AIRCRAFT DESIGN PROJECT -I FIGHTER JETS A PROJECT REPORT

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AIRCRAFT DESIGN PROJECT - I
FIGHTER JETS
A PROJECT REPORT
By:
G.Srilochan 14101073
Sharon George 14101074
Franklin C. 14101091

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ACKNOWLEDGEMENTS
We owe a debt of gratitude to Mr. DILIP A SHAH, 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 assistant professor Mr.MAYAKANNAN 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 HINDUSTAN INSTITUTE OF TECHNOLOGY AND
SCIENCES, for providing us with the amenities and a congenial atmosphere to work in.

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ABSTRACT
Through this project, we intend to design and conceptualize a FIGHTER AIRCRAFT that
can cater to a wide range of needs. Fighter aircraft is a term describing an aircraft, usually of
smaller size, designed for military purposes and combat. The project involves the design of a
fighter jet that can carry heavy weapons, providing the amenities with sophisticated care
while incorporating the design specifications and performance parameters of a fighter jet. The
aircraft allows for 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
a .LIST OF TABLES 5
b.LIST OF GRAPHS 5
c.LIST OF FIGURES 5
1. INTRODUCTION TO DESIGN 8
2. INTRODUCTION TO AIRCRAFT 14
3. COMPARATIVE DETAILS 17
4. COMPARATIVE GRAPHS 23
5. WEIGHT ESTIMATION 33
6. AIRFOIL AND WING SELECTION 36
7. TAIL PLANE 41
8. LANDING GEAR 45
9. POWER PLANT SELECTION 48
10 DRAG ESTIMATION 51
11 .V-N DIAGRAM 55
12. 3 VIEW DIAGRAM 57
13. FINAL PARAMETER 59
14. CONCLUSION 60
15. REFERENCE 61

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LIST OF FIGURES
FIGURE 1:- NACA AEROFOIL 36
FIGURE 2:-LOW WING DIHEDRAL 38
FIGURE 3:-STRAIGHT DIHEDKERAL 39
FIGURE 4:-WING GEOMETRY 40
FIGURE 5:-M201 WITH A RETRACTABLE 46
LIST OF GRAPH
GRAPH 1:- LENGTH VS SPEED 24
GRAPH 2:-WING SPAN VS SPEED 24
GRAPH 3:-WING AREA VS SPEED 25
GRAPH 4:-EMPTY WEIGHT VS SPEED 25
GRAPH 5:-RATE OF CLIMB VS SPEED 26
GRAPH 6:-MAX. TAKE OFF WT VS SPEED 26
GRAPH 7:- LOADED WEIGHT VS SPEED 27
GRAPH 8:-HEIGHT VS SPEED 27
GRAPH 9:-COMBAT RADIUS VS SPEED 28
GRAPH 10:-RANGE VS SPEED 28
GRAPH 11:-SERVICE CEILLING VS SPEED 29
GRAPH 12:-WING LOADING VS SPEED 29
GRAPH 13:-THRUST/WEIGHT VS SPEED 30
GRAPH 14:-THRUST VS SPEED 30
GRAPH 15:-ASPECT RATIO VS SPEED 31
GRAPH 16:-EMPTTY WT VS SPEED 35
GRAPH 17:-CL VS CD & CL VS ALPHA 37
LIST OF TABLE
TABLE 1:-COMPARATIVE DETAILS-1 18
TABLE 6:-OPTIMISE VALUES 32

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List of Symbols and Abbreviations
 - Angle of attack
Β - Climb angle
ρ - Density factor
ϒ - Dihedral angle
Ф - Glide angle
Θ - Turn angle
b - Wing span
c - Chord length
ĉ - Mean chord
CD - Drag coefficient
CD0 - Zero lift drag co-efficient
Cl - Rolling moment coefficient
Clf - Function of airfoil chord over which the flow in laminar
CLmax - Maximum Lift coefficient
Cr - Root chord
cT - Tip chord
D - Drag force
d - Tire diameter
E - Endurance
E - Oswald efficiency factor
G - Acceleration due to gravity
L - Lift force
LE - Leading edge of wing

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Lf - Length of fuselage
Q - Dynamic pressure
R - Turn radius
R/C - Rate of climb
Rr - Rolling radius of tyre
t/c - Wing thickness ratio
T/W - Thrust loading
V - Velocity of air/aircraft
Vcruise - Velocity at cruise
Vf - Volume of fuel
Vstall - Velocity at stall
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
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 2
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

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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 3
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.
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

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and technology. An example of the former is the design of most commercial transports,
starting at least with the Douglas DC-1 in 1932, 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 type(s) and amount of payload
-off distance at the maximum weight

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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
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 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

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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
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.
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

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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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2. INTRODUCTION TO AIRCRAFT
A fighter aircraft is a military aircraft designed primarily for air-to-air combat
against other aircraft, as opposed to bombers and attack aircraft, whose main mission is to
attack ground targets. The hallmarks of a fighter are its speed, maneuverability, and small
size relative to other combat aircraft.
Many fighters have secondary ground-attack capabilities, and some are designed as dual-
purpose fighter-bombers; often aircraft that do not fulfill the standard definition are called
fighters. This may be for political or national security reasons, for advertising purposes, or
other reasons.
A fighter's main purpose is to establish air superiority over a battlefield. Since World War I,
achieving and maintaining air superiority has been considered essential for victory in
conventional warfare .The success or failure of a belligerent's efforts to gain air supremacy
hinges on several factors including the skill of its pilots, the tactical soundness of its doctrine
for deploying its fighters, and the numbers and performance of those fighters. Because of the
importance of air superiority, since the dawn of aerial combat armed forces have constantly
competed to develop technologically superior fighters and to deploy these fighters in greater
numbers, and fielding a viable fighter fleet consumes a substantial proportion of the defense
budgets of modern armed forces
Fighter effectiveness criteria, energy-maneuverability, and OODA loop
Highly effective fighter aircraft design has been recognized since the 1970s to be based upon
four main generation transcending criteria. These criteria in order of importance are:
1. Achieve superiority in the element of surprise, meaning the tendency to surprise the enemy
more often than being surprised by the enemy. Surprise is the most important advantage any
fighter can have since historically in about 80% of air-to-air kills the victim was unaware of
the attacker until too late. Surprise advantage is mostly based upon small visual and radar
signatures, having good visibility out of the cockpit, having little or no engine smoke, and
having higher cruise speed in order to come on the enemy from behind rather than vice-versa.
In more modern fighters the tail-less delta wing configuration provides a higher cruising
speed advantage to better support surprise. Smaller light fighters have tended to statistically

