Ac Design Project 2

8/10/2019 Ac Design Project 2

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AIRCRAFT DESIGN PROJECTII

(MULTIROLE FIGHTER AIRCRAFT)

A PROJECT REPORT

Submitted by

Batch - 18

A. ARIF ABDUL RAHMAN 11908101701

P. SANTHOSH 11908101704

T.V.S SARAVANAN 11908101705

I n partial ful fi llment for the award of the degree

Of

BACHELOR OF ENGINEERING

IN

AERONAUTICAL ENGINEERING

VEL TECH (ENGINEERING COLLEGE)

ANNA UNIVERSITY: CHENNAI 600 025

OCTOBER 2011


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ANNA UNIVERSITY: CHENNAI 600 025

BONAFIDE CERTIFICATE

Certified that this project report titledAIRCRAFTDESIGN PROJECTII(MULTIROLE FIGHTER

AIRCRAFT) is the bonafide work of

A. ARIF ABDUL RAHMAN 11908101701

P. SANTHOSH 11908101704

T.V.S SARAVANAN 11908101705

who carried out the work under my supervision.

SIGNATURE SIGNATURE

Mr. G. BOOPATHY M.E., Mr. M. Ramakrishna M.E.,

HEAD OF THE DEPARTMENT INTERNAL GUIDEDept of Aeronautical Engg. Dept of Aeronautical Engg.

Vel Tech Engg. college Vel Tech Engg. college

No.42 , Avadi - Vel Tech Road, No.42 , Avadi - Vel Tech Road,

Chennai-62 Chennai-62


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CERTIFICATE FOR EVALUATION

College Name : 119 - VEL TECH ENGG COLLEGE

Branch : AERONAUTICAL ENGG.

Semester : VII

The reports of the project work submitted by the above students in

partial fulfillment for the award of Bachelor of Engineering degree In

Aeronautical Engineering of Anna University were evaluated and confirmed

to be the reports of the work done by the above students and then evaluated.

INTERNAL EXAMINER EXTERNAL EXAMINER

S.NO

Name of the Students

Who have done

the project

Title of the project

Name of the

Supervisor with

Designation

1.A. ARIF ABDUL

RAHMANMULTIROLE

FIGHTER

AIRCRAFT

Mr. M. Ramakrishna

M.E.,

Asst. Prof.,

Internal guide

2. P. SANTHOSH

3. T.V.S SARAVANAN


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ACKNOWLEDGEMENT

This project, though done by us would not have been possible, without thesupport of various people, who by their cooperation have helped us in

bringing out this project successfully.

We are grateful to our Chancellor, Dr. R. Rangarajan B.E (Elec), (Mech),

MS (Auto) for his patronage towards our project.

We thank our Principal I/c, Mr. E. Kamalanaban M.E., (Ph.D),who had

always served as an inspiration for us to perform well. We would like to

express our faithful thanks to our head of the department, Mr. G. Boopathy

M.E., for having extended all the department facilities without slightest

hesitation.

We would like to express our unbounded gratefulness to our internal guideand project incharge, Mr. M. Ramakrishna M.E., Asst. Prof., Dept of

Aeronautical Engg. For his extremely valuable guidance and encouragement

throughout the project.

We thank all faculty members and supporting staff for the help they

extended to us for the completion of this project.


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TABLE OF CONTENTS

1.0 Introduction to ADP-II

1.1 Three view drawing

1.2

Data from ADP-I

2.0 V-n diagram for design study

2.1 V-n diagram (Take off, cruise, landing)

3.0 Gust and maneuverability envelopes

3.1 V-n diagram (gust loads)

3.2 Final V-n graph calculation

4.0 Structural design studyTheory approach

5.0 Load estimation of wings

5.1 Force and moment calculation

5.2 Shear force diagram

5.3 Bending moment diagram

5.4

Spar location

6.0 Fuselage Design

6.1

Bulkhead design

6.2

Longeron calculation

6.3

Stringer calculation

7.0 Material Selection

8.0 3 D view of aircraft

9.0 Summary

10.0 References


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1. INTRODUCTION

The structural design of an airplane actually begins with the flight envelope or the

V-n diagram, which clearly limits the maximum load factors that the airplane can

withstand at any particular flight velocity. But, in normal practice the airplane might

experience loads that are much higher than the design loads. Some of the factors that lead

to structural overload of an airplane are high just velocities, sudden movement of the

controls, fatigue loads and in some cases, bird strikes or lightening strikes. So, to add

some inherent ability to withstand these rare but large loads, safety factor of 1.5 is

provided during the structural design.

The two major members than need to be considered for the structural design of an

airplane are the wings and the fuselage. As for as the wing design is concerned, the most

significant load is the bending load. So, the primary load bearing component in the wing

structure is the spar (the front and the rear spars) whose cross section is an I- section.

Apart from the spar to take the bending loads, suitable stringers need to be provided to

take shear loads acting on the wing.

Unlike the wing, which is subjected mainly to unsymmetrical bending load, thefuselage is much simpler for structural analysis due to its symmetrical cross section as

well as symmetrical loading. The main load in case of fuselage is only shear because the

load acting on the wing is transferred to the fuselage skin in the form of shear only. The

structural design of both the wing and the fuselage begin with shear force and bending

moment diagrams for the respective members. The maximum bending stress produced in

each of them is checked to be less than yield stress of the material chosen for the

respective member.


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1.1 THREE VIEW DIAGRAM FROM ADPI

FINAL SCALED PRELIMINARY DESIGN


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1.2IMPORTANT DATA FROM ADP1

AIRCRAFT INITIAL CONSIDERATIONS:

Mach 1.87(max)

Range 3200km

Cruise mach 1.14

Cruise altitude 16500m

W payload 5545kg

W take-off 23500 kg

Power plants F-135 P&W afterburning turbofan

AIRFOIL SELECTION:

NACA 64(3)-618 AIRFOIL

Thickness: 17.9%

Camber: 3.4%

Trailing edge angle: 12.4

o

Lower flatness: 35.9%

Leading edge radius: 2.4%

Max CL: 1.536

Max CLangle: 15.0

Max L/D: 38.202

Max L/D angle: 7.0

Max L/D CL: 1.331

Stall angle: -0.5

Zero-lift angle: -4.5


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Lift for NACA 64(3)-618 AIRFOIL


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2. V-n DIAGRAM

2.1 Introduction

The V-n diagram plays an important role in Aircraft design. The V-n diagram is a plot

between the load factor and the velocity. Load factor is defined as the ratio of the

aerodynamic load to the weight of the aircraft. Aircraft has to perform different loading

conditions at different speeds, controls and high loads due to stormy weather. But at the

same time, it is impossible to investigate all possible loading conditions. There are

structural limitations on the maximum load factor allowed for a given airplane. There are

two categories of structural limitations in airplane design:

1. Limit L oad Factor

This is the boundary associated with permanent structural deformation of one or more

parts of the airplane. If n is less than the limit load factor, the structure may deflect during

maneuver, but it will return to its original state when n = 1. If n is greater than the limit

load factor, then the airplane structure will experience a permanent deformation, i.e., it

will incurstructural damage.

