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AIRCRAFT DESIGN PROJECT - I
Heavy Business Jet
A PROJECT REPORT
Submitted by
VIGNESH. M
VINCENT KEVIN MORRIS
ARAVIND. C
in partial fulfillment for the award of the degree
of
BACHELOR OF ENGINEERING
in
AERONAUTICAL ENGINEERING
RAJALAKSHMI ENGINEERING COLLEGE, THANDALAM
ANNA UNIVERSITY: CHENNAI 600 025
APRIL 2014
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ANNA UNIVERSITY: CHENNAI 600 025
BONAFIDE CERTIFICATE
Certified that this project report “DESIGN OF HEAVY BUSINESS JET” is
the bonafide work of VIGNESH. M (211611101053), VINCENT KEVIN
MORRIS (211611101056) and ARAVIND. C () who carried out the project
work under my supervision.
SIGNATURE SIGNATURE
Mr. Yogesh Kumar Sinha Mr. Surendra Bogadi
HEAD OF THE DEPARTMENT SUPERVISOR Assistant Professor
Aeronautical Engineering Aeronautical Engineering
Rajalakshmi Engineering College, Rajalakshmi Engineering College,
Thandalam, Chennai – 602 105 Thandalam, Chennai- 602 105
Internal Examiner External Examiner
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ACKNOWLEDGEMENTS
We owe a debt of gratitude to Mr. Yogesh Kumar Sinha, Head of the
Department, Department of Aeronautical Engineering, for being a source of
constant encouragement and a pillar of support in all that we do, be it academic
or extracurricular.
We would like to extend our heartfelt thanks to Mr. Surendra Bogadi for his
constant help, erudite guidance and immense passion which enthused us to do
the project better.
A warm token of appreciation to the management at Rajalakshmi Engineering
College, Thandalam for providing us with the amenities and a congenial
atmosphere to work in.
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ABSTRACT
The aim of this project is to design and conceptualize a heavy
corporate/business jet that can cater to a wide range of clientele ranging from
business conglomerates to private organizations and individual parties. Business
jet, private jet or, colloquially bizjet is a term describing a jet aircraft, usually of
smaller size, designed for transporting groups of business people or wealthy
individuals. The project involves the design of a heavy business jet that can
accommodate about 40 passengers at full seating layout, providing the
amenities and level of comfort that a business jet is expected to provide while
incorporating the design specifications and performance parameters of a long
range commercial airliner. The aircraft allows for long range transport with
better efficiency and reduced fuel consumption and noise levels owing to a state
of the art engine and design features.
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TABLE OF CONTENTS
CHAPTER NO. TITLE PAGE NO.
ABSTRACT iv
LIST OF TABLES vii
LIST OF FIGURES viii
LIST OF SYMBOLS AND x
ABBREVIATIONS
1. INTRODUCTION TO DESIGN 1
1.1 Defining a new design 5
1.1.1 Aircraft Purpose 6
1.2 Design Motivation 7
1.3 Design Process 9
1.4 Conceptual Design 11
1.5 Design Process Breakdown 12
2. INTRODUCTION TO BUSINESS JETS 13
2.1 Classification of Business Jets 14
2.2 Need for Business Jets 16
3. COMMON COMPARATIVE STUDY 19
4. COMPARATIVE DATA SHEET 20
5. COMPARATIVE GRAPHS 22
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6. WEIGHT ESTIMATION 32
7. WING LOADING 38
8. AIRFOIL SELECTION 42
9. DRAG ESTIMATION 54
10. POWERPLANT SELECTION 58
11. LANDING GEAR DESIGN 64
12. PERFORMANCE CHARACTERISTICS 68
13. CENTRE OF GRAVITY ESTIMATION 73
14. STABILITY AND CONTROL 83
15. 3-VIEW DIAGRAM 92
16. FINALIZED DESIGN PARAMETERS 94
17. CONCLUSION 95
18. REFERENCES 96
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List of Tables Page no.
1. Design Process breakdown 12
2. Common Comparative study 19
3. Comparative Data Sheet (Business Jets) 20
4. Gross Weight Iteration Table 36
5. Wing Parameters 53
6. Comparison of Turbofan Engines 58
7. Performance Parameters 72
8. Approximate Group Weights Method 75
9. Wing Location and c.g. of the Airplane 81
10. Finalized Design Parameters 94
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List of Figures Page no.
1. Design Process flow chart 10
2. Cruise Speed vs. Range 22
3. Cruise speed vs. Altitude 23
4. Cruise Speed vs. Wing Loading 24
5. Cruise Speed vs. Gross Weight 25
6. Cruise Speed vs. Aspect Ratio 26
7. Range vs. Aspect ratio 27
8. Wing loading vs. Aspect ratio 28
9. Wing Loading vs. Takeoff run 29
10. Wing Loading vs. R/Cmax 30
11. Aspect ratio vs. R/Cmax 31
12. Mission Profile 32
13. Weight Distribution Chart 37
14. Wing loading 41
15. Airfoil Geometry 42
16. Angle of Attack 42
17. NACA 63A-514 (Root airfoil) 45
18. NACA 63-512 (Midspan airfoil) 46
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19. NACA 63-310 (Tip airfoil) 47
20. Maximum Thickness 48
21. Taper Ratio 49
22. Wing Sweep 50
23. Effective Mach no. 51
24. Winglets 52
25. Pratt & Whitney PW1000G 60
26. Landing Gear System 65
27. Performance Characteristics
a. climbing flight 69
b. gliding flight 71
28. Static Longitudinal Stability 82
29. Longitudinal stability 84
30. Fuselage directional stability coefficient 87
31. Directional stability 89
32. Lateral stability 91
33. 3 View Diagram 92
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List of Symbols and Abbreviations
- Angle of attack
- Climb angle
- Density factor
- Density of air
- Dihedral angle
- Glide angle
- Turn angle
- Turn rate
- Wing thickness ratio correction factor
- Yaw angle
1/4 - Quarter chord sweep angle
Cm/e - Elevator control power
Cn/r - Rudder control power
fuel - Density of fuel
(L/D)cruise - Lift-to-drag ratio at cruise
(L/D)loiter - Lift-to-drag ratio at loiter
ac - Aerodynamic centre
APU - Auxiliary Power Unit
AR - Wing aspect ratio
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at - Lift curve slope of tail
av - Lift curve slope of vertical tail
aw - Lift curve slope of wing
b - Wing span
c - Chord length
ĉ - Mean chord
c.g. - Centre of gravity
CAEP - Committee of Aviation Environmental Protection
CD - Drag coefficient
CD0 - Zero lift drag co-efficient
Cfe - Skin friction coefficient
Cl - Rolling moment coefficient
Clf - Function of airfoil chord over which the flow in laminar
CLmax - Maximum Lift coefficient
Cm - Pitching moment coefficient
Cn - Yawing moment coefficient
cR - Root chord
cT - Tip chord
D - Drag force
d - Tire diameter
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E – Endurance
e - Oswald efficiency factor
g - Acceleration due to gravity
Ktf - Factor allowed for trapped fuel
L - Lift force
LE - Leading edge of wing
lf - Length of fuselage
Lt - Load on tyre
lv - Aerodynamic centre of vertical tail to the airplane’s centre of gravity
M - Mach number
MTOW - Maximum Takeoff Weight
N0 - Neutral point
Ne - Number of engines located on top surface of wing
q - Dynamic pressure
R - Turn radius
R/C - Rate of climb
Rr - Rolling radius of tyre
S - Wing area
SFC - Specific Fuel Consumption
SLO - Takeoff run distance
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Sref - Wing reference area
STD - Landing run distance
Swet - Wing wetted area
T - Thrust force
t/c - Wing thickness ratio
T/W - Thrust loading
Tf - A factor which is unity for streamlined shape
V - Velocity of air/aircraft
Vcruise - Velocity at cruise
Vf - Volume of fuel
Vstall - Velocity at stall
w - Tyre width
W/S - Wing loading
W0 - Gross weight of aircraft
Wcrew - Crew weight
We - Empty weight of aircraft
Wf - Weight of fuel
Wpayload – Aircraft payload weight
xlew - Distance of location of wing from nose of the aircraft
λ - Taper ratio of wing
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1. INTRODUCTION TO DESIGN
Modern aircraft are a complex combination of aerodynamic performance,
lightweight durable structures and advanced systems engineering. Air
passengers demand more comfort and more environmentally friendly aircraft.
Hence many technical challenges need to be balanced for an aircraft to
economically achieve its design specification. Aircraft design is a complex and
laborious undertaking with a number of factors and details that are required to
be checked to obtain optimum the final envisioned product. The design process
begins from scratch and involves a number of calculations, logistic planning,
design and real world considerations, and a level head to meet any hurdle head
on.
