The Application of Reliability Methods for Aircraft Design Project ...
Aircraft Design Project 1
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Transcript of Aircraft Design Project 1
1
1. Introduction
1.1 Overview:
Three major types of airplane designs are
1. Conceptual design
2. Preliminary design
3. Detailed design
1. Conceptual design:
It depends on what are the major factors for designing the aircraft.
(a) Power plant Location:
The Power plant must be located in the wings.
(b) Selection of Engine:
The engine should be selected according to the power required i.e.,
thrust required.
(c) Wing selection:
The selection of wing depends upon the selection of
(1) Low wing
(2) Mid wing
(3) High wing
- For a bomber the wing is mostly high wing configuration and
anhedral.
- Sweep may be required in order to reduce wave drag.
2. Preliminary design:
Preliminary is based upon number of factors like Loitering.
2
3. Detailed design:
In the detailed design considers each & every rivets, bolts, paints etc. In
this design the connection & allocations are made.
1.1 Bomber:
A bomber is a military aircraft designed to attack ground and sea targets,
by dropping bombs on them, or – in recent years – by launching cruise missiles
at them.
Strategic bombers are primarily designed for long-range bombing
missions against strategic targets such as supply bases, bridges, factories,
shipyards, and cities themselves, in order to damage an enemy's war effort.
Tactical bombing, aimed at enemy's military units and installations, is
typically assigned to smaller aircraft operating at shorter ranges, typically along
the troops on the ground or sea. This role is filled by various aircraft classes, as
different as light bombers, medium bombers, dive bombers, fighter-bombers,
ground-attack aircraft and multirole combat aircraft among others.
1.1.2 Origin of Bombers:
Bombers evolved at the same time as the fighter aircraft at the start of
World War I. The first use of an air-dropped bomb however, was carried out by
the Italians in their 1911 war for Libya.
Later several number of improvements were made.
3
1.3 Project requirement
1. To design a bomber aircraft
2. Range of 20,000 km with refueling support & must carry 75,000+ kg of
bombs & missiles (possibly nuclear warheads)
3. To operate at subsonic and transonic regimes
4. To operate at regional bases with low cost of operation & maintenance
5. The aircraft must also be capable of single pilot operation scenario.
6. Due to long range pilot work load must be reduced
7. The aircraft must be all weather, all terrain operation capable including
the airbase.
8. To take up a load factor +7.5g to -3.5g.
1.4. Preferred Configuration:
Figure 1.1 High wing Configuration with T tail
4
2. Comparative study of various bomber aircrafts
The first step in the design of aircraft is to collect data of existing aircraft of
similar purpose i.e., bomber. This step is vital in aircraft design as it gives the
designer an insight into the conventional trend in aircraft design.
The designer may, with the help of the data thus acquired, get an idea of the
basic factors that affect the aircraft’s performance viz. Weight, Cruise velocity,
Range, Wing area, Wingspan & Engine thrust. This database will also serve,
during the design process, as a guide for validation of the design parameters that
will be calculated, so that the designer does not deviate unduly from the
conventional design.
The performance data of various bomber aircraft with payload capacity
between 5000 and 56600 kg was collected from the appropriate resources.
5
2.1 Comparative configuration study of bomber airplanes:
Collection of Comparative Data -Dimension
S.No
Name of
the
aircraft
Payload
Capacity
(kg)
Length
(m)
Height
(m)
Wing
span
(m)2
No of
Power
Plant
Loaded
Weight
(kg)
Maximum Takeoff
