Aircraft Design Project - 280 Seat Transport

HINDUSTAN COLLEGE OF ENGINEERING

AIRCRAFT DESIGN PROJECT – 1 INTERNATIONAL MEDIUM-RANGE 280 SEATER PASSENGER AIRCRAFT

SUBMITTED BY: ROBIN RICHARD RAJAN. R SARAVANAN. T RAJESH KUMAR. K

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HINDUSTAN COLLEGE OF ENGINEERING

AIRCRAFT DESIGN PROJECT – 1 REPORT

NAME OF THE STUDENT: NAME OF THE PROJECT : DEPARTMENT :

Certified that this a bonafide record of the work done by of VI semester AERO (B.E.) during the year 2009-2010 on DESIGN OF INTERNATIONAL MEDIUM RANGE 280 SEATER PASSENGER AIRCRAFT.

INT. Examiner

Staff Member Incharge

EXT. Examiner

Name of examination: B.E. DEGREE Registration number: 305071010

Aircraft Design Project - 1

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ACKNOWLEDGEMENT

I would like to extent my heartfelt thanks to Prof . P. K. Dash (Head of Aeronautical Department) for giving me his able support and encouragement. encouragement. At this juncture I must emphasis the point that this DESIGN PROJECT would not have been possible without the highly informative and valuable guidance by Prof. P. S. Venkatanarayanan, whose vast knowledge and experience has must us go about this project with great ease. We have great pleasure in expressing our sincere & whole hearted gratitude to them. It is worth mentioning about my team mates, friends and colleagues of the Aeronautical department, for extending their kind help whenever whenever the necessity arose. I thank one and and all who have directly directly or indirectly helped me in making this design project a great success.


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INDEX Serial No.

Topic

Page No.

1

Aim of the Project

5

2

Abstract

7

3

Introduction

9

4

Comparative Data Sheet

16

5

Graphs

20

6

Mean Design Parameters

39

7

Weight Estimation

41

8

Powerplant Selection

49

9

Fuel Weight Validation

53

10

Wing Selection

55

11

Airfoil Selection

60

12

Lift Estimation

70

13

Drag Estimation

75

14

Landing Gear Arrangement

81

15

Fuselage Design

87

16

Performance Characteristics

94

17

3 – View Diagram

100

18

Conclusion

104

19

Bibliography

106


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ABBREVIATION A.R. B C C root C tip C

Cd Cd,0 Cp CL D E E L (L/D)loiter (L/D)cruise M Mff R Re S Sref Swet Sa Sf Sfr Sg T Tcruise Ttake-off (T/W)loiter (T/W)cruise (T/W)take-off Vcruise Vstall Vt Wcrew Wempty Wfuel Wpayload W0 W/S

 

 R/C

-

Aspect Ratio Wing Span (m) Chord of the Airfoil (m) Chord at Root (m) Chord at Tip (m) Mean Aerodynamic Chord (m) Drag Co-efficient Zero Lift Drag Co-efficient Specific fuel consumption (lbs/hp/hr) Lift Co-efficient Drag (N) Endurance (hr) Oswald efficiency Lift (N) Lift-to-drag ratio at loiter Lift-to-drag ratio at cruise Mach number of aircraft Mission fuel fraction Range (km) Reynolds Number Wing Area (m²) Reference surface area Wetted surface area Approach distance (m) Flare Distance (m) Free roll Distance (m) Ground roll Distance (m) Thrust (N) Thrust at cruise (N) Thrust at take-off (N) Thrust-to-weight ratio at loiter Thrust-to-weight ratio at cruise Thrust-to-weight ratio at take-off Velocity at cruise (m/s) Velocity at stall (m/s) Velocity at touch down (m/s) Crew weight (kg) Empty weight of aircraft (kg) Weight of fuel (kg) Payload of aircraft (kg) Overall weight of aircraft (kg) Wing loading (kg/m²)

-

Density of air (kg/m³) Dynamic viscosity (Ns/m²)

-

Tapered ratio Rate of Climb


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AIM OF THE PROJECT


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AIM OF THE PROJECT The aim of this design project is to design a 280 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 280 passengers and their baggage over a design range of 7200 km at a cruise speed of about 872 km/h.

 To provide the passengers with high levels of safety and comfort.  To operate from regional and international airports.  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.


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ABSTRACT


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ABSTRACT The purpose of the project is to design a 280 seater Medium Range International passenger aircraft. The aircraft will possess a low wing, tricycle landing gear and a conventional tail arrangement. Such an aircraft must possess a wide body configuration to provide sufficient seating capacity. It must possess turbofan engines to provide the required amount of speed, range and fuel economy for the operator. The aircraft will possess three engines.


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INTRODUCTION


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INTRODUCTION At the instant time there are different types of aircrafts with latest technology. Every year there is a great competition for making an aircraft of having higher capacity of members inside the aircraft. So here in this report, We intend to implant the differentiation among the aircrafts having sitting capacity of 250-350 members. This report gives the different aspects of specifications like wing specification, weight specification, power plant specification and performance specification. Airbus started the development of a very large airliner (termed Megaliner by Airbus in the early development stages) in the early 1990s, both to complete its own range of products and to break the dominance that Boeing had enjoyed in this market segment since the early 1970s with its 747. McDonnell Douglas pursued a similar strategy with its ultimately unsuccessful MD-12 design. As each manufacturer looked to build a successor to the 747, they knew there was room for only one new aircraft to be profitable in the 600 to 800 seat market segment. Each knew the risk of splitting such a niche market, as had been demonstrated by the simultaneous debut of the Lockheed L-1011 and the McDonnell Douglas DC-10: both planes met the market’s needs, but the market could profitably sustain only o ne model, eventually resulting in Lockheed's departure from the civil airliner business. In January 1993, Boeing and several companies in the Airbus consortium started a joint feasibility study of an aircraft known as the Very Large Commercial Transport (VLCT), aiming to form a partnership to share the limited market. Airplanes come in many different shapes and sizes depending on the mission of the aircraft, but all modern airplanes have certain components in common. These are the fuselage, wing, tail assembly and control surfaces, landing gear, and powerplant. 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 t o 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. It may also carry armaments of various sorts. Some aircraft carry fuel in the fuselage; others carry the fuel in the wings. In addition, an engine may be housed in the fuselage. 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.


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PURPOSE AND SCOPE OF AIRPLANE DESIGN OBJECTIVES To meet the FUNCTIONAL, OPERATIONAL and SAFETY requirements set out OR acceptable to the USER.



ACTUAL PROCESS OF DESIGN

    



Selection of aircraft type and shape Determination of geometric parameters Selection of power plant Structural design and analysis of various components Determination of aircraft flight and operational characteristics .

How to get the BEST POSSIBLE solution to meet the simultaneous requirements?

 Very complex and long drawn-out process  Meeting higher performance requirements than similar aircraft already in service.

