MULTIROLE FIGHTER AIRCRAFTS

Aircraft Design Project- I

SUBMITTED BY

CHINNAMUTHU M 720711101028

CHINNARAJA A 720711101029

DEIVAMOORTHY B 720711101030

DELHI DURAI G 720711101031

DEPARTMENT OF AERONAUTICAL ENGINEERING

HINDUSTHAN COLLEGE OF ENGINEERING & TECHNOLOGY

COIMBATORE.

HINDUSTHAN

COLLEGE OF ENGINEERING AND TECHNOLOGY

COIMBATORE - 641 032.

DEPARTMENT OF AERONAUTICAL ENGINEERING

Certified that this is the bonafide record of work done by DEIVAMOORTHY B

in the Aircraft Design Project-I (AE 2356 ) of this Institution, as prescribed by the

Anna University for the Sixth Semester during the year 2013-2014.

Place: Coimbatore

Date:

Staff in-charge Head of the Department

University Register no 720711101030

Submitted for the Aircraft Design project - I ( AE2356) Practical Examination of the Anna

University conducted on …08.04.2014…………

Internal Examiner External Examiner

WHEN YOU DESIGN AN AIRPLANE ……… THINK ABOUT HOW YOU WOULD FEEL

IF YOU HAD TO FLY IT! SAFETY FIRST. Sign on the wall of the design office at Douglas

Aircraft Company, 1992.

Aircraft design is an evolutionary process rather than a revolutionary process. Thanks to

Sir George Cayley who is a milestone in the evolutionary process. If anyone wants to design an

aircraft without taking any help from previous designs, it will be a one of two extremes, one a

success with the hectic and long process or a failure even after long duration.

Airplane design is an art and a science. In that respect it is difficult to learn by reading a

book. Airplane design the intellectual engineering process of creating on paper a flying machine

to meet certain specification and requirements established by potential users or to pioneer

innovative, new ideas and technology, like the aircraft to be designed here.

An example of the former is the designer of most commercial transports, starting at least

with the Douglas DC-1 in 1932, which was designed to meet or exceed various specifications by

an airplane company.

An example of the later is the design of Rocket- powered Bell X-1, the first airplane to

exceed the speed of sound in level of climbing or level flight on October 14, 1947. The design

process is indeed an intellectual activity, but a rather a special one that is tempered by good

intuition developed via experience, by attention paid to successful airplane designs that have been

used in past, and by design procedures and databases that are a part of every airplane

manufacturers.

So there is a need to conduct a literature survey related to what sort of aircraft is going to

be designed.

The project is centered towards a design of safe jet transport. The objective of this project

is to provide a better design by manipulating the previous designs.

TABLE OFCONTENTS

Sl.

No.

Date Exercise Name Marks Signature

1 23.1.14 Literature survey

2 23.1.14 Comparative Plots

3 30.1.14 Weight estimation

4 06.2.14 Engine selection

5 20.2.14 Airfoil selection

6 27.2.14 Wing design

7 27.2.14 Wetted area calculation

8 06.3.14 Drag polar

9 20.3.14 Drag calculation

10 27.3.14 Thrust required calculation

11 27.3.14 Rate off climb calculation

LISTOF SYMBOLS

R -Range

V -Velocity

C -specific fuel consumption

E -Loitering time

L/D -lift to drag ratio

𝑉𝑎𝑙𝑡 -Velocity at altitude

𝜌𝑎𝑙𝑡 -Density at altitude

S - wing surface area

b - wing span

µ𝑎𝑙𝑡

-coefficient of viscosity at altitude

𝐶𝐻𝑇 -Horizontal tail volume coefficient

𝐿𝐻𝑇 - Horizontal tail arm length

𝑆𝐻𝑇 - Horizontal tail area

𝑆𝑊 -Wing area

𝐶𝑊 -Wing mean chord

LVT -Vertical tail arm length

SVT –Vertical tail area

CVT -Vertical tail volume coefficient

bW -Wing span

SW -Wing area

VTO - Vertical take-off distance

STO - Take-off distance

FTO - Take-off thrust

VA - Approach Velocity

Sπ – Wetted area

Λ-Sweep angle

λ- Taper ratio

INTRODUCTION

BASIC DESIGN PROCESS:-

An airplane design is both an art and a science. Airplane design is an intellectual

engineering process of creating on paper a flying machine to

Meet specifications established by users

Pioneer innovative, new ideas and technology.

The design process is an intellectual activity developed via experience, by attention paid to

successful airplane designs that have been used in the past and by design procedures and databases

that are a part of every airplane manufacturer.

PHASES OF AIRPLANE DESIGN:-

From the time when an airplane materializes as a new thought to the time the finished product is

ready, the complete design undergoes three distinct phases in perfect sequences which are

Conceptual design

Preliminary design

Detail design

CONCEPTUAL DESIGN:-

The design process starts with a set of specifications or much less frequently to desire to implement

pioneering. There is a concrete goal where we designers are aiming at. The first step towards it is

conceptual design. Within a fuzzy latitude, overall shape, size, weight are determined for the

potential user.

The product of the conceptual design phase is layout of airplane configuration on paper. This

drawing has flexible lines, which can be slightly changed. However we get a detailed account of

the layout configuration at the end of this phase. The major drivers during the conceptual design

process are aerodynamics, propulsion and flight performance.

Structural and control system considerations are not dealt in detail but however they are not dealt

in detail but however they are not totally absent. The designer is influenced by qualitative aspects.

No part of the design process is carried out in total vacuum unrelated to other parts.

PRELIMINARY DESIGN:-

This phase includes only minor changes to be made in the configuration layout. There is serious

control and structural system analysis and design takes place. During this phase substantial wind

tunnel testing will be carried out and major computational fluid dynamics (CFD) calculations. At

the end of the phase, the airplane configuration is frozen and defined. The drawing process is called

lofting. This process makes precise shape of outside skin of airplane making certain all sections fit

together.

The end of the phase is the decision if the airplane is to be manufactured or not. It is no longer a

critical condition where “you – bet your company” on full scale development of a new airplane.

DETAIL DESIGN:-

This phase is literally the ‘nuts and bolts’ phase of airplane design. The aerodynamic, propulsion,

structures, performance, flight control analysis are over in the preliminary phase. The airplane is

to be fabricated and machined. The size, number and location of rivets, fasteners are determined

now. Flight simulators are developed. At the end of this phase, the aircraft is ready to be fabricated.

