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JAWAHARALAL
INSTITUTE OF TECHNOLOGY
(Approved by AICTE & Affiliated to Anna University)
COIMBATORE – 641 105
NAME : J.Dinesh Raja Ruban
REG.NO : 080101134017
SUBJECT : Aircraft Design Lab – I
COURSE : Aeronautical Engineering
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JAWAHARLAL INSTITUTE OF TECHNOLOGY
COIMBATORE – 641 105
DEPARTMENT OF AERONAUTICAL ENGINEERING
Certified that this is the bonafide record work done by
J.Dinesh Raja Ruban in the AIRCRAFT DESIGN LAB – I of this institution as
prescribed by the Anna University, Coimbatore for the Sixth semester during the
year 2010 – 2011.
Staff In charge: Head of the Department
University Register No.:080101134017
Submitted for the Practical Examination of the Anna University conducted on
……………
INTERNAL EXAMINER EXTERNAL EXAMINER
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ACKNOWLEDGEMENT
Firstly I would like to thank the Almighty God for always being by my side and providing me with
strength and capability to face all types of situations during this project tenure.
I extend my fullest and ever owing thanks to Dr.K.K.Babu, Principal, Jawaharlal Institute of
Technology, Coimbatore, for the academic freedom and inspiration.
With deep sense of gratitude, I extend my earnest & sincere thanks to our guideDr.Rajasekar
M.S., Ph.D, Head Of The Department, Aeronautical Engineering , Jawaharlal institute of technology,for
his systematic guidance, encouragement and for providing valuable insights offered over the course of
this project report.
I also thank everyone who lent us support in the completion of this project.
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Introduction:
There 2 classes of fighter aircraft. They are class-1 and class-2 fighter aircrafts. The class
one fighter is officially an air superiority fighter. Most of it can function either as multi role
fighters and ground attack. Air superiority fighter mainly does the function to gain air space
control over the enemy territory so that the bombers can bomb their targets safely, and give
support for the ground units. They literally make the enemy air space home ground for the
invaders aircrafts. Class2 fighters mainly concentrate on electronic warfare and ground attack
along with surveillance.
Today, complex sets of requirements and objectives include specification of airplane
performance, safety, reliability and maintainability, subsystems properties and performance, and
others. Some of these are illustrated in the table below
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--Good neighbor in peace --
Dectability in war
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Today, complex sets of requirements and objectives include specification of airplane
performance, safety, reliability and maintainability, subsystems properties and performance, and
others. Some of these are illustrated in the table below
Aircraft Design Objectives and Constraints
Issue Military
Dominant design criteria
Mission accomplishment and
survivability
Performance
Adequate range and response
Overall mission accomplishment
Airfield environment
Short-to-moderate runways
All types of runway surfaces
Often Spartan ATC, etc.
Limited space available
System complexity and mechanical
design
Low maintenance- availability
issue
Acceptable system cost
Reliability and survivability
Damage tolerance
Government regulations and community
Military standards
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acceptance --Performance and safety --
Reliability oriented
Low noise desirable
--Good neighbor in peace --
Dectability in war
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Mission Profile:
3 4 5
6 7
12 8 9
A
L
T
I
T
U
D
E
Range
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M ission
SegmentDescription Distance Time Al titude
1-2 Ground run 150 Meters 5 Seconds 0 Meter
2-3 Ascent 185 Meters 6 Minutes 14175 Meters
3-4 Cruising 2000 Meters 4 Seconds 14175 Meters
4-5 Aerobatic 1000 Meters 2 Seconds 14175 Meters
5-6 Nose down 4175 Meters 4 Seconds
14175-10000
Meters
6-9 Descent 200 Km 3 Minutes 10000-0 Meters
9-10 Halt 100 Meters 4 Seconds 0 Meter
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[W5/W4]@6 = e-[(700*0.6) / (3186.72*6)]
[W5/W4]@6 = 0.9782
[W5/W4]@8 = e-[(700*0.6) / (3186.72*8)]
[W5/W4]@8 = 0.9836
[W5/W4]@10 = e-[(700*0.6) / (3186.72*10)]
[W5/W4]@10 = 0.9869
[W5/W4]@12 = e-[(700*0.6) / (3186.72*12)]
[W5/W4]@12 = 0.9890
For the nose down maneuver, from the historical data,
W6 / W5 = 0.9860
Next segment in the mission profile is cruise before halting.
Range = 200
G.S.R = 200/1.5
G.S.R = 133.33 Km
Altitude = 9175 m ≈ 10 Km
Pressure = 2.6500×104
W5/W4= e-RC
t/ V×(L/D)
max
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Density = 4.1351×10-1
a = √ (1.4×26500×104) / (4.1351×10-1)
a = 299.53 m/s
a = 1078.30 Km/hr
As mach number is 3,
V∞ = a×3
V∞ = 1078.30×3
V∞ = 3234.924 Km/hr
Time = 133.33 / 3234.924
Time = 0.04121 hours
Head wind = 15 m/s
Head wind = 54 Km/hr
Actual additional distance = 54×0.04121
Actual additional distance = 2.2256 Km
Total rate range = 200 + 2.2256
Total rate range = 202.2256 Km
a = γP/ρ
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For cruising,
[W7/W6]@6 = e-[(202.22*0.6) / (3234.924*6)]
[W7 / W6]@6 = 0.9937
[W7/W6]@8 = e-[(202.22*0.6) / (3234.924*8)]
[W7 / W6]@8 = 0.9953
[W7/W6]@10 = e-[(202.22*0.6) / (3234.924*10)]
[W7 / W6]@10 = 0.9962
[W7/W6]@12 = e-[(202.22*0.6) / (3234.924*12)]
[W7 / W6]@12 = 0.9968
For descending, the aircraft is assumed to consume less amount of fuel.