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enjoy the element of surprise more than heavy fighters due to smaller visual and radar
signatures. A small fighter like the Northrop F-5 with a planform area of about 300 square
feet, or the F-16 at about 400 square feet, compared to about 1050 square feet for the F-15,
has a much lower visual profile. The small fighter is typically invisible to opposing pilots
beyond about 4 miles, whereas a larger fighter such as the F-15 is visible to about 7 miles,
and much farther if the engines smoke.Additionally, smaller targets take longer to visually
acquire even if they are visible. These two factors together give the pilot of the smaller
fighter better statistical odds of seeing a larger fighter first.[44]
Once the small fighter sees and
turns towards the opponent its very small frontal area reduces maximum visual detection
range to about 2.0 to 2.5 miles. If not detected by radar (common when attacking from the
rear), this allows the small fighter to set up a high reliability short range heat-seeking missile
shot by ambush while still invisible to the target. This effect is so pronounced that even
elderly subsonic fighters can often use the element of surprise to defeat more advanced
supersonic fighters. For example, in the 1965 U.S. Featherduster trials the 1st generation F-86
was found to statistically dominate over the 3rd generation Mach 2 F-4 Phantom by superior
surprise.
A surprise advantage that can favor larger fighters is longer radar range. Given similar
technology, smaller fighters typically have about two thirds the radar range against the same
target as heavy fighters. However, this cannot always be counted upon to give the large
fighter a winning advantage, as larger fighters with typical radar cross sectional area of about
10 m² are detectable by a given radar at about 50% farther range than the 2m² to 3m² cross
section of the light fighter. This approximately balances these trade-offs, and can sometimes
favor the lightweight fighter. Also, airborne fighter radars are limited in coverage zone (front
only) and are far from perfect in detecting enemy aircraft. Despite extensive use of radar in
the Vietnam War by the United States, only 18% of North Vietnamese fighters were first
detected by radar of any kind, and only 3% by air-to-air radar on board a fighter aircraft. The
other 82% were visually acquired, which explains why visual signature favoring the smaller
fighters has remained a significant advantage. The modern trend to stealth aircraft is an
attempt to maximize surprise in an era when Beyond Visual Range (BVR) missiles are
becoming more effective than the quite low effectiveness BVR has had in the past.The cost of
stealth has so far in its history limited it to more expensive fighters, but as it becomes more
commoditized and available it will likely become a prominent feature of advanced lower cost

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fighters as well. For example, as of 2016 India is planning a semi-stealthy version of its cost
effective HAL Tejas lightweight fighter.
"The F-22 costs 10 times as much as an early model F-16 fighter and, due to its huge
maintenance load, can fly only half as many sorties per day. Thus, for equal investment, the
F-22 delivers only one-twentieth as many airplanes over enemy territory as the F-16--a
crippling disadvantage, no matter whether the F-22’s stealth and weapons work or don’t
work."
Defense analyst and combat aircraft architect Pierre Sprey.
2. To have numerical superiority in the air, this implies the need for lower procurement cost,
lower maintenance cost, and higher reliability. Having high sortie generation rates, and long
range and combat persistence, are also strong contributors to having superior numbers at the
time and place that combat occurs.
3. To have superior maneuverability, which in maneuvering combat allows getting into
superior position to fire and score the kill. This is a function of achieving lower wing loading,
higher thrust to weight ratio, and superior aerodynamics. The concepts of Energy–
maneuverability theory (see below) are key in modern fighter design in achieving superior
maneuverability.
4. To have superior weapon systems effectiveness, as described in detail below in fighter
weapons. Larger fighters have the benefit of carrying a larger weapons load. However,
combat experience shows that weapons systems "effectiveness" has not been dominated by
the amount of weaponry, but by the ability to reliably achieve split second kills when in
position to do so.

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3. COMPARATIVE DETAILS
In the designer’s perspective it is necessary to compare the existing airplanes that are of
same type as that of our desired airplane. Their important parameters, positive aspects and
pitfalls to be overcome are taken into consideration. The data has been collected for military
transport aircrafts. Several parameters are compared for 15 aircrafts and different parameters
are plotted on the graph. The parameters compared are:
 Seating capacity
 Length
 Wingspan
 Height
 Wing loading
 Aspect ratio
 Service ceiling
 Rate of climb
 Range
 Cruise speed
 Maximum speed
 No of engines
 Maximum thrust capability
 Maximum take-off weight
 Empty weight
 Payload weight
 Engine type

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PARAMETER
UNITS SUKHOI SU-35
EUROFIGHTE
R TYPHOON
SUPERHORNE
T
CREW NUMBER 1 1 2
LENGTH m 21.9 15.96 18.31
WINGSPAN m 15.3 10.95 13.52
HEIGHT m 5.9 5.28 4.88
WING AREA m 2
62 51.2 46.5
EMPTY WT. Kg 18400 11000 14552
LOADED WT. Kg 25300 16000 21320
MAX.TAKEOF
F WT
Kg 34500 23500 29937
RATE OF
CLIMB
m/s 212 318 319
DRY THRUST KN 86.3 60 62.3
MAX SPEED Km/h 2500 2495 19151
COMBAT
RADIUS
Km 1580 1389 1200
RANGE Km 4500 3790 3330
SERVICE
CEILING
M 18000 19812 15001
WING
LOADING
Kg/m2
408 312 459
THRUST/WEIG
HT
- 1.13 1.15 0.93
POWERPLNT
-
AF TURBO
FAN
AF TURBO
FAN
2*TURBOFAN
ASPECTRATIO - 3.77 2.19 4
Table 1

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PARAMETER
UNITS
MITSUBISHI
F-15
CHENGDU J-
20
X-15
CREW NUMBER 1 1 1
LENGTH m 19.43 20 15.45
WINGSPAN m 13.065 13 6.8
HEIGHT m 5.63 4.45 4.12
WING AREA m2
56.5 78 18.6
EMPTY WT. Kg 12700 19391 6620
LOADED WT. Kg 20200 32092 15420
MAX.TAKEOF
F WT
Kg 30845 36288 30000
RATE OF
CLIMB
m/s 254 304 375
DRY THRUST KN 77.62 76.18 167
MAX SPEED Km/h 2665 2100 3794
COMBAT
RADIUS
Km 1524 1252 1230
RANGE Km 4700 3400 450
SERVICE
CEILING
m 20000 1800 25000
WING
LOADING
Kg/m2
73.1 410 170
THRUST/WEIG
HT
- 1.12 0.85 1.85
POWERPLNT
-
AF TURBO
FAN
AF
TURBOFAN
AF TURBO
FAN
ASPECTRATIO - 3.04 2.1 2.35
Table 2:-2

