Aircraft Design Project 1

1

AIM OF THE PROJECT

The aim of this design project is to design a 300 seater passenger

aircraft by comparing the data and specifications of present aircrafts in

this category and to calculate the performance characteristics. Also

necessary graphs need to be plotted and diagrams have to be included

wherever needed.

The following design requirements and research studies are set for the

project:

Design an aircraft that will transport 300 passengers and their

baggage over a design range of 13800 km at a cruise speed of

about 0.85 Mach number.

To provide the passengers with high levels of safety and comfort.

To use advanced and state of the art technologies in order to reduce

the operating costs.

To offer a unique and competitive service to existing scheduled

operations.

To assess the development potential in the primary role of the

aircraft.

To produce a commercial analysis of the aircraft project.

2

ABSTRACT

Aircraft design is an evolutionary process rather than a revolutionary process

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.

Sir George cayley who was pioneer and his revolutionary work has

helped in reaching great heights in aero science. Today our dream for

designing a 300 seater passenger aircraft has come into reality.

The purpose of the project is to design a passenger aircraft comprised of

300 passengers with 5 crew members. Turbofan engines are provided for

the required amount of speed, range, and fuel consumption. There remain

a lot of technical challenges and problems to be met and solved before

sustained, practical passenger aircraft becomes reality. In this project we

use various design parameters. This result of various design process gave

clear view of long range wide body passenger aircrafts. The performance

calculation is done with the normal payload. Two turbofan engines are

used for producing the required thrust.

3

CHAPTER 1

INTRODUCTION OF AIRCRAFT DESIGN

1.1INTRODUCTION

For any airplane to fly, it must be able to lift the weight of the

airplane, its fuel, the passengers, and the cargo. The wings generate most

of the lift to hold the plane in the air. To generate lift, the airplane must

be pushed through the air. The engines, which are usually located

beneath the wings, provide the thrust to push the airplane forward

through the air.

The fuselage is the body of the airplane that holds all the pieces of the

aircraft together and many of the other large components are attached to

it. The fuselage is generally streamlined as much as possible to reduce

drag. Designs for fuselages vary widely. The fuselage houses the cockpit

where the pilot and flight crew sit and it provides areas for passengers

and cargo.

The wing provides the principal lifting force of an airplane. Lift is

obtained from the dynamic action of the wing with respect to the air. The

cross-sectional shape of the wing as viewed from the side is known as the

airfoil section. The planform shape of the wing (the shape of the wing as

viewed from above) and placement of the wing on the fuselage

(including the angle of incidence), as well as the airfoil section shape,

depend upon the airplane mission and the best compromise necessary in

the overall airplane design.

The control surfaces include all those moving surfaces of an airplane

used for attitude, lift, and drag control. They include the tail assembly,

the structures at the rear of the airplane that serve to control and

maneuver the aircraft and structures forming part of the tail and attached

to the wing.

1.2ACTUAL PROCESS OF DESIGN

Selection of aircraft type and shape

4

Determination of geometric parameters

Selection of power plant

Structural design and analysis of various components

Determination of aircraft flight and operational characteristics.

1.3STAGES OF AIRCRAFT DESIGN

Project Feasibility Study

Preliminary Design

Design Project

1.3.1PROJECT FEASIBILITY STUDY

Comprehensive market survey

Studies on operating conditions for the airplane to be designed

Studies on relevant design requirements (specified by

Airworthiness Authorities)

Evaluation of similar existing designs

Studies on possibilities of introducing new concepts

Collection of data on relevant power plants

Laying down preliminary specifications

1.3.2PRELIMINARY DESIGN

It consists of the initial stages of design, resulting in the

presentation of a BROCHURE containing preliminary drawings and

clearly stating the operational capabilities of the airplane being

designed. This Brochure has to be APPROVED by the manufacturer

and/or the customer.

The steps involved:

Layout of the main components

Arrangement of airplane equipment and control systems

Selection of power plant

Aerodynamic and stability calculations

Preliminary structural design of MAJOR components

Weight estimation and c.g. travel

Preliminary and Structural Testing

5

Drafting the preliminary 3-view Drawings

1.3.3DESIGN PROJECT

Internal discussions

Discussions with prospective customers

Discussions with Certification Authorities

Consultations with suppliers of power plant and major accessories

Deciding upon a BROAD OUTLINE to start the ACTUAL

DESIGN, which will consist of Construction of Mock-up

Structural layout of all the individual units, and their stress analysis

Drafting of detailed design drawings

Structural and functional testing

Nomenclature of parts

Supplying key and assembly diagrams

Final power plant calculations

Final weight estimation and c.g. limits

Final performance calculation

1.4THE DESIGN WHEEL

Fig 1.1The Design Wheel

SIZING AND

TRADE

STUDIES

REQUIREMENTS

DESIGN

ANALYSIS

DESIGN

CONCEPT

6

SEVEN INTELLECTUAL POINTS FOR CONCEPTUAL DESIGN

Fig 1.2 Seven Intellectual Points for Conceptual Design

7

1.5DESIGN SEQUENCE

1. Define the mission

2. Compare the past design

3. Parametric selection

a. Geometry

b. Shape

4. Weight Estimation

5. Aerodynamics

a. Wing

b. Speed

c. Altitude

d. Drag

6. Propulsive device

a. Engine selection

b. Location

7. Performance

a. Fuel weight

b. Take-off distance

c. Landing distance

d. Climb

e. Descent

f. Loiter

g. Cruise

8. Configuration

a. Conceptional

b. Preliminary

c. Detailed design

9. Stability and control

a. Tail

b. Flaps

c. Control surfaces

8

10. Structure

a. Primary

b.Secondary

c. Tertiary

11. Construction

a. Truss

b. Semi-monocoque

c. Monocoque

12. Manufacturing Models

a. Mock up model

b. Training model

c. Scale in/out

d. Fake model

e. Test model

f. Prototype model

g. Flying model

13. Life cycle cost Minimize the owning cost

14. Iteration Refine the weight and design

15. Simulation Flight envelope

16. Testing

17. Modification and refinement

18. Design report

a. Executive summary

b. Management summary

c. Design details

d. Manufacturing plan

9

CHAPTER-2

COMPARATIVE DATASHEET

Table 2.1

Comparative Datasheet

Aircrafts

Parameter Units 1 2 3 4 5

Name - 707-320B 757-200 767-200 777-200 787-9

Total Seating

Capacity

- 202 234 290 301 280

Aircraft

Dimensions Length m 46.61 47.32 48.5 63.7 62.8

Height m 12.93 13.56 16.8 18.5 16.9

Fuselage Dia m 3.76 4.1 5.03 6.2 5.9

Wing Span m 44.42 38.05 47.6 60.9 60.1

Chord m 6.25 4.76 5.95 7.02 6.4

Aspect Ratio - 7.1 7.98 7.99 8.67 9.4

Wing Area m2 273.7 181.25 283.3 427.8 325.3

Wing Sweep Degree 35 25 31.5 31.64 32.2

Performance

Cruise Altitude m 10,058 10,668 10,668 10,668 12,192

Ceiling m 11,887 12,802 11,887 13,137 13,106

Range Km 10,650 7,600 7,300 9,695 15,000

Cruise Speed Mach 0.86 0.8 0.8 0.84 0.85

Max Speed Mach 0.97 0.84 0.84 0.87 0.9

No of Engines - 4 2 2 2 2

Max thrust

capability

kN 320.4 193 222 330 320

Design Weights

MTO Weight Kg 151320 115680 142880 247200 248000

Empty Weight Kg 66400 57180 81230 134800 115000

Wing Loading Kg/m2 552.87 638.23 504.34 577.84 762.37

Max Fuel

Capacity

Litre 90,160 43,490 90,770 117,000 127,000

10

Table 2.2

Comparative datasheet 2

Aircrafts


Name - A380-800 B-747-200 B-787-8 B-787-10 B-747-300

Total Seating

Capacity

- 64

4

412 310 313 270

Aircraft

Dimensions Length m 72.