THE SEVEN INTELLECTUAL PIVOT POINTS FOR CONCEPTUAL DESIGN:-

The overall conceptual design is anchored by seven intellectual “pivot points” – seven factors that

anchor the conceptual design thought process. They allow different, detailed thinking to reach out

in all directions from each point.

REQUIREMENTS:-

The requirements are given by the people who are going to buy – the customers. For other aircrafts,

these requirements are usually set by the manufacturer in full appreciation of needs of owner.

Requirements of one airplane are different from the other. There can be no stipulated specific

standard. There must be established requirements that serve as impinge off point for design

process. The requirements that are frequently stipulated are:

Range

Takeoff distance

Stalling velocity

Endurance

Maximum velocity

Rate of climb

For dog fighting combat, maximum turn rate and minimum turn radius

Maximum load factor

Service ceiling

Cost

Reliability and maintainability

Maximum size.

SEVEN INTELEECTUAL PIVOT POINTS FOR DESIGN

NO

YES

REQUIREMENTS

WEIGHT OF AIRPLANE –FIRST ESTIMATE

CRITICAL PERFORMANCE PARAMETER

LIFT COEFFICENT (CLMAX)

LIFT-TO-DRAG RATIO

(L/DMAX)

WING LOADING (W/S)

THRUST TO WEIGHT

RATIO(T/W)

CONFIGURATION LAYOUT – SHAPE

ANDSIZE OF AIRPLANE ON DRAWING

BETTER WEIGHT ESTIMATE

PERFORMANCE ANALYSIS-DOES DESIGN

MEET REQUIREMENTS

S

OPTIMIZATION

AIRCRAFT CONCEPTUAL DESIGN PROCESS

BETTER

REQUIREMENTS

NEW CONCEPT

IDEAS

TECHNOLOGY AVAILABLE

CONCEPT SKETCH

FIRST GUESS SIZING

WEIGHTS

COST AERO

WEIGHT

S INITIAL LAYOUT

AERO

PRELIMNARY DESIGN

REFORMED SIZE PERFORMANCE

OPTIMIZATION

ETC

STRUCTURE

S

SIZING AND

PERFORMANCES

OPTIMIZATION

PROPULSION

LANDING GEAR

PROPULSION

REVISED LAYOUT

CRITICAL PERFORMANCE PARAMETERS:-

Requirements stipulate the performance of the new aircraft. The critical parameters are:

Maximum lift coefficient

Lift to drag ratio (L/D)

Thrust to weight ratio (T/W)

Therefore the next step is to make first estimates of W/S and T/W to achieve the performance as

stipulated by requirements.

CONFIGURATION LAYOUT:-

The configuration layout is a drawing of the shape and size of the airplane as evolved till stage.

The critical performance parameters along with first weight estimate helps to draw the

configuration and approximate the size of the aircraft.

BETTER WEIGHT ESTIMATE:-

The overall size and shape of the airplane are better known now. There is now an improved

estimate of weight based on performance parameters. A more detailed estimate of fuel is required

now.

PERFORMANCE ANALYSIS:-

This is the point where the configuration is judged if it can meet all original specifications. An

interactive process is initiated where the configuration is modified. The critical performance

parameters are adjusted for improving performance. In this stage, some mature decisions should

be made as the specifications or cost or unavailable technology.

Hence some specifications might be relaxed so that others might get higher priority.

OPTIMIZATION:-

When iterative process is over, it has produced a viable airplane. This leads to optimization. The

optimization analysis is carried out may be carried out by a systematic variation of different

parameters T/W, W/S and plotting the performance of graphs which can be found using a sizing

matrix or a carpet plot from which optimum design can be found.

WEIGHT OF AIRPLANCE – FIRST ESTIMATE:-

No airplane can take off the ground unless it produces a lift greater than its weight. There

should be a first estimate of gross takeoff weight. The weight estimate is the next pivot point after

the requirements. Lilienthal, Langley and Wright brothers knew more weight means more drag.

This needed an engine with greater power and hence more weight

CONSTRAINT DIAGRAM:-

A constraint diagram is constructed which identifies allowable solution space for airplane design.

A constraint diagram consists of plots of the sea – level thrust to take off weight ratio versus wing

loading attakeoff weight ratioTO/WO versuswing loading at takeoff WO /S determined by

intellectual pivot point.

THE DESIGN WHEEL

SIZING

AND

TRADE

STUDIES

REQUIREMEN

T

DESIGN

ANALYSIS

DESIGN

CONCEPT

CLASSIFICATION OF AIRPLANES

1. FUNCTIONAL CLASSIFICATIONS

a. Civil Airplanes

b. Military Airplanes

Civil Airplanes Military Airplanes

Cargo transport Strategic fighters

Passenger travel Interceptors

Mail distribution Escort fighters

Agricultural Tactical bombers

Ambulance Strategic bombers

Executive transport Ground attack airplanes

Training Photo-reconnaissance airplanes

Sports Multipurpose airplanes

Air taxi & charter

Forestry

Fish and wildlife sanctuary

Construction

Aerial photography

Off- shore drilling

2. CLASSIFICATION BY POWER PLANT

a. Types of engine

i. Piston Engines

ii. Turbo-Prop Engines

iii. Turbo-jet Engines

iv. Ram-jet Engines

v. Rockets

b. Number of engines

i. Single Engine

ii. Twin Engine

iii. Multi-Engine

c. Location of power plant

i. Engine (with propeller) located in fuselage nose

ii. Pusher Engine located in the rear fuselage

iii. Engines (jet) submerged in the wing

1. At the root

2. Along the span

iv. Engines (jet) in nacelles suspended under the wing(pod mountings)

v. Engines (jet) located on the rear fuselage

vi. Engines (jet) located within the rear fuselage

3. CLASSIFICATION BY CONFIGURATION

a. Shape and position of wing

b. Type of fuselage

c. Location of horizontal tail surfaces

d. Types of Landing gear

Exp.No:1 Date: 23.01.2014

LITERATURE SURVEY

It is very easy to design an aircraft if we have data’s about already existing aircrafts of similar

type. It provides more satisfaction and avoids confusion while choosing some design parameters

for our aircraft. In this detailed survey some many important design drivers like aspect ratio, wing

loading, overall dimensions and engine specifications are determined for our reference. It assists

in proposing a new design and modification in our design which will improve the performance of

the proposed aircraft. This assures the performance of the aircraft as per the design calculations

and easy way of designing an aircraft within particular period of time. So in this literature survey

we have collected some ten already existing 20 seated jet transport aircraft for our reference of

design parameters. Mostly these aircrafts have similar characteristics in many designs aspects

which are shown in the table.

