Fighter aircraft design adp 1

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DESIGN OF FIGHTER AIRCRAFT
AIRCRAFT DESIGN PROJECT- I
Submitted by
DUDEKULA JAMAL (18101147)
In partial fulfilment for the award of the
degree of
BACHELOR OF TECHNOLOGY
IN
AERONAUTICAL ENGINEERING
SCHOOL OF AERONAUTICAL SCIENCES
HINDUSTAN INSTITUTE OF TECHNOLOGY AND SCIENCE
PADUR, CHENNAI – 603103
APRIL 2021

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SCHOOL OF AERONAUTICAL SCIENCES
BONAFIDE CERTIFICATE
Certified that this project report “DESIGN OF FIGHTER AIRCRAFT” is the
bonafide work of “DUDEKULA JAMAL (18101147)”, who carried out the
project work under my supervision. Certified further that to the best of my
knowledge the work reported here does not form part of any other
project/research work on the basis of which a degree or award was conferred on
an earlier occasion on this or any other candidate.
Dr. ASOKAN R Dr. CHANDRASEKAR M
Professor& Head of the department Assistant Professor
School of Aeronautical Sciences School of Aeronautical Sciences
Hindustan Institute of Technology Hindustan Institute of Technology
and Science and Science
Chennai – 603103 Chennai – 603103
Submitted for the project viva voice Examination heldon 17-05-2021
Internal Examiner External Examiner

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ACKNOWLEDGEMENT
It’s my extreme pleasure to thank our chairperson Dr. Elizabeth Verghese,
Hindustan Institute of Technology & Science, for providing me with a good,
pleasing and safe environment in our college which helped me a lot to carry on
with my project.
I wish to express my heartfelt gratitude to Dr. S.N. SRIDHARA, Vice-
Chancellor, Hindustan Institute of Technology & Science for providing me with
an excellent study environment.
I am thankful to Dr. Asokan R, Professor& Head of the Department, School of
Aeronautical Sciences for much of his valuable support, encouragement in
carrying out this work.
I would like to thank my internal guide Dr. CHANDRASEKAR M, for
continuously guiding and actively participating in my project, giving valuable
suggestions to complete the project work.
I would like to thank all the technical and teaching staff of Aeronautical
Department, who extended their support directly or indirectly.
Last, but not the least, I am deeply indebted to my parents who have been the
greatest support while I worked day and night for the project to make it a success.

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TABLE OF CONTENT
CHAPTER TITLE PAGE NO
ABSTRACT V
LIST OF TABLES vi
LIST OF FIGURES vii
LIST OF GRAPHS x
LIST OF SYMBOLS AND ABBREVIATIONS xi
1 INTRODUCTION TO DESIGN 1
2 COMPARATIVE STUDY OF DIFFERENT TYPES OF
AIRPLANES 6
3 COMPARATIVE STUDY ON SPECIFICATIONS AND
PERFORMANCE 10
4 PREPARATION OF COMPARATIVE DATA SHEETS 13
5 COMPARATIVE GRAPHS PREPARATION AND SELECTION
OF MAIN PARAMETERS FOR THE DESIGN 29
6 WEIGHT ESTIMATION 42
7 POWERPLANT SELECTION 53
8 WING, AEROFOIL & TAIL SELECTION 59
9 FUSELAGE AND LANDING GEAR SELECTION 87
10 LIFT AND DRAG CALCULATION 94
11 PERFORMANCE CALCULATION 102
12 THREE VIEWS OF FIGHTER AIRCRAFT 108
13 RESULT AND DISCUSSION 109
14 CONCLUSION AND FUTURE WORK 111
REFERENCE 113

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ABSTRACT
This project is about the design of a multirole supersonic fighter aircraft. A multirole
supersonic fighter aircraft is designed to perform different roles in combat. The air-
to-air combat role has been normally performed by fighter aircraft. In addition, a
multirole fighter has secondary roles such as air-to-surface attack. The term multirole
has been reserved for aircraft designed with the aim of using a common airframe for
multiple tasks where the same basic airframe is adapted to a number of differing
roles. The main motivation for developing multirole aircraft is a cost reduction in
using a common airframe.
Keywords: Fighter aircraft, supersonic, airfoil, turbofan engine.

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LIST OF TABLES
TABLE NO TITLE PAGE NO
6.1 Suggested Fuel Fraction for Several Mission Phases 44
6.2 Suggested value for L/D, Cj, Cp, ηp for several
mission phases
45
6.3 Regression line constant A & B 46
7.1 From Chapter 5, Table.no-5.3 53
7.2 Comparison of different engines 55
8.1 Wing design result 67
8.2
8.3
Comparison of different airfoil
Aerofoil selection for root, tip and mean chord
69
71
8.3 High lift device lift coefficient 77

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LIST OF FIGURES
FIGURE NO TITLE PAGE NO
1.1 Design Methodology 1
1.2 Design process 2
1.3 Conceptual design 4
1.4 Aircraft design configuration 5
4.1 Dassault Rafale 14
4.2 Eurofighter Typhoon 15
4.3 North America X-15 16
4.4 Sukhoi Su- 30MKI 17
4.5 Lockheed Martin F-22 Raptor 18
4.6 Mikoyan MIG-29 19
4.7 Saab JAS 39 Gripen 20
4.8 De Havilland Vampire 21
4.9 McDonnell Douglas F-4 Phantom II 22
4.10 Bell P-39 Airacobra 23
4.11 Dassault Mirage 2000 24
4.12 Grumman F-14 Tomcat 25
4.13 Chengdu J-7 26
4.14 Mitsubishi A6M Zero 27
4.15 Lockheed P-80 Shooting Star 28

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7.1
7.2
Pratt & Whitney F100 Engine
Cross section of engine Pratt & Whitney F100 – PW229
56
57
8.1
8.2
8.3
8.4
Wing types
Wing Position types
Low wing dihedral
Different types of Dihedral
59
60
61
62
8.5 wing planform 63
8.6 Aerofoil 68
8.7 Geometry of S2027 Airfoil 72
8.5 Geometry of GOE 490 Airfoil 72
8.6 Geometry of CLARK X Airfoil 72
8.7 Performance curves for the chosen aerofoil GOE
490
73
8.8 Performance curves for the chosen aerofoil S2027 74
8.9 Performance curves for the chosen aerofoil
CLARK X
75
8.10 Types of flaps 77
8.11 types of tail 81
9.1 MonoCoque Fuselage construction 87
9.2 Semi monocoque fuselage construction 88
9.3 Geodesic Truss Construction 88
9.4 Fixed Landing gear 89
9.5 Retractable Landing gear 89
10.1 lift representation 94
10.2 skin friction drags 97
10.3 form drag 98
10.4 wave drag 98
10.5 Typical streamlining effect 99

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11.1 Take-Off Performance 104
11.2 Landing Performance 105
12.1 Front view,Side view, and top view of Fighter
Aircraft
108
12.4 Isometric view of Fighter aircraft 108

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LIST OF GRAPHS
GRAPH NO TITLE PAGE NO
5.1 Max Speed Vs Aspect Ratio 34
5.2 Max Speed Vs Length 34
5.3 Max Speed Vs Height 35
5.4 Max Speed Vs Wing Area 35
5.5 Max Speed Vs Wing Span 36
5.6 Max Speed Vs Wing Loading 36
5.7 Max Speed Vs Empty Weight 37
5.8 Max Speed Vs Max Take Off Weight 37
5.9 Max Speed Vs Payload Weight 38
5.10 Max Speed Vs Thrust to Weight Ratio 38
5.11 Max Speed Vs Range 38
5.12 Max Speed Vs Rate of Climb 38
5.13 Max Speed Vs Service Ceiling 39
5.14 Max Speed Vs Dry Thrust 39
5.15 Max Speed Vs Afterburner Thrust 40

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LIST OF SYMBOLS & ABBREVIATIONS
A. R - Aspect Ratio
b - Wing span(m)
C - Chord of the Aerofoil (m)
Croot - Chord at Root (m)
Ctip - Chord at Tip (m)
Cd - Drag Co-efficient
Cdo - Zero lift Drag co-efficient
CP
- Specific fuel consumption (lbs /
hp / hr)
CL - Lift Co-efficient
D - Drag(N)
E - Endurance (hr)
e - Oswald efficiency factor
L - Lift (N)
(L/D) Loiter - Lift-to-drag ratio at loiter
(L/D) Cruise - Lift-to-drag ratio at cruise
M - Mach number of aircraft
Mff - Mission fuel fraction
R - Range (km)
Re - Reynolds number
s - Wing area (m2
)
Sref - Reference surface area
Swet - Wetted surface area
Sa - Approach distance (m)
Sf - Flare distance (m)
Sfr - Freeroll distance (m)
S.C - Service ceiling
A.C - Absolute ceiling

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T - Thrust (N)
Tcruise - Thrust at cruise (N)
Ttake-off - Thrust at take-off (N)
(T/W) Loiter - The thrust-to-weight ratio at Loiter
(T/W) Cruise - The thrust-to-weight ratio at cruise
(T/W) Take-off - The thrust-to-weight ratio at take-off
vCruise - velocity at cruise (m/s)
vStall - velocity at stall (m/s)
vt - Velocity at touch down (m/s)
WCrew - Crew weight (kg)
Wempty - Empty weight of the aircraft (kg)
Wfuel - Weight of fuel (kg)
Wpayload - Payload of the aircraft (kg)
W0 - Overall weight (kg)
W/S - Wing loading (kg/m2
)
ρ - Density of air (kg/m3
)
μ - Dynamic viscosity (Ns/m2
)
λ - Tapered ratio
R/C - Rate of Climb
η - Kinematic viscosity (m2
/s)

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CHAPTER 1
INTRODUCTION TO DESIGN
1.1 DESIGN METHODOLOGY
The aircraft design process is the engineering design process by which the
aircrafts are designed. The design process depends on many factors such as
customer and manufacturer demand, safety protocols, physical and economic
constraints etc… For some types of aircraft, the design process is regulated by
national airworthiness authorities. Among the fundamental elements of the
design process are the establishment of objectives and criteria, synthesis,
analysis, construction, testing and evaluation.
Aircraft design is a compromise between many competing factors and
constraints. It accounts for the existing designs and market requirements to
produce the best aircraft. The design method to be followed from the start of the
project to the nominal end falls in three main phases. These phases are illustrated
in Figure 1.1. In some industrial organizations, this phase is referred to as the
‘feasibility study’. At the end of the preliminary design phase, a document is
produced which contains a summary of the technical and geometric details
known about the baseline design. This forms the initial draft of a document that
will be subsequently revised to contain a thorough description of the aircraft. This
is known as the aircraft ‘Type Specification’.
Figure 1.1 Design Methodology

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1.2 DESIGN PROCESS
Figure 1.2 Design process

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1.3 PHASES OF AIRPLANE DESIGN
The complete design process has gone through three distinct phases that are
carried out in sequence. They are
• Conceptual design
• Preliminary design
• Detailed design
CONCEPTUAL DESIGN
This design process starts with a set of specifications (requirements) for a
new airplane or much less frequently as the response to the desire to implement
some pioneering, innovative new ideas and technology. The first steps towards
achieving that goal constitute the conceptual design phase. Here, the overall
shape, size, weight and performance of the new design are determined.
During the conceptual design phase, the designer is influenced by such
qualitative factors such as the increased structural loads imposed by a high
horizontal tail location through the fuselage, and the difficulties associated with
cut-outs in the wing structure if the landing gears are to be retracted into the
wing rather than the fuselage or engine nacelle.
PRELIMINARY DESIGN
In the preliminary design phase, only minor changes are made to the
configuration layout (indeed, if major changes were demanded during this phase,
the conceptual design process is actually flawed, to begin with. It is in the
preliminary design phase that serious structural, control system analysis and
design take place.
DETAIL DESIGN
The detail design phase is literally in great detail to the nuts-and-bolts phase of
airplane design. The aerodynamic, propulsion, structures performance and
flight control analysis have all been finished with the preliminary design
phase. The airplane is now simply a machine to be fabricated. The pressure
design of each individual rib, spar and Section of skin now take place. The size
of number and location of fasteners are determined. At the end of this phase,
the aircraft is ready to be fabricated.