W8/W7 = 0.97
For landing and halting, fuel consumption is very less
W9/W8 = 0.99
[W9/W0]@6 = 0.85*0.960*0.9936*0.9752*0.9782*0.986*0.9937*0.97*0.99
W7 / W6 = e- RC
t/V×(L/D)max
W9/W0=(W1/W0)×(W2/W1)×(W3/W2)×(W4/W3)×
(W5/W4)×(W6/W5)×(W7/W6)×(W8/W7)×(W9/W8)
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[We/Wg]@6 = 1.202[2.202*52323.58]-0.06
We/Wg@6 = 0.5068
[We/Wg]@8 = 1.202[2.202*43132.17]-0.06
0
10000
20000
30000
40000
50000
60000
0 2 4 6 8 10 12 14
W g
(L/D)Max
Wg Vs (L/D)Max
Series1
We/Wg = 1.202 [2.202 Wg]-0.06
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We/Wg@8= 0.5127
[We/Wg]@10 = 1.202[2.202*41401.87]-0.06
We/Wg@10= 0.5140
[We/Wg]@12 = 1.202[2.202*40331.26]-0.06
We/Wg@12= 0.5148
For L/D = 6,
WPay= 0.1146
Wf /Wg = 0.2889
We/Wg= 0.5068
Wg = 52323.58 Kg
For L/D = 8,
WPay= 0.1391
Wf /Wg = 0.2566
We/Wg= 0.5127
Wg = 43132.17 Kg
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For L/D = 10,
WPay= 0.1449
Wf /Wg = 0.2493
We/Wg= 0.5140
Wg = 41401.87 Kg
For L/D = 12,
WPay= 0.1487
Wf /Wg = 0.2445
We/Wg= 0.5148
Wg = 40331.26 Kg
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Where,
ρ = 1.225 Kg/m3
(W/S)Land = 1/2×1.225× (31.40)2×
3
(W/S)Land = 1811.70
(W/S)Land = 1/2×ρV2SCLMax
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LANDING WEIGHT OF AN AIRCRAFT:
Let the landing weight ratio = 0.62
Landing weight = 32500 Kg
SLand = W/1811.70
SLand = 32500/1811.70
SLand = 17.93 m
Stall velocity, Vs@+10% = 31.40+31.40×(10/100)
Stall velocity, Vs@+10% = 34.54 m/s
(W/S)Land = 1/2×1.225× (34.54)2×
3
(W/S)Land = 2192.15
SLand = 32500 / 2192.15
SLand = 14.82 m
Stall velocity, Vs@-10% = 31.40 - 31.40× (10/100)
Stall velocity, Vs@+10% = 28.26 m/s
(W/S)Land = 1/2×ρV2SCLMax
(W/S)Land = 1/2×ρV2SCLMax
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(W/S)Land = 1/2×1.225×(28.26)2×
3
(W/S)Land = 1467.47 Kg
SLand = 32500 / 1467.47
SLand = 22.146 m
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SELECTION OF WING LOADING:
VMax = 1.1× 3186.72
VMax = 3505.392 Km/hr
Log SWet = {0.5+0.5log10[52323.58×2.202]} / (3.29)2
Swet = 39.506 m2
K = 1/πeA
K = 1/ (π×0.7×3)
K = 0.0757
Where,
CD0 = 0.005× (39.506/17.93)
CD0 = 0.0110
CD = 0.0110 + (0.0757×32)
CD = 0.692
CD = CD0 + KCL2
VMax = 1.1× VCruise
CD0 = Cfe × (Swet/S)
T/W = CD×1/2×ρV2Max / (W/S)
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T/W = {0.692×1/2×1.225×(973.72)2}/1811.70
T/W = 221.81
CD0 = 0.005× (39.506/14.82)
CD0 = 0.0133
CD = 0.0133+ (0.0757×32)
CD = 0.694
T/W = {0.694×1/2×1.225×(973.72)2}/2190.15
T/W = 184.01
CD0 = 0.005× (39.506/22.146)
CD0 = 8.91×10-3
CD0 = Cfe × (Swet/S)
T/W = CD×1/2×ρV2Max / (W/S)
CD0 = Cfe × (Swet/S)
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CD = 8.91×10-3
+ (0.0757×32)
CD = 0.690
T/W = {0.690×1/2×1.225 × (973.72)2}/1467.47
T/W = 273.14
T/W = 3090, CD = 0.692
3090 = 0.692 × (1/2) ×1.225 × (3505.39)2 × (S/W)
(W/S) = 1.6848×103
T/W = 2595.5, CD = 0.694
2595.5 = 0.694 × (1/2) ×1.225 × (3505.39)2 × (S/W)
(W/S) = 2.012×10
3
T/W = 3753.71, CD = 0.690
T/W = CD×1/2×ρV2Max / (W/S)
(T/W) = CD× 1/2×ρV2Max(S/W)
(T/W) = CD× 1/2×ρV2Max(S/W)
(T/W) = CD× 1/2×ρV2Max(S/W)
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3753.71 = 0.690 × (1/2) × 1.225 × (3505.39)2 × (S/W)
(W/S) = 6.420×104
WT0 = 52323.58 Kg
C = 0.1, D = 0.5
Log (SWet) = 0.1 + 0.5log(52323.58)
SWet = 39.506 m2
(W/S)take off = 1811.7 / 0.62
(W/S)take off = 2922.07
CD0 =0.005 × (39.506/17.90)
CD0 = 0.0110
Log (SWet) = C + Dlog(WT0)
(W/S)take off = (W/S)landing / 0.62
CD0 = Cfe × (Swet/Stake off )
CD = {Cfe × Swet(W/S)take off }/ {Wtake off × 9.81}
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CD = {0.005×21.458×2438.35} / {52323.58 × 9.81}
CD = 5.095×10-4
CD Clean = 0.0110 + (1/πeA)×32 + 0
CD Clean = 0.0110 + (1/π×0.8×6)×32
CD Clean = 0.6078
CD Take off = 0.0110 + (1/πeA)×32
+ 0.01
CD Take off = 0.0110 + (1/π×0.75×6)×32+ 0.01
CD Take off = 0.6576
CD Landing = 0.0110 + (1/πeA)×32 + 0.05
CD Landing = 0.0110 + (1/π×0.7×6)×32+ 0.05
CD Clean = CD0 + KCL2 + ΔCD0
CD Take off = CD0 + KCL2 + ΔCD0
CD Landing = CD0 + KCL2 + ΔCD0
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CD Landing = 0.7430
CD Landing Gear = 0.0110 + 0.020
CD Landing Gear = 0.031
F1 = 3×0.005
F1 = 0.015
F2 = (0.0110 – 0.015) / 1811.7
F2 = -2.2078×10-6
F3 = 1.8643×10
-7
CD Landing Gear = CD0 + ΔCD0
F1 = 3×Cfe
F2 = (CD0 – F1) / (W/S)
F3 = K/q2
F3 = 1/[πeA(1/2×ρV2Max)]
2
(W/S)Max = F1/F3
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(W/S)Max = √0.015/1.8643×10-7
(W/S)Max = 283187.3 m2
(W/S)Max = 283.18×103 m2
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Airfoils:
One of the difficulties in designing a good airfoil is the requirement for acceptable off-
design performance. While a very low drag section is not too hard to design, it may separate at
angles of attack slightly away from its design point. Airfoils with high lift capability may
perform very poorly at lower angles of attack.One can approach the design of airfoil sections
with multiple design points in a well-defined way. Often it is clear that the upper surface will be
critical at one of the points and we can design the upper surface at this condition. The lower
surface can then be designed to make the section behave properly at the second point. Similarly,
constraints such as Cmo are most affected by airfoil trailing edge geometry.When such a
compromise is not possible, variable geometry can be employed (at some expense) as in the case
of high lift systems.
Airfoil Parameters:
For my Aircraft, the selected airfoil is NACA 6 digit series. That is NACA 64a204.