 Role of Design Laboratories and R&D Institutions.  Trial and Error, in an ingenious fashion.

3 DISTINCT STAGES OF AIRCRAFT DESIGN  Project Feasibility Study  Preliminary Design 

Design Project

PROJECT FEASIBILITY STUDY (to evolve a satisfactory specification)

      

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

PRELIMINARY 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


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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 Drafting the preliminary 3-view Drawings

DESIGN PROJECT

    

Internal discussions

       

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

Discussions with prospective customers Discussions with Certification Authorities Consultations with suppliers of power plant and m ajor accessories Deciding upon a BROAD OUTLINE to start the ACTUAL DESIGN, which will consist of Construction of Mock-up 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


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SEVEN INTELLECTUAL POINTS FOR CONCEPTUAL DESIGN


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DESIGN 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


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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


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COMPARATIVE DATASHEET


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Comparative Datasheet - 1 Airbus Aircrafts Parameter

Units

1

2

3

4

5

Name

(no unit)

A300-600R

A310-300

A330-300

A340-500

A350-800

Total Seating Capacity

(no unit)

266

240

295

313

270

Aircraft Dimensions Length

m

54

46.6

63.6

67.9

60.7

Height

m

16.62

15.8

16.85

17.1

17.2

Fuselage Diameter

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

m

260

219

361.6

439.4

443

Wing Sweep

degree

28°

28°

30°

31.1°

31.9°

2

Performance Cruising Altitude

m

10,668

9,998

10,972

10,972

12,192

Service ceiling

m

12,000

12,497

12,527

12,527

13,137

Range

Km

7,540

9,600

10,500

16,060

15,000

Cruising Speed

Km/h

829

850

871

881

903

Max Speed

Km/h

871

901

913

913

945

Number of Engines

(no unit)

2

2

2

4

2

Max thrust capability

kN

311.4

262.5

320

249

374

Design Weights 3

Kg

171.7

164

233

372

268

3

Kg

90.9

83.1

124.5

170.9

115.7

2

660.38

748.86

644.36

846.61

604.96

68,150

75,470

97,170

2,14,810

1,29,000

MTO Weight

x10

Empty Weight

x10

Wing Loading

Kg/m

Max Fuel Capacity

litre


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Comparative Datasheet - 2 Boeing Aircrafts Parameter

Units

6

7

8

9

10

Name

(no unit)

707-320B

757-200

767-200

777-200

787-9


(no unit)

202

234

290

301

280


m

46.61

47.32

48.5

63.7

62.8

Height

m

12.93

13.56

16.8

18.5

16.9

Fuselage Diameter

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

(no unit)

7.1

7.98

7.99

8.67

9.4

Wing Area

m

273.7

181.25

283.3

427.8

325.3

Wing Sweep

degree

35°

25°

31.5°

31.64°

32.2°

2


m

10,058

10,668

10,668

10,668

12,192

Service ceiling

m

11,887

12,802

11,887

13,137

13,106

Range

Km

10,650

7,600

7,300

9,695

15,000

Cruising Speed

Km/h

972

850

851

905

903

Max Speed

Km/h

1,010

935

913

950

945

Number of Engines

(no unit)

4

2

2

2

2


kN

320.4

193

222

330

320

Design Weights 3

Kg

151.32

115.68

142.88

247.2

248

3

Kg

66.4

57.18

81.23

134.8

115

2

552.87

638.23

504.34

577.84

762.37

90,160

43,490

90,770

117,000

127,000

MTO Weight

x10

Empty Weight

x10

Wing Loading

Kg/m

Max Fuel Capacity

litre


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Comparative Datasheet - 3 Other Aircrafts Parameter

Units

Name

(no unit)


(no unit)

11 Lockheed L-1011-200 263

12 Ilyushin IL-96-300 300

13 Tupolev Tu-204-100 210

14 Douglas DC-8-63CF 259

15 Tupolev Tu-114 220


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

6.78

5.82

4.40

6.01

6.08

Aspect Ratio

(no unit)

6.98

10.32

9.48

7.52

8.39

Wing Area

m

321.1

350

184.2

271.9

311.1

Wing Sweep

degree

35°

30°

30°

32°

35°

2


m

10,257

10,668

12,100

10,668

8,991

Service ceiling

m

10,668

13,106

12,588

12,497

11,887

Range

Km

7,420

10,400

5,650

3,445

6,200

Cruising Speed

Km/h

935

860

830

876

770

Max Speed

Km/h

990

900

900

965

870

Number of Engines

(no unit)

3

4

2

4

4


kN

222.4

157

158.3

84.5

60

Design Weights 3

Kg

211

250

103

161

175

3

Kg

105.1

120.4

60

66.36

91 to 93

2

657.11

714.28

559.17

592.12

562.52

99,935

152,620

41,000

66,243

71,615

MTO Weight

x10

Empty Weight

x10

Wing Loading

Kg/m

Max Fuel Capacity

litre


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GRAPHS


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Graph 1 Cruising Speed vs. Length

Length = 55.0m


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Graph 2 Cruising Speed vs. Height

Height = 15.7m


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Graph 3 Cruising Speed vs. Fuselage Diameter

Fuselage Diameter = 5.26m


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Graph 4 Cruising Speed vs. Wing Span

Wing Span = 51.5m


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Graph 5 Cruising Speed vs. Chord

Chord = 6.0m


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Graph 6 Cruising Speed vs. Aspect Ratio

Aspect Ratio = 8.6


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Graph 7 Cruising Speed vs. Wing Area

Wing Area = 348m2


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Graph 8 Cruising Speed vs. Wing Sweep

Wing Sweep = 31.5°


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Graph 9 Cruising Speed vs. Cruising Altitude

Cruising Altitude = 10800m


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Graph 10 Cruising Speed vs. Service Ceiling

Service Ceiling = 12000m


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Graph 11 Cruising Speed vs. Range

Range = 7200m


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Graph 12 Cruising Speed vs. Maximum Speed

Max Speed = 940km/h


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Graph 13 Cruising Speed vs. Number of Engines

Number of Engines = 3


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Graph 14 Cruising Speed vs. Maximum Thrust Capability

Maximum Thrust Capability = 265kN


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Graph 15 Cruising Speed vs. Maximum Take Off Weight

Maximum Take Off Weight = 272000 kg


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Graph 16 Cruising Speed vs. Empty Weight

Empty Weight = 85000 kg


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Graph 17 Cruising Speed vs. Wing Loading

Wing Loading = 710 kg/m3


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Graph 18 Cruising Speed vs. Maximum Fuel Capacity

Maximum Fuel Capacity = 100000 litre


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MEAN DESIGN PARAMETERS


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Mean Design Parameters

S. No.

Design Parameter

Value

Unit

1

Cruising Speed

872

km/h

2

Length

55.0

m

3

Height

15.7

m

4

Fuselage Diameter

5.26

m

5

Wing Span

51.5

m

6

Chord

6.0

m

7

Aspect Ratio

8.6

(no unit)