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BLOCK ARRAY FOR CONCEPTUAL DESIGN
Figure 1.3 Conceptual design

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Figure 1.4 Aircraft design configuration

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CHAPTER 2
COMPARATIVE STUDY OF DIFFERENT TYPES OF
AIRPLANES
The following types of aircraft are taken for the study
 HOMEBUILT PROPELLER DRIVEN
 SINGLE ENGINE PROPELLER DRIVEN
 TWIN ENGINE PROPELLER DRIVEN
 AGRICULTURAL AIRPLANES
 BUSINESS JETS
 REGIONAL TURBO PROPELLER DRIVEN AIRPLANE
 TRANSPORT JETS
 MILITARY TRAINERS
 FIGHTERS
 MILITARY PATROL BOMB AND TRANSPORT AIRPLANES
 FLYING BOATS, AMPHIBIANS AND FLOAT AIRPLANES
 SUPERSONIC CRUISE AIRPLANES
Among these one aircraft is chosen for the study on its specification
and performance
2.1 HOMEBUILT AIRCRAFT
Homebuilt aircraft, also known as amateur-built aircraft or kit planes, are
constructed by anyone who may or may not be a professional in the aerospace
field. These aircraft may be constructed from “scratch,” from plans, or from
assembly kits.
Homebuilt aircraft are generally small, one to four- seat sports planes
which employ simple methods of construction. Fabric-covered wood or metal
frames and plywood are common in the aircraft structure. Fiberglass and other
composites as well as full aluminium construction techniques are also being used.

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2.2 SINGLE ENGINE PROPELLER DRIVEN AIRCRAFT
Single engine propeller aircraft are well-suited for short missions under
300miles. They can easily access smaller airports with shorter
runways. They are also known as light aircrafts. They are mainly used for
freight transport, sightseeing, photography and other similar roles as well as
personal use.
These aircrafts are nowadays used for training of pilots for the commercial
passenger aircrafts. Using these aircrafts, pilots can acquire license after
completing certain training requirements.
2.3 TWIN ENGINE PROPELLER DRIVEN AIRCRAFT
Causal observation of twin-engine propeller aircraft reveals that most
configurations consist of a forward wing with nacelle-mounted engines on each
side and a single tail empennage. However, about a third of the aircrafts do have
various engine and airframe arrangements. Aircrafts are arranged in nine
categories, as much as possible, with similar configuration traits. Each
configuration category is identified with a sample aircraft.
2.4 AGRICULTURAL AIRPLANES
An agricultural aircraft is an aircraft that was built for agricultural use
usually for the aerial application of pesticides (crop-dusting) or fertilizer in
these roles, they are referred to as “crop dusters” or “top dressers”. Agricultural
aircraft are also used for hydro-seeding. Agricultural aircraft are typically small,
simple, and rugged. Most have spraying systems attached to the trailing edges
of their wings, and pumps are usually driven by wind turbines.
2.5 BUSINESS JETS
A business jet, private jet or bizjet is a jet aircraft designed for transporting
small groups of people. Business jets may be adapted for other roles, such as
evacuation of causalities or express parcel deliveries, and some are used by public
bodies, government officials or the armed forces.

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2.6 REGIONAL TURBO PROPELLER DRIVEN AIRPLANE
A regional airliner or a feeder liner is a small airliner that is designed to fly up to
100 passengers on short-haul flights, usually feeding larger carriers’ airline hubs
from small markets. This class of airliners is typically flown by the regional
airlines that are either contracted by or subsidiaries of the larger airlines.
Regional airliners are used for short trips between smaller towns or from a larger
city to a smaller city. Feeder line, commuter, and local service are all alternative
terms for the same class of flight operations.
2.7 COMMERCIAL TRANSPORT AIRPLANE
A transport aircraft is used for transporting the passenger and air cargo.
Such aircrafts are most often operated by airlines. An airliner is typically
defined as an airplane intended for carrying multiple passengers or cargo in
commercial service. The largest of them are wide-body jets which are called
also twin-aisle. These are usually used for long-haul flights between airline hubs
and major cities. A smaller, more common class of airliners is the narrow-body
or single- aisle used for short to medium-distance flights with fewer passengers
than their wide-body counterparts.
2.8 MILITARY TRAINER
A trainer is a class of aircraft designed specifically to facilitate flight
training of pilot and aircrew. The use of a dedicated trainer aircraft with
additional safety features—such as tandem flight controls, forgiving flight
characteristics and a simplified cockpit arrangement—allows pilots-in-training to
safely advance their real-time piloting, navigation and warfighting skills without
the danger of overextending their abilities alone in a fully featuredaircraft.
2.9 FIGHTER 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.

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2.10 MILITARY PATROL BOMB AND TRANSPORT AIRPLANES
Military transport aircraft or military cargo aircraft are typically fixed wing
and rotary wing cargo aircraft which are used to airlift troops, weapons and other
military equipment by a variety of methods to any area of military operations
around the surface of the planet, usually outside the commercial flight routes in
uncontrolled airspace.
Originally derived from bombers, military transport aircraft were used for
delivering airborne forces during World War II and towing military gliders. Some
military transport aircraft are tasked to perform multi-role duties such as aerial
re-fuelling and, rescue missions, tactical, operational and strategic airlifts onto
unprepared runways, or those constructed by engineers.
2.11 FLYING BOATS, AMPHIBIANS AND FLOAT AIRPLANES
A flying boat is a fixed-winged seaplane with a hull, allowing it to land on
water, that usually has no type of landing gear to allow operation on land. It
differs from a floatplane as it uses a purpose-designed fuselage which can float,
granting the aircraft buoyancy. Flying boats may be stabilized by under wing
floats or by wing-like projections (called sponsons) from the fuselage. Their
advantage lay in using water instead of expensive land-based runways, making
them the basis for international airlines in the interwar period. They were also
commonly used for maritime patrol and air-sea rescue.
2.12 SUPER CRUISE AIRCRAFT
Supercruise is sustained supersonic flight of a supersonic aircraft with a
useful cargo, passenger, or weapons load performed efficiently, which typically
precludes the use of highly inefficient afterburners or “reheat”. Many well-known
supersonic military aircraft not capable of supercruise must maintain supersonic
flight in short bursts typically with afterburners. Aircraft such as the SR-71
Blackbird is designed to cruise at supersonic speed with afterburners enabled.

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CHAPTER 3
COMPARATIVE STUDY ON SPECIFICATIONS AND
PERFORMANCE
CREW
A group of people who works during the flight mission and operate an aircraft.
PASSENGERS
They are travellers on a public or private conveyance other than the pilot and crew.
EMPTY WEIGHT
The empty weight of an aircraft is the weight of the aircraft without including
passengers, baggage, or fuel.
PAYLOAD
The payload is what the airplane is intended to transport – passengers, baggage,
freight etc.
TAKE OFF WEIGHT
It is the maximum weight at which the pilot is allowed to attempt to take off due
to structural or other limits.
LANDING WEIGHT
It is the maximum aircraft gross weight due to design or operational limitations
at which an aircraft is permitted to land.
WING LOADING
It is the total weight of an aircraft divided by the area of its wing.

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WING AREA
It is the projected area of the wing planform and is bounded by the leading trailing
edges and the wing tips.
WING SPAN
The maximum distance between the two wing tips is wing span and is denoted by b.
THRUST TO WEIGHT RATIO
It is a dimensionless ratio of thrust to weight or a vehicle propelled by such an
engine that indicates the performance of the engine or vehicle.
WINGSWEEP BACK ANGLE
The angle at which a wing is either swept backward or occasionally forward from
its root.
ASPECT RATIO
It is the ratio of wing span to its mean chord. It is also equal to the square of the
wing span divided by the wing area.
Aspect ratio = b2
/s
THRUST
It is the force exerted by the engines on the airframe to overcome drag and is
measured in Newton (N).
POWER
It is the rate at which work is done.
WET THRUST
It is the augmented thrust with the usage of afterburners or liquid injection.
CRUISE SPEED
The speed at which combustion engines have an optimum efficiency level for fuel
consumption and power output.

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RATE OF ASCENT (CLIMB)
The rate of positive altitude changes with respect to time or distance.
RATE OF DESCENT (SINK)
The rate of negative altitude changes with respect to time or distance.
ABSOLUTE CEILING
It is the altitude where maximum rate of climb is zero is the highest altitude
achievable in steady, level flight.
SERVICE CEILING
It is the altitude where the maximum rate of climb is 100 ft/min and it’s
represented the practical upper limit for steady, level flight.
RANGE
It is the maximum distance an aircraft can fly between take-off and landing, as
limited by fuel capacity in powered aircraft.
ENDURANCE
It is the maximum length of time that an aircraft can spend in cruising flight as
long as the fuel is available.
STALLING VELOCITY
It is the velocity below which an aircraft will descend, or ‘stall’, regardless ofits
angle of attack.
TAKEOFF DISTANCE
It consists of two parts, the ground run and the distance from where the vehicle
leaves the ground until it reaches 50 ft or 15 m. The sum of these two distances
is considered the take-off distance.
LANDING DISTANCE
It is the distance required to bring the aircraft to a stop under ideal
conditions, assuming the aircraft crosses the runway threshold at a height of
50 ft, at the correct speed.

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CHAPTER 4
PREPARATION OF COMPARATIVE DATA SHEETS
4.1 INTRODUCTION
It’s the collection of data of various airplanes related to the concept taken.
Around 10 to 15 aircraft data with their design parameters are compared.
4.2 AIRCRAFT FOR REFERENCE
1. Dassault Rafale
2. Eurofighter Typhoon
3. North American X- 15
4. Sukhoi Su-30MKI
5. Lockheed Martin F-22 Raptor
6. Mikoyan MiG-29
7. Saab JAS 39 Gripen
8. De Havilland Vampire
9. McDonnell Douglas F-4 Phantom II
10.Bell p-39 Airacobra
11.Dassault Mirage 2000
12.Grumman F-14 Tomcat
13.Chengdu J-7
14.Mitsubishi A6M Zero
15.Lockheed P-80 Shooting Star

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DASSAULT RAFALE – SPECIFICATION
PARAMETER Values
Crew 1 or 2
Length (m) 15.27
Height (m) 5.34
Wing Area (m²) 45.7
Wing Span (m) 10.90
Aspect Ratio 2.59
Max Take Off Weight (Kg) 24,500
Empty weight (Kg) 10,000
Payload Weight (Kg) 15,000
Thrust to Weight Ratio 0.98
Max Speed (Km/h) 2,223
Service Ceiling (m) 15,835
Range (Km) 3,700
Rate of Climb (m/s) 304.8
Wing loading (Kg/m²) 328
Dry Thrust (KN) 50.04
Afterburner Thrust (KN) 75
Engine Type 2 × Snecma M88-4e turbofans

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4.2.2 EUROFIGHTER TYPHOON – SPECIFICATION
PARAMETER Values
Crew 1or 2
Length (m) 15.96
Height (m) 5.28
Wing Span (m) 10.95
Aspect Ratio 2.205
Range (Km) 2,900
Rate of Climb (m/s) 315
Dry Thrust (KN) 60
Engine Type Eurojet EJ200, Turbofan