The taper ratio of the wing is 0.295
Leading edge sweep angle is 46º
Root chord of the wing is 8.61m
Tip chord of the wing is 2.067m
Finess ratio is 0.24
Planform area is 171.098 m
2
Aspect Ratio is 6
Span of the wing is 32.04 m
Half span is 16.02 m
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Aerofoil:
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Wing Design:
There are essentially two approaches to wing design. In the direct approach, one finds the
planform and twist that minimize some combination of structural weight, drag, and CLmax
constraints. The other approach involves selecting a desirable lift distribution and then
computing the twist, taper, and thickness distributions that are required to achieve this
distribution. The latter approach is generally used to obtain analytic solutions and insight into the
important aspects of the design problem, but is is difficult to incorporate certain constraints and
off-design considerations in this approach. The direct method, often combined with numerical
optimization is often used in the latter stages of wing design, with the starting point established
from simple (even analytic) results.
Wing lift distributions play a key role in wing design. The lift distribution is directly related to
the wing geometry and determines such wing performance characteristics as induced drag,
structural weight, and stalling characteristics. The determination of a reasonable lift and Cl
distribution, combined with a way of relating the wing twist to this distribution provides a good
starting point for a wing design. Subsequent analysis of this baseline design will quickly show
what might be changed in the original design to avoid problems such as high induced drag or
large variations in Cl at off-design conditions.
Parameters:
Span:
Selecting the wing span is one of the most basic decisions to make in the design of a wing. The
span is sometimes constrained by contest rules, hangar size, or ground facilities but when it is not
we might decide to use the largest span consistent with structural dynamic constraints (flutter).
This would reduce the induced drag directly.However, as the span is increased, the wing
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structural weight also increases and at some point the weight increase offsets the induced drag
savings. This point is rarely reached, though, for several reasons.
1. The optimum is quite flat and one must stretch the span a great deal to reach the actual
optimum.
2. Concerns about wing bending as it affects stability and flutter mount as span is increased.
3. The cost of the wing itself increases as the structural weight increases. This must be
included so that we do not spend 10% more on the wing in order to save .001% in fuel
consumption.
4. The volume of the wing in which fuel can be stored is reduced. It is more difficult to
locate the main landing gear at the root of the wing.
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The selection of my aircraft wing is based on the following. They are
Landing speed/landing distance
Maximum speed VMax
Absolute ceiling
Rate of climb
Based on range
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SELECTION OF WING BASED ON
(A)
LANDING SPEED/LANDING DISTANCE:
Landing ground run:
It is the actual distance the airplane travels from the time the wheel first touch to the time
the airplane comes to a halt.
Landing distance:
It is the horizontal distance the airplane covers from being at the screen height (15m) till
comes to a halt.
Sland = Landing distance
Sland = 500m
VA - Approach Velocity
VA= 1.71× √500 m/s
VA= 40.8 m/s
Vs - Stalling Speed
Vs = VA/1.3
Vs = 31.40 m/s
CLmax – Maximum lift coefficient
CLmax – To be taken from various reference airplanes.
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The value of CLmax depends on the following.
(a) Wing geometry i.e. aspect ratio (A), taper ratio (λ) andsweep (Λ).
(b)Airfoil shape.
(c) Flap type, ratio of flap area to wing area (Sflap/S) and flap deflection (δflap).
(d) Type of leading edge slat and its deflection.
(e) Reynolds number.
(f) Surface texture.
(g) Interference effect from fuselage, nacelle and pylons.
(h) Influence of propeller slip stream, if present.
Density ρ at the landing airport = 1.225 kg/m3
Wing loading (W/S) based on landing distance = CLmax× (ρ×Vs
2
)/2 (N/m2
)
= (3×1.225×402)/2
= 1811.7 N/m2
To be estimated from mission,
Fuel consumption and disposed weight = 0.62
Wland/Wtake-off = [1- (Wfuel/ Wtake-off ) - (Wdisposible/Wtake-off )]
(W/S)take-off = (W/S)landing / (Wland/Wtake-off )
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(W/S)take-off = 1811.7/ (32500/52323.58)
(W/S)take-off = 1811.7/0.62
(W/S)take-off = 2922.09 N/m2
0
5
10
15
20
25
30
35
40
0 500 1000 1500 2000 2500
V S t a l l
W/S
VStall Vs W/S
Series1
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B) MAXIMUM SPEED VMax:
The optimization from Vmax consideration aims at finding out the wing loading which will
result in the lowest thrust requirement for a chosen Vmax at Hcr (cruise altitude).
1)
Estimation of Drag polar
(A)
Estimation of CD0
C and d are constant based on type of airplane
Wtake-off = 52323.58 kg
C = 0.1
d = 0.5
log10 Swet = (0.1) + (0.5) log10 (52323.58×2.205)
Swet = 39.506
Cfe = Equivalent skin friction drag coefficient (varies from 0.0025 to 0.0065 based on type of
aero plane) = 0.005
Swet = Wetted area
Swet= 39.506 m2
Log10Swet = C + d log10Wtake-off
CD0 = CfeSwet / S = CfeSwet (W/S)take-off /Wtake-off
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CD0= 0.005 × 39.506 × 2922.09 / 52323.58×9.81
CD0 = 0.00124
(B)
Estimation of K:
For low speed airplane K = 1/(π×A×e)
Where,
A = Aspect ratio
Aspect ratio = 6
e = Oswald efficiency factor (lies between 0.8 to 0.88 with unswept wing)
e swept wing = e unswept wing × cos (Λ0-5)
e = 0.7
K = 1/(π×A×e)
K = 1/(π×6×0.7)
K =0.0757
1)
Estimation of drag polar:
CD= 1.2 × 10-3
× 0.0757 × CL2
CD0 = Cfe× Swet / Stake-off = Cfe × Swet × (W/S)take-off / Wtake-off
CD = CD0 + k CL2
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CD Clean cleans drag polar can be used while calculating maximum rate of climb and
subsonic cruise cases.