GEOMETRIC SPECIFICATIONS

Sl.

No.

Name of theAircraft Aspect Ratio Wing Span

(m)

Length

(m)

Wing Area

(m2)

Wing Loading

(Kg/m2)

1 Chengdu J-10 2.87 9.75 15.49 5.43 33.1

2 EurofighterTyphoon 2.09 11.61 20.83 6.45 64.57

3 F/A-18 Hornet 3.98 12.3 17.1 4.7 38

4

F-16 Fighting

Falcon 3.56 9.96 15.06 4.88 27.87

5 F-35 Lighting II 2.68 10.7 154.67 4.33 42.7

6 HAL Tejas 1.75 8.2 13.2 4.4 38.4

7 JAS 39 Gripen 2.35 8.4 14.1 4.5 30

8 JF-17 Thunder 3.66 9.45 14.93 4.72 24.4

9

Lockheed F-22

Raptor 2.36 13.56 18.9 5.08 78.04

10 MiG-29 3.42 11.4 17.37 4.73 38

11 MiG-29K 3.34 11.99 17.3 4.4 43

12 MiG-29M 3.42 11.4 17.37 4.73 38

13 Mirage 2000 2.03 9.13 14.36 5.2 41

14 Mitsubishi F-2 3.56 11.13 15.52 4.69 34.84

15 Rafale 2.55 10.8 15.27 5.34 45.7

16 Su-27m 3.78 15.3 21.9 5.9 62

17 Su-35 3.78 15.3 21.9 5.9 62

18 Sukhoi Su -47 3.71 15.16 22.6 6.3 61.87

19

Sukhoi T-50 PAK-

FA 3.49 14.7 21.935 6.36 62

20 T-50 Golden Eagle 2.49 14 19.8 6.05 78.8

21 Tornado IDS 7.27 13.91 16.72 5.95 26.6

WEIGHT SPECIFICATIONS

Sl.

No. Name of the Aircraft

Empty

Weight

(Kg)

Gross Weight

(Kg)

Maximum Take-off Weight

(Kg)

1 Chengdu J-10 9,750 14,250 19,277

2 EurofighterTyphoon 8777 11,346 14300

3 F/A-18 Hornet 10400 13,013 23500

4 F-16 Fighting Falcon 8570 11,675 19200

5 F-35 Lighting II 13300 17,490 31800

6 HAL Tejas 6500 9,500 13200

7 JAS 39 Gripen 6800 9,068 14000

8 JF-17 Thunder 6586 9,586 12383

9

Lockheed F-22

Raptor 19700 22,822 38000

10 MiG-29 13380 16,720 22400

11 MiG-29K 18550 20,950 24500

12 MiG-29M 11000 13,100 20000

13 Mirage 2000 7500 10,100 17000

14 Mitsubishi F-2 9527 13,927 22090

15 Rafale 9500 14,200 24500

16 Su-27m 8400 11,215 34500

17 Su-35 8400 14,000 34500

18 Sukhoi Su -47 16375 21,645 35000

19 Sukhoi T-50 PAK-FA 8400 13,490 38800

20 T-50 Golden Eagle 18500 23,300 37000

21 Tornado IDS 13890 17,140 28000

POWERPLANT SPECIFICATIONS

Sl.

No. Name of the Aircraft Type of Engine

Number

of

Engines

Power or Thrust per Engine

(KN)

1 Chengdu J-10 Turbofan 1 79.43

2 EurofighterTyphoon Turbofan 1 44

3 F/A-18 Hornet Turbofan 2 48.98

4 F-16 Fighting Falcon Turbofan 1 76.3

5 F-35 Lighting II Turbofan 1 125

6 HAL Tejas Turbofan 1 53.9

7 JAS 39 Gripen Turbofan 1 54

8 JF-17 Thunder Turbofan 1 49.4

9

Lockheed F-22

Raptor

Turbofan

2 104

10 MiG-29 Turbofan 2 88.26

11 MiG-29K Turbofan 2 88.3

12 MiG-29M Turbofan 2 81.4

13 Mirage 2000 Turbofan 1 64.3

14 Mitsubishi F-2 Turbofan 1 76 15 Rafale Turbofan 2 50.04

16 Su-27m Turbofan 2 86.3

17 Su-35 Turbofan 2 86.3

18 Sukhoi Su -47 Turbofan 2 83.4

19 Sukhoi T-50 PAK-FA Turbofan 2 123

20 T-50 Golden Eagle Turbofan 2 93.1

21 Tornado IDS Turbofan 2 71.53

PERFORMANCE SPECIFICATIONS

Sl.

No. Name of the Aircraft

Maximum

speed

(m/s)

Cruising

speed (m/s)

Service ceiling

(Km)

Range

(Km)

1 Chengdu J-10 2695 386.11 18,000 1149

2 EurofighterTyphoon 2450 588.89 16765 2900

3 F/A-18 Hornet 1190 347.22 15240 2000

4 F-16 Fighting Falcon 2120 670.54 15240 1950

5 F-35 Lighting II 1930 536.43 18288 2220

6 HAL Tejas 1350 383.33 15000 850

7 JAS 39 Gripen 2,204 388.89 15240 1865

8 JF-17 Thunder 1960 544.44 16920 1689

9

Lockheed F-22

Raptor 2410 670.56

19812 2960

10 MiG-29 2400 666.21 18013 1430

11 MiG-29K 2200 610.56 17500 1500

12 MiG-29M 2600 694.5 17500 1600

13 Mirage 2000 2530 649.44 17060 1550

14 Mitsubishi F-2 2469.6 590 18000 834

15 Rafale 1,912 385.83 15,235 3,700

16 Su-27m 2390 375 18000 3600

17 Su-35 2390 510.42 18000 3600

18 Sukhoi Su -47 1717 500 18000 3300

19 Sukhoi T-50 PAK-FA 2500 416.67 17300 3000

20 T-50 Golden Eagle 1770 246.39 14630 1851

21 Tornado IDS 2400 268.06 15240 1390

Exp.No:2 Date:23.01.2014

COMPARATIVE GRAPHS

Speed Vs aspect ratio

0.00

2.50

5.00

7.50

0 200 400 600 800

ASP

ECT

RA

TIO

SPEED (m/s)

