression after impact strengths of five carbon fibre /

toughened matrix composites,.This investigation found that all 26 five materials

were stronger and more impact damage tolerant than more brittle carbon/epoxy

composite materials currently used in aircraft structures.

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CHAPTER-3

FABRICATION OF MODEL

The fabrication process of the aircraft model can be sub-divided into 3 basic

steps viz.

3.1 Carving:

Carving of the aircraft model means precise shaping the wood into the desired

without using any powered tools. The wood used for the fabrication of the

model is the Balsa wood, which are lightweight, simple to construct and

inexpensive to gather materials for. Extreme accuracy has to be maintained in

making the model as the whole success of the project depends on it. Various

tools that were used are wooden files, sand paper, hacksaw blade, bench knives,

straight chisels, skew chisels etc.

3.2 Fixing:

The second stage of the fabrication is to fix the various parts of the aircraft more

or less like assembly. The parts that were fixed to the fuselage were the wings,

propeller, vertical stabilizer and the horizontal stabilizer. Various adhesives

were used in this process like fevicol, anabond and m-seal.

3.3 Primer Coating / Artwork:

Once the adhesives have dried then comes the final stage in fabrication process

– the artwork. Before the model is painted primer coating has to be given to

model. A primer is a preparatory coating put on materials before painting.

Priming ensures better adhesion of paint to the surface, increases paint

durability, and provides additional protection for the material being painted.

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Fig 3.1 FABRICATION OF MODEL

The above figure gives us a pictorial description as how the model looks with

primer coated over it. Once the primer has dried off the model has been painted

with the desired colors.

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Fig 3.2 Tools for fabrication

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3.4 Fabrication methods

There are numerous methods for fabricating composite components. Some

methods have been borrowed (injection molding, for example), but many were

developed to meet specific design or manufacturing challenges. Selection of a

method for a particular part, therefore, will depend on the materials, the part

design and end-use or application.

Autoclaves cure composites with heat and pressure and are important tools for

curing parts made with thermoset resins. Improvements in control software are

helping autoclave operators such as this one at Helicomb International (Tulsa,

Okla.) increase throughput by 35 to 40 percent. At the same time, new resin

formulations are being developed for out-of-autoclave cure processing.

Fig 3.3 Fabrication process

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The most basic fabrication method for thermoset composites is hand layup,

which typically involves laying dry plies or prepreg plies by hand onto a tool to

form a laminate stack. Here, technicians at Liberty Aerospace (Melbourne, Fla.)

hand lay carbon/epoxy prepreg for a general aviation part.

There are numerous methods for fabricating composite components. Some

methods have been borrowed (injection molding, for example), but many were

developed to meet specific design or manufacturing challenges. Selection of a

method for a particular part, therefore, will depend on the materials, the part

design and end-use or application.

Composite fabrication processes involve some form of molding, to shape the

resin and reinforcement. A mold tool is required to give the unformed resin

/fiber combination its shape prior to and during cure. For an overview of mold

types and materials and methods used to make mold tools.

The most basic fabrication method for thermoset composites is hand layup,

which typically consists of laying dry fabric layers, or ―plies,‖ or prepreg plies,

by hand onto a tool to form a laminate stack. Resin is applied to the dry plies

after layup is complete (e.g., by means of resin infusion). In a variation known

as wet layup, each ply is coated with resin and ―debulked‖ or compacted after it

is placed.

Several curing methods are available. The most basic is simply to allow cure to

occur at room temperature. Cure can be accelerated, however, by applying heat,

typically with an oven, and pressure, by means of a vacuum. For the latter, a

vacuum bag, with breather assemblies, is placed over the layup and attached to

the tool, then evacuated using a vacuum pump before cure.

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The vacuum bagging process consolidates the plies of material and significantly

reduces voids due to the off-gassing that occurs as the matrix progresses

through its chemical curing stages.

Many high-performance thermoset parts require heat and high consolidation

pressure to cure — conditions that require the use of an autoclave. Autoclaves,

generally, are expensive to buy and operate. Manufacturers that are equipped

with autoclaves usually cure a number of parts simultaneously. Computer

systems monitor and control autoclave temperature, pressure, vacuum and inert

atmosphere, which allows unattended and/or remote supervision of the cure

process and maximizes efficient use of the technique.

When heat is required for cure, the part temperature is ―ramped up‖ in small

increments, maintained at cure level for a specified period of time defined by

the resin system, then ―ramped down‖ to room temperature, to avoid part

distortion or warp caused by uneven expansion and contraction.

When this curing cycle is complete and after parts are demolded, some parts go

through a secondary freestanding postcure, during which they are subjected for

a specific period of time to a temperature higher than that of the initial cure to

enhance chemical crosslink density.