Configuration ΔCD0 e
Landing gear 0.015 to 0.025 No effect
Landing flaps 0.05 to 0.075 0.7 to 0.75
Take-off flaps 0.05 to 0.075 0.75 to 0.8
Clean - 0.8 to 0.85
Take-off flaps,
ΔCD0 (1) =0.01
e(1) = 0.75
Landing flaps,
ΔCD0 (2) = 0.05
e(2) = 0.7
Landing gears,
ΔCD0(3) = 0.015
K 1 = 1/(π×6×0.75)
K 1 = 1/(π×A×e)
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K 1 = 0.0707
K 2 = 1/(π×6×0.7)
K 2 = 0.075
Take-off flaps, landing gear up
ΔCD0 (4) = 0.01
Take-off flaps, landing gear up
ΔCD0(5) = 0.025
Landing flaps, gears up
ΔCD0(6) = 0.05
Landing flaps, gears down
ΔCD0(7) = 0.065
DRAG POLAR FOR DIFFERENT CONFIGURATION:
K 2 = 1/(π×A×e)
ΔCD0 (4) = ΔCD0 (1)
ΔCD0 (5) = ΔCD0 (1) + ΔCD0(3)
ΔCD0 (6) = ΔCD0 (2)
ΔCD0 (7) = ΔCD0 (2) + ΔCD0 (3)
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Clean Configuration,
CD = 1.2 ×10-3+0.0757CL2
Take-off flaps, landing gear up
CD = 0.011 + 0.0707 CL2
Take-off flaps, landing gear down
CD = 0.02612 + 0.0707 CL2
Landing flaps, gear up
CD= 0.05112 + 0.075 CL2
Landing flaps, gear down
CD = 0.01112 + 0.07 CL2
BREAK-UP OF DRAG POLAR:
CD = CD0+KCL2
CD = CD0 + ΔCD0 (4) + K 1 × CL2
CD = CD0 + ΔCD0 (5) + K 1×CL2
CD = CD0 + ΔCD0 (6) + K 2×CL2
CD = CD0 + ΔCD0 (7) + K2×CL2
CD = F1 + F2 × (W/S) + F3 × (W/S)2
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Mmax = 3.04
Vmax = 1033.6 m/s
F3=1/[(π×6×0.7)×(0.5×1.225×1033.62)
2]
F3=1.77×10-13
Wing loading for the lowest thrust requirement (TVmax) for chosen at a given ceiling Hcr
(W/S) = (0.5×ρ×Vmax2)×√(F1× π A e)
(W/S) = (0.5×1.2165×10-1
×88.192
) × (F1× πAe)
(W/S) = 22662.38 N/m2
Vmax= Mmax × speed of sound
F3 =(π×A×e(0.5×ρ×Vmax2)
-1
(W/S) = CD × qmax/(TVmax/W)
(W/S) = (F1/F3)
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0
50
100
150
200
250
300
350
400
450
0 1000 2000 3000 4000 5000 6000
V m a x
W/S
Vmax Vs (W/S)
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C) ABSOLUTE CEILING (Hmax):
At absolute ceiling the flight is possible only at one speed at which
Thrust required = Thrust minimum = Drag minimum
If the flight velocity (VHmax) at the absolute ceiling
Thrust Loading,
Both VHmax and Hmax are prescribed, then
CD0 =1.12×10-3
K =0.0757
ρHmax = 1.2165×10-1
kg/m3
= (√1.12×10-2
/0.0757)×(0.5×1.2165×10-1
×885.192)
= 7.3984×10-3
×VHmax2
Thrust Loading (T/W) = √(4CD0 K) = √(4k(F1+F2(w/s))
(T/W) = (0.5×ρHmax×VHmax2)×2CD0/(W/S)
(W/S) = (CD0/K)×(0.5×ρHmax×VHmax2)
(W/S) = (CD0/k)×(0.5×ρHmax×VHmax2)
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200
300
400
500
600
700
800
900
0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000
V H M a x
W/S
VH Max Vs (W/S)
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(D) RATE OF CLIMB:
Wing loading is such that thrust required is minimum for the specified rate of climb (Vc).
For chosen flight velocity V, the optimum wing loading = (W/S)opt = (F1/F3) 0.5
(tR/C)v = (Vc/V) + q(F1/PR/C + F3 + F3PR/C)
For each V, we have an optimum ρ
Each curve corresponding to each V has a minimum. To get the minimum of the minima,
we draw an envelope which is tangential to all the curves. Minimum of this envelop gives the
optimum wing loading from the rate of climb consideration and the corresponding minimum
thrust loading (tR/Cmax)min.
For Jet Airplanes,
(L/D)max = (4 CD0 K)-0.5
(L/D)max = (4×1.12×10-3
× 0.0757)-0.5
(L/D)max = 54.3016
For propeller Airplanes,
V(R/C)max ={(T/W)(W/S)/3ρ∞CD.0[1+ (1+(3/(L/D)max2(T/W)
2)]}
1/2
V(R/C)max = {2/ ρ∞ ((K/3CD0)(W/S))}1/2
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V(R/C)max = [(1*(W/S))/ (3*0.22785*1.12*10-3) {1+√ (1+ (3/54.30162
))}]1/2
V(R/C)max = [2613.32 (W/S)]1/2
(W/S) = V(R/C)2/2613.32
0
100
200
300
400
500
600
700
800
900
0 50 100 150 200 250 300
V R / C M
a x
W/S
VR/C Max Vs (W/S)
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(E) BASED ON RANGE:
For jet airplanes,
The density on altitude and the TSFC depends on flight velocity and altitude. Optimum
wing loading for a chosen Vcr is obtained at different altitude. Minimum of this curve gives the
optimum wing loading and corresponding cruising altitude.
For Jet airplanes,
V(Cl1/2
/CD)max = [2/0.227885√[(3×0.0757)/(1.12×10-3
)] (W/S)]1/2
V(Cl1/2
/CD)max= 117.17 × (W/S)1/2
W/S = 3000 N/m2
WTake off =52323.58 Kg
L/D = CL/CD = 6
CLMax = 3
CD0 = 112×10-3
K = 0.0757
WFuel/WTake off = 0.288
W/S = (F1+CD0)×(R/3.6× (ρ×q/2)×(TSFC)/Wtake-off ×(Wtake-off /Wf)
V(Cl1/2
/CD)max = {2/ ρ∞× (3K/CD0) × (W/S)}1/2
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Range = 2000 Km
Mach number = 3
Cruise altitude = 14000 m
Absolute ceiling = 18000 m
Rate of climb, V∞ = 720 m/s
Rate of climb angle = sin-1
(Ve/V∞) = 14º
Climb velocity = 174.18 m/s
Landing distance = 600 m
ηMax = (3×1/2×0.2278×(720)2)/3000
ηMax = 59.058
Where,
C - Constant (C= 0.6)
Maximum efficiency,
ηMax = CL Max×1/2×ρV2/(W/S)
WWing = C×Sw×AR×(t/c)-0.4 (1+taper ratio)0.1
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Sw - Wing area
AR - Aspect ratio
t/c - Finess ratio
W/S = 3000
W = 52323.58 Kg
W = 513294.31 N
S = 513294.31/3000
S = 171.09 m
2
Aspect ratio = 6
Span2 = 171.09×6
Span2 = 1026.54
Span = 32.04
Sweep angle = 60º
Aspect ratio = span2/plan form area
Taper ratio = Tip chord/Root chord
Span2 = Plan form area × Aspect ratio
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Cos(Sweep)-1
= Cos(60º)-1
Cos(Sweep)-1
= 2
Mean chord =171.09/32.04
Mean chord = 5.33
(Ct+Cr )/2 = 10.67
Ct+Cr = 10.67
Ct/Cr = 0.24
Ct = 0.24Cr
0.24Cr +Cr = 10.67
1.24Cr = 10.67
Cr = 8.61m
Ct = 0.24 × 8.61
Ct = 2.067m
Plan form area = Span × Mean chord
Mean chord = (Ct+Cr)/2
Taper ratio = Ct/Cr
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0.30 × 52323.58= 0.6×28.14×6 (t/c)-0.4
×(1.24)0.1
×2
(t/c)-0.4
= 151.66
t/c = 7.06×10-6
WWing = C×Sw×AR×(t/c)-0.4
(1+taper
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100
200
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400
500
600
700
800
900
1000
0 1000 2000 3000 4000 5000 6000 7000
( V C L / C D ) 1 / 2
WS
(VCL/CD)1/2 Vs WS
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Wing with centre of gravity
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Cabin Layout and Fuselage Geometry:
The design of the fuselage is based on payload requirements, aerodynamics, and
structures. The overall dimensions of the fuselage affect the drag through several factors.