8

Wing Area

348

m2

9

Wing Sweep

31.5°

degree

10

Cruising Altitude

10800

m

11

Service Ceiling

12000

m

12

Range

7200

km

13

Maximum Speed

940

km/h

14

Number of Engines

3

(no unit)

15

Maximum Thrust Capability

265

kN

16

Maximum Take Off Weight

272000

kg

17

Empty Weight

85000

kg

18

Wing Loading

710

kg/m2

19

Maximum Fuel Capacity

100000

litre


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WEIGHT ESTIMATION


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WEIGHT ESTIMATION FIRST WEIGHT ESTIMATION: The design take off gross weight W o 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. W0 = { Wcrew +Wpayload + Wfuel + Wempty } Wfuel - weight of the fuel load at beginning of the flight

W0 =

        −  −  +

1

0

  

0

- Fuel weight fraction

0

- Empty weight fraction

0

ESTIMATION OF

:

In the plot of W0 vs. We /W0 for the aircrafts shown in the comparative data sheet the values of We /W0 tend to cluster around a horizontal line at W e /W0


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Estimation of

:

The amount of fuel to carry out the mission depends critically on the efficiency of the propulsion device, the engine specific fuel consumption. It also depends on L/D ratio.

Normal mission profile for passenger aircraft Cruise 2

Loiter

3 Glide

Climb

Landing

Take off 0

4

1

The fuel weight ratio

 

5

can be obtained from the product of mission segment weight at the

0

end of the segment divided by the weight at the beginning of segment.

Suggested Fuel Fractions For Several Mission Phases Table 1 Airplane Type

Take Off

Climb

Descent

Landing

Business Jets

0.995

0.980

0.990

0.992

Transport

0.970

0.985

1.000

0.995

Military Trainers

0.990

0.980

0.990

0.995

Supersonic Cruise

0.995

0.92-0.87

0.985

0.992

From Table 1, we get the following values: For takeoff, segment 0- 1 historical data’s shows that,

 

1

= 0.97

0

For climb, segment 1-2 historical data shows that, 2

= 0.985

1


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For loiter, segment 3-4 ignoring the fuel consumption during descent we assume,



4

=1

3

For landing, segment 4-5 based on historical data we assume that,



5

= 0.995

4

The Brequet’s range equation is used to calculate the value of



3

. As we all know that

2

maximum range is covered during cruise we considering this equation,

R=

 ∞    ln

2 3

Initial Estimates of Lift/Drag Ratio (L/D) Table 2 Aircrafts

cruise

loiter

Homebuilt & single-engine

8 - 10

10 - 12

Business jets

10 – 12

12 - 14

Regional turboprops

11 – 13

14 – 16

Transport jets

13 – 15

14 - 18

Military trainers

8 – 10

10 - 14

Fighters

4 – 7

6 – 9

Military patrol, bombers & transports

13 – 15

14 – 18

Supersonic cruise

4-6

7 – 9

From the Table 2, L/D values of similar type of aircrafts we come to know that the approximate the value of L/D for our aircraft to be 15. So,

 

= 15


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Specific Fuel Consumption Table 3 Aircrafts

Cruise

Loiter

Business & transport jets

0.5 - 0.9

0.4 - 0.6

Military trainers

0.5 - 1.0

0.4 - 0.6

Fighters

0.6 - 1.4

0.6 - 0.8

Military patrol, bombers, transports, flying boats

0.5 – 0.9

0.4 - 0.6

Supersonic cruise

0.7 – 1.5

0.6 - 0.8

From the comparative data sheet,

∞

V = 872 km/hr R = 7200 km From Table 3, we found the values of c j as 0.6 hr -1

So now substituting these values in the Brequet’s range equation,

R=

 

2

 ∞    ln

2 3

= 1.39135

3

3

= 0.718726

2

Now using all the fuel fractions,

              5 0

=

1 0

x

5

2 1

x

3 2

x

= 0.68327

0


4 3

x

5 4

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If at end of the flight, the fuel tanks are not completely empty, making six percent of allowance for reserve and trapped fuel,

   −  0

= 1.06 1

 

0

5 0

= 0.33573

Wpayload + Wcrew= 0.256W=69,632 kg

(Or) We assume that the airplane occupies 280 passengers (with average weight of 180kg per passenger including baggage) and 12 crew (with average weight 100kg). Wpayload + Wcrew= 280(180) + 12(100) =51,600 kg

From the graph we get values of



as 0.475

0

By substituting these values in:

 −  −    − −   +

W0 =

1

We get W0 as,

W0 =

0

0

280(180) + 12(100)

1

0.33573

0.475

= 272626.4 kg

This is only the first estimation. Now by doing iterations, we can get a fairly accurate value of the Maximum Take Off Weight (W0).


P a g e | 47

ITERATION PROCESS (W0) For the iteration process, we use the given formula, We



= 1.02 ×

0

− 0

0.06

FIRST:

We



−

= 1.02 × 272626.4

0

We



0.06

= 0.481355676

0

W0 = 282099.285 SECOND:

−

We



= 1.02 × 282099.285

0

We



0

0.06

= 0.4803702

W0 = 280587.572 THIRD:

−

We



= 1.02 × 280587.572

0

We



0

0.06

= 0.4805251

W0 = 280824.1 FOURTH:

We



−

= 1.02 × 280824.1

0

We



0

= 0.4805008

W0 = 280786.977


0.06

P a g e | 48

FIFTH:

We



−

= 1.02 × 280786.977

0

We



0.06

= 0.4805046

0

W0 = 280792.801 SIXTH:

We



−

= 1.02 × 280792.801

0

We



0

0.06

=0.480504

W0 = 280791.887 SEVENTH:

We



−

= 1.02 × 280791.887

0

We



0

0.06

=0.480504

W0 = 280792

After doing seven iterations, we can see that the value of

We Wo

starts to converge on 0.480504.

So we can take the value W 0 = 280792 as the final estimate of the W 0.

Max Take Off Weight (W 0) = 280,792 kg.

We know that,

 

= 0.33573

0

So, substituting the value of W 0, we get the first estimation value of W f , Wf = 0.3375 × 280792 = 94767.3 kg Weight of the Fuel (Wf )= 94,767.3 kg.


P a g e | 49

POWERPLANT SELECTION


P a g e | 50

POWERPLANT SELECTION •

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.