Electron-beam (E-beam) curing has been explored as an efficient curing method

for thin laminates. In E-beam curing, the composite layup is exposed to a stream

of electrons that provide ionizing radiation, causing polymerization and

crosslinking in radiation-sensitive resins. X-ray and microwave curing

technologies work in a similar manner. A fourth alternative, ultraviolet (UV)

curing, involves the use of UV radiation to activate a photoinitiator added to a

thermoset resin, which, when activated, sets off a crosslinking reaction. UV

curing requires light-permeable resin and reinforcements.

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An emerging technology is the monitoring of the cure itself. Dielectric

cure monitors measure the extent of cure by gauging the conductivity of ions —

small, polarized, relatively insignificant impurities that are resident in resins.

Ions tend to migrate toward an electrode of opposite polarity, but the speed of

migration is limited by the viscosity of the resin — the higher the viscosity, the

slower the speed. As crosslinking proceeds during cure, resin viscosity

increases. Other methods include dipole monitoring within the resin, the

monitoring of microvoltage produced by the crosslinking, monitoring of the

exothermic reaction in the polymer during cure and, potentially, the use of

infrared monitoring via fiber-optic technology.

3.5 Open molding

Open contact molding in one-sided molds is a low-cost, common process

for making fiberglass composite products. Typically used for boat hulls and

decks, RV components, truck cabs and fenders, spas, bathtubs, shower stalls and

other relatively large, noncomplex shapes, open molding involves either hand

layup or a semi-automated alternative, sprayup.

In an open-mold sprayup application, the mold is first treated with mold

release. If a gel coat is used, it is typically sprayed into the mold after the mold

release has been applied. The gel coat then is cured and the mold is ready for

fabrication to begin. In the sprayup process, catalyzed resin (viscosity from 500

to 1,000 cps) and glass fiber are sprayed into the mold using a chopper gun,

which chops continuous fiber into short lengths, then blows the short fibers

directly into the sprayed resin stream so that both materials are applied

simultaneously. To reduce VOCs, piston pump-activated, non-atomizing spray

guns and fluid impingement spray heads dispense gel coats and resins in larger

droplets at low pressure. Another option is a roller impregnator, which pumps

resin into a roller similar to a paint roller.

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In the final steps of the sprayup process, workers compact the laminate by

hand with rollers. Wood, foam or other core material may then be added, and a

second sprayup layer imbeds the core between the laminate skins. The part is

then cured, cooled and removed from the reusable mold.

Hand layup and sprayup methods are often used in tandem to reduce

labor. For example, fabric might first be placed in an area exposed to high

stress; then, a spray gun might be used to apply chopped glass and resin to build

up the rest of the laminate.

Balsa or foam cores may be inserted between the laminate layers in either

process. Typical glass fiber volume is 15 percent with sprayup and 25 percent

with hand layup.

Sprayup processing, once a very prevalent manufacturing method, has

begun to fall out of favor. Federal regulations in the U.S. and similar rules in the

EU have mandated limits on worker exposure to, and emission into the

environment of VOCs and hazardous air pollutants (HAPs). Styrene, the most

common monomer used as a diluent in thermoset resins, is on both lists.

Because worker exposure to and emission of styrene is difficult and expensive

to control in the sprayup process, many composites manufacturers have

migrated to closed mold, infusion-based processes, which better contain and

manage styrenes.

Although open molding via hand layup is being replaced by faster and

more technically precise methods (as the following makes clear), it is still

widely used in the repair of composite parts. For more information about

―Composites repair‖ see the so-named article under "Editor's Picks."

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3.6 Resin infusion processes

Ever-increasing demand for faster production rates has pressed the

industry to replace hand layup with alternative fabrication processes and has

encouraged fabricators to automate those processes wherever possible.

A common alternative is resin transfer molding (RTM), sometimes

referred to as liquid molding. RTM is a fairly simple process: It begins with a

two-part, matched, closed mold that is made of metal or composite material.

Dry reinforcement (typically a preform) is placed into the mold and the mold is

closed. Resin and catalyst are metered and mixed in dispensing equipment, then

pumped into the mold under low to moderate pressure through injection ports,

following predesigned paths through the preform. Extremely low-viscosity resin

is used in RTM applications for thick parts to permeate preforms quickly and

evenly before cure. Both mold and resin can be heated, as necessary, for

particular applications. RTM produces parts without an autoclave. However,

when cured and demolded, a part destined for a high-temperature application

usually undergoes postcure. Most RTM applications use a two-part epoxy

formulation. The two parts are mixed just before they are injected.

Bismaleimide and polyimide resins also are available in RTM formulations.

Light RTM is a variant of RTM that is growing in popularity. In Light RTM,

low injection pressure, coupled with vacuum, allow the use of less-expensive,

lightweight two-part molds or a very lightweight, flexible upper mold.

The benefits of RTM are impressive. Generally, the dry preforms and

resins used in RTM are less expensive than prepreg material and can be stored

at room temperature. The process can produce thick, near-net shape parts,

eliminating most post-fabrication work. It also yields dimensionally accurate

complex parts with good surface detail and delivers a smooth finish on all

exposed surfaces.