Fuselages with smaller fineness ratios have less wetted area to enclose a given volume, but more
wetted area when the diameter and length of the cabin are fixed. The higher Reynolds number
and increased tail length generally lead to improved aerodynamics for long, thin fuselages, at the
expense of structural weight. Selection of the best layout requires a detailed study of these trade-
offs, but to start the design process, something must be chosen. This is generally done by
selecting a value not too different from existing aircraft with similar requirements, for which
such a detailed study has presumably been done. In the absence of such guidance, one selects an
initial layout that satisfies the payload requirements.The following sections are divided into
several parts: the selection of cabin cross-section dimensions, determination of fuselage length
and shape, FAR's related to fuselage design and seating, and finally considerations related to
supersonic aircraft.
Cross-Section Shape:
It is often reasonable to start the fuselage layout with a specification of the cross-section:
its shape and dimensions.
Most fuselage cross-sections are relatively circular in shape. This is done for two reasons
By eliminating corners, the flow will not separate at moderate angles of attack or
sideslip
When the fuselage is pressurized, a circular fuselage can resist the loads with
tension stresses, rather than the more severe bending loads that arise on non-
circular shapes.
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Fuselage sizing:
We can fine the size of the fuselage by using the formula,
where,
a – 1.0 – 1.8
b – 0.5 – 0.25
lf = 0.4 × (52323.58)0.4
lf = 0.4 × 77.17
lf = 30.87 feet
lf = 9.4092
lf = 9.41
Cockpit:
The cockpit must be designed in such a way that, the pilot can visible the runway clearly.
The angle given for our aircraft is 11º. This is because the pilot will not get the sight problem to
look the ground while taxing.
lf = aW0
c
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Seats in the cockpit:
The seats used for our aircraft is the ejection type seat.
As this is a fighter aircraft to save the life of the pilot we are using the ejector seat model.
As it is important to give him enough comfort to the pilot we have to select the space in
cockpit and other things as per the comfortness of pilot.
The seat pitch is 90cm with the seat width of 55cm.
The head room must be in 165cm above.
The aisle height is greater than 193cm.
Volume per passenger is 0.14m
3.
Tail sizing:
We are using the following formula to find the horizontal tail size,
Cht = lht×sht
Formula to find the horizontal tail size is,
Cvt = lvt×svt
Cht/cw = (lht/lw) × (sht/sw)
Cvt/cw = (lvt/lw) × (svt/sw)
Cht/cw+ Cvt/cw = (lht/lw) × (sht/sw) + (lvt/lw) × (svt/sw)
Sht/Sw = (Cht/Cw) × (lw/lht)
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Svt/Sw = (Cvt/Cw) × (lv/lvt)
For horizontal tail,
Aspect ratio, AR = 3
Taper ratio, TR = 0.3
For vertical tail,
Aspect ratio, AR = 1.5
Taper ratio, TR = 0.5
(t/c)Tail = 6.354×10-6
Weight:
The structural weight of our aircraft is 1150Kg
The propulsion weight of our aircraft is 1500Kg
The fixed equipment weight of our aircraft is 350Kg
The empty weight of our aircraft is 3000Kg
The weapons weight of our aircraft is 1500 Kg
(t/c)Tail = 0.9×(t/c)Wing
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Engine Selection:
As our aircraft is the fighter aircraft flying at a speed of Mach 3, we selected the turbo fan
engines
Two engines are located in the wing
The dry thrust is assumed to be of 76.4 KN (17,185 lbf)
The thrust with after burner is assumed to be of 109.8 KN (24,675 lbf)
Engine Placement:
The arrangement of engines influences the aircraft in many important ways. Safety,
structural weight, flutter, drag, control, maximum lift, propulsive efficiency, maintainability, and
aircraft growth potential are all affected. Engines may be placed in the wings, on the wings,
above the wings, or suspended on pylons below the wings. They may be mounted on the aft
fuselage, on top of the fuselage, or on the sides of the fuselage. Wherever the nacelles are placed,
the detailed spacing with respect to wing, tail, fuselage, or other nacelles is crucial. Engines
buried in the wing root have minimum parasite drag and probably minimum weight. Their
inboard location minimizes the yawing moment due to asymmetric thrust after engine failure.
However, they pose a threat to the basic wing structure in the event of a blade or turbine disk
failure, make it very difficult to maximize inlet efficiency, and make accessibility for
maintenance more difficult.