SPEED Vs ASPECT RATIO

Speed Vs rate of climb

Speed Vs range

0

150

300

450

0 200 400 600 800

R/C

(m

/s)

SPEED (m/s)

SPEED Vs R/C

Speed Vs altitude

Speed Vs wing loading

0

2000

4000

6000

0 200 400 600 800

RA

NG

E (K

m)

SPEED (m/s)

SPEED Vs RANGE

0

10,000

20,000

30,000

0 200 400 600 800

ALT

ITU

DE

(m)

SPEED (m/s)

SPEED Vs ALTITUDE

Speed Vs b/l

0

200

400

600

800

0 200 400 600 800

WIN

GLO

AD

ING

(K

g/m

2 )

SPEED (m/s)

SPEED Vs WINGLOADING

0.000

0.250

0.500

0.750

1.000

0 200 400 600 800

b/l

SPEED (m/s)

SPEED Vs b/l

RESULT:

From the above comparative graphs and calculation,

1. Velocity Vs Aspect ratio

Velocity =660m/s

Aspect ratio =3.0

2. Velocity Vs Rate of climb

Velocity =640 m/s

Rate of climb=300m/s

3. Velocity Vs Range

Velocity = 650m/s

Range = 2000 Km

4. Velocity Vs altitude

Velocity =650m/s

Altitude =18000 Km

5. Velocity Vs Wing loading

Velocity =642m/s

Wing loading =355Kg/m2

6. Velocity Vs b/l

Velocity=640m/s

b/l=0.63

Average velocity = 647m/s =2.19 Mach

Exp.No:3 Date: 30.01.2014

PRIMARY WEIGHT ESTIMATION

The purpose of this section is to introduce a technique to obtain the first estimate of the maximum

take-off weight for an aircraft before it is designed and built. The word estimation is intentionally

selected to indicate the degree of the accuracy and reliability of the output. Hence, the value for

the maximum take-off weight is not final and must be revised in the later design phases. The result

of this step may have up to about 20% inaccuracies, since it is not based on its own aircraft data.

But the calculation relies on the other aircraft data with similar configuration and mission. Thus,

we are adopting the past history as the major source of the information for the calculation in this

step. At the end of the preliminary design phase, the take-off weight estimation is repeated by

using another more accurate technique.

An aircraft has a range of weights from minimum to maximum depending upon the number of pilots

and crew, fuel, and payloads (passengers, loads, luggage, and cargo). As the aircraft flies, the fuel is

burning and the aircraft weight is decreasing. The most important weight in the design of an aircraft

is the maximum allowable weight of the aircraft during take-off operation. It is also referred to as

all up weight. The design maximum take-off weight (MTOW or WTO) is the total weight of an

aircraft when it begins the mission for which it is designed. The maximum design take-off weight

is not necessarily the same as the maximum nominal take-off weight, since some aircraft can be

overloaded beyond design weight in an emergency situation, but will suffer a reduced performance

and reduced stability. Unless specifically stated, maximum take-off weight is the design weight. It

means every aircraft component (e.g. wing, tail) is designed to support this weight.

The major factor that determines the whole design of aircraft especially the selection of

overall weight, airfoil and power plant of the aircraft.

Total weight of an airplane is given by,

WTO =WC+WPL+WF+WE

Where,

WTO = Design takeoff weight of the aircraft

WC = crew weight

WPL = weight of the payload

WF = weight of the fuel

WE = empty weight

To simplify the calculation, both fuel and empty weights can be expressed as fractions of the total

takeoff weight, i.e., Wf/WO. Equation

WO = WC+WPL+ ( )WTO+( )WTO

This can be solved for WTO as follows:

WTO ─ ( ) WTO ─ ( ) WTO = WC+WPL

WTO = ( )

Now WTO can be determined if (WF/WTO) and (WE/WTO) can be estimated.

These are described below.

WPL=WPASSENGERS+WBAGGAGE

Assuming that each passenger with baggage weight is 90kg then the payload weight is,

W Pay Load = 3000 kg

Assuming that each crew with baggage weight is 90kg then,

W Crew =(1*90 ) = 90kg

So,

Wpl+Wc

W TO = ---------------------------------

1-(W f/WTO) – (WE / WTO )

(90+3000)

= --------------------------------

1-(0.287)-(0.6)

= 27345.13kg

MISSION PROFILE:-

From the figure the various stages of aircraft during mission is as follows,

1 start &warm up

2 Taxiing in the runway

3 Takeoff

4 Climb

5 Cruising

6 Loiter

7 Descent

8 Dush out

9 Drop bombs

10 Strafe

11 Dash in

12 Climb

13 Crusing

14 Decent

15 Landing.

For subsonic jet transport aircraft weight fuel fraction is,

(W15/W0) = ( W1/W0) * ( W2/W1) * ( W3/W2) * ( W4/W3) * ( W5/W4) * ( W6/W5) * ( W7/W6) *

(W8/W7) * ( W9/W8 ) * ( W10/W9) * ( W11/W10 ) * ( W12/W11 ) * ( W13/W12 ) * ( W14/W13 ) * (

W15/W14 )

APPROXIMATE WEIGHT ESTIMATION:

Weight fraction for each profile in mission segment,

For Warm up,

(W1/W0) =0.990

For Taxy,

(W2/W1) =0.990.

For Takeoff,

(W3/W2) =0.990.

For Climb,

(W4/W3) =0.971.

For Cruising,

(W5/W4) = 0.954

For loiter,

(W6/W5) =0.967

For descent,

(W7/W6) = 0.990

For Dush Out,

( W8/W7 ) = 0.951

For Drop Bombs,

( W9/W8 ) = 0.990

For Strafe,

( W10/W9 ) = 0.967

For Dash in,

( W11/W10 ) = 0.954

For Clime,

( W12/W11 ) = 0.971

For Cruise in,

( W13/W12 ) = 0.990

For Decent,

( W14/W13 ) = 0.990

For landing,

(W15/W14) =0.990

Then,

(WF/WTO) = (1-W15/W0))

=0.287

Assume Empty Weight fraction,

So, overall weight,

WPL + WC

W TO = ----------------------------------

1-(W f/WTO) – (WE / WTO )

Approximate Overall weight = 27345.13 kg

RESULT:

Thus the final Takeoff weight of the proposed aircraft was estimated using fuel fraction method

were as follows,

WTO (APPROXIMATE) =27345.13 kg.

