aircraft design project-1(50 seated aircraft)

8/11/2019 Aircraft Design Project-1(50 Seated Aircraft)

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DE

T

AN AIRCR

RENJITH.R

CHARLES.

INDERJITH

in partial

BAC

AE

DHANALAKSHMI

ANNA U

SIGN OF A 50 SEATED

ANSPORT AIRCRAFT

AFT DESIGN PROJECT REPOR

Submitted by

- (7219111

.PHILIPOSE - (7219111

. V - (7219111

fulfilment for the award of the degre

of

ELOR OF ENGINEERING

IN

ONAUTICAL ENGINEERING

SRINIVASAN COLLEGE OF EN

COIMBATORE

IVERSITY:: CHENNAI 600 0

APRIL 2014

- I

1015)

1003)

1301)

INEERING

25


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50 SEATED TRANSPORT AIRCRAFT


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ANNA U

B

Certified that this pro

TRANSPORT AIRC

(721911101015),CHAR

(721911101301) who car

SIGNATURE

Mr.P.DHARMADURAI, B.E

SUPERVISOR

LECTURER Department of Aeronautical

Dhanalakshmi Srinivasan Co

Engineering,Coimbatore

Submitted for the Aircraf

at Dhanalakshmi Srinivas

INTERNAL EXAMINER

NIVERSITY:: CHENNAI 600 0

NAFIDE CERTIFICATE

ject report on "DESIGN OF A

AFT is the bonafide work o

ES.C.PHILIPOSE(721911101003)I

ried out the project work under my su

,(M.E) Mr.S.RAMESHBAB

HEAD OF THE DE

Department of Aeronngg, Dhanalakshmi Sriniv

llege Engineering,Coimbat

Design ProjectI VivaVoce held

an College of Engineering ,Coimbator

EXTERNAL

5

50 SEATED

RENJITH.R

DERJITH.V,

ervision.

SIGNATURE

,M.E, (Ph.D)

ARTMENT

utical Engg san College of

re

n ..................

e641105.

EXAMINER


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ACKNOWLEDGEMENT

Firstly I would like to thank the Almighty god for always being by myside and providing me with strength and capability to face all types of situations

during this project tensure

I thank our beloved Chairman A.Srinivasan , Dhanalakshmi Srinivasan

Groups of Institution, Coimbatore for providing the facilities

I extend my fullest and ever owing thanks to Dr.S.Charles Principal,

Dhanalakshmi Srinivasan College of Engineering and technology, Coimbatore,

for the academic freedom and inspiration

We also thank our Professor and Head of the department,

Mr.S.RameshBabu,M.E,(Ph.D,) Our Lecturer Mr.P.Dharmadurai.B.E

(M.E),and staff members of Aeronautical department of Dhanalakshmi

Srinivasan College of Engineering for leading their support to this project.

I also thank everyone who lent us support in the completion of this

project.


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i

ABSTRACT

The aim of this design project is to design a 50 Seated Transport

Aircraft by comparing the data and specifications of present transport aircrafts

and to calculate performance details. The aircraft designed is such that the

landing and take-off field lengths they require are accordingly shorter than

those for the larger transport aircraft minimum drag and maximum thrust is also

taken into consideration. Then the necessary graphs have to be plotted for

further performance calculation. Required diagrams are also drawn.


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ii

TABLE OF CONTENTS

CHAPTER NO. TITLE PAGE NO.

ABSTRACT i

TABLE OF CONTENTS ii

LIST OF TABLES vi

LIST OF FIGURES vii

LIST OF ABBREVIATIONSx

01 INTRODUCTION 1

1.1 Preliminary Design 2

1.2 Project Design 3

1.3 Detail Design 4

1.4 Manufacturing 9

1.5 Testing 10

02 COMPARATIVE DATA SHEET 12

2.1 Specification 14

03 GRAPHS 19

3.1 Graphs for Comparison of Contemporary 19

Aircraft

3.2 Mean Design Parameter 28

04 WEIGHT ESTIMATION 29

4.1 First Weight Estimation 29

4.2 Estimation of We/Wo 30

4.3 Estimation of Wf/Wo 31

4.4 Mission Profile 32


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4.5 Calculation of Wo 37

4.6 Iteration 38

05 POWER PLANT SELECTION 40

5.1 Required Engine 41

5.2 Engine Specification 43

06 FUEL WEIGHT VALIDATION 45

6.1 Calculation 46

07 WING SELECTION 47

7.1 Introduction 47

7.2 Wing Geometry Design 47

7.3 Wing Chord Design 49

08 AIRFOIL SELECTION 52

8.1 Introduction 52

8.2 Estimation of the Critical Performance 54

Parameter

8.3 Airfoil Geometry 57

09 FLAP SELECTION 62

9.1 Introduction 62

9.2 Types of Flaps 62

9.3 Selected Flap 65

10 FUSELAGE AND CABIN LAYOUT 67

10.1 Introduction 67

10.2 Fuselage Layout 68

10.3 Fuselage Sizing 69

10.4 Passenger Cabin Layout 71


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10.5 Rear Fuselage 73

11 TAIL SELECTION 75

11.1 Tail Surface 75

11.2 T-Tail 75

11.3 Horizontal and Vertical Tail Calculation 77

12 C.G CALCULATION 79

12.1 Center Of Gravity 79

13 LANDING GEAR SELECTION 81

13.1 Introduction 81

13.2 Landing Gear Design Requirement 81

13.3 Landing Gear Configuration 82

13.4 Retractable Landing Gear 83

13.5 Tyre Sizing 85

13.6 Landing Gear Height 85

13.7 Landing Gear Attachment 86

14 LIFT ESTIMATION 87

14.1 Lift 87

14.2 Lift Coefficient [CL] 87

14.3 Generation of Lift 87

14.4 Calculation 90

15 DRAG ESTIMATION 91

15.1 Drag 91

15.2 Drag Coefficient 91

15.3 Drag Calculation 93

16 PERFORMANCE CHARACTERISTICS 95

16.1 Takeoff Performance 95


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16.2 Ground Roll Takeoff Distance 95

16.3 Climbing Performance 96

16.4 Manoeuvres/Turning Performance 99

16.5 Gliding Performance 100

16.6 Landing Performance 101

16.7 Endurance Calculation 102

17 THREE VIEW DIAGRAM OF AIRCRAFT 103

17.1 Surface Model 104

18 CONCLUSION 107

18.1 Design Data 108

19 BIBLIOGRAPHY 110


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vi

LIST OF TABLES

TABLE NO. TABLE PAGE

4.1 FUEL FRACTION 33

4.2 LIFT/DRAG RATIO 34

4.3 SPECIFIC FUEL CONSUMPTION 36

5.1 ENGINE SELECTION 41

7.1 DIHEDRAL ANGLE () 50

8.1 NACA 6 SERIES AIRFOILS 56

8.2 SELECTED AIRFOIL 57

9.1 CL MAX DUE TO FLAP 66

11.1 HORIZONTAL AND VERTICAL TAIL 77

CALCULATION


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vii

LIST OF FIGURES

FIGURE NO. FIGURE PAGE NO.

1.1 PHASE OF DESIGN 2

1.2 PRELIMINARY DESIGN CONCEPT 3

1.3 DESIGN CRITERIA 5

1.4 LIFT & DRAG IN AIRFOIL 7

3.1 THRUST VS ASPECT RATIO 19

3.2 THRUST VS CRUISE SPEED 19

3.3 THRUST VS EMPTY WEIGHT 20

3.4 THRUST VS GROSS WEIGHT 20

3.5 THRUST VS HEIGHT 21

3.6 THRUST VS LENGTH 21

3.7 THRUST VS MAX. TAKEOFF WEIGHT 22

3.8 THRUST VS PROPELLER POWER 22

3.9 THRUST VS RANGE 23

3.10 THRUST VS RATE OF CLIMB 23

3.11 THRUST VS SERVICE CEILING 24

3.12 THRUST VS SPEED 24

3.13 THRUST VS THRUST LOADING 25

3.14 THRUST VS USEFUL LOAD 25

3.15 THRUST VS WING AREA 26

3.16 THRUST VS WING SPAN 26

3.17 THRUST VS WING LOADING 27


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4.1 MISSION PROFILE 32

5.1 ENGINE LAYOUT 42

5.2 CROSS-SECTIONAL VIEW 43

7.1 WING GEOMETRY DESIGN 47

7.2 WING LAYOUT IN AIRCRAFTS 51

8.1 AIRFOIL LAYOUT 52

8.2 AIRFOIL GEOMETRY 57

8.3 ANGLE OF ATTACK VS LIFT COEFFICIENT 59

FOR NACA 65-410

8.4 ANGLE OF ATTACK VS LIFT COEFFICIENT 59

FOR NACA 65(2)-415

8.5 PERFORMANCE CURVE FOR CHOSEN 60

AIRCRAFT

9.1 TYPES OF FLAPS 64

9.2 DOUBLE FLOWER-SLOTTED 65

10.1 CABIN LAYOUT 67

10.2 COCKPIT LAYOUT 70

10.3 HONEYWELLS AVIONIC SUITE 70

10.4 COCKPIT INSTRUMENT LAYOUT 71

10.5 PASSENGER CABIN LAYOUT 71

11.1 TYPES OF AIRCRAFT TAIL 75

11.2 STABILITY DUE TO HORIZONTAL TAIL 76

12.1 C.G INDICATION 79

12.2 C.G LAYOUT 80


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13.1 MAIN LANDING GEAR ASSEMBLY 81

