This document provides details on the design of a 1-seater military aircraft. It discusses the aircraft's specifications including its weight, performance characteristics, and dimensions of the wing. It also summarizes the structural analysis and material selection for the fuselage and wings. Several chapters describe the preliminary and detailed design of the aircraft's wing, fuselage, and tail section. Load distributions and structural components of each section are analyzed.

1. AIRCRAFT DESIGN
PROJECT-II
1-SEATER MILITARY AIRCRAFT
Prepared By:
Md Tarique Mahmood/Manish Kumar Dubey/Manoj Rasaily

2. INTRODUCTION
Aircraft Design Project-II is a continuation of Aircraft Design Project-I.
As mentioned in our earlier project, military aviation is of prime importance for
every country and with this in mind we have designed a supersonic 1-seater military fighter
aircraft.
The purpose of ADP-II is to enhance the knowledge in continuation of
the design project given in ADP-I. Also, Aircraft Design Project-II deals with a
more in-depth study and analysis of aircraft performance and structural
characteristics. In the following pages we have carried out structural analysis of
fuselage and wings and the appropriate materials have been chosen to give our
aircraft adequate structural integrity. The determination the landing gear
position, retraction and other accompanying systems and mechanisms have also
been done.

3. REQUIRED DATAS FROM ADP-I
PARAMETERS SHAPES
FUSELAGE Conventional
WING Delta Wing, Mid Wing
TAIL V- Tail
LANDING GEAR Tricycle
ENGINE Turbo Fan Engine
PARAMETERS VALUES
CREW 1 (ONE)
ROLE Air Superiority, Bomber
ENDURANCE 5 hrs.
PAYLOAD 15,000 kg
TYPE OF PAYLOAD Missiles, Bombs &
Rockets
COMBAT TIME 2 hrs.

4. SPECIFICATIONS OF DESIGNED AIRCRAFT
DESIGN POINT PARAMETERS VALUES
Weight
Overall, Weight 66711.88 lbs.
Empty Weight 35518.95 lbs.
Fuel Weight 22041.82 lbs.
Payload Weight 31069.33 lbs.
Performance
Cruise speed 1621.28 ft/s
Maximum Speed 2100 ft/s
Range 16400000 ft
Thrust Required 76549.127 lbs.
Wing
Area 252.72ft
Span 49.87 ft
Root Chord 25.60 ft
Tip Chord 7.67 ft
Aerofoil NACA 632-215
PARAMETERS VALUES
MAXIMUM SPEED 2.1 Mach
CRUISE SPEED 2300 ft/s
SERVICE CEILING 60000 ft
RATE OF CLIMB 1000 ft/s

5. CHAPTER-1
PRELIMINARY DESIGN OF AN AIRCRAFT WING
WING LOAD DISTRIBUTION LOAD ACTING ON WING
As both the wings are symmetric, let us
consider the starboard wing at first. There are three
primary loads acting on a wing structure in transverse
direction which can cause considerable shear forces
and bending moments on it.
•They are as follows:
•Lift force (given by Schrenk’s curve)
•Self-weight of the wing &Weight of the power
plant.
While performing a structural analysis of the aircraft, it is
necessary to investigate all the various loads acting on the aircraft
that will help us in determining the shear force and bending
moment distribution.
Wing load distribution is an important phase in the structural design
of the aircraft. This is because, wing is the component that enables the
aircraft to fly and any damages to the wing during flight due to over
stressing can drastically reduce the lift by sections of wing or the
entire wing being ripped off and the aircraft plummets into the ground
or sea.