Exp.No:4 Date: 06.02.2014

ENGINE SELECTION

• Thrust to weight ratio

• Thrust matching

• Engine rating

• Rubber sizing of the engine

• Number of the engines

Thrust to weight ratio:

T/W directly affects the performance of the aircraft. An aircraft with a higher T/W will

accelerate more quickly, climb more rapidly, reach a higher maximum speed, and sustain higher

turn rates. On the other hand, the larger engines will consume more fuel throughout the mission,

which will drive up the aircraft up the aircraft’s takeoff gross weight to perform the design mission.

T/W is not a constant. The weight of the aircraft varies during the flight as fuel is burned.

Also, the engine’s thrust varies with altitude and velocity (as does the horsepower and propeller

efficiency, (ηp).When the designers speak of an aircraft’s thrust-to-weight ratio they generally refer

to the T/W during sea-level static (zero velocity), standard-day conditions.

T/WTO Ratio for General Aviation- single engine is 0.60

Overall weight of aircraft WTO =27345.13 kg =268.255 KN.

Then,

T=0.60×268.255

=160.95 KN

So, the thrust needed=160.95 KN

From the literature survey the nearest value of the thrust corresponding aircraft is Jet engine

The Jet engine has the following characteristics,

• Thrust per engine =160.95 KN

• Number of engine = 1

• Type of engine = Turbofan

• Total thrust =160.95 KN

RESULT:

• Name of engine selected = Turbofan

• Number of engine = 2

• Total thrust = 160.95 KN

Exp.No:5 Date:20.02.2014

AIRFOIL SELECTION

Wing design:

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

important component of an aircraft, since a fixed-wing aircraft is not able to fly without it. Since

the wing geometry and its features are influencing all other aircraft components, we begin the

detail design process by wing design. The primary function of the wing is to generate sufficient

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

drag (D) and nose-down pitching moment (M). While a wing designer is looking to maximize the

lift, the other two (drag and pitching moment) must be minimized. In fact, wing is assumed ad a

lifting surface that lift is produced due to the pressure difference between lower and upper surfaces.

During the wing design process, eighteen parameters must be determined. They are as follows:

1. Wing reference (or planform) area (SW or Sref or S)

2. Number of the wings

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

4. Horizontal position relative to the fuselage

5. Cross section (or airfoil)

6. Aspect ratio (AR)

7. Taper ratio

8. Tip chord (Ct)

9. Root chord (Cr)

10. Mean Aerodynamic Chord (MAC or C)

11. Span (b)

12. Twist angle

13. Sweep angle

14. Dihedral angle

15. Incidence (iw)

16. High lifting devices such as flap

17. Aileron

18. Other wing accessories

The airfoil, in many respects, is the heart of the airplane. The airfoil affects the cruise speed, take-

off and landing distances, stall speed, handling qualities, and overall aerodynamic efficiency

during all phases of flight. The design of the airfoil is a complex and time consuming process.

Much of the Wright brothers success can be traced to their development of airfoils using a wind

tunnel of their own design, and the in-flight validation of those airfoils in their glider experiments

if 1901-1902. More recently, the low speed airfoils develop by peter Lissaman contributed much

to the success of the man-powered Gosssmer Condor, and the airfoils designed by John Rontz

were instrumental to the success of Burt Rutan’s radical designs.

Cruising Reynolds number (Re) as follows,

Density*Vcr*C

Recr = --------------------------

Viscosityalt

=282.98*106

=Velocity at altitude

= Density at altitude

C = (s/b)

= 5.07m

S = wing surface area

b = wing span

And, from standard air table at altitude 18000 m,

Temp = 216.16 k.

Density = 0.819 kg/m2

= 1.79*10-5

VCR = M * a

a = ( 1.4*287*216.16 )0.5

=294.71 m/s

Aspect ratio of our aircraft=3.0

From the literature survey for that aspect ratio,

Area=77.03 m2

Span=15.2 m

And, c =s/b =5.07m

For the Reynolds’s number approximately, from the THEORY OF WING SECTION by ABBOT

following data can be obtained.

Airfoil type Maximum lift coefficient Minimum drag coefficient

NACA 63-006

NACA 63-009

NACA 63-206

NACA 64-006

NACA 64-009

0.83

1.18

1.02

0.83

1.119

0.004

0.0042

0.04

0.038

0.038

Fig: NACA 64-009 Aerofoil

RESULT:

From the above analysis NACA 64-009 series type airfoil was selected for our aircraft design.

Exp.No:6 Date:27.02.2014

Wing and Tail Calculations

Fuselage:

Once the takeoff gross weight has been estimated, the fuselage, the wing. And tail can be sized.

Many methods exist to initially estimate the required fuselage size. For certain types of aircraft,

the fuselage size is determined strictly by “real world constraints”. For example, a large passenger

aircraft devotes most of its length to the passenger compartment. Once the number of passengers

is known and the number of seats across is selected, the fuselage length and diameter are essentially

determined.

Wing:

Actual wing size can now be determined simply as the takeoff weight divided by takeoff wing

loading. Remember that this reference area of the theoretical, trapezoidal wing, and includes the

area extending into the aircraft center line.

Tail Volume Co-efficient:

For the initial layout, the historical approach is used for the estimation of the tail size. The

effectiveness of a tail in generating a moment about the centre of gravity is proportional to the

force produced by the tail and to the tail moment arm. The primary purpose of the tail is to counter

the moments produced by the wing.

1. Length of fuselage:

LFU = a woc

= 15.2/0.63

= 24.13 m.

2. Surface area:

Aspect ratio of our aircraft=3.0

From the literature survey for that aspect ratio,

Area=77.03 m2

Span=15.20 m.

3. Taper ratio

Taper ratio is defined as the ratio between the tip chord (Ct) to the root chord (Cr). This

definition is applied to the wing, as well as the horizontal tail, and the vertical tail.General, the

taper ratio varies between zero and one. 0 ≤ λ ≤ 1

The taper ratio can be defined as,

λ=tip chord

root chord

And the value for the taper ratio in general from design book is0.4

So, C root chord =2s

b(1+𝛌)

=2×77.03

15.2(1+0.3) =7.796 m.