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4.2.3 NORTH AMERICAN X-15 – SPECIFICATION
PARAMETER Values
Crew 1
Length (m) 15.47
Height (m) 4.04
Wing Area (m²) 19
Wing Span (m) 6.81
Aspect Ratio 2.44
Max Speed (Km/h) 7270
Range (Km) 450
Dry Thrust (KN) 253
Engine Type 1 × Reaction Motors XLR99-RM-2 liquid-
fuelled rocket engine

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4.2.4 SUKHOI SU-30MKI – SPECIFICATION
PARAMETER Values
Crew 2
Length (m) 21.935
Height (m) 6.36
Wing Span (m) 14.7
Aspect Ratio 3.48
Range (Km) 3000 km
Afterburner Thrust (KN) 122.58
Engine Type 2 × Lyulka AL-31FP turbofans

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4.2.5 LOCKHEED MARTIN F-22 RAPTOR – SPECIFICATION
PARAMETER Values
Crew 1
Length (m) 18.09
Height (m) 5.09
Wing Span (m) 13.56
Aspect Ratio 2.36
Range (Km) 3000
Dry Thrust (KN) 116
Engine Type 2× Pratt & Whitney F119-PW-100
augumemted turbofan engines

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4.2.6 MIKOYAN MIG-29 – SPECIFICATION
PARAMETER Values
Crew 1
Length (m) 17.32
Height (m) 4.73
Wing Area (m²) 38
Wing Span (m) 11.36
Aspect Ratio 3.39
Range (Km) 1,430
Dry Thrust (KN) 50
Engine Type 2 × Klimov RD-33 afterburning turbofan
engine

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4.2.7 SAAB JAS 39 GRIOEN – SPECIFICATION
PARAMETER Values
Crew 1
Length (m) 14.1 m -39C / 14.8 m -39D
Height (m) 4.5
Wing Area (m²) 30
Wing Span (m) 8.4
Aspect Ratio 2.35
Range (Km) 3,200
Dry Thrust (KN) 54
Engine Type 1 × Volvo RM12 afterburning turbofan
engine

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4.2.8 MCDONNELL DOUGLAS F-4 PHANTOM II -SPECIFICATION
PARAMETER Values
Crew 2
Length (m) 19.2
Height (m) 5
Wing Span (m) 11.7
Aspect Ratio 2.78
Range (Km) 2,699
Engine Type 2 × General Electric J79-GE-17A after-
burning turbojet engines

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4.2.9 DE HAVILLAND VAMPIRE – SPECIFICATION
PARAMETER Values
Crew 1
Length (m) 9.37
Height (m) 2.69
Wing Span (m) 12
Aspect Ratio 5.92
Thrust to Weight Ratio 3
Range (Km) 1,960
Wing loading (Kg/m²) 283.74
Engine Type 1 × de Havilland Goblin 3 centrifugal-flow
turbojet engine

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4.2.10 BELL P-39 AIRACOBRA– SPECIFICATION
PARAMETER Values
Crew 1
Length (m) 9.19
Height (m) 3.78
Wing Span (m) 10.36
Aspect Ratio 5.42
Thrust to Weight Ratio 4
Range (Km) 854
Engine Type Allison V-1710-85 V-12 liquid-cooled
piston engine

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4.2.11 DASSAULT MIRAGE 2000 -SPECIFICATION
PARAMETER Values
Crew 1
Length (m) 14.36
Height (m) 5.2
Wing Area (m²) 41
Wing Span (m) 9.13
Aspect Ratio 2.03
Range (Km) 1,550
Engine Type 1 × SNECMA M53-P2 afterburning
turbofan engine

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4.2.12 GRUMMAN F-14 TOMCAT SPECIFICATION
PARAMETER Values
Crew 2
Length (m) 19.13
Height (m) 1.545
Wing Span (m) 19.545
Aspect Ratio 7.27
Range (Km) 3,000
Dry Thrust (KN) 61
Engine Type 2 × General Electric F110-GE-400
afterburning turbofans

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4.2.13 CHENGDU J-7 – SPECIFICATION
PARAMETER Values
Crew 1
Length (m) 14.884
Height (m) 4.11
Wing Span (m) 8.32
Aspect Ratio 2.78
Range (Km) 850
Engine Type Powerplant: 1 × Liyang Wopen-13F
afterburning turbojet

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4.2.14 MITSUBISHI A6M ZERO – SPECIFICATION
PARAMETER Values
Crew 1
Length (m) 9.06
Height (m) 3.05
Wing Span (m) 12
Aspect Ratio 6.4
Range (Km) 1,870
Engine Type 1 × Nakajima NK1C Sakae-12 14-cylinder
air-cooled radial piston engine

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4.2.15 LOCKHEED P-80 SHOOTING STAR – SPECIFICATIONS
PARAMETER Values
Crew 1
Length (m) 10.49
Height (m) 3.43
Wing Span (m) 11.81
Aspect Ratio 6.37
Range (Km) 1,328
Rate of Climb (m/s) 6,870
Dry Thrust (KN) 20
Engine Type 1 × Allison J33-A-35 centrifugal
compressor turbojet

29
CHAPTER 5
COMPARATIVE GRAPHS PREPARATION AND SELECTION
OF MAIN PARAMETERS FOR THE DESIGN
5.1 CONSOLIDATION OF DATA
Consolidation of data is the comparison of collected data. We have made the
comparison of three aircrafts each with 15 selected aircraft in the following table
PARAMETER DASSAULT
RAFALE
EUROFIGHTER
TYPHOON
NORTHAMERICAN
X- 15
Crew 1 Or 2 1 or2 1
Length (m) 15.27 15.96 15.47
Height (m) 5.34 5.28 4.04
Wing Area (m²) 45.7 51.2 19
Wing Span (m) 10.9 10.95 6.81
Aspect Ratio 2.59 2.205 2.44
Max Take Off
Weight (Kg)
24,500 23,500 15,420
Empty weight (Kg) 10,000 11,000 6,622
Payload Weight
(Kg)
15,000 16,000 15,422
Thrust to Weight
Ratio
0.98 1.15 2.07
Max Speed (Km/h) 2,223 2,125 7,270
Service Ceiling (m) 15,835 19,812 108,000
Range (km) 3,700 2,900 450
Rate of Climb (m/s) 304.8 315 300
Wing loading
(Kg/m²)
328 312 829
Dry Thrust (KN) 50.04 60 253
Afterburner Thrust
(KN)
75 90 313
Engine Type 2Xsnecma M88-4e
Turbofans
Eurojet EJ200,
Turbofan
1x Reaction Motors
XLR99-RM-2 liquid
fueled rocket engine

30
PARAMETER SUKHOI D SU-
30MKI
LOCKHEED
MARTIN F-22
RAPTOR
MIKOYAN MIG-29
Crew 2 1 1
Length (m) 21.935 18.09 17.32
Height (m) 6.36 5.09 4.73
Wing Area (m²) 62.0 78.4 38
Wing Span (m) 14.7 13.56 11.36
Aspect Ratio 3.48 2.36 3.39
Max Take Off
Weight (Kg)
38,800 38,000 18,000
Empty weight (Kg) 18,400 19,700 11,000
Payload Weight
(Kg)
24,900 29,410 14,900
Thrust to Weight
Ratio
1.1 1.08 1.09
Max Speed (Km/h) 2,500 2,414 2,400
Range (km) 3000 3000 1,430
Rate of Climb (m/s) 300 350 330
Wing loading
(Kg/m²)
401 377 403
Dry Thrust (KN) 74.5 116 50
Afterburner Thrust
(KN)
122.58 156 81.59
Engine Type 2 x Lyulka AL-31FP
turbofans
2 x Pratt & Whitney
F119-PW-100
agumemted turbofan
engines
2 x Klimov RD-33
afterburning
turbofan engine

31
PARAMETER SAAB JAS 39
GRIOEN
MCDONNELL
DOUGLAS F-4
PHANTOM II
DE HAVILLAND
VAMPIRE
Crew 1 2 1
Length (m) 14.1 or 14.8 19.2 9.37
Height (m) 4.5 5 2.69
Wing Area (m²) 30 49.2 24.3
Wing Span (m) 8.4 11.7 12
Max Take Off
Weight (Kg)
14,000 28,030 5,620
Empty weight (Kg) 6,800 13,757 3,304
Payload Weight
(Kg)
6,500 18,824 6,895
Thrust to Weight
Ratio
0.97 0.86 3
Max Speed (Km/h) 2,460 2,370 882
Range (km) 3,200 2,699 1,960
Rate of Climb (m/s) 245 210 24
Wing loading
(Kg/m²)
283 380 283.74
Dry Thrust (KN) 54 52.96 10.2
Afterburner Thrust
(KN)
80.05 79.38 14.9
Engine Type 1 x VolvoRM12
afterburning
turbofan engine
2x General Electric
J79-GE-17A
engines
1x de Havilland
Goblin 3 centrifugal-
flow Turbojet engine

32
PARAMETER BELL P-39
AIRACOBRA
DASSAULT
MIRAGE 2000
GRUMMAN F-14
TOMCAT
Crew 1 1 2
Length (m) 9.19 14.36 19.13
Height (m) 3.78 5.2 4.9
Wing Area (m²) 19.8 41 52.5
Wing Span (m) 10.36 9.13 19.545
Max Take Off
Weight (Kg)
3,180 17,000 33,725
Empty weight (Kg) 2,956 7,500 19,833
Payload Weight
(Kg)
3,434 13,800 27,669
Thrust to Weight
Ratio
4 0.7 0.89
Max Speed (Km/h) 626 2,336 2,485
Range (km) 854 1,550 3,000
Rate of Climb (m/s) 19.33 285 230
Wing loading
(Kg/m²)
169 337 470
Dry Thrust (KN) 53.37 64.3 61
Afterburner Thrust
(KN)
75.61 95.1 104
Engine Type Allison V-1710-85 V-
12 liquid-cooled
piston engine
1 × SNECMA M53-
P2 afterburning
turbofan engine
2 × General Electric
F110-GE-400
afterburning
turbofans

33
PARAMETER CHENGDU J-7 MITSUBISHI A6M
ZERO
LOCKHEED P-80
SHOTING STAR
Crew 1 1 1
Length (m) 14.884 9.06 10.49
Height (m) 4.11 3.05 3.43
Wing Area (m²) 24.88 22.44 22.07
Wing Span (m) 8.32 12 11.81
Max Take Off
Weight (Kg)
9,100 2,796 7,646
Empty weight (Kg) 5,292 1,680 3,819
Payload Weight
(Kg)
7,540 2,796 5,534
Thrust to Weight
Ratio
0.507 0.28 0.364
Max Speed (Km/h) 2,200 533 956
Range (km) 850 1,870 1,328
Rate of Climb (m/s) 195 15.7 34.9
Wing loading
(Kg/m²)
441.85 107.4 250
Dry Thrust (KN) 44.1 41.97 20
Afterburner Thrust
(KN)
64.7 60.8 24
Engine Type Powerplant: 1 ×
Liyang Wopen-13F
1 × Nakajima NK1C
Sakae-12 14-cylinder
air-cooled radial
piston engine
1 × Allison J33-A-35
centrifugal
compressor turbojet

34
5.2 COMPARATIVE GRAPHS PREPARATION
MAX SPEED vs ASPECT RATIO
Graph 5.1 Max Speed Vs Aspect Ratio
Aspect ratio: 2.8
MAX SPEED vs LENGTH
Graph 5.2 Max Speed Vs Length
Length: 18m

35
MAX SPEED vs HEIGHT
Graph 5.3 Max Speed Vs Height
Height: 5.1m
MAX SPEED vs WING AREA
Graph 5.4 Max Speed Vs Wing Area
Wing Area: 48 m^2

36
MAX SPEED vs WING SPAN
Graph 5.5 Max Speed Vs Wing Span
Wing span: 12.5m.
MAX SPEED vs WING LOADING
Graph 5.6 Max Speed Vs Wing Loading
Wing loading: 440 Kg/m^2.