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PARAMETER
UNITS F-22 RAPTOR
LOCKHEED
YF-12
DASSAULT
RAFALE
CREW NUMBER 1 2 2
LENGTH m 18.92 30.97 1527
WINGSPAN m 13.56 16.9 10.8
HEIGHT m 5.08 5.65 5.34
WING AREA m2
78.04 167 45.7
EMPTY WT. Kg 19700 27604 10300
LOADED WT. Kg 294410 56200 15000
MAX.TAKEOF
F WT
Kg 38000 63504 24500
RATE OF
CLIMB
m/s 325 200 245
DRY THRUST KN 116 91.2 50.4
MAX SPEED Km/h 2410 3661 1912
COMBAT
RADIUS
Km 1188 1475 1852
RANGE Km 3220 4800 3700
SERVICE
CEILING
m 20000 27400 15235
WING
LOADING
Kg/m2
377 336.58 328
THRUST/WEIG
HT
- 1.08 0.44 0.988
POWERPLNT
- AFTURBOFAN
AFTURBO
FAN
2TURBOFAN
ASPECTRATIO - 2.35 1.7 2.55
Table 3

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PARAMETER
UNITS HAL TEJAS
MIKOYAN
MIG-31
F-35
LIGHTING-2
CREW NUMBER 2 2 1
LENGTH m 13.2 22.69 15.67
WINGSPAN m 8.2 13.49 10.7
HEIGHT m 4.4 6.15 4.33
WING AREA m2
38.4 61.6 42.7
EMPTY WT. Kg 6560 21821 13199
LOADED WT. Kg 9500 41000 22470
MAX.TAKEOF
F WT
Kg 13500 46200 31800
RATE OF
CLIMB
m/s 150 350 350
DRY THRUST KN 53.6 93 125
MAX SPEED Km/h 2205 3000 1930
COMBAT
RADIUS
Km 11650 1320 1158
RANGE Km 3000 3000 2220
SERVICE
CEILING
m 16000 20600 18288
WING
LOADING
Kg/m2
2.47 665 107.7
THRUST/WEIG
HT
- 1.07 0.85 0.87
POWERPLNT
- TURBBOFAN
AF
TURBOFAN
AF TURBO
FAN
ASPECTRATIO - 2.74 2.94 2.68
Table 4

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PARAMETER
UNITS F-16 SAAB JAS-39
SUKHOI SU-
30MKI
CREW NUMBER 1 2 2
LENGTH m 15.06 14.1 21.935
WINGSPAN m 9.96 8.4 14.7
HEIGHT m 4.8 4.5 6.36
WING AREA m 27.87 30 62
EMPTY WT. Kg 8570 6800 18400
LOADED WT. Kg 12000 8500 26090
MAX.TAKEOF
F WT
Kg 19200 14000 38800
RATE OF
CLIMB
m/s 250 250 300
DRY THRUST KN 76.3 54 64
MAX SPEED Kmph 2414 2204 2100
COMBAT
RADIUS
Km 550 800 963
RANGE Km 4220 3200 3000
SERVICE
CEILING
m 15240 15240 17300
WING
LOADING
Kg/m2
431 283 401
THRUST/WEIG
HT
N/A 1.095 0.97 0.96
POWERPLNT
N/A
AF TURBO
FAN
AF TURBO
FAN
AF
TURBOFAN
ASPECTRATIO N/A 3.09 1.8 3.5
Table 5

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4. COMPARITIVE GRAPHS
The comparative graphs are plotted for various parameters from the comparative sheets of
various military transport aircrafts. Comparison of data’s of similar aircrafts of same
classification are necessary to know the strengths and weakness of every aircraft and develop
an aircraft such that it would be more superior than other aircrafts of the same classification.
Using these graphs, the mean value from each graph is recorded with the help of a circle. The
circle should be plotted such that most no of points are covered by the circle. The circle with
same dimensions as that of the first graph is to be used in all other graphs. The co-ordinates
of the centre of the circle gives us the mean value of the parameters that are plotted. Here,
using the cruise speed of the aircraft, other parameters are compared. The list of graphs that
are used to compare the parameters are:
1. Length vs. max speed
2. Height vs. max speed
3. Fuselage diameter vs. max speed
4. Wing span vs. max speed
5. Wing area vs. max speed
6. Wing loading vs. max speed
7. Aspect ratio vs. max speed
8. Service ceiling vs. max speed
9. Rate of climb vs. max speed
10. Range vs. max speed
11. Maximum speed vs. max speed
12. No of engines vs. max speed
13. Max thrust capability vs. max speed
14. Max take-off weight vs. max speed
15. Empty weight vs. max speed
16. Payload weight vs. max speed

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Graph 1
Graph 2
0
5
10
15
20
0 1000 2000 3000 4000
WING
SPAN
SPEED
WING SPAN VS SPEED
WING SPAN

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Graph 3
Graph 4
0
20
40
60
80
100
120
140
160
180
0 1000 2000 3000 4000
WING
AREA
SPEED
WING AREA VS SPEED
WING AREA
0
5000
10000
15000
20000
25000
30000
0 1000 2000 3000 4000
EMPTY
WEIGHT
SPEED
EMPTY WEIGHT VS SPEED
EMPTY WEIGHT

26
Graph 5
Graph 6
0
50
100
150
200
250
300
350
400
0 500 1000 1500 2000 2500 3000 3500 4000
RATE
OF
CLIMB
SPEED
RATE OF CLIMB
RATE OF CLIMB
0
10000
20000
30000
40000
50000
60000
70000
0 500 1000 1500 2000 2500 3000 3500 4000
MAX
TAKE
OFF
WT
SPEED
MAX TAKE OFF WT VS SPEED

27
Graph 7
Graph 8
0
10000
20000
30000
40000
50000
60000
0 1000 2000 3000 4000
LOADED
WT
SPEED
LOADED WEIGHT VS SPEED
LOADED WEIGHT
0
1
2
3
4
5
6
7
0 1000 2000 3000 4000
HEIGHT
SPEED
HEIGHT VS SPEED