Weight
(kg)
Empty
weight
(kg)
1 Mirage IIIE 5000 15 4.5 8.22 1 12200 13500 7050
2
Mirage
IVA 7264 23.49 5.4 11.85 2 31600 33475 14500
3 F-111F 14300 22.4 5.22 19.2 2 37600 45300 21400
4 F-111F Swept 14300 22.4 5.22 9.75 2 37600 45300 21400
5 Tu-22R 9000 41.6 10.13 23.17 2 85000 92000
6 Tu-85/1 18000 39.306 11.358 55.96 4 76000 107292 54711
7 YB-60 33000 52.1 18.4 62.8 8 73000 140000 69407
8 B-2A 23000 21 5.18 52.4 4 152200 170600 71700
9
Tu-
142M3 15000 49.5 12.12 51.1 4 170000 185000 90000
10 Tu-95MS 15000 46.2 12.12 50.1 4 171000 188000 90000
11 B-1B 56600 44.5 10.4 41.8 4 148000 216400 87100
12
B-1B
Swept 56600 44.5 10.4 24.1 4 148000 216400 87100
13 B-52H 31500 48.5 12.4 56.4 8 120000 220000 83250
14 Tu-160 40000 54.1 13.1 55.7 4 267600 275000 110000
15
Tu-160
Swept 40000 54.1 13.1 35.6 4 267600 275000 110000
Table 2.1 Collection of Comparative Data -Dimension
6
Collection of Comparative Data -Performance parameters
S.No
Name of
the
aircraft
Thrust
to
weight
ratio
Wing
loading
(N/m2)
Aspect
ratio Power Plant
1
Mirage
IIIE 0.50 3796.47 1.939 SNECMA Atar 09C turbojet
2
Mirage
IVA 0.60 5949.29 1.8
SNECMA Atar 9K-50[13] turbojets Dry thrust: 49.03 kN (11,023 lbf) each Thrust with afterburner: 70.61 kN (15,873 lbf) each
3 F-111F 0.61 6035.11 7.56
Pratt & Whitney TF30-P-100 turbofans Thrust with afterburner: 25,100 lbf (112 kN) each Dry thrust: 17,900 lbf (79.6 kN) each
4
F-111F
Swept 0.61 7563.51 1.95
Pratt & Whitney TF30-P-100 turbofans Thrust with afterburner: 25,100 lbf (112 kN) each Dry thrust: 17,900 lbf (79.6 kN) each
5 Tu-22R 0.38 5150.25 3.314
Dobrynin RD-7M-2 turbojets Dry thrust: rated 107.9 kN (24,250 lbf) each Thrust with afterburner: 161.9 kN (36,376 lbf) each
6 Tu-85/1 0.44 2717.37 11.45 Dobrynin VD-4K turbo-compound radial engines, 3,200 kW (4,300 hp) each
7 YB-60 0.44 1471.50 8.1 Pratt & Whitney J57-P-3 turbojets, (38 kN) each
8 B-2A 0.21 3227.49 5.74 General Electric F118-GE-100 non-afterburning turbofans, 17,300 lbf (77 kN) each
9 Tu-142M3 0.29 9112.02 8.394
Kuznetsov NK-12MV turboprops, 11,033 kW (14,795 shp) each
10 Tu-95MS 0.40 5944.86 8.097 Kuznetsov NK-12M turboprops, 11,000 kW (14,800 shp)[23] each
11 B-1B 0.38 8004.96 9.65
General Electric F101-GE-102 augmented turbofans Dry thrust: 14,600 lbf (64.9 kN) each Thrust with afterburner: 30,780 lbf (136.92 kN) each
12
B-1B
Swept 0.38 8004.96 9.65
General Electric F101-GE-102 augmented turbofans Dry thrust: 14,600 lbf (64.9 kN) each Thrust with afterburner: 30,780 lbf (136.92 kN) each
13 B-52H 0.31 5836.95 8.56 Pratt & Whitney TF33-P-3/103 turbofans, (76 kN) each
14 Tu-160 0.37 7269.21 7.757
Samara NK-321 turbofans Dry thrust: 137.3 kN (30,865 lbf) each Thrust with afterburner: 245 kN (55,115 lbf) each
15
Tu-160
Swept 0.37 7279.02 3.52
Samara NK-321 turbofans Dry thrust: 137.3 kN (30,865 lbf) each Thrust with afterburner: 245 kN (55,115 lbf) each
Table 2.2 Collection of Comparative Data -Performance parameters
7
Collection of Comparative Data -Performance parameters
S.No
Name of the
aircraft
Maximum
Speed
(m/s)
Range
(km)
Service
ceiling (m)
Rate of
Climb
(m/s)
Combat
Radius
(km)
Payload
Capacity
(kg)2
1 Mirage IIIE 652.8 2400 17000 83.3 1200 5000
2 Mirage IVA 650.0 4000 20000 43.13 1240 7264
3 F-111F 737.5 6760 20100 131.5 2140 14300
4 F-111F Swept 737.5 6760 20100 131.5 2140 14300
5 Tu-22R 419.4 4900 13300 2450 9000
6 Tu-85/1 177.2 12000 11700 17 5850 18000
7 YB-60 227.2 13000 16200 5.38 4700 33000
8 B-2A 270.0 11100 15200 5550 23000
9 Tu-142M3 256.9 12000 6500 15000
10 Tu-95MS 255.6 15000 13716 10 7500 15000
11 B-1B 372.2 11998 18000 5543 56600
12 B-1B Swept 372.2 11998 18000 5543 56600
13 B-52H 277.8 16232 15000 31.85 7210 31500
14 Tu-160 616.7 12300 15000 70 7300 40000
15
Tu-160
Swept 616.7 12300 15000 70 7300 40000
Table 2.3 Collection of Comparative Data -Performance parameters (Cont.)