7

70.6 68.3 67.9 60.7

Height m 24.45 19.3 16.9 17.1 17.2

Fuselage Dia m 5.6

4

5.64 5.64 5.64 5.96

Wing Span m 79.75 59.6 60.1 63.45 64.8

Chord m 5.8 5.64 6.5 6.8 7

Aspect Ratio - 7.

7

7.7 9.3 9.3 9.25

Wing Area m2 84

5

219 325 439.4 443

Wing Sweep Degree

33.5

37.5 32.2 31.1 31.9

Performance


Ceiling m 12,000 12,497 12,000 12,527 13,137

Range Km 15700 12690 14,500 16,060 15,000


Max Speed Mach 0.89 0.89 0.90 0.86 0.9


Max thrust

capability

kN 369 244 28

0

249 374

Design Weights

MTO Weight Kg 575000 377842 228000 372000 268000

Empty Weight Kg 276000 174000 118000 170900 115700

Wing Loading Kg/m2 660.38 748.86 644.36 846.61 604.96

Max Fuel

Capacity

Litre 323,546 200 126.210 2,14,810 1,29,000

11

Table 3

Comparative Datasheet 3

Aircrafts


Name

- Lockheed

L-1011-200

Ilyushin

IL-96-300

Tupolev

Tu-204-100

Douglas

DC-8-63CF

Tupolev

Tu-114

Total Seating

Capacity

- 26

3

300 210 259 220

Aircraft Dimensions

Lengt

h

m 54.15 55.3 46.1 57.1 54.1

Height m 16.87 17.5 13.9 13.11 15.44

Fuselage

Diameter

m 6.0 6.08 4.1 3.73 4.2

Wing Span m 47.35 60.11 41.8 45.24 51.1

Chord m 6.7

8

5.82 4.40 6.01 6.08

Aspect Ratio - 6.9

8

10.32 9.48 7.52 8.39

Wing Area m

2

321.1 350 184.2 271.9 311.1

Wing Sweep Degree 35 30 30 32 35

Performance


Ceiling m 10,668 13,106 12,588 12,497 11,887

Range Km 7,420 10,400 5,650 3,445 6,200


Max Speed Mach 0.95 0.84 0.85 0.8 0.82


Max thrust

capability

kN 222.4 157 158.3 84.5 60

Design Weights

MTO Weight Kg 211000 250000 103000 161000 175000

Empty Weight Kg 105100 120400 60000 66360 93000

Wing Loading Kg/m2 657.11 714.28 559.17 592.12 562.52

Max Fuel

Capacity

Litre 99,935 152,620 41,000 66,243 71,615

12

Table 4


Aircrafts


Name - B-767-400-ER Ilyushin

IL-86

Ilyushin

IL-96M

Ilyushin

IL-96T

Ilyushin

IL-96-400

Total Seating - 304 320 340 313 386

Aircraft Dimensions

Length m 61.4 60.21 64.7 63.9 63.9

Height m 16.62 15.8 15.7 15.7 15.7

Fuselage Dia m 5.64 6.08 6.08 6.08 6.08

Wing Span m 51.9 48.06 60.11 60.11 60.11

Chord m 5.8 5.64 6.5 6.8 7

Aspect Ratio - 7.7 7 7 7 7

Wing Area m2

290 300 350 350 350

Wing Sweep Degree 28 35 35 35 35

Performance


Ceiling m 13,100 13,100 13,100 13,100 13,100

Range Km 10,418 3,400 12,800 5,000 10,000


Max Speed Mach 0.86 0.84 0.84 0.84 0.84


Max thrust kN 282 128 167 167 171

Design Weights

MTO Weight Kg 204000 215000 270000 270000 265000

Empty Weight Kg 104000 117.5 132.4 116.4 122.3

Wing Loading Kg/m2 660.38 748.86 644.36 846.61 604.96

Max Fuel

Capacity

Litre 91,400 75,470 152,260 152,260 152,260

13

Table 5


Aircrafts


Name (no unit) A300-B4 A310-200 A330-300 A340-500 A350-800

Total Seating

Capacity

(no unit) 266 240 295 313 270

Aircraft Dimensions

Length m 53.62 46.6 63.6 67.9 60.7

Height m 16.62 15.8 16.85 17.1 17.2

Fuselage Dia m 5.64 5.64 5.64 5.64 5.96

Wing Span m 44.85 43.9 60.3 63.45 64.8

Chord m 5.8 5.64 6.5 6.8 7

Aspect Ratio (no unit) 7.7 7.78 9.3 9.3 9.25

Wing Area m2 260 219 361.6 439.4 443

Wing Sweep degree 28 28 30 31.1 31.9

Performance


Ceiling m 12,000 12,497 12,527 12,527 13,137

Range Km 7,540 9,600 10,500 16,060 15,000


Max Speed Mach 0.86 0.84 0.86 0.86 0.9

No of Engines (no unit) 2 2 2 4 2

Max thrust

capability

kN 311.4 262.5 320 249 374

Design Weights

MTO Weight

Kg 171700 164000 233000 372000 268000

Empty Weight Kg 90900 83100 124500 170900 115700

Wing Loading Kg/m2 660.38 748.86 644.36 846.61 604.96

Max Fuel

Capacity

Litre 68,150 75,470 97,170 2,14,810 1,29,000

14

CHAPTER 3

LIST OF GRAPHS

3.1.1CRUISE SPEED VS CARGO CAPACITY

Cruise Speed (Mach)

Graph1

3.1.2CRUISE SPEED VS OVERALL LENGTH

Cruise Speed (Mach)

Graph 2

0

100

200

300

400

500

600

0.77 0.82 0.87 0.92 0.97

0

10

20

30

40

50

60

70

80

90

0.77 0.82 0.87 0.92 0.97

Cargo capacity = 146 m3

Car

go

cap

acit

y (

m3)

Overall Length = 55.5 m

Over

all

Len

gth

(m

)

15

3.1.3CRUISE SPEED VS WING SPAN

Cruise Speed (Mach)

Graph 3

3.1.4CRUISE SPEED VS WING AREA

Cruise Speed (Mach)

Graph 4

0

10

20

30

40

50

60

70

80

90

0.77 0.81 0.85 0.89 0.93 0.97

0

100

200

300

400

500

600

700

800

900

0.77 0.82 0.87 0.92 0.97

Win

g s

pan

(m

) W

ing

are

a (m

2)

Wing span = 54 m

Wing Area = 350 m2

16

3.1.5CRUISE SPEED VS OVERALL HEIGHT

Cruise Speed (Mach)

Graph 5

3.1.6CRUISE SPEED VS CABIN WIDTH

Cruise Speed (Mach)

Graph 6

0

5

10

15

20

25

30

0.77 0.82 0.87 0.92 0.97

0

1

2

3

4

5

6

7

0.77 0.82 0.87 0.92 0.97

Over

all

hei

gh

t (m

) C

abin

wid

th (

m)

Cabin Width = 5.6 m

Overall Height = 18 m

17

3.1.7CRUISE SPEED VS OPERATING WEIGHT

Cruise Speed (Mach)

Graph 7

3.1.8CRUISE SPEED VS FUSELAGE WIDTH

Cruise Speed (Mach)