28
Graph 9
Graph 10
0
200
400
600
800
1000
1200
1400
1600
1800
2000
0 1000 2000 3000 4000
COMBAT
RADIUS
SPEED
COMBATRADIUS VS SPEED
0
1000
2000
3000
4000
5000
6000
0 500 1000 1500 2000 2500 3000 3500 4000
RANGE
SPEED
RANGE VS SPEED
RANGE

29
Graph 11
Graph 12
0
5000
10000
15000
20000
25000
30000
0 1000 2000 3000 4000
SERVICE
CEILING
SPEED
SERVICE CEILING VS SPEED
SERVICE CEILING
0
100
200
300
400
500
600
700
0 1000 2000 3000 4000
WING
LOADING
SPEED
WINGLOADING VS SPEED

30
Graph 13
Graph 14
0
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
2
0 1000 2000 3000 4000
THRUST/
WT
SPEED
THRUST/WT VS SPEED
THRUST/WEIGHT
0
20
40
60
80
100
120
140
160
180
0 500 1000 1500 2000 2500 3000 3500 4000
T
H
R
U
S
T
SPEED
THRUST VS SPEED

31
Graph 15
0
0.5
1
1.5
2
2.5
3
3.5
4
4.5
0 1000 2000 3000 4000
ASPECT
RATIO
SPEED
ASPECTRATIO VS SPEED

32
OPTIMISED VALUES
PARAMETERS
FROM GRAPH UNITS
HEIGHT
5
m
WING AREA
50 m2
EMPTY WEIGHT
15000 Kg
LOADED WT
20000 Kg
MWEIGHTAX. TAKE
OFF WT.
30000 Kg
RATE OF CLIMB
275 m/s
DRY THRUST 76.07 KN
MAX. SPEED 2400 Km/h
COMBAT RADIUS 1000 Km
RANGE 3700 Km
SERVICE CEILING 18000 m
WING LOADING 360 Kg/m2
THRUST/WEIGHT 0.9 -
POWERPLANT TURBOFAN -
ASPECT RATIO 2.05 -
COST 240 M
CREW 2 Number
LENGTH 18 m
Table 6

33
5. WEIGHT ESTIMATION:
The estimation of the weight of a conceptual aircraft is a critical part of the design process.
The weights engineer interfaces with all other engineering groups, and serves as the "referee"
during the design evolution. Weights analysis per se does not form part of the aerospace
engineering curriculum at most universities. It requires a broad background in aerospace
structures, mechanical engineering, statistics, and other engineering disciplines. There are
many levels of weights analysis. Previous chapters have presented crude statistical techniques
for estimating the empty weight for a given take off weight. These techniques estimate the
empty weight directly and are only suitable for "first-pass" analysis. More sophisticated
weights methods estimate the weight of the various components of the aircraft and then sum
for the total empty weight. In this chapter, two levels of component weights analysis will be
presented. The first is a crude component builds up based upon plan form areas, wetted areas,
and per cents of gross weight. This technique is useful for initial balance calculations and can
be used to check the results of the more detailed statistical methods. The second uses detailed
statistical equations for the various components. This technique is sufficiently detailed to
provide a credible estimate of the weights of the major component groups.
The takeoff gross weight-the sum of the empty weight and the useful load-reflects the weight
at takeoff for the normal design mission. The flight design gross weight represents the aircraft
weight at which the structure will withstand the design load factors. Usually this is the same
as the takeoff weight, but some aircraft are designed assuming that maximum loads will not
be reached until the aircraft has taken off and climbed to altitude, burning off some fuel in the
process.
Gross weight W0 = Wcrew + Wpayload + Wfuel + Wempty
Wcrew =200lbs
=92kg
Wpayload=8000kgs
(Operating Empty weight) WOE = WE + WINST (10%*WTO)
(Fuel weight) WF = (USED)+ (RESERVED )(10-25%of used)
(Fuel Fraction) MFF= ( * )

34
(USED)=(1-MFF)*WTD
(RESERVED)=0.25* (USED)
WF= (1-MFF) *1.25 WTO
Payload Weight WPL 8000kgs
(Crew weight) W crew = Cockpit + Cabin crew
=200lbs
Error percentage= (WOE th-WOE grp/WOE th )*100
MFF= ( * )
PHASE 5(Cruise Brequet Eqtn)
Jet =(V/CJ)CR(L/D)CR ln(W4/W5)
W5/W4= 0.999
PHASE 6(Loiter)
Jet ELTR=(1/CJ)LTR*(L/D)LTR ln(W5/W6)LTR
W6/W5 = 0.94
MISSION PROFILE
1- Warm-up
2- Taxi
3- Take-off
4- Accelerate
5- Climb to altitude
6- Outbound Cruise

35
7- Descent
8- Manoeuvre
9- Weapon drop
10- Climb to altitude
11- Descent
12- Loiter
13- Landing
14- Shut Down
15- Towing
MFF= ( * )
WOE=WTO-WPL-WF-WCREW-Winst
WE = 30000-3300-8000-92-3000
WETh =15608kg
WEGRH= 15000kg
Percentage error= (15608-15000)/15608 *100
=3.89 %
Graph 16
RESULT:
WETh =15608kg
WEGRH= 15000kg
% error=3.98%
0
5000
10000
15000
20000
25000
30000
0 20000 40000 60000 80000
EMPTY
WT
MAX TAKE OFF WT
EMPTY WT VS MAX TAKE OFF WT

36
6.AIRFOIL AND WING SELECTION
Airfoil selection:
W = L =  v2
S CL Wing
CL Wing = [

] =

where, V = stall velocity
S = wing area
 = density at altitude (service ceiling)
= wing loading
CL Wing = Cl Airfoil = [

]
= 1.43
Figure 1
The first family of NACA airfoil sections, developed in the 1930s, was the "four-digit" series.
Following are some definitions of airfoil section characteristics, followed by a description of
how the NACA "four-digit" series specifies these characteristics.

