40504526-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

P a g e | 1

Aircraft Design Project - 1

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

P a g e | 2


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. 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 the necessity arose. I

thank one and all who have directly or indirectly helped me in making this design

project a great success.

P a g e | 3


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

P a g e | 4


ABBREVIATION

A.R. - Aspect Ratio

B - Wing Span (m)

C - Chord of the Airfoil (m)

C root - Chord at Root (m)

C tip - Chord at Tip (m)

C - Mean Aerodynamic Chord (m)

Cd - Drag Co-efficient

Cd,0 - Zero Lift Drag Co-efficient

Cp - Specific fuel consumption (lbs/hp/hr)

CL - Lift Co-efficient

D - Drag (N)

E - Endurance (hr)

E - Oswald efficiency

L - Lift (N)

(L/D)loiter - Lift-to-drag ratio at loiter

(L/D)cruise - Lift-to-drag ratio at cruise

M - Mach number of aircraft

Mff - Mission fuel fraction

R - Range (km)

Re - Reynolds Number

S - Wing Area (m²)

Sref - Reference surface area

Swet - Wetted surface area

Sa - Approach distance (m)

Sf - Flare Distance (m)

Sfr - Free roll Distance (m)

Sg - Ground roll Distance (m)

T - Thrust (N)

Tcruise - Thrust at cruise (N)

Ttake-off - Thrust at take-off (N)

(T/W)loiter - Thrust-to-weight ratio at loiter

(T/W)cruise - Thrust-to-weight ratio at cruise

(T/W)take-off - Thrust-to-weight ratio at take-off

Vcruise - Velocity at cruise (m/s)

Vstall - Velocity at stall (m/s)

Vt - Velocity at touch down (m/s)

Wcrew - Crew weight (kg)

Wempty - Empty weight of aircraft (kg)

Wfuel - Weight of fuel (kg)

Wpayload - Payload of aircraft (kg)

W0 - Overall weight of aircraft (kg)

W/S - Wing loading (kg/m²)

- Density of air (kg/m³)

- Dynamic viscosity (Ns/m²)

- Tapered ratio

R/C - Rate of Climb

P a g e | 5


AIM OF THE PROJECT

P a g e | 6


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.

P a g e | 7


ABSTRACT

P a g e | 8


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.

P a g e | 9


INTRODUCTION

P a g e | 10


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 one

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 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. 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.