8
Collection of Comparative Data -Performance parameters
S.No Name of the aircraft Airfoil
Span to
length ratio
Span to
height ratio Introduction
Retirement Remarks
1 Mirage IIIE 0.30 1.83 1961 In service Good
2 Mirage IVA 0.23 2.19 1959 2005 Good
3 F-111F
NACA 64-210.68 root, NACA 64-209.80 tip 0.23 3.68 1967 2010 Good
4 F-111F Swept
NACA 64-210.68 root, NACA 64-209.80 tip 0.23 1.87 1967 2010 Good
5 Tu-22R 0.24 2.29 1962 1990 Good
6 Tu-85/1 0.29 4.93 Prototype Good
7 YB-60 0.35 3.41 1954 In service Good
8 B-2A 0.25 10.12 1997 In service Good
9 Tu-142M3 0.24 4.22 1953 In service Awesome
10 Tu-95MS 0.26 4.13 1956 In service Awesome
11 B-1B NA69-190-1 0.23 4.02 1986 In service Good
12 B-1B Swept NA69-190-2 0.23 2.32 1986 In service Good
13 B-52H
NACA 63A219.3 mod root, NACA 65A209.5 tip 0.26 4.55 1961 In service Awesome
14 Tu-160 0.24 4.25 2005 In service Good
15 Tu-160 Swept 0.24 2.72 2005 In service Good
Table 2.4 Collection of Comparative Data -Performance parameters (Cont.)
9
2.2 Comparative graphs for determining optimum value:
S.No Name of the aircraft Wing loading (N/m2) Maximum Speed (m/s)
1 Mirage IIIE 3796.47 652.8
2 Mirage IVA 5949.29 650.0
3 F-111F 6035.11 737.5
4 F-111F Swept 7563.51 737.5
5 Tu-22R 5150.25 419.4
6 Tu-85/1 2717.37 177.2
7 YB-60 1471.50 227.2
8 B-2A 3227.49 270.0
9 Tu-142M3 9112.02 256.9
10 Tu-95MS 5944.86 255.6
11 B-1B 8004.96 372.2
12 B-1B Swept 8004.96 372.2
13 B-52H 5836.95 277.8
14 Tu-160 7269.21 616.7
15 Tu-160 Swept 7279.02 616.7
Table 2.5 Wing loading vs. Maximum Speed
Figure 2.1 Wing loading vs. Maximum Speed
0.00
1000.00
2000.00
3000.00
4000.00
5000.00
6000.00
7000.00
8000.00
9000.00
10000.00
0.0 100.0 200.0 300.0 400.0 500.0 600.0 700.0 800.0
Win
g Lo
adin
g (N
/sq
.m)
Maximum Speed (m/s)
Wing Loading (N/sq.m) Vs Maximum Speed (m/s)
Wing loading (N/m2)
10
S.No Name of the aircraft Span to length ratio Maximum Speed (m/s)
1 Mirage IIIE 0.30 652.8
2 Mirage IVA 0.23 650.0
3 F-111F 0.23 737.5
4 F-111F Swept 0.23 737.5
5 Tu-22R 0.00 419.4
6 Tu-85/1 0.29 177.2
7 YB-60 0.35 227.2
8 B-2A 0.25 270.0
9 Tu-142M3 0.24 256.9
10 Tu-95MS 0.26 255.6
11 B-1B 0.23 372.2
12 B-1B Swept 0.23 372.2
13 B-52H 0.26 277.8
14 Tu-160 0.24 616.7
15 Tu-160 Swept 0.24 616.7
Table 2.6 Span to length Ratio vs. Maximum Speed
Figure 2.2 Span to length Ratio vs. Maximum Speed
0.00
0.50
1.00
1.50
2.00
2.50
3.00
0.0 100.0 200.0 300.0 400.0 500.0 600.0 700.0 800.0
Span
to
len
gth
Rat
io
Maximum Speed (m/s)
Span to length ratio Vs Maximum Speed (m/s)
Span tolength ratio
11
S.No Name of the aircraft Aspect ratio Maximum Speed (m/s)
1 Mirage IIIE 1.94 652.8
2 Mirage IVA 1.80 650.0
3 F-111F 7.56 737.5
4 F-111F Swept 1.95 737.5
5 Tu-22R 3.31 419.4
6 Tu-85/1 11.45 177.2
7 YB-60 8.10 227.2
8 B-2A 5.74 270.0