6. SHRENK’S CURVE
Lift varies along the wing span due to the variation in
chord length, angle of attack and sweep along the span.
Schrenk’s curve defines this lift distribution over the wing span
of an aircraft, also called simply as Lift Distribution Curve.
Schrenk’s curve is an approximation for the lift distribution along
the span for the wing. The equation of the curve is obtained by
taking the average of the trapezoidal and elliptic lift distributions.
Schrenk’s Curve is given by
Where;
y1 is Linear Variation of lift along semi wing span also
named as L1 y2 is Elliptic Lift Distribution along the
wing span also named as L2

7. CONST RUCT ION OF SCHRENK’S CURVE :
Schrenk’s Curve is given by
`
Y=
WING SPAN Y2 Y2 Y
0 37952.4 6581.28 22266.8
0.28536 36244.9 6573.05 21409
0.57073 34537.4 6548.3 20542.8
0.85609 32829.8 6506.82 19668.3
1.14146 31122.3 6448.32 18785.3
1.42682 29414.8 6372.30 17893.5
1.71219 27707.3 6278.15 16992.7
1.99755 25999.7 6165.02 16082.4
2.28292 24292.2 6031.85 15162.0
2.56828 22584.7 5877.27 14231.0
2.85365 20877.2 5699.56 13288.3
3.13901 19169.6 5496.46 12333.0
3.42438 17462.1 5265.03 11363.6
3.70974 15754.6 5001.34 10377.9
3.99511 14047.1 4699.98 9373.54
4.28047 12339.5 4353.11 8346.35
4.56584 10632.0 3948.77 7290.41
4.85120 8924.53 3466.90 6195.72
5.13657 7217.01 2868.21 5042.86
5.42193 5509.48 2055.00 3782.24

8. LINEAR LIFT DISTRIBUTION(Y1):
LINEAR LIFT DISTRIBUTION:
Lift at root
Lift at tip
LINEAR LIFT DISTRIBUTION:
Density, D = 1.225 kg/m3
Root chord of the wing, Cr = 8.054 m
Tip chord of the wing, Ct = 2.013 m
Area of the wing planform
S=126.66 m2
Wing span, b=25.16 m
By representing this lift at sections of root and tip we can get the
equation for the wing.
y1 = Lroot – * x
y1 = 37952.46532 – * x
y1 = 35517.42964 – 5983.653 x

9. ELLIPTICAL LIFT DISTRIBUTION(Y2):
Twice the area under the curve or line will give the lift
which will be required to overcome weight.
Considering an elliptic lift distribution, we get
Where,
b1=
Actual
lift at
root a=
wing
semi
span
W = Gross Weight in kg

10. SELF-WEIGHT OF WING(Y3)
For this preliminary analysis, the structural weight of the
wing is assumed to vary parabolic-ally along the span, with zero
weight per unit span at the wingtips. Again, the area enclosed
between the weight distribution curve and the semi span axis
should be equal to wing structural weight. Self-weight of the
wing,
W(wing) = 0.0759 x 29500.64726 x 9.81
W(wing) = 21965.56244N
W(Port wing)= - 10982.78122 N (Acting
Downwards)
W(Starboard)= - 10982.78122 N (Acting
Downwards)
Assuming parabolic weight distribution

11. FUEL WEIGHT(Yf):
This design has fuel in the wing so we have to
consider the weight of the fuel in one of the wing.
Again, by using general formula for straight line
y= mx + c
we get,
dy = 1.426825 x
Dy= (5210.148434-dy)
Where, m =
Solving this equation,

12. CHAPTER 2
DETAIL DESIGN OF WING
The aircraft wings are the primary lift producing device for an aircraft.
The aircraft wings are designed aerodynamically to generate lift force
which is required in order for an aircraft to fly. Besides generating the
necessary lift force, the aircraft wings are used to carry the fuel
required for the mission by the aircraft, can have mounted engines or
can carry extra fuel tanks or other armaments. The basic goal of the
wing is to generate lift and minimize drag as far as possible. When the
airflow passes the wing at any suitable angle of attack, a pressure
differential is created. A region of lower pressure is created over the
top surface of the wing while, a region of higher pressure is created
below the surface of the wing. This difference in pressure creates a
differential force which acts upward which is called lift. For most
aircrafts, where, the wings are the primary structures to generate lift,
the aircrafts wings must generate sufficient lift to carry the entire
weight of an aircraft. In modern commercial, fighter and jet aircrafts,
the aircraft wings are not only designed to provide the necessary lift
during the different phases of flight, but also have a variety of other
roles and functions