Graph 8

0

50000

100000

150000

200000

250000

300000

0.77 0.82 0.87 0.92 0.97

2

2.5

3

3.5

4

4.5

5

5.5

6

6.5

7

0.72 0.77 0.82 0.87 0.92

Oper

atin

g w

eight

(kg)

Operating Weight = 120,000 kg

Fu

sela

ge

wid

th (

m)

Fuselage Width = 6 m

18

3.1.9CRUISE SPEED VS FUSELAGE HEIGHT

Cruise Speed (Mach)

Graph 9

3.1.10CRUISE SPEED VS FUSELAGE DIAMETER

Cruise Speed (Mach)

Graph 10

2

2.5

3

3.5

4

4.5

5

5.5

6

6.5

7

0.77 0.82 0.87 0.92 0.97

2

2.5

3

3.5

4

4.5

5

5.5

6

6.5

7

0.77 0.82 0.87 0.92 0.97

Fuselage Height = 5.8 m

Fuselage Diameter = 6 m

Fu

sela

ge

Hei

ght

(m)

19

3.1.11CRUISE SPEED VS MAXIMUM TAKEOFF WEIGHT

Cruise Speed (Mach)

Graph 11

3.1.12CRUISE SPEED VS TAKEOFF FIELD LENGTH

Cruise Speed (Mach)

Graph 12

0

100000

200000

300000

400000

500000

600000

700000

0.77 0.82 0.87 0.92 0.97

0

500

1000

1500

2000

2500

3000

3500

4000

0.77 0.82 0.87 0.92 0.97

Maximum Takeoff Weight =230,000 kg

Takeoff Field Length=2430 m

Tak

eoff

Fie

ld L

eng

th

(m)

Max

imu

m T

akeo

ff

Wei

ght

(kg

)

20

3.1.13CRUISE SPEED VS MAX SPEED

Cruise Speed (Mach)

Graph 13

3.1.14CRUISE SPEED VS RANGE

Cruise Speed (Mach)

Graph 14

0.7

0.75

0.8

0.85

0.9

0.95

1

0.77 0.82 0.87 0.92 0.97

0

5000

10000

15000

20000

25000

30000

35000

40000

0.77 0.82 0.87 0.92 0.97

Max speed = 0.871

Range = 13,800 km

Ran

ge

(km

)

Max

sp

eed

(m

ach

)

21

3.1.15CRUISE SPEED VS FUEL CAPACITY

Cruise Speed (Mach)

Graph 15

3.1.16CRUISE SPEED VS CEILING

Cruise Speed (Mach)

Graph 16

0

50000

100000

150000

200000

250000

300000

350000

0.77 0.82 0.87 0.92 0.97

10000

10500

11000

11500

12000

12500

13000

13500

14000

14500

15000

0.77 0.82 0.87 0.92 0.97

Fuel Capacity = 136,000 litres

Ceiling = 12750 m

Cei

lin

g (

m)

Fu

el C

apac

ity

(lit

ers)

22

3.1.17CRUISE SPEED VS WING SWEEP ANGLE

Cruise Speed (Mach)

Graph 17

3.1.18CRUISE SPEED VS ASPECT RATIO

Cruise Speed (Mach)

Graph 18

25

26

27

28

29

30

31

32

33

34

35

0.77 0.82 0.87 0.92 0.97

5

6

7

8

9

10

11

0.77 0.82 0.87 0.92 0.97

Win

g S

wee

p A

ngle

(Deg

ree)

0

Wing Sweep Angle = 30.6

Aspect Ratio = 8

Asp

ect

Rat

io (

no

un

it)

23

3.1.19CRUISE SPEED VS PAYLOAD

Cruise Speed (Mach)

Graph 19

3.1.20CRUISE SPEED VS THRUST

Cruise Speed (Mach)

Graph 20

0

10000

20000

30000

40000

50000

60000

70000

80000

90000

100000

0.77 0.82 0.87 0.92 0.97

0

100

200

300

400

500

600

0.77 0.82 0.87 0.92 0.97

Payload = 46,000 kg

Pay

load

(k

g)

Thrust = 185 KN

Th

rust

(K

N)

24

3.1.21CRUISE SPEED VS MAXIMUM LANDING WEIGHT

Cruise Speed (Mach)

Graph 21

3.1.22CRUISE SPEED VS MAXIMUM ZERO FUEL WEIGHT

Cruise Speed (Mach)

Graph 22

0

50000

100000

150000

200000

250000

300000

350000

400000

450000

0.77 0.82 0.87 0.92 0.97

0

50000

100000

150000

200000

250000

300000

350000

400000

450000

0.75 0.8 0.85 0.9 0.95 1

Max. Landing Weight = 185,000 kg

Max

. L

andin

g W

eight

(kg

)

Max. Zero Fuel Weight = 170,000 kg

Max

. Z

ero

Fuel

Wei

gh

t (k

g)

25

3.2 MEAN DESIGN PARAMETERS

S.No Design Parameter Value Unit

1. Cruising Speed Mach 0.85 (no unit)

2. Length 55.5 m

3. Wing Span 54 m

4. Wing Area 350 m2

5. Height 18 m

6. Cabin Width 5.6 m

7. Seating Capacity 300 (Passengers) (no unit)

8. Cargo Capacity 146 m3

9. Fuselage Width 6 m

10. Fuselage Height 5.8 m

11. Fuselage Diameter 6 m

12. Takeoff Field Length 2430 m

13. Maximum Speed Mach 0.871 (no unit)

14. Range 13800 Km

15. Maximum Fuel Capacity 136,000 Litre

16. Service Ceiling 12,750 m

17. Wing Sweep Angle 30.6 (degree)

18. Aspect Ratio 8 (no unit)

19. Thrust 185 kN

20. Empty Weight (Operating) 120,000 Kg

21. Maximum Takeoff Weight 230,000 Kg

22. Maximum Payload 46,000 Kg

23. Maximum Zero Fuel

Weight

170,000 Kg

24. Maximum Landing Weight 185,000 Kg

25. Engine 2 (no unit)

26

CHAPTER-4

FLIGHT MISSION PATH

Fig4.1

The above plan one of the most basic and would generally correspond

to a Commercial aircrafts. It consists of flight phases made up of engine

start up and take-off, climb and accelerate to cruise altitude and speed,

cruise out to destination, and landing.

4.1 ENGINE START-UP AND TAKE-OFF

The Engine start-up and Take-off is the first phase in any flight plan. It

consists of starting the engines, taxiing to the take-off position, take-off,

and climb out. A good empirical estimate for the weight of fuel used in

this phase is from 2.5 to 3 % of the total take-off weight.

1. Engine Starts Warm-Up

2. Taxi

3. Take-off

4. Climb

5. Cruise

6. Loiter

7. Descent

8. Landing

27

4.2 ACCELERATION TO CRUISE VELOCITY AND ALTITUDE

After the take-off the aircraft will generally climb to cruise altitude and

accelerate to cruise speed. The estimate for the weight fraction for this

phase of the flight is also found from the empirical data.

4.3 CRUISE OUT TO DESTINATION

For a cruising aircraft the fuel weight fraction can be determined quite

well from an analytical formulation called the Brequet range equation.

4.4 ACCELERATION TO HIGH SPEED (INTERCEPT)

The flight phase involves accelerating from the cruise Mach no to a

maximum flight Mach no as part of a high speed intercept.

4.5 RETURN CRUISE

Return cruise refers to a flight plan in which the aircraft return to its

point of origin to land for a flight plans in which the landing destination is

different from where it took-off, return cruise can be viewed as the second

half of the cruise phase. In either case return cruise treated exactly like

cruise out with two possible exceptions: Loss of fuel weight, Increase in

altitude due to decrease in weight.