Engine Name

Dry weight (kg)

Max Thrust (kN)

Thrust to Weight ratio

Rolls Royce Trent 772B-60

4788

320

6.8:1

Pratt & Whitney PW4000-100

4270

310

7.4:1

CFM International CFM56-5C4

3990

151

3.9:1

General Electric CF6-50

4104

240

6:1

Pratt & Whitney JT9D-7R4H1

4030

250

6.3:1

Bypass Ratio

5

5

6.4

4.4

4.8

The preferable choice of engine, from those listed above would be the Rolls Royce Trent 772B-60 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.


P a g e | 51

Details about the selected engine: Rolls Royce Trent 772B-60

Since its launch with Cathay Pacific in 1995, the Trent 700 has built up the greatest service experience on the A330. As the only engine specifically designed for the A330 it delivers the greatest performance over the widest range of operational and environmental conditions.

The Trent 700 marked the birth of a new family of engines; it incorporates revolutionary advances in wide chord hollow titanium fan blade technology, Full Authority Digital Engine Control (FADEC) and 3-D aerodynamics, whilst maintaining the three-shaft design characteristics of low weight, high strength and exceptional performance retention. As part of a successful and expanding family, the Trent 700 has benefited through continuous improvement as technology has flowed from later generation family members. Incorporation of the HP module from the Trent 800 enabled the Trent 700 to deliver the best performance of any engine on the A330 whilst delivering long on-wing life and low maintenance costs. Improvements in the LP turbine and other technology flowed from the Trent 1000 will ensure the Trent 700 delivers the lowest fuel burn on the A330. Having been selected by over 40 operators of the A330, the Trent 700 is the most popular engine on the aircraft. This is apparent in China where 100 per cent of A330 operators have selected the Trent 700 and in


P a g e | 52

the Middle East it has 80 per cent market share. The engine’s unrivalled high and hot performance gives Trent 700 customers a distinct operating advantage. All this contributes to a leading market share of around 50 per cent. In addition to its capability the Trent 700 has superb environmental credentials as the cleanest and quietest engine on the A330.

As a complete package the Trent 700 provides any customer with the greatest flexibility.

Technical Details Engine Thrust Bypass ratio Inlet mass flow Fan diameter Length Stages Certification EIS

: Trent 772B-60 : 71,100lb : 5.0 : 2030lb/sec : 97.4in : 154in : Fan, 8 IPC, 6 HPC, 1 HPT, 1 IPT, 4 LPT : Jan 1994 : Mar 1995


P a g e | 53

FUEL WEIGHT VALIDATION


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FUEL WEIGHT VALIDATION The choice of a suitable engine, having been made, it is now possible to estimate the amount of fuel required for a flight at the given cruising speed for the given range.

Wfuel

=

∗∗∗∗   .

The factor of 1.2 is provided for reserve fuel.

Thrust at altitude is calculated using the relation:

T 





T 0 * 

1.2

 a lt 

 0

Altitude = 10800m = 35433ft

  =

0

= 0.3715 / 1.225 = 0.303

Cruise velocity = 872km/hr = 242.2m/s To = 320kN

 

= 320×0.303 1.2 = 76.363kN = 7784.2kg

SFC = 0.4hr -1 (at medium thrust setting) Number of engines = 3

CALCULATION:

Wfuel =

3×7784.2×7200×0.4×1.2 872 Wfuel = 92,553.42 kg


P a g e | 55

WING SELECTION


P a g e | 56

WING SELECTION 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 critical Mach number can well be delayed by the use of an appropriate Sweepback angle to the wing structure. The natural choice of the standard series is the 65 series which is designed specifically for use in high-speeds.

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 (λ) and the Sweepback angle at quarter chord (Λqc).

•

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. C root 

2 S b  (1   )

The tip chord is given by,

Ctip    C root POSITION OF WING The location of the wing in the fuselage (along the vertical axis) is very important. Each configuration (Low, High and mid) has its own advantages but in this design, the Lowwing offers significant advantages such as

 Uninterrupted Passenger’s cabin.  Placement of Landing gear in the wing structure itself.  Location of the engine on a low-wing makes Engine-overhaul easier.


P a g e | 57

 Permits usage of the Wing carry through box which alone can admit the amount of fuel that we require to carry.

 Landing gear usually becomes high in such wing configurations and therefore, provides greater ground clearance ad reduces the amount of fuselage upsweep that is to be provided.



Low wing affects the flow over the horizontal tail to mi nimum extent.

 The low-wing requires that some-amount of dihedral angle is provided for lateral stability. As of now, the dihedral angle is assumed to be 5 degrees, but it may be subject to change in the stability analysis.

WING PLANFORM

WING SETTING ANGLE The wing has to be set at angle to the fuselage center line such that during cruise, the fuselage is in a level condition (parallel to the direction of the velocity vector). This requires that the wing setting angle correspond to the angle which will produce the desired C L for cruise. The C L that will be obtainable from an airfoil section (for a given angle of attack) is given by: CL =0.9 x C l x cosΛ.

    

Cl=

2×

× 2×


P a g e | 58

DESIGN CALCULATION

(First Estimation) Croot Calculation: C root

Croot =

2 S



b

 (1   )

2 × 395.48 58.32× (1+0.25)

= 10.85m

Ctip Calculation:

Ctip



  C root

Ctip = 0.25 x 10.85 = 2.7m

Cmean Calculation: Cm =

2 3

Cm =

× 2 3



×

λ λλ

(1+ + 2 ) (1+ )

×10.85×1.05 = 7.6m

Coefficient of Lift Calculation:

Section Lift Coefficient:

    

Cl= Cl =

2×

× 2×

2×710×9.8

0.3715×242.2 2

= 0.638

Wing Lift Coefficient: CL =0.9 x C l x cosΛ CL =0.9 x 0.638 x cos35 o = 0.47


P a g e | 59

It is to be found graphically the following parameters were estimated for the aircraft designed.

DESIGN CHARACTERISTICS 2

VALUES 710

W/S (kg/m ) 2

Wing area S (m )

395.48

Aspect Ratio

8.6

Span b (m)

58.32

Taper ratio ( )

0.25

Root Chord (m)

10.85

Mean Chord (m)

7.6

Tip chord (m)

2.7

Lift coefficient (CL)

0.47

Sweepback Angle( )

35°

 ∆


P a g e | 60

AIRFOIL SELECTION


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AIRFOIL SELECTION 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. The following are the airfoil geometry and definition: Chord line: It is the straight line

connecting leading edge (LE) and trailing edge (TE). Chord (c): It is the length of

chord line. Thickness (t): measured perpendicular to chord line as a % of it (subsonic typically 12%). Camber (d): It is the curvature of section, perpendicular distance of section mid-points from

chord line as a % of it (sub sonically typically 3%). Angle of attack (α): It is the angular difference between chord line and airflow direction.

The following are airfoil categories: Early it was based on trial & error. NACA 4 digit is introduced during 1930’s.