13. WING DESIGN
WING COMPONENT-SPAR
The wing spars are the main load carrying structural
member of the aircraft wing. The wing spars are used to carry the
loads that occur during the flight (Flight loads) as well as carry
the weight of the aircraft wing while on the ground (ground
loads). The wing spars run throughout the root to the tip and can
be placed perpendicularly or at an angle. Commercial aircrafts
sometimes have a smaller number of wing spars than fighter
aircrafts, this is due to the fact that, the fighter aircrafts have to
deal with much higher flight loads. Therefore, the analysis has to
be very accurate. The structural analysis of the wing by defining
the primary load carrying member Spars is done below.
Spars are members which are basically used to carry the bending and
shear loads acting on the wing during flight. There are two spars, one
located at 15-20% of the chord known as the front spar, the other
located at 60-70% of the chord known as the rear spar. Some of the
functions of the spar include:
WING COMPONENET-STRINGER
The thickness of the skin determined above is too high for
the skin of an aircraft. Therefore, in order to reduce skin
thickness and redistribute the shear flow in the wing skin,
stringers are added. The number of stringers can be determined
by evaluating the amount by which the skin thickness should be
reduced.
The section selected for the stringers is the Z‐section with
end tabs. This section is selected as it gives the maximum area
moment of inertia for the minimum cross section area. The
properties of the stringer section selected are given below:
h=0.2
406m
t=0.0
1m
A = 0.004812 m2

14. CHAPTER-3
PRELIMINARY DESIGN OF AN AIRCRAFT FUSELAGE
Loads acting on fuselage:
s.no Components Distance
from
reference
line (m)
load
1 Crew 3.04 7.22 KN/m
2 nose landing gear 6.8 122.176 KN
3 Pay load 12.17 158.82 KN
4 Fuselage mass 15.88 109 KN
5 Main landing gear 18.52 1099.9 KN
6 Horizontal
Stabilizer
21.32 20.356KN/m
7 Vertical stabilizer 24.82 24.435KN/m
The fuselage is the main structure or body of the
fixed-wing aircraft. It provides space for cargo, controls,
accessories, passengers, and other equipment. In single-engine
aircraft, the fuselage houses the power plant. In multi engine
aircraft, the engines may be either in the fuselage, attached to the
fuselage, or suspended from the wing structure. There are two
general types of fuselage construction: truss and monocoque. The
monocoque (single shell) fuselage relies largely on the strength of
the skin or covering to carry the primary loads. The design may be
divided into two classes: 1. Monocoque and 2.Semi-monocoque

15. CHAPTER-4
DETAILED DESIGN OF AN AIRCRAFT FUSELAGE
BULK HEAD SHEAR FLOW DISTRIBUTION
From reference aircraft
Radius (R) = 2m
Fuselage section length = 32.84 m
Longeron height = 33mm
Longeron width = 20mm
Longeron thickness = 1mm
(top) no. followers = 8
(bottom) no.of.longerons = 8
Fineness ratio= (length of the body)/(max width of body)
Fineness ratio=32.84/41.16
Fineness ratio=0.7978

16. ----------
BULK HEAD SHEAR FLOW DISTRIBUTION

17. CHAPTER-5
BALANCING AND MANOEUVRING LOADS (TAIL PLANE,
RUDDER AND AILERON)
BALANCING LOADS:
A horizontal surface balancing load is a load
necessary to maintain equilibrium in any specified flight
condition with no pitching acceleration horizontal balancing
surface must be designed for the balancing loads occurring at
any point on the limit maneuvering envelope and in the flap
Conditions it is not required to balance the rudder because it
will not deflect due to gravity aileron will defect in vice
versa direction so it doesn’t require balancing load.
MANOEUVERING LOADS:
Each horizontal surface and its supporting structure, and the
main wing of a canard or tandem wing configuration, if that
surface has pitch control, must be designed for the maneuvering
loads imposed by the certain conditions:
>A sudden movement of the pitching control, at the speed VA,
to the maximum aft movement, and the maximum forward
movement, as limited by the control stops, or pilot effort,
whichever is critical.
>A sudden aft movement of the pitching control at speeds above VA,
followed by a forward movement of the pitching control resulting in
the foll owing combinations of normal and angular acceleration.

