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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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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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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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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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
NO.OF.ENGINES - 2 2 2
PROPELLER POWER KW 1,864 1,425 1,945
THRUST POWER KN 29.6 26.79 30.46
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
MAX.TAKE OFF WEIGHT Kg 20,820 19,818 20,600
CREW MEMBERS - 2 2 2
RANGE Km 2,055 2,632 2,445
CRUISE SPEED Km/hr 530 435 740
SPEED Km/hr 560 654 833
SERVICE CEILING m 7,620 8,140 11,277.60
RATE OF CLIMB m/s 6.43 7.9 9.12
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
NO.OF.ENGINES - 2 2 2
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
CREW MEMBERS - 2 2 2
RANGE Km 2,445 2,185 4,355
CRUISE SPEED Km/hr 740 682 450
SPEED Km/hr 833 594 514
SERVICE CEILING m 9448.8 7,620 7,620
RATE OF CLIMB m/s 6.96 7.8 6.15
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
NO.OF.ENGINES - 2 2 2
PROPELLER POWER KW 2,120 1,756 1,340
THRUST POWER KN 33.8 28.8 24.53
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
MAX.TAKE OFF WEIGHT Kg 21,092 20,234 16,765
CREW MEMBERS - 2 2 2
RANGE Km 1,715 2,034 1,284
CRUISE SPEED Km/hr 452 528 458
SPEED Km/hr 494 435 528
SERVICE CEILING m 7,620 11,430 6,4005.9
RATE OF CLIMB m/s 5.9 6.81 6.2
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
NO.OF.ENGINES - 2 2 2
PROPELLER POWER KW 1,682 1,902 1,468
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
CREW MEMBERS - 2 2 2
RANGE Km 1,022 2,761 1,558
CRUISE SPEED Km/hr 286 450 528
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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THRUST vs EMPTY WEIGHT
THRUST vs GROSS WEIGHT
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THRUST vs HEIGHT
THRUST vs LENGTH
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THRUST vs MAX.TAKE OFF WEIGHT
THRUST vs PROPELLER POWER
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THRUST vs RANGE
THRUST vs RATE OFF CLIMB
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THRUST vs SERVICE CEILING
THRUST vs SPEED
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THRUST vs THRUST LOADING
THRUST vs USEFUL LOAD
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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 de- sign 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.
Rolls-Royce AE-3007A
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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Rolls-Royce AE-3007A
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).
0.58 bar (8.5 psi) for aircraft at 13.1 km (43,000 ft).
0.65 bar (9.4 psi) for aircraft at 19.8 km (65,000 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
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