13.2 TYPES OF LANDING GEAR 82

13.3 MAIN LANDING GEAR IN AIRCRAFT 83

13.4 LANDING GEAR MARKING 84

13.5 NOSE LANDING GEAR DEPOYED 84

14.1 GENERATION OF LIFT 87

14.2 AERODYNAMIC FORCES DUE TO LIFT 88

14.3 PRESSURE VARIATION 89

14.4 LIFT AT DIFFERENT ANGLES 89

14.5 LIFT CURVE 90

15.1 DRAG SEPARATION 91

15.2 FORM DRAG 92

15.3 DRAG AT DIFFERENT MACH NUMBERS 93

15.4 TYPICAL STREAMLINING EFFECT 93

16.1 TAKEOFF FOR AIRCRAFT 95

16.2 WEIGHT COMPONENT INDICATION 97

16.3 THRUST VS CLIMB ANGLE 98

16.4 GLIDING LAYOUT 101

17.1 AIRCRAFT FRONT VIEW 103

17.2 AIRCRAFT TOP VIEW 103

17.3 AIRCRAFT SIDE VIEW 103

17.4 SURFACE VIEW OF AIRCRAFT 104


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x

LIST OF SYMBOLS & 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)

Cm - Mean Aerodynamic Chord (m) C

CD - Drag Co-efficient

CD o - Zero Lift Drag Co-efficient

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

CL - Lift Co-efficient

D - Drag (N)

E - Endurance (hr)

E - Oswald efficiencyL - Lift (N)

M - Mach number of aircraft

Mff - Mission fuel fraction

R - Range (km)

Re - Reynolds Number

S - Wing Area (m)

Sref - Reference surface area

Swet - Wetted surface area

Sa - Approach distance (m)

Sg - Ground roll Distance (m)

T - Thrust (N)

Tcruise - Thrust at cruise (N)

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


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Vcruise - Velocity at cruise (m/s)

Vstall - Velocity at stall (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/m)

- Density of air (kg/m)

- Dynamic viscosity (Ns/m)

- Tapered ratio


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1

INTRODUCTION

The start of the design process requires the recognition of a need. Thisnormally comes from a project brief or a request for proposals (RFP). Such

documents may come from various sources:

Established or potential customers

Government defense agencies.

Analysis of the market and the corresponding trends from aircraft demand

Development of an existing product (e.g. aircraft stretch or engine

change).

Exploitation of new technologies and other innovations from research and

development.

It is essential to understand at the start of the study where the project

originated and to recognize what external factors are influential to the design

before the design process is started.

At the end of the design process, the design team will have fully specified

their design configuration and released all the drawings to the manufacturers. In

reality, the design process never ends as the designers have responsibility for

the aircraft throughout its operational life. This entails the issue of modifications

that are found essential during service and any repairs and maintenance

instructions that are necessary to keep the aircraft in an airworthy condition. The

design method to be followed from the start of the project to the nominal end

can be considered to fall into three main phases. These phases are illustrated in

Figure 2.0.

Chapter-1


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1.1 PRELIMINARY DESIGN

The preliminary phase (sometimes called the conceptual design stage)

starts with the project brief and ends when the designers have found and refined

a feasible baseline design layout. In some industrial organizations, this phase is

referred to as the feasibility study. At the end of the preliminary design phase,

a document is produced which contains a summary of the technical and

geometric details known about the baseline design n. This forms the initial draft

of a document that will be subsequently revised to contain a thorough

description of the aircraft. This is known as the aircraft Type Specification.

The ultimate objective during preliminary design is to ready the company

for the detail design stage, also called full-scale development. Thus, the end of

preliminary design usually involves a full scale development proposal. In

todays environment, this can result in a situation jokingly referred to as you -

bet-your-company. The possible loss on an overrun contrast o from lack of

sales can exceed the net worth of the company! Preliminary design must

establish confidence that the airplane can be built in time and at the estimated

cost.


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1.2 PROJECT DESIGN

The next phase (project design) takes the aircraft configuration defined

towards the end of the preliminary design phase and involves conducting

detailed analysis to improve the technical confidence in the design. Wind tunnel

tests and computational fluid dynamic analysis are used to refine the

aerodynamic shape of the aircraft. Finite element analysis is used to understand

the structural integrity. Stability and control analysis and simulations will be

used to appreciate the flying characteristics. Mass and balance estimations willbe performed in increasingly fine detail. Operational factors (cost, maintenance

and marketing) and manufacturing processes will be investigated

1.2.1 Introduction to the project

1) Project brief

2) Problem definition

3) Design concepts

Fig 1.2 Preliminary design concept


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4) Initial sizing and layout

5) Initial estimates

6) Constraint analysis and trade-offs

7) Revised baseline layout

8) Further work

9) Study review

Design project work, as taught at most universities, concentrates on

the preliminary phase of the design process. The project brief, or request for

proposal, is often used to define the design problem. Alternatively, the problem

may originate as a design topic in a student competition sponsored by industry,

a government agency, or a technical society. Or the design project may be

proposed locally by a professor or a team of students. Such design project

assignments range from highly detailed lists of design objectives and

performance requirements to rather vague calls for a new and better

replacement for existing aircraft. In some cases student teams may even beasked to develop their own design objectives under the guidance of their design

professor.

1.3 DETAIL DESIGN

The process of designing an aircraft, generally divided into three

distinct phases: conceptual design, preliminary design, and detail design. Each

phase has its own unique characteristics and influence on the final product.These phases all involve aerodynamic, propulsion, and structural design, and

the design of aircraft systems.


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1.3.1. Design phases:

`Conceptual design activities are characterized by the definition and

comparative evaluation of numerous alternative design concepts potentially

satisfying an initial statement of design requirements. The conceptual design

phase is iterative in nature. Design concepts are evaluated, compared to the

requirements, revised, reevaluated, and so on until convergence to one or more

satisfactory concepts is achieved.

Fig 1.3 Design criteria


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During this process, inconsistencies in the requirements are often

exposed, so that the products of conceptual design frequently include a set of

revised requirements. During preliminary design, one or more promising

concepts from the conceptual design phase are subjected to more rigorous

analysis and evaluation in order to define and validate the design that best meets

the requirements. Extensive experimental efforts, including wind-tunnel testing

and evaluation of any unique materials or structural concepts, are conducted

during preliminary design. The end product of preliminary design is a complete

aircraft design description including all systems and subsystems.

During detail design the selected aircraft design is translated into the

detailed engineering data required to support tooling and manufacturing

activities.

1.3.2. Requirements

The requirements used to guide the design of a new aircraft are

established either by an emerging need or by the possibilities offered by some

new technical concept or invention. Requirements can be divided into twogeneral classes: technical requirements (speed, range, payload, and so forth) and

economic requirements (costs, maintenance characteristics, and so forth).

1.3.3. Aerodynamic design

Initial aerodynamic design centers on defining the external geometry and

general aerodynamic configuration of the new aircraft.

The aerodynamic forces that determine aircraft performance capabilitiesare drag and lift. The basic, low-speed drag level of the aircraft is

conventionally expressed as a term at zero lift composed of friction and pressure

drag forces plus a term associated with the generation of lift, the drag due to lift

or the induced drag. Since wings generally operate at a positive angle to the

relative wind (angle of attack) in order to generate the necessary life forces, the

wing lift vector is tilted aft, resulting in a component of the lift vector in the

drag direction (see illustration).