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It is possible to place inserts inside the preform before the mold is closed,

allowing the RTM process to accommodate core materials and integrate

―molded in‖ fittings and other hardware into the part structure. Moreover, void

content on RTM’d parts is low, measuring in the 0 to 2 percent range. Finally,

RTM significantly cuts cycle times and can be adapted for use as one stage in

an automated, repeatable manufacturing process for even greater efficiency,

reducing cycle time from what can be several days, typical of hand layup, to just

hours — or even minutes. A recent variant of RTM, called high-pressure RTM

(HP-RTM), is gaining attention for its potential to quickly produce automotive

parts. Typically designed as a completely automated system including mold

shuttles, the ability to rapidly fill a mold loaded with a preform with a very fast

curing resin shows promise for high production.

In contrast to RTM, where resin and catalyst are premixed prior to

injection under pressure into the mold, reaction injection molding (RIM) injects

a rapid-cure resin and a catalyst into the mold in two separate streams. Mixing

and the resulting chemical reaction occur in the mold instead of in a dispensing

head. Automotive industry suppliers combine structural RIM (SRIM) with rapid

preforming methods to fabricate structural parts that don’t require a Class A

finish. Programmable robots have become a common means to spray a chopped

fiberglass/binder combination onto a vacuum-equipped preform screen or mold.

Robotic sprayup can be directed to control fiber orientation. A related

technology, dry fiber placement, combines stitched preforms and RTM. Fiber

volumes of up to 68 percent are possible, and automated controls ensure low

voids and consistent preform reproduction, without the need for trimming.

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3.7 Vacuum-assisted resin transfer molding (VARTM)

It refers to a variety of related processes that represent the fastest-

growing new molding technology. The salient difference between VARTM-

type processes and RTM is that in VARTM, resin is drawn into a preform

through use of a vacuum only, rather than pumped in under pressure. VARTM

does not require high heat or pressure. For that reason, VARTM operates with

low-cost tooling, making it possible to inexpensively produce large, complex

parts in one shot.

In the VARTM process, fiber reinforcements are placed in a one-sided

mold, and a cover (typically a plastic bagging film) is placed over the top to

form a vacuum-tight seal. The resin typically enters the structure through

strategically placed ports and feed lines, termed a ―manifold.‖ It is drawn by

vacuum through the reinforcements by means of a series of designed-in

channels that facilitate wetout of the fibers. Fiber content in the finished part

can run as high as 70 percent.

Current applications include marine, ground transportation and

infrastructure parts. A twist on the VARTM process is the use of two bags,

termed double-bag infusion, which uses one vacuum pump attached to the inner

bag to extract volatiles and entrapped air, and a second vacuum pump on the

outer bag to compact the laminate. This method has been employed by The

Boeing Co. (Chicago, Ill.) and NASA, as well as small fabricating firms, to

produce aerospace-quality laminates without an autoclave.

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3.8 Resin film infusion (RF)

It is a hybrid process in which a dry preform is placed in a mold on top

of a layer, or interleaved with multiple layers, of high-viscosity resin film.

Under applied heat, vacuum and pressure, the resin liquefies and is drawn into

the preform, resulting in uniform resin distribution, even with high-viscosity,

toughened resins, because of the short flow distance.

3.9 High-volume molding methods

3.9.1 Compression molding

It is a high-volume thermoset molding process that employs expensive

but very durable metal dies. It is an appropriate choice when production

quantities exceed 10,000 parts. As many as 200,000 parts can be turned out on a

set of forged steel dies, using sheet molding compound (SMC), a composite

sheet material made by sandwiching chopped fiberglass between two layers of

thick resin paste. To form the sheet, the resin paste transfers from a metering

device onto a moving film carrier. Chopped glass fibers drop onto the paste, and

a second film carrier places another layer of resin on top of the glass. Rollers

compact the sheet to saturate the glass with resin and squeeze out entrapped air.

Fig 3.4 composite layers

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The resin paste initially is the consistency of molasses (between 20,000

and 40,000 cps); over the next three to five days, its viscosity increases and the

sheet becomes leather-like (about 25 million cps), ideal for handling.

When the SMC is ready for molding, it is cut into smaller sheets and the

charge pattern (ply schedule) is assembled on a heated mold (121°C to 262°C or

250°F to 325°F). The mold is closed and clamped, and pressure is applied at

24.5 to 172.4 bar (500 to 2,500 psi). As material viscosity drops, the SMC flows

to fill the mold cavity. After cure, the part is demolded manually or by integral

ejector pins.

A typical low-profile (less than 0.05 percent shrinkage) SMC formulation for a

Class A finish consists, by weight, of 25 percent polyester resin, 25 percent

chopped glass, 45 percent fillers and 5 percent additives. Fiberglass thermoset

SMC cures in 30 to 150 seconds and overall cycle time can be as low as 60

seconds. Other grades of SMC include low-density, flexible and pigmented

formulations. Low-pressure SMC formulations that are now on the market offer

open molders low-capital-investment entry into closed-mold processing with

near-zero VOC emissions and the potential for very high-quality surface finish.