And, Ctip chord = λ× C root chord

=2.338 m.

4. Aerodynamic mean chord:

𝐶𝑊=2

3× C root chord×(

1+λ+λ2

(1+λ))

= 2

3×7.796×

1+0.3+0.32

(1+0.3)

𝐶𝑊=5.56 m.

Location of mean chord is, x = 0.25x5.56 = 1.39 m.

And, y = b

6

(1+2λ)

(1+λ)

= 15.2

6

(1+0.6)

(1+0.3)

=3.117 m.

5. Vertical and horizontal volume coefficient:

CHT =𝐿𝐻𝑇×𝑆𝐻𝑇

𝐶𝑊×𝑆𝑊

Where,

𝐶𝐻𝑇-Horizontal tail volume coefficient

𝐿𝐻𝑇- Horizontal tail arm length

𝑆𝐻𝑇- Horizontal tail area

𝑆𝑊 -Wing area

𝐶𝑊 -Wing mean chord

Since,𝐿𝐻𝑇is 25% of the fuselage length,

𝐿𝐻𝑇= 0.25×𝐿𝐹𝑈

= 0.25×24.13

= 6.0325 m.

For our design,

𝑆𝑊=77.03m2

𝐶𝑊=5.56 m.

From “Aircraft design: A Conceptual approach” by Daniel.P.Raymer 3rd Ed,

𝐶𝐻𝑇=0.40 So,

SHT=𝐶𝑊×𝑆𝑊×𝐶𝐻𝑇

𝐿𝐻𝑇

SHT= 0.40×5.56×77.03

6.0325= 28.39 m2

And,

𝐶𝑉𝑇=𝐿𝑉𝑇×𝑆𝑉𝑇

𝑏𝑊×𝑆𝑊

Where,

LVT -Vertical tail arm length

SVT –Vertical tail area

CVT -Vertical tail volume coefficient

bW -Wing span

SW -Wing area

Since,𝐿𝑉𝑇is 50% of the fuselage length,

𝐿𝑉𝑇= 0.5×𝐿𝐹𝑈

= 0.5×24.13

=10.8585 m.

For our design,

𝑆𝑊= 77.03m2.

𝑏𝑊= 15.2m.

From “Aircraft design: A Conceptual approach” by Daniel.P.Raymer 3rd Ed,

𝐶𝑉𝑇=0.07 m.

So,

𝑆𝑉𝑇=𝑏𝑊×𝑆𝑊×𝐶𝑉𝑇

𝐿𝑉𝑇

= 15.2×77.03×0.07

10.8585

= 7.547 m2

Fig: Geometry of wing

RESULT:

The dimensional parameters are,

Wing span, bw=15.2m

Wing area, Sw=77.06m2

Root chord, Cr=7.8 m

Tip chord, Ct=2.34m

Mean aerodynamic chord length, Cw=5.56m

Horizontal tail Surface, SHT=28.39m2

Vertical tail surface, SVT=7.547m2

Exp.No:7 Date:27.02.2014

Wetted area calculations

Aircraft wetted area (Swet), the total exposed surface area, can be visualized as the area of the

external parts of the aircraft that would get wet if it were dipped into water. The wetted area must

be calculated for drag estimation, as it the major contributor to friction drag.

The wing and tail wetted areas can be approximated from their platforms. The wetted area is

estimated by multiplying the true view exposed plan form area is estimated by multiplying the true

view exposed planform area (S exposed) times a factor based upon the wing or tail thickness ratio.

If a wing or tail were paper –thin, the wetted area would be exactly twice the true plan form area.

The effect of finite thickness id to increase the wetted area, as approximated by the following

equations.

Note that the true exposed plan form area is the projected area divided by the cosine of the dihedral

angle.

If t/c ˂ 0.05,

S wet =2.003 S exposed

If t/c ˃0.05,

S wet= S exposed [1.977 + 0.52(t/c)]

The exposed area can be measured from the drawing in several ways. A professional designer will

have access to a “planimeter” a mechanical device for measuring areas. Use of the planimeter is a

dying art as the computer replaces the drafting board. Alternatively the area can be measured by

tracing onto graph paper and “counting squares”.

The wetted area of the fuselage can be initially estimated using just the side and top views of the

aircraft. The side and top view projected areas of the fuselage are measured from the drawing, and

the values are averaged.

For a long, thin body circular in cross section, this average projected area times Π will yield the

surface wetted area. If the body is rectangular in cross section, the wetted area will be four times

the average projected area. For typical aircraft the following equation provides a reasonable

approximation.

S wet=3.4 [(A top + A side) / 2) ]

A more accurate estimation of wetted area can be obtained by graphical integration using a number

of fuselage cross sections. If the perimeters of the cross sections are measured and plotted Vs

longitudinal locations, using the same units on the graph, then the integrated area under the

resulting curve gives the wetted area.

Perimeters can be measured using a professional’s “map-measure,” or approximated using a piece

of scrap paper. Simply follow around the perimeter measurements should not include the portions

where components join, such as at the wing –fuselage intersection. These areas are not “wetted”.

CALCULATIONS

1) For fuselage

𝑠𝜋𝑓= πdf

2

4

Π denotes its wetted calculation

From Airplane Design Part II by Dr.Johnroskam, lf

dffor Single Engine Aircraft is 6.5,

From wing design calculation Lf =24.13 m,

Now, df= 24.13

8.5 =2.84 m,

𝑠𝜋𝑓= πdf

2

4=

π×2.842

4=6.335 m2

2) For wing

sπw=𝑡𝑤 × 𝑏𝑤

A known relation,tw

croot = 0.09(from aerofoil t/c max)

From wing design calculation,crootis 7.8 m,

𝑡w=0.09×5.56 = 0.5004 m.

𝑠𝜋𝑤= 0.5004×15.2 =7.161 m2

3)For horizontal tail

sπht= 𝑡ht×𝑏ht=9.97×0.05004 =0.5 m2

𝑡ht= 𝑡vt= 10 percent 𝑡w=0.1×0.5004 =0.05004

From “Aircraft design: A Conceptual approach” by Daniel P.Raymer,

(AR)ht=bht

2

sht

= 3.5

Now,

bht2

= 3.5×28.4 = 99.4 m

4) For vertical tail

(AR)vt=bvt

2

svt

= 1.1

𝑠𝜋𝑣𝑡=𝑏vt×tvt= 2.881×0.05004 =0.1112 m2.

