400 Seater Passenger Aircraft

AIRCRAFT DESIGN PROJECT-I

DESIGN PROJECT OF FOUR HUNDRED SEATER TWIN

ENGINE PASSENGER AIRCRAFT

A PROJECT REPORT

Submitted by

In partial fulfilment for the awards of the degree

Of

BACHELOR OF ENGINEERING

IN

AERONAUTICAL ENGINEERING

DHANALAKSHMI SRINIVASAN COLLEGE OF ENGG & TECHNOLOGY,

CHENNAI.

ANNA UNIVERSITY: CHENNAI 600 025

APRIL 2011

BONAFIDE CERTIFICATE

Certified that this report “A DESIGN PROJECT OF FOUR HUNDRED

SEATER TWIN ENGINE PASSENGER AIRCRAFT’’ is the bonafide work of

.................................... who carried out the project work under my supervision.

1


SIGNATURE SIGNATURE

SARAVANAN.G .........................

HEAD OF THE DEPARTMENT SUPERVISOR

Aeronautical engineering, Lecturer,

D.S. College of Engineering & Technology, D.S. College of Engineering&

Chennai- 603104. Technology, Chennai-603104.

Internal Examiner External Examiner

This report for the Design of FOUR HUNDRED SEATER TWIN ENGINE

PASSENGER AIRCRAFT is prepared on the basis of Anna University Syllabus.

This is prepared by references attached in this report.

For getting interested in this subject and nurturing my knowledge base, I would

like to thank my beloved teachers. Mr. Saravanan.G , Head of the Department

and ......................... , lecture who deserve all credit

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Last, but not least, I am thankful to all of my Department staffs.

Dedicated toBeloved Parents, Department Staffs

& Management

TABLE OF CONTENTS

CHAPTER NO TITLE PAGE

ABSTRACT

LIST OF SYMBOLS

INTRODUCTION

1. COMPARITIVE STUDY OF 12

1.1 DIMENSIONS

1.2 WEIGHT CONFIGURATION

3


1.3 PERFORMANCE

1.4 ENGINE CONFIGURATION

2. SELECTION OF MAIN PARAMETERS 16

2.1 SELECTION OF PARAMETERS

2.1.1 Airfoil selection

2.1.2 Co-efficient of lift Vs Angle of attack

2.1.3 Co-efficient of lift Vs Drag

2.1.4 Max L/D Vs Velocity or Mach no.

2.1.5 Range Vs Velocity

2.1.6 Altitude Vs Velocity

2.1.7 Aspect ratio Vs Velocity

2.1.8 Wing loading Vs Velocity

2.1.9 SFC Vs Mach number

2.1.10 T/W Vs Velocity

3. WEIGHT ESTIMATION 25

3.1 WEIGHT CALCULATION

3.2 MISSION PROFILE

3.3 APPROXIMATE WEIGHT ESTIMATION

4. ENGINE SELECTION 29

4.1 LOCATION OF ENGINE

4.2 THRUST CALCULATION

4.3 ENGINE CONFIGURATION

4.4 CONFIGURATION

4.4.1 Advantages of Buried Engine

4.4.2 Disadvantages of Buried Engine

4.4.3 Advantages of Low wing

4


4.4.4 Disadvantages of Low wing

5. AIRFOIL SELECTION 36

5.1 CALCULATION OF CL

5.1.1 Reynolds’s Number

5.1.2 Maximum CL

5.1.3 Skin friction Drag for turbulent flow

5.1.4 Required CL max

5.1.5 NACA-63-215

6. WING SELECTION 46

6.1 EQUIVALENT ASPECT RATIO

6.2 STRUCTURAL WEIGHT FOR VARYING

THE THICKNESS OF AIRFOIL

6.3 LOCATION OF CENTRE OF GRAVITY

7. WETTED SURFACE AREA AND DRAG ESTIMATION 49

7.1 CALCULATION OF WETTED SURFACE AREA

7.1.1Fuselage

7.1.2 Wing area

7.1.3 Horizontal Tail

7.1.4 Vertical Tail

7.1.5 Engine

12. THREE VIEWS OF SUPERSONIC FIGHTER AIRCRAFT 52

5


12.1 FRONT VIEW

12.2 TOP VIEW

12.3 SIDE VIEW

\

13. CONCLUSION 58

13.1 BIBLIOGRAPHY

ABSTRACT

6


All of the airliners aim at building an aircraft with large capacity

and long range at a higher velocity and with low fuel consumption. Our mini

project conceptualizes this aim. So in our mini project we have concentrated on a

400 seater passenger aircraft with twin engine which can travel at a cruise mach

number of 0.84 and a minimum range of 1200km at an optimum altitude. For the

propulsion system we have chosen an existing engine for reference. Historic data is

being used wherever necessary to make our project more precise

LIST OF SYMBOLS

W Weight of aircraft

W0 Overall weight

Wf Weight of fuel

We Empty weight

L Lift of aircraft

7


D Drag of the aircraft

CL Coefficient of lift

CD Coefficient of drag

S Wing area

B Wing span

T Thrust

T/W Thrust loading

W/S Wing loading

A.R Aspect ratio

Cr,Ct Chord length of root,tip

Tr,tt thickness of root, tip

Sπ Wetted surface area

CDπ Coefficient of drag of wetted surface area

ΛL.E Sweep angle of the leading edge

ß Dihedral angle

Α Angle of attack

Ρ Density(kg/m3)