P a g e | 11


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

P a g e | 12


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

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

P a g e | 13


SEVEN INTELLECTUAL POINTS

FOR CONCEPTUAL DESIGN

P a g e | 14


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

P a g e | 15


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

P a g e | 16


COMPARATIVE DATASHEET

P a g e | 17


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 m2

260 219 361.6 439.4 443

Wing Sweep degree 28° 28° 30° 31.1° 31.9°

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

MTO Weight x103

Kg 171.7 164 233 372 268

Empty Weight x103

Kg 90.9 83.1 124.5 170.9 115.7

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

P a g e | 18



Boeing Aircrafts


Name (no unit) 707-320B 757-200 767-200 777-200 787-9


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


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


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



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


Max Speed Km/h 1,010 935 913 950 945


Max thrust capability kN 320.4 193 222 330 320

Design Weights

MTO Weight x103

Kg 151.32 115.68 142.88 247.2 248

Empty Weight x103

Kg 66.4 57.18 81.23 134.8 115

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

P a g e | 19



Other Aircrafts


Name (no unit) Lockheed

L-1011-200

Ilyushin

IL-96-300

Tupolev

Tu-204-100

Douglas

DC-8-63CF

Tupolev

Tu-114


Aircraft Dimensions

Length m 54.15 55.3 46.1 57.1 54.1

Height m 16.87 17.5 13.9 13.11 15.44


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


Wing Area m2 321.1 350 184.2 271.9 311.1

Wing Sweep degree 35° 30° 30° 32° 35°

Performance



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


Max Speed Km/h 990 900 900 965 870


Max thrust capability kN 222.4 157 158.3 84.5 60

Design Weights

MTO Weight x103

Kg 211 250 103 161 175

Empty Weight x103

Kg 105.1 120.4 60 66.36 91 to 93

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

P a g e | 20


GRAPHS

P a g e | 21


Graph 1

Cruising Speed vs. Length

Length = 55.0m

P a g e | 22


Graph 2

Cruising Speed vs. Height

Height = 15.7m

P a g e | 23


Graph 3

Cruising Speed vs. Fuselage Diameter

Fuselage Diameter = 5.26m

P a g e | 24


Graph 4

Cruising Speed vs. Wing Span

Wing Span = 51.5m

P a g e | 25


Graph 5

Cruising Speed vs. Chord

Chord = 6.0m

P a g e | 26


Graph 6

Cruising Speed vs. Aspect Ratio

Aspect Ratio = 8.6

P a g e | 27


Graph 7

Cruising Speed vs. Wing Area

Wing Area = 348m2

P a g e | 28


Graph 8

Cruising Speed vs. Wing Sweep

Wing Sweep = 31.5°

P a g e | 29


Graph 9

Cruising Speed vs. Cruising Altitude

Cruising Altitude = 10800m

P a g e | 30


Graph 10

Cruising Speed vs. Service Ceiling

Service Ceiling = 12000m

P a g e | 31


Graph 11

Cruising Speed vs. Range

Range = 7200m

P a g e | 32


Graph 12

Cruising Speed vs. Maximum Speed

Max Speed = 940km/h

P a g e | 33


Graph 13

Cruising Speed vs. Number of Engines

Number of Engines = 3

P a g e | 34


Graph 14

Cruising Speed vs. Maximum Thrust Capability

Maximum Thrust Capability = 265kN

P a g e | 35


Graph 15

Cruising Speed vs. Maximum Take Off Weight

Maximum Take Off Weight = 272000 kg

P a g e | 36


Graph 16

Cruising Speed vs. Empty Weight

Empty Weight = 85000 kg

P a g e | 37


Graph 17

Cruising Speed vs. Wing Loading

Wing Loading = 710 kg/m3

P a g e | 38


Graph 18

Cruising Speed vs. Maximum Fuel Capacity

Maximum Fuel Capacity = 100000 litre

P a g e | 39


MEAN DESIGN PARAMETERS

P a g e | 40


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

P a g e | 41


WEIGHT ESTIMATION

P a g e | 42


WEIGHT ESTIMATION

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.

W0 = { Wcrew +Wpayload + Wfuel + Wempty }

Wfuel - weight of the fuel load at beginning of the flight

W0 = 𝑤𝑐𝑟𝑒𝑤 +𝑊𝑝𝑎𝑦𝑙𝑜𝑎𝑑

1− 𝑊𝑓𝑢𝑒𝑙

𝑊0 −

𝑊𝑒𝑚𝑝𝑡𝑦

𝑊0

𝑊𝑓

𝑊0 - Fuel weight fraction

𝑊𝑒

𝑊0 - Empty weight fraction

ESTIMATION OF We /W0:

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 We /W0

P a g e | 43


Estimation of Wf / W0:

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

The fuel weight ratio 𝑊𝑓

𝑊0 can be obtained from the product of mission segment weight at the

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,

𝑊1

𝑊0 = 0.97

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

𝑊2

𝑊1 = 0.985

Take off

Climb

Cruise

Glide

Landing

0 1

2 3

4 5

Loiter

P a g e | 44


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

𝑊4

𝑊3 = 1

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

𝑊5

𝑊4 = 0.995

The Brequet’s range equation is used to calculate the value of 𝑤3

𝑤2. As we all know that

maximum range is covered during cruise we considering this equation,

R = 𝑣∞

𝑐𝑗

𝐿

𝐷ln

𝑤2

𝑤3

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,

𝐿

𝐷 = 15

P a g e | 45


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 cj as 0.6 hr-1

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

R = 𝑣∞

𝑐𝑗

𝐿

𝐷ln

𝑤2

𝑤3

𝑤2

𝑤3 = 1.39135

𝑤3

𝑤2 = 0.718726

Now using all the fuel fractions,

𝑤5

𝑤0 =

𝑤1

𝑤0 x

𝑤2

𝑤1 x

𝑤3

𝑤2 x

𝑤4

𝑤3 x

𝑤5

𝑤4

𝑤5

𝑤0 = 0.68327

P a g e | 46


If at end of the flight, the fuel tanks are not completely empty, making six percent of

allowance for reserve and trapped fuel,

𝑤𝑓

𝑤0 = 1.06 1 −

𝑤5

𝑤0

𝑤𝑓

𝑤0= 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 𝑊𝑒

𝑊0 as 0.475

By substituting these values in:

W0 = 𝑤𝑐𝑟𝑒𝑤 +𝑊𝑝𝑎𝑦𝑙𝑜𝑎𝑑

1− 𝑊𝑓𝑢𝑒𝑙

𝑊0 −

𝑊𝑒𝑚𝑝𝑡𝑦

𝑊0

We get W0 as,

W0 = 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

𝑊0= 1.02 × 𝑊0

−0.06

FIRST:

We

𝑊0= 1.02 × 272626.4−0.06

We

𝑊0= 0.481355676

W0 = 282099.285

SECOND:

We

𝑊0= 1.02 × 282099.285−0.06

We

𝑊0= 0.4803702

W0 = 280587.572

THIRD:

We

𝑊0= 1.02 × 280587.572−0.06

We

𝑊0= 0.4805251

W0 = 280824.1

FOURTH:

We

𝑊0= 1.02 × 280824.1−0.06

We

𝑊0= 0.4805008

W0 = 280786.977

P a g e | 48


FIFTH:

We

𝑊0= 1.02 × 280786.977−0.06

We

𝑊0= 0.4805046

W0 = 280792.801

SIXTH:

We

𝑊0= 1.02 × 280792.801−0.06

We

𝑊0=0.480504

W0 = 280791.887

SEVENTH:

We

𝑊0= 1.02 × 280791.887−0.06

We

𝑊0=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 W0 = 280792 as the final estimate of the W0.

Max Take Off Weight (W0) = 280,792 kg.

We know that,

𝑤𝑓

𝑤0 = 0.33573

So, substituting the value of W0, we get the first estimation value of Wf,

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

Bypass

Ratio

Rolls Royce

Trent 772B-60 4788 320 6.8:1

5

Pratt & Whitney

PW4000-100 4270 310 7.4:1

5

CFM

International

CFM56-5C4

3990 151 3.9:1 6.4

General Electric

CF6-50 4104 240 6:1

4.4

Pratt & Whitney

JT9D-7R4H1 4030 250 6.3:1

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 : Trent 772B-60

Thrust : 71,100lb

Bypass ratio : 5.0

Inlet mass flow : 2030lb/sec

Fan diameter : 97.4in

Length : 154in

Stages : Fan, 8 IPC, 6 HPC, 1 HPT, 1 IPT, 4 LPT

Certification : Jan 1994

EIS : Mar 1995

P a g e | 53


FUEL WEIGHT VALIDATION

P a g e | 54


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 = 𝑵𝒖𝒎𝒃𝒆𝒓 𝒐𝒇 𝒆𝒏𝒈𝒊𝒏𝒆𝒔∗𝑻𝒉𝒓𝒖𝒔𝒕 𝒂𝒕 𝒂𝒍𝒕𝒊𝒕𝒖𝒅𝒆∗𝑹𝒂𝒏𝒈𝒆∗𝑺𝑭𝑪∗𝟏.𝟐

𝑪𝒓𝒖𝒊𝒔𝒆 𝑽𝒆𝒍𝒐𝒄𝒊𝒕𝒚

The factor of 1.2 is provided for reserve fuel.

Thrust at altitude is calculated using the relation:

Altitude = 10800m = 35433ft

𝜍 = 𝜌𝑎𝑙𝑡

𝜌0 = 0.3715/1.225 = 0.303

Cruise velocity = 872km/hr = 242.2m/s

To = 320kN

𝑇𝜍= 320×0.3031.2

𝑇𝜍= 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

2.1

0 * TT

0

alt

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 Sweep-

back 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.

The tip chord is given by,

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 Low-

wing 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.

)1(

2

b

SCroot

tip rootC C

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 minimum 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 CL for

cruise. The CL that will be obtainable from an airfoil section (for a given angle of attack) is

given by:

CL =0.9 x Cl x cosΛ.

Cl= 2×𝑊

𝜌×𝑣2×𝑆

P a g e | 58


DESIGN CALCULATION

(First Estimation)

Croot Calculation:

Croot = 2 × 395.48

58.32× (1+0.25) = 10.85m

Ctip Calculation:

Ctip = 0.25 x 10.85 = 2.7m

Cmean Calculation:

Cm = 2

3× 𝐶𝑟𝑜𝑜𝑡 ×

(1+λ+λ2 )

(1+λ)

Cm = 2

3×10.85×1.05 = 7.6m

Coefficient of Lift Calculation:

Section Lift Coefficient:

Cl= 2×𝑊

𝜌×𝑣2×𝑆

Cl = 2×710×9.8

0.3715×242.22 = 0.638

Wing Lift Coefficient:

CL =0.9 x Cl x cosΛ

CL =0.9 x 0.638 x cos35o = 0.47

)1(

2

b

SCroot

tip rootC C

P a g e | 59


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

designed.