Automakers are exploring carbon fiber-reinforced SMC, hoping to take

advantage of carbon’s high strength- and stiffness-to-weight ratios in exterior

body panels and other parts. Newer, toughened SMC formulations help prevent

microcracking, a phenomenon that previously caused paint ―pops‖ during the

painting process (surface craters caused by outgassing, the release of gasses

trapped in the microcracks during oven cure).

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Composites manufacturers in industrial markets are formulating their own

resins and compounding SMC in-house to meet needs in specific applications

that require UV, impact and moisture resistance and have surface-quality

demands that drive the need for customized material development.

3.9.2 Injection molding

It is a fast, high-volume, low-pressure, closed process using, most

commonly, filled thermoplastics, such as nylon with chopped glass fiber. In the

past 20 years, however, automated injection molding of BMC has taken over

some markets previously held by thermoplastic and metal casting

manufacturers.

In the BMC injection molding process, a ram- or screw-type plunger forces a

metered shot of material through a heated barrel and injects it (at 5,000 to

12,000 psi) into a closed, heated mold. In the mold, the liquefied BMC flows

easily along runner channels and into the closed mold. After cure and ejection,

parts need only minimal finishing. Injection speeds are typically one to five

seconds, and as many as 2,000 small parts can be produced per hour in some

multiple-cavity molds.

Parts with thick cross-sections can be compression molded or transfer molded

with BMC. Transfer molding is a closed-mold process wherein a measured

charge of BMC is placed in a pot with runners that lead to the mold cavities. A

plunger forces the material into the cavities, where the product cures under heat

and pressure.

76

3.9.4 Filament winding

It is a continuous fabrication method that can be highly automated and

repeatable, with relatively low material costs. A long, cylindrical tool called a

mandrel is suspended horizontally between end supports, while the ―head‖ —

the fiber application instrument — moves back and forth along the length of a

rotating mandrel, placing fiber onto the tool in a predetermined configuration.

Computer-controlled filament-winding machines are available, equipped with

from 2 to 12 axes of motion.

In most thermoset applications, the filament winding apparatus passes the fiber

material through a resin ―bath‖ just before the material touches the mandrel.

This is called wet winding. However, a variation uses towpreg, that is,

continuous fiber pre-impregnated with resin. This eliminates the need for an

onsite resin bath. In a slightly different process, fiber is wound without resin

(dry winding). The dry shape is then used as a preform in another molding

process, such as RTM.

Following oven or autoclave curing, the mandrel either remains in place to

become part of the wound component or, typically, it is removed. One-piece

cylindrical or tapered mandrels, usually of simple shape, are pulled out of the

part with mandrel extraction equipment. Some mandrels, particularly in more

complex parts, are made of soluble material and may be dissolved and washed

out of the part.

Others are collapsible or built from several parts that allow its disassembly and

removal in smaller pieces. Filament-winding manufacturers often ―tweak‖ or

slightly modify off-the-shelf resin to meet specific application requirements.

Some composite part manufacturers develop their own resin formulations.

77

In thermoplastics winding, all material is in prepreg form, so a resin bath

is not needed. Material is heated as it is wound onto the mandrel — a process

known as curing ―on the fly‖ or in-situ consolidation. The prepreg is heated,

layed down, compacted, consolidated and cooled in a single, continuous

operation. Thermoplastic prepregs eliminate autoclave curing (cutting costs and

size limitations) and reduce raw material costs, and the resulting parts can be

reprocessed to correct flaws.

Filament winding yields parts with exceptional circumferential or ―hoop‖

strength. The highest-volume single application of filament winding is golf club

shafts. Fishing rods, pipe, pressure vessels and other cylindrical parts comprise

most of the remaining business.

3.10 Pultrusion

Its like RTM, has been used for decades with glass fiber and polyester resins,

but in the last 10 years the process also has found application in advanced

composites applications. In this relatively simple, low-cost, continuous process,

the reinforcing fiber (usually roving, tow or continuous mat) is typically pulled

through a heated resin bath and then formed into specific shapes as it passes

through one or more forming guides or bushings. The material then moves

through a heated die, where it takes its net shape and cures. Further

downstream, after cooling, the resulting profile is cut to desired length.

Pultrusion yields smooth finished parts that typically do not require post

processing. A wide range of continuous, consistent, solid and hollow profiles

are pultruded, and the process can be custom-tailored to fit specific applications.

78

3.11 Tube rolling

It is a longstanding composites manufacturing process that can produce

finite-length tubes and rods. It is particularly applicable to small-diameter

cylindrical or tapered tubes in lengths as great as 20 ft/6.2m. Tubing diameters

up to 6 inches/152 mm can be rolled efficiently. Typically, a tacky prepreg

fabric or unidirectional tape is used, depending on the part. The material is

precut in patterns that have been designed to achieve the requisite ply schedule

and fiber architecture for the application. The pattern pieces are laid out on a flat

surface and a mandrel is rolled over each one under applied pressure, which

compacts and debulks the material. When rolling a tapered mandrel — e.g., for

a fishing rod or golf shaft — only the first row of longitudinal fibers falls on the

true 0° axis. To impart bending strength to the tube, therefore, the fibers must be

continuously reoriented by repositioning the pattern pieces at regular intervals.