2. Ul timate Load Factor

This is the boundary associated with outright structural failure. If n is greater than the

ultimate load factor, parts of the airplane willbreak.There arefour main critical conditions:

High Angle of Attack (+)

Low Angle of Attack (+)

Low Angle of Attack (-)

High Angle of Attack (-)

For airplane design, the limit load factor depends on the type of the aircraft.


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A typical V-n diagram looks like this:

2.2 Calculation of v-n diagram :

2.2.1 Takeoff, cruise and landing:

For positive curve,

From CLvs graph,

CL=1.56 and = 14

CL=0 and = -2

Lift slope for airfoil,a0= dCL / d

= 1.560 / 14(- 2)

a0=0.0975

Lift slope for wing, a = a0 / 1 + (64.72 a0/ e AR)


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= 0.0975 / 1+ (64.72(0.0975) / (0.8)(3.72))

a = 0.0658

co-efficient of lift , CL= a ( CLmax - CL = 0)

= 0.0658 ( 14( - 2 ) )

CL= 1.052

n = L/W = 1\2( V2S CL) / W

n = 1.624310-3V2

similarly for negative curve,

From CLvs graph,

CL= - 0.1 and = - 6

CL= - 0.4 and = -12

Lift slope for airfoil,a0= dCL / d

a0= 0.025

Lift slope for wing, a = a0 / 1 + (64.72 a0/ e AR)

a = 0.021311

co-efficient of lift , CL= a ( CLmax - CL = 0)

CL= - 0.1278

n = L/W = 1\2( V2S CL) / W

n = - 2.0290410-4

V2


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-5

-4

-3

-2

-1

0

1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

0 25 50 75 100 125 150 175 200 225 250 275 300 325 350 375 400 425 450

Load

Factorn

Velocity (m)

V-n Diagram

vA

V

A'

Vc VD

Vc'

VD'

Stall area

Stall

area

Flight velocity (V) (m/s) npositive nnegative

0 0 0

10 0.162 0.09

20 0.64 0.36

30 1.46 0.81

40 4.06 1.44

50 6.80 2.25

64.72 7.95 3.03

388.88 9 3.03


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3.GUST ENVELOPE

TO FIND GUST LOAD FACTOR:

Gust loads are unsteady aerodynamic loads that are produced by atmospheric

turbulence. They represent a load factor that is added to the aerodynamic loads.

The effect of turbulence is to produce a short time change in the effective angle of

attack. This change can be either positive or negative, thereby producing an increase or

decrease in the wing lift and change in the load factor,=

.

The incremental load factor is then (from Design of aircraft by

Thomas.C.Corke,n =

u V CL2 W

S

(i)

Where,

n- Change in the load factor, -density at sea level (1.2256 kg/m3),u- Gust velocity

(m/s)V- Flight velocity (m/s), CL co efficient of lift, at max = 141.56 at ( =

0) =0.4,

Then,

npeak =nlevel flight+ n.. (ii)

The gusts that result from atmospheric turbulence occur in a fairly large band of

frequencies. Therefore, their effect on an aircraft depends on factors that affect its

frequency response. In particular, the frequency response is governed by an equivalent

mass ratio, ,defined as (from Design of aircraft by Thomas.C.Corke,

=

2WS

g c CL

Where,

-density at sea level (1.2256 kg/m3),

C L max = 1.56 (at flap deflection =14),

C L max = 0.4 (at flap deflection =0),

c- Average cord from wing design (12.815 m)


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the mass ratio , is a parameter in a response coefficient, K, which is defined differently

for supersonic,

K=1.03

6.95+1.03

The normal component of gust velocity, u, is the product of the normal average of values

taken from flight data, (u), and the response coefficient is (from Design of aircraft by

Thomas.C.Corke,

u=K u

GUST VELOCITY

At altitude of below 20000 ft the gust velocity is (from Design of aircraft by

Thomas.C.Corke,

Flight condition Altitude (m)Gust velocity ()

(m/)

K u (m/s)

High angle of

attack


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LOAD FACTOR

Gust envelop is,

-8

-7

-6-5

-4

-3

-2

-1

0

1

2

3

4

5

67

8

9

10

11

12

13

14

15

0 25 50 75 100 125 150 175 200 225 250 275 300 325 350 375 400 425 450

Load

Factorn

Velocity m

Gust EnvelopeVg

Vcruise

Vdive

Vg'V

cruise

'Vdive'

Dive condition 415.707 2.668628 11.66863

Flight condition Flight velocity (v) (m/), -

High angle of attack 112.09 1.07398 4.07398

Level flight 388.88 0.47341 3.47341

Dive condition 415.707 0.21677 3.21677


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The final V-n diagram is a combination of gust loads, maneuvering load factors

with the ultimate load factor. The ultimate load factor of any fighter aircraft is given by

multiplying the limit load factor with factor of safety. The final V-n diagram which

shows the possible operating region, limit load factor, ultimate load factor is shown

below the plot.

-8-7-6-5-4-3-2-101234

56789

10111213141516

0 25 50 75 100 125 150 175 200 225 250 275 300 325 350 375 400 425 450

Load

Fa

ctorn

Velocity m

Gust Envelope

vA

VA'

Vc VD

V

c'

VD'

Stall area

Stall areaVg'

Vg

Vcruise

Vdive

Vcruise'Vdive'

Flight velocity (V) (m/), Load factor (n) Load factor (n)

0 1 1

112.09 - 4.073

194.16 13.645 -

388.88 14.81 3.47

415.7 11.66 3.21

415.7 0 0


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Pulldown Maneuver

The pulldown maneuver is the inverted form of the pullup maneuver.

In this case, both lift and weight are contributing to the pitch rate - as a result, simply

rolling an airplane inverted will initiate this maneuver.

By convention, we will still consider lift and thus load factor positive as sketched.

Since the lift and weight are now in the same direction, the downward acting centripetal

forces is:

F =L +W =W(n +1) r

And the flight path radius of curvature and pitch down rate are given by:

=

2

=

2

(+1) =

=

(+1)

Note that as long as n > -1, R and are positive - for n=-1 the plane is in level flight;

for n < -1 the plane begins climbing in a pushup maneuver.

CACULATION OF PULL-UP AND PULL-DOWN V-n DIAGRAM

PULL-UP

The turn radius is given by: =

2

(1)

For the corner speed =64.72 m/s and limit loar factor n=10

=64.722

9.81(101)

R=53.37 m


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V in m n R

30 1.46187 198.8368

35 1.989768 126.2923

40 2.59888 102.1123

45 3.289208 90.26384

50 4.06075 83.34625

55 4.913508 78.87387

60 5.84748 75.78101

PULL-DOWN:

The turn radius is given by: =

2

(+1)

For the corner speed =64.72 m/s and limit loar factor n=10

=64.722

9.81(10+1)

R= 42.69m

V in m n R

30 1.46187 37.30365

35 1.989768 41.80927

40 2.59888 45.36559

45 3.289208 48.17502

50 4.06075 50.40795

55 4.913508 52.19803

60 5.84748 53.64703


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4. Structural design studyTheory approach

The structural design study of any aircraft includes the following

Wing design

The wing design and calculation of the lift distribution.

Drawing the bending moment and shear flow diagram.

Design of spars.

Fuselage design

Design of bulk heads.

Design of longerons.