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Aircraft that fly near or above the speed of sound must be designed to

minimize aerodynamic compressibility effects, evidenced by the formation of

shock waves and significant changes in all aerodynamic forces and moments.

Compressibility effects are mediated by the use of thin airfoils, wing and tail

surface sweepback angles, and detailed attention to the lengthwise variation of

the cross-sectional area of the configuration.

1.3.4. Propulsion design

Propulsion design comprises the selection of an engine from among theavailable models and the design of the engine's installation on or in the aircraft.

Selection of the best propulsion concept involves choosing from among a wide

variety of types ranging from reciprocating engine-propeller power plants

through turboprops, turbojets, turbofans, and ducted and undusted fan engine

developments. The selection process involves aircraft performance analyses

comparing flight performance with the various candidate engines installed. In

the cases where the new aircraft design is being based on a propulsion system

which is still in development, the selection process is more complicated.

1.3.5. Structural design

Structural design begins when the first complete, integrated aerodynamic

and propulsion concept is formulated. The process starts with preliminary

estimates of design air loads and inertial loads (loads due to the mass of the

aircraft being accelerated during maneuvers).

Fig 1.4 Lift & Drag in airfoil


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During conceptual design, the structural design effort centers on a first-

order structural arrangement which defines major structural components and

establishes the most direct load paths through the structure that are possible

within the constraints of the aerodynamic configuration. An initial

determination of structural and material concepts to be used is made at this time,

for example, deciding whether the wing should be constructed from built up

sheet metal details, or by using machined skins with integral stiffeners, or from

fiber in forced

composite materials.

1.3.6. Aircraft systems design

Aircraft systems include all of those systems and subsystems required for

the aircraft to operate. Mission systems are those additional systems and

subsystems peculiar to the role of military combat aircraft. The major systems

are power systems, flight-control systems, navigation and communication

systems, crew systems, the landing-gear system, and fuel systems.

Design of these major subsystems must begin relatively early in theconceptual design phase, because they represent large dimensional and volume

requirements which can influence overall aircraft size and shape or because they

interact directly with the aerodynamic concept (as in the case of flight-control

systems) or propulsion selection (as in the case of power systems).

DESIGN SEQUENCE

1. Define the mission

2. Compare the past design

3. Parametric selection

a. Geometry

b. Shape


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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. Stability and control

a. Tail

b. Flaps

c. Control surfaces

1.4 MANUFACTURING

Businesses in this industry do one or more of the following:

manufacture complete aircraft; manufacture aircraft engines, propulsion units

and other related equipment or parts; develop and make prototypes of aircraft;

aircraft conversions (i.e. major modification to systems); and complete aircraft


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overhaul and rebuilding (i.e. periodic restoration of aircraft to original design

specifications).

Industry Products

Aircraft

Aircraft engines and engine parts

Other aircraft parts and auxiliary equipment

Industry Activities

Manufacturing and rebuilding of aircraft

Developing and producing prototypes for aircraft

blimps, gliders, hand gliders, ultra light aircraft and helicopters

Manufacturing aircraft engines and engine parts

Developing and producing prototypes for aircraft engines and engine

Parts

Manufacturing aircraft assemblies, subassemblies, propellers, joints, and

other parts

Manufacturing aircraft auxiliary parts Developing and producing prototypes for aircraft parts and auxiliary

equipment

1.5 TESTING

Flight testing is a branch of aeronautical engineering that develops and

gathers data during flight of an aircraft and then analyzes the data to evaluate

the flight characteristics of the aircraft and validate its design, including safetyaspects.

The flight test phase accomplishes two major tasks:

Finding and fixing any aircraft design problems and then

Verifying and documenting the aircraft capabilities for government

certification or customer acceptance

.


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The flight test phase can range from the test of a single new system for an

existing aircraft to the complete development and certification of a new aircraft.

Therefore the duration of a flight test program can vary from a few weeks to

many years.

Examples of some subsystems we have performed aerospace testing on

include:

Airframes: Structural, Fatigue,

Antennas

Avionics

Power Inverters,

Communications

Flight Control Surfaces, Winglets

Landing Gear

Oxygen Systems

Passenger Service Units (PSU's) Rotor Systems

Windows and doors

Etc..

3


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COMPARATIVE DATA SHEET

In the designers perspective it is necessary to compare the existingairplanes that are of the same type as that of our desired airplane. Their

important parameters, positive aspects to b e considered and pitfalls to be

overcome are taken into consideration.

The data have been collected from various sites from the internet for 50

seated TRANSPORT AIRCRAFT design.