37
The mean camber line is the locus of points midway between the upper and lower surfaces as
measured perpendicular to the chord line.
The most forward point of the mean camber line is the leading edge.
The most rearward point of the mean camber line is the trailing edge.
The straight line connecting the leading and trailing edges is the chord line of the airfoil.
The actual distance between the leading and trailing edges, measured along the chord line, is
the chord, c.
The maximum camber is the maximum distance between the mean camber line and the chord
line, measured perpendicular to the chord line.
Graph 17
The thickness is the distance between the upper and lower surfaces, also measured
perpendicular to the chord line. Having defined these fundamental characteristics, additional
properties are now defined, before proceeding to an explanation of the NACA system.
Because airfoil sections vary in size, the following properties are generally stated in terms of
the chord, c. The shape of the airfoil section at the leading edge is usually circular, with a
radius of approximately 0.02c.
The digits in NACA's four digit numbering system are defined as follows:

38
 the first digit denotes the maximum camber, Cmax, as a percent of the chord;
 the second digit denotes the chordwise position of the maximum camber, XCmax, in
tenths of the chord;
 the last two digits denotes the maximum thickness of the airfoil section, t, as a percent
of the chord.
WING SELECTION
Low wing with dihedral
Figure 2
Low wing
A low wing is one which is located on or near the base of the fuselage.
Placing the wing low down allows good visibility upwards and frees up the central fuselage
from the wing spar carry-through. By reducing pendulum stability, it makes the aircraft more
manoeuvrable, as on the Spitfire; but aircraft that value stability over manoeuvrability may
then need some dihedral. A low wing allows a lighter structure because the fuselage sides
carry no additional loads, and the main undercarriage legs can be made shorter.
A feature of the low wing position is its significant ground effect, giving the plane a tendency
to float further before landing. Conversely, this very ground effect permits shorter takeoffs.
The low wing configuration has proved particularly suitable for passenger jetliners
Dihedral angle and dihedral effect

39
Dihedral Angle is the upward angle from horizontal of the wings of a fixed-wing aircraft, or
of any paired nominally-horizontal surfaces on any aircraft. The term can also apply to the
wings of a bird. Dihedral Angle is also used in some types of kites such as box kites. Wings
with more than one Angle change along the full span are said to be polyhedral.
Dihedral Angle has important stabilizing effects on flying bodies because it has a strong
influence on the dihedral effect.
Dihedral effect of an aircraft is a rolling moment resulting from the vehicle having a non-
zero angle of sideslip. Increasing the dihedral angle of an aircraft increases the dihedral effect
on it. However, many other aircraft parameters also have a strong influence on dihedral
effect. Some of these important factors are: wing sweep, vertical center of gravity, and the
height and size of anything on an aircraft that changes it’s sidewards force
as sideslip changes.
Figure 3
Aspect ratio, AR = = 2.64
Taper ratio, λ =
For rectangle wing, λ = 1
For elliptical wing, λ = 0.9
For tapered wing, λ = 0.5
Croot =
=5.79m
Ctip = λ (Croot)

40
= 0.5 *5.79
= 2.89m
Sweep angle (from graph) = 31.8º
Mean aerodynamic centre
ŷ = [ ]
= 5.75m
ĉ = Croot
= 4.503m
Figure 4

41
7.TAIL PLANE SELECTION
A tail plane, also known as a horizontal stabiliser, is a small lifting surface located on
the tail (empennage) behind the main lifting surfaces of a fixed-wing aircraft as well as other
non-fixed-wing aircraft such as helicopters and gyroplanes. Not all fixed-wing aircraft have
tail planes. Canards, tailless and flying wing aircraft have no separate tail plane, while in v-
tail aircraft the vertical stabilizer, rudder, and the tail-plane and elevator are combined to
form two diagonal surfaces in a V layout.
The function of the tail plane is to provide stability and control. In particular, the tail
plane helps adjust for changes in the centre of pressure or centre of gravity caused by changes
in speed and attitude, fuel consumption, or dropping cargo or payload.
9.2 Types of tail configuration:
Fig. 9.1: Types of tail configurations

42
Conventional tail:
The vertical stabilizer is mounted exactly vertically, and the horizontal stabilizer is
directly mounted to the empennage (the rear fuselage). This is the most common vertical
stabilizer configuration.
T-tail:
A T-tail has the horizontal stabilizer mounted at the top of the vertical stabilizer. It is
commonly seen on rear-engine aircraft, such as the Bombardier CRJ200, the Fokker 70,
the Boeing 727, the Vickers VC10 and Douglas DC-9, and most high-performance gliders.T-
tails are often incorporated on configurations with fuselage mounted engines to keep the
horizontal stabilizer away from the engine exhaust plume.T-tail aircraft are more susceptible
to pitch-up at high angles of attack. This pitch-up results from a reduction in the horizontal
stabilizer's lifting capability as it passes through the wake of the wing at moderate angles of
attack. This can also result in a deep stall condition.T-tails present structural challenges since
loads on the horizontal stabilizer must be transmitted through the vertical tail.
Cruciform tail:
The cruciform tail is arranged like a cross, the most common configuration having the
horizontal stabilizer intersecting the vertical tail somewhere near the middle. The PBY
Catalina uses this configuration. The "push-pull" twin engine Dornier Do 335 World War II
German fighter used a cruciform tail consisting of four separate surfaces, arranged in dorsal,
ventral, and both horizontal locations, to form its cruciform tail, just forward of the rear
propeller.Falcon jets from Dassault always have cruciform tail.
Twin tail:
Rather than a single vertical stabilizer, a twin tail has two. These are vertically
arranged, and intersect or are mounted to the ends of the horizontal stabilizer. The Beechcraft
Model 18 and many modern military aircraft such as the American F-14, F-15, and F/A-
18 use this configuration. The F/A-18, F-22 Raptor, and F-35 Lightning II have tailfins that
are canted outward, to the point that they have some authority as horizontal control surfaces;
both aircraft are designed to deflect their rudders inward during takeoff to increase pitching
moment. A twin tail may be either H-tail, twin fin/rudder construction attached to a single
fuselage such as North American B-25 Mitchell or Avro Lancaster, or twin boom tail, the
rear airframe consisting of two separate fuselages each sporting one single fin/rudder, such as
Lockheed P-38 Lightning or C-119 Boxcar.