4.6 LOITER

The loiter consist of cruising for specified amount of time over a small

region. Loiter is usually built into the flight plan to allow for delays prior

landing. For this phase the fuel weight fraction is derived an analytical

expression called the Endurance equation.

4.7 LANDING

The final phase of the flight plan is landing. As an estimate of the fuel

weight fraction used at landing, we use the same empirical formula that

was used for start- up and take-off.

28

CHAPTER 5

WEIGHT ESTIMATION

5.1 FIRST WEIGHT ESTIMATION

The design take off gross weight Wo is the weight of the airplane

at the instant it begins its mission. It includes the weight of all the fuel

on board at the beginning of the flight.

Wo = Wcrew + Wpayload + Wfuel + Wempty

5.2 CREW WEIGHT

The two pilots and three cabin attendants at 175 lbs each and 30 lbs

baggage each

Therefore,

No.of Crew = 5

Wcrew = (5*175 lbs) + (5*30 lbs)

=1025 lbs.

Wcrew = 1025 lbs.

5.3 PAYLOAD WEIGHT

The 150 passengers at 175 lbs each and 30 lbs of baggage each.

Therefore,

No.of passengers = 300

Wpayload = (300*175 lbs) + (300*30 lbs)

=52,830 lbs

Wpayload = 52,830 lbs

29

5.4 FUEL WEIGHT

Mission Profile

Phase 1:

The Engine starts warm-up Weight Ratio is W1 /W2

Phase 2:

The Taxi Weight Ratio is W2 / W1

Phase 3:

The Take-off Weight Ratio is W3 / W2

Phase 4:

The Climb Weight Ratio is W4 / W3

Phase 7:

The Descent Weight Ratio is W7 / W6

Phase 8:

The Landing Weight Ratio is W8 / W7

1. Engine Starts Warm-Up

2. Taxi

3. Take-off

4. Climb

5. Cruise

6. Loiter

7. Descent

8. Landing

30

The phase 1, 2, 3, 4, 7, 8 are refer Table 6

Table 5.1

S

No

Aircraft

W1 / W0

W2 / W1

W3 / W2

W4 / W3

W7 / W6

W8 / W7

1

Transport

jet

0.990

0.990

0.995

0.980

0.990

0.992

Phase 5:

The Cruise Weight Ratio is W5 / W4

By using formula,

Rcr = (V / Cj)cr x ( L /D )cr x In ( W4 / W5 )

Rcr = 7,452 Nautical miles

Vcr = 849.7 kmph

Table 5.2

Cruise

Loiter

L/D 14 17

Cj 0.75 0.5

Rcr = (V / Cj)cr * ( L /D )cr * In ( W4 / W5 )

7452 = (849700 / 0.75) * (14) * In (W4 / W5)

In (W4 / W5) = 4.69 * 10-4

(W4 / W5) = 1.000469

( W5 / W4 ) = 1.0001

31

Phase 6:

The Loiter Weight Ratio is W6 / W5

Eltr = 25 min = 0.417 hrs.

Eltr = (1 / Cj)ltr x( L /D )ltr x In ( W5 / W6 )

0.417 = (1/ 0.5) x (17) x In (W5 / W6)

In (W5 / W6) = 0.0122

(W5 / W6) = 1.00122

The total weight ratio is,

( W8 / W0 ) = ( W1 / W0 ) ( W2 / W1 ) (W3 / W2 ) ( W4 / W3 )

( W5 / W4 )( W6 / W5 ) ( W7 / W6 ) ( W8 / W7 )

(W8 / W0) = 0.990 x 0.990 x 0.995 x 0.9980 x 0.9995 x 0.9878

x0.990 x0.992

5.5 EMPTY WEIGHT

The formula is,

WE = Antilog10 ( (log10 WTo A) / B)

Table 5.3

Eltr = (1 / Cj)ltr x ( L /D )ltr x In ( W5 / W6 )

( W6 / W5 ) = 1.000

( W8 / W0 ) = 0.9558

Wfuel = (1 Mff ) WTo

Aircraft A B

Transport Jet 0.0833 1.0383

32

By refer the graph,

Graph 23

WTo = 5.07 x 105 lbs,

WE = 270000 lbs

ITERATION

We put approximate Value

WTo = 508,560 lbs

WE = Antilog10 (5.454849)

Therefore,

WE = 270000 lbs

33

The Take-off Weight is

WTo = 527698 lbs

Wfuel = (1 Mff ) WTo

Wfuel = 0.0564 (558,560)

The total weight estimation is,

Wo = Wcrew + Wpayload + Wfuel + Wempty

Wo= 1,025 + 52,830 + 270,000 + 23,324.25

Wo = 347179 lbs

Wfuel =51421.164 lbs

34

CHAPTER 6

POWERPLANT SELECTION

6.1 INTRODUCTION

From the first weight estimate, we can have a rough idea of the

weight of the power-plant that is to be used.

The total weight of the power-plant (0.055W) requires being

approximately 15,443.5 kg.

Choice of engine is a Turbofan for obvious reasons such as

higher operating fuel economy & efficiency for high payloads.

Engines can be used in combination of 2 x 7721.8 kg engines. Or

3 x 5147.85 kg engines Or 4 x 3860.6 kg engines providing enough

thrust for Take-off.

Most of the aircraft in the 250-350 passenger category were found

to have 2 engines and 4 engines. Hence the preference is towards

having three engines (Trijet).

A list of engines with weight and thrust matching our requirements are

chosen and are tabulated below.

Table 6.1

Engine

name

Rolls

Royce

Trent

772B

Pratt &

whitney

PW400

CFM

CF

M56

General

Electric

CF6-50

Pratt &

Whitney

JT9D

Dry weight 478

8

4270 3990 4104 4030

Max thrust 320 310 151 240 250

Bypass

ratio

5 5 6.4 4.4 4.8

35

The preferable choice of engine, from those listed above would be

the pratt & whitney pw JT9D engine which meets our demand of

weight and powers. Airbus A330 and Boeing 777 aircrafts uses these

engines which are similar in payload capabilities such as the one

under design.

6.2 DETAILS ABOUT THE SELECTED ENGINE

6.2.1PRATT & WHITNEY PW JT9D

Since its launch with Cathay Pacific in 1995, PW JT9D has built up

the greatest service experience on the A330. As the only engine

specifically designed for the BOEING 777 it delivers the greatest

performance over the widest range of operational and environmental

conditions.

Fig 6.1 JT9D Turbofan engine

6.3 DESCRIPTION

High bypass turbo-fan engine

Bypass ratio is 5.0 : 1

6.3.1COMPRESSOR

36

Single Stage low pressure fan

3 Stage low pressure axial flow compressor

11 Stage high pressure axial flow compressor

6.3.2 COMBUSTION CHAMBER

Annular combuster

6.3.3 PRESSURE RATIO (OVERALL)

Nominal at sea-level ISA condition 23.4 : 1

6.3.4 TURBINE

4 Stage low pressure turbine

2 Stage high pressure turbine

6.3.5 DIMENSIONS

Overall length 3260mm

Maximum Radius 1670mm

6.3.6 DRY WEIGHT

The dry power plant weight less intake, cowl doors & cowl door

support structure is 3905kg (8608lbs).