Several parameters are compared for over 15 aircrafts and different

critical parameters were plotted on graph. They are

Cruise speed

Range

Wing area

Thrust loading

Empty weight

Maximum take-off weight

Length

Wing span

Aspect ratio

Thrust

Power plant

Service ceiling

Speed

Wing area

Wing loading

Thrust power

Chapter-2


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No of engines

Crew member

Types of Engine

Endurance

Height


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SPECIFICATION

TABLE-1.1

SPECIFICATIONS UNITS NAME OF THE AIRCRAFTS

BOMBARDIER

CRJ100

ANTONOV

AN-140

ATR 42-200

ENGINE NAME

- GE CF34-3A1 KlimovTV3-

117VMA-

SBM1

Pratt&Whitney

Canada

PW120

NO.OF.ENGINES - 2 2 2

PROPELLER POWER KW 1,446 1,838 1,300

THRUST POWER KN 26.2 29.8 24.4

THRUST LOADING - 0.424 0.665 0.609

LENGTH m 26.77 22.6 22.67

HEIGHT m 6.22 8.23 7.59

WING SPAN m 21.21 26.4 24.57

WING AREA m2 48.35 51 54.5

ASPECT RATIO - 9.30 13.665 11.07

WING LOADING Kg/m2 126.7 104.74 87.26

EMPTY WEIGHT Kg 13,655 12,810 10,500

GROSS WEIGHT Kg 19,781 18,152 15,256

MAX.TAKE OFF WEIGHT Kg 24,041 21,500 15,550

CREW MEMBERS - 2 2 2

RANGE Km 3,000 1,380 1,885

CRUISE SPEED Km/hr 510 460 494

SPEED Km/hr 860 575 754

SERVICE CEILING m 12,496 7,600 7,600

RATE OF CLIMB m/s 9.27 6.83 6.89

USEFULL LOAD Kg 6,126 5,342 4,756

Chapter-3


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TABLE-1.2

SPECIFICATIONS UNITS NAME OF THE AIRCRAFTSFOKKER 50 HANDLEY

PAGE DART

HERALD

EMBRAER

ERJ-145

ENGINE NAME

- Pratt & Whitney

Canada

PW125B

Rolls-Royce

Dart Mk.527

Rolls Royce

AE-3007A




THRUST LOADING - 0.605 0.594 0.459

LENGTH m 25.25 23.01 29.87

HEIGHT m 8.32 7.32 6.75

WING SPAN m 29 28.9 20

WING AREA m2 70 82.3 51.2

ASPECT RATIO - 12.01 10.14 8.12

WING LOADING Kg/m2 73.14 55.62 112.59

EMPTY WEIGHT Kg 12,250 11,345 11,667

GROSS WEIGHT Kg 17,370 15,923 17,432



RANGE Km 2,055 2,632 2,445


SPEED Km/hr 560 654 833

SERVICE CEILING m 7,620 8,140 11,277.60


USEFULL LOAD Kg 5,120 4,578 5,765


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TABLE-1.3

SPECIFICATIONS UNITS NAME OF THE AIRCRAFTSSAAB 2000 CASA CN-235 XIAN MA60

ENGINE NAME - Allison AE-

2100A

General Electric

CT7-9C3

Pratt &Whitney

Canada PW127J


PROPELLER

POWER

KW 3,096 1,305 2,051

THRUST POWER KN 39.43 24.44 32.17

THRUST LOADING - 0.459 0.7206 0.6913

LENGTH m 29.87 27.28 21.4

HEIGHT m 6.75 7.73 8.18

WING SPAN m 20 24.76 25.81

WING AREA m2 51.2 55.7 59.1

ASPECT RATIO - 8.12 11.00 11.27

WING LOADING Kg/m 112.59 95.87 78.68

EMPTY WEIGHT Kg 11,667 13,800 9,800

GROSS WEIGHT Kg 17,432 19,140 14,450

MAX.TAKE OFF

WEIGHT

Kg 20,600 22,800 15,100


RANGE Km 2,445 2,185 4,355


SPEED Km/hr 833 594 514

SERVICE CEILING m 9448.8 7,620 7,620


USEFULL LOAD Kg 5,340 4,650 4,785


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TABLE-1.4

SPECIFICATIONS UNITS NAME OF THE AIRCRAFTS

AVRO 748 BOMBARDIER

DASH 8

DE

HAVILLAND

CANADA

DASH 7

ENGINE NAME

- Rolls-royce

dart Rda 7 mk

536-2

2PW 123B Pratt&Whitney

Canada PT6A-

50




THRUST LOADING - 0.7965 0.583 0.628

LENGTH m 24.56 22.07 27.1

HEIGHT m 7.57 8.3 7.98

WING SPAN m 31.23 27.43 28.35

WING AREA m2 77 56.2 79.9

ASPECT RATIO - 12.66 13.37 10.05

WING LOADING Kg/m2 66.7 88.7 72.04

EMPTY WEIGHT Kg 12,327 11,791 12,560

GROSS WEIGHT Kg 19,456 16,860 15,560



RANGE Km 1,715 2,034 1,284


SPEED Km/hr 494 435 528

SERVICE CEILING m 7,620 11,430 6,4005.9


USEFULL LOAD Kg 5,136 4,986 5,756


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TABLE-1.5

SPECIFICATIONS UNITS NAME OF THE AIRCRAFTSMARTIN 2-0-2 ANTONOV

AN-24

DHC-8-

300SERIESENGINE NAME - Pratt &Whitney R-

2800 CA-18

Ivcenko AI-24A 2PW 123B



THRUST POWER KN 27.17 29.216 26.07

THRUST LOADING - 0.648 0.871 0.516

LENGTH m 26.47 24.77 23.34

HEIGHT m 8.66 8.32 7.49

WING SPAN m 28.42 29.2 27.43

WING AREA m2 80.3 75 56.2

ASPECT RATIO - 10.05 11.36 13.37

WING LOADING Kg/m2 61.967 68.266 91.24

EMPTY WEIGHT Kg 11,379 13,300 11,791

GROSS WEIGHT Kg 18,460 21,000 17,654

MAX.TAKE OFF

WEIGHT

Kg 18,756 17,450 19,500


RANGE Km 1,022 2,761 1,558


SPEED Km/hr 311 684 765

SERVICE CEILING m 10,058 8,400 9,450

RATE OF CLIMB m/s 6.8 6 8

USEFULL LOAD Kg 4,976 5,120 5,138


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GRAPHS

THRUST vs ASPECT RATIO

THRUST vs CRUISE SPEED


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20

THRUST vs EMPTY WEIGHT

THRUST vs GROSS WEIGHT


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21

THRUST vs HEIGHT

THRUST vs LENGTH


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22

THRUST vs MAX.TAKE OFF WEIGHT

THRUST vs PROPELLER POWER


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23

THRUST vs RANGE

THRUST vs RATE OFF CLIMB


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24

THRUST vs SERVICE CEILING

THRUST vs SPEED


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THRUST vs THRUST LOADING

THRUST vs USEFUL LOAD


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26

THRUST vs WING AREA

THRUST vs WING SPAN


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THRUST vs WING LOADING


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MEAN DESIGN PARAMETER

SL.NO MEAN DESIGN PARAMETER MAGNITUDE UNIT

1 Propeller power 1450 KW

2 Thrust power 30.5 KN

3 Thrust loading 0.6 -

4 Length 24 m

5 Height 7.5 m

6 Wing span 28 m

7 Wing area 53 m

8 Aspect ratio 10 -

9 Wing loading 3946.99 kg/m

10 Wempty weight 11100 Kg

11 Gross weight 19000 Kg12 Max.Take- off weight 20000 Kg

13 Crew member 2 -

14 Range 1800 Km

15 Cruise speed 510 Km/h

16 Speed 810 Km/h

17 Service ceiling 7100 m

18 Rate of climb 7.2 m/s

19 Useful load 4900 Kg


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WEIGHT ESTIMATION

4.1 THE WEIGHT OF AN AIRCRAFT AND ITS FIRS

T ESTIMATELet us discuss the nature of the weight of an airplane in detail. There are

various types ways to subdivide and categorize the weight components of an

airplane. The following is a common choice.

1. Crew weight Wcrew. The crew comprises the people necessary to operate the

air plane in flight. For our airplane, the crew is simply the pilot.

2. Payload weight Wpayload . The payload is what the airplane is intended to

transport passenger, baggage, freight, etc. If airplane is intended for military

combat use, the payload includes bombs, rockets, and other disposable

ordnance.

3. Fuel weight Wfuel. This is the weight of the fuel in the fuel tanks. Since fuel

is consumed during the course of the flight, Wfuel is a variable, decreasing with

time during the course of the flight.

4. Empty weight Wempty. This is the weight of everything else-the structure,

engines( with all accessory), electronic equipment (including radar computers,

communication device,etc.),landing gear, fixed equipment(seats, galleys, etc.),

and anything else that is not crew, payload, or fuel.

The sum of these weights is the total weight of the airplane W. Again, W

varies throughout the fight because fuel is being consumed, and for a military

combat airplane, ordnance may be dropped or expended, leading to a decrease

in the payload weight.

The design takeoff gross weight W0 is the weight of airplane at the instant

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

beginning of the flight.

Chapter-4


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Hence,

W0 = Wcrew + Wpayload + Wfuel + Wempty [4.1]

In Eq. (4.1), Wfuel is the weight of the full fuel load at the beginning of the

flight.

In Eq. (4.1), W0 is the important quantity for which we want a first estimate; W0

is the desired result from graph. To help make this estimate, Eq. (4.1) can be

rearranged as follows. If we denote Wfuel by Wfand Wempty by We (for notational

simplicity), Eq. (4.1) can be written as

W0 = Wcrew + Wpayload + Wf+ We [4.2]

W0=Wcrew+Wpayload+ W0+ W0 [4.3]

Solving Eq. (4.3) for W0, we have

W0= [4.4]

The form of Eq. (4.4) is particularly useful. Although at this stage we donot have a value of W0, we can fairly readily obtain values of the ratios Wf/W0

and We/W0, as we will see next. Then Eq. (4.4) provides a relation from which

W0 can be obtained in an iterative fashion.[The iteration is required because in

Eq.(4.4) Wf/W0 and We/W0 may themselves be functions of W0.]

4.2 ESTIMATION OF We/W0

Most airplane design are evolutionary rather than revolutionary; that is, anew design is usually an evolutionary change from previously existing

airplanes. For this reason, historical, statistical data on previous airplanes

provides a starting point for the conceptual design of a new airplane. We will

use such data here. In particular, Graph of We/W0 versus W0 for a number of

Turbofan engine, jet aircrafts.


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As a result of the data shown in graph. we choose for our first estimate

= 0.56 [4.5]

4.3 ESTIMATION OF Wf/ W0

The amount of fuel required to carry out the mission depends critically on

theefficiency of the propulsion device-the engine specific fuel consumption and

the propeller efficiency. It also depends critically on the aerodynamic

efficiency-the lift-to-drag ratio. These factors are principal players in theBrequet range equation, represented here:

R= ln [4.6]

Equation (4.6) is very important in our estimation of Wf/W0, as defined

below. The total fuel consumed during the mission is that mission is that

consumed from the moment the engines are turned on at the airport to the

moment they are shut down at the end of the flight. Between these times, the

flight of the airplane can be described by a mission profile, a conceptual sketch

of altitude versus time such as shown in (figure 4.1).As stated in the

specifications. The mission profile is that for a simple cruise from one location

to another. This is the mission profile shown in Figure. It starts at the point

labeled0, when the engines are first turned on. The takeoff segment is denotedby the line segment0-1, which includes warm-up, taxing, and takeoff. Segment

1-2 denotes the climb to cruise altitude (the use of a straight line here is only

schematic and is not meant to imply a constant rate of climb to altitude).

Segment 2-3 represents the cruise, which is by far the largest segment of the

mission. Segment 2-3 shows an increase in altitude during cruise, consistent

with an attempt to keep CL

(and hence L/D) constant as the aircraft weightdecreases because of the consumption of fuel. Segment 3-4 denotes the descent,


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which generally includes loiter time to account for air traffic delays; for design

purposes, a loiter time of 20 min is commonly used. Segment 4-5 represents

landing .The mission profile shown in Figure is particularly simple. For other

types of missions, especially those associated with military combat aircraft, the

mission profile with include such aspects as combat dog fighting, weapons

drop, in-flight refueling etc. For a discussion of such combat mission profiles,

see, for example, Raymer book. For our purpose, we will deal only with the

simple cruise mission profile sketched in Figure (Fig.4.1)

4.4 MISSION PROFILE

The mission profile is a useful book keeping tool to help us estimate fuel

weight. Each segment of the mission profile is associated with a weight fraction,

defined as the aircraft weight at the end of the segment divided by the weight at

the beginning of the segment.