DESIGN CHARACTERISTICS VALUES

W/S (kg/m2) 710

Wing area S (m2) 395.48

Aspect Ratio 8.6

Span b (m) 58.32

Taper ratio (𝛌) 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

P a g e | 61


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 CLmax.

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

– 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).

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

– 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 (as tenths of CL).

– 4th

digit (x0.1): design CL.

– 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:

𝑡

𝑐 =

0.3

𝑀

1

𝑀𝑐𝑜𝑠∆− 𝑀𝑐𝑜𝑠∆

13

[1 − 5 + 𝑀𝑐𝑜𝑠∆ 2

5 + (𝑀#)2

3.5

]23

Taking 𝑀# = 1.05 - 0.25 CL (cruise)

Where,

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,

𝑡

𝑐 = 0.12

P a g e | 63


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

Name

Thickness

(%)

Camber

(%)

Lift Coeff.

(CL)

Lift-to-Drag

(L/D)

Stall Angle

(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 (𝑪𝒍)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 (𝐿

𝐷)

CALCULATIONS:

Available 𝐶𝑙𝑚𝑎𝑥 =

1.238

3+

1.107

3+

1.015

3 = 1.12

𝐶𝐿max 𝑎𝑣𝑎𝑖𝑙= 0.9 × 𝐶𝑙max 𝑎𝑣𝑎𝑖𝑙

= 0.9 × 1.12 = 1.008

Flaps Selection

For the current design, double slotted flap is selected. ∆𝐶𝑙 𝑚𝑎𝑥 of the double slotted flap

for different configurations is given in the table below:

FLAPS TAKE OFF LANDING

Double slotted flap 20o 40o

∆𝑪𝒍 𝒎𝒂𝒙/𝒄𝒐𝒔∆ 1.825 2.5

∆𝑪𝒍 𝒎𝒂𝒙 1.5 2.05

P a g e | 68


∆𝐶𝑙𝑚𝑎𝑥 = 𝐶𝐿 max 𝑟𝑒𝑞 - 𝐶𝐿 max 𝑎𝑣𝑎𝑖𝑙𝑎𝑏𝑙𝑒

𝐶𝐿 max 𝑟𝑒𝑞 = 𝐶𝐿 max 𝑎𝑣𝑎𝑖𝑙𝑎𝑏𝑙𝑒 + ∆𝐶𝑙𝑚𝑎𝑥

𝐶𝐿 max 𝑟𝑒𝑞 (Take Off) = 1.008+1.5 = 2.508

𝐶𝐿 max 𝑟𝑒𝑞 (Landing) = 1.008+2.05 = 3.058

𝑉𝑆𝑡𝑎𝑙𝑙 = 0.25× 𝑉𝐶𝑟𝑢𝑖𝑠𝑒 = 60.55 m/s

We Have,

W/S=700.722 kg/m2

From this,

S = 400.72 𝑚2

b = 58.32 m (from table)

DESIGN CALCULATION

(Second Estimation)

Croot Calculation:

Croot = 2 × 400.72

58.32× (1+0.25) = 11m

Ctip Calculation:

Ctip = 0.25 x 11 = 2.75m

)1(

2

b

SCroot

tip rootC C

P a g e | 69


Cmean Calculation:

Cm = 2

3× 𝐶𝑟𝑜𝑜𝑡 ×

(1+λ+λ2 )

(1+λ)

Cm = 2

3×11×1.05 = 7.7m

Coefficient of Lift Calculation:

Section Lift Coefficient:

Cl= 2×𝑊

𝜌×𝑣2×𝑆

Cl = 2×700.72×9.8

0.3715×242.22 = 0.63022

Wing Lift Coefficient:

CL =0.9 x Cl x cosΛ

CL =0.9 x 0.63022 x cos35o = 0.4646


W/S (kg/m2) 700.72

Wing area S (m2) 400.72

Aspect Ratio 8.6

Span b (m) 58.32

Taper ratio (𝛌) 0.25

Root Chord (m) 11

Tip chord (m) 2.75

Mean chord (m) 7.7

Sweepback Angle(∆) 35°

Cruise Lift Coefficient (𝑪𝒍) 0.63022

P a g e | 70


LIFT ESTIMATION

P a g e | 71


LIFT ESTIMATION

LIFT:

Component of aerodynamic force generated on aircraft perpendicular to flight

direction.