18. BALANCING AND MANEUVERING LOADS
Maximum pitch contr ol d isp lacemen t a t VA:
The airplane is assumed to be flying in steady level flight
and the cockpit pitch control is suddenly moved to obtain
extreme nose up pitching acceleration. In defining the tail load,
the response of the airplane must be taken into account.
Airplane loads that occur subsequent to the time when normal
acceleration at the
c.g. exceeds the positive limit maneuvering load or the resulting
tail plane normal load reaches its maximum, whichever occurs
first, need not be considered.
Specified contr ol d isp lacemen t:
A checked maneuver, based on a rational pitching control
motion vs. time profile, must be established in which the design
limit load factor will not be exceeded. Unless lesser values
cannot be exceeded, the airplane response must result in pitching
accelerations not less than the following:
a) A positive pitching acceleration (nose up) is assumed to be
reached concurrently with the airplane load factor of 1.0. The
positive acceleration must be equal to at least 39n(n-1)/v,
(rad/sec).Where, “n” is the positive load factor at the speed
under consideration; and V is the airplane equivalent speed in
knots.
A negative pitching acceleration (nose down) is assumed to be
reached on currently with the positive maneuvering load factor. This
negative pitching acceleration must be equal to at least -26n(n-1)/v,
(rad/sec).
These are the following conditions;
(a)Symmetric maneuvering conditions:
Where sudden displacement of a control is specified, the
assumed rate of control surface displacement may not be less
than the rate that could be applied by the pilot through the control
system. In determining elevator angles and chord wise load
distribution in the maneuvering conditions, the effect of
corresponding pitching velocities must be taken into account.
The in-trim and out-of-trim flight conditions must be considered.
(b)Maneuvering balanced conditions:
Assuming the airplane to be in equilibrium with zero pitching
acceleration, the maneuvering conditions on the maneuvering
envelope must be investigated.
(c) Pitch maneuver conditions:
The movement of the pitch control surfaces may be adjusted to take
into account limitations imposed by the maximum pilot effort, control
system stop and any indirect effect imposed by limitations in the
output side of the control system (for example, stalling torque or
maximum rate) obtainable by a power control system.

19. CHAPTER-6
WING ROOT ATTACHMENT
The fairings solve the three problems mentioned at the beginning.
A small attitude change no longer causes sudden
deterioration of airspeed.
Power-off glide is stable and the sink rate is much
reduced.
The plane no longer requires nose-up attitude in level
flight when heavily load.
There are additional benefits, such as improved climb rate,
increased cruise speed and reduced stall speed.
If the fuselage expands and contracts over the wings, as it
does for the 601 HDS, then these increases drag and reduces lift,
particularly for large payloads or high angle of attack. The
negative effects can be mitigated by a fairing that simulates a
constant width fuselage.
The fairings significantly improve climb rate, ceiling,
stall speed, sink rate at low speeds, optimum glide ratio,
minimum power to stay aloft, and stability of the plane when
CG is close to the rear limit.
Generally, the required angle of attack in level flight is significantly
reduced at or below cruise power settings. This manifests itself by a
much lower nose, particularly at gross weight
WING ROOT FAIRINGS:
Wing root fairings have substantially improved
low speed and high speed – load flying Characteristics.
The design process and result are described in hidden
drag.
For us low and slow fliers, it is convenient to
consider the total drag of an airplane to be composed of
parasite drag and induced drag. Parasite drag is the
resistance produced by irregular surfaces. The airflow is
disrupted by such surface and becomes turbulent.
Bending of smooth airflows creates induced drag.
It is easy to see the causes of parasite drag. For example,
un faired gear legs and external antennas are indicators. Induced
drag is harder to identify. A sleek-looking airplane may have lot
of induced drag and thus may not fly fast.
This is a story about hidden drag.