Wing mean chord

Μ Ground friction

Ν Kinematics viscosity

Λ Taper ratio

C.G Center of gravity

R range

E Endurance

8


Free stream velocity

C Chord

Lf Length of fuselage

VT Vertical tail

HT Horizontal tail

Θ Angle of flap deflection

η0,ηi Span station of flap

G Gravity

S Distance

H Height

H altitude

INTRODUCTION

Airplane Design – Introduction

Three major types of airplane design are

9


1. Conceptual deign

2. Preliminary design

3 Detailed designs

1. CONCEPTUAL DESIGN :

It depends on what are the major factors for the designing the aircraft

A. power plant location

The power plant location is either padded or buried type

engines are more preferred .Rear location is preferred for low drag, reduced shock

and to use whole thrust.

B. selection of engine:

The engine to be used is selected according to the power

required.

Wing selection:

The selection of wing depends upon the selection of

low wing

mid wing

high wing

10


2. PRELIMINARY DESIGN :

Preliminary design is based only on loitering; U is the

mathematical method of skinning the aircraft after skinning the aircraft looks like a

masked body.

Preliminary design is done with the help of FORTRAN software.

3. DETAILED DESIGN :

In the detailed design considers each and every rivets, nuts, bolts,

paints, etc. In this design the connection and allocation are made.

1. COMPARATIVE STUDY OF PASSENGER AIRCRAFT

SPECIFICATION :

1.1.WING SPECIFICATIONS:

S.NO. NAME OF

A/C

WING

SPAN (m)

LENGTH

(m)

HEIGHT

(m)

WING

AREA

11


(m2 )

1. A330-300 60.30 63.30 16.70 361.6

2. A340-

600/600HGW

63.43 75.30 17.30 475.5

3. A350-900R 64.8 67.0 17.2 480.5

4. 777-300ER 64.8 73.9 18.7 477.6

5. 747-400 64.4 70.6 19.4 378.5

12


1.2.WEIGHT SPECIFICATION:

S.NO NAME OF

A/C

EMPTY

WEIGHT

MAX

(T/W)

GROSS

WEIGHT

1. A330-300 173000 0.7350 233000

2. A340-

600/600HGW

177000 0.70027 368000

3. A350-900R 176000 0.8990 301000

4. 777-300ER 175000 0.8517 347540

5. 747-400 178750 0.7180 396890

13


1.3.POWER PLANT SPECIFICATIONS:

14

S.NO. NAME OF

A/C

TYPE OF

ENGINE

NO.OF

ENGINE

THRUST

(KN)

1. A330-300 Pratt&Whitney

pw 4170

2 320

2. A340-

600/600HG

W

RR Trent 500 2 257.7

3. A350-900R PR Trent x WB 2 270.6

4. 777-300ER GE 90-110B 2 296

5. 747-400 GECF6-80CB5F 2 282


1.4. PERFORMANCE SPECIFICATIONS:

S.NO. NAME OF

A/C

MAX.

SPEED

(km/hr)

CRUISING

SPEED

(km/hr)

SERVIC

E

CEILING

(m)

RANGE

(km)

CREW

1. A330-300 900 871 12,643 10,501 2

2. A340-

600/600HG

W

905 854 11,887 14,350 2

3. A350-900R 850 805 11,490 9250 2

4. 777-300ER 840 810 11,680 14,630 2

5. 747-400 912 870 12.863 14,205 2

15


2. SELECTION OF MAIN PARAMETERS FOR AIRCRAFT DESIGN

2.1. SELECTION OF MAIN PARAMETERS:

2.1.1 C o-efficient of lift Vs Angle of attack

The experimental data indicate that CL varies linearly with over a

large range of angle of attack. Thin airfoil theory, which is the subject by

more advanced books on aerodynamics also predicts the same type of linear

variation. The slope of the linear position of the lift curve is designed as lift

slope there is still a positive value of CL that is there is still a positive value

of CL that is, there is still some lift even when the airfoil is at zero angle of

attack.

2.1.2 C o-efficient of lift Vs Co-efficient of drag

The drag polar is a parabola with its axis on the zero-lift axis and its

vortex is CD

CD0-is the parasite drag co-efficient at zero lift

and includes both induced drag and the contribution to parasite drag

due to lift in our redefined e, which now includes the effect from is called

the Oswald efficiency factor. The basic aerodynamic properties of the

airplane are and we consider both CDo and e as

known aerodynamic qualities obtained from the aerodynamicist.

16


2.1.3.C o-efficient of lift Vs Mach number

At low mach number less than Mcr, CD is virtually constant and

is equal to its low speed values. The free stream mach number at which CD

begins to increase rapidly is defined as the drag divergence mach number.