37
5.2.7 Max Speed Vs Empty Weight
Graph 5.7 Max Speed Vs Empty Weight
Empty Weight: 9,000 Kg.
5.2.8 Max Speed Vs Max Take-off Weight
Graph 5.8 Max Speed Vs Max Take-off Weight
Max Take-off Weight: 16,000 Kg.

38
5.2.9 Max Speed Vs Payload Weight
Graph 5.9 Max Speed Vs Payload Weight
Payload Weight: 7,500 Kg.
5.2.10 Max Speed Vs Thrust to Weight Ratio
Graph 5.10 Max Speed Thrust to Weight Ratio
Thrust to Weight Ratio: 1.4.

38
5.2.11 Max Speed Vs Range
Graph 5.11 Max Speed Vs Range
Range: 2,250 Km.
5.2.12 Max Speed Vs Rate of Climb
Graph 5.12 Max Speed Vs Rate of Climb
Rate of Climb: 275 m/s.

39
5.2.13 Max Speed Vs Service Ceiling
Graph 5.13 Max Speed Vs Service Ceiling
Service ceiling: 17,000 m.
5.2.14 Max Speed Vs Dry Thrust
Graph 5.14 Max Speed Vs Dry Thrust
Dry Thrust: 110 KN.

40
5.2.15 Max Speed Vs Afterburner Thrust
Graph 5.15 Max Speed Vs Afterburner Thrust
Afterburner Thrust: 130 KN.

41
5.3 DESIGN PARAMETERS FROM GRAPH
FLIGHT PARAMETER SI UNIT VALUE IMPEREAL
UNIT
VALUE
Length m 18 ft 59.05
Height m 5.1 ft 16.73
Wing Area m² 48 ft² 516.66
Wing Span (m) m 12.5 ft 41.01
Aspect Ratio 2.8 2.8
Max Take Off
Weight
Kg 16,000 lb 35,273.9
6
Empty weight Kg 9,000 lb 19,841
Payload Weight Kg 7,500 lb 16,534.67
Thrust to Weight Ratio 1.4 1.4
Max Speed Km/h 2,200 Mile/hr 1,367.01
Service Ceiling m 17,000 Miles 10.56
Range Km 2,250 Miles 1,398.08
Rate of Climb m/s 275 Miles/hr 615.157
Wing loading Kg/m² 440 lb/ft² 90.11
Dry Thrust KN 110 lbf 24,728.93
Afterburner Thrust KN 130 lbf 29,225.16

42
Rcr = 1378.90 nm
CHAPTER 6
WEIGHT ESTIMATION
6.1 INTRODUCTION
To find the weight of the following parameters of an aircraft.
 Takeoff Weight (WTO)
 Fuel Weight (WF)
 Empty Weight (WE)
The following are the data which is obtained from the graph to proceed for the
Weight estimation.
 Max Speed = 1,367.01 miles/hr
 Takeoff weight = 35,273.96 lbs
 Service ceiling (S.C) = 10.56 miles
 Range = 1398.08 miles
 Takeoff Distance (T.D) = 0.4319miles
 Landing Distance (L.D) = 0.4048 miles
 Payload = 16,534.67 lbs
RCR = R – [T.D + L.D + 2 x (S.C)]
Rcr = [1398.08 – (0.4319 + 0.4048 + (2 X 9.17))]
Where,
R – Total range = 1398.08 nm
T.D – Take off distance = 0.4319 nm
L.D – Landing distance = 0.4048 nm
Service ceiling = 9.17 nm

43
7.6MISSION PROFILE
Description:
0 – 1 – Engine Start & Warm up 6 – 7 - Descent
1 – 2 - Taxing 7 – 8 - Drop bombs
2 – 3 - Take off 8 – 9 - Strafe
3 – 4 - Climb 9 – 10 - Climb
4 – 5 - Cruise out 10 – 11 - Cruise in
5 – 6 - Loitering 11 – 12 - Descent
12 – 13 - Landing, Taxi, Shutdown

44
7.6MISSION FUEL FRACTION
The following tables 6.1, 6.2, 6.3 will be used for getting the values for the specified aircraft types.
Table 6.1 Suggested Fuel Fraction for Several Mission Phases

45
Table 6.2 Suggested value for L/D, Cj, Cp, ηp for several mission phases

46
Table 6.3 Regression line constant A & B

47
6.4 CALCULATION
Phase 1: Engine start and Warm-up
Begin weight is W0. End weight is W1. The ratio = 0.990
Phase 2: Taxi
Phase 3: Take-Off
Phase 4: Climb
Phase 5: Cruise – out
Begin weight is W4. End weight is W5. The amount of fuel used during cruise can be
found from Brequet’s range equation mentioned below.
Rcr = [ ] cr [ ] cr ln [ ]
Rcr = [R – (T + L + (2 X service ceiling))] = 1378.90 nm
Rcr = [ ] cr [ ] cr ln [ ]
1378.90 = [
.
] [ 7 ] ln [ ]
Where,
V – Speed (from graph) = 1367 mph
𝐶𝑗 = 0.6
= 7
𝑾𝟓
𝑾𝟒
= 1.09

48
Phase 6: Loitering
Begin weight is W5. End weight is W6. The ratio W6/W5 can be estimate from the
Brequet’s endurance equation which is mentioned below.
Elt = [ ] lt [ ] lt ln [ ]
0.5 = [
.
] [ 9 ] ln [ ]
Where,
Elt = 30 mins of loitering = 0.5 hrs
The mission profile assumes no range credit during loiter. Loiter time is 30 minutes.
Cj = 0.6
= 9
𝑾𝟔
𝑾𝟓
= 0.947
Phase 7: Descent
Begin Weight is W6. End Weight is W7. No credit is taken for range. However, a
penalty for fuel used during descents from high altitudes needs to be assessed.
Typically, the ratio
= 0.99
Phase 8: Drop Bombs
Begin Weight is W8. End Weight is W9. Typically, the ratio
= 1
Phase 9: Strafe
= 0.986
Phase 10: Climb
= 0.969

50
Phase 11: Cruise – in
= 0.959
Phase 12: Descent
Begin Weight is W13. End Weight is W14. No credit is taken for range. However, a
penalty for fuel used during descents from high altitudes needs to be assessed. Typically,
the ratio
= 0.99
Phase 13: Landing, Taxi and shutdown
= 0.995
Mission Fuel – Fraction (Mf f)
The Overall mission fuel-fraction, Mff can now be computed as
Mf f =
Mf f = [(0.990) (0.990) (0.990) (0.971) (1.09) (0.947) (0.99) (1) (0.986) (0.969) (0.959)
(0.99) (0.995)]
Mf f = 0.947
Maximum Take-Off Weight (WTO)
WTO = 35,273.96 lbs
Payload Weight (Wpayload)
Wpayload = Weight of Number of Passengers + Military loads
= (242 * 0) + 16,534.67
Wpayload = 16,534.67 lbs

50
Crew Weight (Wcrew)
Wcrew =(220+ 22) *Number of crew
= (220+22) * 1
Wcrew = 242 lbs
Weight of Fuel (Wf)
Wf = Wf used + Wres
Wf used:
Wf used = (1 – 𝑀𝑓𝑓) * WTO
Where,
𝑀𝑓𝑓 = 0.947
WTO = 35,273.96 lbs
Wf used = (1 – 0.947) * 35,273.96
Wf used = 1,869.51 lbs
Weight of Fuel Reserve (Wres):
Wres = 10-15 % of fuel used
= 10 % of fuel used
Wres = 186.95 lbs
Wf = Wfused + Wres
Wf = 1,869.51 + 186.95
Wf= 2056.46 lbs
Weight of Trapped Fuel Oil (WTFO)
WTFO = 0.5 % of WTO
WTFO = 176.36 lbs

51
Weight of Operative Empty (WOE Tent)
WOE Tent = WTO – Wf – Wpayload
= 35,273.96 – 2056.46 – 16,534.67
WOE Tent = 16,682.83 lbs
Weight of Empty (WE Tent)
WE Tent= WOE Tent – WTFO – Wcrew
= 16,682.83 – 176.36 – 242
WE Tent = 16,264.47 lbs
Weight of Actual (WE Actual)
WE Actual= inv log10 [
10 , . –
], Where A = 0.5091; B = 0.9505
= inv log10 [
10 , . – .
.
]
WE Actual= 17,728.13 lbs
Difference between WE Actual and WE Tent
WE Actual – WE Tent= 17,728.13 – 16,264.47
WE Actual – WE Tent= 1463.66 lbs
Percentage of Error
% Error = [
E Actual – E Tent
E Actual
]* 100
= [
.
.
]* 100
% Error = 8.25 %

52
RESULT
PARAMETERS SI UNIT (Kg) IMPERIAL UNIT (lbs)
Take-off Weight (WTO ) 16,000 35,273.96
Fuel Weight (WF ) 932.79 2056.46
Empty Weight (WE ) 8041.34 17728.13
Payload Weight (Wpayload ) 7,500 16,534.67

53
CHAPTER 7
POWERPLANT SELECTION
7.1 INTRODUCTION
An airplane, an object which is Airborne. It is the multidisciplinary area where
Aerodynamics, Structures, Propulsion, control & stability place a major role in the
formation of an aircraft. Unlike automobile engines, these engines are Air-breathing
engines which use atmospheric air as the medium for airborne. There is a different kind
of engines equipped with an aircraft,
7.2 TYPES OF ENGINES
1. Piston Engine
2. Turbofan
3. Turboprop
4. Turbojet
5. Ramjet
6. Scramjet
7.3 THRUST REQUIRED CALCULATION
𝐓𝐑
𝐖𝟎
= 𝒂(𝑴𝒎𝒂𝒙)^𝑪
Table 7.1 From Chapter 5, Table.no-5.3
𝐓𝐑
𝐖𝟎
= 𝒂(𝑴𝒎𝒂𝒙)^𝑪
𝒂 C
Jet trainer 0.488 0.728
Jet fighter (dogfighter) 0.648 0.594
Jet fighter (other) 0.514 0.141
Military cargo/ bomber 0.244 0.341
Jet transport 0.267 0.363

54
From above table for Jet Fighter,
a =0.514; c =0.141
From Result of Weight Estimation, W0 = 156.91 𝐾𝑁
From Graph, umax = 611.11 m/s
T@17000m = 216.7 K
W. K. T,
Mmax=
@
Mmax=
.
√ . × × .
Mmax=
.
.
Mmax= 2.07

𝑻𝑹
𝑾𝑶
= 0.514 × 2.07 .
TR= 89.36 KN
𝑻
𝑾
CALCULATION
TR = W𝑇0 ( )
⇒ =
⇒ =
. ×
𝑇
𝑊
= 0.569
The thrust produced should be 10% more than the required thrust.
Hence, thrust required is TR= 98.29 KN
Therefore, Thrust required for single engine is 98.29 KN.