9 Tu-142M3 8.39 256.9
10 Tu-95MS 8.10 255.6
11 B-1B 9.65 372.2
12 B-1B Swept 9.65 372.2
13 B-52H 8.56 277.8
14 Tu-160 7.76 616.7
15 Tu-160 Swept 3.52 616.7
Table 2.7 Aspect Ratio vs. Maximum Speed
Figure 2.3 Aspect Ratio vs. Maximum Speed
0
2
4
6
8
10
12
14
0.0 100.0 200.0 300.0 400.0 500.0 600.0 700.0 800.0
Asp
ect
Rat
io
Maximum Speed (m/s)
Aspect ratio Vs Maximum Speed (m/s)
Aspect ratio
12
S.No Name of the aircraft Wing Area (m2)
Maximum Speed (m/s)
1 Mirage IIIE 34.85 652.8
2 Mirage IVA 78 650.0
3 F-111F 61.07 737.5
4 F-111F Swept 48.77 737.5
5 Tu-22R 162 419.4
6 Tu-85/1 273.6 177.2
7 YB-60 486.7 227.2
8 B-2A 478 270.0
9 Tu-142M3 311.1 256.9
10 Tu-95MS 310 255.6
11 B-1B 180.2 372.2
12 B-1B Swept 181.2 372.2
13 B-52H 370 277.8
14 Tu-160 400 616.7
15 Tu-160 Swept 360 616.7
Table 2.8 Wing Area vs. Maximum Speed
Figure 2.4 Wing area vs. Maximum Speed
0
100
200
300
400
500
600
0.0 100.0 200.0 300.0 400.0 500.0 600.0 700.0 800.0
Win
g A
rea
(sq
.m)
Maximum Speed (m/s)
Wing Area (sq.m) Vs Maximum Speed (m/s)
Wing Area (m2)
13
S.No Name of the aircraft Combat Radius (km) Maximum Speed (m/s)
1 Mirage IIIE 1200 652.8
2 Mirage IVA 1240 650.0
3 F-111F 2140 737.5
4 F-111F Swept 2140 737.5
5 Tu-22R 2450 419.4
6 Tu-85/1 5850 177.2
7 YB-60 4700 227.2
8 B-2A 5550 270.0
9 Tu-142M3 6500 256.9
10 Tu-95MS 7500 255.6
11 B-1B 5543 372.2
12 B-1B Swept 5543 372.2
13 B-52H 7210 277.8
14 Tu-160 7300 616.7
15 Tu-160 Swept 7300 616.7
Table 2.9 Span to length Ratio vs. Maximum Speed
Figure 2.5 Combat radius vs. Maximum Speed
0
1000
2000
3000
4000
5000
6000
7000
8000
0.0 100.0 200.0 300.0 400.0 500.0 600.0 700.0 800.0
Co
mb
at r
adiu
s (k
m)
Maximum Speed (m/s)
Combat Radius (km) Vs Maximum Speed (m/s)
Combat Radius (km)
14
S.No Name of the aircraft Payload Capacity (kg)2
Maximum Speed (m/s)
1 Mirage IIIE 5000 652.8
2 Mirage IVA 7264 650.0
3 F-111F 14300 737.5
4 F-111F Swept 14300 737.5
5 Tu-22R 9000 419.4
6 Tu-85/1 18000 177.2
7 YB-60 33000 227.2
8 B-2A 23000 270.0
9 Tu-142M3 15000 256.9
10 Tu-95MS 15000 255.6
11 B-1B 56600 372.2
12 B-1B Swept 56600 372.2
13 B-52H 31500 277.8
14 Tu-160 40000 616.7
15 Tu-160 Swept 40000 616.7
Table 2.10 Payload capacity vs. Maximum Speed
Figure 2.6 Payload Capacity vs. Maximum Speed
0
10000
20000
30000
40000
50000
60000
0.0 100.0 200.0 300.0 400.0 500.0 600.0 700.0 800.0
Pay
load
Cap
acit
y (k
g)
Maximum Speed (m/s)
Payload Capacity (kg) Vs Maximum Speed (m/s)
Payload Capacity (kg)
15
S.No Name of the aircraft
Thrust to weight
ratio
Maximum Speed
(m/s)
1 Mirage IIIE 0.5 652.8
2 Mirage IVA 0.6 650.0
3 F-111F 0.61 737.5
4 F-111F Swept 0.61 737.5
5 Tu-22R 0.38 419.4
6 Tu-85/1 0.44 177.2
7 YB-60 0.44 227.2
8 B-2A 0.205 270.0
9 Tu-142M3 0.29 256.9
10 Tu-95MS 0.4 255.6
11 B-1B 0.38 372.2
12 B-1B Swept 0.38 372.2