2.1.4.Dihedral

Dihedral is the design feature of the airplane that provides

lateral stability. Dihedral effect is always a coupling between yawing and

rolling motion, so that one doesn’t occur without the other.

2.1.5.L/Dmax Vs Mach number

To design the aircraft we should better understand the L/D Vs

Velocity. Because for passenger aircraft L/D should be maximum and is a

key parameter in design. Usually the (L/D) is maximum for the cruise flight

of most of the commercial aircraft.

17


2.1.6. Range Vs Velocity

It is plot the between the range of the aircraft and the velocity. The plot of

different aircraft is drawn. The Range is the total distance traversed by an

airplane on one load of fuel.

Range equation

Ct - specific fuel consumption

W0 – Gross weight of the airplane including everything, fuel load, payload, crew,

structure etc

W1 – Weight of the airplane when the fuel tanks are empty

18


830 840 850 860 870 880 890 900 910 9200

2000400060008000

10000120001400016000

RANGE Vs VELOCITY

A330-300

A340- 600/600HGW

A350-900R

777-300ER

747-400

Velocity (km/hr)

Rang

e (k

m)

2.1.7. Altitude Vs Velocity

The graph is drawn between the altitude and velocity. It is main design parameters.

830 840 850 860 870 880 890 900 910 920155016001650170017501800185019001950

Altitude Vs Velocity

A330-300A340- 600/600HGWA350-900R777-300ER747-400

Velocity (km/hr)

Altit

ude

(m)

19


2.1.8. Weight Vs velocity

The weight Vs velocity is drawn in the graph. For the various aircraft weight

is considered for the various aircraft weight is considered for drawing the graph.

The optimum weight is calculated.

830 840 850 860 870 880 890 900 910 920174500175000175500176000176500177000177500178000178500179000

Weight Vs Velocity

A330-300A340- 600/600HGWA350-900R777-300ER747-400

Velocity(km/hr)

Wei

ght (

kg)

20


2.1.10 Wing loading Vs Velocity

Wing loading effect on climb

Wing loading selection is important parameter for design of aircraft to find the

optimum wing loading by drawing graph.

21


830 840 850 860 870 880 890 900 910 920690

695

700

705

710

715

720

725

730

735

Wing Loading Vs Velocity

A330-300A340- 600/600HGWA350-900R777-300ER747-400

Velocity (km/hr)

(W/S

)

2.1.11. Thrust loading Vs Velocity

The drawn the graph between thrust loading for different passenger

aircraft with velocity of that aircraft. We find the optimum thrust loading of certain

category of aircraft.

22


830 840 850 860 870 880 890 900 910 9200

0.2

0.4

0.6

0.8

1

Thrust Loading Vs Velocity

A330-300

A340- 600/600HGW

A350-900R

777-300ER

747-400Velocity (km/hr)

(T/W

)

2.1.12 Aspect Ratio Vs Velocity

The graph is drawn between the aspect ratio and velocity. The

choice of low aspect ratio is driven by supersonic performance and high aspect

ratio for subsonic aircraft.

23


830 840 850 860 870 880 890 900 910 9200

2

4

6

8

10

12

Aspect ratio Vs Velocity

A330-300A340- 600/600HGWA350-900R777-300ER747-400

Velocity (km/hr)

Aspe

ct ra

tio

2.2. SELECTION OF PARAMETERS

Optimum VELOCITY 910km/hr

Optimum RANGE 12000km

Optimum ALTITUDE 1820m

Optimum WEIGHT 176000kg

Optimum ASPECT RATIO 9.30

24


Optimum W/S WING LOADING

713kg/m2

Optimum T/W THRUST LOADING

0.801

3. WEIGHT ESTIMATION:

3.1.Weight of aircraft:

Overall weight

From the specification =0.5742

25


3.2. Mission profile

2 3

0 1 4 5

3.3. Approximate Weight Estimation:

Estimation of :

Warm up & Take off:

=0.995

26


Climbing:

=0.985

Cruising:

R=6479.48nm

C=0.6-0.9lb/hr/lb

V=491.36 knots

=10-13

W 3

W 2

=exp ¿

= exp (−6479∗0.6491.36∗10

¿

W 3

W 2

=0.4533

Loitering:

E=18min

27


C=0.6-0.9lb/hr/lb

=10-13

W 4

W 3

=exp [−0.3∗0.610

]

W 4

W 3

=0.9821

Descending:

=0.99

Landing:

=0.9628

Mff =0.995*0.985*0.4533*0.9821*0.99*0.9628

Mff =0.415

= (1-Mff) = (1-0.415)

=0.5841

W 0=

W crew+W payload

1−W f

W 0

−W e

W 0

W 0=[2∗120 ]+[400∗160]1−0.584−0.5742

28


=330325.71kg

4. ENGINE SELECTION

Types of turbine engines

Turbojets:

29


The basic working principle of the turbojet engine is that air from outside is taken

into the front of the engine. Then it is compressed to achieve 3 to 12 times more

than its original pressure by a compressor. It passes into a combustion chamber

where fuel is added to the air. There it is ignited and the temperature is raised to

between 1,100°F to 1,300° F. The hot air is then pushed through a turbine, which

is used to drive the compressor. For a typical turbojet engine, the pressure at the

turbine discharge is nearly twice the atmospheric pressure, this high pressure gas

can be sent to the nozzle, where the velocity of the gas can be increased. In order to

increases the thrust, an afterburner can be placed after the turbine and before the

nozzle. This is basically another combustion chamber and it can substantially

increase the gas temperature before the nozzle. This increases by about 40 percent

in the thrust at takeoff and by a much larger percentage at high speeds once the

plane is in the air. The turbojet engine is a reaction engine. It sucks air in and

compresses it. The gas then passes through the turbine and escapes from the back

of the engine.

Turboprops:

30


The turboprop engine is a jet engine, which is attached to a propeller. Hot gases

pass through the turbine and the turbine is turned. The propeller is then turned by

the gas turbine by means of a drive shaft. It is very similar to the turbojet engine,

the turboprop engine consists of a compressor, combustion chamber, and turbine.

The turbine is turned by the passing gases, and then the turbine is used to drive the

compressor and propeller. Compared to a turbojet engine, the turboprop has higher

propulsion efficiency at flight speeds below about 500 miles per hour. In the

modern turboprop engine, in order to gain high efficiency at high speed, the

propellers have a smaller diameter but use a larger number of blades. To adapt to

the higher flight speeds, scimitar-shaped blades with swept-back leading edges at

the blade tips are used. Nowadays, turboprop engines are used in some small

airliners and transport aircraft.

Turbofans

The turbofan engine is a jet engine with a large fan at the front. The fan sucks in air

and most of the air flows around the outside of the engine, which make it operate

quietly and provides more thrust at low speeds. Nowadays, most airliners are

powered by turbofan engines. Compared to the turbojet, the turbofan engine has

31


many advantages. In a turbojet all the air passes through the compressor,

combustion chamber, and turbine. In a turbofan engine only a proportion of the

incoming air goes into the gas generator. The rest of the air is directly ejected out

of the engine, or mixed with the gas generator exhaust to produce a "hot" jet. The

aim of this system is to increase the thrust without increasing fuel consumption. It

achieves this by increasing the total mass of air that passes through the engine and

reduces the velocity within the same total energy supply.

Turbo shafts:

The Turbo-shaft engine is another form of gas-turbine engine, which is widely

used in helicopters. It operates like a turboprop system. However, it does not have

a propeller but drives the helicopter rotor instead. The turbo-shaft engine is

designed to keep the speed of the helicopter rotor independent from the rotating

speed of the gas generator. It allows the rotating speed of the rotor to remain

constant even when the rotating speed of the generator is varied to adjust the

amount of power it produces.

Ramjets

The Ramjet engine is the simplest jet engine. It has no moving parts. It is

essentially a turbojet engine without the rotating machinery inside the engine. So

32


its compression ratio depends wholly on its forward speed. Because of this fact, it

can not produce static thrust and it produces very little thrust, when the speed is

below the speed of the sound. Consequently, a ramjet vehicle cannot take off by

itself. So, other means, such as another aircraft may be needed to help it to take off.

This engine is used in guided-missile system, and space vehicles.

4.1. Location

The engines are padded under the wing of the aircraft.

4.2. Thrust Loading

Where,

0 - estimated weight

T/W-optimum thrust loading value

Total thrust required =330325.710.3 9.81

33


=972148.56 N

=99097.71kg

Total thrust required = 218473.05lbf

For single engine the thrust is 218473.05lbf,

After examination of available engine, meeting our requirement have been short listed and the engine “GE90-115B” was chooses to be used in this design.

4.4. Padded Engines

Advantages

There engines produce less noise in the cabin because the engine and exhaust are away from the fuel age.

It has higher wetted area than build engine and jet exhaust can be directed downward by flaps which greatly increases lift and short takeoff.

Disadvantages

Increases the drag due to presence of pylons.

34


UNDER WING Advantages

Length of landing gear can be less. Lateral stability is more.

Disadvantages

Ground clearance is low.

ENGINE CONFIGURATION:

35

Wei

ght

1826

0

Wid

th/

Dia

met

er(i

nch)

134

Len

gth

(inc

h)

192.

8

Fan

D

iam

eter

(inc

h)

123

SF

C

0.32

4

Thr

ust

1150

00 lb

f

Mod

el

GE

90-

110B

Man

ufac

ture

Gen

eral

Ele

ctri

cs


5. AIRFOIL SELECTION

We have to keep in mind that the airfoil of our flying surfaces is only one

variable of the many components which makes our airplanes fly well - or not so

well - in a range of possible configurations. When we do an investigation of any

part of our aircraft we must not look at this part as THE solution, rather we must

always remember that it is only one part of a whole. Analysis is necessary; but

only a synthetic view will give us the whole picture. It is a bit like somebody trying

to understand the human body by studying the skeleton only, or the chemicals of

36


the body only, etc.: the failure of modern medicine comes from this fact. Scientists

look at the parts of a corpse and decide they know something about a living body!