6.3.7 ENGINE RATINGS

The ISA Sea-level static thrust ratings are

Take-off thrust - 222.41 KN

Thrust to weight ratio - 5.8

Fuel type - Jet A-1

Oil system - pressure spray with scavenge

37

CHAPTER-7

AIROFOIL SELECTION

7.1 AIROFOIL

The airfoil is the main aspect and is the heart of the airplane. The

airfoils affects the cruise speed landing distance and take off, stall speed

and handling qualities and aerodynamic efficiency during the all phases

of flight.

Aerofoil Selection is based on the factors of Geometry & definitions,

design/selection, families/types, design lift coefficient, thickness/chord

ratio, lift curve slope, characteristic curves.

An airfoils shape is defined by several parameters, which are shown in

the following figure:

Fig 7.1 Airfoil section

7.2 DEFINITIONS

7.2.1 CHORD LINE

Straight line drawn from the leading edge to the trailing edge

7.2.2 CHORD LENGTH (C)

Length of the chord line

38

7.2.3 MEAN CAMBER LINE

Curved line from the leading edge to the trailing edge, which is

equidistant between the upper and lower surfaces of the airfoil.

7.2.4 MAXIMUM CAMBER

Maximum distance between the chord line and the mean camber line

7.2.5 MAXIMUM THICKNESS

Maximum distance between the upper and lower surfaces of the airfoil

normal to the chord line

7.2.6 SPAN

Width of the airfoil

7.2.7 ANGLE OF ATTACK

Angle between the chord line and the stream wise flow direction

7.2.8 ZERO LIFT ANGLE OF ATTACK

Angle of Attack that no lift is produced. For our symmetric wedge this

would be an angle of attack of zero.

7.2.9 STALL ANGLE OF ATTACK

Angle of attack at which there is maximum lift (or lift coefficient).

Fig 7.2 Flow around the airfoil

7.2.10 SYMMETRIC OR UNCAMBERED AIRFOIL

Upper and lower surfaces are mirror images, which leads to the mean

camber line to be coincident with the chord line. A symmetric airfoil will

also have a just camber of zero.

7.2.11 CAMBERED AIRFOIL

An asymmetric airfoil for which the mean camber line will be above the

chord line.

39

Fig 7.3 Difference between uncambered and cambered airfoil

7.2.12 PITCHING MOMENT

Torque or moment created on the wing due to net lift and drag forces.

Tends to rotate the leading edge either up or down.

7.2.13 PITCHING MOMENT COEFFICIENT

Cm = (M) / (0.5 V2 S c)

Where,

M- Pitching moment (will depend on the moment reference center)

c- Chord length

7.2.14 CENTER OF PRESSURE

The moment reference center for which the moment is zero.Depends on

the angle of attack.

7.2.15 AERODYNAMIC CENTER

The moment reference center for which the moment does not vary with

angle of attack

7.3 NACA CLASSIFICATION

Airfoils have been classified by the National Advisory Committee for

Aeronautics (NACA), the forerunner of NASA, and have been cataloged

using a four digit code. Hence a specific airfoil can be identified by

NACA WXYZ

Where, W: maximum camber as % of the chord length

X: Location of the maximum camber form the leading edge along

the chord line in tenths of chord length

Y&Z: Maximum thickness in % of the chord length

40

7.3.1 NACA AIRFOIL CHARTS

Every NACA airfoil has two charts to present the lift, drag, and moment

coefficient data for the airfoil. The first chart will have curves of lift

coefficient versus angle of attack at various Reynolds numbers and

curves of moment coefficient at the quarter chord point versus angle of

attack at various Reynolds numbers. See the chart below. In addition to

the lift and moment coefficients, the stall angle of attack and zero lift

angle of attack can be determined.

The second chart will have curves of drag coefficient versus lift

coefficient at various Reynolds numbers and curves of moment

coefficient at the aerodynamic center versus lift coefficient at various

Reynolds numbers. In addition to smooth airfoils, it is common for data

for an airfoil whose leading edge has a sandpaper surface texture to be

included. The second chart also has an insert picture of the air foil

geometry and the aerodynamic center for the airfoil at different Reynolds

numbers is provided in tabular form.

7.3.2 COMPRESSIBILITY EFFECTS

For Mach number less than 0.3, we may assume that our flow is

incompressible and the standard airfoil charts work very well. For Mach

numbers greater than 0.3, we must correct the lift coefficient using the

Prandtl-Glauert correction which gives

CL = CL chart / SQRT (1-M2)

This is valid for Mach numbers up to 0.7.

7.3.3 AIRFOIL CATEGORIES

The following are airfoil categories:

NACA 4 Digit

1st digit: maximum camber (as % of chord).

2nd digit (x10): location of maximum camber (as % of

chord from leading edge (LE)).

3rd & 4th digits: maximum section thickness (as % of chord).

41

NACA 5 Digit

1st digit (x0.15): design lift coefficient.

2nd & 3rd digits (x0.5): location of maximum camber (as % of

chord from LE).

4th & 5th digits: maximum section thickness (as % of chord).

NACA 6 Digit

1st digit: identifies series type.

2nd digit (x10): location of minimum pressure (as % of chord from

leading edge (le)).

3rd digit: indicates acceptable range of cl above/below design

value for satisfactory low drag performance.

4th digit (x0.1): design cl.

5th & 6th digits: maximum section thickness (%c)

7.4 SELECTED AEROFOIL

Airfoil: NACA 23012

Fig8.4 Airfoil NACA 23012

CL = 0.3 at angle of attack 0

CL max = 1.6

42

CHAPTER-8

WING SELECTION AND WING LOADING

8.1 INTRODUCTION

After the final weight estimation of the aircraft, the primary

component of the aircraft to be designed is the wing. The wing weight

and its lifting capabilities are in general, a function of the thickness of the

airfoil section that is used in the wing structure. The first step towards

designing the wing is the thickness estimation. The thickness of the wing,

in turn depends on the critical mach number of the airfoil or rather, the

drag divergence Mach number corresponding to the wing section.

The wing may be considered as the most important component of an

aircraft, since a fixed-wing aircraft is not able to fly without it. Since the

wing geometry and its features are influencing all other aircraft

components, we begin the detail design process by wing design. The

primary function of the wing is to generate sufficient lift force or simply

lift (L).

However, the wing has two other productions, namely drag force or

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

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

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

lift is produced due to the pressure difference between lower and upper

surfaces. Aerodynamics textbooks are a good source to consult for

information about mathematical techniques for calculating the pressure

distribution over the wing and for determining the flow variables.

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

43

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 (or washout) (t)

13. Sweep angle ()

14. Dihedral angle ()

15. Incidence (iw) (or setting angle, set)

16. High lifting devices such as flap

17. Aileron

18. Other wing accessories

8.2 NUMBER OF WINGS

One of the decisions a designer must make is to select the number of

wings. The options are:

1. Monoplane (i.e. one wing)

2. Two wings (i.e. biplane)

3. Three wings

Fig 8.1Configuration of wing

44

8.3 WING VERTICAL LOCATION

One of the wing parameters that could be determined at the early

stages of wing design process is the wing vertical location relative to the

fuselage centerline. This wing parameter will directly influence the

design of other aircraft components including aircraft tail design, landing

gear design, and center of gravity. In principle, there are four options for

the vertical location of the wing. They are:

Fig 8.2 Position of wing

8.4 SELECTED WING IS LOW WING

8.4.1LOW WING

In this section, advantages and disadvantages of a low wing

configuration (Figure 9.2-c) will be presented. Since the reasons for

several items are similar with the reasons for a high wing configuration,

the reasons are not repeated here. In the majority of cases, the

specifications of low wing are compared with a high wing configuration.

8.4.1.1 ADVANTAGES

1. The aircraft take off performance is better; compared with a high

wing configuration; due to the ground effect.