Every airplane goes through many changes in design before it is finally built in
a factory. These steps between the first ideas for an airplane and the time when
it is actually flown make up the design process. Along the way, engineers think
about four main areas of aeronautics: Aerodynamics, Propulsion, Structures and
Materials, and Stability and Control.
Aerodynamics is the study of how air flows around an airplane. In order for an
airplane to fly at all, air must flow over and under its wings. The more
aerodynamic, or streamlined the airplane is, the less resistance it has against the
air. If air can move around the airplane easier, the airplane's engines have less
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work to do. This means the engines do not have to be as big or eat up as much
fuel which makes the airplane more lightweight and easier to fly. Engineers
have to think about what type of airplane they are designing because certain
airplanes need to be aerodynamic in certain ways. For example, fighter jets
maneuver and turn quickly and fly faster than sound (supersonic flight) over
short distances. Most passenger airplanes, on the other hand, fly below the
speed of sound (subsonic flight) for long periods of time.
Propulsion is the study of what kind of engine and power an airplane needs. An
airplane needs to have the right kind of engine for the kind of job that it has. A
passenger jet carries many passengers and a lot of heavy cargo over long
distances so its engines need to use fuel very efficiently. Engineers are also
trying to make airplane engines quieter so they do not bother the passengers
onboard or the neighborhoods they are flying over. Another important concern
is making the exhaust cleaner and more environmentally friendly. Just like
automobiles, airplane exhaust contains chemicals that can damage the earth's
environment.
Structures and Materials is the study of how strong the airplane is and what
materials will be used to build it. It is really important for an airplane to be as
lightweight as possible. The less weight an airplane has, the less work the
engines have to do and the farther it can fly. It is tough designing an airplane
that is lightweight and strong at the same time. In the past, airplanes were
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usually made out of lightweight metals like aluminum, but today a lot of
engineers are thinking about using composites in their designs. Composites look
and feel like plastic, but are stronger than most metals. Engineers also need to
make sure that airplanes not only fly well, but are also easy to build and
maintain.
Stability and Control is the study of how an airplane handles and interacts to
pilot input and feed. Pilots in the cockpit have a lot of data to read from the
airplane's computers or displays. Some of this information could include the
airplane's speed, altitude, direction, and fuel levels as well as upcoming weather
conditions and other instructions from ground control. The pilot needs to be
able to process the correct data quickly, to think about what kind of action needs
to be taken, and to react in an appropriate way. Meanwhile, the airplane should
display information to the pilot in an easy-to-read and easy-to-understand way.
The controls in the cockpit should be within easy reach and just where the pilot
expects them to be. It is also important that the airplane responds quickly and
accurately to the pilot's instructions and maneuvers.
“A beautiful aircraft is the expression of the genius of a great engineer who is
also a great artist.”
Neville Shute,
British Aeronautical Engineer and Novelist,
From, No Highway, 1947.
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When you look at aircraft, it is easy to observe that they have a number of
common features: wings, a tail with vertical and horizontal wing sections,
engines to propel them through the air, and a fuselage to carry passengers or
cargo. If, however, you take a more critical look beyond the gross features, you
also can see subtle, and sometimes not so subtle, differences. This is where
design comes into play. Each and every aircraft is built for a specific task, and
the design is worked around the requirement and need of the aircraft. The
design is modeled about the aircraft role and type and not the other way around.
Thus, this is why airplanes differ from each other and are conceptualized
differently. Aircrafts that fall in the same category may have similar
specifications and performance parameters, albeit with a few design changes.
Design is a pivotal part of any operation. Without a fixed idea or knowledge of
required aircraft, it is not possible to conceive the end product. Airplane design
is both an art and a science. In that respect it is difficult to learn by reading a
book; rather, it must be experienced and practiced. However, we can offer the
following definition and then attempt to explain it. Airplane design is the
intellectual engineering process of creating on paper (or on a computer screen) a
flying machine to (1) meet certain specifications and requirements established
by potential users (or as perceived by the manufacturer) and/or (2) pioneer
innovative, new ideas and technology. An example of the former is the design
of most commercial transports, starting at least with the Douglas DC-1 in 1932,
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which was designed to meet or exceed various specifications by an airplane
company. (The airline was TWA, named Transcontinental and Western Air at
that time.) An example of the latter is the design of the rocket-powered Bell X-
1, the first airplane to exceed the speed of sound in level or climbing flight
(October 14, 1947). The design process is indeed an intellectual activity, but a
rather special one that is tempered by good intuition developed via experience,
by attention paid to successful airplane designs that have been used in the past,
and by (generally proprietary) design procedures and databases (handbooks, etc)
that are a part of every airplane manufacturer.
1.1 Defining a new design
The design of an aircraft draws on a number of basic areas of aerospace
engineering. These include aerodynamics, propulsion, light-weight structures
and control. Each of these areas involves parameters that govern the size, shape,
weight and performance of an aircraft. Although we generally try to seek
optimum in all these aspects, with an aircraft, this is practically impossible to
achieve. The reason is that in many cases, optimizing one characteristic
degrades another.
There are many performance aspects that can be specified by the mission
requirements. These include:
The aircraft purpose or mission profile
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The type(s) and amount of payload
The cruise and maximum speeds
The normal cruise altitude
The range or radius with normal payload
The endurance
The take-off distance at the maximum weight
The purchase cost
1.1.1 Aircraft Purpose
The starting point of any new aircraft is to clearly identify its purpose. With
this, it is often possible to place a design into a general category. Such
categories include combat aircraft, passenger or cargo transports, and general
aviation aircraft. These may also be further refined into subcategories based on
particular design objectives such as range (short or long), take-off or landing
distances, maximum speed, etc. The process of categorizing is useful in
identifying any existing aircraft that might be used in making comparisons to a
proposed design. With modern military aircraft, the purpose for a new aircraft
generally comes from a military program office. For example, the mission
specifications for the X-29 pictured in figure 1.1 came from a 1977 request for
proposals from the U.S. Air Force Flight Dynamics Laboratory in which they
were seeking a research aircraft that would explore the forward swept wing
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concept and validate studies that indicated such a design could provide better
control and lift qualities in extreme maneuvers. With modern commercial
aircraft, a proposal for a new design usually comes as the response to internal
studies that aim to project future market needs. For example, the specifications
for the Boeing commercial aircraft (B-777) were based on the interest of
commercial airlines to have a twin-engine aircraft with a payload and range in
between those of the existing B-767 and B-747 aircraft. Since it is not usually
possible to optimize all of the performance aspects in an aircraft, defining the
purpose leads the way in setting which of these aspects will be the “design
drivers.” For example, with the B-777, two of the prominent design drivers
were range and payload.
1.2 Design Motivation
Fundamentally, an aircraft is a structure. Aircraft designers design structures.
The structures are shaped to give them desired aerodynamic characteristics, and
the materials and structures of their engines are chosen and shaped so they can
provide needed thrust. Even seats, control sticks, and windows are structures,
all of which must be designed for optimum performance. Designing aircraft
structures is particularly challenging, because their weight must be kept to a
minimum. There is always a tradeoff between structural strength and weight.
A good aircraft structure is one which provides all the strength and rigidity to
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allow the aircraft to meet all its design requirements, but which weighs no more
than necessary. Any excess structural weight often makes the aircraft cost more
to build and almost always makes it cost more to operate. As with small
excesses of aircraft drag, a small percentage of total aircraft weight used for
structure instead of payload can make the difference between a profitable
airliner or successful tactical fighter and a failure. Designing aircraft structures
involves determining the loads on the structure, planning the general shape and
layout, choosing materials, and then shaping, sizing and optimizing its many
components to give every part just enough strength without excess weight.
Since aircraft structures have relatively low densities, much of their interiors are
typically empty space which in the complete aircraft is filled with equipment,
payload, and fuel. Careful layout of the aircraft structure ensures structural
components are placed within the interior of the structure so they carry the
required loads efficiently and do not interfere with placement of other
components and payload within the space. Choice of materials for the structure
can profoundly influence weight, cost, and manufacturing difficulty. The
extreme complexity of modern aircraft structures makes optimal sizing of
individual components particularly challenging. An understanding of basic
structural concepts and techniques for designing efficient structures is essential
to every aircraft designer
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1.3 Design Process
The process of designing an aircraft and taking it to the point of a flight test
article consists of a sequence of steps, as illustrated in the figure. It starts by
identifying a need or capability for a new aircraft that is brought about by (1) a
perceived market potential and (2) technological advances made through
research and development. The former will include a market-share forecast,
which attempts to examine factors that might impact future sales of a new
design. These factors include the need for a new design of a specific size and
performance, the number of competing designs, and the commonality of
features with existing aircraft. As a rule, a new design with competitive
performance and cost will have an equal share of new sales with existing
competitors. The needs and capabilities of a new aircraft that are determined in
a market survey go to define the mission requirements for a conceptual aircraft.