But, let us go back to something less serious (!?!) and look at the airfoil or wing

section of our airplane in such a way that we will have a little better understanding

of how our aircraft flies.

Relative Motion

Today it is universally accepted that an airfoil in motion through still air and air

blowing over a stationary airfoil have the same effects. This was not the case in

scientific circles some 120 years ago, but now is common knowledge, and justifies

the wind tunnel tests where true air flows over an airfoil and from which we can

predict characteristics of an airplane moving through still air. The important thing

is the relative speed of airfoil and air.

Reynolds Numbers

Early investigations into the theory of fluid dynamics have predicted a certain

number of constants to which similar disturbances (and an airfoil in the air is a

disturbance) produce similar effects - in hydrodynamics, these are referred to as

'Froude Numbers" (hulls of boats); in high speed aerodynamics the "Mach Number'

are other examples. For our smaller and slower aircraft, the only "number" which

really needs to be considered is the "Reynolds Number" and it is defined as:

Re = V x I / v

Where:

V = Relative speed (m/sec)

37


I = typical "length" of a solid body (M)

v = kinematic viscosity of the air (sec/m2)

Re is a dimensionless number, which makes it independent of the measuring

systems. The kinematic viscosity is to a certain extent dependent on the density of

the air, but for our aircraft flying below 12,000 ft., it can be assumed constant

(equivalent to 15 x 106 sec/m2 in metric).

The speed can easily be converted to metric:

1 mph = 1.15 Kts. = 1.61 km/h = 1.61 / 3.6 m/s = .45 m/sec.

The same applies to the length:

1 ft. = .305 m.

Our small aircraft have a wing chord, which is the "length" to use when talking

about airfoils, of some 5 ft. equivalent to 1.5 m.

Thus the Reynolds number simplifies to:

Re = (.45 x vmph x 1.5) / (15 x 10-6) = 4.5 vmph

or at stall speed of 50 mph: Re = 1.8 x 106 (you know that 106 = 1,000,000 = 1

million).

Keep in mind the above values are for a 5 ft. chord. For a 2-1/2 ft. chord typical of

tail surfaces or the tip of a tapered wing, the Re will be only 1/2 above values.

38


If the air is looked at, not as a continuous medium, but composed of small balls

(the molecules of modern physics), there is obviously an average distance between

those balls. The Reynolds number is then nothing else than the relation between

the typical solid body length to this average distance between the molecules of the

air in which the solid is moving.

As long as this Reynolds number is between the values of .4 x 106 (400,000) and

some 10 X 106 (ten million) what we will say about airfoils will apply.

Note that for smaller Re (say 10,000 to 400,000, which is the range for radio

controlled models and smaller windmills), other lows apply; however, we will not

consider these numbers in this present set of articles which deal with light planes.

The same applies at very large Reynolds numbers, which are practically associated

with Mach numbers larger than .3, where the compressibility of the air can no

longer be neglected as it is in classic aerodynamics which considers the air as an

incompressible, continuous medium.

Boundary Layer

When the air hits the airfoil leading edge it will separate into the upper and lower

airstream, which meets again at the trailing edge.

39


It is obvious that the air very close to the airfoil "rubs" against the solid surface and

is slowed down. In other words, starting downstream of the impact point, the air

loses some of its momentum, or velocity. And it loses more and more as we follow

it along the path close to the solid airfoil. We can see that friction creates an area

where there is less speed. The reduced speed area just outside of the airfoil

becomes thicker and thicker as we follow it from the leading edge to the trailing

edge. This area is called the boundary layer. Its thickness is increasing as described

and is defined as the thickness at which the local free stream speed is finally

reached. A typical boundary layer thickness is 1/2" near the trailing edge. The

friction, which obviously, is a loss, results in the friction drag of the airfoil.

40


Again the theory of fluid dynamics shows that there are two possible types of

stable boundary layers:

1. The first, to build up, is called 'laminar" because the flow is nice and steady

and the friction drag is relatively low.

2. The second is called 'turbulent" because the flow is rather rough and the

friction drag is higher.

The unfortunate thing is that the "laminar boundary layer" will automatically

become turbulent (with associated higher drag) close to the leading edge of the

airfoil unless very special precautions are taken. These precautions are:

a. A very smooth airfoil surface: Slight construction defects (or bugs as they

stick to the airfoil leading edge) will change the laminar boundary layer into

a turbulent one. Unless you have a perfect airfoil and keep it this way forget

about the gain possible with a laminar flow!

b. A special shape of the airfoil: The pressure distribution on the airfoil is

related to the airfoil shape. Today we can calculate (with high speed

computers) airfoils which maximize the length of the laminar boundary

layer. Still, what is mentioned in a) applies. But, do not get desperate. The

friction drag of the airfoil with a laminar boundary layer is .08, whereas in

turbulent flow it becomes .12. Sure, this is a 50% increase but only on the

friction drag of the airfoil. The other drag contributions are airfoil shape,

wind induced drag, tail drag, fuselage and landing gear drag, interference

drag, cooling drag and a few more. Your aircraft will never go 50% faster

just by changing the airfoil - at the very best, you may gain a few (3 to 5)

percentage points.