Structural design criteria

The structural criteria define the types of maneuvers, speed, useful loads, and

gross weights which are to be considered for structural design analysis. These are items

which are under the control of the airplane operator. In addition, the structural criteria

must consider such items as inadvertent maneuvers, effects of turbulent air, and severity

of ground contact during landing. The basic structural design criteria, from which the

loadings are determined, are based largely on the type of the airplane and its intended

use.


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5. Load estimation of wings

Lift distribution over a delta wing

Lift distribution in a delta wing is a complex phenomenon involving vortex lift

and lift due to flow over the wing. In supersonic flow elliptic wing does not perform

optimally due to shock waves. So delta wings are used and so lift produced are in

accordance with the planform type, hence lift distribution in delta wing is linear without

vortex , but with vortex, lift becomes non-linear.

Lift distribution analysis

For our analysis ,the lift produced is assumed to be linear, this is reasonable for

analysis because vortex lift is upto 20% of total lift.

This can be given approximately by;

L(x)clc(y)

Where, x along span

y along chord

LINEAR LIFT DISTRIBUTION


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SHEAR FORCE AND BENDING MOMENT ANALYSIS OF WING

The more difficult loading is the linear distribution loading due to the lift on the

wing. The shear force distribution on the wing can be calculated by integrating the

loading function, the lift distribution, from the left free end to any point, x, on the wing,

The integration constant, C1, can be determined by using the boundary condition

that the shear force on the wing tip (x = -b) must be zero.

The calculated shear force values are,

X with Respect to

Wing rootShear force (KN)

1.8 108.4689

3.6 86.29647

4.95 71.329

5.4 66.65639

6.3 57.78599

7.2 49.54869

8.1 41.94449

9 34.97


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The shear force graph is plotted as,

Bending-Moments

The bending-moment is the force at each location on the spar that bends the wing

upward during normal non-inverted flight. The bending-moment is zero at the wing-tip

and maximum at the root. But its value is not proportional across the span. In other

words, it is not half as much at the wing mid-point as it is at the root. In fact, the mid-

point bending-moment is only about a 1/4 of the root value.

Thus, Similarly, the bending moment can be found by integrating the shear force

distribution, as

The bending moment force on the wing tip must be zero which is used to

determine the integration constant. The moment function become

0

20

40

60

80

100

120

140

-10 -5 0 5 10

Shear Force Diagram For Delta Wing

Wing Span m

SF ( KN)


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The calculated bending moment values are,

X with Respect to

Wing rootBending moment

(Nm)

1 563.3

2 426.24

3 287.46

4 62.35

5 26.75

6 12.68

7 4.68

8 0.327

The bending moment graph is plotted as,

0

100

200

300

400

500

600

700

-10 -5 0 5 10

Bending Moment Diagram for Delta

wing

BM (Nm)


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6. FUSELAGE DESIGN

Theory

The fuselage includes the cabin and/or cockpit, which contains seats for the occupants

and the controls for the airplane. In addition, the fuselage may also provide room for

cargo and attachment points for the other major airplane components. Some aircraft

utilize an open truss structure. The truss-type fuselage is constructed of steel or aluminum

tubing. Strength and rigidity is achieved by welding the tubing together into a series of

triangular shapes, called trusses.

Construction of the Warren truss features longerons, as well as diagonal and

vertical web members. To reduce weight, small airplanes generally utilize aluminum

alloy tubing, which may be riveted or bolted into one piece with cross-bracing members.

As technology progressed, aircraft designers began to enclose the truss members to

streamline the airplane and improve performance. This was originally accomplished with

cloth fabric, which eventually gave way to lightweight metals such as aluminum. In some

cases, the outside skin can support all or a major portion of the flight loads. Most modern

aircraft use a form of this stressed skin structure known as monocoque or

semimonocoque construction.


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The monocoque design uses stressed skin to support almost all imposed loads.

This structure can be very strong but cannot tolerate dents or deformation of the surface.

This characteristic is easily demonstrated by a thin aluminum beverage can. You can

exert considerable force to the ends of the can without causing any damage.

Partially Completed Structural Layout

Since no bracing members are present, the skin must be strong enough to keep the

fuselage rigid. Thus, a significant problem involved in monocoque construction is

maintaining enough strength while keeping the weight within allowable limits. Due to the

limitations of the monocoque design, a semi-monocoque structure is used on many of

todays aircraft.

The semi-monocoque system uses a substructure to which the airplanes skin is attached.

The substructure, which consists of bulkheads and/or formers of various sizes and

stringers, reinforces the stressed skin by taking some of the bending stress from the

fuselage.


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BULKHEADS

Any major vertical structural member of a semimonocoque fuselage, hull, or

float may be considered a bulkhead. Bulkheads serve to maintain the required external

contour at the station.Rib repair by patching. Where they are located. They also

give rigidity and strength to the structure. Bulkhead construction is similar to that used

for wing ribs. It consists of a web reinforced by angle stiffeners. The web is attached

to the skin by formed flanges or extruded angles, which serve as cap strips. Non-

watertight bulkheads may have lightening holes, and most bulkheads are cut out to

give clearance for stringers. The stringers are usually attached to the bulkhead by angle

clips. For our fighter,

The forward fuselage is 5.2m long and 1.7m wide.

The canopy is about 356 cm long and 114 cm wide.

The mid fuselage in 5.2 m long and 2m wide.

The aft fuselage is 5.8m long and 3.6m wide.


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LONGERONS

Most aircraft fuselages are constructed in sections and are of the

semimonocoque design. A longeron is a fore-and-aft member of the fuselage or nacelle

and is usually continuous across a number of points of support, such as frames and

bulkheads. The longerons, along with the stringers, are the major load-carrying members

and stiffeners.

The cross section we have chosen is I section, the no of longerons is given as

[(Max dia of fuselage)/spacings of longerons] = no of longerons

For our fighter,

Max dia =2 m

Spacing of longerons (referred to Jon Roskam) = 10 inch (.254m)

[2/.254]=7.87

Hence No. of longerons =8 (app).


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7. WING DESIGN

Wings develop the major portion of the lift of a heavier-than-air aircraft.Wing

structures carry some of the heavier loads found in the aircraft structure. The particular

design of a wing depends on many factors, such as the size, weight, speed, rate of climb,

and use of the aircraft. The wing must be constructed so that it holds its aerodynamics

shape under the extreme stresses of combat maneuvers or wing loading. Wing

construction is similar in most modern aircraft. In its simplest form, the wing is a

framework made up of spars and ribs and covered with metal. The construction of an

aircraft wing is shown in fig.Spars are the main structural members of the wing. They

extend from the fuselage to the tip of the wing. All the load carried by the wing is taken

up by the spars. The spars are designed to have great bending strength. Ribs give the

wing section its shape, and they transmit the air load from the wing covering to the spars.

Ribs extend from the leading edge to the trailing edge of the wing. In addition to the main

spars, some wings have a false spar to support the ailerons and flaps. Most aircraft wings

have a removable tip, which streamlines the outer end of the wing. Most Navy aircraft are

designed with a wing referred to as a wet wing. This term describes the wing that is

constructed so it can be used as a fuel cell. The wet wing is sealed with a fuel-resistant

compound as it is built. The wing holds fuel without the usual rubber cells or tanks. The

wings of most naval aircraft are of all metal, full cantilever construction. Often, they maybe folded for carrier use. A full cantilever wing structure is very strong. The wing can be

fastened to the fuselage without the use of external bracing, such as wires or struts. A

complete wing assembly consists of the surface providing lift for the support of the

aircraft. It also provides the necessary flight control surfaces.