These are compiled in the form of a design proposal that includes (1) the
motivation for initiating a new design and (2) the “technology readiness” of new
technology for incorporation into a new design. It is essential that the mission
requirements be defined before the design can be started. Based on these, the
most important performance aspects or “design drivers” can be identified and
optimized above all others. Following the design proposal, the next step is to
produce a conceptual design. The conceptual design develops the first general
size and configuration for a new aircraft. It involves the estimates of the weights
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and the choice of aerodynamic characteristics that will be best suited to the
mission requirements stated in the design proposal.
No
Requirements Satisfied
Yes
Stop
Final Evaluation
Go
Design Process flow chart
Research, Development and Market Analysis
Mission Requirements
Conceptual Design
Preliminary Design
Detailed Design
Test Article Fabrication
Flight Test
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The conceptual design is driven by the mission requirements, which are set in
the design proposal. In some cases, these may not be attainable so that the
requirement may need to be relaxed in one or more areas. This is shown in the
iterative loop in the flow chart. When the mission requirements are satisfied, the
design moves to the next phase, which is the preliminary design.
1.4 Conceptual Design
This article deals with the steps involved in the conceptual design of an aircraft.
It is broken down in to several elements, which are followed in order. These
consist of:
1. Literature survey
2. Preliminary data acquisition
3. Estimation of aircraft weight
a. Maximum take-off weight
b. Empty weight of the aircraft
c. Weight of the fuel
d. Fuel tank capacity
4. Estimation of critical performance parameters
a. Wing area
b. Lift and drag coefficients
c. Wing loading
d. Power loading
e. Thrust to weight ratio
5. Engine selection
6. Performance curves
7. 3 View diagrams
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1.5 Design Process Breakdown
• Conceptual Design:
- Competing concepts evaluated What drives the design?
- Performance goals established Will it work/meet requirement?
- Preferred concept selected What does it look like?
• Preliminary Design:
- Refined sizing of preferred concept Do serious wind tunnel tests
tests
- Design examined data/establish parameters Make actual cost estimate
- Some changes allowed
• Detail Design:
- Final detail design Certification process
- Drawings released Component/systems tests
- Detailed performance Manufacturing
- Only “tweaking” of design allowed Flight control system design
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2. INTRODUCTION TO BUSINESS JETS
A business jet is a jet that is owned by a private company or individual that is
used primarily for transporting the people who own the aircraft. That being said
a lot of planes that were developed to be used as business jets are also used for
other purposes. In addition there are also companies that are set up exclusively
to operate business jets. Therefore the lines between a business aircraft and a
commercial one have become somewhat blurred.
Over the last few years business jets have become a very popular way to travel.
They offer great comfort of travel and service, with the option of having the
aircraft at your beck and call whenever you require it. A private business jet
trumps regular commercial transport in a number of areas. Nowadays,
organizations and individuals who can afford the heavy expenses that a private
jet entails are willing to invest in one. Greater ease of travel, ease of access,
faster and hassle free transit and high comfort levels are some of the advantages
of business jet transport.
In most cases a business jet will be quite a bit smaller than a commercial jet.
The most common ones carry fewer than twenty passengers since this allows
them to operate under a different set of rules from the ones that are required for
airliners. There are however now quite a few business jets that are the size of
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airliners and in many cases they are airliners that have been adapted for the
purpose. Nevertheless most business jets are quite small and only carry a small
number of people.
Business jets have a much more luxurious interior, with a number of amenities
and services that a normal airliner would not have. Airliners are designed to
carry large numbers of people, most of who are looking for the lowest cost
possible. Business jets on the other hand are designed to carry people in a much
higher level of comfort. The people who travel by business jet are almost
always quite well off and expect this level of comfort when they travel.
2.1 Classification of Business Jets
The business jet industry groups these jets into four loosely-defined classes
Mid-sized jets
Combining flight distance, speed and comfort, these mid-sized jets are ideal for
intimate trips.
Number of Passengers: 8 - 10
Sample Aircraft:
Gulfstream 200, Embraer Legacy 450, Cessna Citation X, Bombardier
Challenger 605
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Large-cabin jets
These aircraft are fast, comfortable, and can accommodate a medium-sized
group.
Number of Passengers: 8 - 15
Sample Aircraft:
Gulfstream 550, Embraer Legacy 650, Dassault Falcon 7X
Light jets
Light jets have been a staple of the business jet industry since the advent of
the Learjet 23 in the early 1960s. They provide access to small airports and the
speed to be an effective air travel tool.
Number of Passengers: 3 – 10
Sample Aircraft:
Learjet 40, Cessna Citation CJ1, Dassault Falcon 10, Beechcraft Premier I
VIP business jets / Heavy airliners
With a variety of potential configurations, jets in this category have the capacity
for dining rooms, bedrooms and offices.
Number of Passengers: 18 - 40/50 – 250
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These heavy airliners are an ideal choice for larger groups, corporate meetings
and special events.
Sample Aircraft:
Boeing BBJ, Airbus AGJ, Embraer Lineage 1000
2.2 Need for Business Jets
The following list details some of the primary reasons companies utilize
business aviation as a solution to some of their transportation challenges:
Accessing communities with little or no airline service
Business aviation serves ten times the number of communities (more than 5,000
airports) served by commercial airlines (about 500 airports). This means
business aviation can allow companies to locate plants or facilities in small
towns or rural communities with little or no commercial airline service. With
nearly 100 communities having lost airline service, this is important.
Reaching multiple destinations quickly and efficiently.
Companies that need to reach multiple destinations in a single day may elect to
use business aviation because that type of mission could be hard or impossible
to complete with other modes of transportation.
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Supporting the travel needs of many types of company employees.
An NBAA survey revealed that 72 percent of passengers aboard business
airplanes are non-executive employees. Companies often send teams of
employees to a given destination because it is the most cost-effective means of
transport.
Moving equipment.
When companies need to immediately move sensitive or critical equipment,
business aviation is often the best solution.
Ensuring flexibility.
Businesses don’t always know in advance where or when opportunities will
present themselves. In today’s business environment, companies need to be
nimble enough to move quickly. Business aviation provides flexibility for
companies that need to ensure employees can respond to changing demands and
circumstances
Increasing employee productivity and providing security.
Business aviation is a productivity tool – when traveling aboard business
aircraft, employees can meet, plan and work en route. Business aviation also
allows employees to discuss proprietary information in a secure environment
without fear of eavesdropping, industrial espionage or physical threat.
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Keeping in contact.
Many aircraft have technologies that allow employees to remain in
communication throughout the duration of their flight. This can be critical for
companies managing a rapidly changing situation.
Providing a return to shareholders.
Studies have found that businesses which use business aviation as a solution to
some of their transportation challenges return more to shareholders than
companies in the same industry that do not utilize business aviation.
Schedule Predictability.
More than 3 percent of all commercial airline flights are cancelled. Nearly one
quarter are delayed. Today, because of record load factors on commercial
airlines, if your flight is cancelled or a delay causes you to miss your
connection, the odds of you getting on the next flight are significantly reduced.
When the future of a company and its employees is dependent upon you
arriving on time, business aviation is an important tool.