Mission segment weight fraction =

For example, the cruise weight fraction is W3/W2, where W3 is the aircraft

weight at the end of the cruise and W2 is the weight at the beginning of cruise.

The fuel weight ratio Wf/W0,can be obtained from the product of the mission

segment weight fractions as follows. Using the mission profile in Figure, the

Fig 4.1 Mission Profile


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ratio of the aircraft weight at the end of the mission to the initial gross weight is

W5/W0. In turn,

= [4.7]

SUGGESTED FUEL FRACTIONS FOR SEVERAL MISSION PHASES:

TABLE 4.1

AIRPLANE TYPE TAKE OFF CLIMB DESCENT LANDING

Business Jet 0.995 0.980 0.990 0.992Transport 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

The right side of Eq. (4.7) is simply the product of the individual mission

segment weight fractions. Also, keep in mind that for the simple cruise mission

shown in Figure, the change in weight during each segment is due to the

consumption of fuel. It, at the end of the flight, the fuel tanks were completely

empty, then

Wf= W0-W5 [4.8]

Or

=1-

However, at the end of the mission, the fuel tanks are not completely

empty-by design .There should be some fuel left in reserve at the end of the

mission in case weather conditions or traffic problems require that the pilot of

the aircraft divert to another airport, or spend a longer-than-normal time in a

holding pattern. Also, the geometric design of the fuel tanks and the fuel system

leads to some trapped fuel that is unavailable at the end of the flight. Typically,


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a 6% allowance is made for reserve and trapped fuel. Modifying Eq. (4.8)for

this allowance, we have

=1.06 [4.9]

Hence, the sequence for the calculation of Wf/W0 that appears in the

denominator of Eq. (4.9) is as follows:

a. Calculate each individual mission segment weight fraction W1/ W0, W2

etc., that appears in Eq. (4.7).

b. Calculate W5/ W0 from Eq. (4.7).

c. Calculate Wf

/ W0

from Eq. (4.9).Let us proceed to make this calculation for our transport fifty seated aircraft.

For takeoff, segment 0-1, historical data show that W1/ W0 are small, on the

order of 0.97. Hence, we assume

= 0.970 [4.10]

For climb, segment 1-2. we again rely on historical data for a first

estimate which indicate that W2/ W1 is also small, on the order of 0.985. Hence,

we assume

= 0.985 [4.11]

INITIAL ESTIMATES OF LIFT/DRAG RATIO (L/D):

TABLE 4.2

AIRCRAFTS CRUISE LOITER

Homebuilt & Single Engine 8-10 10-12

Business Jet 10-12 12-14

Regional Turboprop 11-13 14-16

Transport Jets 13-15 14-18

Military Trainers 8-10 10-14

Fighters 4-7 6-9


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Military Patrol, Bombers &

Transports

13-15 14-18

Supersonic Cruise 4-6 7-9

For cruise, segment 2-3, we make use of the Brequet range equation. This

requires an estimate of L/D. At this stage of our design process, we cannot carry

out a detailed aerodynamics analysis to predict L/D- we have not even laid out

the shape of the airplane yet. Therefore, we can only make a crude

approximation, again based on data from existing aircraft. Loft in has tabulated

the values of (L/D)max for a number of famous aircraft over the past century.

Hence, a reasonable first approximation for our aircraft is

(L/D)max =14 [4.12]

Also needed in the range equation, are the specific fuel

consumption c and velocity Vcr.

A typical value of specific fuel consumption for aircraft turbo fan engine is 0.6

lb of fuel consumed per horse power per hour. In consistent units, noting that 1

hp = 550 ft-lb/s, we have

c = 0.7 [4.13]

A reasonable value for the velocity, assuming a variable- pitch engine

Vcr = 510 km/hr [4.14]

The ratio W0/W1 in that equation is replaced for the mission segment 2-3 by

W2/W3. Hence,for range equation

R= ln [4.15]


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SPECIFIC FUEL CONSUMPTION:

TABLE 4.3

AIRCRAFTS CRUISE LOITER

Business Jets & 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

Supersonic Cruise 0.7-1.5 0.6-0.8

Solving Eq. (4.15) for W2/W3, we have

= . [4.16]

The loiter segment 3-4 in figure is essentially the descent from cruise

altitude to the landing approach. For our approximate calculation here, we will

ignore the detail of fuel consumption during descent is part of the required

3221.13-mi range, Hence, for this assumption

= 1.00 [4.17]

Finally, the fuel consumed during the landing process, segment 4-5, is

also based on historical data. The amount of fuel used for landing is small, and

based on previous aircraft, the value of W5/W4 is approximately 0.995. Hence,

we assume for our airplane that

= 0.995 [4.18]

Collecting the various segment weight fractions form Eq. (4.10), (4.11), (4.16),


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W0 =

=22933.33 Kg

We know that,

= 0.56

=12842.67 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).

4.6 ITERATION PROCESS (W0):

For the iteration process, we use the given formula,

= 1.02 0-0.06 [4.23]

FIRST:

= 1.02 25671.64-0.06

=0 .578

W0 = 22211.324 Kg

SECOND:

W0 = 22953.53 Kg

THIRD:

W0 = 22998.83Kg

FOURTH:

W0 = 23001.57Kg


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FIFTH:

W0 = 23001.71Kg

SIXTH:

W0 = 23001.74Kg

After doing sixth iterations, we can take the value W0 =23001.74 Kg as the final

estimate of the W0.

Max Takeoff Weight (W0) = 23001.74 Kg [4.24]

We know that,

= 0.215

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

Wf= 23001.74 0.215

Wf= 4945.37 K

Weight of the Fuel Wf= 4945.37 Kg [4.25]

The weight of aviation gasoline is 5.64 lb/gal. Hence, the capacity of the fuel

tank (or tanks) should be

Tank capacity = .

.

Tank capacity = 1933.2955 gal


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POWER PLANT SELECTION

From the first weight estimate, we can have rough idea of the weight ofthe power plant that is to be used.

The total weight of the power plant is found to be 0.25W.

The literature survey indicated a thrust to weight ratio of0.25 was

appropriate.

The choice of engine is a turbofan for the following reasons such as:

1) High operating fuel economy2) Efficiency for high payloads

3) Short take-off roll due to increased thrust at low speeds

Most of the aircraft in the business category were found to have 2

engines & hence the preference is towards having twin engines

Max. take off weight ,W0 = 23001.74 kg

=23001.749.81

=225.65 KN

Wpowerplant =0.25W0

=0.25225.65103

=61.62 KN [5.1]

Engines can be used in a combination of 230.8 KN

A choice of engines from different manufacturers is always the preferred

commercial position for the airframe manufacturer. This ensures that the engine

price and availability is more competitive. It also provides the potential airline

Chapter-5


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customer with more bargaining power when selecting the aircraft/engine

purchase.

There are several available engines that would suit our requirement. All of them

are currently used on civil aircraft operations therefore considerable experience

is available.

The engines below are typical options:

TABLE:5.1

SL.NO NAME OF THE ENGINE TYPE THRUST

1 Rolls-Royce AE-3007A Turbofan 31.3KN

2 Klimov TV3-117VMA-SBM1 Turboprop 27.6KN

3 Allison AE-2100A Turboprop 35.7.2KN

5.1 REQUIRED ENGINE

Calculated thrust and weight of the engine are satisfied with the General

Rolls-Royce AE-3007A therefore chosen this engine.

Rolls-Royce AE-3007A

Manufactured by Rolls-Royce in Indianapolis, Indiana the AE 3007

turbofan entered into service in 1995 as a leader in its class, meeting the

meticulous requirements of regional, corporate and military customers. With a

common core among the Rolls-Royce AE family of engines, including the AE


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2100 turboprop and the AE 1107 turboshaft, the AE 3007 allows operators to

benefit from worldwide usage, military qualifications and international civil

certification.

Safety and reliability are strong values of the AE 3007, supported by the Rolls-

Royce global customer support and maintenance network. Rolls-Royce offers

both TotalCare and CorporateCare maintenance plans for the AE 3007

family of engines, allowing worry-free management and cost predictability for

operators.