Strength of Wings

Uniformly built spars, having the same structure and dimensions from the root to

even just halfway to the wing-tip, add weight with little benefit in strength, and thisweight is closer to the wing-tips which inhibit rolls in both fighter planes, as well as

sailplanes. While spar design may be an interesting engineering challenge. Uniformly

built spars are simple to construct.


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Maximum Wing-Load

First let's make it clear that maximum wing-load is not the same as wing-loading.

The maximum wing-load is the force when the plane is performing a maximum-G

manoeuvre. Fighter planes may experience maximum-G during a high speed turn, or pull-

up from a high speed dive. Figure shows the load forces in banked turns.

Figure G-load vs Bank Angle

Figure-7.3 shows the maximum load forces at the bottom of a verticle loop. At the top of

the loop, gravity is aiding the plane in pulling toward the center of the loop, while at the

bottom, it is pulling it out of the loop. The figure shows how the loads increase with

speed, and with smaller diameter loops. Even if the plane is not being flown in a

complete loop, these numbers still apply if the plane is simply being pulled out of a dive.

G-load in verticle loops

The point is that the maximum load is far greater than simply the weight of the plane.

Three things need to be considered:

The weight of the plane less the weight of the wings, W sub b.

The maximum-Gs the plane experiences in flight, N.

A safety-factor accounting for material and construction flaws, S.


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The wings are assumed to support themselves, therefore, the wing root only supports

the weight of the fuselage and tail. However, this weight may be multiplied many times

determined by the maximum-Gs experienced during a flight maneuver. The value of N

would be 3, if a 3G flight maneuever is the worst case. This would result in a maximum

of N* W b.

Bending Strength

Knowing the bending-moment across the span determines the required bending-

strength of the spar. The bending-strength is determined by the structure of the spar, its

shape, and the materials used. In the same way that the bending-moment of the wing was

not as simple as our simple approximation for a rectangular wing, determining the

equations for the bending-strength of the spar is also not simple and straight-forward. The

problem is that the force is linearly varying from zero at the neutral-axis to its maximum

at the surface. However, the equations are well known for I-beam and comparable

structures. The bdimension represents the total empty space within the beam and may be

split, as in the case of an I-beam.

Figure 7.5 - Spar Structure

B H3- b h

3

M = -----------

6 H

Figure-7.6 compares the accurate moment-of-inertia calculation to several

approximations. The approximations are based on the cross-sectional area of the spar-

cap, A = b Tc. Each approximation multiplies the area, A, by: 1) the full spar-height, H,2) the distance between the midpoints of both spar caps, (H-h)/2, and 3) the distance

between the spar-caps, h = H-2Tc. The curves illustrate how all approximations become

increasingly accurate as the spar-caps become thinner.


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Figure 7.6 - Comparision of Moment-of-Inertia calculation and approximations vs spar-

cap thickness (1" spar)

Only the last approximation, using the distance between the spar-caps, is

conservative in that it errors on the side of under estimating the strength. It is also

relatively accurate. This approximation has less than 5% error for the case where the

spar-caps are less-than 15% of the spar thickness. This is certainly the case for 1/8" spar-

caps in a 7/8" spar, more typical in wooden construction. High-performance structures

using thin (~3%) carbon-fiber laminates can use the most convenient approximation

accurately.

Shear-Strength

The forces involved in resisting the bending-moment of the wing are parallel to

the spar. The spar must also support the weight of the plane as it is distributed across thewing. Like the previous calculations, it must consider the lift distribution across the wing.

A simple conservative approximation assumes a proportionally decreasing load from the

maximum load at the root, to zero load at the wing-tip. Figure-6 shows actual values for

several planforms. It shows that for an elliptical planform, the shear is actually 40% of

the max at the midpoint of the halfspan.


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Spar definition

The maximum bending moment from previous section was found to be as

583035.21 Nm. Therefore we define 3 Spars with front spar at 15% of chord, middle spar

at 45% of chord and rear spar at 70% of chord. The position of the three spars from the

leading edge of the root chord is given below as follows

Possible Spar Location

Front spar - 15% of chord = 2.25 m

Middle spar - 45% of chord = 6.75 m

Rear spar - 70% of chord = 10.5 m

Bending moment M = Max BM * FOS * n

= 583035.21 1.5 8.932

= 7811851.79Nm

The Structural load bearing members in the wing are the Spars and Stringers. Thebending moment carried by the Spars is 70% and that of Stringers is 30% of the total

Bending Moment.

Bending Moment taken by Spars is = 0.7 x 7811851.79Nm = 5468296.25 Nm


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WING CONSTRUCTION:

1. WING FUEL TANKS

In addition to providing the required strength and stiffness, the structural

box almost always has to provide fuel space. Integral tanks, as opposed to separate

internally supported types, are preferred since their use enables the maximum advantage

to be taken of the available volume. Integrally machined or molded constructions, which

use a small amount of large components, are obviously an advantage since sealing is

reduced to a minimum. The major problem occurs at tank end ribs, particularly in the

corners of the spar web and skins, and at lower surface access panels. The corner

difficulty is overcome by using special suitcase corner fittings.

Access panels should be large enough for a person to get through so that

the inside can be inspected and resealed if necessary. On shallow section wings, the

access has to be in the lower surface so that the operator can work in an acceptable way

even if the depth is insufficient to climb in completely. Apart from the sealing problems,

lower surface access panels are in what is primarily a tension skin and so introduce stress

concentrations in an area where crack propagation is a major consideration. The access

panels are arranged in a span-wise line so the edge reinforcing can be continuous andminimum stress concentration due to the cut-outs. Access panels are often designed to

carry only shear and pressure loads, the wing bending being reacted by the edge

reinforcing members. A deep wing can avoid these problems by using upper surface

access panels but this is not a preferred aerodynamic solution.

2. RIB LOCATION AND DIRECTION

The span-wise location of ribs is of some consequence. Ideally, the rib

spacing should be determined to ensure adequate overall buckling support to the

distributed flanges. This requirement may be considered to give a maximum pitch of the

ribs. In practice other considerations are likely to determine the actual rib locations such

as:


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a) Hinge positions for control surfaces and attachment/operating points for flaps,

slats, and spoilers.

b) Attachment locations of power plants, stores and landing gear structure.

c) A need to prevent or postpone skin local shear or compression buckling, as

opposed to overall buckling. This is especially true in a mass boom form of

construction.

d) Ends of integral fuel tanks where a closing rib is required. When the wing is

sweep back, it is usual for the ribs to be arranged in the flight direction and

thereby define the aerofoil section. While the unswept wing does give

torsional stiffness, the ribs are heavier, connections are more complex and in

general the disadvantages overweigh the gains.

3. HORIZONTAL STABILISER

When the horizontal stabilizer is constructed as a single component across the

centerline of the aircraft, the basic structural requirements are very similar to those of a

wing.

4. VERTICAL STABILISER

The vertical stabilizer presents a set of issues which are different from those of the

main plane or horizontal stabilizer. Relevant matters are:

a) It is not unusual to build the vertical stabilizer integrally with the rear fuselage.