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3. COMMON COMPARATIVE STUDY
Parameters Lancair IV
Boeing 777 Antonov
An-70
Gulfstream
G550
F-16
Dimensions
Length 7.62 m 63.7 m 40.7 m 29.39 m 15.06 m
Height 2.44 m 18.51 m 16.38 m 7.87 m 4.88 m
Wing span 9.93 m 60.93 m 44.06 m 28.50 m 9.96 m
Aspect ratio 9.3 8.7 8.2 7.7 3.2
Wing area 9.1 m2 427.8 m2 394 m2 105.63 m2 27.86 m2
Specifications
Empty weight 907 kg 134800 kg 66230 kg 17917 kg 8570 kg
MTOW 1610 kg 247200 kg 145000 kg 41277 kg 19200 kg
Fuel weight 703 kg 94210 kg 38000 kg 18819 kg 7797 kg
Performance
Max. speed 595 kph 950 kph 780 kph 1086kph 2120 kph
Range 2494 km 9700 km 6600 km 12501 km 4220 km
Max. (R/C) 13.2 m/s 26.33 m/s 24.9 m/s 21 m/s 254 m/s
Max. (W/S) 176.9kg/m2 536.5kg/m2 498 kg/m2 390.8kg/m2 431 kg/m2
Service ceiling 8840 m 13140 m 12000 m 15545 m 15240 m
Takeoff run 457 m 2440 m 1800 m 1801 m -
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4. COMPARATIVE DATA SHEET (BUSINESS JETS)
Parameters Airbus A340
Prestige
Boeing B777
VIP
Embraer Lineage
1000
Type Long range
corporate jet
Wide body
business jet
Long range
business jet
Dimensions
Length 63.68 m 73.86 m 36.24 m
Height 16.74 m 18.49 m 10.28 m
Wing span 60.3 m 60.9 m 28.72 m
Aspect ratio 10.1 8.7 8.9
Wing area 361.6 m2
427 m2
92.5 m2
Wing
Sweep angle 300
31.60
250
Root chord 10.6 m 9.57 m 4.76 m
Tip chord 2.6 m 2.29 m 1.22 m
Mean chord 7.26 m 7.01 m 3.22 m
Flaps Single slotted Double slotted Double slotted
Taper ratio 0.25 0.24 0.25
Wing loading 760.5 m2
699.8 m2 594.6 m
2
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Weights
Payload 50900 kg 59422 kg 12400 kg
Empty weight 130900 kg 167800 kg 28080 kg
MTOW 275000 kg 299370 kg 55000 kg
Fuel weight 114615 kg 135845 kg 21871 kg
Powerplant
Name CFM56-5C3 GE90-115B GE CF34-
10E
Thrust rating 145 kN (x4) 514 kN (x2) 82.3 kN (x2)
SFC 0.58 0.55 0.65
Dry weight 3990 kg 8283 kg 1700 kg
Performance
Max speed 914 kph 927 kph 890 kph
Cruise speed 880 kph 895 kph 847 kph
Max R/C 23 m/s 24.38 m/s 16.19 m/s
Service ceiling 12500 m 13140 m 12496 m
Range 14300 km 18700 km 8334 km
Takeoff run 3125 m 3045 m 1852 m
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5. COMPARATIVE GRAPHS
1. Cruise Speed vs. Range
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2. Cruise speed vs. Altitude
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3. Cruise Speed vs. Wing Loading
Page 38
25
4. Cruise Speed vs. Gross Weight
Page 39
26
5. Cruise Speed vs. Aspect Ratio
Page 40
27
6. Range vs. Aspect ratio
Page 41
28
7. Wing loading vs. Aspect ratio
Page 42
29
8. Wing Loading vs. Takeoff run
Page 43
30
9. Wing Loading vs. R/Cmax
Page 44
31
10. Aspect ratio vs. R/Cmax
Page 45
32
6. WEIGHT ESTIMATION
MISSION PROFILE
Gross weight W0 = Wcrew + Wpayload + Wfuel + Wempty
Wcrew = 800 kg ( 2 pilots + 6 cabin crew)
Wpayload = 4000 kg (40 passengers max)
Gross weight W0 = Wcrew + Wpayload [1 – (WfW0) – (WeW0)]
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33
Estimation of empty weight fraction (WeW0)
WeW0 = A W0c
= 1.0595 x W0-0.0598
Estimation of fuel fraction (WfW0)
WfW0 = Ktf x (1 -WnW0)
WnW0 = W1W0 x W2W1 x …. x Wn-1wn-2 x WnWn-1
Fuel fraction for warm up, taxing and take-off (W1W0) = 0.98
Fuel fraction for climb (W2W1) = 0.98
Fuel fraction for cruise (W3W2)
From Breguet range equation:
W3W2 = exp { -R x TSFC / (3.6 x V x L/D) }
To calculate L/D
(L/D)max = 1/ (4 Cd0 k)
Cd0 = 0.005 Rw Tf S-0.1
{1- (Clf/Rw)} [1- 0.2M + 0.12{M (cos1/4)1/2
/ Af -
t/c}20
]
We have,
M= 0.79
Page 47
34
AR= 9.45
t/c = 0.14
Taper ratio, = 0.25
Sweep angle, 1/4 = 250
Ne = 0 (no. of engines located on top of the wings)
Clf = 0 (assuming no laminar flow over the wing in cruise)
Rw = Swet/S = 5.5
Tf = 1.1 (a factor which is unity for streamlined shape)
Af = 0.93 (airfoil factor)
= 1.013 (wing thickness ratio correction factor)
f () = 0.00592
Substituting,
Cd0 = 0.01875 [0.8868]
Cd0 = 0.0166
To calculate K
K = 1/πAR { 1+ 0.12M6 [((1+ {0.142 + f()A(10 t/c)
0.33})/ (cos1/4 )
2 ) +{0.1
(3Ne + 1)/ (4+ AR)0.8
}]
K = 0.03368 {1.029 [1.249 + 0.0125]}
Page 48
35
K = 0.04372
(L/D)max = 1 / (4Cd0k)
(L/D)max = 18.56
(L/D)cruise =86.6 % (L/D)max
(L/D)cruise = 16.07
W3W2 = exp { -R x TSFC / (3.6 x V x L/D) }
= exp { - 11000 x 0.51 / (3.6 x 236.67 x 16.07)
W3/W2 = 0.663
Fuel fraction for loiter (W4/W3)
W4/W3 = exp { - E x TSFC / (L/D)}
= exp { - 0.5 x 0.41 / 18.56 }
W4/W3 = 0.989
Fuel fraction for descent, landing and taxing (W5/W4) = 0.98
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36
W5/W0 = W1W0 x W2W1 x W3/W2 x W4/W3 x W5/w4
= 0.98 x 0.98 x 0.663 x 0.989 x 0.98
W5/W0 = 0.617
WfW0 = Ktf x (1 -WnW0)
Wf/W0 = 1.06 x 1 – (0.617)
Wf/W0 = 0.406
Gross Weight, W0 = Wcrew + Wpayload [1 – (WfW0) – (WeW0)]
W0 = 4800 / [1 – 0.406 – 1.0595(W0)-0.0598
]
W0 Guessed (kg) We/W0 W0 Calculated (kg)
80000 0.539 87272
85000 0.537 84210
84500 0.5377 85257
84800 0.5375 84995
84900 0.53746 84907
Page 50
37
The estimated Gross weight (W0) is 84907 kg
Wf = 0.406 x 84907
= 34472 kg is the maximum fuel weight onboard.
Maximum fuel capacity,
Vf = Wf / fuel
= 34472 / 0.809; fuel = 0.809 kg/l, (value obtained from Jane’s All the World
Aircraft)
Vf = 42610 L (in wings + 3-9 belly tanks)
54%
6%
40% Empty weightPaylodFuel weight
Weight Distribution Chart
Page 51
38
7. WING LOADING
THRUST LOADING (T/W)
T/W0 = AMmaxC
= 0.267 (0.82)0.363
(T/W)takeoff = 0.248
(T/W)cruise = 1/(L/D)cruise
(T/W)cruise= 0.0622
WING LOADING (W/S)
Stall:
Vstall = Vapproach / 1.3
= 72.5 / 1.3
Vstall = 55.77 m/s
CLmax = CL =0 cos 1/4
= 3.4 cos 250
= 3.08
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39
W/S = (Vstall2 CLmax )/2
= (1.225 x 55.772 x 3.08) / 2
W/S = 598.12 kg/m2 (at sea level)
Landing:
Ground roll distance, S = 80 (W/S) / (CLmax )
(W/S) = S (CLmax )/ 80
= 662 x 0.82 x 3.08 / 80
(W/S)landing = 21 kg/m2
Cruise:
Skin friction coefficient, Cfe = 0.003 (subsonic)
Assuming Swet/S = 5.5
Parasite drag CDo = Cfe (Swet / Sref)
CDo = 0.0165
Oswald efficiency factor,
1/e = 1/ewing + 1/efuselage + 0.05
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40
ewing = 0.84 for an unswept wing of A = 9.45 and λ = 0.25
ewing for a swept wing is,
ewing = ewing=0 cos (-5)
= 0.84 cos (25-5)
= 0.7893
1/efuselage = 0.1
1/e = 1.267 + 0.1+ 0.05
Therefore, e = 0.707
At 10 km altitude, V= 236.67, = 0.41 kg/m2
q0 = ℓ V2 / 2 = 1170.5 kg/m
2
(W/S)optimum cruise = q0 (πeARCD0 /3)0.5
= 1170.5 ( 0.432)
(W/S)optimum cruise = 506.2 kg/m2
(W/S)takeoff = (W/S)opt cruise x (W1/W0)-1
x (W2/W1)-1
= 506.2 x 0.98-1
x 0.98-1
(W/S)takeoff = 527 kg/m2
Page 54
41
0
100
200
300
400
500
600
takeoff cruise landing stall
W/S
Wing loading
Page 55
42
8. AIRFOIL SELECTION
AIRFOIL GEOMETRY
An airfoil is a surface designed to obtain a desirable reaction from the air
through which it moves.