55
7.4 SELECTION OF ENGINE
Choice of the engine is a Turbofan for obvious reasons such as higher operating fuel
economy & efficiency for high payloads.
A list of engines with weight and thrust matching our requirements are chosen and are
tabulated below
Table 7.2 Comparison of Different Engines
S.NO Name of the engine Engine
Type
Dry Weight
(kg)
SFC
(kg/kN.h)
Total Thrust
(kN)
1. GTRE GTX-35VS
Kaveri
Turbofan 1,236 207 81
2. Pratt & Whitney
F100
Turbofan 1,737 197 129.7
3. Volvo RM8 Turbofan 2,350 257 125
4. Tumansky R- 15 Turbojet 2,454 275 100
5. Eurojet EJ200 Turbofan 988 176 90
6. Snecma M88 Turbofan 897 169.5 75
7.5 DETAILS ABOUT THE ENGINE
The Pratt & Whitney F100 is a two-spool afterburning turbofan engine. The F100 has
been selected by the U.S. Air Force (USAF), Navy, Air Force Reserve, Air National
Guard (ANG) and 22 foreign nations for the Boeing F-15 Eagle/F-15E Strike Eagle and
the Lockheed Martin F-16 Fighting Falcon multi-role fighters.
F100 engines power 99% of all USAF F-15 aircraft and 62% of the world’s inventory of
F-16 fighters. The F100 has a record of dependability, performance and safety. To date,
more than 7,000 F100 engines have been produced. As of July 2018, the F100 engine
fleet has accumulated more than 28 million flight hours and 3,800 engines remain in
service with 23 customers.
The latest model in the F100 Series, the F100-PW-229 (introduced in 1992), is an
improved high-thrust improvement of the older F100-PW-220 (introduced in 1986). The
F100-PW-229 incorporates proven technological innovations and generates more than
29,000 pounds of thrust with afterburner.

56
The modular maintenance concept, coupled with a state-of-the art FADEC (Full
Authority Digital Engine Control) system with improved, real-time engine monitoring
and fault isolation capability, promotes the highest level of operational readiness.
Figure 7.1 Pratt & Whitney F100 Engine
The newest engine in the Pratt & Whitney 229-Series, the F100-PW-229 Engine
Enhancement Package (EEP) – launched in 2004 – has raised the engine depot
inspection interval from 4,300 to 6,000 Total Accumulated Cycles (TAC), effectively
extending the typical depot interval from 7 to 10 years and, at the same time, providing a
30% engine life-cycle cost reduction.
Furthermore, the F100-PW-229 engine is the only fighter engine funded and qualified by
the U.S. Air Force to the 6,000-cycle capability. The F100-PW-229 EEP includes
advanced hot section technology developed for the F119-PW-100 turbofan used on the
F-22 Raptor and the F135 engine used on the F-35 Lightning II.
The EEP configuration was incorporated into all production F100-PW-229 engines in
2009 and has been specifically designed to be easily installed in all existing pre-2009
F100-PW-229 engines.
Another variant of the F100, the F100-PW-220U, powered Northrop Grumman’s X-47B
flight test aircraft for the U.S. Navy’s Unmanned Combat Air System Carrier
Demonstration (UCAS-D) program. The F100-PW-220U provides up to 16,000 pounds
of thrust and is designed for operation in a maritime environment such as on aircraft
carriers.

57
Figure 7.2 Cross section of the engine Pratt & Whitney F100-PW229
7.6 TECHNICAL SPECIFICATIONS
 Type: Afterburning Turbofan
 Length: 191 inches (4.90 m)
 Diameter: 34.8 inches (0.88 m) inlet, 46.5 inches (1.18 m) maximum external
 Dry weight: 3,849 pounds(1,737 kg)
 Compressor: Dual spool Axial-flow Compressor with 3 fan and 10 compressor
stages
 Combustors: Annular Combustion chamber

58
 Turbine: 2 low-pressure and 2 high-pressure stages
 Manufacturer: Pratt & Whitney (United Technologies)
 Maximum thrust:
 17,800 pounds-force (79 kN) dry thrust
 29,160 pounds-force (129.7 kN) with afterburner
 Overall pressure ratio: 32:1
 Bypass ratio: 0.36:1
 Turbine inlet temperature: 2,460 °F (1,350 °C)
 Specific fuel consumption: Military thrust: 0.76 lb/(lbf·h) (77.5 kg/(kN·h)) Full
afterburner: 1.94 lb/(lbf·h) (197.8 kg/(kN·h))
 Thrust-to-weight ratio: 7.8:1
 First Flight: 1989
CONCLUSION
The preferable choice of engine, from the above, would be Pratt & Whitney F100
engine since the engine thrust is 129 KN. It is a Single afterburning turbofan engine
equipped and also it meets our thrust required calculation 98.29 KN which also suits
our demand of weight and power.

59
CHAPTER 8
WING, AEROFOIL & TAIL SELECTION
8.1 INTRODUCTION
This chapter explains the selection of wing, types of wing and calculation of wing
design parameter.
8.1.1 WING SELECTION
After the final weight estimation of the aircraft, the primary components of the
aircraft to be designed is the wing. The wing weight and it’s lifting capabilities are
in general, a function of the thickness of the aerofoil section that is used in the
wing structure. The first step towards designing the wing is the thickness
estimation. The thickness of the wing in turn depends on the critical Mach number
of the aerofoil or rather, the drag divergence Mach number corresponding to the
wing section.
8.1.2 TYPES OF WING
Wings are differentiated from there wing configuration by the following
 Swept back wing
 Delta wing
 Tapered wing
 Based on the aspect ratio
Figure 8.1 Wing types

60
8.1.3 THE POSITION OF WING (Low Wing- Dihedral)
The location of the wing in the fuselage (along with the vertical axis) is very
important. Each configuration (Low, High and mid) has its own advantages, but in
this design, the Low wing with dihedral offers significant advantages.
Figure 8.2 Types of Wing Position
High wing:
 Stable in roll and therefore anhedral is generally required to reduce stability.
 Affords good ground clearance for external stores/engine pods/propellers.
 Short landing gear.
 High wings can block upwards pilot visibility (bad for turning and climbing)
Low wing:
 Unstable in roll and therefore dihedral is generally required to increase stability.
 Very bad ground clearance for external stores/engine pods/propellers.
 Long landing gear if external stores/engine pods/propellers are necessary.
Mid wing:
 Neutral roll stability, the dihedral can be zero.
 OK ground clearance for external stores/engine pods/propellers.
 Short landing gear.

61
Note:
o Sweep can have an effect similar to dihedral. According to Raymer, 10
degrees of sweep have the same effect as 1 degree of dihedral.
o Supersonic aircraft have sweep angles up to 60 degrees, i.e. effective dih
edral of up to 6 degrees.
Figure 8.3 Low wing dihedral
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 Fighter Aircrafts.

62
Dihedral angle and dihedral effect:
 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 nonzero 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 8.4 Different types of Dihedral
WING AERODYNAMIC CENTER
 Position of trapezoidal and swept wing aerodynamic center:
Subsonic conditions: ¼ of the mean aerodynamic chord
Supersonic conditions: 0.4 of the mean aerodynamic chord (Raymer).
 For Delta wings,
the aerodynamic center lies at 2/3 of the root chord from the apex of the
wing. Cropping the wing tends to move the aerodynamic center forward.

63
8.1.4 WING GEOMETRY DESIGN
 The geometry of the wing is a function of four parameters, namely the Wing
loading (W/S), Aspect Ratio (b2 /S), Taper ratio (λ) and the Sweepback angle
at quarter chord (Λqc).
 The Take-off Weight that was estimated in the previous analysis is used to
find the Wing Area S (from W/S). The value of S also enables us to calculate
the Wingspan b (using the Aspect ratio). The root chord can now be found
using the equation.
Croot =
𝟐∗𝑺
𝒃∗(𝟏 𝛌)
The tip chord is given by,
Ctip= λ*Croot
 WING PLANFORM
Figure 8.5 wing planform
the shape of the wing as viewed from directly above - deals with airflow in three
dimensions and is very important to understanding wing performance and aeroplane
flight characteristics. Aspect ratio, taper ratio, and sweepback are factors in planform
design that are very important to the overall aerodynamic characteristic of a wing.

64
8.2 WING DESIGN CALCULATION
8.2.1 WING AREA
Area, S =
𝐖𝐓𝐎
𝑾𝒊𝒏𝒈 𝑳𝒐𝒂𝒅𝒊𝒏𝒈
=
𝟏𝟔𝟎𝟎𝟎
𝟒𝟒𝟎
S = 36.36 𝐦𝟐
Where,
𝑊𝑇𝑂 = 16,000 kg (From Chapter 6 Weight estimation)
Wing Loading = 440 kg/m
8.2.2 ASPECT RATIO
Aspect Ratio = = 2.8 (From Graph 5.1)
Wing Span, b = 𝐴. 𝑅 × 𝑊𝑖𝑛𝑔 𝐴𝑟𝑒𝑎 = √2.8 × 36.36
b = 10.08 m
Where,
Wing Area = 36.36 m
8.2.3 ROOT CHORD (CRoot)
CR = =
.
.
Where,
Wing Span = 10.08 m
Aspect Ratio = 2.8
Croot = 3.6 m

65
8.2.4 TAPER RATIO (𝝀)
𝜆 =
( )
( R)
Where,
Taper Ratio 𝜆 = 0.5
Root Chord CR = 3.6 m
Tip Chord Ct = 3.6× 0.5 = 1.8 m
Ct = 1.8 m
Note:
For rectangle wing, λ = 1
For elliptical wing, λ = 0.9
For tapered wing, λ = 0.5
8.2.5 MEAN AERODYNAMIC CHORD – MAC (𝑪)
MAC (𝐶) = × 𝐶𝑅 ×
Where,
CR = 3.6 m
λ = 0.5
MAC (𝑪) = 2.8 m
8.2.6 VOLUME OF FUEL WEIGHT
Volume of fuel weight = =
.
Volume of fuel weight = 1.164 m
Where,
Weight of fuel = 932.79 Kg
Density of fuel = 801 Kg/m

66
8.2.7 THICKNESS OF ROOT CHORD (CR) AND TIP CHORD (Ct)
20 % of Volume of fuel weight = × 𝐶 × 0.375 × 𝑏
Where,
20 % of Volume of fuel weight = 0.2328 m
𝐶 = 7.84
b = 10.08 m
𝒕
𝒄
= 0.00785 m
Thickness of Root chord (TR) = 0.00785 × CR = 0.00785 × 3.6
TR = 0.02826 m
Thickness of Tip chord (Tt) = 0.00785 × Ct = 0.00785 × 1.188
Tt = 0.0105 m
Figure 8.5 Wing Design

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Table 8.1 Wing design Result
S.NO DESIGN CHARACTERISTICS VALUES
1 Wing loading (Kg/m ) 440
2 Wing Area S (m ) 36.36
3 Aspect Ratio 2.8
4 Span b (m) 10.08
5 Taper ratio (λ) 0.5
6 Root Chord (m) 3.6
7 Tip chord (m) 1.8
8 Mean chord (𝑪) 2.8

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8.3 AIRFOIL SELECTION
8.3.1 AIRFOIL NOMENCLATURE
The aerofoil is the main aspect and is the heart of the aeroplane. The aerofoil affects the
cruise speed, landing distance and take off distance, stall speed and handling qualities
and aerodynamic efficiency during all phases of flight.
Aerofoil Selection is based on the factors of Geometry & definitions, design/selection,
families/types, design lift coefficient, thickness/chord ratio, lift curve slope,
characteristic curves.
Figure 8.6 Aerofoil
The following are the aerofoil geometry and definition:
Chord line:
It is the straight line connecting leading edge (LE) and trailing edge (TE).
Chord (c):
It is the length of chord line.
Thickness (t):
measured perpendicular to chord line as a % of it (subsonic typically 12%).
Camber (d):
It is the curvature of the section, perpendicular distance of section mid-points from chord
line as a % of it (sub sonically typically 3%).
The angle of attack (α):
It is the angular difference between the chord line and airflow direction.