2. The pilot has a better higher-than-horizon view, since he/she is

above the wing.

3. The retraction system inside the wing is an option along with inside

the fuselage.

45

4. Landing gear is shorter if connected to the wing. This makes the

landing gear lighter and requires less space inside the wing for

retraction system. This will further make the wing structure lighter.

5. In a light GA aircraft, the pilot can walk on the wing in order to get

into the cockpit.

6. The aircraft is lighter compared with a high wing structure.

7. Aircraft frontal area is less.

8. The application of wing strut is usually no longer an option for the

wing structure.

9. Item 8 implies that the aircraft structure is lighter since no strut is

utilized.

10. Due to item 8, the aircraft drag is lower.

11. The wing has less induced drag.

12. It is more attractive to the eyes of a regular viewer.

13. The aircraft has higher lateral control compared with a high wing

configuration, since the aircraft has less lateral static stability, due to

the fuselage contribution to the wing dihedral effect.

14. The wing has less downwash on the tail, so the tail is more effective.

15. The tail is lighter; compared with a high wing configuration.

16. The wing drag is producing a nose-down pitching moment, so a low

wing is longitudinally stabilizing. This is due to the lower position

of the wing drag line relative to the aircraft center of gravity.

8.4.1.2 DISADVANTAGES

1. The wing generates less lift; compared with a high wing

configuration; since the wing has two separate sections.

2. With the same token to item 1, the aircraft will have higher stall

speed; compared with a high wing configuration; due to a lower

CLmax.

3. Due to item 2, the take-off run is longer.

4. The aircraft has lower airworthiness due to a higher stall speed.

5. Due to item 1, wing is producing less induced drag.

46

6. The wing has less contribution to the aircraft dihedral effect, thus

the aircraft is laterally dynamically less stable.

7. Due to item 6, the aircraft is laterally more controllable, and thus

more maneuverable.

8. The aircraft has a lower landing performance, since it needs more

landing run.

9. The pilot has a lower lower-than-horizon view. The wing below

the pilot will obscure part of the sky for a fighter pilot.

8.5 WING LOADING

In aerodynamics, wing loading is the loaded weight of the aircraft

divided by the area of the wing. The faster an aircraft flies, the more lift

is produced by each unit area of wing, so a smaller wing can carry the

same weight in level flight, operating at a higher wing loading.

L = W = (1/2) V2

S CL

Vstall = SQRT ((2xW) / ( S CL))

(W/S) = V2

stall CL / 2

= (289.228x0.25)2x

(1.225) x (0.3)/2

(W/S) = 961.17 N/m2

8.6 WING GEOMETRY DESIGN

The geometry of the wing is a function of four parameters, namely the

Wing loading (W/S), Aspect Ratio (b2/S), Taper ratio () and the

Sweepback angle at quarter chord.

The Take-off Weight that was estimated in the previous analysis is

used to find the Wing area S (from W/S).The value of S also enables us

to calculate the Wingspan b (using the Aspect ratio). The root chord can

now be found using the equation.

The root chord is given by,

Croot = (2 x S) / b (1+)

The tip chord is given by,

Ctip = x Croot

The mean chord is given by,

47

Cmean = (2/3) Croot x (1++2) / (1+)

8.6.1 Croot CALCULATION

Croot = (2 x S) / b x (1+)

= (2 x 350) / (54 x 1.25)

Croot = 10.37m

8.6.2 Ctip CALCULATION

Ctip = x Croot

Ctip = 2.59m

8.6.3 Cmean CALCULATION

Cmean = (2/3) Croot x (1++2) / (1+)

= (2/3) x10.37 x (1.05)

Cmean = 7.259m

8.7 LIFT ESTIMATION

8.7.1 LIFT

Component of aerodynamic force generated on aircraft perpendicular

to flight direction.

Fig 8.3 Forces acting in aircraft

8.7.2 LIFT COEFFICIENT (CL)

Amount of lift generated depends on:

Planform area (S),

48

Air density (),

Flight speed (V),

Lift coefficient (CL)

Lift is given by,

Lift = (1/2)V2SCL

CL is a measure of lifting effectiveness and mainly depends upon:

Section shape,

Planform geometry,

Angle of attack (),

Compressibility effects,

Viscous effects (Reynolds number).

8.7.3 GENERATION OF LIFT

Aerodynamic force arises from two natural sources:

o Variable pressure distribution.

o Shear stress distribution.

Shear stress primarily contributes to overall drag force on aircraft.

Lift mainly due to pressure distribution, especially on main lifting

surfaces, i.e. wing.

Require (relatively) low pressure on upper surface and higher

pressure on lower surface.

Any shape can be made to produce lift if either cambered or

inclined to flow direction.

Classical aerofoil section is optimum for high subsonic lift/drag ratio.

49

Fig 8.4 Pressure distribution on airfoil

8.7.4 PRESSURE VARIATIONS WITH ANGLE OF ATTACK

Negative (nose-down) pitching moment at zero-lift (negative ).

Positive lift at = 0.

Highest pressure at LE stagnation point, lowest pressure at crest on

upper surface.

Peak suction pressure on upper surface strengthens and moves

forwards with increasing .

Most lift from near LE on upper surface due to suction.

Fig8.5 Airfoils at different angle of attack

50

8.7.5 LIFT CURVES

Fig8.6 Lift curve

8.7.6 LIFT CALCULATION:

General Lift equation is given by,

Lift = (1/2) V2SCL

8.7.6.1 LIFT AT CRUISE

= 0.27641 kg/m3 (at the cruising altitude of 12750m)

V = 289.228 m/s

S = 350m2

CL (cruise) = 0.6 (from the wing and airfoil estimation)

Substituting all these values in the general lift equation,

L (cruise) = 1/20.27859 (289.228)2 350 0.6

Lift at cruise = 2427.860 kN

8.7.6.2 LIFT AT TAKEOFF

= 1.225 kg/m3 (at sea-level)

V = 0.7 x Vlo

51

= 0.7 x 1.2 x Vstall

= 60.738 m/s

S = 350m2

CL(take-off) = 1.2251(flaps kept at the take-off position of 20)


L(take-off) = 1/2 1.225 (60.738)2 350 1.2251

Lift at take-off = 968.873 kN

8.7.6.3 LIFT AT LANDING


V = 0.7 x Vlo


= 65.799 m/s

S = 350m2

CL(Landing) = 1.6 (flaps kept at the take-off position of 40)


L(Landing) = 1/2 1.225 (65.799)2 350 1.6

Lift at Landing = 1485.0214 kN

52

CHAPTER-9

DRAG ESTIMATION

9.1 DRAG

Drag is the resolved component of the complete aerodynamic force

which is parallel to the flight direction (or relative oncoming

airflow).

It always acts to oppose the direction of motion.

It is the undesirable component of the aerodynamic force while lift

is the desirable component.

9.2 DRAG COEFFICIENT (CD)

Amount of drag generated depends on:

Planform area (S), air density (), flight speed (V), drag coefficient

(CD)

CD is a measure of aerodynamic efficiency and mainly depends

upon:

Section shape, planform geometry, angle of attack (),

compressibility effects (Mach number), viscous effects (Reynolds

number).

9.3 DRAG COMPONENTS

9.3.1 SKIN FRICTION

Due to shear stresses produced in boundary layer.

Significantly more for turbulent than laminar types of boundary

layers.

Fig9.1 Skin friction drag

53

9.3.2 FORM (PRESSURE) DRAG

Due to static pressure distribution around body - component

resolved in direction of motion.

Sometimes considered separately as fore body and rear (base) drag

components.