Chord line: Straight line connecting leading edge and trailing edge.
Thickness: Measured perpendicular to chord line as a % of it.
Camber: Curvature of section – perpendicular distance of section mid-points
from chord line as a % of it.
ANGLE OF ATTACK ()
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43
Angle of attack () is the angle between the free stream and the chord line.
Aerofoil Selection is based on the factors of Geometry & definitions,
design/selection, families/types, design lift coefficient, thickness/chord ratio, lift
curve slope, characteristic curves.
The following are airfoil categories:
Early on, airfoil selection was based on trial & error.
NACA 4 digit was introduced during the 1930’s.
NACA 5-digit is aimed at pushing position of max camber forwards for
increased Clmax.
NACA 6-digit is designed for lower drag by increasing region of laminar flow.
The modern airfoil 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)).
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44
– 3rd & 4th digits: maximum section thickness (as % of chord).
NACA 5 Digit:
– 1st digit (x0.15): design lift coefficient.
– 2nd & 3rd digits (x0.5): location of maximum camber (as % of chord
from LE).
– 4th & 5th digits: maximum section thickness (as % of chord).
NACA 6 Digit:
– 1st digit: identifies series type.
– 2nd digit (x10): location of minimum pressure (as % of chord from
leading edge (LE)).
– 3rd digit: indicates acceptable range of CL above/below design value
for satisfactory low drag performance (as tenths of CL).
– 4th digit (x0.1): design CL.
– 5th & 6th digits: maximum section thickness (%c)
It becomes necessary to use high speed airfoils, i.e., the 6x series, which have
been designed to suit high subsonic cruise Mach numbers.
Page 58
45
NACA 63A-514 (Root airfoil)
Max thickness 14%
Max camber 3.2%
Page 59
46
NACA 63-512 (Midspan airfoil)
Max thickness 12.5%
Max camber 2.2%
Page 60
47
NACA 63-310 (Tip airfoil)
Max thickness 10%
Max camber 1.1%
(JavaFoil – airfoil generator)
Page 61
48
MAXIMUM THICKNESS (T/C)
Maximum thickness of the airfoil desired to produce max Cl is 14%
With a wing sweep angle of 250, the max lift coefficient can be obtained from
Clmax = Cl=0 cos 1/4
= 3.4 cos 250
= 3.08 (at 400 flap settings)
Clcruise = 2W/ (V2S)
= 2 x 598 x 9.81 / 0.41 x 236.672)
= 0.511
Clreq,takeoff = 1.5 (at 120 angle of attack) ….. (From the plots above)
Page 62
49
.
Wing area, S = 84907 / 598.12 = 142 m2
Wing span, b = (9.45 x 142) = 36.63 m
Root chord, cR = 2 x 142 / [36.63 x (1 + 0.25)] …. ( = 0.25)
= 6.2 m
Tip chord, cT = 0.25 x 6.2
= 1.55 m
Mean aerodynamic chord (mac) = (2/3) [(1++2)/(1+)] c = 4.34 m
Page 63
50
LE = c/4 + [(1-) / AR (1+)]
= 250 + 0.75 / (9.45 x 1.25)
= 25.060
Meff = McosLE
= 0.79 cos25.060
= 0.715
Wing sweep reduces effective Mach number over the wing.
= (1-M2
eff)
= 0.7
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51
Dihedral () is the angle of the wing with respect to the horizontal plane when
seen in the front view. Dihedral of the wing affects the lateral stability of the
airplane. A value of Γ = 50 is chosen.
Wing sweep effect on dCL/d
dCL/d = 2.π.AR / [2+ {4+(AR.)2. (1+tan
2t/c / 2)}]
= 2 x π x 9.45 / [2+ {4+(9.45 x 0.7)2 x (1+tan
225 / 0.7
2)}]
= 59.376 / 5.269
= 11.27
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52
WINGLETS
Blended winglets are used in this heavy business jet.
A blended winglet is attached to the wing with smooth curve instead of a sharp
angle and is intended to reduce interference drag at the wing/winglet junction.
These winglets which stand 2.5m tall each offers 5 to 7% reduction in cruise
drag (induced drag) and increase in wing area and aspect ratio without
geometrically increasing the wing span which results in 8 to 10% increase in
range.
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53
WING PARAMETERS:
Design Parameters Values
Wing loading (W/S) 598 kg/m2
Wing area (S) 142 m2
Aspect ratio (AR) 9.45
Wing span (b) 36.63 m
Taper ratio () 0.25
Root chord (cR) 6.2 m
Tip chord (cT) 1.55 m
Mean chord (cm) 4.34 m
Design CL 0.511
Sweepback angle () 250
Dihedral angle () 50
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54
9. DRAG ESTIMATION
The drag polar is expressed as
CD = CD0 + KCL2
Where K = 1 / πAe
e = Oswald efficiency factor
Parasite drag CDo = Cfe (Swet / Sref)
Where, Cfe = equivalent skin friction drag coefficient ;
Swet = Wetted area of the airplane.
Swet/Sref = 5.5
The estimation of K is carried out next and then the value of CD0 is deduced
using the earlier calculation that (L/D)max = 18.56
ESTIMATION OF K:
Oswald efficiency factor,
1/e = 1/ewing + 1/efuselage + 0.05
ewing = 0.84 for an unswept wing of A = 9.45 and λ = 0.25
ewing for a swept wing is,
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55
ewing = ewing=0 cos (-5)
= 0.84 cos (25-5)
= 0.7893
1/efuselage = 0.1
1/e = 1.267 + 0.1+ 0.05
Therefore, e = 0.707
K = 1 / πAe
= 1 / π x 9.45 x 0.707
= 0.0476
(L/D)max = 1 / 2(CD0 K)
CD0 = 1 / 4K (L/D)2
max
= 1 / 4 x 0.0476 x 18.562
= 0.0165
Cfe = 0.0165 / 5.5
= 0.003
The drag polar is:
CD = 0.0165 + 0.0476 CL2
Drag, D = (1/2) V2SCD
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56
Takeoff:
= 1.225 kg/m3
V = 1.15 Vstall
= 1.15 (55.77)
= 64.13 m/s
S = 142 m2
Drag, D = 0.5 x 1.225 x 64.132 x 142 x (0.0165 + 0.0476 x 1.5
2)
Dtakeoff = 44.21 kN
Landing:
= 1.225 kg/m3
V = 1.3 Vstall
= 72.5 m/s
Drag, D = 0.5 x 1.225 x 72.52 x 142 x (0.0165 + 0.0476 x 3.08
2 )
DLanding = 213.9 kN
Cruise:
= 0.466 (at 10 km altitude)
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57
V = 236.67 m/s
S = 142 m2
Drag, D = 0.5 x 0.41 x 236.672 x 142 x (0.0165 + 0.0476 x 0.511
2 )
Dcruise = 47.17 kN
also, (T/W)cruise = 1 / (L/D)cruise
T/W = 0.0622
T = 0.0622 x (84907 x 9.81 )
Tcruise = 51.81 kN
In straight and level flight, D ~ T
Here, Dcruise and Tcruise calculated are almost equal.
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10. POWERPLANT SELECTION
COMPARISON OF TURBOFAN ENGINES:
Engine Length
(m)
Diameter
(m)
Thrust
(kN)
Weight
(kg)
T/W
Bypass
ratio
Pressure
ratio
Sfc
(hr-1
)
CFM56-
7B
2.5
1.55
86.7
2366
3.7
5.5:1
32.8:1
0.56
PW6000
2.74
1.44
99-106
2289
4.7
5:1
28.2:1
0.68
CFM leap
-1A
3.5
1.9
123.5-
133.4
3762
3.6
11:1
40:1
0.44
PW1000G
3
2.1
110 -
150
3796
4
10:1 -
12:1
38:1
0.39
ENGINE SELECTION:
The thrust loading based on sea level static thrust is:
T/W = 0.248 ……. (from thrust loading calculation)
Thus, the thrust required is,
Treq = 0.248 x 84907 x 9.81 = 206.57 kN
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59
It is observed that the maximum thrust requirements occurs from Vmax
consideration i.e. Tmax = 206.57 kN.
As a twin engine configuration has been adopted, the above requirement implies
a thrust per engine of 103.28 kN.