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Sukhoi Su-27
Manufacturer Soviet Union
First flight 20 May 1977
Introduced
ENGINE
Model Saturn/ Lyul'ka AL-31F afterburning turbofansThrust
No. Of engines 2
Dry thrust 33,510 lb (149.06 kN)
Thrust with afterburner 55,116 lb (245.18 kN)
Fuel capacity
ACCOMMODATION
Crew 1
WEIGHT RATIOS
Empty Weight 51,015 lb (23,140 kg)
Gross Weight
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Loaded Weight 62,390 lb (28,300 kg)
Maximum Take Off Weight 72,750 lb (33,000 kg)
Max Payload 8,820 lb (4,000 kg)
Armaments
Fuel Capacity 20,725 lb (9,400 kg)
FUSELAGE
Length 71.92 ft (21.94 m)
Height 19.42 ft (5.92 m)
WING
Area (m2) 667.8 ft² (62.04 m
2)
Span 48.17 ft (14.70 m)
PERFORMANCE
Maximum SpeedMach 2.35 1,555 mph (2,500 km/h) at 36,090
ft (11,000 m)
Endurance
Service Ceiling 18,500 m (62,523 ft)
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Range 3,530 Km (2,070 mi)
Rate of climb 300 m/s
Wing loading 371 Kg/m2
Thrust/weight 1.09
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Sukhoi Su-35
Manufacturer Soviet Union
First flight -
Introduced -
ENGINE
Model Saturn 117S with TVC nozzle turbofan
No. Of engines 2
Dry thrust 8,800 kgf (86.3 kN, 19,400 lbf) each
Thrust with afterburner 14,500 kgf (142 kN, 31,900 lbf) each
Fuel capacity -
ACCOMMODATION
Crew 1
WEIGHT RATIOS
Empty Weight 18,400 kg (40,570 lb)
Gross Weight
Loaded Weight 25,300 kg (56,660 lb)
Maximum Take Off Weight 34,500 kg (76,060 lb)
Armaments -
FUSELAGE
Length 21.9 m (72.9 ft)
Height 5.90 m (19.4 ft)
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WING
Area 62.0 m² (667 ft²)
Span 15.3 m (50.2 ft)
PERFORMANCE
Maximum Speed Mach 2.25 (2,390 km/h, 1,490 mph) at altitude
Endurance -
Service Ceiling 18,000 m (59,100 ft)
Range3,600 km (1,940 nmi) ; (1,580 km, 850 nmi
near ground level)
Ferry Range 4,500 km (2,430 nmi) with external fuel tanks
Rate of climb >280 m/s (>55,100 ft/min)
Wing loading 408 kg/m² (84.9 lb/ft²)
Thrust/weight 1.1
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Sukhoi Su-25
Manufacturer Soviet Union
First flight 22 February 1975
Introduced -
ENGINE
Model Tumansky R-195 turbojets
No. Of engines 2
Thrust 44.18 kN (9,480 lbf) each
ACCOMMODATION
Crew 1
WEIGHT RATIOS
Empty Weight 10,740 kg (23,677 lb)
Gross Weight -
Loaded Weight 16,990 kg (37,456 lb)
Maximum Take Off Weight 20,500 kg (45,194 lb)
Armaments 4,400 kg (9,700 lb)
FUSELAGE
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Length 15.33 m (50 ft 11)
Height 4.80 m (15 ft 9 in)
WING
Area 30.1 m² (324 ft²)
Span 14.36 m (47 ft 1 in)
PERFORMANCE
Maximum Speed 950 km/h (590 mph, Mach 0.77)
Endurance -
Service Ceiling 10,000 m (22,200 ft)
Range 2,500 km (1,553 mi)
Rate of climb 58 m/s (11,400 ft/min)
Wing loading 584 kg/m² (119 lb/ft²)
Thrust/weight 0.51
Combat radius 375 km (235 mi)
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Sukhoi Su-24
Manufacturer Soviet Union
First flight -
Introduced -
ENGINE
Model
Saturn/Lyulka AL-21F-3A
afterburningturbojet engines
No. Of engines 2
Dry thrust 75 kN (16,860 lbf) each
Thrust with afterburner 109.8 kN (24,675 lbf) each
Fuel capacity 11,100 kg (24,470 lb)
ACCOMMODATION
Crew Two (pilot and weapons system operator)
WEIGHT RATIOS
Empty Weight 22,300 kg (49,165 lb)
Gross Weight -
Loaded Weight 38,040 kg (83,865 lb)
Maximum Take Off Weight 43,755 kg (96,505 lb)
Armaments Up to 8,000 kg (17,640 lb) hard points
FUSELAGE
Length 22.53 m (73 ft 11 in)
Height 6.19 m (20 ft 4 in)
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WING
Area 55.2 m² (594 ft²)
Span
17.64 m extended, 10.37 m maximum sweep
(57 ft 10 in / 34 ft 0 in)
PERFORMANCE
Maximum Speed 2,320 km/h (1,440 mph)
Endurance
Service Ceiling 11,000 m (36,090 ft)
Range 2,775 km (1,500 nm, 1,725 mi)
Rate of climb 150 m/s (29,530 ft/min)
Wing loading 651 kg/m² (133 lb/ft²)
Thrust/weight 0.60
Takeoff roll 1,550 m (5,085 ft
Landing roll 1,100 m (3,610 ft)
G-Force limit 6
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Sukhoi Su-17
Manufacturer Soviet Union
First flight -
Introduced 1970
ENGINE
Model Lyulka AL-21F-3 afterburningturbojet
No. Of engines 1
Dry thrust 76.4 kN (17,185 lbf)
Thrust with afterburner 109.8 kN (24,675 lbf)
Fuel capacity 3,770 kg (8,310 lb)
ACCOMMODATION
Crew 1
WEIGHT RATIOS
Empty Weight 12,160 kg (26,810 lb)
Gross Weight
Loaded Weight 16,400 kg (36,155 lb)
Maximum Take Off Weight
Armaments Up to 4000 kg (8,820 lb)
FUSELAGE
Length 19.02 m (62 ft 5 in)
Height 5.12 m (16 ft 10 in)
WING
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Area
Spread: 38.5 m² (415 ft²)
Swept: 34.5 m² (370 ft²)
Span
Spread: 13.68 m (44 ft 11 in)
Swept: 10.02 m (32 ft 10 in)
PERFORMANCE
Maximum Speed
Sea level: 1,400 km/h (755 knots, 870
mph)
Altitude: 1,860 km/h (1,005 knots,
1,155 mph, Mach 1.7)
Endurance -
Service Ceiling 14,200 m (46,590 ft)
Ferry Range 2,300 km (1,240 nmi, 1,430 mi)
Combat Range1,150 km (620 nm, 715 mi) in hi-lo-hi attack
with 2,000 kg (4,410 lb) war load
Rate of climb 230 m/s (45,275 ft/min)
Wing loading 443 kg/m² (90.77 lb/ft²)
Thrust/weight 0.68
G-force limit 7
Airframe lifespan 2,000 flying hours, 20 years
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Sukhoi Su-7
Manufacturer Soviet Union
First flight -
Introduced 1955
ENGINE
Model Lyulka AL-7
No. Of engines 1
Dry thrust 66.6 kN (14,980 lbf)
Thrust with afterburner 94.1 kN (22,150 lbf)
Fuel capacity 3,220 kg (7,100 lb)
ACCOMMODATION
Crew 1
WEIGHT RATIOS
Empty Weight 8937 kg (lb)
Gross Weight