NACA 5-digit is aimed at pushing position of max camber forwards for increased C Lmax. NACA 6-digit is designed for lower drag by increasing region of laminar flow. Modern it is mainly based upon need for improved aerodynamic characteristics at speeds just below speed of sound. NACA 4 Digit st

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

– 2 digit (x10): location of maximum camber (as % of chord from leading edge (LE)). rd

th

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


P a g e | 62

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 st

– 1 digit: identifies series type. nd

– 2 digit (x10): location of minimum pressure (as % of chord from leading edge (LE)). – 3rd digit: indicates acceptable range of C L above/below design value for satisfactory low drag performance (as tenths of C L). – 4th digit (x0.1): design C L. – 5th & 6th digits: maximum section thickness (%c)

The airfoil that is to be used is now selected. As indicated earlier during the calculation of the lift coefficient value, it becomes necessary to use high speed airfoils, i.e., the 6x series, which have been designed to suit high subsonic cruise Mach numbers.

t/c Calculation:

   ∆ −∆ −  ∆   =

Taking Where,



#

0.3

1 3

1

[1

5+ 5+(

= 1.05 - 0.25 C L (cruise)

M = Drag Divergence Cruise Mach Number = 0.85

∆

= Sweep Back Angle = 35 ° at Quarter Chord

CL (cruise) = 0.47



Substituting the values in the above equation, we get, = 0.12


2

# )2

3.5

2 ]3

P a g e | 63

NACA 6-series Airfoils having t/c ratio of 0.12

Name

Thickness Camber Lift Coeff. Lift-to-Drag Stall Angle (%) (%) (CL) (L/D) (deg)

TE Angle (deg)

LE Radius (%)

NACA 63-212

12

1.1

1.035

36.2

5.5

11.7

1.5

NACA 63-412

12

2.2

1.159

44.3

5.5

11.6

1.5

NACA 64(1)-112

12

0.6

0.936

32.1

4.5

9

1.5

NACA 64(1)-212

12

1.1

1.008

37.5

4.5

12.3

1.5

NACA 65(1)-212

12

1.1

0.971

31.7

3.5

10.8

1.3

NACA 65(1)-412

12

2.2

1.107

44.8

4

10.8

1.3

NACA 66(1)-212

12

1.1

0.957

32.5

-0.5

14

1.3

From the above list of airfoils, the one chosen is the 65(1)-412 airfoil which has the suitable lift coefficient for the current design. In order to obtain better span-wise distribution of lift and to have better stalling characteristics (the root should stall before the tip so that the pilot may realize and avoid a stall by sensing the vibrations on his control stick), it is usually necessary to provide a lower t/c to the tip section and a higher t/c to the root section. Hence, Section used at the mean aerodynamic chord

- 65(1)-412

Section used at the tip

- 65-410

Section used at the root

- 65(2)-415



CHORD

AIRFOIL

( )max

ROOT

65(2)-415

1.238

MEAN

65(1)-412

1.107

TIP

65-410

1.015


P a g e | 64

Airfoil Geometry

NACA 65-410 (tip)

NACA 65(1)-412 (mean)

NACA 65(2)-415 (root)


P a g e | 65

Angle of Attack (vs) Lift Coefficient of NACA 65-410

Angle of Attack (vs) Lift Coefficient of NACA 65(2)-415


P a g e | 66

Performance curves for the chosen airfoil NACA 65(1)-412

Angle of Attack (α) vs Coefficient of Lift (CL)

Angle of Attack (α) vs Coefficient of Drag (CD)


P a g e | 67



Angle of Attack (α) vs Lift-to-Drag ratio ( )

CALCULATIONS:

 

Available

=





max

 = 0.9 ×

1.238 3

max

+

1.107 3

+

1.015 3

= 1.12

 = 0.9 × 1.12 = 1.008

Flaps Selection For the current design, double slotted flap is selected. for different configurations is given in the table below:

∆

of the double slotted flap

FLAPS

TAKE OFF

Double slotted flap

20

40

1.825

2.5

1.5

2.05

∆ ∆ ∆ /

o


LANDING o

P a g e | 68

∆         ∆       =

max

-

max

max

=

max

(Take Off) = 1.008+1.5 = 2.508

max

(Landing) = 1.008+2.05 = 3.058

+

max

= 0.25×

= 60.55 m/s

We Have,

W/S=700.722 kg/m2 From this, S = 400.72



2

b = 58.32 m (from table)

DESIGN CALCULATION

(Second Estimation)

Croot Calculation: C root 

Croot =

2 S b  (1   ) 2 × 400.72

58.32× (1+0.25)

= 11m

Ctip Calculation:

Ctip



  C root

Ctip = 0.25 x 11 = 2.75m


P a g e | 69

Cmean Calculation: Cm =

2 3

×

Cm =

2



×

λ λλ

(1+ + 2 ) (1+ )

×11×1.05 = 7.7m

3

Coefficient of Lift Calculation:

Section Lift Coefficient:

    

Cl= Cl =

2×

× 2×

2×700.72×9.8

0.3715×242.2 2

= 0.63022

Wing Lift Coefficient: CL =0.9 x C l x cosΛ CL =0.9 x 0.63022 x cos35 o = 0.4646

DESIGN CHARACTERISTICS 2

VALUES 700.72

W/S (kg/m ) 2

Wing area S (m )

400.72

Aspect Ratio

8.6

Span b (m)

58.32

Taper ratio ( )

0.25

Root Chord (m)

11

Tip chord (m)

2.75

Mean chord (m)

7.7

 ∆

Sweepback Angle( )



Cruise Lift Coefficient (

35° )


0.63022

P a g e | 70

LIFT ESTIMATION


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LIFT ESTIMATION LIFT: Component of aerodynamic force generated on aircraft perpendicular to flight direction.

Lift Coefficient (C L) •

Amount of lift generated depends on: – Planform area (S), air density ( ), flight speed (V), lift coefficient (CL) Lift

•



(

1 2

2  V ) SCL



qSCL

CL is a measure of lifting effectiveness and mainly depends upon: – Section shape, planform geometry, angle of attack ( ), compressibility effects (Mach number), viscous effects (Reynolds’ number).

Generation of Lift

•

Aerodynamic force arises from two natural sources: – Variable pressure distribution. – 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.


P a g e | 72

Pressure variations with angle of attack

–

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

–

Positive lift at  = 0o.

–

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.


P a g e | 73

Lift Curves of Cambered and Symmetrical airfoils

CALCULATION: General Lift equation is given by, Lift



(

1 2

 V

2

) SCL



qSCL

Lift at Cruise



= 0.3715 (at the cruising altitude of 10800m)

V = 242.2 m/s S = 400.72 kg/m 2 CL(cruise) = 0.63022 (from the wing and airfoil estimation)

Substituting all these values in the general lift equation, L(cruise) =

1 2

× 0.3715 × 242.22 × 400.72 × 0.63022 Lift at cruise =


P a g e | 74

Lift at Take-Off



= 1.225 (at sea altitude)

V = 0.7 x V lo = 0.7 x 1.2 x V stall S = 400.72 kg/m 2 CL(take-off) = 2.508 (flaps extended and kept at the take-off position of 20 o)

Substituting all these values in the general lift equation, 1

L(take-off) = × 1.225 × (0.7 × 1.2 × 66.86)2 × 400.72 × 2.508 2

Lift at take-off = Lift at Landing



  .