The above comparison of high bypass turbofan engines shows the competition
between CFM LEAP-1C and PW1000-G in various parameters. Unlike fighter
aircraft, business jets or any airliner in that case looks for an important
parameter which is lowest specific fuel consumption. Though LEAP-1C gives a
pressure ratio higher than PW1000-G, it contains more number of stages which
adds weight to the aircraft. On the other hand, PW1000G has a lowest TSFC of
0.39/hr.
PW1000-G will be designed with a variable inlet duct and a Gearing system
(Geared turbofan), that will allow changes in bypass ratio by controlling the
rpm of the fan, whenever required as per the flight phase. In addition to the
geared turbofan, the current design includes a variable-area nozzle, which offers
reduction in noise. It also offers 15% reduction in CO2 emission and 55%
reduced NOx margin in accordance with CAEP/6.
Taking these advantages in consideration, PW1000-G is selected.
Selected Engine series of PW1000G family:
PW1124G
PW1127G
PW1133G
Page 73
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DETAILS OF THE SELECTED ENGINE:
PRATT & WHITNEY PW1000G
The Pratt & Whitney PW1000G is a high-bypass geared turbofan engine family,
currently selected as the exclusive engine for the Bombardier CSeries,
Mitsubishi Regional Jet (MRJ), Embraer's second generation E-Jets, and as an
ultra efficient option for the Airbus A320neo.
Page 74
61
FAN:
A large, light-weight fan moves well over 90% of air around the core, delivering
a very quiet engine with very low fuel burn.
COMPRESSORS AND TURBINES:
A compact, high-speed low-pressure system accomplishes the same work in
fewer stages. That means fewer airfoils, fewer life-limited parts, and ultimately
lower maintenance costs.
CORE:
The supercharged low-pressure system allows the advanced PurePower engine
core – optimized for high-cycle durability – to run cooler than the closest
competition, with fewer stages, and without expensive materials. That means
longer time on wing and lower maintenance costs
Page 75
62
PW1000G
COMPONENTS:
COMPRESSOR: Axial flow, 1-stage geared fan, 3-stage LP, 8-stage HP
COMBUSTORS: Annular combustion chamber
TURBINE: Axial, 2-stage HP, 3-stage LP
Page 76
63
GEARED TURBOFAN:
In a conventional turbofan engine, a single shaft (the "low-pressure" or LP
shaft) connects the fan, the low-pressure compressor and the low-pressure
turbine. A second concentric shaft connects the high-pressure compressor and
high-pressure turbine.
In this configuration, the maximum tip speed for the fan limits the rotational
speed for the LP shaft and thus the LP compressor and turbine. At high bypass
ratios (and thus high radius ratios) the tip speeds of the LP turbine and LP
compressor must be relatively low, which means extra compressor and turbine
stages are required to keep the average stage loadings and, therefore, overall
component efficiencies to an acceptable level.
In a geared turbofan, a reduction gearbox between the fan and the LP shaft
allows the latter to run at a higher rotational speed thus enabling fewer stages to
be used in both the LP turbine and the HP compressor, increasing efficiency and
reducing weight. Also the weight saved on turbine and compressor stages is
offset to some extent by the mass of the gearbox.
The Pure Power engine allows for a more efficient arrangement: a big, slow fan
shoving air into a small, fast turbine. The result is a shorter, lighter engine that
can produce the same amount of power as a larger conventional turbofan, while
burning 15 percent less fuel and emitting 15 percent less carbon dioxide.
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64
11. LANDING GEAR DESIGN
The landing gear supports the aircraft when it is not flying, allowing it to take
off, land and usually to taxi without damage. Landing gear placement is
essential for ground stability and controllability. A good landing gear position
must provide superior handling characteristics and must not allow over-
balancing during takeoff or landing.
Landing gear arrangement:
Landing gears normally come in two types: conventional or "taildragger"
landing gear, where there are two main wheels towards the front of the aircraft
and a single, much smaller, wheel or skid at the rear; or tricycle landing gear,
where there are two main wheels (or wheel assemblies) under the wings and a
third smaller wheel in the nose.
To decrease drag in flight some undercarriages retract into the wings and/or
fuselage with wheels flush against the surface or concealed behind doors; this is
called retractable gear.
With a tricycle landing gear, the c.g is ahead of the main wheels, so the aircraft
is stable on the ground. It improves forward visibility on the ground and permits
a flat cabin floor for passengers and cargo loading.
Thus retractable tricycle landing gear system is selected.
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65
Tyre sizing:
The “wheel” is the circular metal object upon which the rubber “tyre” is
mounted. The “brake” inside the wheel slows the aircraft by increasing the
rolling friction. However, the term “wheel” is frequently used to mean the entire
wheel/brake/tyre assembly.
The tyres are sized to carry the weight of the weight of the aircraft. Typically
the main tyres carry about 90% of the total aircraft weight. Nose tyres carry
only about 10% of the static load but experience higher dynamic loads during
landing.
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66
The nose gear is of double‐bogey type with two wheels. The main gear consists
of two sets of wheels (wing‐retracted) each of multi‐bogey type with 4 wheels
each.
Nose gear:
Load on nose gear = 0.1W0
= 8490.7 kg
Load per tyre, Lt = 4245.35 kg = 9359.4 lb
Wheel diameter = 2.69Lt0.251
………..(from Raymer)
= 21.9 in = 0.56 m
Wheel width = 1.17Lt0.216
………(from Raymer)
= 7.11 in = 0.18 m
Tyre size, 27 x 7.75
Tyre diameter, d = 27 in = 0.686 m
Tyre width, w = 7.75 in = 0.197 m
Rolling radius, Rr = 11.5 in = 0.3 m
Pavement contact area, Ap = 2.3 x(wd) x (0.5d – Rr)
= 66.54 in2
Tyre pressure = 9359.4 / 66.54 = 140.6 psi
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67
Main gear:
Load on main gear = 0.9 W0
= 76416.3 kg
Load per tyre, Lt = 19104 kg = 42117.3 lb
Wheel diameter = 2.69Lt0.251
………..(from Raymer)
= 31.94 in = 0.81 m
Wheel width = 1.17Lt0.216
………(from Raymer)
= 9.83 in = 0.25 m
Selecting Goodrich tyre of size 40x14.5
Tyre diameter d = 40 in = 1 m
Tyre width, w = 14.5 in = 0.37 m
Rolling radius, Rr = 16.3 in = 0.414 m
Pavement contact area, Ap = 2.3 x(wd) x (0.5d – Rr)
= 205 in2
Tyre pressure = 42117.3 / 205 = 205 psi
Wheel base = 17.17 m
Wheel track = 5.71 m (incl. shock struts)
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12. PERFORMANCE CHARACTERISTICS
TAKEOFF PERFORMANCE:
Distance from rest to clearance of obstacle in flight path and usually considered
in two parts:
- Ground roll - rest to lift-off (SLO)
- Airborne distance – lift off to specified height of 50ft
The aircraft will accelerate up to lift-off speed (VLO = about 1.2 x Vstall) when it
will then be rotated.
Ground roll take-off distance is given by
SLO = 1.21W / gSCLmax(T/W) ….. (from ‘Aircraft Performance & Design’
by John. D Anderson)
= 1.21 x 84907x9.81 / (9.81 x 1.225 x 142 x1.5 x 0.248)
…(CLmax for takeoff = 1.5, from airfoil selection)
SLO= 1587.67 m
CLIMBING:
Consider aircraft in a steady unaccelerated climb with vertical climb speed of Vc
𝐿=𝑊 𝑐𝑜𝑠𝛾C
𝑇=𝐷+𝑊 𝑠𝑖𝑛𝛾C
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69
VC = (T - D)Vstall / W
R/Cmax = VC = (2x206570 – 44210) x 55.77 / (84907 x 9.81)
R/Cmax = 24.7 m/s
LEVEL TURN:
In the case of a commercial transport aircraft, it is capable of performing only a
constant altitude banked turn and not any vertical pull-up or pull-down
manoeuvres.
In steady condition: T = D
Force balance gives:
W = Lcos
Fr = mV2 / r = Lsin
tan = V2 / Rg
So for given speed and turn radius there is only one correct bank angle for a co-
ordinate (no sideslip) turn.
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70
In the turn, n = L/W = sec > 1 and is therefore determined by bank angle.
Turn radius (R) and turn rate () are good indicators of aircraft
manoeuvrability.
V2 / (Rg) = tan = (sec
2 - 1) = (n2 - 1)
R = V2/ (g (n2 - 1))
And = V/R = (g (n2 - 1)) / V
W = Lcos
Let = 600
n = 𝐿/𝑊 = 2
R = V2/ (g (n
2 - 1))
= 236.672 / (9.81 x (2
2 – 1))
R = 3296.5 m
= V/R
= 236.67 / 3296.5
= 0.072 rad/s
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71
GLIDING:
The thrust can be assumed to be zero while the aircraft is gliding.