The above engine is a high by pass ratio,two-spool axial flow turbofan

engine.The mean design features include

A single stage fan

A 14-stage axial flow compressor with inlet guide vanes and five variable

geometry stator stages

A 2-stage high pressure turbine to drive the compressor

A 3-stage low pressure turbine to drive the fan.

Dual redundant ,full Authority Digital Electronic Controls

Accessory gearbox

Fig:5.1


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Air system for aircraft pressurization and engine starting

5.2 ENGINE SPECIFICATIONS

GENERAL CHARACTERISTICS

Length :306cm

Width :155cm

Diameter :0.98m

Weight :436kg

COMPONENTS

Compressor : 1LP,14HP

Turbine : 2HP,3LP

PERFORMANCE

Thrust : 28.9-42kn

Inlet mass flow : 240-280 lb/sec

Turbine inlet temperature : 9940c

Thrust to weight ratio : 4.1-5.6

Exhaust nozzle area :0.4323m2

Fan bypass : 40.8kg/min

Rotor speeds :16270 - 8700

Fuel inlet pressure :379.2kpa

Bypass ratio : 5

Pressure ratio :23


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Engine Position

Fig 5.1 Cross sectional View

Fig:5.2

Fig:5.3


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FUEL WEIGHT VALIDATION

The choice of a suitable engine, having been made, it is now possibleto 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 =T01.2

=

Altitude = 10200 m = 33465 ft

=

= .

.= 0.326 [6.1]

Cruise velocity = 510 Km/hr = 141.66m/s

T0 = 31.3 KN

= 31.30.326 .

= 8.15 KN = 831.26kg [6.2]

SFC = 0.7 (at medium thrust setting)

Number of engines = 2

Chapter-6


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6.1 CALCULATION:

Wfuel = . . .

Wfuel= 4928.87 kg [6.3]


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WING DESIGN

7.1 INTRODUCTION

After the final weight estimation of the aircraft, the primary component of

the aircraft to be designed is the wing. The wing may be considered as the most

important component of an aircraft, since a fixed-wing aircraft is not able to fly

without it. Since the wing geometry and its features are influencing all other

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

primary function of the wing is to generate sufficient lift force or simply lift (L).However, the wing has two other productions, namely drag force or drag (D)

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

maximize the lift, the other two (drag and pitching moment) must be minimized.

The wing must produce sufficient lift while generate minimum drag, and

minimum pitching moment. These design goals must be collectively satisfied

throughout all flight operations and missions.

7.2 WING GEOMETRY DESIGN

Chapter-7

Fig:7.1 Wing Geometry Design


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

Wing loading (W/S),

Sweepback angle at quarter chord (qc).

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

to find the

Aspect Ratio (b2/S),

The value of S also enables us to calculate the Taper ratio ()

Form Raymer book we choose our, Taper Ratio ) = 0.6

The root chord is given by,

Root chord (Cr) =( )

The tip chord is given by,

Tip chord (Ct) = Croot

Mean Aerodynamic Chord,

Mean chord = Croot( )

( )

Where,S = Reference wing area

C = Chord

b = Wing span

= Taper ratio

A= Aspect ratio = b2/S

Sweep back angle () can be obtained approximately using a taper ratio() of 0.6

7.2.1. WING AREA:

Wing planform area (S) =

= . .

.

= 57.16m2 [7.1]


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7.2.2. WING SPAN (b):

Aspect ratio = 10 (from the graph)

Aspect ratio =

Span (b) = (Wing planform area Aspect ratio)0.5

= (57.1610)0.5

=23.9m [7.2]

7.3 WING CHORD DESIGN

7.3.1. ROOT CHORD, Cr

The root chord is given by,

Root chord (Cr) =( )

= 2.989m [7.3]

7.3.2. TIP CHORD, Ct

Tip chord (Ct) = Croot

Tip chord (Ct) = 0.62.989

= 1.79m [7.4]

DETERMINATION OF THE MEAN AERODYNAMIC CHORD

Mean chord = Croot( )

( )

= 3.487m [7.5]

7.3.3. Distance of the Mean Chord from the Aircraft Centre line

= ( )

( )

= 5.47m [7.6]


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7.3.4. SWEEP ANGLE ():

Sweep back angle at Leading edge

=tan

= . [7.7]

7.3.5. DIHEDRAL ANGLE ( )

TABLE:7.1

From the above table the Dihederal angle of different 50 seated transport

aircraft are range between 2-50.we take our design consideration

Dihedral Angle ( ) = . [7.8]

7.4 WING VERTICAL LOCATION

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

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

centerline. This wing parameter will directly influence the design of other


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aircraft components including aircraft tail design, landing gear design, and

center of gravity. In principle, there are four options for the vertical location of

the wing.

7.4.1 Low Wing

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

configuration; due to the ground effect

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

wing.

The retraction system inside the wing is an option along with inside thefuselage

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

gear lighter and requires less space inside the wing for retraction system.

This will further make the wing structure lighter

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

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

The wing has less induced drag.

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

Fig:7.2 Low Wing Arrangement


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AIRFOIL SELECTION

8.1 INTRODUCTION: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, liftcurve 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.

Chapter-8

Fig:8.1 Airfoil Layout


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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 : 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 1930s.

NACA 5-digit is aimed at pushing position of max camber forwards for

increased

CL max.

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).

NACA 5 Digit

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


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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 highsubsonic cruise Mach numbers.

8.2 ESTIMATION OF THE CRITICAL PERFORMANCE

PARAMETERS

We now move to pivot point 3, namely, an estimation of critical

performance (CL) max, L/D, W/S, and T/W. These parameters are directed by the

requirements; that is, they will be determined by such aspects as maximum

speed, range, and ceiling, rate of climb, stalling speed, landing gear, and takeoff

distance.


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Maximum Lift Coefficient

This is the stage in the design process where we make an initial choice for

the airfoil shape for the wing. Historically, general aviation airplanes have

employed the NACA four digit, and 6-series airfoil sections-the laminar-flow

series.

L=W=0.5V2stallSCL cruise [8.1]

VStall = 0.25 Vcruise [8.2]

VStall = 0.25 141.66

VStall = 35.416 m/s [8.3]

sub, the value Eq.(7.3) in (7.1)

= 0.5 0.4 ( . ) 57.16CL cruise

CL cruise = 0.972 [8.4]

t/c CALCULATION:

= .

cos 1

( )

{ ( #) }

.

Taking # = 1.05 - 0.25 CLcruise=0.80

Where,


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M = Drag Divergence Cruise Mach Number = 0.83

= Sweep Back Angle = 2.87 at Quarter Chord

CL cruise = 0.972

Substituting the values in the above equation, we get,

= 0.12 [8.5]

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


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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

TABLE:8.1

CHORD AIRFOIL CL max

ROOT 65(2)-415 1.238

MEAN 65(1)-412 1.107

TIP 65-410 1.015

8.3 AIRFOIL GEOMETRY

Fig:8.2 Airfoil Geometry


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Fig:8.2 Airfoil Geometry


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Fig:8.3

Fig:8.4


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Performance curves for the chosen airfoil NACA 65(1)-412Fig:8.5


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CALCULATIONS:

(CL max ) = .

+ .

+ .

= 1.12

max avail = 0.9 CL max = 0.9 1.12 = 1.008 [8.6]


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FLAP SELECTION

9.1 INTRODUCTIONDuring takeoff and landing the airplane's velocity is relatively low. To

keep the lift high (to avoid objects on the ground!), airplane designers try to

increase the wing area and change the airfoil shape by putting some moving

parts on the wings' leading and trailing edges. The part on the leading edge is

called a slat, while the part on the trailing edge is called a flap. The flaps and

slats move along metal tracks built into the wings. Moving the flaps aft (toward

the tail) and the slats forward increases the wing area. Pivoting the leading edge

of the slat and the trailing edge of the flap downward increases the effective

camber of the airfoil, which increases the lift. In addition, the large aft projected

area of the flap increases the drag of the aircraft. This helps the airplane slow

down for landing.

9.2 TYPES OF FLAP

Types of flap systems include:

Krueger flap: hinged flap on the leading edge. Often called a "droop".

Plain flap: rotates on a simple hinge.

Split flap: upper and lower surfaces are separate, the lower surface

operates like a plain flap, but the upper surface stays immobile or moves

only slightly.

Gouge flap: a cylindrical or conical aerofoil section which rotates

backwards and downwards about an imaginary axis below the wing,

increasing wing area and chord without affecting trim. Invented by

Arthur Gouge for Short Brothers in 1936.