The spars are extended to form fuselage frames or bulkheads. A root rib is

made to coincide with the upper surface of the fuselage and is used to transmit

the fin root skin shears directly into the fuselage skin. Fin span-wise bending

results in fuselage torsion. Often it is logical to incline the rear spar bulkhead to

continue the line of the rear spar since it is usually the end of the main fuselage

structure. On the other hand, the front spar and any intermediate attachment

frames are often best kept perpendicular to fuselage fore and aft datum. The

change in direction being made at the fin root rib. Otherwise the structural form

can follow that of a wing.


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b) Sometimes on smaller aircraft the fin is designed as a separate component which

may readily be detached. The fin attachment lugs are arranged in both lateral and

fore and aft directions so that in addition to vertical loads they react side and drag

loads.

HIGH LIFT SYSTEMS:

There is a wide variety of leading and trailing edge high-lift systems. Some

types are simply hinged to the wing, but many require some degree of chord-wise

extension. This can be achieved by utilizing a linkage, a mechanism, a pivot located

outside the aerofoil contour or, perhaps most commonly, by some from of track. Trailing

edge flaps may consist of two or more separate chord-wise segments, or slats, to give a

slotted surface and these often move on tracts attached to the main wing structure.

The majority of flaps and slats are split into span wise segments of no greater

lengths than can be supported at two or three locations. As with control surfaces, the

locations of the support points are established so as to minimize local deformations since

the various slots are critical in determining the aerodynamic performance. Sometimes the

actuation may be located at a different pan wise position from the support points. This is

often a matter of convenience, layout clearances, and the like.

The structural design of flaps is similar to that of control surfaces but its simpler

as there is no requirement for mass balance, the operating mechanisms normally being

irreversible. On large trailing edge flap components, there is often more than one spar

member. Especially when this assists in reacting the support or operating loading. There

may be a bending stiffness problem in the case of relatively small chord slat segments

and full depth honey combs can be used to deal with this. Figure shows a cross section of

a typical slotted flap of metal construction but the same layout applies if composite

materials are used.


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Delta Wings:

A delta wing is awing whose shape when viewed from above looks like a

triangle, often with its tip cut off. It sweeps sharply back from the fuselage with the angle

between the leading edge (the front) of the wing often as high as 60 degrees and the angle

between the fuselage and the trailing edge of the wing at around 90 degrees. Often delta-

wing airplanes lack horizontal stabilizers. Despite the fact that paper airplanes have delta

wings and appear to fly quite well when launched from a height, delta wings actually

perform poorly at low speeds and often are unstable (i.e., they do not stay in level flight

on their own). Their primary advantage is efficiency in high-speed flight.

By 1953, Convair's engineers had developed the YF-102 Delta Dagger, a radical

design that lacked a horizontal tail and featured a large, sharply swept delta wing. Wind

tunnel tests of small-scale models indicated that the aircraft could accelerate through

Mach 1 (the speed of sound) with relative ease, rather than "punching" through it like

earlier experimental planes that had to burn a lot of fuel to go faster than Mach 1.

However, the first prototype unexpectedly encountered immense drag as it approached

Mach 1. This so-called "transonic" region presented a major problem for the aircraft.

Near the same time, Richard T. Whitcomb, an aeronautical scientist at

theNational Advisory Committee for Aeronautics (NACA), was studying transonic drag.

Whitcomb developed what he called the "supersonic area rule." This theory stated that

aircraft that would fly at supersonic speed should increase in cross-sectional area from a

pointed nose. Anything that protruded into the airstream, such as the canopy over the

cockpit, wings, or tail, should be accompanied by a reduction in cross-section elsewhere.

In 1954, Whitcomb, who was then only 33-years old, was awarded the prestigiousCollier

Trophy for this contribution to aeronautics.

Convair's designers quickly applied the supersonic area rule to a new aircraft, the

YF-102A, pinching the fuselage near its mid-point to give it a slightly hourglass (or

Coke-bottle) appearance. This was a compromise for an existing aircraft; later airplanes

included the area rule in their designs in much less obvious ways. When the first YF-

102A with this new design took flight, it easily accelerated through Mach 1.
http://www.centennialofflight.gov/essay/Dictionary/airfoil/DI2.htmhttp://www.centennialofflight.gov/essay/Evolution_of_Technology/NACA/Tech1.htmhttp://www.centennialofflight.gov/essay/Dictionary/Area_Rule/DI103.htmhttp://www.centennialofflight.gov/essay/Dictionary/Collier_Trophy/DI60.htmhttp://www.centennialofflight.gov/essay/Dictionary/Collier_Trophy/DI60.htmhttp://www.centennialofflight.gov/essay/Dictionary/Collier_Trophy/DI60.htmhttp://www.centennialofflight.gov/essay/Dictionary/Collier_Trophy/DI60.htmhttp://www.centennialofflight.gov/essay/Dictionary/Collier_Trophy/DI60.htmhttp://www.centennialofflight.gov/essay/Dictionary/Area_Rule/DI103.htmhttp://www.centennialofflight.gov/essay/Evolution_of_Technology/NACA/Tech1.htmhttp://www.centennialofflight.gov/essay/Dictionary/airfoil/DI2.htm


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The two most famous current aircraft to use the delta wing are theConcorde and

the Space Shuttle. The Concorde's delta wing made the plane's sustained cruising speed

of Mach 2 possible. The Space Shuttle's wing, known as a "cranked delta" because the

leading edge of the wing has a slight bend near its midpoint, is used for a different

purpose. The Space Shuttle originally had what was known as a "high crossrange"

requirement, which was the ability to glide for thousands of miles to either side of its

flight path when landing. Conventional straight wings did not provide enough lift at high

speeds and altitudes to achieve this type of range, and so the large delta wing was

necessary.

While delta wings are critical to achieving high lift for supersonic flight, they also

have a number of disadvantages for less high-performing aircraft. They require high

landing and takeoff speeds and long takeoff and landing runs, are unstable at high angles

of attack, and produce tremendous drag when "trimmed" to keep the plane level. Of these

disadvantages, pilots and designers usually consider the high landing and takeoff speeds

the most important because they make flying the plane dangerous. Indeed, when the

Concorde had its first ever crash in 2000, after two decades of safe operations, the high-

speed takeoff was a factor in this terrible accident, for the plane's high ground speed

before becoming airborne placed major stress upon the aircraft's tires, which exploded

upon striking an object on the runway.

Computer-controlled "fly-by-wire" flight control systems have allowed designers

to compensate for some of the delta wing's poor control qualities. Canards are small

horizontal fins (or small wings) mounted on the fuselage in front of an aircraft's main

wings to provide greater control, particularly during high angles of attack. When they are

part of a delta-wing aircraft, they improve its stability and maneuverability.
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8. MATERIALS

8.1 Materials used in aircraft manufacturing:

For many years, aircraft designers could propose theoretical designs that they

could not build because the materials needed to construct them did not exist. (The term

"unobtainium" is sometimes used to identify materials that are desired but not yet

available.)

The techniques for building aircraft evolved gradually during the years between

the wars. Wood and canvas changed to aluminum as the principal structural material

while designs improved and records were set and broken. Monoplanes (single wing

aircraft) were becoming more popular than biplanes (two wing aircraft).