Loaded Weight 13,570 kg (29,915)
Maximum Take Off Weight 15,210 kg (33,530 lb)
Armaments 2,500 kg (4,410 lb)
FUSELAGE
Length 16.80 m (55 ft 1 in)
Height 4.99 m (16 ft 4 in)
WING
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Area (m ) 34 m² (366 ft²)
Span 9.31 m (30 ft 7 in)
PERFORMANCE
Maximum Speed
1,150 km/h (620 kn, 715 mph, Mach
0.94) at sea level
2,150 km/h (1,160 kn, 1,335 mph) at
high altitude
Endurance
Service Ceiling 17,600 m (57,740 ft)
Range 1,650 km (890 nmi, 1,025 mi)
Rate of climb 160 m/s (31,500 ft/min)
Wing loading 434.8 kg/m² (89.05 lb/ft²)
Thrust/weight 0.71
Takeoff roll 950 m (3,120 ft)
Landing roll 700 m (2,300 ft)
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Sukhoi Su-2
Manufacturer Soviet Union
First flight -
Introduced -
ENGINE
Model Lyulka AL-21F-3 afterburningturbofan
No. Of engines 2
Dry thrust 60 kN (13,000 lbf) each
Thrust with afterburner 89 kN (20,000 lbf) each
Fuel capacity 4,500 kg (9,900 lb) internal
ACCOMMODATION
Crew 1
WEIGHT RATIOS
Empty Weight 11,150 kg (24,600 lb)
Gross Weight -
Loaded Weight 16,000 kg (35,000 lb)
Maximum Take Off Weight 23,500 kg (52,000 lb)
Armaments 5,700 Kg (12450 lb)
FUSELAGE
Length 15.96 m (52.4 ft)
Height 5.28 m (17.3 ft)
WING
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Area 51.2 m (551 sq ft)
Span 10.95 m (35.9 ft)
PERFORMANCE
Maximum Speed
At altitude: Mach 2
(2,495 km/h/1,550 mph)
At sea level: Mach 1.2
(1,470 km/h/910 mph)
Endurance -
Service Ceiling 19,810 m (64,990 ft)
Range 2,900 km (1,800 mi)
Rate of climb 315 m/s (62,000 ft/min)
Wing loading 312 kg/m (64.0 lb/ft )
Thrust/weight 1.15
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Sukhoi PAK FA
Manufacturer Soviet Union
First flight -
Introduced -
ENGINE
Model
New unnamed engine by NPO Saturn and
FNPTS MMPP
No. Of engines 2
Dry thrust AL-41F1 of 147 kN
Thrust with afterburner 157 kN
Fuel capacity 10,300 kg (22,711 lb)
ACCOMMODATION
Crew 1
WEIGHT RATIOS
Empty Weight 18,500 kg (40,785 lb)
Gross Weight -
Loaded Weight 26,000 kg (57,320 lb)
Maximum Take Off Weight 37,000 kg (81,570 lb)
Armaments
FUSELAGE
Length 19.8 m (65.9 ft)
Height 6.05 m (19.8 ft)
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WING
Area 78.8 m (848.1 ft )
Span 14 m (46.6 ft)
PERFORMANCE
Maximum Speed2,100 – 2,500 km/h (Mach 2+) (1,300 –
1,560 mph) ; at 17,000 m (45,000 ft) altitude
Endurance -
Service Ceiling 20,000 m (65,616 ft)
Range -
Rate of climb 350 m/sec (68,900 ft/min)
Wing loading
330(normal) - 470(maximum) kg/m
(67(normal) - 96(maximum) lb/ft2)
Thrust/weight 1.19
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Rafale B
Manufacturer Dassault Aviation
First flight 4 July 1986
Introduced 4 December 2000
ENGINE
Model Snecma M88-2 turbofans
No. Of engines 2
Dry thrust 50.04 kN (11,250 lbf) each
Thrust with afterburner 75.62 kN (17,000 lbf) each
Fuel capacity -
ACCOMMODATION
Crew 2
WEIGHT RATIOS
Empty Weight 10,196 kg (22,431 lb)
Gross Weight
Loaded Weight 14,016 kg (30,900 lb)
Maximum Take Off Weight 24,500 kg (53,900 lb)
Armaments 13,350 Kg (29,370 lb)
FUSELAGE
Length 15.27 m (50.1 ft)
Height 5.34 m (17.5 ft)
WING
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Area 45.7 m² (492 ft²)
Span 10.80 m (35.4 ft)
Airfoil -
PERFORMANCE
Maximum Speed
High altitude: Mach 2 (2,390 km/h,
1,290 knots)
Low altitude: 1,390 km/h, 750 knots
Endurance
Service Ceiling 16,800 m (55,000 ft)
Combat Radius 1,852+ km
Range 3,700+ km (2,000+ nmi)
Ferry Range -
Rate of climb 304.8+ m/s (60,000+ ft/min)
Wing loading 306 kg/m² (62.8 lb/ft²)
Thrust/weight 1.10
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Mirage G8-02
Manufacturer Dassault Aviation
First flight 18 November, 1967
Introduced 1960
ENGINE
Model SNECMA Atar 9K50turbojets
No. Of engines 2
Dry thrust 70.1 kN (15,800 lbf) each
Thrust with afterburner -
Fuel capacity -
ACCOMMODATION
Crew 1
WEIGHT RATIOS
Empty Weight 14,740 kg (32,500 lb)
Gross Weight -
Loaded Weight -
Maximum Take Off Weight 19,340 Kg (42,548)
Armaments 7,350 Kg (16,170 lb)
FUSELAGE
Length 18.80 m (61 ft 8 in)
Height 5.35 m (17 ft 7 in)
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WING
Area -
Span
Extended: 15.40 m (50 ft 6 in)
Swept: 8.70 m (28 ft 7 in)
PERFORMANCE
Maximum Speed 2.2 Mach
Endurance -
Service Ceiling 18,500 m (60,700 ft)
Range 3,850 km (2,080 nm, 2,390 mi)
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MiG-31
Manufacturer Soviet Union
First flight -
Introduced -
ENGINE
Model Soloviev D-30F6 afterburning turbofans
No. Of engines 2
Dry thrust 93 kN (20,900 lbf) each
Thrust with afterburner 152 kN (34,172 lbf) each
Fuel capacity -
ACCOMMODATION
Crew 2
WEIGHT RATIOS
Empty Weight 21,820 kg (48,100 lb)
Gross Weight -
Loaded Weight 41,000 kg (90,400 lb)
Maximum Take Off Weight 46,200 kg (101,900 lb)
Armaments 15,600 Kg (34,320lb)
FUSELAGE
Length 22.69 m (74 ft 5 in)
Height 6.15 m (20 ft 2 in)
WING
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Area 61.6 m² (663 ft²)
Span 13.46 m (44 ft 2 in)
PERFORMANCE
Maximum Speed
High altitude: Mach 2.83 (3,000 km/h,
1,860 mph)
Low altitude: Mach 1.2 (1,500 km/h,
930 mph)
Endurance -
Service Ceiling 20,600 m (67,600 ft)
Range
Rate of climb 208 m/s (41,000 ft/min)
Wing loading 665 kg/m² (136 lb/ft²)
Thrust/weight 0.85
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MiG-15bis
Manufacturer Soviet Union
First flight -
Introduced -
ENGINE
Model Klimov VK-1turbojet
No. Of engines 1
Dry thrust -
Thrust with afterburner 26.5 kN (5,950 lbf)
Fuel capacity 1,400 L (364 US gal)
ACCOMMODATION