3.12 Automated fiber placement (AFP)

The fiber placement process automatically places multiple individual

prepreg tows onto a mandrel at high speed, using a numerically controlled,

articulating robotic placement head to dispense, clamp, cut and restart as many

as 32 tows simultaneously. Minimum cut length (the shortest tow length a

machine can lay down) is the essential ply-shape determinant. The fiber

placement heads can be attached to a 5-axis gantry, retrofitted to a filament

winder or delivered as a turnkey custom system. Machines are available with

dual mandrel stations to increase productivity.

Advantages of fiber placement include processing speed, reduced

material scrap and labor costs, parts consolidation and improved part-to-part

uniformity. Often, the process is used to produce large thermoset parts with

complex shapes.

79

3.13 Automated tape laying (ATL)

ATL an even speedier automated process in which prepreg tape, rather

than single tows, is laid down continuously to form parts. It is often used for

parts with highly complex contours or angles. Tape layup is versatile, allowing

breaks in the process and easy direction changes, and it can be adapted for both

thermoset and thermoplastic materials. The head includes a spool or spools of

tape, a winder, winder guides, a compaction shoe, a position sensor and a tape

cutter or slitter. In either case, the head may be located on the end of a multiaxis

articulating robot that moves around the tool or mandrel to which material is

being applied, or the head may be located on a gantry suspended above the tool.

Alternatively, the tool or mandrel can be moved or rotated to provide the head

access to different sections of the tool. Tape or fiber is applied to a tool in

courses, which consist of one row of material of any length at any angle.

Multiple courses are usually applied together over an area or pattern and are

defined and controlled by machine-control software that is programmed with

numerical input derived from part design and analysis. Capital expenditures for

computer-driven, automated equipment can be significant.

Although ATL generally is faster than AFP and can place more material

over longer distances, AFP is better suited to shorter courses and can place

material more effectively over contoured surfaces.

3.14 Centrifugal casting

Pipe from 1 inch/25 mm to 14 inches/356 mm in diameter is an

alternative to filament winding for high-performance, corrosion-resistant

service. In cast pipe, 0°/90° woven fiberglass provides both longitudinal and

hoop strength throughout the pipe wall and brings greater strength at equal wall

thickness compared to multiaxial fiberglass wound pipe.

80

In the casting process, epoxy or vinyl ester resin is injected into a 150G

centrifugally spinning mold, permeating the woven fabric wrapped around the

mold’s interior surface. The centrifugal force pushes the resin through the layers

of fabric, creating a smooth finish on the outside of the pipe, and excess resin

pumped into the mold creates a resin-rich, corrosion- and abrasion-resistant

interior liner.

Fiber-reinforced thermoplastic components now can be produced by

extrusion, as well. Breakthrough material and process technology has been

developed with long-fiber glass-reinforced thermoplastic (ABS, PVC or

polypropylene) composites to provide profiles that offer a tough, low-cost

alternative to wood, metal and injection-molded plastic parts used in office

furniture, appliances, semitrailers and sporting goods. A huge market has

emerged in the past decade for extruded thermoplastic/wood flour (or other

additives, such as bast fibers or fly ash) composites. These wood plastic

composites, or WPCs, used to simulate wood decking, siding, window and door

frames, and fencing.

81

CHAPTER-4

TESTING

4.1 WIND TUNNEL

The "Wind tunnel" is a facility, by artificially producing airflow relative

to a stationary body, that measures aerodynamic force and pressure distribution

to simulate the actual flight of airplane or orbiting plane in the air.

4.2 TYPES:

Wind tunnels are often denoted by the speed in the test section relative to

the speed of sound. The ratio of the air speed to the speed of sound is called the

Mach number.

Tunnels are classified as

• Subsonic (M < 0.8),

• Transonic (0.8 < M < 1.2) ,

• Supersonic (1.2 < M < 5.0) , or

• Hypersonic (M > 5.0).

82

OPEN CIRCUIT SUBSONIC WIND TUNNEL:

Fig 4.1 Open Circuit Wind Tunnel

4.3 Honey comb:

Honey comb along with the wire mesh protects the wind tunnel from

foreign objects. It also provides laminar flow for the wind tunnel test section.

4.4 Effuser:

It converts the available pressure energy to kinetic energy which is

located upstream of the test.

4.5 Test section:

The models to be tested are placed inside the test section by means of

supports and balances. The instruments necessary for recording the data are also

fixed in the wind tunnel.

83

4.6 Diffuser:

Diffuser is locates at the downstream of the test section, it converts the

kinetic energy to pressure energy.

4.7 Propeller driving unit:

A fan or a propeller is fitted with electric motor to drive airflow to the test

section.