41


5.1. CALCULATION OF CL:

5.1.2. COEFFICIENT OF LIFT: ( )

CL max=[ W

S]

q

q = 12

ρatmV stall2

V stall=0.25∗V cruise

= 0.25 * 248

V stall = 62.5 ms

At 1800 altitude

T = 272.57 K

P = 7.563*104 N

m2

𝜌 = 1.02 Kg

m3

q = 12

ρatmV stall2

q = 0.5 * 1.02* 62.52

q = 1922.2 Kg

m3

WS optimum = 4074

N

m2

42


CL max=[ W

S]

q

CL max = 4074

1922.2 CL max = 2.112

Coefficient of drag

CD = CDO + K CL2

K = 1

π∗e∗AR

K = 1

π∗9.3∗0.77

K = 0.04445

CD =0.219

5.1.4. Calculation of

ΔC Lmax= 2.112 – 1.745 ΔC Lmax = 0.367

43


This extra lift can be obtained by the use of flap. Our required is

0.367. Hence we can use split flap which meets our lift requirement also spoilers.

NACA 63-215 AIRFOIL CO-ORDINATES:

44

UPPER SURFACE LOWER SURFACE


45

0.000000 0.000000 0.003990 0.012500 0.006370 0.015280 0.011200 0.019800 0.023480 0.027920 0.048290 0.039600 0.073230 0.048470 0.098230 0.055690 0.148340 0.066820 0.198520 0.074870 0.248750 0.080490 0.299000 0.083920 0.349260 0.085300 0.399520 0.084570 0.449770 0.081940 0.500000 0.077680 0.550190 0.072030 0.600350 0.065240 0.650470 0.057510 0.700530 0.049060 0.750550 0.040140 0.800510 0.031050 0.850430 0.022130 0.900300 0.013680 0.950140 0.006160 1.000000 0.000000 0.000000 0.000000 0.006010 -0.011500 0.008630 -0.013880 0.013800 -0.017660 0.026520 -0.024200 0.051710 -0.033280 0.076770 -0.039990 0.101770 -0.045350 0.151660 -0.053360 0.201480 -0.058950 0.251250 -0.062590 0.301000 -0.064480 0.350740 -0.064700 0.400480 -0.063150 0.450230 -0.060040 0.500000 -0.055620 0.549810 -0.050130 0.599650 -0.043820 0.649530 -0.036910 0.699470 -0.029620 0.749450 -0.022240


DIAGRAM OF NACA63-215

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6. WING SELECTION

Equivalent Aspect RatioAspect ratio = a× Mc

Aspect ratio = 6.5*(0.84)

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Aspect ratio = 9.3

W/S=415.29 N/m2

S=330325.71/415.29 S=795.41m2

Aspect ratio =b/s b2=AR×S

b=√(AR×S)

b=√(9.3×795.41)

b=86.007 b/c=8.9m cr=¿ ¿8.94/8.9 cr=1.00m

Taper ratio=0.394 c t/cr=0.41 c t=0.41×cr

= 0.41m

= 2 cr (1+λ+ λ 2 ) 3 1+ λ

=2×13.21 (1+0.4+0.4 2 ) 1+0.4

= 9.81 m

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Cr=2b

A . R (1+ λ)

=2∗86.007

9.3(1+0.4 )

=13.21 m

Ct/Cr=0.4

Ct=0.4*Cr

=0.4*13.21

=5.284 m

= 2 cr (1+λ+ λ 2 ) 3 1+ λ

=2×13.21 (1+0.4+0.4 2 ) 3 [1+0.4]

= 9.8131

From historical data

λLE=350 (Leading edge sweep angle)

CALCULATION OF THICKNESS TO CHORD RATIO:

Volume of fuel

Volume of fuel = weight of fuel/800

Wf /W0 = 0.374 Wf =123541.8155 kg

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Volume of fuel Vf,

= 154.42 m3

V =[ [tc]* * 0.5 * * 0.5 b * 0.75] * 2

tc = 0.0497

tr/Cr=0.0497

tr=0.6565 m

tt/Ct=0.0497

tt= 0.2626 m

CALCULATION OF CENTRE OF GRAVITY(C.G)

X=Ct2 =

5.2842

= 2.642

Y=b6 (

1+2 λ1+ λ )

=86.007

6 (1+2∗0.4

1+0.4 )

=18.43 m

C.G of wing(2.642, 18.43)

7.CALCULATION OF WETTED SURFACE AREA:

FRONTAL AREA:

Frontal area=π4 d2

50


=π4 *6.0722

=28.95m2

Length of the fuselage lf=70.08m

7.1 WING AREA:

Wing area

Wing area,

S=86.007*.6565

=56.46 m2

7.2 HORIZONTAL TAIL:

Bht=√Sh t∗A . R

=√99.20∗9.3

=30.37m

tht/Cht=.1823

Sht=bht*Chr[λ+¿)]