V = 0.7 x V t = 0.7 x 1.3 x V stall S = 400.72 kg/m 2 CL(landing) = 3.058 (flaps extended and kept at the landing position of 40 o)

Substituting all these values in the general lift equation, 1

L(landing) = × 1.225 × (0.7 × 1.3 × 60.55)2 × 400.72 × 3.058 2

Lift at landing =

 


.

P a g e | 75

DRAG ESTIMATION


P a g e | 76

DRAG ESTIMATION 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.

Drag Coefficient (CD)

 Amount of drag generated depends on: o Planform area (S), air density ( ), flight speed (V), drag coefficient (C D)  CD is a measure of aerodynamic efficiency and mainly depends upon: o Section shape, planform geometry, angle of attack ( ), compressibility effects (Mach number), viscous effects (Reynolds’ number).

Drag Components

 Skin Friction: o o

Due to shear stresses produced in boundary layer . Significantly more for turbulent than laminar types of boundary layers.

 Form (Pressure) Drag o

o

Due to static pressure distribution around body - component resolved in direction of motion. Sometimes considered separately as forebody and rear ( base) drag components.


P a g e | 77

 Wave Drag o o

o

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. Often decomposed into portions related to: Lift. Thickness or Volume.  


P a g e | 78

Typical streamlining effect

Lift induced (or) trailing vortex drag

The lift induced drag is the component which has to be included to account for the 3-D nature of the flow (finite span) and generation of wing lift.


P a g e | 79

CALCULATION: Generally for jet aircrafts, it is given that CD,0 = 0.0030 e = 0.8

The general drag equation is given by,

   ∅  1 = 2

2

2

,0

+

For calculating Ø, we use the formula,

   16

Ø=

1+

16

2

2

Where h = height above ground, b = wing span. h = 2m b = 58.32m

Ø=





2 2 16× 58.32 = 0.2314 2 1 + (16× )2 58.32

Drag at Cruise



= 0.3715 (at the cruising altitude of 10800m)

V = 242.2 m/s S = 400.72 kg/m 2 CL(cruise) = 0.63022 (from the wing and airfoil estimation)

Substituting all these values in the general drag equation, 1

2

D(cruise) = × 0.3715 × (242.2) × 400.72 (0.0030 + 2

Drag at cruise = 31674.846 N


0.2314×0.63022 2 3.14×8.6×0.8

)

P a g e | 80

Drag at Take-off




V = 0.7 x V lo = 0.7 x 1.2 x V stall S = 400.72 kg/m 2 CL(take-off) = 2.508 (flaps extended and kept at the take-off position of 20 o)


2

D = × 1.225 × (0.7 × 1.2 × 66.86) × 400.72 (0.0030 + 2

0.2314×2.5082 3.14×8.6×0.8

)

Drag at take-off = 54482.6 N

Drag at Landing




V = 0.7 x V t = 0.7 x 1.3 x V stall S = 400.72 kg/m 2 CL(landing) = 3.058 (flaps extended and kept at the landing position of 40 o)


2

D = × 1.225 × (0.7 × 1.3 × 60.55) × 400.72 (0.0030 + 2

Drag at landing

= 76876.7 N


0.2314×3.0582 3.14×8.6×0.8

)

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LANDING GEAR ARRANGEMENT


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LANDING GEAR SELECTION In aviation, the undercarriage or landing gear is the structure (usually wheels) that supports an aircraft and allows it to move across the surface of the earth when it is not in flying. So more importance is to be given as it carries the entire l oad on the ground.

OVERVIEW The design and positioning of the landing gear are determined by the unique characteristics associated with each aircraft, i.e., geometry, weight, and mission requirements. Given the weight and cg range of the aircraft, suitable configurations are identified and reviewed to determine how well they match the airframe structure, flotation, and operational requirements. The essential features, e.g., the number and size of tires and wheels, brakes, and shock absorption mechanism, must be selected in accordance with industry and federal standards discussed in the following chapters before an aircraft design progresses past the concept formulation phase, after which it is often very difficult and expensive to change the design. Three examples of significant changes made after the initial design include the DC-10-30, which added the third main gear to the fuselage, the Airbus A340, where the main gear center bogie increased from two to four wheels in the -400 series, and the Airbus A-300, where the wheels were spread further apart on the bogie to meet LaGuardia Airport flotation limits for US operators. The purpose of Landing Gears is to move the aircraft on ground. After take-off the landing gear is retracted, before landing it is extended and locked into position. Liebherr provides system architecture for gear actuation control, steering control, wheel and brake integration and position and status control, as well as system integration, series production and of course product support. Liebherr acquired knowledge and experience based on the realization of different landing gear programs. The integration of various technologies and use of new material for individual landing gear concepts lead to competitive products: 

Landing Gear Systems



Nose Landing Gear Subsystem



Main Landing Gear Subsystem



Brake and Brake Control Subsystem



Research and Development Technology


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TYPES OF GEAR ARRANGEMENTS Wheeled undercarriage comes in two types: conventional or tail dragger undercarriage, where there are two main wheels towards the front of the aircraft and a single, much smaller, wheel or skid at rear; tricycle undercarriage where there are two main wheels under the wings and a third smaller wheel in the nose. Most modern aircraft have tricycle undercarriage. Sometimes a small tail wheel or skid is added to aircraft with tricycle undercarriage arrangements.

RETRACTABLE GEAR To decrease drag in flight some undercarriages retract into the wings and/or fuselage with wheels flush against or concealed behind doors, this is called retractable gear. It was in late 1920s and 1930s that such retractable landing gear became common. This type of gear arrangement increased the performance of aircraft by reducing the drag.

LARGE AIRCRAFT As the size of aircraft grows larger, they employ more wheels to with the increasing weight. The airbus A340-500/-600 has an additional four wheel undercarriage bogie on the fuselage centerline. The Boeing 747 has five sets of wheels, a nose-wheel and four sets of four wheel bogies. A set is located under each wing, and two inner sets located in the fuselage, a little rearward of outer bogies.

MAIN FUNCTIONS •

Carry aircraft max gross weight to take off runway

•

Withstand braking during aborted take off

•

Retract into compact landing gear bay


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•

Damp touchdown at maximum weight.

Total LG weight typically 3% of MTOW for commercial airliners.

STEERING The steering mechanism used on the ground with wheeled landing gear varies by aircraft, but there are several types of steering.