Fig9.2 Form drag

9.3.3 WAVE DRAG

Due to the presence of shock waves at transonic and supersonic

speeds.

Result of both direct shock losses and the influence of shock waves

on the boundary layer.

Fig9.3 Shock formation over wedge

54

9.4 TYPICAL STREAMLINING EFFECT

Fig9.4 Flow around different shapes

9.5 LIFT INDUCED (OR) TRAILING VORTEX DRAG

Fig9.5 Downwash region in wings

9.6 CALCULATION

Generally for jet aircrafts, it is given that

CD,0 = 0.0030

e = 0.8

55

The general drag equation is given by,

= ( ) 2 (,0 + ( / 2))

For calculating , we use the formula,

= (16 h/b)2 / (1 + (16 h/b)2 )

Where,

h= 2m

b= 65m

= 0.2599

9.6.1 DRAG AT CRUISE

= 0.27641 kg/m3 (at the cruising altitude of 12750m)

V = 289.228 m/s

S = 350m2

CL(cruise) = 0.523 (from the wing and airfoil estimation)

Substituting all these values in the general drag equation,

D(cruise) = 1/2x0.27641 x (289.228)2 x 350 x 5.65333 x 10

-3

Drag at cruise = 30.992 kN

9.6.2 DRAG AT TAKEOFF


V = 0.7 x Vlo


= 60.738 m/s

S = 350m2

CL(take-off) = 0.6257(flaps kept at the take-off position of 20)


D(take-off) = 1/2x 1.225 x (60.738)2 x 350 x (6.7583 x 10

-3 )

Drag at take-off = 17.7235 kN

9.6.3 DRAG AT LANDING


V = 0.7 x Vlo


= 65.799 m/s

56

S = 350 m2

CL(Landing) = 0.65 (flaps kept at the take-off position of 40)


D(Landing) = 1/2x 1.225 x (65.799)2 x 350 x (7.54656 x10

-3 )

Drag at Landing = 33.513 kN

57

CHAPTER-10

PERFORMANCE CHARACTERISTICS

10.1 TAKE-OFF PERFORMANCE

Distance from rest to clearance of obstacle in flight path and

usually considered in two parts:

o Ground roll - rest to lift-off (SLO)

o Airborne distance - lift-off to specified height (35 ft FAR, 50

ft others).

The aircraft will accelerate up to lift-off speed (Vlo = about 1.2 x

Vstall) when it will then be rotated.

A first-order approximation for ground roll take-off distance may be

made from:

SLO = (1.44 x W2) / ( g S CLmax T)

Slo may be reduced by increasing T, S or Cl,max (high lift devices

relate to latter two).

An improved approximation for ground roll take-off distance may

be made by including drag, rolling resistance and ground effect

terms.

The bracketed term will vary with speed but an approximation may

be made by using an instantaneous value for when V = 0.7 x Vlo.

In the above equation:

Where,

accounts for drag reduction when in ground effect

58

is calculated by using the following formula,

Where,

h = height above ground,

b = wing span.

r = 0.02 for smooth paved surface, 0.1 for grass.

10.1.1 CALCULATION

= (3.8777635 x 1012

) / (1392870392)

SLO = 1685.159 m

Take-off runway Distance = 1685.159 m

10.2 CLIMBING

Consider aircraft in a steady unaccelerated climb with vertical climb

speed of Vc.

Fig10.1 Aircraft during climb

59

Force balance gives:

= (53427635 x 10

3) / (167278.608 x 9.81)

R/Cmax = 19.989 m/s 10.3 MANOEUVRES / TURNING FLIGHT

An aircraft is capable of performing many different types of turns and

maneuvers.

Three of the more common turns will be considered here in simplistic

terms:

Constant altitude banked turn.

Vertical pull-up maneuver.

Vertical pull-down maneuver.

In the case of a commercial transport aircraft, it is capable of

performing only a constant altitude banked turn and not any vertical pull-

up or pull-down maneuver.

10.3.1 CONSTANT ALTITUDE BANKED TURN

In steady condition:

- T = D


So for given speed and turn radius there is only one correct bank angle

for a co-ordinate (no sideslip) turn.

Maneuverability equations simplified through use of normal load factor

(n) = L/W.

60

In the turn, n = L/W = sec > 1 and is therefore determined by bank

angle.

Turn radius (R) and turn rate () are good indicators of aircraft

maneuverability.

10.3.2 CALCULATION

W = L cos

Let = 30

n = (L/W)

= 1.572

R = 7030.37 m

= (V/R)

= 0.0411 rad/sec

10.4 GLIDING

Similar to the steady unaccelerated case but with T = 0.

Fig10.2 Aircraft at gliding


= 4.085

61

10.5 LANDING PERFORMANCE:

Landing performance consists of three phases:

Airborne approach at constant glide angle (around 3o) and

constant speed.

Flare - transitional maneuvers with airspeed reduced from

about 1.3 Vstall down to touch-down speed.

Ground roll - from touch-down to rest.

Ground roll landing distance (s3 or sl) estimated from:

Where,

Vav may be taken as 0.7 x touch-down speed (Vt or V2)

Vt is assumed as 1.3 x Vstall.

r is higher than for take-off since brakes are applied - use r = 0.4

for paved surface.

If thrust reversers (Tr) are applied, use:

10.5.1 CALCULATION

= (4074.34 + 0.4 (167278.608 x 9.81 373084)

= 11193.283 m

Landing Runway Distance = 1193.283 m

62

CHAPTER 11

CALCULATION OF CENTER OF GRAVITY

11.1 INTRODUCTION

The precise location of the aircraft cg is essential in the positioning of

the landing gear, as well as for other MDO applications, e.g., flight

mechanics, stability and control, and Performance. Primarily, the aircraft

cg location is needed to position the landing gear such that ground

stability, maneuverability, and clearance requirements are met. Given the

fact that none of the existing conceptual design-level cg estimation

procedures has the degree of responsiveness and accuracy required for

MDO applications, a new approach is formulated to provide a reliable

range of cg locations that is better suited for MDO applications.

11.2 CURRENT CAPABILITIES

Although not expected to determine the location of the aircraft cg,

current aircraft sizing programs, as typified by Jayaram et al. and

McCullers, do provide some rudimentary estimates. These codes use

estimated component weights obtained from statistical weight equations,

and either user-specified or default component cg locations to arrive at

the overall aircraft cg location. The lack of responsiveness and accuracy

has rendered current approaches inadequate for MDO application.

Fig 11.1 Cg Range

The lack of responsiveness is attributed to the fact that each aircraft

component is assigned a specific location within the airframe. Typically,

63

these approaches do not estimate the operational range of cg locations.

The cg location is a complicated function of the configuration, loading,

and fuel state, with an allowable range limited by a number of

operational factors. Although a range of cg locations can be established

by varying the configuration, equipment arrangement, and payload and

fuel states individually, the process is difficult. The accuracy limitations

arise because the codes assume that the user has the experience and

knowledge required to make adjustments to the component weight and cg

estimates. Unfortunately, this approach is not suitable for use in

automated procedures required in MDO.

Evidently, what is needed is a new approach which is capable of

establishing a maximum permissible cg range for a given configuration.

This available cg range can then be compared with the desired

operational cg range obtained from performance, control, and operational

requirements. If the desired cg range is within the available cg range, the

concept is viable and can be balanced. If not, the configuration must be

changed, either by the designer or an MDO procedure if an automated

process is being used.