99.20=21.259Chr

Chr=4.666m

=0.85061*30.37

=25.83m2

VERTICAL TAIL:

Svt = tvt *bvt

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bvt=√Svt∗A . R

svt=53.67m2

tvt/cr = 0.1823

bvt=√53.67∗9.3

bvt=22.34m

tvt=0.1823*3.98

=0.7555

svt = 0.7255*27.3

=16.20m2

ENGINE:

= π4 * 3.542

=9.84m2

For two engines

=2*9.84

=19.68m2

Under carriage or landing gear:

Assuming 90% of engine area for main landing gear

So, = 0.9*19.68m2

= 17.712m2

Neglect the drag of the nose wheel landing gear

¼ of the flap

Sπ = θ

360 * π*r2

r = 0.2*Cr

=0.2*13.21

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=2.642m

Sπ = 15

360 *π*2.6422

Sπ=0.9137m2

Full flap

(Wetted area where full flap is deflected)

Sπ = θ

360 * π*r2

r =2.642

θ=38.5o

Sπ=2.345m2

8. THREE VIEWS OF THE AICRAFT :

8.1 FRONT VIEW

53

SL:N

O

WETTED AREA

1 Fuselage 0.03 28.95 0.8685

2 Horizontal tail 0.008 56.46 0.45168

3 Vertical tail 0.008 25.86 0.20688

4 Wing 0.008 16.20 0.1296

5 Engine 0.01 17.712 0.17712

6 Landing gear 0.04 19.68 0.7872

7 ¼ flap 0.035 0.9137 0.03197

8 full flap 0.0504 2.345 0.1181


8.2 SIDE VIEW

8.3 TOP VIEW

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MATERIAL SELECTION

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Several factors influence the choice of a material for different parts of an aircraft. High Strength to weight is the chief among them. Other factors include stiffness, toughness, resistance to fatigue, corrosion resistance, ease of fabrication, availability, consistency of supply and of course cost. The main groups of materials used have been wood, steel, aluminium alloys, titanium alloys and fibre reinforced composites. Let us have a bird’s eye view of the different categories of materials used.

WOOD:The first generation of aircrafts was fabricated with wood and canvas. The

strength to weight ratio of the Spruce and birch varieties of wood used was moderately high and equal to that of the present day heat treated aluminium alloys. The effect of moisture and humidity made the use of wood less advisable as it caused inconsistency in the properties of the material. Changes in shape and dimensions also resulted. Though wood was made use in the manufacture of wing spars for its good properties, the increased wing loadings and complex structural forms of turbo jets has brought its usage to an end.

STEEL:Steel delivered high modulus of elasticity, high proof stress and high tensile strength to the manufacturer. However, it exhibited very high specific gravity which limited its usage. Thin walled, box section spars were fabricated using steel. Carbon present in steel though produces necessary hardening, causes brittleness and distortion. So, a new family called maraging steels were manufactured involving either no or very less carbon content in it. Typical maraging steel would have these elements present in the proportions: nickel 17-19%, cobalt 8-9%, molybdenum 3-3.5% and titanium 0.15-0.25%. The cost of manufacture of maraging steel is very high, about three times that of the conventional one. Arrestor hooks, rocket motor casings, helicopter under carriages, gears and ejector seats are few components manufactured using maraging steel.

ALUMINIUM ALLOYS:The three major groups of aluminium alloys used for airframe construction

56


are Nickel free duralumin, derivatives of Y alloy and the aluminium-zinc-magnesium group. The type of alloy used varies for different requirements of the aircraft and also on the type of aircraft used. But the major disadvantage of aluminium alloys is that one property is increased by sacrificing many other properties. For instance, the duralumin alloys possess a lower static strength than the zinc-bearing alloy, but are preferred for portions of the structure where fatigue considerations are of primary importance such as the under-surfaces of wings where tensile fatigue loads predominate.

TITANIUM ALLOYS:Titanium alloys are mostly used in combat aircrafts than in transport

aircrafts. They possess high fatigue strength to tensile strength ratio, good corrosion and fatigue resistance. But exposure to temperature and presence of salt environment greatly affect these properties. Moreover high density imposes weight constraints on the material.

COMPOSITE MATERIALS:Composite materials consist of strong fibers such as glass and carbon set

in a matrix of plastic or resin. They are mechanically and chemically protective. They have very high strength to weight ratios. Weight saving is a major advantage while using composite materials. However, failure of a composite is not clearly defined yet and also repair of this class of materials is still a topic under study. This is an emerging class of materials.

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Material AISI alloy steel 4340

5 Cr-Mo-V Steel

6A1-4V Titanium alloy

Inconel X Nickel Alloy

8 Mn Titanium alloy

A261A Mg Alloy

7075 Al Alloy

Ftu ksi 260 280 130 155 120 39 79Fty ksi 217 240 120 100 110 24 69Fcu ksi 242 260 125 105 110 14 69Fsy ksi 149 170 80 108 84 19 47e % 10 7 10 20 10 9 6E *106 psi

29 30 16 31 15.5 6.3 10.3

Ec *106 psi

29 30 16.4 31 16 6.3 10.5

w lb in-3 0.283 0.281 0.16 0.304 0.171 0.0647 0.101Form Bar Bar Bar Sheet Sheet,

plateExtruded bar

Sheet, plate

From the above table, based on the strength to weight ratio, 7075 Al Alloy is the best suited material for the wing spar design as well as the aircraft skin.