Chapter-9


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Fowler flap: slides backwards before hinging downwards, thereby

increasing both camber and chord, creating a larger wing surface better

tuned for lower speeds. It also provides some slot effect. The Fowler flap

was invented by Harlan D. Fowler .

Fairey-Youngman flap: moves body down before moving aft and

rotating.

Slotted flap: a slot (or gap) between the flap and the wing enables high

pressure air from below the wing to re-energize the boundary layer over

the flap. This helps the airflow to stay attached to the flap, delaying the

stall.

Blown flaps: systems that blow engine air over the upper surface of the

flap at certain angles to improve lift characteristics.


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Fig:9.1 Types Of Flapes


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9.3 SELECTED FLAP

A wing designed for efficient high-speed flight is often quite different

from one designed solely for take-off and landing. Take-off and landing

distances are strongly influenced by aircraft stalling speed, with lower stall

speeds requiring lower acceleration or deceleration and correspondingly shorter

field lengths. It is always possible to reduce stall speed by increasing wing area,

but it is not desirable to cruise with hundreds of square feet of extra wing area

(and the associated weight and drag), area that is only needed for a few minutes.

It is also possible to reduce stalling speed by reducing weight, increasing

air density, or increasing wing CLmax. The latter parameter is the most

interesting. One can design a wing airfoil that compromises cruise efficiency to

obtain a good CLmax, but it is usually more efficient to include movable leading

and/or trailing edges so that one may obtain good high speed performance while

achieving a high CLmax at take-off and landing. The primary goal of a high lift

system is a high CLmax; however, it may also be desirable to maintain low drag

at take-off, or high drag on approach. It is also necessary to do this with a

system that has low weight and high reliability.

Fig:9.2


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CL max INCREASES DUE TO FLAP

TABLE:9.1

Our flap is Double fowler flap the required value is at above.

TAKE-OFF CL max DUE TO FLAP

During Take-off Flap deflection up to 200

(CL max ) = 0.5 + 1.008

= 1.508 [9.1]

LANDING CL max DUE TO FLAP

During Landing Flap deflection up to 500

(CL max ) = 0.9 + 1.008

= 1.908 [9.2]


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FUSELAGE AND CABIN LAYOUT

10.1 INTRODUCTIONThe interiors of business aircraft are laid out more flexibly than are

commercial transports. Interior appointments often cost millions of dollars and

can be very luxurious, especially for the large long range aircraft such as the

Gulfstream V or Global Express. Business aircraft based on commercial

transports such as Boeing Business Jet provide even greater possibilities.

Cabine layout of of 50 seater transport aircraft

Cabin parameters obtained from similar transport aircrafts

Seat pitch = 0.9652m

Seat width = 0.7m

Aisle width=0.61m

Seats abreast=2

No. of aisles=1

Chapter-10

Fig:10.1 Cabin Layout


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10.2 FUSELAGE LAYOUT-INTRODUCTION

The fuselage layout is important as the length of the entire aircraft

depends on this. The length and diameter of the fuselage is related to the seating

arrangement. The fuselage of a passenger aircraft is divided into a number of

sections:

a. Nose

b. Cockpit

c. Cabin

d. Tail fuselage

Functions of fuselage:

provides of volume for payload

provide overall structural integrity

Possible mounting of landing gear and power plant

Once fundamental configuration is establishment, fuselage layout proceeds

almost

independent of other design aspects.

Pressurization

If required, it has a major impact upon the overall shape.

Overall effect depends on the level of pressurization.

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.


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Pressure compartment has to avoid use of flat surfaces.

Normal (High) Differential Pressurisation:

Usual requirement is for effective altitude to be no more than 11 km

(32000 ft) ISA for passenger transports.

Implied pressure differentials are:

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.

10.3 FUSELAGE SIZING:

The required value of Fuselage size is taken from the graph

LFUSELAGE = 19.5 m [10.1]

10.3.1 NOSE AND COCKPIT-FRONT FUSELAGE:

The layout of the flight deck and specified pilot window geometry is

often the starting point of the overall fuselage layout. For the current design,

flight decks of various airplanes are considered and the value of

is found to be 0.03 [10.2]

Lnos = 0.03 19.5

Lnos = 0.58 m [10.3]

The cockpit length for a 2 member crew is given by RAYMER


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Honeywells avionics suite is designed for commercial airline applications

Fig:10.2 Cockpit Layout

Fig:10.3 Honeywell's Avionic Suite


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Cockpit instument layoutt

9.4 PASSENGER CABIN LAYOUT:

Fig:10.4 Cockpit Instrument Layout

Fig:10.5 Passenger Cabin Layout


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Two major geometric parameters that specify the passenger cabin are

Cabin Diameter and Cabin Length. These are in turn decided by more specific

details like number of seats, seat width, seating arrangement (number abreast),

seat pitch, aisle width and number of aisles.

We choose a circular cross section for the fuselage. The overall size must

be kept small to reduce aircraft weight and drag, yet the resulting shape must

provide a comfortable and flexible cabin interior which will appeal to the

customer airlines. The main decision to be taken is the number of seats abreast

and the aisle arrangement. The number of seats across will fix the number of

rows in the cabin and thereby the fuselage length. Design of the cabin cross

section is further complicated by the need to provide different classes like first

class, business class, economy class etc.

10.4.1 CABIN LENGTH:

The total number of seats (50) is distributed as 4 seats abreast. Cabin

parameters are chosen based on standards of similar airplanes.

The various parameters chosen are as follows

Seat pitch =0.86m

Seat width =0.93m

Aisle width =0.43m

Seats abreast =2

No. of aisles =1

Hence, the total cabin length will be = seat pitch rows

Fig:10.6 Cabin Length


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=0.86 19 + additional space

Total cabin length =18m [10.4]

10.4.5 CABIN DIAMETER:

Using the number of seats abreast, seat width, aisle width we calculate the

internal diameter of the cabin.

dfus (internal) = 2.10m [10.5]

According to the standards prescribed by Raymer, chapter 9, the structural

thickness is given by

t = 0.02df + 1 inch [10.6]

= 0.02 2.10 + 0.0254

t = 0.067 m

Therefore the external diameter of the fuselage is obtained as

= 2.10 + 0.0672

External diameter = 2.235 m [10.7]

10.5 REAR FUSELAGE:

The rear fuselage profile is chosen to provide a smooth, low drag shape

which supports the tail surfaces. The lower side of the provide adequate

clearance for aircraft when rotation during takeoff. The rear fuselage should


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also house the auxiliary power unit (APU). Based on data collected for similar

aircraft we choose the ratio Ltail/ dfus as 4.

Ltail = 6m [10.8]

10.5.1 Total fuselage length:

Various parts of the fuselage are indicated below

Cockpit length = 3.9

Cabin length = 18m

Total = 27.93m [10.9]

Fig:10.7 Overall Layout


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TAIL SELECTION

11.1 TAIL SURFACES:The type and area of the tail surfaces are important in determining the

stability of the airplane. A conventional tail arrangement is chosen. Some of the

important parameters that decide the aerodynamic characteristics of the tail are

area ratio (St/S), tail volume ratio(VH and Vv), tail arm, tail span etc. All this

parameters have to be decided for both the horizontal and vertical tail.

From the above list of tail types, the T-tail unit type is chosen which the most

suitable configuration for the current design.

11.2. T-TAIL

A T-tail is an aft tail configuration (see figure. 34) that looks like the

letter T;which implies the vertical tail is located on top of the horizontal tail.

The T-tail

Chapter-11

Fig 11.1 Types of aircraft tail


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configuration is another aft tail configuration that provides a few advantages,

while it has a few disadvantages. The major advantage of a T-tail configuration

is that it is out of the regions of wing wake, wing downwash, wing vortices, and

engine exit flow (i.e. hot and turbulent high speed gas). This allows the

horizontal tail to provide a higher efficiency, and a safer structure. The lower

influence from the wing results in a smaller horizontal tail area; and the lower

effect from the engine leads in a less tail vibration and buffet. The less tail

vibration increases the life of the tail with a lower fatigue problem.

On the other hand, the disadvantages that associated with a T-tail are:

1. vertical tail structure,

2. deep stall.

The bending moment created by the horizontal tail must be transferred to the

fuselage through the vertical tail. This structural behavior requires the vertical

tail main spar to be stronger; which cause the vertical tail to get heavier.