20. CHAPTER-7
DESIGN OF LANDING GEAR
LANDING GEAR ARRANGEMENT:
Landing gears normally come in two types: conventional
or "taildragger" landing gear, where there are two main wheels
towards the front of the aircraft and a single, much smaller,
wheel or skid at the rear; or tricycle landing gear, where there are
two main wheels (or wheel assemblies) under the wings and a
third smaller wheel in the nose.
To decrease drag in flight some undercarriages retract into
the wings and/or fuselage with wheels flush against the surface or
concealed behind doors; this is called retractable gear. With a
tricycle landing gear, the c.g is ahead of the main wheels, so the
aircraft is stable on the ground. It improves forward visibility on
the ground and permits a flat cabin floor for passengers and cargo
loading.
Thus, retractable tricycle landing gear system is selected
TYRE SIZING:
The “wheel” is the circular metal object upon which the
rubber “tyre” is mounted. The “brake” inside the wheel slows the
aircraft by increasing the rolling friction. However, the term
“wheel” is frequently used to mean the entire wheel/brake/tyre
assembly. The tyres are sized to carry the weight of the aircraft.
Typically, the main tyres carry about 90% of the total aircraft
weight. Nose tyres carry only about 10% of the static load but
experience higher dynamic loads during landing.
The nose gear is of double‐bogey type with two wheels.
The main gear consists of two sets of wheels (wing‐retracted)
each of multi‐bogey type with 4 wheels each.

21. LANDING GEAR CALCULATIONS
Wheel diameter = A WW
B
d = 1.59(56548)0.315
d = 51.19 inch = 1.30 m
Wheel Width = A WW
B
w = 0.1043(56548)0.48
w = 19.92 inch = 0.506 m
C on tact Ar ea :
WW = Ap * P
AP = 0.04103 m2
Rt = 0.628 m
T yr e Selection – Main Wh eel
Main wheel load = 254468.5 N
Wheel diameter = A WW
B
d = 1.63(254468.5)0.315
d = 82.21 inch = 2.08 m
Wheel Width = A WW
B
w = 0.1043(254468.5)0.48
w = 41.01 inch = 1.04 m
C on tact Ar ea :
WW = Ap * P
AP = 0.1846 m2
Rt = 0.985 m
R UNWAY L OADING:
For main wheel,
NOSE GEAR:
Load on nose gear = 0.1W0
=
0.1*65000
= 56500 kg
MAIN LANDING GEAR:
Load on main gear = 0.9W0
= 0.9*65000
= 58500kg
Tyre Selection – Nose Wheel
Nose wheel load WW = 56548 N
From Reymer book,
For fighter aircraft Diameter Width
A 1.59 0.0989
B 0.302 0.467
Nose landing gear Main landing gear
No. of wheel 1 2
Total load 5700 51300
Load by each wheel 5700 25650
Pressure type (psi) 200 200
Pressure type (bar) 13.78 13.78

22. CHAPTER-8
DETAIL DESIGN OF AN AIRCRAFT
Front view:
Front view
Side view:
Side view
Top view:
Top view

23. CONCLUSION
Design is a fine blend of science, presence of mind and the
application of each one of them at the appropriate time.
Design of anything needs experience and an optimistic
progress toward the ideal system the scientific society
always looks for the best product design.
This involves a strong fundamental and their skill full
application which is a tough job endowed upon the
designer, we had put enough had work to best of our
knowledge for this design.
A design never gets completed in a flutter sense but it is one further
step towards the ideal system. But during the design of this fighter
aircraft, we learned a lot about AERONAUTICS and its implication
when applied to an aircraft design.