4.8 Measurement of aerodynamic forces

Ways that air velocity and pressures are measured in wind tunnels:

Air velocity through the test section (called the throat) is determined

by Bernoulli's principle. Measurement of the dynamic pressure, the static

pressure, and (for compressible flow only) the temperature rise in the airflow

Direction of airflow around a model can be determined by tufts of yarn

attached to the aerodynamic surfaces

Direction of airflow approaching an aerodynamic surface can be visualized

by mounting threads in the airflow ahead of and aft of the test model

Dye, smoke, or bubbles of liquid can be introduced into the airflow upstream

of the test model, and their path around the model can be photographed

4.9 Force and moment measurements:

With the model mounted on a force balance, one can measure lift, drag, lateral

forces, yaw, roll, and pitching moments over a range of angle of attack. This

allows one to produce common curves such as lift coefficient versus angle of

attack.

84

The force balance itself creates drag and potential turbulence that will

affect the model and introduce errors into the measurements. The supporting

structures are therefore typically smoothly shaped to minimize turbulence.

4.10 Flow visualization:

In general, flow visualization is an experimental means of

examining the flow pattern around a body or over its surface. The flow is

"visualized" by introducing Yarn Tufts, smoke or pigment to the flow in the

area under investigation. The primary advantage of such a method is the ability

to provide a description of a flow over a model without complicated data

reduction and analysis. Smoke flow visualization involves the injection of

streams of vapor into the flow. The vapor follows filament lines (lines made up

of all the fluid particles passing through the injection point). In steady flow the

filament lines are identical to streamlines (lines everywhere tangent to the

velocity vector). Flow visualization can thus reveal the entire flow pattern

around a body.

4.11 TUFT WANDS:

The least expensive method for flow visualization is a tuft wand. This method is

very much versatile and at the same time the flow pattern around the test object

is visible. A long tuft on a pole is useful for tracking the flow near the object.

Flow visualization foe the moment is possible if the trace particles location can

be identified at any time in the flow field.

85

4.12 MODEL TESTING IN WIND TUNNEL

The wind tunnel is calibrated initially. The model is mounted in the wind

tunnel force balance with the help of a strut fixed at its center of gravity. After

ensuring that all the connections are proper the tunnel is started with an initial

velocity. The velocity is increased gradually; the lift and drag values are noted

simultaneously for corresponding velocities. The model is tested with landing

gear and then without the landing gear. In order to fix a retractable landing gear

mechanism we have proposed another wing with a thicker airfoil. The model

with a newly proposed wing is tested in the wind tunnel and the corresponding

values are noted. From the tabulations it is observed that the drag in the airplane

is reduced to a certain percentage without the landing gear. The flow over the

wings is observed in all the three cases by tuft flow visualization technique.

4.13 DIFFICULTIES FACED DURING TESTING

The propeller in the airplane did not run during the testing due to its

misalignment during fabrication. We used a white tape to tighten and hence we

could rectify the problem. The strut fixed to the airplane was slightly improper

causing certain vibrations; hence we welded the strut to a plate and then fixed

the model.

86

CHAPTER-5

WING INSTALLATION

5.1 Objective

the procedure for installing left and right wing is the same with the expectation

that some of the parts are handled.where the parts are handled the part number

for the left hand wing is called up first, and then the parts for the right wing

follows it in bracket.

Although not absolutely necessary, it is recommended that you to manufacture

two wing supports so that wing assembly may be supported while positioning

and riveting the wing skin to frame

Total assembly for each wing is approximately 30 hours.

5.2 Installation of wing components

Mark to position of pitot tube doubler 880 on bottom wing skin 690-3l.

Drill 3 holes 4.5mm, and cut out.

Install pitot tube 87 and doubler 880 with screws AN526-838RB,

washers AN960-8 and nuts MS20365-832.

Install the pitot line 881 through grommets AN931-6-10 in the wing and

connect it to the pitot tube 878 with 882-2.

Install navigation light cable 1600-L2J1 and strobe light cable 1600-

L4E16 through grommets AN931-6-10 in the centre ribs of the wing.

Tie the pitot line 881-11 together with the to electrical ables, clip them to

the cable tie base 229-560,with tie warps MS3347-5.

Install and tighten the fitting AN832-3D and nut AN924-3D through

doubler 691-10.

87

Fig 5.1 Rear beam assemby 699L(699R)

88

Fig 5.2 AFT RIB ASSY 698-7L(698-7R)

89

5.3 Assemble the wing tips

Draw a line .34‖ from the edge of the wing tip. Locate the wing tip

assembly 925L(925R) on the end of the wing, and use a hole finder to

locate screw holes in end rib. Make sure ring tips is positioned far enough

and the wing so that the holes a line that you have drawn and drill holes.

Posiiton paper template on the outboard edge of the wing tip 925-L.

Drill a 1.25‖hole with a holes saw, and then file the hole to 1.38‖to

hallow strobe light to fit through.

Put the strobe light through the hole. Mark and then drill three hole no.12

and install rivetus A6K75

Fit the stroke and position light base plate to the wing tip 925-L with

screws.