69
The following are aerofoil categories:
1. Early it was based on trial & error.
2. NACA 4 digit is introduced during 1930’s.
3. NACA 5-digit is aimed at pushing position of max camber forwards for increased
CLmax.
4. NACA 6-digit is designed for lower drag by increasing region of laminar flow.
5. Modern it is mainly based on the need for improved aerodynamic characteristics
at speeds just below the speed of sound.
Table 8.2 Comparison of different airfoil
S.NO Name Thickness
(%)
Camber
(%)
Lift
Coefficie
nt
(CL)
Lift to
Drag
(L/D)
Stall
Angle
(deg)
TE
Angle
(deg)
LE
Radius
(%)
1 GOE 546
AIRFOIL
10.4 3.5 1.326 60.2 3 12.9 2.6
2 GOE 490
AIRFOIL
8.8 3.8 1.358 72.9 8 15.7 2.5
3 NACA
64(1)-212
MOD A
12 2.4 1.302 49.9 -0.5 6.7 3
4 CLARK X
AIRFOIL
11.7 3.3 1.308 57 8 18.2 3
5 S2027 14.5 2.7 1.303 46.4 12 7.2 2.6

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NACA 4 Digit
 1st digit: maximum camber (as % of chord).
 2nd digit (x10): location of maximum camber (as % of chord from leading edge
(LE)).
 3rd & 4th digits: maximum section thickness (as % of chord).
NACA 5 Digit
 1st digit (x0.15): design lift coefficient.
 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 the series type.
 2nd digit (x10): location of minimum pressure (as % of chord from leading edge
(LE)).
 3rd digit: indicates an 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).
From the above list of aerofoil, the one chosen is the GOE 490 which have the suitable
lift coefficient for the current design.
Min Lift Coefficient can be calculated using below formula
W = L = 𝑣 SCL Wing
CL max = [
× × ×
]

71
where,
V = stall velocity = 51.4 m/sec
S = wing area 36.36 m
 = density of air = kg/m
CL = 0.27
CL Available = 90% Of CL = 0.9 × 0.27
CL Available = 0.243 at 0 AOA
Hence,
Section used at the mean aerodynamic chord – GOE 490 AIRFOIL
The section used at the tip – CLARK X
The section used at the root – S2027
Table 8.3 Aerofoil selection for root, tip and mean chord
CHORD AIRFOIL CLmax
ROOT S2027 1.303
MEAN GOE 490 AIRFOIL 1.358
TIP CLARK X 1.303
CLmax = (1.303+1.358+1.303) / 3= 1.321
CLmax Available = 0.9 ∗ CLmax = 1.189

72
8.3.2 AEROFOIL GEOMETRY SELECTION
8.3.2.1 S2027
Figure 8.4 Geometry of S2027 Airfoil
8.3.2.2 GOE 490 AIRFOIL
Figure 8.5 Geometry of GOE 490 Airfoil
8.3.2.3 CLARK X
Figure 8.6 Geometry of CLARK X Airfoil

73
8.3.3 Performance curves for the chosen aerofoil GOE490
Figure 8.7 Performance curves for the chosen aerofoil GOE 490

74
8.3.4 Performance curves for the chosen aerofoil S2027
Figure 8.8 Performance curves for the chosen aerofoil S 2027

75
8.3.5 Performance curves for the chosen aerofoil CLARK X
Figure 8.9 Performance curves for the chosen aerofoil CLARK X

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8.4 HIGH LIFTING DEVICES
In aircraft design and aerospace engineering, a high-lift device is a component or
mechanism on an aircraft's wing that increases the amount of lift produced by the
wing. The device may be a fixed component or a movable mechanism which is
deployed when required. Common movable high-lift devices include wing flaps
and slats. Fixed devices include leading-edge root extensions and boundary layer
control systems, which are less commonly used.
8.4.1 TYPES OF DEVICES
 Flaps
 Slots & Slats
 Boundary layer control and blown flaps
 Leading edge root extension
For the current design, the Double slotted flap is selected. ∆ of the slotted flap for
different configurations is given in the table below:
Table 8.4 High lift device lift coefficient
S.NO HIGH LIFT DEVICE CL
1 Plain flap 0.7-0.9
2 Split flap 0.7-0.9
3 Flower flap 1.0-1.3
4 Slotted flap 1.3Cf/C
5 Double Slotted flap 1.6Cf/C
6 Triple Slotted flap 1.9Cf/C
7 Leading edge flap 0.2-0.3
8 Leading edge slat 0.3-0.4
9 Kruger flap 0.3-0.4
Note: Cf/C – Flap chord
CALCULATIONS
CLmax Required (takeoff) = 1.189 + 1.6 = 2.789
CLmax Required (landing) = 1.189 + 1.5 = 2.689

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8.5 TAIL PLANE SELECTION
A tail plane, also known as a horizontal stabilizer, is a small lifting surface located on the
tail (empennage) behind the main lifting surfaces of a fixed-wing aircraft. The tail plane
of an aircraft is classified as follows,
 Conventional Tail
 T – Tail
 V – Tail
 Inverted V-tail
 Cruciform tail
 Twin- tail
 Triple-Tail
8.5.1 Conventional tail:
The conventional tail design is the most common form. It has one vertical stabilizer
placed at the tapered tail section of the fuselage and one horizontal stabilizer divided into
two parts, one on each side of the vertical stabilizer. For many airplanes, the
conventional arrangement provides adequate stability and control.
8.5.2 T - tail:
The horizontal stabilizer is mounted on top of the fin, creating a "T" shape when viewed
from the front. T-tails keep the stabilizers out of the engine wake, and give better pitch
control. T-tails have a good glide ratio, and are more efficient on low-speed aircraft.

80
8.5.3 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. A V-tail can be lighter than a
conventional tail in some situations and produce less drag. A V-tail may also have a
smaller radar signature.
8.5.4 Inverted V tail:
The inverted V-tail is similar to V-tail but it is inverted and it provides more stability and
manoeuvrability. It is mostly used in Unmanned Aerial Vehicles.
8.5.5 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.

80
8.5.6 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/A18 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.

82
8.5.7 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.
Volume Coefficient’s:
Horizontal tail volume coefficient for fighter aircraft is =𝑉 = 0.4
Vertical tail volume coefficient for fighter aircraft is = 𝑉 = 0.07

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SELECTION: (TWIN TAIL)
 Reason for selection of Twin Tail
There are three reasons for twin vertical stabilizers:
1. Need to have lots of vertical stabilizer power without the plane getting too
tall. (F-14, F-15, Su-27, Mig-29, F-22, J-20, Su-57). Twin-engined
aircraft needs lots of stabilizing power in case of one engine shuts down.
There are however some twin-engined planes with single vertical
stabilizer, but in most of those planes the engines are very close to each
other and the stabilizer is quite big (Rafale, EF Typhoon, Panavia
Tornado, F–111)
2. High-alpha performance. (F-35, F-18, F-22, Su-27, Mig-29, J-20, Su-
57?). One vertical stabilizer would be directly behind the body of the
plane, and would not get undisturbed air when flying at high alpha angle.
Two vertical stabilizers that are canted outwards allow them to get clean
air that is not obstructed by the body of the plane when flying at high
alpha angle. This makes the plane more controllable at high alpha angles.
3. Stealth. (F-35, F-22, J-20, Su-57,). Single vertical stabilizer would easily
make a 90-degree angle with the wings or horizontal stabilizers, and 90-
degree angles make reflections go back where they came from, so they
need to be avoided on a stealth aeroplane. With two stabilizers that are
canted outwards, the angles can easily be something else than 90 degrees.
Also, single vertical stabilizer would reflect radio waves directly on the
side.
The Twin – Tail empennage is selected for the aircraft due to the following
advantages,
1. In the case of multi engine propeller driven aircraft, twin vertical stabilizers give the
advantage of prop wash directly over the rudder surfaces, increasing low speed
controllability, and control in the event of an engine failure.
2. Weight can often be saved when incorporating twin tails, as they split the
aerodynamic load, and in being shorter, require much less structural support than a
single massive tail would.
3. Twin tails save vertical space for large aircraft, which would otherwise require a
larger hanger, or smaller aircraft stowed below deck in the case of an aircraft carrier.

84
4. Aircraft like the F-15 with a wide semi-lifting body fuselage, benefit from control in
high AoA situations, where the fuselage would otherwise block sufficient airflow to a
single centrally mounted vertical stabilizer.
5. Two smaller tails have a smaller radar cross section, which is definitely a
consideration for military aircraft. Twin tails can also be angled, as seen on jets like the
F-18, F-22 and F-35, to further reduce radar cross section, and scatter radar waves. In the
case of the A-10, the twin tails hide the hot engine exhaust from certain angles, making it
harder for heat seeking missiles, and other thermal based tracking systems to maintain a
lock on them.
6. The F-18, F-22 and F-35 also take advantage of twin vertical stabilizers by their
ability to use the rudder surfaces as air brakes.
AEROFOIL SELECTION:
NACA0009 Symmetrical Aerofoil is selected due to Good stall characteristics, small
centre of pressure movement across large speed range, and High pitching moment.
Geometry of NACA0009 Aerofoil

85
Performance curves of NACA0009 Aerofoil:

86
CONCLUSION:
The aerofoil which I have selected for root, mean, tip chords are all with 11% thickness
to chord ratio.
The Double slotted flaps taken for high lifting device, will provide manoeuvrability for
the fighter aircraft.
In my Design Project aircraft are equipped with Twin Tail Design. As mentioned above
the Twin tail design will give better stability performance to the aircraft.

87
CHAPTER 9
FUSELAGE AND LANDING GEAR SELECTION
9.1 FUSELAGE SELECTION
The fuselage construction plays a major role in reducing the total weight of the aircraft.
The fuselage construction of an aircraft is classified as follows,
 Monocoque
 Semi-Monocoque
 Geodesic
9.1.1 MONOCOQUE
The Monocoque (single shell) fuselage relies largely on the strength of the skin or
covering to carry the primary loads. Monocoque construction uses stressed skin to
support almost all loads much like an aluminium beverage can. Because most twisting
and bending stresses are carried by the external skin rather than by an open framework,
the need for internal bracing was eliminated or reduced, saving weight and maximizing
space.
Figure 9.1 MONOCOQUE
9.1.2 SEMI-MONOCOQUE
To overcome the strength/weight problem of Monocoque construction, a modification
called semi Monocoque construction was developed. It also consists of frame
assemblies, bulkheads, and formers as used in the Monocoque design but, additionally,
the skin is reinforced by longitudinal members called longerons.
Figure 9.2 SEMI-MONOCOQUE

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9.1.3 GEODESIC TRUSS
Geodesic airframe is a type of construction for the airframes of aircraft developed by
British aeronautical engineer Barnes Wallis in the 1930s. It makes use of a space frame
formed from a spirally crossing basket-weave of load-bearing members. The principle is
that two geodesic arcs can be drawn to intersect on the fuselage in a manner that the
torsional load on each cancel out that on the other
Figure 9.3 GEODESIC TRUSS
SELECTION
In a jet fighter the fuselage consists of a cockpit large enough only for the controls and
pilot, but in a jet airliner it includes a much larger cockpit as well as a cabin that has
separate decks for passengers and cargo.
The predominant types of fuselage structures are the monocoque (i.e., kind of
construction in which the outer skin bears a major part or all of the stresses) and semi-
monocoque. These structures provide better strength-to-weight ratios for the fuselage
covering than the truss-type construction used in earlier planes.