Material requirement for aircraft building:

small weight

high specific strength

heat resistance

fatigue load resistance

crack and corrosion resistance

8.2 Titanium alloy

Titanium alloysaremetallicmaterials which contain a mixture oftitanium and

otherchemical elements.Such alloys have very hightensile strength andtoughness (even

at extreme temperatures), light weight, extraordinary corrosion resistance, and ability to

withstand extreme temperatures. However, the high cost of both raw materials and

processing limit their use tomilitary applications,aircraft, spacecraft, medical

devices,connecting rods on expensivesports cars and some premiumsports

equipment andconsumer electronics. Auto manufacturers Porsche and Ferrari also use

titanium alloys in engine components due to its durable properties in these high stress

engine environments.
http://en.wikipedia.org/wiki/Metalhttp://en.wikipedia.org/wiki/Materialhttp://en.wikipedia.org/wiki/Titaniumhttp://en.wikipedia.org/wiki/Chemical_elementhttp://en.wikipedia.org/wiki/Tensile_strengthhttp://en.wikipedia.org/wiki/Toughnesshttp://en.wikipedia.org/wiki/Militaryhttp://en.wikipedia.org/wiki/Aircrafthttp://en.wikipedia.org/wiki/Spacecrafthttp://en.wikipedia.org/wiki/Connecting_rodhttp://en.wikipedia.org/wiki/Sports_carhttp://en.wikipedia.org/wiki/Sports_equipmenthttp://en.wikipedia.org/wiki/Sports_equipmenthttp://en.wikipedia.org/wiki/Consumer_electronicshttp://en.wikipedia.org/wiki/Consumer_electronicshttp://en.wikipedia.org/wiki/Sports_equipmenthttp://en.wikipedia.org/wiki/Sports_equipmenthttp://en.wikipedia.org/wiki/Sports_equipmenthttp://en.wikipedia.org/wiki/Sports_carhttp://en.wikipedia.org/wiki/Connecting_rodhttp://en.wikipedia.org/wiki/Spacecrafthttp://en.wikipedia.org/wiki/Aircrafthttp://en.wikipedia.org/wiki/Militaryhttp://en.wikipedia.org/wiki/Toughnesshttp://en.wikipedia.org/wiki/Tensile_strengthhttp://en.wikipedia.org/wiki/Chemical_elementhttp://en.wikipedia.org/wiki/Titaniumhttp://en.wikipedia.org/wiki/Materialhttp://en.wikipedia.org/wiki/Metal


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8.3 ALUMINIUM ALLOY:

Aluminium alloys arealloys in whichaluminium (Al) is the predominant metal.

The typical alloying elements arecopper,magnesium, manganese,silicon andzinc.There

are two principal classifications, namelycasting alloys and wrought alloys, both of which

are further subdivided into the categoriesheat-treatable and non-heat-treatable. About

85% of aluminium is used for wrought products, for example rolled plate, foils

andextrusions. Cast aluminium alloys yield cost effective products due to the low

melting point, although they generally have lowertensile strengths than wrought alloys.

The most important cast aluminium alloy system is Al-Si, where the high levels of silicon

(4.0% to 13%) contribute to give good casting characteristics. Aluminium alloys are

widely used in engineering structures and components where light weight or corrosion

resistance is required.

Wrought alloys:

The International Alloy Designation System is the most widely accepted naming scheme

forwrought alloys.Each alloy is given a four-digit number, where the first digit indicates

the major alloying elements.

1000 series are essentially pure aluminium with a minimum 99% aluminium content by

weight and can bework hardened.

2000 series are alloyed with copper, can beprecipitation hardened to strengths

comparable tosteel.

3000 series are alloyed with manganese, and can bework hardened.

4000 series are alloyed with silicon. They are also known assilumin.

5000 series are alloyed with magnesium.

6000 series are alloyed with magnesium and silicon, are easy to machine, and can be

precipitation hardened, but not to the high strengths that 2000 and 7000 can reach.

7000 series are alloyed withzinc, and can be precipitation hardened to the highest

strengths of any aluminium alloy.

8000 series is a category mainly used for lithiumalloys.
http://en.wikipedia.org/wiki/Alloyshttp://en.wikipedia.org/wiki/Aluminiumhttp://en.wikipedia.org/wiki/Copperhttp://en.wikipedia.org/wiki/Siliconhttp://en.wikipedia.org/wiki/Zinchttp://en.wikipedia.org/wiki/Castinghttp://en.wikipedia.org/wiki/Heat_treatmenthttp://en.wikipedia.org/wiki/Extrudinghttp://en.wikipedia.org/wiki/Tensile_strengthhttp://en.wikipedia.org/w/index.php?title=Wrought_alloy&action=edit&redlink=1http://en.wikipedia.org/wiki/Work_hardenedhttp://en.wikipedia.org/wiki/Precipitation_hardenedhttp://en.wikipedia.org/wiki/Steelhttp://en.wikipedia.org/wiki/Work_hardenedhttp://en.wikipedia.org/wiki/Siluminhttp://en.wikipedia.org/wiki/Siluminhttp://en.wikipedia.org/wiki/Work_hardenedhttp://en.wikipedia.org/wiki/Steelhttp://en.wikipedia.org/wiki/Precipitation_hardenedhttp://en.wikipedia.org/wiki/Work_hardenedhttp://en.wikipedia.org/w/index.php?title=Wrought_alloy&action=edit&redlink=1http://en.wikipedia.org/wiki/Tensile_strengthhttp://en.wikipedia.org/wiki/Extrudinghttp://en.wikipedia.org/wiki/Heat_treatmenthttp://en.wikipedia.org/wiki/Castinghttp://en.wikipedia.org/wiki/Zinchttp://en.wikipedia.org/wiki/Siliconhttp://en.wikipedia.org/wiki/Copperhttp://en.wikipedia.org/wiki/Aluminiumhttp://en.wikipedia.org/wiki/Alloys


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Cast alloys:

The Aluminium Association (AA) has adopted a nomenclature similar to that of wrought

alloys.British Standard and DIN have different designations. In the AA system, the

second two digits reveal the minimum percentage of aluminium, e.g. 150.x corresponds

to a minimum of 99.50% aluminium. The digit after the decimal point takes a value of 0

or 1, denoting casting and ingot respectively. The main alloying elements in the AA

system are as follows:

1xx.x series are minimum 99% aluminium

2xx.x series copper

3xx.x series silicon, copper and/or magnesium

4xx.x series silicon

5xx.x series magnesium

7xx.x series zinc

8xx.x series lithium

Named alloys:

Alclad Aluminium sheet formed from high-purity aluminium surface layers bonded to

high strength aluminium alloy core material

Birmabright (aluminium, magnesium) a product of The Birmetals Company, basically

equivalent to 5251

Duralumin (copper, aluminium)

Magnalium

Magnox (magnesium, aluminium)

Silumin (aluminium, silicon)

Titanal (aluminium, zinc, magnesium, copper, zirconium) a product ofAustria Metall

AG.Commonly used in high performance sports products, particularly snowboards and

skis.