Crew 1
WEIGHT RATIOS
Empty Weight 3,580 kg (7,900 lb)
Gross Weight -
Loaded Weight 4,960 kg (10,935 lb)
Maximum Take Off Weight 6,105 kg (13,460 lb)
Armaments 1,650 Kg (3,630 lb)
FUSELAGE
Length 10.11 m (33 ft 2 in)
Height 3.70 m (12 ft 2 in)
WING
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Area 20.6 m² (221.74 ft²)
Span 10.08 m (33 ft 1 in)
PERFORMANCE
Maximum Speed 1,075 km/h (668 mph)
Endurance -
Service Ceiling 15,500 m (50,850 ft)
Range
1,200 km, 1,975 km with external tanks (745
mi / 1,225 mi)
Rate of climb 50 m/s (9,840 ft/min)
Wing loading 240.8 kg/m² (49.3 lb/ft²)
Thrust/weight 0.54
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Lockheed Martin F -22 Raptor
Manufacturer Lockheed Martin
First flight 2009
Introduced 2004
ENGINE
Model Pratt & Whitney F119-PW-100
No. Of engines 2
Dry thrust 23,500 lb (104 kN) each
Thrust with afterburner 35,000+ lb (156+ kN) each
Fuel capacity
18,000 lb (8,200 kg) internally, or 26,000 lb
(11,900 kg) with two external fuel tanks
ACCOMMODATION
Crew 1
WEIGHT RATIOS
Empty Weight 43,430 lb (19,700 kg)
Gross Weight
Loaded Weight 64,460 lb (29,300 kg)
Maximum Take Off Weight 83,500 lb (38,000 kg)
Armaments -
FUSELAGE
Length 62 ft 1 in (18.90 m)
Height 16 ft 8 in (5.08 m)
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WING
Area 840 ft² (78.04 m²)
Span 44 ft 6 in (13.56 m)
Airfoil NACA 64A05.92 root, NACA 64A?04.29 tip
PERFORMANCE
Maximum Speed
At altitude:Mach 2.25 (1,500 mph, 2,410
km/h)
Super cruise: Mach 1.82 (1,220 mph, 1,963
km/h)
Endurance -
Service Ceiling 65,000 ft (19,812 m)
Combat Radius 410 nmi (471 mi, 759 km)
Range
1,600 nmi (1,840 mi, 2,960 km) with 2
external fuel tanks
Ferry Range 2,000 mi (1,738 nmi, 3,219 km)
Rate of climb -
Wing loading 77 lb/ft² (375 kg/m²)
Thrust/weight 1.08 (1.26 with loaded weight & 50% fuel)
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Jaguar A
Manufacturer SEPECAT (Breguet/BAC)
First flight 8sep1968
Introduced 1973
ENGINE
Model Rolls-Royce/Turbomeca Adour Mk
102turbofans
No. Of engines 2
Dry thrust 22.75 kN (5,115 lbf) each
Thrust with afterburner 32.5 kN (7,305 lbf) each
Fuel capacity -
ACCOMMODATION
Crew 1
WEIGHT RATIOS
Empty Weight 7,000 kg (15,432 lb)
Gross Weight -
Loaded Weight 10,954 kg (24,149 lb)
Maximum Take Off Weight 15,700 kg (34,612 lb)
Armaments 11,200 Kg (24,640 lb)
FUSELAGE
Length 16.83 m (55 ft 2½ in)
Height 4.89 m (16 ft 0½ in)
WING
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Area 24.2 m² (220 ft²)
Span 8.69 m (28 ft 6 in)
Aspect Ratio 3.12:1
PERFORMANCE
Maximum Speed
Mach 1.6 (1,699 km/h, 917 knots, 1,056 mph)
at 11,000 m (36,000 ft)
Endurance -
Service Ceiling 14,000 m (45,900 ft)
Combat Radius
908 km (490 nmi, 564 mi) (lo-lo-lo, externalfuel)
Range -
Ferry Range 3,524 km (1,902 nmi, 2,190 mi)
Rate of climb -
Wing loading -
Thrust/weight -
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F-111 Aardvark
Manufacturer General Dynamics
First flight 21 December, 1964
Introduced 18 July, 1967
ENGINE
Model Pratt & Whitney TF30-P-100turbofans
No. Of engines 2
Dry thrust 17,900 lbf (79.6 kN) each
Thrust with afterburner 25,100 lbf (112 kN) each
Fuel capacity -
ACCOMMODATION
Crew 2 (pilot and weapons system operator)
WEIGHT RATIOS
Empty Weight 47,200 lb (21,400 kg)
Gross Weight -
Loaded Weight 82,800 lb (37,600 kg)
Maximum Take Off Weight 100,000 lb (45,300 kg)
Armaments 13,050 lb (5919.38 Kg)
FUSELAGE
Length 73 ft 6 in (22.4 m)
Height 17.13 ft (5.22 m)
WING
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Area
Spread: 657.4 ft² (61.07 m²)
Swept: 525 ft² (48.77 m²)
Span
Spread: 63 ft (19.2 m)
Swept: 32 ft (9.75 m)
Aspect Ratio spread: 7.56, swept: 1.95
Drag Area 9.36 ft² (0.87 m²)
Zero-Lift Drag Coefficient 0.0186
Airfoil NACA 64-210.68 root, NACA 64-209.80 tip
PERFORMANCE
Maximum Speed Mach 2.5 (1,650 mph, 2,655 km/h)
Endurance
Combat radius 1,330 mi (1,160 nmi, 2,140 km)
Service Ceiling 66,000 ft (20,100 m)
Ferry Range 4,200 mi (3,700 nmi, 6,760 km)
Rate of climb 25,890 ft/min (131.5 m/s)
Wing loading
Spread: 126.0 lb/ft² (615.2 kg/m²)
Swept: 158 lb/ft² (771 kg/m²)
Thrust/weight 0.61
Lift-to-drag ratio 15.8
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MiG-35
Manufacturer Soviet Union
First flight -
Introduced -
ENGINE
Model Klimov RD-33MKafterburningturbofans
No. Of engines 2
Dry thrust 5,400 kgf, 53.0 kN (11,900 lbf) each
Thrust with afterburner 9,000 kgf, 88.3 kN (19,800 lbf) each
Fuel capacity -
ACCOMMODATION
Crew 1 or 2
WEIGHT RATIOS
Empty Weight 11,000 kg (24,250 lb)
Gross Weight
Loaded Weight 17,500 kg (38,600 lb)
Maximum Take Off Weight 29,700 kg (65,500 lb)
Armaments 8,300 Kg (18,260lb)
FUSELAGE
Length 17.3 m (56 ft 9 in)
Height 4.7 m (15 ft 5 in)
WING
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Area 38 m (124 ft )
Span 12 m (39 ft 4 in)
PERFORMANCE
Maximum Speed
Mach 2.25 (2,400 km/h, 1,491 mph) at
altitude;[20]
1,450 km/h (901 mph) at low-level
Endurance -
Service Ceiling 17,500 m (57,400 ft)
Range 2,000 km (1,240 mi)
Rate of climb 330 m/s (65,000 ft/min)
Wing loading -
Thrust/weight 1.14
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Conclusion:
Thus the Fighter Aircraft, Class-1 is studied for it performance. The aerofoil has been
selected and the wing has been drawn. The various weight ratios are determined. The centre of
gravity the wing and the fuselage are determined. The center of gravity for the aircraft with the
engine mounted are also determined. And the various aircraft reference data also placed here.