Lift Coefficient (CL)

• Amount of lift generated depends on:

– Planform area (S), air density (), flight speed (V), lift coefficient (CL)

• 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.

21Lift ( )

2L LV SC qSC

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 at Cruise

𝜌 = 0.3715 (at the cruising altitude of 10800m)

V = 242.2 m/s

S = 400.72 kg/m2

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 = 2751761.6 N

21Lift ( )

2L LV SC qSC

P a g e | 74


Lift at Take-Off

𝜌 = 1.225 (at sea altitude)

V = 0.7 x Vlo = 0.7 x 1.2 x Vstall

S = 400.72 kg/m2

CL(take-off) = 2.508 (flaps extended and kept at the take-off position of 20o)


L(take-off) = 1

2× 1.225 × (0.7 × 1.2 × 66.86)2 × 400.72 × 2.508

Lift at take-off = 𝟏𝟗𝟒𝟏𝟔𝟐𝟕.𝟕 𝐍

Lift at Landing


V = 0.7 x Vt = 0.7 x 1.3 x Vstall

S = 400.72 kg/m2

CL(landing) = 3.058 (flaps extended and kept at the landing position of 40o)


L(landing) = 1

2× 1.225 × (0.7 × 1.3 × 60.55)2 × 400.72 × 3.058

Lift at landing = 𝟐𝟐𝟕𝟖𝟕𝟒𝟒.𝟕 𝐍

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 (CD)

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 Due to shear stresses produced in boundary layer.

o Significantly more for turbulent than laminar types of boundary layers.

Form (Pressure) Drag

o Due to static pressure distribution around body - component resolved in

direction of motion.

o Sometimes considered separately as forebody and rear (base) drag

components.

P a g e | 77


Wave Drag

o Due to the presence of shock waves at transonic and supersonic speeds.

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

boundary layer.

o 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,

𝐷 = 1

2𝜌𝑉2𝑆 𝐶𝐷,0 +

∅𝐶𝐿2

𝜋𝐴𝑒

For calculating Ø, we use the formula,

Ø =

16𝑕

𝑏

2

1 + 16𝑕

𝑏

2

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

h = 2m

b = 58.32m

Ø = 16×

2

58.32

2

1 + (16×2

58.32)2

= 0.2314

Drag at Cruise

𝜌 = 0.3715 (at the cruising altitude of 10800m)

V = 242.2 m/s

S = 400.72 kg/m2

CL(cruise) = 0.63022 (from the wing and airfoil estimation)

Substituting all these values in the general drag equation,

D(cruise) = 1

2× 0.3715 × (242.2)2 × 400.72 (0.0030 +

0.2314 ×0.63022 2

3.14×8.6×0.8)

Drag at cruise = 31674.846 N

P a g e | 80


Drag at Take-off


V = 0.7 x Vlo = 0.7 x 1.2 x Vstall

S = 400.72 kg/m2

CL(take-off) = 2.508 (flaps extended and kept at the take-off position of 20o)


D = 1

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

0.2314×2.5082

3.14×8.6×0.8)

Drag at take-off = 54482.6 N

Drag at Landing


V = 0.7 x Vt = 0.7 x 1.3 x Vstall

S = 400.72 kg/m2

CL(landing) = 3.058 (flaps extended and kept at the landing position of 40o)


D = 1

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

0.2314×3.0582

3.14×8.6×0.8)

Drag at landing = 76876.7 N

P a g e | 81


LANDING GEAR

ARRANGEMENT

P a g e | 82


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 load 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

P a g e | 83


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

P a g e | 84


• 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.

P a g e | 85


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.

P a g e | 86


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.

P a g e | 87


FUSELAGE DESIGN

P a g e | 88


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).

P a g e | 89


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).



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:

P a g e | 90


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 m2 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 15

o.

P a g e | 91


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 flap 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.

P a g e | 92


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

o 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.

TYPICAL FLIGHT DECK LAYOUT:

P a g e | 93


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.


Overall Length (m) 55.0

Fuselage Width (m) 5.26

Cabin Width (m) 5.0

Length/Width 10.456

SEATING ARRANGEMENT:

P a g e | 94


PERFORMANCE

CHARACTERISTICS

P a g e | 95


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 (SLO)

– 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 =1.44𝑊2

𝑔𝜌𝑆𝐶𝐿,𝑚𝑎𝑥 𝑇

• This shows its sensitivity to W (W2) and (1/

2 since T also varies with ).