89
9.2 LANDING GEAR SELECTION
Landing gear is the undercarriage of an aircraft or spacecraft and may be used for either
takeoff or landing. For aircraft it is generally both. For aircraft, the landing gear supports
the craft when it is not flying, allowing it to take off, land, and taxi without damage.
9.2.1 LANDING GEAR TYPES
The landing gears are classified as follows,
1. Fixed
2. Retractable
Fixed Landing Gear
Landing gear employing a rear-mounted wheel is called fixed landing gear. Fixed gear is
designed to simplify design and operation. The advantages are that it is always deployed
and its initial instalments cost is low. Whereas its disadvantage is that produces constant
drag.
Figure 9.4 FIXED LANDING GEAR
Retractable Landing Gear
A retractable gear is designed to streamline the airplane by allowing the landing gear to
be stowed inside the structure during cruising flight. Retractable landing gear systems
may be operated either hydraulically or electrically, or may employ a combination of the
two systems.
Figure 9.5 RETRACTABLE LANDING GEAR

90
SELECTION:
The Retractable landing gear is implemented in the aircraft due to the following reasons,
1. There will be less drag during cruise as the landing gear will be retracted
2. It helps in higher cruise speeds and increased climb performance
3. The primary benefits of being able to retract the landing gear are increased climb
performance and higher cruise airspeeds due to the resulting decrease in drag.
4. Retractable landing gear systems may be operated either hydraulically,
electrically, or may employ a combination of the two systems.
5. Warning indicators are provided in the cockpit to show the pilot when the wheels
are down and locked and when they are up and locked or if they are in
intermediate positions.
6. Emergency operation systems provide additional security.
9.2.2 LANDING GEAR CONFIGURRATIONS
The landing gears have different configurations based on the number of wheels and their
arrangement. They are classified as follows,
a. Single wheel
b. Bicycle
c. Tricycle
d. Quadricycle
e. Multi-bogey
Single wheel Landing Gear
The single-wheel configuration, defined as a main gear of having a total of two wheels,
one on each strut, the dual-wheel configuration, defined as a main gear of having a total
of four wheels, two on each strut, and the dual-tandem configuration, defined as two sets
of wheels on each strut.
Figure 9.6 SINGLE WHEEL LANDING GEAR

91
Bicycle Landing Gear
A relatively uncommon landing gear option is the bicycle undercarriage. Bicycle gear
features two main gear along the center line of the aircraft, one forward and one aft of
the center of gravity. Preventing the plane from tilting over sideways are two small
outrigger gear mounted along the wing.
Figure 9.7 BICYCLE LANDING GEAR
Tricycle
The most commonly used landing gear arrangement is the tricycle-type landing gear. It
is comprised of main gear and nose gear. Tricycle-type landing gear is used on large and
small aircraft. It allows more forceful application of the brakes without nosing over
when braking, which enables higher landing speeds.
Figure 9.8 TRICYCLE LANDING GEAR

92
Quadricycle
Quadricycle gear are also very similar to the bicycle arrangement except there are four
main gear roughly equal in size and mounted along the fuselage. Like bicycle gear, the
Quadricycle undercarriage also requires a very flat attitude during take-off and landing.
This arrangement is also very sensitive to roll, crosswinds, and proper alignment with
the runway.
Figure 9.9 QUADRICYCLE LANDING GEAR
Multi-bogey
A final variation that is worth mentioning is the use of multiple wheels per landing gear
strut. This additional tire is particularly useful on carrier-based aircraft where two nose
wheels are a requirement. Multiple wheels are also often used on main gear units for
added safety, especially on commercial airliners.
Figure 9.10 MULTI-BOGEY LANDING GEAR

93
SELECTION:
The Tricycle landing gear configuration is implemented for the fighter aircraft.
In tricycle landing gear
 Landing gear employing a front-mounted wheel is called tricycle landing gear.
 Tricycle landing gear aircraft have two main wheels attached to the airframe
behind it's CG that support most of the weight of the structure.
 Additionally, a nose wheel will typically provide some sort of nose wheel steering
control.
Tricycle Gear Benefits:
1. It allows the more forceful application of the brakes during landings at high
speeds without causing the aircraft to nose over.
2. It permits better forward visibility for the pilot during take-off, landing, and
taxiing.
3. It tends to prevent ground looping (swerving) by providing more directional
stability during ground operation since the aircraft's CG is forward of the main
wheels [The forward CG keeps the airplane moving forward in a straight line
rather than ground looping.
CONCLUSION
Semi-monocoque and Tri-cycle type landing gear has been selected for my fighter
aircraft

94
CHAPTER 10
LIFT AND DRAG CALCULATION
11.1 LIFT ESTIMATION
Component of aerodynamic force generated on aircraft perpendicular to the flight
direction.
Figure 10.1 lift representation
LIFT COEFFICIENT (CL)
Amount of lift generated depends on:
– Planform area (S), air density (𝜌), flight speed (V), lift coefficient(CL)
L =
𝟏
𝟐
𝜌 𝑉 𝑆
CL is a measure of lifting effectiveness and mainly depends upon:
– Section shape, planform geometry, the angle of attack (𝛼), compressibility effects
(Mach number), viscous effects (Reynolds’ number).
GENERATION OF LIFT
Aerodynamic force arises from two natural sources:
– Variable pressure distribution.
– Shear stress distribution.

95
L=159.362KN
 Shear stress primarily contributes to overall drag force on aircraft.
 Lift mainly due to pressure distribution, especially on main lifting surfaces, i.e., wing.
 Require (relatively) low pressure on upper surface and higher pressure on the lower
surface.
 Any shape can be made to produce lift if either cambered or inclined to flow direction.
 Classical aerofoil section is optimum for high subsonic lift/drag ratio.
10.1.1 LIFT AT TAKE-OFF
L =
𝟏
𝟐
𝜌 𝑉 𝑆𝐶𝐿𝑚𝑎x
ρ = Density at sea level = 1.225 Kg/m3
Vstall = stalling speed = 57.21 m/s
V = 0.7*1.3*Vstall = 52.06 m/s
V =
√
2wTo/
𝜌𝑠𝐶𝐿
52.06 = √
2 𝑋 146169
1.225 𝑋 43.82 𝑋 𝐶𝑙
s = wing area = 48 m2
CLmax = coefficient of lift = 2
𝐿 =0.5X 1.225 X (52.06)2 X 48 X 2
L = 159362.32 N
10.1.2 LIFT AT CRUISE
L =
𝟏
𝟐
ρ = Density at 16500 m = 0.1539 Kg/m3
V = Vcruise = cruising speed = 597.2 m/s
V = √2𝑊𝑇𝑂
𝜌𝑠𝐶𝐿
597.2 = √ 2 𝑋 156906.4
0.1539 𝑋 48 𝑋 𝐶𝑙

96
L=159.365 KN
L = 159.515 KN
CLmax
Cruising Lift Coefficient = 0.121
𝐿 = 0.5X 01539 X (597.2)2 X 48 X 0.121
L = 159365.04 N
10.1.3 LIFT AT LANDING
L =
𝟏
𝟐
V = 0.7*1.2*Vstall = 48.05 m/s
V = √
2𝑊𝑇𝑂
𝜌𝑠𝐶𝐿
48.05 = √ 2 𝑋 156906.4
1..225 𝑋 48 𝑋 𝐶𝑙
CLmax = Maximum coefficient of lift = 2.35
L= 0.5 X 1.225 X (48.05)2
X 48 X 2.35
L = 159515.16 N
10.1 DRAG ESTIMATION
DRAG:
• Drag is the resolved component of the complete aerodynamic force which is
parallel to the flight direction (or relative oncomingairflow).
• It always acts to oppose the direction of motion.
• It is the undesirable component of the aerodynamic force while the lift is the
desired component.

97
DRAG COEFFICIENT (CD)
• Amount of drag generated depends on:
1.Planform area (S), air density (𝜌), flight speed (V), drag
coefficient (CD)
2.CD is a measure of aerodynamic efficiency and mainly depends
upon:
i. Section shape, planform geometry, the angle of attack,
compressibility effects
(Mach number), viscous effects (Reynolds’ number).
DRAG COMPONENTS
• SKIN FRICTION
1.Due to shear stresses produced in the boundary layer.
2.Significantly more for turbulent than laminar types of
boundary layers.
Figure 10.2 skin friction drags
• FORM (PRESSURE) DRAG
1. Due to static pressure distribution around the body -
component resolved in direction of motion.
2. Sometimes considered separately as forebody and rear
(base) drag components.

98
Figure 10.3 form drag
WAVE DRAG
1. Due to the presence of shock waves at transonic and
supersonic speeds.
2. The result of both direct shock losses and the influence of
shock waves on the boundary layer.
Figure 10.4 wave drag

99
Figure 10.5 Typical streamlining effect
∅ = (16h/b)^2/1+ (16h/b)^2
h = height = 5.1 m
b = Wing span = 12.5 m
10.2.1 DRAG AT TAKE-OFF
𝐷 = 1/2 𝜌𝑉^2 𝑆[𝐶𝐷𝑂 + (∅𝐶𝐿𝑚𝑎𝑥^2/𝜋𝑒(𝐴𝑅)) ]
V = 0.7*1.3*Vstall = 52.06 m/s
CLmax = coefficient of lift = 2
CDO = 0.003
∅= 0.97
∅ = 0.97

100
Oswald efficiency factor e = 0.971
AR = Aspect Ratio = 2.8
D= 0.5 𝑋 1.225 𝑋(52.06)2
𝑋43.82 [0.003 +(0.97 X 2^2/𝜋 𝑋 0.971(2.8)) ]
10.2.2 DRAG AT CRUISE
𝐷 = 0.5 𝜌𝑉𝑐𝑟𝑢𝑖𝑠𝑒^2 𝑆[𝐶𝐷𝑂 + ∅𝐶𝐿𝑚𝑎𝑥/ 2 𝜋𝑒(𝐴𝑅) ]
ρ = Density at max altitude = 0.1539 Kg/m3
Vcruise = cruising speed = 597.2 m/s
CL = cruising lift coefficient = 0.121
CDO = 0.003
∅ = 0.9
D = 0.5 𝑋 0.1539 𝑋(597.2)^2𝑋 48 [0.003 + (0.97 𝑋 0.1212^2/ 𝜋 𝑋 0.971(2.8) ]
D = 6143.36 N
10.2.3 DRAG AT LANDING
D = 𝜌𝑉2𝑆[𝐶𝐷𝑂 +∅𝐶𝐿𝑚𝑎𝑥^2/ 𝜋𝑒(𝐴𝑅)]
V = 0.7*1.2*Vstall = 48.05 m/s
D = 36.453 KN

101
CLmax
Maximum coefficient of lift = 2.35
CDO = 0.003
∅ = 0.97
D = 0.5 𝑋 1.225 𝑋(48.05)2𝑋 48 [0.003+(0.97 𝑋 2.352 /𝜋 𝑋 0.971(2.8) ]
RESULT
CONDITION LIFT (N) DRAG (N)
TAKE OFF 159362 36453
CRUISE 159365 6143.36
LANDING 159515 42796
D = 42.796 KN

102
CHAPTER 11
PERFORMANCE CALCULATION
INTRODUCTION
Our study of static performance (no acceleration) answered a number of questions about
the capabilities of a given airplane-how fast it can fly, how far it can go, etc. However,
there are more questions to be asked: How fast can it turn? How high can it "zoom"?
What ground distances are covered during takeoff and landing? The answers to these
questions ,involve accelerated flight, the subject of this chapter.
11.1 THRUST AND POWER
THRUST
By definition, the thrust available, denoted by TA, is the thrust the power plant of the
airplane. The various propulsion devices are described at length. The single purpose of
these propulsion devices is to reliably provide thrust in order to propel the aircraft.
Unlike the thrust TR, which has almost everything to do with the airframe of the airplane
and virtually nothing to do with the power plant, the thrust available TA has almost
everything to do with the power plant and virtually nothing to do with the airframe.
Thrust Available (TA)
TA = 98.29 KN
Thrust required (TR)
TR= TA × ( )^m = 98.29 X
.
.
^ 1.14 = 14.08 KN
POWER
By definition, the power available, denoted by PA, is the power provided by the
powerplant of the airplane. The maximum power available compared with the power
required allows the calculation of the maximum velocity of the airplane.