43
Triple tail:
A variation on the twin tail, it has three vertical stabilizers. An example of this
configuration is the Lockheed Constellation. On the Constellation it was done to give the
airplane maximum vertical stabilizer area while keeping the overall height low enough so that
it could fit into maintenance hangars.
V-tail:
A V-tail has no distinct vertical or horizontal stabilizers. Rather, they are merged into
control surfaces known as ruddevators which control both pitch and yaw. The arrangement
looks like the letter V, and is also known as a butterfly tail. The Beechcraft Bonanza Model
35 uses this configuration, as does the F-117 Nighthawk, and many of Richard Schreder's HP
series of homebuilt gliders.
TWIN TAIL
A twin tail is a specific type of vertical stabilizer arrangement found on the
empennage of some aircraft. Two vertical stabilizers—often smaller on their own than a
single conventional tail would be—are mounted at the outside of the aircraft's horizontal
stabilizer. This arrangement is also known as an H-tail,[1]
as it resembles a capital "H" when
viewed from rear - these were used on a wide variety of World War II multi-engine designs
that saw mass production, especially on the American B-24 Liberator and B-25 Mitchell
bombers, the British Avro Lancaster and Handley-Page Halifax heavy bombers, and on the
Soviet Union's Petlyakov Pe-2 attack bomber.
A special case of twin tail is twin boom tail or double tail where the aft airframe
consists of two separate fuselages, "tail booms", which each have a rudder but are usually
connected by a single horizontal stabilizer. Examples of this construction are the twin-
engined Lockheed P-38 Lightning; Northrop P-61 Black Widow; Focke-Wulf Fw 189; the
single jet-engined de Havilland Vampire; cargo-carrying Fairchild C-119 Flying Boxcar and
the little known Transavia PL-12 Airtruk
DESIGN
Separating the control surfaces allows for additional rudder area or vertical surface
without requiring a massive single tail. On multi-engine propeller designs twin fin and
rudders operating in the propeller slipstream give greater rudder authority and improved
control at low airspeeds, and when taxiing. A twin tail can also simplify hangar requirements,

44
give dorsal gunners enhanced firing area, and in some cases reduce the aircraft's weight. It
also affords a degree of redundancy—if one tail is damaged, the other may remain functional.
Most often, the twin vertical surfaces are attached to the ends of the horizontal stabilizer, but
a few aircraft in aviation history—like the Armstrong Whitworth Whitley, Mitsubishi G3M
and Dornier Do 19 bombers, had their twin vertical surfaces mounted to the upper surface of
the fixed stabilizer instead, at some distance inwards from the horizontal stabilizer's tips.
Many canard aircraft designs incorporate twin tails on the tips of the main wing. Very
occasionally, three or more tails are used, as on the Breguet Deux-Ponts, Lockheed
Constellation and Boeing 314 Clipper. A very unusual design can be seen on the E-2
Hawkeye, which has two additional vertical tails fixed to the horizontal stabilizer between the
normal vertical twin-tail surfaces. This arrangement was chosen for the stringent size
limitations of carrier-based aircraft. Significant aircraft with twin tails include the
Consolidated B-24 Liberator, Handley-Page Halifax, Avro Lancaster, and P-38 Lightning.
The arrangement is not limited to World War II-vintage aircraft, however. Many fighter
aircraft, like the F-14 Tomcat, F-15 Eagle, Sukhoi Su-27, Mig-29, and A-10 Thunderbolt II,
make use of twin tail configurations, as do civilian and cargo designs like the Antonov An-
14, Antonov An-22, Antonov An-28, Antonov An-38.

45
8.LANDING GEAR SELECTION
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 overbalancing during take-off 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.
Tricycle gear is a type of aircraft undercarriage, or landing gear, arranged in a tricycle
fashion. The tricycle arrangement has a single nose wheel in the front, and two or more main
wheels slightly aft of the center of gravity. Tricycle gear aircraft are the easiest to take-off,
land and taxi, and consequently the configuration is the most widely used on aircraft.
Several early aircraft had primitive tricycle gear, notably very early Antoinette planes and the
Curtiss Pushers of the pre-World War I Pioneer Era of aviation. Waldo Waterman's 1929
tailless Whatsit was one of the first to have a steerable nose wheel

46
Figure 5
A Mooney M20J with a retractable tricycle landing gear
Polish 3Xtrim 3X55 Trener with a fixed tricycle landing gear taxiing..
Tricycle gear and taildraggers compared
Tricycle gear is essentially the reverse of conventional landing gear or taildragger. On the
ground, tricycle aircraft have a visibility advantage for the pilot as the nose of the aircraft is
level, whereas the high nose of the taildragger can block the view ahead. Tricycle gear
aircraft are much less liable to 'nose over' as can happen if a taildragger hits a bump or has the
brakes heavily applied. In a nose-over, the aircraft's tail rises and the propeller strikes the
ground, causing damage. The tricycle layout reduces the possibility of a ground loop, because
the main gear lies behind the center of mass. However, tricycle aircraft can be susceptible to
wheel-barrowing. The nosewheel equipped aircraft also is easier to handle on the ground in

47
high winds due to its wing negative angle of attack. Student pilots are able to safely master
nosewheel equipped aircraft more quickly.[2]
Tricycle gear aircraft are easier to land because the attitude required to land on the main gear
is the same as that required in the flare, and they are less vulnerable to crosswinds. As a
result, the majority of modern aircraft are fitted with tricycle gear. Almost all jet-powered
aircraft have been fitted with tricycle landing gear, to avoid the blast of hot, high-speed gases
causing damage to the ground surface, in particular runways and taxiways. The few
exceptions have included the Yakovlev Yak-15, the Supermarine Attacker, and prototypes
such as the Heinkel He 178, the first four prototypes (V1 through V4) of the Messerschmitt
Me 262, and the Nene powered version of the Vickers VC.1 Viking. Outside of the United
States — where the tricycle undercarriage had solidly begun to take root with its aircraft
firms before that nation's World War II involvement at the end of 1941 — the Heinkel firm in
World War II Germany began building airframe designs meant to use tricycle undercarriage
systems from their beginnings, as early as late 1939 with the Heinkel He 280 pioneering jet
fighter demonstrator series, and the unexpectedly successful Heinkel He 219 twin-engined
night fighter of 1942 origin.[4]
A Cessna 150 taildragger.
The taildragger configuration has its own advantages, and is arguably more suited to rougher
landing strips. The tailwheel makes the plane sit naturally in a nose-up attitude when on the
ground, which is useful for operations on unpaved gravel surfaces where debris could damage
the propeller. The tailwheel also transmits loads to the airframe in a way much less likely to
cause airframe damage when operating on rough fields. The small tailwheel is much lighter
and much less vulnerable than a nosewheel. Also, a fixed-gear taildragger exhibits less
interference drag and form drag in flight than a fixed-gear tricycle aircraft whose nosewheel
may sit directly in the propeller's slipstream. Tailwheels are smaller and cheaper to buy and
to maintain, and manhandling a tailwheel aircraft on the ground is easier. Most tailwheel
aircraft are lower in overall height and thus may fit in lower hangars. Tailwheel aircraft are
also more suitable for fitting with skis in wintertime.[2]