DETAILED WING DESIGNSPAR DESIGN:

Spars are members which are basically used to carry the bending and shear loads acting on the wing during flight. There are two spars, one located at 15-25% of the chord known as the front spar, the other located at 60-70% of the chord known as the rear spar. Some of the functions of the spar include:

They form the boundary to the fuel tank located in the wing. The spar flange takes up the bending loads whereas the web carries the

58


shear loads. The rear spar provides a means of attaching the control surfaces on the

wing.

Considering these functions, the locations of the front and rear spar are fixed at 0.175c and 0.758985c respectively. The AG04 airfoil is drawn to scale using any design software and the chord thickness at the front and rear spar locations are found to be 0.356m and 0.134m respectively.

The spar design for the wing root has been taken because the maximum bending moment and shear force are at the root. It is assumed that the flanges take up all the bending and the web takes all the shear effect. The maximum bending moment for high angle of attack condition is 750527.6785 Nm. the ratio in which the spars take up the bending moment is given as

(Mfr/Mr) = (h12/h2

2) = (0.3562/0.1342) = 7.05814 Mfr+Mr = 750527.6785 NmFrom the above two equations, Mfr = 657388.6069 Nm, Mr = 93139.07161 NmThe yield tensile stress σy for 7075 Al Alloy is 455.053962 MPa. The area of the flanges is determined using the relationσy = Mz/(A*z2)where M is bending moment taken up by each spar,

A is the flange area of each spar,z is the centroidal distance of the area = h/2.

Using the available values, Area of front spar Afr = 81.15948 cm2, Area of rear spar Ar = 30.5488 cm2

Each flange of the spar is made of two angle sections. For the front spar, the length of the angle is 6t, angle height is 5t with angle thickness t. Area for each angle of front spar is found to be 20.288 cm2 and hence value of t is found to be 1.455947 cm. Front spar - Dimensions of each angle:Length = 8.735682 cmHeight = 7.27974 cm Thickness = 1.455947 cm.

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CONCLUSION:

We have come to a completion of the conceptual design of an

aircraft. Aircaft design involves a variety of faculties of the field of Aeronautical

engineering structures, performance, aerodynamics, stability etc. this project has

enabled us to get a taste of what it is to design a real aircraft. The fantasies of the

flying world seem to be much more than what we thought. With this design

project as the base, we will strive to progress in the field of airplane design and

maintenance. We convey our heartfelt gratitude to all of them who have

provided their helping hand in the completion of this project.

BIBLOGRAPHY:

REFERENCES:

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Aerodynamic design:

1. Jane’s All the world’s aircraft

2. Aircraft design – a conceptual approach – Daniel P. Raymer

3. Design of aircraft – Thomas Corke

4. Aircraft Performance – J.D. Anderson

5. Aircraft performance, Stability and control – Perkins and Hage

6. Fluid dynamic Drag - Hoerner

7. Summary of airfoil data – Abbott, Doenhoff and Stivers

8. www.airliners.net

9. www.wikipedia.org

10. www.aerospaceweb.org

Structural design:

1. Analysis of Aircraft structures – Bruhn

2. Aircraft Structures for engineering students – T.H.G Megson

3. Aircraft structures – Peery and Azar

4. Airplane design – Jan Roskam

5. Airframe Stress Analysis and Sizing – Niu

1. Analysis of Aircraft structures – Bruhn

2. Aircraft Structures for engineering students – T.H.G Megson

3. Aircraft structures – Peery and Azar

4. Airplane design – Jan Roskam

5. Airframe Stress Analysis and Sizing – Niu

www.NASA.org

61

http://www.NASA.org/


www.ZAP16.com

www. AIRLINERS.COM

Few websites followed,

www.Propulsion.org

www.ADL.GETCH.edu

www.wikipedia.org

62

http://www.wikipedia.org/

http://www.ADL.GETCH.edu/

http://www.Propulsion/

http://www.combataircraft.com/

http://www.ZAP16.com/





ANOOJ.M (32207101001) who carried out the project work under my

supervision.

SIGNATURE SIGNATURE

SARAVANAN.G CHOCKAPPAN.N






63





DILIP SANKAR.S (32207101006) who carried out the project work under my

supervision.

SIGNATURE SIGNATURE







64





DURAI KESAVAN.D (32207101007) who carried out the project work under my

supervision.

SIGNATURE SIGNATURE








65




ARIF AHAMED.B (32207101002) who carried out the project work under my

supervision.

SIGNATURE SIGNATURE








66




CHANDRA MOHAN.R (32207101004) who carried out the project work under

my supervision.

SIGNATURE SIGNATURE







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Documents

400 Seater Passenger Aircraft