103
𝑃𝑅 =8408.57 KW
𝑃𝐴 =54494.5KW
R/C=293.71 m/s
 Power Available
𝑃𝐴 = 𝑇𝐴 ∗ 𝑉∞ = 98.29 ∗ 597.2
 Power Required
𝑃𝑅 = 𝑇𝑅 ∗ 𝑉∞ = 8.396 ∗ 597.2
11.1 RATE OF CLIMB AND RATE OF SINK
Rate of Climb
R/C = PA-PR / WTO
R/C = 54494.5 – 8408.57
156.91
Rate of Sink
R/S = (2W/P∞)^2 X ( 𝐶𝐷/𝐶𝐿 )^ 1.5
𝐶𝐷=𝐶𝐷𝑂 +𝐾(𝐶𝐿)2=0.003+0.117∗(2)2 = 0.471
R/S=57.84m/s

104
11.2 TAKE-OFF PERFORMANCE
Figure 11.1 Take-Off Performance
Distance from rest to clearance of obstacle in flight path and usually
considered in two parts:
• Ground roll - rest to lift-off (SLO)
• Airborne distance - lift-off to specified height (35 ft FAR, 50 ft others).
The aircraft will accelerate up to lift-off speed (Vlo = about 1.2 x VStall) when it
will then be rotated.
A first-order approximation for ground roll take-off distance may be made from:
S LO= 144𝑊2
𝑔 𝑋 𝜌 𝑋 𝑆 𝑋 𝐶𝐿𝑚𝑎𝑥 X T
This shows its sensitivity to W (W2
) and 𝜌 (1/ 𝜌2 since T also varies with
𝜌). Slo may reduce by increasing T, S or Cl, max (high lift devices relate to latter
two).
An improved approximation for ground roll take-off distance may be made by
including drag, rolling resistance and ground effect terms.
S LO = 144𝑊2
𝑔 𝑋 𝜌 𝑋 𝑆 𝑋 𝐶𝐿𝑚𝑎𝑥 𝑋 {𝑇 − [𝐷 + 𝜇𝑟(𝑊 − 𝐿)]}𝑎𝑣
The bracketed term will vary with speed but an approximation may be made by
using an instantaneous value for when V = 0.7 x Vlo in the above equation:

105
𝐷 = 0.5 𝜌𝑉^2𝑆[𝐶𝐷𝑂 + (∅𝐶𝐿𝑚𝑎𝑥 ^2 /𝜋𝑒(𝐴𝑅)) ]
∅ = (16h/b)^2/1+ (16h/b)^2
Where ∅ accounts for drag reduction when in ground effect:
Where h = height above ground, b = wingspan.
𝜇𝑟 = 0.02 for smooth paved surface, 0.1 for grass.
CALCULATION
𝑆𝐿𝑂 =
144 𝑋 (36453) 2
9..81 𝑋 1..225 𝑋48 𝑋 2 𝑋 {91250 − [98290 + 0.02(156910 − 159362)]}𝑎𝑣
SLO = 2680.15
11.3 LANDING PERFORMANCE
Figure 11.2 Landing Performance

106
APPROACH & LANDING
 Consists of three phases:
 Airborne approach at constant glide angle (around 30
) and constant speed.
 Flare - transitional manoeuver with airspeed reduced from about 1.3 VStall
down to touch-down speed.
 Ground roll - from touch-down to rest.
 Ground roll landing distance (s3 or s1) estimated from:
Where Vav may be taken as 0.7 x touch-down speed (Vt or V2) and Vt is
assumed as 1.3 x Vstall
S L = 1.69𝑊2
𝑔 𝑋 𝜌 𝑋 𝑆 𝑋 𝐶𝐿𝑚𝑎𝑥 X T𝑋 { [𝐷 + 𝜇𝑟(𝑊 − 𝐿)]}𝑎𝑣
 𝜇𝑟 is higher than for take-off since brakes are applied - use 𝜇𝑟 = 0.4 for the
paved surface.
 If thrust reversers (Tr) are applied, use:
S L = 1.69𝑊2
𝑔 𝑋 𝜌 𝑋 𝑆 𝑋 𝐶𝐿𝑚𝑎𝑥 X T𝑋 {𝑇 + [𝐷 + 𝜇𝑟(𝑊 − 𝐿)]}𝑎𝑣
CALCULATION
𝑆𝐿=
1.63 𝑋 (42796) 2
9..81 𝑋 1..225 𝑋48 𝑋 2.35 𝑋 {98290 + [42796+ 0.02(156910 − 159362)]}𝑎𝑣
SL = 1618.99 m

107
RESULT
PARAMETER VALUE
THRUST REQUIRED 14.08 KN
THRUST AVAILABLE 98.29 KN
POWER REQUIRED 8408.57 KW
POWER AVAILABLE 54494.5 KW
RATE OF CLIMB 293.71 m/sec
RATE OF SINK 57.84 m/sec
TAKE – OFF DISTANCE 2680.15 m
LANDING DISTANCE 1618.99 m

108
CHAPTER 12
THREE VIEWS OF FIGHTER AIRCRAFT
Figure 12.1 Front view, Top view and Side view of fighter aircraft
Figure 12.2 Isometric View of fighter aircraft

110
CHAPTER 13
RESULT AND CONCLUSION
WEIGHT
PARAMETERS SI UNIT (Kg) IMPERIAL UNIT (lbs)
Take-off Weight (WTO ) 16,000 35,273.96
Fuel Weight (WF ) 932.79 2056.46
Empty Weight (WE ) 8041.34 17728.13
Payload Weight (Wpayload ) 7,500 16,534.67
ENGINE TYPE
 The preferable choice of engine is Pratt & Whitney F100 engine since the engine
thrust is 129 KN. It is a Single afterburning turbofan engine.
 Thrust required calculation 98.29 KN.
WING TYPE
Tapered wing with dihedral monoplane configuration mounted as a low wing.
AIRFOIL CHOSEN
1. Section used at the mean aerodynamic chord - GOE 490 AIRFOIL
2. The section used at the tip - CLARK X
3. The section used at the root - S2027.
FUSELAGE TYPE
A semi-monocoque fuselage has been constructed.
EMPENNAGE TYPE
Twin Tail Plane configuration is used because Twin tail design will give better stability
performance to the aircraft
LANDING GEAR
Retractable Tri-cyclic landing gears is constructed.

110
LIFT AND DRAG CALCULATION
CONDITION LIFT (N) DRAG (N)
TAKE-OFF 159362 36453
CRUISE 159365 6143.36
LANDING 159515 42796
PERFORMANCE CALCULATION
PARAMETER VALUE
THRUST REQUIRED 14.08 KN
THRUST AVAILABLE 98.29 KN
POWER REQUIRED 8408.57 KW
POWER AVAILABLE 54494.5 KW
RATE OF CLIMB 293.71 m/sec
RATE OF SINK 57.84 m/sec
TAKE – OFF DISTANCE 2680.15 m
LANDING DISTANCE 1618.99 m

111
CHAPTER 14
CONCLUSION AND FUTURE WORK
CONCLUSION
The preliminary design of a Multirole Fighter aircraft 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 aero plane’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 design is a fine blend of science, creativity, the presence of mind and the application
of each one of them at the appropriate time. Design of anything needs experience and an
optimistic progress towards the ideal system.
The scientific society always looks for the best product design. This involves the strong
fundamentals of science and mathematics and their skillful applications, which is a tough
job endowed upon the designer. We have enough hard work for this design project.
A design never gets completed in a fluttering sense but it is one step further towards the
ideal system. But during the design of this aircraft, we learnt a lot about aeronautics and
its implications when applied to an aircraft design.
The challenges we faced at various phases of the project made clear the fact that
experience plays a vital role in the 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.

112
FUTURE WORK
The above work will enhance the knowledge in continuation of the design given in
Aircraft Design project-I.
In Design Project – II will be studied for the design with Gust and manoeuvrability
envelopes. Performance of Critical loading and the final calculation of V-n graph. A
theoretical approach to Study of structural design will be undertaken. To estimate loads
of wings, to estimate loads of fuselage. Balancing and maneuvering loads on the
tailplane, Aileron and Rudder load are started. Designing the structural layout of the
aeroplane. Even some of the components like wings, the fuselage is designed. Finally,
detailed design report will be prepared with sketches or drawings.

113
REFERENCE
1. Ajay Kumar Kundu “Aircraft Design”, Cambridge University Press, ISBN -
9780511844652 (2010)
2. Daniel P. Raymer “Aircraft Design: A Conceptual Approach” American Institute
of Aeronautics and Astronautics, Inc. 2nd Edition ISBN 0-930403-51-7 (1992)
3. Egbert Molenbeek “Synthesis of subsonic Airplane design”, Delft University
Press, 1st Edition ISBN - 9789024727247, 90-247-2724-3 (1982)
4. Gerard Frawley & Jim Thorn “International Directory of Military Aircraft”
Australian Aviation, ISBN – 187567120X, 9781875671205 (1996)
5. Ira H. Abbott & Albert E. von Danhof, “Theory of Wing Section – Including
summary of Airfoil data” Dover publications, ISBN - 0486605868 9780486605869
(1959)
6. Dr Jan Roskam “Airplane Design” Roskam aviation & engineering, Vol. 8 ISBN -
127-126-125-1 (2009)
7. John D.Anderson Jr.“Aircraft Performance and Design” WCB/McGraw-Hill,
ISBN - 9780070019713, 0-07-001971-1 (1999)
8. John D. Anderson Jr. “Fundamentals of Aerodynamics” McGraw-Hill Education,
6th Edition ISBN - 1259129918, 9781259129919 (2016)
9. John D. Anderson Jr. “Introduction to Flight” McGraw-Hill Education, ISBN
0073380245, 9780073380247 (2011)
10. Lloyd R. Jenkinson & James F. Marchman III “Aircraft Design Projects: for
engineering students”, Butterworth-Heinemann, ISBN - 0-75065772-3 (2003)
11. Paul Jackson (Editor-in-Chief), “Jane’s All the World’s Aircraft” 119
12. Jane's Publishing, 95th Sub Edition, ISBN - 9780710626141, 0710626142 (2004)
13. William Green “The Observer’s World Aircraft Directory” Frederick Warne &
Co. Ltd, ISBN - 978-1125857120 (1961)

114
14.http://www.airliners.net/aircraft-data
15.https://www.faa.gov/airports/engineering/aircraft_char_database/
16.http://airfoiltools.com/
17.http://www.combataircraft.com/
18.https://www.faa.gov/regulations_policies/handbooks_manuals/air craft/a mt
airframe handbook/media/ama_Ch13.pdf
19.https://www.military.com/equipment/fighter-aircraft
20. http://www.airfoildb.com/
21.https://www.dassault-aviation.com/en/
22.https://www.militaryfactory.com/
23.https://www.lockheedmartin.com/
24.https://www.geaviation.com/commercial/engines
25.https://www.rolls-royce.com/products-and-services/civilaerospace.aspx
26.http://getdrawings.com/aircraft-drawing
27.https://en.wikipedia.org/wiki/Rolls-Royce/Snecma_Olympus_593

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