5) Engine area

𝑠𝜋𝑒𝑛𝑔𝑖𝑛𝑒=𝜋𝑑e

2

4

=𝜋×1.422

4 Since de=

df

2= 2.84/2 =1.42 m

=1.583 m2.

6) 1/4 flap deflection

𝜃 =15˚

For Single Engine range, (0.05 to 0.1)

The below is average of above range,

sπ= 0.075 m2

7) 3/4 flap deflection

𝜃 =45˚

For Single Engine range, (0.15 to 0.2)

The below is average of above range,

sπ= 0.175 m2

8) Undercarriage

sπu=1.1×sπengine

=1.1×1.583

=1.741 m2

RESULT:

The wetted area details are,

S.No Component sπ (m2)

1 Fuselage 6.335

2 Wing 7.161

3 Horizontal tail 0.500

4 Vertical tail 0.111

5 Engine 1.583

6 1/4 flap 0.075

7 3/4 flap 0.175

8 Undercarriage 1.741

Exp.No:8 Date:06.03.2014

DRAG POLAR

CDt =CDO+K(CL)2

Where,

K=1

𝛱∗𝑒∗𝐴𝑅

=1

𝛱×0.7×3.0

=0.055

1.At SEA LEVEL, (h=0)

Where,

ρ =1.225 kg/m3

a = (γ×R×T) ^0.5 = (1.4×287×288.16) ^0.5 =340.268 m/s.

CL =2×𝑊

𝜌×𝑆×𝑉^2 =

2×27345.13×9.81

1.225×77.03×129.4^2 = 0.3395

S.No

V

(m/s)

CL CDT =(𝐶𝐷𝑂 + (𝑘CL)2

1

2

3

4

5

129.4

258.8

388.2

517.6

647

0.34

0.085

0.038

0.0212

0.014

0.0475

0.0311

0.0302

0.0301

0.0300

2. At Altitude, (h=18.0 km)

T=281.66 K,

ρ =0.12165 kg/m3

a = (γ×R×T) ^0.5 = (1.4×287×281.66) ^0.5 =336.40 m/s

CL =2×𝑊

𝜌×𝑆×𝑉^2 =

2×27345.13×9.8

0.12165×77.03×129.4^2 = 3.42

S.No

V

(m/s)

CL CDT =(𝐶𝐷𝑂 + (𝑘CL)2

1

2

3

4

5

129.4

258.8

388.2

517.6

647

3.42

0.85

0.38

0.21

0.136

1.8032

0.1395

0.0519

0.0367

0.0328

0

0.01

0.02

0.03

0.04

0.05

0.06

0 0.1 0.2 0.3 0.4

co e

ffic

ien

t o

f d

rag

co efficient of lift

Series1

0.02995

0.03

0.03005

0.0301

0.03015

0.0302

0.03025

0 0.01 0.02 0.03 0.04

co e

ffic

ien

t o

f d

rag

co efficient of lift

Series1

RESULT: The graph drawn b/w lift coefficient and drag coefficient for different stages of aircraft.

And the variation of trend was observed.

Exp.No:9 Date: 20.03.2014

CALCULATION OF DRAG

Aerodynamic forces that split into two forces: Lift force or lift, and Drag force or drag. A pre-

requisite to aircraft performance analysis is the ability to calculate the aircraft drag at various flight

conditions. Drag force is the summation of all forces that resist against aircraft motion.

The drag coefficient is non-dimensional parameter, but it takes into account every aerodynamic

configuration of the aircraft including, wing, tail, fuselage and landing gear. This coefficient has two

main parts. The first part is referred to as lift-related drag coefficient or induced drag coefficient (CDi)

and the second part is called zero-lift drag coefficient (CDo).

Calculation of CDo

The CDoof an aircraft is simply the summation of CDoof all contributing components.

CDof, CDow, CDoht, CDovt, CDoLG, CDoN, CDoS, CDoHLD, are respectively representing

fuselage, wing, horizontal tail, vertical tail, landing gear, nacelle, strut, high lift device (such as

flap).

CDoOTHERS is components such as antenna, pitot tube, wire, and wiper

Fuselage

The zero-lift drag coefficient of fuselage is given by the following equation:

where, Cf is skin friction coefficient and is non-dimensional number. It is determined based on the

Prandtl relationship as follows:

(for turbulent and laminar flow)

Where ρ is the air density, V is aircraft true airspeed, µ is air viscosity, and L is the length of the

component in the direction of flight. For the fuselage, L it the fuselage length. The second

parameter (fLD) is a function of length to diameter ratio

The third parameter (fM) is a function of Mach number (M).

The last two parameters Swetf and S, where are respectively the wetted area of the fuselage and

the wing reference area.

Wing, Horizontal Tail, and Vertical Tail

In these equations, Cfw, Cfht, Cfvt are similar to what we defined for fuselage. The only difference

is that the equivalent value of L in Reynolds number) for wing, horizontal tail, and vertical tail are

their mean aerodynamic chord (MAC).

High lift devices

The δfis the flap deflection in degrees (usually less than 50 degrees).

Landing gear

Engine (cooling drag)

where P is the engine power (hp), T is the air temperature (K), σ is the relative density of the air,

V is the aircraft velocity (m/sec), and S is the wing reference area (m2). Th parameter Ke is a

coefficient that depends on the type of engine. It varies between 1 and 3.

Overall CDo

whereKc is a correction factor and depends on several factors such as the type, year of fabrication

and configuration of the aircraft.