 RUDDER STEERING  DIRECT STEERING 

TILLER STEERING

Configuration Selection The nose wheel tricycle undercarriage has long been the preferred configuration for passenger transports. It leads to a nearly level fuselage and consequently the cabin floor when the aircraft is on the ground. The most attractive feature of this type of undercarriages is the improved stability during braking and ground maneuvers. Under normal landing attitude, the relative location of the main assembly to the aircraft cg produces a nose-down pitching moment upon touchdown. This moment helps to reduce the angle of attack of the aircraft and thus the lift generated by the wing. In addition, the braking forces, which act behind the aircraft cg, have a stabilizing effect and thus enable the pilot to make full use of the brakes. These factors all contribute to a shorter landing field length requirement. The primary drawback of the nose wheel tricycle configuration is the restriction placed upon the location where the main landing gear can be attached. With the steady increase in the aircraft takeoff weight, the number of main assembly struts has grown from two to four to accommodate the number of tires required to distribute the weight over a greater area.

Landing Gear Disposition: The positioning of the landing gear is based primarily on stability considerations during taxiing, liftoff and touchdown, i.e., the aircraft should be in no danger of turning over on its side once it is on the ground. Compliance with this requirement can be determined by examining the takeoff/landing performance characteristics and the relationships between the locations of the landing gear and the aircraft cg.


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Stability at Touchdown and During Taxiing Static stability of an aircraft at touchdown and during taxiing can be determined by examining the location of the applied forces and the triangle formed by connecting the attachment locations of the nose and main assemblies. Whenever the resultant of air and mass forces intersects the ground at a point outside this triangle, the ground will not be able to exert a reaction force which prevents the aircraft from falling over. As a result, the aircraft will cant over about the side of the triangle that is closest to the resultant force/ground intersect.

Braking and Steering Qualities The nose assembly is located as far forward as possible to maximize the flotation and stability characteristics of the aircraft. However, a proper balance in terms of load distribution between the nose and main assembly must be maintained. When the load on the nose wheel is less than about eight percent of the maximum takeoff weight (MTOW),controllability on the ground will become marginal, particularly in cross-wind 21 conditions. This value also allows for fuselage length increase with aircraft growth. On the other hand, when the static load on the nose wheel exceeds about 15 percent of the MTOW, braking quality will suffer, the dynamic braking load on the nose assembly may become excessive, and a greater effort may be required for steering.

Ground Operation Characteristics: Besides ground stability and controllability considerations, the high costs associated with airside infrastructure improvements, e.g., runway and taxiway extensions and pavement reinforcements, have made airfield compatibility issues one of the primary considerations in the design of the landing gear. In particular, the aircraft must be able to maneuver within a pre-defined space as it taxies between the runway and passenger terminal. For large aircraft, this requirement effectively places an upper limit on the dimension of the wheelbase and track.


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LANDING GEAR TYPES During landing and take-off, the undercarriage supports the total weight of the airplane. Undercarriage is of three types

 Bicycle type  Tricycle type  Tricycle tail wheel type

From the above list of landing gear types, the tricycle type is chosen which is the most suitable configuration for the current design.


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FUSELAGE DESIGN


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FUSELAGE DESIGN INTRODUCTION The fuselage is an aircraft's main body section that holds crew and passengers or cargo. In single-engine aircraft it will usually contain an engine, although in some amphibious aircraft the single engine is mounted on a pylon attached to the fuselage which in turn is used as a floating hull. The fuselage also serves to position control and stabilization surfaces in specific relationships to lifting surfaces, required for aircraft stability and manoeuvrability. Common practice to modularise layout:

 Crew compartment, power plant system, payload configuration, fuel volume, landing 

gear stowage, wing carry-through structure, empennage, etc. Or simply into front, centre and rear fuselage section designs.

Functions of fuselage:

 Provision of volume for payload.  Provide overall structural integrity.  Possible mounting of landing gear and power plant. Once fundamental configuration is established, fuselage layout proceeds almost independently of other design aspects.

PRIMARY CONSIDERATIONS Most of the fuselage volume is occupied by the payload, except for:

   

Single and two-seat light aircraft. Trainer and light strike aircraft. Combat aircraft with weapons carried on outer fuselage & wing. High performance combat aircraft.

Payload includes:

      

Passengers and associated baggage. Freight. Internal weapons (guns, free-fall bombs, bay-housed guided weapons). Crew (significant for anti-sub and early-warning aircraft). Avionics equipment. Flight test instrumentation (experimental aircraft). Fuel (often interchangeable with other payload items on a mass basis).


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Pressurisation:

 If required, has a major impact upon overall shape.  Overall effect depends on level of pressurisation required. Low Differential Pressurisation:

 Defined as no greater than 0.27 bar (4 psi).  Mainly applicable to fighters where crew are also equipped with pressure suits.  Cockpit pressurisation primarily provides survivable environment in case of suit failure at high altitude.

 Also used on some general aviation aircraft to improve passenger comfort at moderate altitude.

 Pressure compartment has to avoid use of flat surfaces. Normal (High) Differential Pressurisation:

 Usual requirement is for effective altitude to be no more than 2.44 km (8000 ft) ISA for passenger transports.

 Implied pressure differentials are: o

0.37 bar (5.5 psi) for aircraft at 7.6 km (25,000 ft).

o


o


 High pressure differential required across most of fuselage for passenger transports so often over-riding fuselage structural design requirement.

 Particular need to base outer shell cross-section on circular arcs to avoid significant mass penalties.

 Pure circular sections best structurally but “double-bubbles” sometimes give best compromise with internal layout.

Circular Section Examples:


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Fuselage Aerodynamics:  Aim is to achieve reasonably streamlined form together with minimum surface area to meet required internal volume.

   

Both drag and mass heavily influenced by surface area. Require absence of steps and minimum number of excrescences. Fundamental differences between subsonic and supersonic applications. Concerned with: cross-section shape, nose shape & length, tail shape/length, overall length.

Cross-Section Shape – Subsonic Aircraft:

 Not too critical aerodynamically, but should: o

avoid sharp corners

o

provide fairings for protuberances

 Constant cross-section preferable for optimized volume utilization and ease of manufacture.

Nose Shape:

 Should not be unduly “bluff”.  Local changes in cross-section needed to accommodate windscreen panels.  Windscreen angle involves compromise between aerodynamics, bird-strike, reflection and visibility requirements.

 Windscreen panel sizes should be less than 0.5 m 2 each.  Starting point for front fuselage layout is often satisfactory position for pilot’s eye.  Reasonable nose length is about: o

1.1 to 2.0 x fuselage diameter (subsonic).

o

4 x fuselage diameter (supersonic).

Tail Shape:

 Smooth change in section required, from maximum section area to ideally zero.  Minimisation of base area especially important for transonic/supersonic aircraft.  Important parameter for determining tail upsweep angle is ground clearance required for take-off and landing rotation.