Y alloy,Hiduminium,R.R. alloys: pre-warnickel-aluminium alloys, used in

aerospace and engine pistons, for their ability to retain strength at elevated temperature.
http://en.wikipedia.org/wiki/British_Standardhttp://en.wikipedia.org/wiki/Alcladhttp://en.wikipedia.org/wiki/Birmabrighthttp://en.wikipedia.org/wiki/Duraluminhttp://en.wikipedia.org/wiki/Magnaliumhttp://en.wikipedia.org/wiki/Magnox_(alloy)http://en.wikipedia.org/wiki/Siluminhttp://en.wikipedia.org/w/index.php?title=Titanal&action=edit&redlink=1http://en.wikipedia.org/wiki/Austria_Metall_AGhttp://en.wikipedia.org/wiki/Austria_Metall_AGhttp://en.wikipedia.org/wiki/Y_alloyhttp://en.wikipedia.org/wiki/Hiduminiumhttp://en.wikipedia.org/wiki/R.R._alloyshttp://en.wikipedia.org/wiki/Nickel-aluminium_alloyhttp://en.wikipedia.org/wiki/Nickel-aluminium_alloyhttp://en.wikipedia.org/wiki/R.R._alloyshttp://en.wikipedia.org/wiki/Hiduminiumhttp://en.wikipedia.org/wiki/Y_alloyhttp://en.wikipedia.org/wiki/Austria_Metall_AGhttp://en.wikipedia.org/wiki/Austria_Metall_AGhttp://en.wikipedia.org/wiki/Austria_Metall_AGhttp://en.wikipedia.org/w/index.php?title=Titanal&action=edit&redlink=1http://en.wikipedia.org/wiki/Siluminhttp://en.wikipedia.org/wiki/Magnox_(alloy)http://en.wikipedia.org/wiki/Magnaliumhttp://en.wikipedia.org/wiki/Duraluminhttp://en.wikipedia.org/wiki/Birmabrighthttp://en.wikipedia.org/wiki/Alcladhttp://en.wikipedia.org/wiki/British_Standard


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Applications

Aerospace alloys

Scandium-Aluminium

The main application of metallic scandium by weight is inaluminium-scandium

alloys for minor aerospace industry components. These alloys contain between 0.1% and

0.5% (by weight) of scandium. They were used in the Russian military aircraft Mig

21 andMig 29.

List of aerospace Aluminium alloys:

The following aluminium alloys are commonly used inaircraft and

otheraerospace structures:

7075 aluminium

6061 aluminium

6063 aluminium

2024 aluminium

5052 aluminium

8.4. Composites:

For many years, aircraft designers could propose theoretical designs that they

could not build because the materials needed to construct them did not exist. (The term

"unobtainium" is sometimes used to identify materials that are desired but not yet

available.) For instance, large space planes like the Space Shuttle would have proven

extremely difficult, if not impossible, to build without heat-resistant ceramic tiles to

protect them during re-entry.

Fiberglass is the most common composite material, and consists of glass fibers

embedded in a resin matrix. Fiberglass was first used widely in the 1950s for boats and

automobiles, and today most cars have fiberglass bumpers covering a steel frame.

Fiberglass was first used in the Boeing 707 passenger jet in the 1950s, where it comprised

about two percent of the structure.
http://en.wikipedia.org/wiki/Aluminium-scandium_alloyhttp://en.wikipedia.org/wiki/Aluminium-scandium_alloyhttp://en.wikipedia.org/wiki/Mig_21http://en.wikipedia.org/wiki/Mig_21http://en.wikipedia.org/wiki/Mig_29http://en.wikipedia.org/wiki/Aircrafthttp://en.wikipedia.org/wiki/Aerospacehttp://en.wikipedia.org/wiki/7075_aluminium_alloyhttp://en.wikipedia.org/wiki/6061_aluminium_alloyhttp://en.wikipedia.org/wiki/6063_aluminium_alloyhttp://en.wikipedia.org/wiki/2024_aluminium_alloyhttp://en.wikipedia.org/w/index.php?title=5052_aluminium_alloy&action=edit&redlink=1http://en.wikipedia.org/w/index.php?title=5052_aluminium_alloy&action=edit&redlink=1http://en.wikipedia.org/wiki/2024_aluminium_alloyhttp://en.wikipedia.org/wiki/6063_aluminium_alloyhttp://en.wikipedia.org/wiki/6061_aluminium_alloyhttp://en.wikipedia.org/wiki/7075_aluminium_alloyhttp://en.wikipedia.org/wiki/Aerospacehttp://en.wikipedia.org/wiki/Aircrafthttp://en.wikipedia.org/wiki/Mig_29http://en.wikipedia.org/wiki/Mig_21http://en.wikipedia.org/wiki/Mig_21http://en.wikipedia.org/wiki/Mig_21http://en.wikipedia.org/wiki/Aluminium-scandium_alloyhttp://en.wikipedia.org/wiki/Aluminium-scandium_alloy


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8.4.1 Carbon Fiber:

Carbon fibre is the reinforcement material of choice for "advanced" composites,

particularly following significant price reductions over the past decade. A major

advantage of carbon fibres is their higher fatigue resistance compared to glass orAramid.

Unlike these last two materials, carbon fibres do not suffer from stress rupture. Carbon

fibres are supplied in tows and may vary from 1000 fibres per tow to hundreds of

thousands of fibres per tow.

8.4.2 KEVLAR:

Kevlar is the registered trademark for a para-aramid synthetic fiber, related to

other aramids such as Nomex and Technora. Developed at DuPont in 1965, this high

strength material was first commercially used in the early 1970s as a replacement for

steel in racing tires. Typically it is spun into ropes or fabric sheets that can be used as

such or as an ingredient in composite material components.

8.4.3 FIBERGLASS:

Fiberglass, (also called fibreglass and glass fibre), is material made from

extremely fine fibers of glass. It is used as a reinforcing agent for many polymer

products; the resulting composite material, properly known as fiber-reinforced polymer

(FRP) or glass-reinforced plastic (GRP), is called "fiberglass" in popular usage.

Glassmakers throughout history have experimented with glass fibers, but mass

manufacture of fiberglass was only made possible with the invention of finer machine

tooling.

Just some applications for aerospace fiberglass:

Aircraft Enclosures for Controls

Antenna Enclosures

Storage Bins

Luggage Racks
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8.4.4 BORON FIBER:

Specialty Materials boron fiber is particularly desirable for use in aerospace

industry applications where high compression loads are present. Some examples of these

applications include the following aircraft: F-15 Fighter, F-14 Fighter, B1 Bomber,

Blackhawk Helicopter, Predator B UAV, and the Space Shuttle. Additionally, Specialty

Materials boron fiber has been used to repair aircraft structures in both military and

commercial aircraft including the B-52, C-130, F-4, F-5, F-111 and Boeing 727, 747, 757

and MD-11. Please see the chart below for reference.

Aircraft Manufacturer MaterialApplication

BoeingBoron/Epoxy

Horizontal

and vertical

tail skins

rudder

Grumman

AerospaceCorp.