Sl.No. Aircraft type Kc

1 Passenger 1.1

2 Agriculture 1.5

3 Cargo 1.2

4 Single engine piston 1.3

5 General Aviation 1.2

6 Fighter 1.1

No. Component CDo of

component

Percent from

total CDo (%)

1 Wing 0.0053 23.4

2 Fuselage 0.0063 27.8

3 Wing tip tank 0.0021 9.3

4 Nacelle 0.0012 5.3

5 Engine strut 0.0003 1.3

6 Horizontal tail 0.0016 7.1

7 Vertical tail 0.0011 4.8

8 Other components 0.0046 20.4

9 Total CDo 0.0226 100

No. Aircraft type CDo E

1 Subsonic jet 0.014-0.02 0.75-0.85

2 Large turboprop 0.018-0.024 0.8-0.85

3 Twin-engine piton prop 0.022-0.028 0.75-0.8

4 Small GA with

retractable landing gear

0.02-0.03 0.75-0.8

5 Small GA with fixed

landing gear

0.025-0.04 0.65-0.8

6 Agricultural

8 Supersonic jet 0.02-0.04 0.6-0.8 Typical values of CDoand e

for several aircraft

without crop duster

0.06-0.065 0.65-0.75

For our wing, k= 1

𝛱𝑒𝐴𝑅 =0.1516

1.At SEA LEVEL, (h=0)

Where,

T=288.16 K,

ρ =1.225 kg/m3

a = (γ×R×T) ^0.5 = (1.4×287×288.16) ^0.5 =340.268 m/s.

CL =2×𝑊

𝜌×𝑆×𝑉^2 =

2×27345.13×9.81

1.225×77.03×129.4^2 = 0.3395

S.No

V

(m/s)

CL CDo

CDT =(𝐶𝐷𝑂 + (𝑘CL)2 D=(( CDT

×W)/ CL)

(N)

1

2

3

129.4

258.8

388.2

0.34

0.085

0.038

0.03

0.03

0.03

0.0475

0.0311

0.0302

37.48

98.15

213.19

4

5

517.6

647

0.0212

0.014

0.03

0.03

0.0301

0.0300

380.87

596.12

2. At Altitude, (h=18.0 km)

T=281.66 K,

ρ =0.12165 kg/m3

a = (γ×R×T) ^0.5 = (1.4×287×281.66) ^0.5 =336.40 m/s

CL =2×𝑊

𝜌×𝑆×𝑉^2 =

2×27345.13×9.8

0.12165×77.03×129.4^2 = 3.42

S.No

V

(m/s)

CL CDo

CDT =(𝐶𝐷𝑂 + (𝑘CL)2 D=(( CDT

×W)/ CL)

(kN)

1

2

3

4

5

129.4

258.8

388.2

517.6

647

3.42

0.85

0.38

0.21

0.136

0.03

0.03

0.03

0.03

0.03

1.8032

0.1395

0.0519

0.0367

0.0328

141.44

44.03

36.64

46.88

64.70

.

GRAPH BETWEEN VELOCITY & DRAG:

0

20

40

60

80

100

120

140

160

0 100 200 300 400 500 600 700

dra

g(N

)

velocity(m/s)

0

100

200

300

400

500

600

700

0 100 200 300 400 500 600 700

dra

g(N

)

velocity(m/s)

RESULT:

From the above tables and graphs, drag and velocity at different altitudes are obtained.

Exp.No:10 Date:27.03.2014

THRUST REQUIRED CALCULATION

Thrust available, from the engine selection calculation, F = 160.95 KN

Freq = F ×σ1.15

For sea level,

Freq = F [(20 – h) / (20+h)] 1.15

= 160.95 [(20-0) / (20+0)] 1.15

= 160.95 KN

For h = 3 km,

Freq = 160.65 [(20- 3) / (20+3)] 1.15

=113.69 KN

For h=6 km,

Freq = 160.65 [20-6) / (20+6)] 1.15

=78.98 KN

For h=9 km,

Freq = 160.95 [(20 – 9) / (20+9)] 1.15

= 52.79 KN.

For h=12 km,

Freq = 160.95 [(20-12) / (20+ 12) ] 1.15

= 32.68 KN.

For h=15 km,

Freq= 160.95 (( 20-15) / (20 +15))1.15 = 17.17 KN

For h = 18 km,

Freq = 160.95 ( (20-18) / (20+18) )1.15

= 5.45 KN

S.NO ALTITUDE

(Km)

THRUST or POWER

( KN )

1 0 160.95

2 3 113.69

3 6 78.98

4 9 52.79

5 12 32.68

6 15 17.17

7 18 5.45

RESULT:

Thus the thrust required for multirole fighter aircraft has been

calculated.

h = 0 km, Freq = 160.95 KN

h = 18 km, Freq = 5.45 KN

Exp no. 11 Date : 27.03.14

RATE OF CLIMB CALCULATION

Rate of climb is defined as the aircraft speed in the vertical axis or the vertical component of the

aircraft airspeed. Hence rate of climb is about how fast an aircraft gains height.

Jet aircraft:

In general, the Rate of Climb (ROC) is defined as the ratio between excess power and the aircraft

weight

Prop-driven Aircraft:

The available power is the engine power times the propulsive efficiency.

1. At SEA LEVEL (h=0)

S.No V(m/s) T

(KN)

D

(KN)

RATE OF

CLIMB (m/s)

1

2

3

4

5

129.4

258.8

388.2

517.6

647

160.95

160.95

160.95

160.95

160.95

37.48

98.15

213.19

380.87

596.12

59.56

60.59

-75.60

-424.34

-1049.58

2. At h= 18 km.

S.No V(m/s) T

(KN)

D

(KN)

RATE OF

CLIMB (m/s)

1

2

129.4

258.8

5.45

5.45

141.44

44.03

-65.55

-37.22

3

4

5

388.2

517.2

647

5.45

5.45

5.45

36.64

46.88

64.70

-45.136

-79.88

-142.90

1.At h=0km

2. At h=18km

-1200

-1000

-800

-600

-400

-200

0

200

0 200 400 600 800

R/C

(m/s

)

Velocity(m/s)

Series1

RESULT:

From the above analysis, two graphs – rate of climb Vs velocity for different altitudes and rate of

climb Vs altitude drawn and the trend in rate of climb was observed.

TOP VIEW, SIDE VIEW, FRONT VIEW (CAD DRAWING)

-160

-140

-120

-100

-80

-60

-40

-20

0

0 200 400 600 800R

/C(m

/s)

Velocity(m/s)

Series1

REFERENCE

TEXTS:

1. Theory of wing section by IRA H.ABBOT and ALBERT E.VON DOENHOFF.

2. Aircraft performance and design by JOHN D.ANDERSON JR

3. Aircraft design: A conceptual Approach by DANIEL P.RAYMER

4. Aircraft design by THOMAS CORK

5. Aircraft design by MOHAMMAD SADRAEY

6. Aircraft design by JOHN ROSKAM.

7. JANES All the World Aircrafts