Fit the strobe and position light assembly 156/003 to the base plate,with

screw.

Connect the plug to the iring socket, and attach wing tip assembly 925-L

to wing assembly, with screws 8Z*1/2‖.

5.4 WING ASSEMBLY

We chose to assemble the wing using the sawhorse method. It goes something

like this:

Level both sawhorses so that they are level and parallel.

Lay main spar across.

Install ribs.

Install rear spar.

Verify vertical and horizontal alignment of ribs using "plumb bob"

method.

90

Attach aft top skin to rear spar using 3/32" cleco's.

Align ribs to vertical axis and secure aft top skin to ribs using 1/8"

cleco's.

After this aft top skin is clecoed into its location you have essentially locked in

the alignment of the wing assembly.

Start by aligning the centerline mark of the rib through the 3/32" holes in the

spar web. It is not critical at this point to have the rib aligned top to bottom we

will take care of that in a minute.The wingtip skin (white part) is composite

with coordination holes drilled using a fixture. Ribs and spars (black parts) join

the skin via these holes, which act as mates. Additional fastener holes are added

later and act as contacts. In conventional aircraft assembly, these parts would be

positioned relative to each other by another fixture; fastener holes would be

drilled through all the parts, using this fixture as a guide, while the parts were

clamped, and fasteners would be put through these holes.

Fig 5.3 Drilling holes

91

Proper design of a new process also includes adequate understanding of

the process it is intended to replace, so that the new one will do everything

necessary that the original one did, only better.

5.5 PROCEDURE:

Now align the edge of the web spacers to our marks and center them

vertically

Clamp rib into position making sure that the rib is aligned for a smooth

transition of the spar cap to the rib flange as described

Turn the spar assembly over and drill all holes, insert 3/32" Cleco in #2

and #5 holes

Do this for all rear ribs not including the root rib. Check all of your

alignments now. After this there is no more adjustment.

When attaching the rear spar to the ribs I used two pieces of thick scrap

material and clamped it to the ribs. This in effect clamped the rear spar

into location for drilling.

Fig 5.4 Rear spar attachment

This picture shows how the bow in the main spar affects the shape of the

rear spar attachment. No worries here because when you attached the aft skin

things will straighten out easily.

92

It is alright to drill the rear spar mounting holes to final size for the

rivets. After the rear spar is fitted remove all of the ribs and the rear spar. De-

burr all holes, including those in the front of the rear ribs.

Now using the same technique drill and fit all the front ribs to the spar

using only the 3/32" cleco's making sure there is a smooth transition from spar

caps to rib flanges as shown in Detail K on page W-14. Remove and de-burr.

Re-attach the rear ribs and rear spar, again only putting a 3/32" cleco into

holes #2 and #5. Attach the front ribs by first relaxing the cleco in hole #2 and

inserting the front rib. Rotate the rib into position over the cleco in position #5

and again relax the cleco allowing the rib to clamp into position. Repeat for all

ribs.

5.5 Critical point

Using two very good and sturdy sawhorses lay the spar across them with

the rear ribs pointing vertically. Position the first sawhorse on the root section

of the main spar and the second in the bay between ribs 9 and 10.

Using the wire plumb bob method align the two 1/4" alignment holes in

the rear ribs. If you can get the majority of the ribs positioned so that the wire

will hang straight from the top hole to the bottom hole you are good. Clamp the

spar into position from the back or bottom side. Using a good quality square,

align the rear ribs to be square with the spar. Use a straight piece of aluminum

from the root rib to the main spar to clamp into place.

With the spar and ribs aligned you can now hang the top aft skin into

position. Align the skin to the rear spar and using several clamps secure

it. Make sure you have the rear edge of the skin flush to the rear edge of the

rear spar. Using 3/32" cleco's drill and secure into position

93

0.00

0.10

0.20

0.30

0.40

0.50

0.60

0.70

0.80

0.90

1.00

0 1 2 3 4 5 6 7 8

LIFT vs DRAG

CHAPTER-6

RESULTS

6.1OBSERVATIONS

WING 1: WITH PROPELLER AND LANDING GEAR

Table 6.1

LIFT(N) DRAG(N) VELOCITY(m/sec) L/D

1.5 0.2 5 7.5

2.8 0.3 10 9.33

3.2 0.4 15 8

4.4 0.5 20 8.8

5.8 0.7 25 8.28

6.6 0.8 30 8.25

7.2 0.9 35 8

94

WING 1: WITHOUT LANDING GEAR

Table 6.2

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0 1 2 3 4 5 6 7 8 9

LIFT vs DRAG

LIFT(N) DRAG(N) VELOCITY(m/sec) L/D

1.6 0.1 5 16

2.9 0.2 10 14.5

4.1 0.3 15 13.6

5.5 0.4 20 13.75

6.8 0.5 25 13.6

7.3 0.6 30 12.16

8.2 0.7 35 11.71

95

WING 2: WITHOUT LANDING GEAR

LIFT(N) DRAG(N) VELOCITY(m/sec) L/D

1.7 0.1 5 17

3 0.2 10 15

4.3 0.3 15 14.23

5.6 0.11 20 14

7.1 0.5 25 14.2

9.4 0.7 30 13.42

11.6 0.8 35 14.5

Table 6.3

Graph 6.3

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

0 2 4 6 8 10 12 14

LIFT vs DRAG

96

6.2DRAG DIFFERENCE:

Graph 6.4

00.10.20.30.40.50.60.70.80.9

1

Wing 1 withlanding gear

Wing 2Withoutlanding gear

97

6.3 Material specification

6.3.1Carbon fabric construction data-Hexcel fibers

Style Weave Count warp

Count fill Warp yarn

Fill yarn Fabric weight (g/m2) GSM

F3A282 Plain 11.5 11.5 AS4GP 3K

AS4GP 3K

5.70 193

F3A286 4H Satin 11.5 11.5 AS4GP 3K

AS4GP 3K

5.64 191

F3B262(GP) Plain 8 8 AS4CGP 3K

AS4CGP 3K

4.72 160

F3B262(J) Plain 8 8 AS4CJ 3K AS4CJ 3K 4.72 160

F3B282(J) Plain 12.5 12.5 AS4CGP 3K

AS4CGP 3K

5.72 194

F3B284(GP) 2/2 twill 12.5 12.5 AS4CJ3K AS4CJ3K 5.72 194

F3B286(GP) 2/2 twill 12.5 12.5 AS4CGP 3K

AS4CGP 3K

5.76 195

F3B284(J) 4H Satin 12.5 12.5 AS4CJ3K AS4CJ3K 5.76 195

F3B286(GP) 4H Satin 12.5 12.5 AS4CGP 3K

AS4CGP 3K

5.70 193

F3B286(J) Plain 12.5 12.5 AS4CGP 3K

AS4CGP 3K

5.70 193

F4M282 Plain 12.5 12.5 IM7GP 6K

IM7GP 6K

5.80 197

F4M466 5H Satin 16 16 IM7GP 6K

IM7GP 6K

8.40 285

Table 6.4 Carbon fabric construction data-Hexcel fibers

98

6.3.2 PAN Carbon fibers data

Produce

r

Fiber

name

Availability Tensile

Strength

(Ksi)

Tensile

Modul

us

(msi)

Elongatio

n

(%)

GSM

H excel AS2C 3K 644 32.0 1.90 1.80

AS4 3K,6K,12K 647/626/6

49

33.5 1.80 1.78/1.7

9

AS4C 3K,6K,12K 647/626/6

34

33.5 1.80 1.78

AS4D 12K 689 33.5 1.80 1.79

AS7 12K 700 35.0 1.80 1.79

IM6 12K 833 40.5 1.90 1.76

IM7 6K,12K 770/822 40.0 1.80/1.90 1.78

IM8 12K 885 44.0 1.80 1.78

IM9 12K 890 44.0 1.90 1.80

IM10 1K,3K,6K,12

K

1010 44.0 2.10 1.79

Cytec T300 3K,6K,12K 545 33.5 1.60 1.76

T650/3

5

12K 620 37.0 1.70 1.76

Table 6.5 PAN Carbon fibers data

99

Fig 6.1 cross sectional view of carbon fiber wing

100

fig 6.2 Ribs and spar

101

Conclusion

From this we have implemented a new methodology for implementing

carbon fiber materials in light sport aircraft. This will increase the elasticity

and high withstanding temperature.This will attain the pilot to reach the

maximum G-level.

Thus it may be concluded that if the Thorpedo T211 aircraft is provided with

provisions for retractable landing gear, drag reduction occurs. The reduction

would directly affect the fuel consumption, carbon emission and the range of

aircraft. Fuel consumption will be reduced which would help to improve the

range. CO2 emissions are also reduced thus good for environment

102

REFERENCE

Brady, George S., Henry R. Clauser, and John A. Vaccari. Materials Handbook.

McGraw-Hill, 1997.

Kroschwitz, Jacqueline I. and Mary Howe-Grant, ed. Encyclopedia of Chemical

Technology. John Wiley and Sons, Inc., 1993.

Ebbesen, T.W. "Carbon Nanotubes." Physics Today (June 1996): 26-32.

American Carbon Society website. http://www.ems.psu.edulcarbon .

Carbon Composites website. http://www.carb.com .

— Chris Cavette

Investigations for Mechanical Properties of Metal Matrix Composite Prepared

by Combining FDM, Vacuum Moulding and Stir Casting

-Rupinder Singh, Sunpreet Singh and Sardar Singh

Hysteresis Heating of Polypropylene Based Composites

-Ravi Shukla, JohnneyMertens and S Senthilvelan

Frictional Heat Generation in Selective Ceramic Reinforced Polymer

Composites - Effect of Particle Size

-

C. Gurunathan, R. Gnanamoorthy and S. Jayavel