48
9.ENGINE SPECIFICATIONS
F16 FIGHTING FALCON (GE F110) :
Type: Turbofan
Natural Origin: United States
Manufacturer: General Electric
First run: 1980
Major Application: General Dynamics
F16 Fighting Falcon
Grumman F14 Tomcat
McDonnel Douglas F15E
Strike Eagle
Developed from: General Electric F101
Variants: General Electric F118
Specifications
o Type: Afterburning Turbofan
o Length: 463-590 cms
o Diameter: 118 cms
o Dry weight: 1778-1996 kg
o Compressor: 2 spool:3fan,9hp stage
o Combustors: annular
o Turbine: 2 LP and 1 HP stages
Performance:-
o Max Thrust: 16610lbf (76.3 KN)
o Turbine inlet temperature: 2750F (1510C)
o Thrust to weight: 129:7.29
DISCRIPTION:-
F-14
The F-14A entered service with the United States Navy in 1973 powered by Pratt & Whitney
TF30s. By the end of the decade, following numerous problems with the original engine (and
similar problems with the Pratt & Whitney F100 on the F-15 and F-16), the DoD began
procuring the upgraded TF30-P-414As. While these engines solved the serviceability

49
problems, the fuel consumption and thrust was comparable to the initial model–considerably
less than what the F-14 had been designed for.
In 1979, a derivative of the GE F101 turbofan called the F101-X was selected to power the F-
14 and was later designated the F110-GE-400. The primary difference between the F110-GE-
400 and the F110-GE-100 is length - the F110-GE-400 has a 50-inch (1.3 m) tailpipe
extension to suit the F-14 airframe, which is fitted downstream of the augmentor (afterburner
section). The F110-GE-400 engine produced 23,400 lbf (104 kN) of thrust with afterburner at
sea level, which rose to 30,200 lbf (134 kN) at Mach 0.9.[3]
This provided a significant
increase over the TF30's maximum thrust of 20,900 lbf (93 kN).[4]
These upgraded jets were
known as F-14Bs, as were production aircraft powered by the F110. The same engine also
powers the final variant of the aircraft, the F-14D.
F-16
The F-16 Fighting Falcon entered service powered by the Pratt & Whitney F100 afterburning
turbofan. Seeking a way to drive unit costs down, the USAF implemented the Alternative
Fighter Engine (AFE) program in 1984, under which the engine contract would be awarded
through competition. The F110 currently powers 86% of the USAF F-16C/Ds (June 2005).
The F110-GE-100 provides around 4,000 lbf (17.8 kN) more thrust than the F100-PW-200
and requires more air, which led to the increase in the area of the engine intake. The F-16C/D
Block 30/32s were the first to be built with a common engine bay, able to accept both
engines, with block 30s having the bigger intake (known as "Big Mouth") and block 32s
retaining the standard intake.
Initial orders were for the F110-GE-100 rated at 28,000 lbf (125 kN). Later versions of the
F110 include the F110-GE-129 delivering 29,400 lbf (131 kN) thrust and the F110-GE-132
delivering 32,000 lbS

50
F-15
An F110 engine undergoes performance testing at the Air Force's Arnold Engineering
Development Center.
Two F110-GE-129 engines, with 29,400 lbf (131 kN) of thrust, power 40 F-15K fighters of
South Korea. This is the first time production F-15s will be powered by a GE engine, since
all previous F-15 models were powered by Pratt and Whitney. The GE engines will be
manufactured through a joint licensing agreement with Samsung Techwin Company. It has
also been chosen by the Republic of Singapore Air Force (RSAF) to power its F-15SG, and
Saudi Arabia to power its F-15SA.

51
10.V-N DIAGRAM:-
V– n Diagram Flight regime of any aircraft includes all permissible combinations of
speeds, altitudes, weights, centres of gravity, and configurations. This regime is shaped by
aerodynamics, propulsion, structure, and dynamics of aircraft. The borders of this flight
regime are called flight envelope or manoeuvring envelope. The safety of human onboard is
guaranteed by aircraft designer and manufacturer. Pilots are always trained and warned
through flight instruction manual not to fly out of flight envelope, since the aircraft is not
stable, or not controllable or not structurally strong enough outside the boundaries of flight
envelope. A mishap or crash is expected, if an aircraft is flown outside flight envelope.
Pilots are using several graphs and charts in their flight operations. Four important envelopes
are as follows:
1. Diagram of variations of aircraft lift coefficient versus Mach number (CL – M)
2. Diagram of variations of airspeed versus altitude (V – h)
3. Diagram of variations of centre of gravity versus aircraft weight (X cg – W)
4. Diagram of variations of airspeed versus load factor (V – n)
One of the most important diagrams is referred to as flight envelope. This envelope
demonstrates the variations of airspeed versus load factor (V – n). In another word, it depicts
the aircraft limit load factor as a function of airspeed. One of the primary reasons that this
diagram is highly important is that, the maximum load factor; that is extracted from this
graph is a reference number in aircraft structural design. If the maximum load factor is under-
calculated, the aircraft cannot withstand flight load safely. For this reason, it is recommended
to structural engineers to recalculate the V-n diagram on their own as a safety factor.
Load Factor
The load to the aircraft on the ground is naturally produced by the gravity (i.e. 1
times g). But, there are other sources of load to the aircraft during flight; one of which is the
acceleration load. This load is usually normalized through load factor (i.e. "n" times g). In
another word, aircraft load is expressed as a multiple of the standard acceleration due to
gravity (g = 9.81 m/sec2
= 32.17 ft/sec2
). Recall that we defined the load factor as the ratio
between lift and weight.
where "a" is the centrifugal acceleration (V2
/R). As this acceleration increases; i.e.
airspeed increases or radius of turn decreases; the load factor will increase too. For other