11.3 ALTERNATE METHOD

Component location flexibility at the conceptual design phase is

actively exploited as a means to improve the responsiveness and accuracy

of current cg estimation procedures. In the proposed procedure, aircraft

components are assigned a range of cg locations based on the geometry,

as well as physical and functional considerations, associated with each

component. By arranging the cg of the components at their fore- and aft-

most limits, the maximum permissible cg range of a particular layout can

be established. This cg range can then be used by an MDO procedure to

determine the forward and aft aircraft cg limits required to meet

performance and stability and control considerations. Adjusted for

uncertainty, this maximum permissible cg range can be used as a

constraint for the operational cg range during the optimization.

64

11.4 ESTABLISHMENT OF COMPONENT CG RANGE

The assignment of component cg range is based on the geometry,

planform, and the type of components involved. In the case of the

primary components, e.g., fuselage, wing, and empennage, the location of

these items remains relatively unchanged once the concept is frozen.

Consequently, the cg range is expected to be centered near the volumetric

center of the component and is unlikely to shift too much. For ease of

identification, the primary components will be referred to as the

constrained items.

As for secondary components, e.g., equipment and operational items,

the location of each component varies from one aircraft concept to

another, depending on the philosophy and preference of the airframe

manufacturer. Note that as long as the stowage and functionality

constraints are not violated, these components can be assigned to any

available space throughout the aircraft due to their compactness.

Consequently, the corresponding cg range is defined by the forward and

aft boundaries of the stowage space within which the item is located.

Accordingly, these components are termed the unconstrained items.

Although the payload and passenger amenity, i.e., furnishings and

services, are confined within the cargo holds and cabin, operational

experience has shown that the cg location of these items varies according

to the loading condition and cabin layout as specified by the airlines,

respectively. Similarly, the cg location of the fuel varies as a function of

time as the fuel is being consumed during the duration of the mission.

Given the added freedom in terms of the loading pattern, these

components are also classified as unconstrained items.

11.5 GENERIC COMPONENT LAYOUT

The ranges are based on the layout of existing commercial transports

and can be modified to accommodate any unique layout of the aircraft

concept under consideration. The locations of the front and rear spar for

the wing and empennage are dictated by space required for housing the

control surfaces and the associated actuation systems, where values of 15

65

and 65 percent chord, respectively, are typically used. As in the

conventional cantilever wing and empennage construction, the majority

of the structure, i.e., bulkheads, ribs, and fuel tanks, are located between

the front and rear spars. Thus, it can be expected that the cg of the wing is

most likely to be located between the two, along the respective mean

aerodynamic chords (mac). In addition, given the physical arrangement

of the fuel tanks, the cg of the fuel and the fuel system can be expected to

be located near the same vicinity.

The cg of the fuselage depends on the structural arrangement of the

pressure bulkheads, Frames, and the aft-body taper ratio. Other factors

include local structural reinforcement around the landing gear wheel

wells, cargo holds, and the layout of the cabin, e.g., a forward upper-deck

as found on the Boeing Model 777 or a double-decker as found on the

proposed ultra-high-capacity transports. Taking these factors into

consideration, the proposed procedure assumes that the cg of the fuselage

is most likely to be located between 40 and 50 percent of the fuselage

length.

11.6 WEIGHT AND BALANCE

When the weight of the aircraft is at or below the allowable limit(s) for

its configuration (parked, ground movement, takeoff, landing, etc.) and

its center of gravity is within the allowable range, and both will remain so

for the duration of the flight, the aircraft is said to be within weight and

balance. Different maximum weights may be defined for different

situations For example, large aircraft may have maximum landing

weights that are lower than maximum takeoff weights (because some

weight is expected to be lost as fuel is burned during the flight). The

center of gravity may change over the duration of the flight as the

aircraft's weight changes due to fuel burn or by passengers moving

forward or aft in the cabin.

11.7 ARM

The arm is the horizontal distance from the reference datum to the

center of gravity (CG) of an item. The algebraic sign is plus (+) if

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measured aft of the datum or to the right side of the center line when

considering a lateral calculation. The algebraic sign is minus (-) if

measured forward of the datum or the left side of the center line when

considering a lateral calculation.

11.8 MOMENT

The moment is the moment of force, or torque that results from an

objects weight acting through an arc that is centered on the zero point of

the reference datum distance. Moment is also referred to as the tendency

of an object to rotate or pivot about a point (the zero point of the datum,

in this case). The further an object is from this point, the greater the force

it exerts. Moment is calculated by multiplying the weight of an object by

its arm.

11.9 MEAN AERODYNAMIC CHORD (MAC)

A specific chord line of a tapered wing. At the mean aerodynamic

chord, the center of pressure has the same aerodynamic force, position,

and area as it does on the rest of the wing. The MAC represents the width

of an equivalent rectangular wing in given conditions. On some aircraft,

the center of gravity is expressed as a percentage of the length of the

MAC. In order to make such a calculation, the position of the leading

edge of the MAC must be known ahead of time. This position is defined

as a distance from the reference datum and is found in the aircraft's flight

manual and also on the aircraft's type certificate data sheet. If a general

MAC is not given but a LeMAC (leading edge mean aerodynamic chord)

and a TeMAC (trailing edge mean aerodynamic chord) are given (both of

which would be referenced as an arm measured out from the datum line)

then your MAC can be found by finding the difference between your

LeMAC and your TeMAC.

67

11.10 CALCULATION

Fig11.2

CG WITHOUT WING

X = (Wcrew x 5.89) + (Wengine x 25) + (Wpassengers x 36) + (Wbaggage x 46)

Wcrew + Wengine + Wpassengers + Wbaggage

Where,

Wcrew = 464.94 kg

Wengine = 7810 kg (for two engine)

Wpassengers = 23814 kg

Wbaggage = 680.4 kg

X = (464.94 x 5.41) + (12320 x 30) + (19845 x 36) + (3402 x 46)

464.94 + 12320 + 19845 + 3402

Xwithout wing = 34.67 m

CG WITHOUT WING

Wtotal = Wcrew + Wengine + Wpassengers + Wbaggage

Wtotal = 32768.34 kg

CG WITH WING

Xtotal = (Wtotal x 34.67) + (Wwing x (34.67 + Xwing ))

Wtotal + Wwing

5.89m 20 m 11.99m 8.72m

Wcrew Wengine Wpassengers Wbaggage

68

CG for wing

Xwing = (Ctip + Croot + Cmean ) x S

3 x S

(S = 350 m2)

Xwing = (2.59 + 7.259 + 10.37) x (350)

3 x 350

Xwing = 6.739 m

Wwing = 2.5 x S

Wwing = 875 kg

Xtotal = (32769.34 x 34.67) + (875 x (34.67 + 6.739))

32769 .34 + 875

Xtotal = 34.845 m

69

CHAPTER-12

VIEWS OF DESIGNED AIRCRAFT

12.1. TOP VIEW

Fig12.1Top view

65 m

70

12.2. FRONT VIEW

Fig12.2 Front view

17 m

71

12.3. SIDE VIEW

Fig 12.3 side view

54 m

72

CONCLUSION

Design is a fine blend of science, presence of mind and the application of

each one of them at the appropriate time. Design of anything needs experience

and an optimistic progress towards the ideal system. The scientific society

always look for the best product design .This involves a strong fundamental in

science and mathematics and their skill full application which is a tough job

endowed upon the designer . We had put enough hard work to the best of our

knowledge for this design project. A design never gets completed in a flutter

sense but it is one further step towards the ideal system. But during the design of

this passenger aircraft we learnt about aeronautics and its implications when

applied to an aircraft.Thus a conceptual design of a 250 seater passenger aircraft

has been successfully done. The Aircraft is a twin engine configuration. It uses

two RR Trent 768 engines which fulfills the power requirement. The wing is

B737 A/il airfoil.

73

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