= tan-1
[1/ (L/D)]
= tan-1
[ 1 / 18.57]
= 3.080 is the glide angle.
LANDING PERFORMANCE:
APPROACH & LANDING:
- Airborne approach at constant glide angle (around 30 ) and at constant
speed.
- Flare - transitional maneuver with airspeed reduced from about 1.3 Vstall
down to touch-down speed.
- Ground roll - from touch-down to rest.
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72
Ground roll distance (STD):
STD = 1.69 W2 / 𝑔𝜌𝑆𝐶𝐿𝑚𝑎𝑥 [D + r (W-L)]
r is higher than for take-off since brakes are applied - use r = 0.4 for paved
surface.
STD = 1.69 x (84907 x 9.81)2 / (9.81 x 1.225 x 142 x 3.08 x [213976 +
0.4(832937.67 – 0.8 x 832937.67)]
Landing distance is STD = 795 m
Performance
Parameters
Values Units
Takeoff distance 1587.67 m
(R/C)max 24.7 m/s
Turn radius 3296.5 m
Turn rate 0.072 rad/s
Glide angle 3.08 deg
Landing distance 795 m
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73
13. CENTRE OF GRAVITY ESTIMATION
The weight of an airplane changes in the flight due to consumption of fuel and
dropping off / release of armament or supplies. Further, the payload and the
amount of fuel carried by the airplane may vary from flight to flight.
These factors lead to change in the location of the centre of gravity (c.g.) of the
airplane. The shift in the c.g location affects the stability and controllability of
the airplane.
The weight of entire airplane can be sub divided into empty weight and useful
load. The empty weight can be further subdivided into:
(i) structures group
(ii) propulsion group and
(iii) equipment group.
The structures group consists of the following components:
- wing
- horizontal tail /canard
- vertical tail
- fuselage
- landing gear - main and nose/tail wheel
- nacelle, engine pod and air intake
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The propulsion group consists of the following components:
- engine as installed
- reduction gear
- propeller for piston and turboprop engines
- cooling provisions
- engine controls
- fuel system and tanks
The equipment group consists of the following items:
- flight controls
- auxiliary power unit (APU)
- instruments
- hydraulic, pneumatic, electrical, armament, air conditioning, anti-icing
- avionics
- furnishings in passenger airplanes
The useful load consists of:
(i) Crew
(ii) Fuel - usable and trapped
(iii) Oil
(iv) Payload - passengers, cargo and baggage in transport airplane;
ammunition, expendable weapons and other items in military airplanes.
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Approximate group weights method:
(Reference: “Aircraft design: A Conceptual Approach” by Daniel P. Raymer)
The aim of estimating the weights of individual components and their c.g. is to
obtain the location of the c.g. of the airplane. Then, the shift in the airplane c.g.
is examined under various conditions.
At this stage of preliminary design, the weights of individual components are
estimated using simpler method like using the table above.
The gross weight of the airplane estimated is 84907 kg
The weights and c.g. locations of various components are estimated below:
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Wing:
S = 142 m2
b = 36.63 m
bsemi = (36.63/2) – (3.76/2) = 16.435 m
cR = 6.2 m
cT = 1.55 m
Fuselage width = 3.76 m
mac = 4.34 m
Location of L.E of mac from L.E of wing = 0.45 m
S(exposed)wing = 142 – (6.2x3.76) = 118.7 m2
From the table, the weight of the wing is,
Wwing = 118.7 x 49 = 5816.3 kg
Wwing / W0 = 6.8%
From the table, c.g of the wing is at 40% of mac
Hence, the location of the c.g. of wing from the leading edge of the root chord
is, 0.45 + 0.4x4.34 = 2.186 m
Horizontal tail:
btail = 14.35 m
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Stail = 32.8 m2
cR ,h.tail = 3.38 m
cT, htail = 1.18 m
mac, htail = 2.29 m
S(exposed)tail = 27.88 m2
From the table,
Weight of Horizontal tail = 27 x 27.88 = 752.76 kg
Wht / W0 = 0.887%
From the table, c.g of the h.tail is at 40% of mac.
Hence, the location of the c.g. of h.tail from the leading edge of the root chord
of h.tail is, 0.45 + 0.4 x 2.29 = 1.37 m
Vertical tail:
The contribution of dorsal fin to the weight of v.tail is ignored at this stage of
preliminary design.
Sv.tail = 26.40 m2
From the table,
Weight of vertical tail = 27 x 26.40 = 712.8 kg
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Wvt/W0 = 0.84%
From the table, c.g of the v.tail is at 40% of mac.
Mean aerodynamic chord of v.tail = 3.7 m
Hence, the location of the c.g. of v.tail from the leading edge of the root chord
of v.tail is 1.45 + 0.4 x 3.7 = 2.93 m
Engine:
The weight of each engine is 3796 kg
From the table, the installed weight of two engines is,
Wengine = 1.3 x (2 x 3796) = 9869.6 kg
Wengine / W0 = 11.62 %
For gas turbine engines the location of c.g. from the engine inlet is between 30
to 45% of engine length.
In the present case the engine length is 3 m.
The engines are located at 3.06 m from the wing root and the inlet is at 2.5 m
from wing leading edge.
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Hence, the location of c.g of engine from L.E of the wing is,
= -2.5 + (0.4 x 3) = -1.3 m i.e., 1.3 m ahead of the L.E of root chord of wing.
Landing gear:
From the table, the weight of the nose wheel plus the main landing gear is 4.3%
of W0. i.e., 0.043 x 84907 = 3651 kg
Out of this total weight, the nose wheel and main wheel account for 15% and
85% respectively.
Hence, nose wheel weighs 0.15 x 3651 = 547.65 kg
And the main wheels weigh 0.85 x 3651 = 3103.35 kg.
With regard to the locations of the c.g.’s of nose wheel and main wheels, it is
recalled that the nose wheel and main wheels share respectively 10 % and 90 %
of the airplane weight.
Wheel base is 17.17 m.
Hence, the c.g. of the nose wheel is 0.9 x 17.17 = 15.45 m ahead of the c.g. of
the airplane.
The c.g. of the main wheels, as a group, is:
= 0.1 x 17.17 = 1.717 m behind the c.g. of the airplane.
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Fuselage:
An approximate estimation of fuselage wetted area is,
= 0.75 x perimeter of fuselage x length of fuselage
= 0.75 x π x 3.76 x 42
= 372.1 m2
(Wfuse + Wsyst)/ W0 = (Wempty/W0) – [(Ww/W0) + (Wht/W0) + (Wvt/W0) + (We/W0)
+ (Wlg/W0)]
(Wfuse + Wsyst)/ W0 = 0.5375 – (0.068 + 0.00887 + 0.0084 + 0.1162 + 0.043)
(Wfuse + Wsyst) = 0.2932 x 84907 = 24894.73 kg
From the table, the c.g of fuselage and systems is, 0.45 x length of fuselage
= 0.45 x 42 =18.9 m from the nose of the fuselage.
Wing location and c.g. of the airplane:
The wing is chosen such that the c.g. of the entire airplane with the gross weight
is at 25% of the mean aerodynamic chord of the wing.
The distance of the leading edge of the root chord of the wing from the nose of
the fuselage is denoted by xlew. The 25% of the mean aerodynamic chord (mac)
of wing is 1.5 m behind xlew.
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Hence, the chosen location of the c.g. of the entire airplane is at (xlew + 1.5) m
from the nose of fuselage.
Item W (kg) x (m) W.x (kg.m)
Wing
5816.3
xlew +2.186
5816.3xlew +
12714.43
H.tail
752.76
xlew+1.37
752.76xlew
+1031.28
V.tail
712.8
xlew + 2.93
712.8xlew +
2088.5
Engines
9869.6
xlew - 1.3
9869.6xlew –
12830.48
Nose wheel
547.65
xlew – 15.45
547.65xlew –
8461.2
Main wheel
3103.35
xlew + 1.717
3103.35xlew
+5328.45
Fuselage & syst
24894.73
18.9
470510.4
45696.89
20800.9xlew +
512964.34
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The location of wing from nose of airplane is,
45696.89(xlew + 1.5) = 20800.9xlew + 512964.34
xlew = 17.85 m
Hence, the c.g. of the airplane is at 17.85 + 1.5 = 19.3 m from nose of fuselage.
Since, c.g lies at a.c of wing, aerodynamic centre also lies at 19.3 m from nose
of fuselage.