Aircraft with T-tail are subject to a dangerous condition known as the

deep stall (Ref. 6); which is a stalled condition at an angle of attack far abovethe original stall angle.T-tail Aircraft often suffer a sever pitching moment

instability at angles well above the initial stall angle of about 13 degrees,

without wing leading edge high lift device, or about 18 degrees, with wing

leading edge high lift device. If the pilot allows the aircraft to enter to this

unstable region, it might rapidly pitch up to a higher angle of about 40 degrees.

Fig 11.2 Stability due to Horizontal Tail


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11.3.1 TAIL AREA:

The areas of the horizontal and vertical tail (SH and Sv) are calculated as,

SH = 0.31 57.16

SH = 17.71 m2

[11.1]

Sv = 0.21 57.16

SV = 12 m2

[11.2]

11.3.2 TAIL SPAN:

The span of the horizontal and vertical tail (bh and bv) are given as,

bh = (AhSH)0.5

[11.3]

bv= ((AhSB))0.5

[11.4]

Taking ARH = 5 and ARV = 1.7, we get

bh = 9.4 m [11.5]

bv = 4.5 m [11.6]

Fig 11.3 Tail Section


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CENTRE OF GRAVITY

The center-of-gravity (CG) is the point at which an aircraft wouldbalance if it were possible to suspend it at that point. It is the mass center of the

aircraft, or the theoretical point at which the entire weight of the aircraft is

assumed to be concentrated. Its distance from the reference datum is determined

by dividing the total moment by the total weight of the aircraft. The center-of-

gravity point affects the stability of the aircraft. To ensure the aircraft is safe to

fly, the center-of gravity must fall within specified limits.

12.1 CENTER OF GRAVITY IS CALCULATED AS FOLLOWS:

Determine the weights and arms of all mass within the aircraft.

Multiply weights by arms for all mass to calculate moments.

Add the moments of all mass together.

Divide the total moment by the total weight of the aircraft to give an

overall arm.

Chapter-12

Fig:12.1 Center Of Gravity Indication


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The arm that results from this calculation must be within the arm limits for

the center of gravity. If it is not, weight in the aircraft must be removed, added

(rarely), or redistributed until the center of gravity falls within the prescribed

limits.

For the sake of simplicity, centre of gravity calculations are usually

performed along only a single line from the zero point of the reference datum.

Weight is calculated simply by adding up all weight in the aircraft. This

weight must be within the allowable weight limits for the aircraft.

First estimate weight components for which we have some idea of their

location of the engine, the passengers and pilot, and the baggage.

Considering the forces to be acting at middle each part, and hence taking

moment about the nose, we get the centre of gravity.

CG =

( )+( )+ ( ) + ( )

( )

= 14.4 m [12.1]

12.2 Layout

Fig:12.2 Center Of Gravity Layout


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LANDING GEAR SELECTION

13.1 INTRODUCTIONEvery aircraft maintained in todays Aerospace Company is equipped

with a landing gear system. Most Aerospace company aircraft also use arresting

and catapult gear. The landing gear is that portion of the aircraft that supports

the weight of the aircraft while it is on the ground. The landing gear contains

components that are necessary for taking off and landing the aircraft safely.

Some of these components are landing gear struts that absorb landing and

taxiing shocks; brakes that are used to stop and, in some cases, steer the aircraft;

nose wheel steering for steering the aircraft; and in some cases, nose catapult

components that provide the aircraft with carrier deck takeoff capabilities.

13.2 LANDING GEAR DESIGN REQUIREMENTS

The following design requirements are identified to be satisfied: ground

clearance requirement, tip-back (or tip-forward angle if tail gear) angle

requirement, take-off rotation requirement, overturn angel requirement,

structural integrity, aircraft ground stability, aircraft ground controllability, low

cost, maintainable, and manufacturable.

Chapter-13

Fig:13.1 Main Landind Gear Assembly


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11.3 LANDING GEAR CONFIGURATION

This is a transport aircraft, and the passengers comfort is an important

requirement. So, the tail gear, bicycle, single main configurations would not

satisfy this requirement.

Three viable configurations are:

1. Tricycle or nose-gear,

2. Quadricycle, and

3. Multi-bogey.

4. Ski type gear

5. Float type gear

Since the aircraft weight is not very high, both quadricycle, and multi-

bogey configurations are set aside due to their cost and weight. Therefore the

best landing gear configuration for this aircraft is Nose gear or tricycle. An

attractive feature for this configuration is that the aircraft will be horizontal at

the ground. The passengers do not have to climb during boarding period. The

Fig:13.2 Types Of Landing Gear


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nose gear also decreases the take-off run, and at the same time, the aircraft will

take-off sooner.

13.3 FIXED OR RETRACTABLE

The aircraft must compete with other transport aircraft in the market, and

it must have a fairly high performance, so a retractable landing gear (see figure)

is the best option. The cost of this configuration covered by the customers

(passengers). Then, this will reduce the aircraft drag during flight and therefore

the aircraft will feature a higher performance. The higher landing gear weight

due to retraction system will be paid off compared with the other advantages of

a retractable landing gear

Fig:13.3 Main Landing Gear In Aircraft


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Main landing gear deployed

Nose landing gear deployed

13.4.1 STEERING OF LANDING GEAR

The steering mechanism used on the ground with wheeled landing gear varies

by aircraft, but there are several types of steering.

Fig:13.4 Landing Gear Marking

Fig:13.5 Nose Landing Gear Deployed


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RUDDER STEERING

DIRECT STEERING

TILLER STEERING

Maximum Takeoff Weight of the aircraft (from Weight Estimation)

= 23001.74Kg

13.5 TYRE SIZING

During landing and takeoff, the undercarriage supports the total weight of

the airplane. Undercarriage is of three types

Bicycle type

Tricycle type

Tricycle tail wheel type

13.6 LANDING GEAR HEIGHT

The aircraft cg is at the same height as the wing mid-plane. The landing

gear height is a function of its attachment location. The nose gear will be

naturally attached to the fuselage. But, the main gear attachment tends to have

two main alternatives: 1. Attach to the fuselage, 2. Attach to the wing. As soon

the wheel track is determined, we are able to decide about landing gear

attachment; and then the landing gear height may be determined.

13.6.1ATTACH MAIN GEAR TO THE FUSELAGE:

HLG = Haircraft( Dfuse +H tail ) [13.1]

apply eq.(14.2) and (9.5) in (14.3)

= 6.5(2.28+2.64)

HLG = 1.581 m [13.2]


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13.7 LANDING GEAR ATTACHMENT

As a natural selection, the nose gear is attached to the fuselage nose.

However, for the main gear, we need to compare the fuselage diameter with the

wheel track. It is observed that the fuselage diameter (2.78m) is smaller than the

wheel track (29.22 m). Hence, the main gear cannot be attached to the fuselage.

Thus, main gear may be either attached directly to the wing; or attached under

the nacelle. In order to determine the best location, several design requirements

must be examined, which is beyond the scope of this example. For the time

being, it is decided to attach the landing gear to the wing. Thus, the landing gear

height

will be:

HLG = 1.581m [13.3]

Tyre sizes 309.5-14(main) ,19.56 .75-8(nose)

Tyre pressure 8.60-9.00 bars

Minimum ground turning radius nose wheel 12.51m ,Minimum turning circle29.22m

(The above measurements are collected from similar aircraft with given

landing gear)

15


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14.1 LIFT:Component of aero

flight direction

14.2. Lift Coefficient (C

Amount of lift gen

Planform area (S),

L =

CL is a measure of

Section shape, plan

Effect (mach numb

14.3 GENERATION O

87

LIFT ESTIMATION

dynamic force generated on aircraft p

)

rated depends on:

air density (p), flight speed (V), lift co

2SCL

lifting effectiveness and mainly depen

form geometry, angle of attack (), c

er), viscous effects (Reynolds number

LIFT

rpendicular to

fficient (CL)

[14.1]

ds upon:

mpressibility

Chapter-14

Fig:14.1 Generation Of Lift


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Aerodynamic force a rises from two natural sources:

Variable pressure distribution.

Shear stress distribution.

Shear stress primarily contributes to overalldrag force on aircraft.

Lift mainly due topressure 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 eithercamberedorinclinedto

flow direction.

Classicalaerofoil section is optimum for high subsonic lift/drag ratio.

Fig:14.2 Aerodynamic Forces Due To Lift


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Pressure variations with angle of attack

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

positive lift at =00

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.

Fig:14.3 Pressure Variation

Fig:14.4 Lift At Different Angles

8/11/2

aircraft design project-1(50 seated aircraft)

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