• 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.

𝑆𝐿𝑂 =1.44𝑊2

𝑔𝜌𝑆𝐶𝐿,𝑚𝑎𝑥 𝑇 − 𝐷 + 𝜇𝑟 𝑊 − 𝐿 𝑎𝑣

• 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:

𝐷 = 1

2𝜌𝑉2𝑆 𝐶𝐷,0 +

∅𝐶𝐿2

𝜋𝐴𝑒

• Where accounts for drag reduction when in ground effect:

Ø =

16𝑕

𝑏

2

1 + 16𝑕

𝑏

2

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

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

P a g e | 96


CALCULATION:

Ø = 16×

2

58.32

2

1 + (16×2

58.32)2

= 0.2314

D = 1

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

0.2314×2.5082

3.14×8.6×0.8) = 54482.6 N

Slo = 1.44×(280792 ×9.81)2

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 Vc.

• Force balance gives:

𝐿 = 𝑊 𝑐𝑜𝑠𝛾𝑐

𝑇 = 𝐷 + 𝑊 𝑠𝑖𝑛𝛾𝑐

𝑉𝑐 =(𝑇 − 𝐷) × 𝑉

𝑊

R/Cmax = (𝑇𝑉𝑠𝑡𝑎𝑙𝑙 −𝐷𝑉𝑠𝑡𝑎𝑙𝑙 )

𝑊 =

3×320000 ×60.55 −(40382 .15×60.55)

280792 ×9.81

R/Cmax = 20.2 m/s

P a g e | 97


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

– Fr = mV2/r = Lsin

• tan = V2/(Rg)

• 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.

• 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 = (sec

2 - 1) = (n

2 - 1)

• R = V2 / (g (n

2 - 1)) and = V/R = (g (n

2 - 1)) / V

CALCULATION:

W = Lcos

Let = 300

n = 𝐿

𝑊 = 1.1547

R = V2

g n2 − 1 = 10357.16 m

= 𝑉

𝑅 = 0.0234 rad/s

P a g e | 98


GLIDING

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

Force balance gives:

D

L

1tan 1

15

1tan 1

3.814o

LANDING PERFORMANCE

APPROACH & LANDING

• Consists of three phases:

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

– Flare - transitional manoeuvre with airspeed reduced from about 1.3 Vstall

down to touch-down speed.

– Ground roll - from touch-down to rest.

P a g e | 99


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

𝑆𝐿𝑂 =1.69𝑊2

𝑔𝜌𝑆𝐶𝐿,𝑚𝑎𝑥 𝐷 + 𝜇𝑟 𝑊 − 𝐿 𝑎𝑣

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

𝑆𝐿𝑂 =1.69𝑊2

𝑔𝜌𝑆𝐶𝐿,𝑚𝑎𝑥 𝑇𝑅 + 𝐷 + 𝜇𝑟 𝑊 − 𝐿 𝑎𝑣

CALCULATION:

D = 1

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

0.2314×3.0582

3.14×8.6×0.8) = 76876.7 N

Sl = 1.69×(280792 ×9.81)2

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

P a g e | 100


3-VIEW DIAGRAM

P a g e | 101


3-VIEW DIAGRAM

58

.32m

1

5.7

m

55m

P a g e | 102


P a g e | 103


P a g e | 104


CONCLUSION

P a g e | 105


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.

P a g e | 106


BIBLIOGRAPHY

P a g e | 107


BIBLIOGRAPHY

1. Introduction to Flight by J.D.Anderson

2. Aerodynamics by Clancy

3. Fundamentals of Aerodynamics by J.D.Anderson

4. The Design of the Aeroplane by Darrol Stinton

5. Jane’s All the World’s Aircraft

6. Aircraft Design: A Conceptual Approach by Daniel. P. Raymer

WEBSITE REFERENCES

1. www.wikipedia.org

2. www.naca/aerofoil.gov

3. www.worldaircraftdierctory.com

4. www.boeing.com

5. www.airbus.com

6. www.airliners.net

7. And other websites related to design of aircrafts.

40504526-Aircraft-Design-Project-280-Seat-Transport

Documents

Transcript of 40504526-Aircraft-Design-Project-280-Seat-Transport