52
flight operations, similar expressions can be drawn. In some instances; especially for
missiles; this load factor may get as high as 30. Hence, the structure must carry this huge load
safely. The aircraft structure must be strong enough to carry other loads including
acceleration load such that aircraft is able to perform its mission safely.
AIRCRAFT TYPE
MAXIMUM POSITIVE
LOAD FACTOR
MAXIMUM NEGATIVE
LOAD FACTOR
Acrobatic 6 -3
Homebuilt 5 -2
Transport 3 to 4 -1 to -2
Highly manoeuvrable 6.5 to 12 -3 to -6
Bomber 2 to 4 -1 to -2
Velocity:
Velocity is the major parameter we need to consider while estimating the safety of an
aircraft. Aircraft has different velocities at different stages of flight. These velocities plays a
crucial role in structural damages while in flight. The following are the velocities:
 Maximum velocity (VMAX)
 Corner velocity (V*
)
 Dive velocity (VD)
V-n diagram is the diagram between the various stages of velocity of aircraft in flight and the
load factors of an aircraft.
11.1 Plotting of V-n Diagram For Cargo Aircraft:
11.1.1 Calculation:

53
The positive load factor value is taken as n = 3.5 from the table
(CL)MAX = 1.5 (From airfoil tools website)
(CL)MAX (negative) = -1.0(From airfoil tools website)
n max(positive) = 3.5
The negative load factor value is given as
n max(negative) = 0.4 * nmax(positive)
= 0.4 *3.5
n max(negative) = 1.4
Maximum velocity for positive load factor is given as
Vmax = (2Wg/ρ S CLmax)1/2
Vmax = (2 x30000 x9.81/1.225 x50 x1.5)1/2
Vmax(positive) = 115.4m/s
Maximum velocity for negative load factor is given as
Vmax = (-2Wg/ρ S(-CLmax))1/2
Vmax = (2 x30000 x9.81/1.225 x50 x(-1.0))1/2
Vmax(negative) = 141.2m/s
The corner velocity for positive maximum value of load factor is given as
n max = (0.5ρ(V*)2
SCLmax/W g)
3.5 = (0.5x1.225x(V*)2
x50x1.5 /30000x9.81)
V*(positive) = 216.6 m/s
The corner velocity for negative maximum value of load factor is given as
nmax(negative) = (0.5ρ(V*)2
S(-CLmax)/Wg)
-1.4 = (0.5x1.225x(V*)2
x50x(-1)/30000x9.81)
V*(negative) = 167.7m/s
Dive velocity
VD = 1.4xVc (for a commercial aircraft)
Where, Vc is the cruise velocity,
Vc = 230m/s
Therefore
VD = 1.4x 230
VD (+ve,-ve) = 323.6m/s
In aerodynamics, the flight envelope, service envelope, or performance envelope of an
aircraft refers to the capabilities of a design in terms of airspeed and load factor or altitude.

54
The term is somewhat loosely applied, and can also refer to other measurements such as
maneuverability. When a plane is pushed, for instance by diving it at high speeds, it is said to
be flown "outside the envelope", something considered rather dangerous.
1. It indicates the graph between the velocity and the load factor
2. Upper part of the graph shows the positive limit load factor
3. Beyond the point A in the graph it shows the structural damage of the air craft
4. Point inside the graph indicates the CL<CL max
5. Point on the graph indicates CL=CLmax
6. Point outside the graph indicates CL>CLmax it leads to the stall region that is unstable
region
7. Lower part the graph shows the negative limit load factor
8. Point B in the graph shows the maximum negative lift capability
9. Beyond the point B or beyond the negative limit load factor it leads to the negative
ultimate load factor and may cause he structural damage
10. Beyond the negative ultimate load factor it leads to structural failure
Velocity a point A can be given by the formula V*=√
The rate of climb can also be written as t=∫

55
11.DRAG ESTIMATION:
DRAG:
In fluid dynamics drag is a force acting opposite to the relative motion of any object
moving with respect to A surrounding fluid this can adjust between two fluid layers or a fluid
and a solid surface unlike other resistive forces, such as dry friction , which are nearly
independent of velocity, drag forces, depend on velocity.
Types of drag are generally divided into the following catagories:
1. Parasitic Drag : consists of
a) Form Drag
b) Skin friction
c) Interference Drag
2. Wave Drag
Accurate drag estimation is critical in making computational design studies. Drag may
be estimated thousands of times during a multidisciplinary design optimization, and
computational fluid dynamics is not yet possible in these studies. The current model
has been developed as part of an air-vehicle conceptual-design multidisciplinary
design optimization framework. Its use for subsonic and transonic aircraft
configurations is presented and validated. We present our parametric geometry
definition, followed by the drag model description. The drag model includes induced
friction.
D = ρ S
= +
=
CALCULATION:
= =0.24
( )max =
( )max =150, k=0.301
=0.00003
=0.000031+0.240 =0.24

56
=0.24
TAKE OFF DRAG:
D = S
V=1.15
=1.225 kg , S=50 , =0.54
Take off drag=90.706 KN
LANDING DRAG:
D = S
V=1.3
=1.225 kg , S=50 , =0.24
Landing drag=115.911KN
CRUISE DRAG:
L = ρ v2
SCL
=68.87 KN
V=√
=178.194 m/s
D = ρ S
=68.586 KN
RESULT:
LANDING DRAG=115.911KN
CRUISE DRAG =68.586KN
TAKE OFF DRAG =90.706KN

57
12.TWO DIMENSIONAL VIEW DIAGRAM
FRONT VIEW:-
SIDE VIEW:-
All dimensions are in m

58
TOP VIEW:-
All dimensions are in m

59
13.FINAL VALUES:-
Table.7
PARAMETERS UNITS VALUE
EMPTY WT(THEO) Kg 15608
AIR FOIL SELECTION N/A NACA 4415
TAIL N/A TWIN TAIL
LANDING GEAR N/A
TRICYCLE LANDING
GEAR
ENGINE N/A
F16-FIGHTING FALCON
(GE F110)
DRAG KN 68.586
TAKEOFF DRAG KN 90.76
LANDING DRAG KN 115.911
CRUISE DRAG KN 68.873

60
14.CONCLUSION
The preliminary design of a modern efficient fighter 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.

61
15. 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
7. Boeing technical characteristics, viewed 2 March 2014
http://www.boeing.com/boeing/commercial/737family/specs.page
8. Engine selection and technical Information, viewed 25 March 2014
www.purepowerengine.com
http://en.wikipedia.org/wiki/Pratt_%26_Whitney_PW1000G
9. JavaFoil – Analysis of airfoil, viewed 29 March 2014
http://www.mhaerotools.de/aerofoils/javafoil.htm
10.https://en.wikipedia.org/wiki
11.https://booksite.elsevier.com
12.www.boeing.com
13.www.airbus.com
14.www.airliner.net
15.www.airfoiltools.com

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