Boron/Epoxy

Horizontal tail

skins

Rockwell

International Boron/Epoxy Dorsal Longeron

Sikorsky Boron/Epoxy Rotor Blades and

Stabilator

General AtomicsBoron/ Graphit e

Epoxy Top Beam Cap


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8.5. Radar-absorbent material

Theory

Radar-absorbent material, or RAM, is a class of materials used in stealth

technology to disguise a vehicle or structure from radar detection. A material's

absorbency at a given frequency of radar wave depends upon its composition. RAM

cannot perfectly absorb radar at any frequency, but any given composition does have

greater absorbency at some frequencies than others; there is no one RAM that is suited to

absorption of all radar frequencies. A common misunderstanding is that RAM makes an

object invisible to radar. A radar absorbent material can significantly reduce an object's

radar cross-section in specific radar frequencies, but it does not result in "invisibility" on

any frequency. Bad weather may contribute to deficiencies in stealth capability. A

particularly disastrous example occurred during the Kosovo war, in which moisture on

the surface of an F- 117 Nighthawk allowed long-wavelength radar to track and shoot it

down. RAM is only a part of achieving stealth.

Types of RAM

Iron ball paint

One of the most commonly known types of RAM is iron ball paint. It contains

tiny spheres coated with carbonyl iron or ferrite. Radar waves induce molecularoscillations from the alternating magnetic field in this paint, which leads to conversion of

the radar energy into heat. The heat is then transferred to the aircraft and dissipated. The

iron particles in the paint are obtained by decomposition of iron pentacarbonyl and may

contain traces of carbon, oxygen and nitrogen.

Foam absorber

Foam absorber is used as lining of anechoic chambers for electromagnetic

radiation measurements. This material typically consists of fireproofed urethane foam

loaded with carbon black, and cut into long pyramids. The length from base to tip of the

pyramid structure is chosen based on the lowest expected frequency and the amount of

absorption required. For low frequency damping, this distance is often 24 inches, while

high frequency panels are as short as 3-4 inches. Panels of RAM are installed with the

tips pointing inward to the chamber.


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Jaumann absorber

A Jaumann absorber or Jaumann layer is a radar absorbent device. When first

introduced in 1943, the Jaumann layer consisted of two equally-spaced reflective surfaces

and a conductive ground plane. One can think of it as a generalized, multilayered

Salisbury screen as the principles are similar.

More elaborate Jaumann absorbers use series of dielectric surfaces that separate

conductive sheets. The conductivity of those sheets increases with proximity to the

ground plane.


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9. 3D DIAGRAM


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9. Summary

V-n characteristics for positive load factor [n limpositive= 9]

VA(Design Maneuvering Speed) = 195.72 m/s.

VC(Design Cruising Speed) = 388.88 m/s.VD(Design Diving Speed) = VC + 26.82m/s =412.7m/s.

V-n characteristics for negative load factor [n limnegative= -3]

VS (Stalling speed) = 64.72m/s.

VA = 61.83=112.02 m/s.

VC (Cruise speed) = 388.88 m/s.

VD (Diving speed) = VC + 26.82m/s = 412.72m/s.

V-n characteristics for gust load

Maximum gust load factor = 8.439.

Minimum gust load factor = -3.2.

Load estimation of wings

Maximum shear force =108468.9N

Maximum bending moment=532378.9Nm.

Fuselage design

The forward fuselage is 5.2m long and 1.7m wide.

The canopy is about 356 cm long and 114 cm wide.

The mid fuselage in 5.2 m long and 2m wide.

The aft fuselage is 5.8m long and 3.6m wide.


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Materials used

Aluminium: Fuselage (because they are corrosion resistant and have high

strength to weight ratio.

Transparent Plastic: Transparent plastic is used in canopies, windshields, and

other transparent enclosures.

Kevlar: for racing tires.

Boron fiber: Particularly desirable for use in aerospace industry applications

where high compression loads are present.

Reinforced Plastic: Reinforced plastic is used in the construction of radomes,

wingtips, stabilizer tips, antenna covers, and flight controls. It has high strength-

to-weight ratio and is resistant to mildew and rot. Suitable for other parts of the

aircraft.

Design

Thus the final sketches of Aircraft design project II are made using the three view

diagram of our Aircraft design project I.


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10. REFERENCES

Raymer, D.P., Aircraft Design: A Conceptual Approach, Third Edition, AIAA,

Inc., Reston, VA, pp. 229-270, 379-401, 406, 408-412, 426-446, 2006.

Roskam, Jan., Airplane Design, Roskam Aviation and Engineering Corporation,

Ottawa, KS, 2007.

Raymer, D.P., Aircraft Design: A Conceptual Approach, Second Edition, AIAA,

Inc., Reston, 2006.

Lloyd R. Jenkinson and James F. Marchman III, Aircraft Design Projects forengineering students, 2003.

John D. Anderson, Aircraft Performance and Design, Tata McGraw-Hill Edition

2010.

John D. Anderson, Fundamentals of Aerodynamics, Tata McGrraw-Hill Edition

2010.

Airfoil Investigation Database,http://www.worldofkrauss.com/foils.

F136 Engine- Best Of Both Worlds, http://f136engine.com

Stability Performances and Analysis, Wing Loading Estimation

http://adamone.rchomepage.com

Composites

http://hsc.csu.edu.au/engineering_studies/aero_eng/2579/polymer_composites.ht

ml

V-n Diagramhttp://adg.stanford.edu/aa241/structures/ivn.html

Shear Force and Bending Moment Diagrams

http://www.transtutors.com/homework-help/Mechanical+Engineering/Bending+Moment+and+Shear+Force/cantilever-

subjected-to-uniformly-varying-load.aspx
http://adamone.rchomepage.com/http://adamone.rchomepage.com/http://hsc.csu.edu.au/engineering_studies/aero_eng/2579/polymer_composites.htmlhttp://hsc.csu.edu.au/engineering_studies/aero_eng/2579/polymer_composites.htmlhttp://hsc.csu.edu.au/engineering_studies/aero_eng/2579/polymer_composites.htmlhttp://adg.stanford.edu/aa241/structures/ivn.htmlhttp://adg.stanford.edu/aa241/structures/ivn.htmlhttp://www.transtutors.com/homework-help/Mechanical+Engineering/Bending+Moment+and+Shear+Force/cantilever-subjected-to-uniformly-varying-load.aspxhttp://www.transtutors.com/homework-help/Mechanical+Engineering/Bending+Moment+and+Shear+Force/cantilever-subjected-to-uniformly-varying-load.aspxhttp://www.transtutors.com/homework-help/Mechanical+Engineering/Bending+Moment+and+Shear+Force/cantilever-subjected-to-uniformly-varying-load.aspxhttp://www.transtutors.com/homework-help/Mechanical+Engineering/Bending+Moment+and+Shear+Force/cantilever-subjected-to-uniformly-varying-load.aspxhttp://www.transtutors.com/homework-help/Mechanical+Engineering/Bending+Moment+and+Shear+Force/cantilever-subjected-to-uniformly-varying-load.aspxhttp://www.transtutors.com/homework-help/Mechanical+Engineering/Bending+Moment+and+Shear+Force/cantilever-subjected-to-uniformly-varying-load.aspxhttp://www.transtutors.com/homework-help/Mechanical+Engineering/Bending+Moment+and+Shear+Force/cantilever-subjected-to-uniformly-varying-load.aspxhttp://adg.stanford.edu/aa241/structures/ivn.htmlhttp://hsc.csu.edu.au/engineering_studies/aero_eng/2579/polymer_composites.htmlhttp://hsc.csu.edu.au/engineering_studies/aero_eng/2579/polymer_composites.htmlhttp://adamone.rchomepage.com/

Ac Design Project 2

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Transcript of Ac Design Project 2