 Typically 12o to 15o.


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 Typical tail section lengths are: o

2.5 to 3.0 x diameter (subsonic)

o

6 to 7 x diameter (supersonic)

Centre Fuselage & Overall Length - Subsonic Aircraft:

 Theoretically minimum drag for streamlined body with fineness ratio (length/diameter) of 3.

 In reality, typical value is around 10, due to: o

Need to utilise internal volume efficiently.

o

Requirement for sufficiently large moment arm for stability/control purposes.

o

Suitable placement of overall CG.

Wing Location - Aerodynamics Considerations:  Mid-wing position gives lowest interference drag, especially well for supersonic aircraft.

 Top-mounted wing minimises trailing vortex drag, especially good for low-speed aircraft.

 Low wing gives improved landing gear stowage & more usable flap area. From the above given locations of wings, the one chosen is the Low wing configuration which gives improved landing gear stowage & more usable fl ap area.

Empennage Layout Vertical Surface:  Single, central fin most common arrangement, positioned as far aft as possible. Horizontal Surface:  Efficiency affected by wing downwash, thus vertical location relative to wing important.

 Usually mounted higher than wing except on high wing design or with small moment arm – low tail can give ground clearance problems.


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Avionics & APU:

 Including navigation, communications and flight control/management equipment.

 Provision necessary for adequate volume in correct location with ease of access.

 Location of radar, aerials, etc also important Sensors often have to face forward/down in aircraft nose. o Long range search & early warning scanners sometimes located on fuselage.  Auxiliary power unit (APU) commonly located at extreme rear of fuselage on transport aircraft. o

TYPICAL FLIGHT DECK LAYOUT:


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SEATING ARRANGEMENTS:

 Typical split of classes: o

8% first, 13% business, 79% economy

BAGGAGE AND FREIGHT:

It is to be found graphically the following parameters were estimated for the aircraft designed.

DESIGN CHARACTERISTICS

VALUES

Overall Length (m)

55.0

Fuselage Width (m)

5.26

Cabin Width (m)

5.0

Length/Width

10.456

SEATING ARRANGEMENT:


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PERFORMANCE CHARACTERISTICS


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PERFORMANCE CHARACTERISTICS TAKE-OFF PERFORMANCE: •

•

Distance from rest to clearance of obstacle in flight path and usually considered in two parts: – Ground roll - rest to lift-off (S LO) – Airborne distance - lift-off to specified height (35 ft FAR, 50 ft others). The aircraft will accelerate up to lift-off speed (V lo = about 1.2 x V stall) when it will then be rotated.

•

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

   2

1.44

=

,

•

This shows its sensitivity to W (W 2) and  (1/ 2 since T also varies with ).

•

Slo may be reduced by increasing T, S or C l,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.

   −  − =

+

,

•

2

1.44

The bracketed term will vary with speed but an approximation may be made by using an instantaneous value for when V = 0.7 x V lo.

•

In the above equation:

   ∅  1 = 2

•

2

2

+

,0

Where  accounts for drag reduction when in ground effect:

   16

Ø=

1+

2

16

•

Where h = height above ground, b = wing span.

•

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


2

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CALCULATION:

Ø=





2 2 58.32 2 1 + (16× )2 58.32

16×

= 0.2314

1

0.2314×2.5082

2

3.14×8.6×0.8

D = × 1.225 × (0.7 × 1.2 × 66.86)2 × 400.72 (0.0030 +

Slo =

1.44×(280792×9.81)2



−

) = 54482.6 N

−



9.8×1.225×400.72×2.508×{(3×320000 ) [54482.6+.02 280792×9.81 1941627 .7 ]}

Take-off runway distance = 1018.38m

CLIMBING •

Consider aircraft in a steady unaccelerated climb with vertical climb speed of V c.

•

Force balance gives:

   −   =

=

=

R/Cmax =

+

(

)×

 − 

(

)

=

−

3×320000 ×60.55

(40382.15×60.55)

280792×9.81

R/Cmax = 20.2 m/s


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MANOEUVRES / TURNING FLIGHT An aircraft is capable of performing many different types of turns and manoeuvres. •

Three of the more common turns will be considered here in simplistic terms: – Constant altitude banked turn. – Vertical pull-up manoeuvre. – Vertical pull-down manoeuvre.

•

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 manoeuvre.

CONSTANT ALTITUDE BANKED TURN •

In steady condition: – T = D

•

Force balance gives: – W = mg = Lcos  2

– Fr = mV /r = Lsin •

 tan  = V2 /(Rg)

•

So for given speed and turn radius there is only one correct bank angle for a coordinate (no sideslip) turn.

•

Maneuverability equations simplified through use of normal load factor (n) = L/W.

•

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.

•

V2 / (Rg) = tan  = (sec2  - 1) = (n2 - 1)

•

 R = V2 / (g (n2 - 1)) and  = V/R = (g (n2 - 1)) / V

CALCULATION: W = Lcos  Let  = 300

  −    n=

R=

= 1.1547

V2

g n2

=

1

= 10357.16 m

= 0.0234 rad/s


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

Force balance gives:

    1    tan 1    L   D  1   tan 1   15 



3.814

o

LANDING PERFORMANCE APPROACH & LANDING •

Consists of three phases: – Airborne approach at constant glide angle (around 3 o) and constant speed. – Flare - transitional manoeuvre with airspeed reduced from about 1.3 V stall

down to touch-down speed. – Ground roll - from touch-down to rest.


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•

Ground roll landing distance (s 3 or sl) estimated from:

    − 1.69 +

=

,

•

2

Where Vav may be taken as 0.7 x touch-down speed (V t or V 2) and V t 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 (T r) are applied, use:

     − 1.69 2 + +

=

,

CALCULATION: 1

0.2314×3.0582

2

3.14×8.6×0.8

D = × 1.225 × (0.7 × 1.3 × 60.55)2 × 400.72 (0.0030 +

Sl =



1.69×(280792×9.81)2





) = 76876.7 N

−



9.8×1.225×400.72×3.058×{ 3×320000 +[76876.7+0.4 280792 ×9.81 2278744 .7 ]}

Landing Runway distance = 710.3 m


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3-VIEW DIAGRAM


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3-VIEW DIAGRAM

5 8 .3 2 m

55m

1 5 .7 m


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CONCLUSION


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CONCLUSION Design is a fine blend of science, creativity, 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 looks for the best product design. This involves the strong fundamentals in science and mathematics and their skilful applications, which is a tough job endowed upon the designer. We have enough hard work for this design project. A design never gets completed in a flutter sense but it is one step further towards ideal system. But during the design of this aircraft, we learnt a lot about aeronautics and its implications when applied to an aircraft design.


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BIBLIOGRAPHY


Aircraft Design Project - 280 Seat Transport

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