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14. STABILITY & CONTROL
STATIC LONGITUDINAL STABILITY:
For the aircraft configuration adopted,
xc.g = 1.5/4.34 = 0.3456ĉ;
ĉ = 4.34 m
(Cm/CL)fuse,nac = 0.042 ;
/ = 0.4
vHT = 0.48 ;
t = 0.9
For the wing, airfoil selected is NACA 63A-514
CL/ = 0.125 deg-1
….. (refer lift curve plot in airfoil selection)
xac = 0.25ĉ
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For the tail, airfoil selected is NACA 0012
CL/ = 0.1 deg-1
For the longitudinal stick‐fixed static stability of the aircraft, we have
(Cm/CL)stickfixed = (xcg – xac) + (Cm/CL)fuse,nac – [(at vHT t/aw) (1- /)]
= (0.3456 – 0.25) + 0.042 – 0.20736
(Cm/CL)stickfixed = - 0.06976
The negative value of (Cm/CL)stickfixed indicates that the airplane has
longitudinal stick-fixed static stability.
Longitudinal stability
(Ref: “Airplane Performance Stability & Control” by Perkins)
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Stick fixed Neutral point:
When the value of (Cm/CL)stickfixed reduces to zero, the location of the
C.G is called the Neutral Point.
N0 = (xac/ĉ) - (Cm/CL)fuse,nac + [(at vHT t/aw) (1- /)]
= 0.25 – 0.042 + 0.20736
N0 =0.41536
Static margin, = N0 – (xcg/ĉ)
= 0.41536 – 0.3456
Static margin = 0.07
STICK FIXED LONGITUDINAL CONTROL:
e = 0.4
vHT = 0.48
t = 0.9
The rate at which the pitching moment coefficient of the aircraft changes with
change in elevator deflection is called the Elevator control power.
(Cm/e) = - at vHT t e
= - 0.1 x 0.48 x 0.9 x 0.4 (Cm/e) = - 0.01728 deg-1
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DIRECTIONAL STABILITY:
Directional stability of the airplane is a measure of its tendency to produce
restoring moments when disturbed from an equilibrium angle of sideslip –
usually taken as zero. It is measured quantitatively by the variation of yawing
moment coefficient with sideslip angle.
Cn = N / q S b
In general Cn / should be negative for the airplane to have static directional
stability. All the components of the aircraft contribute to the stability coefficient Cn /
.
Contribution from Wing:
The wing contribution to directional stability is quite small, as the cross wind
effects on the wing are very small. The critical factor is the sweepback () of
the wing.
(Cn / )wing = - 0.00006
For the aircraft, = 250
(Cn / )wing = - 0.0003 deg-1
Contribution from Fuselage and Nacelle:
(Cn / )fuse,nace = (0.96K / 57.3) (Sf/Sw) (Lf/b) (h1/h2)1/2
(w2/w1)1/3
where, K is an empirical constant which is obtained from the graph below,
K = 0.12
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Sf = 157.92 m2 = Projected area of fuselage
h1, w1 : Height and width of fuselage at Lf /4 in m
h2, w2 : Height and width of fuselage at 3Lf /4 in m
h1 = h2 = 3.76 m
w1 = w2 = 3.76 m
(Cn / )fuse,nace = (0.96K / 57.3) (Sf/Sw) (Lf/b) (h1/h2)1/2
(w2/w1)1/3
(Cn / )fuse,nace = 0.00256 deg-1
Thus the fuselage is destabilizing the aircraft.
(Reference: “Airplane Performance Stability & Control” by Perkins)
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Contribution of Vertical Tail:
The vertical tail is the stabilizing component in the aircraft as far as directional
stability is concerned.
(Cn / )v.tail = - av vv v (1 - )
(Cn / )v.tail = (Cn / )v.t + (Cn / )slipstream interference
For the vertical tail,
av = 0.1 deg-1
vv = 0.05
v = 0.9
(Cn / )v.t = - 0.0045
(Cn / )slipstream interference = -0.0003 (Perkins)
(Cn / )v.tail = -0.0048 deg-1
(Cn / )airplane = (Cn / )wing + (Cn / )fuse,nace + (Cn / )v.tail
(Cn / )airplane = - 0.00254 deg-1
Directional Control:
The rate at which the yawing moment coefficient of the aircraft changes with
change in rudder deflection is called the Rudder control power.
(Cn/r) = - av vv v r
For rudder, r = 0.4
(Cn/r) = - 0.0018 deg-1
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Directional stability
(Reference: “Airplane Performance Stability & Control” by Perkins)
LATERAL STABILITY:
When a small vertical disturbance causes the aircraft to roll to one side, as such,
the airplane will continue to roll at the same constant velocity. As such, the
airplane is neutrally stable in roll. However, due to the development of sideslip,
the lift distribution over the wings is altered, tending to produce restoring
moments which restore the aircraft to its original state. This effect is generally
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called the Dihedral Effect. Lateral stability of the airplane is a measure of this
tendency to produce restoring moments when disturbed in roll.
The rolling moment coefficient is:
Cl = L / (q s b)
In general, Cl / should be positive for the airplane to have static lateral
stability. All the components of the aircraft contribute to the stability coefficient
Cl / .
Contribution from Wing:
The wing dihedral angle has a linear variation with the stability coefficient.
There is also an additional component due to the tip shape. The stability
coefficient is given by
(Cl / )wing = 0.0002 + (Cl / )tip-shape + (Cl / )sweepback
where - dihedral angle in degrees = 50
(Cl / )tip-shape = 0.0002 deg-1
(Perkins)
(Cl / )sweepback = -0.5 (Cn / )wing
= 0.00015 deg-1
(Cl / )wing = 0.00135 deg-1
Since the wing configuration is low‐wing,
(Cl / )interference = - 0.0008 deg-1
(Perkins)
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Contribution of Vertical Tail:
The vertical tail is stabilizing as far as directional stability is concerned.
(Cl / )v.tail = av vv v (Zv/lv)
For the vertical tail,
Zv = 4.78 m
lv = 17.75 m
(Cl / )v.tail = 0.00121183
Since the wing configuration is low‐wing,
(Cl / )interference = 0.00016 deg-1
(Cl / )airplane = (Cl / )wing + (Cl / )v.tail + (Cl / )interference
(Cl / )airplane = 0.001922
Lateral stability
(Reference: “Airplane Performance Stability & Control” by Perkins)
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15. 3-VIEW DIAGRAM
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93
All dimensions are in m
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16. FINALIZED DESIGN PARAMETERS
Parameters Values
Seating capacity 12 - 40
Length 43.1 m
Height 12.5 m
Wing span 36.63 m
Wing area 142 m2
Wing sweep angle 250
Cabin width 3.53 m
Fuselage width 3.76 m
Empty weight 45634 kg
Max. Takeoff weight 84907 kg
Max. Fuel capacity 42610 L
Max. speed 0.82 M
Cruise speed 0.79 M
Range 11000 km
Service ceiling 12500 m
Wing loading 598 kg/m2
Engines (x2) PW1000G geared turbofan
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17. CONCLUSION
The preliminary design of a heavy business jet is done and the various design
considerations and performance parameters required are calculated and found
out. The obtained design values are not necessarily a definite reflection of the
airplane's true and conceptualized design, but the basic outlay of development
has been obtained.
The final design stays true to the desired considerations of a long range aircraft
that can provide high fuel efficiency as well. There is no ideal design as such
and continuous changes, improvements and innovations serve to make the
design as ideal as possible, while always looking to achieve optimum
performance.
The challenges we faced at various phases of the project made clear the fact that
experience plays a vital role in successful design of any aircraft or aircraft
component. A lot of effort has been put into this project and as much as we have
worked, we have learnt in turn.
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18. REFERENCES
1. Anderson, John D. Jr., (1999) Aircraft Performance and Design,
McGraw-Hill, New York
2. Anderson, John D. Jr., (2001) Introduction to Flight, McGraw-Hill ,
New York
3. Perkins, C. and Hage, R. (1949) Airplane Performance, Stability and
Control, Wiley, New York
4. Raymer, Daniel P. (1992) Aircraft Design: A Conceptual Approach,
AIAA Education series, Washington, DC
5. Roskam, J. (1985) Airplane Design, Roskam Aviation and Engineering
Corp., Ottawa, Kansas
6. Taylor, J. (2004) Jane’s All the World’s Aircraft, Jane’s, London, UK
WEBSITES
1. Boeing technical characteristics, viewed 2 March 2014
http://www.boeing.com/boeing/commercial/737family/specs.page
2. Engine selection and technical Information, viewed 25 March 2014
www.purepowerengine.com
http://en.wikipedia.org/wiki/Pratt_%26_Whitney_PW1000G
3. JavaFoil – Analysis of airfoil, viewed 29 March 2014
http://www.mhaerotools.de/aerofoils/javafoil.htm