13 B-52H 0.31 277.8
14 Tu-160 0.37 616.7
15 Tu-160 Swept 0.37 616.7
Table 2.11 Thrust to weight Ratio vs. Maximum Speed
Figure 2.7 Thrust to weight Ratio vs. Maximum Speed
0.00
0.10
0.20
0.30
0.40
0.50
0.60
0.70
0.0 200.0 400.0 600.0 800.0
Thru
st t
o w
eig
ht
rati
o
Maximum Speed (m/s)
Thrust to Weight Ratio Vs Maximum Speed (m/s)
Thrust to weight ratio
16
S.No Name of the aircraft Span to height ratio Maximum Speed (m/s)
1 Mirage IIIE 1.826666667 652.8
2 Mirage IVA 2.194444444 650.0
3 F-111F 3.67816092 737.5
4 F-111F Swept 1.867816092 737.5
5 Tu-22R 2.287265548 419.4
6 Tu-85/1 4.926923754 177.2
7 YB-60 3.413043478 227.2
8 B-2A 10.11583012 270.0
9 Tu-142M3 4.216171617 256.9
10 Tu-95MS 4.133663366 255.6
11 B-1B 4.019230769 372.2
12 B-1B Swept 2.317307692 372.2
13 B-52H 4.548387097 277.8
14 Tu-160 4.251908397 616.7
15 Tu-160 Swept 2.717557252 616.7
Table 2.12 Span to height Ratio vs. Maximum Speed
Figure 2.8 Span to height Ratio vs. Maximum Speed
0.00
2.00
4.00
6.00
8.00
10.00
12.00
0.0 100.0 200.0 300.0 400.0 500.0 600.0 700.0 800.0
Span
to
He
igh
t R
atio
Maximum Speed (m/s)
Span to Height Ratio Vs Maximum Speed (m/s)
Span to height ratio
17
S.No Name of the aircraft
Maximum Takeoff
Weight (kg) Maximum Speed (m/s)
1 Mirage IIIE 13500 652.8
2 Mirage IVA 33475 650.0
3 F-111F 45300 737.5
4 F-111F Swept 45300 737.5
5 Tu-22R 92000 419.4
6 Tu-85/1 107292 177.2
7 YB-60 140000 227.2
8 B-2A 170600 270.0
9 Tu-142M3 185000 256.9
10 Tu-95MS 188000 255.6
11 B-1B 216400 372.2
12 B-1B Swept 216400 372.2
13 B-52H 220000 277.8
14 Tu-160 275000 616.7
15 Tu-160 Swept 275000 616.7
Table 2.13 Maximum takeoff weight vs. Maximum Speed
Figure 2.9 Maximum takeoff weight vs. Maximum Speed
0
50000
100000
150000
200000
250000
300000
0.0 200.0 400.0 600.0 800.0
Max
imu
m T
ake
off
We
igh
t (k
g)
Maximum Speed (m/s)
Maximum Takeoff Weight (kg) Vs Maximum Speed (m/s)
Maximum TakeoffWeight (kg)
18
2.2 Parameter Selection:
From Comparison (Assumed and extrapolated values from graph)
Maximum takeoff weight (kg) 500000
Thrust to weight ratio 0.28
Aspect ratio 8.4
Wing loading (N/sq.m) 7848
Span to height ratio 5
Span to length ratio 1.5
Combat radius (km) 5000
Pay load capacity (kg) 75000
Maximum speed (kmph) 1000
Service ceiling (m) 15000
Maximum speed (m/s) 277.777
Table 2.14 Parameter Selection
19
3. Rough Weight Estimate
Optimal values of mass fraction for bombers
Parameter Range of values Notation
Empty Mass Ratio 0.37-0.32 ME/MTO
Total Fuel Mass Ratio 0.40-0.62 MF/MTO
Payload Ratio 0.14-0.19 MPay/MTO
Wing Loading 4385-7848 N/m2
Wo/S
Thrust to Weight Ratio 0.26-0.40 T/Wo
Table 3.1 Mass Fraction Parameters for bomber
3.1 General rough weight estimate:
Figure 3.1 Payload mass Fraction vs. Maximum Takeoff weight
0
0.05
0.1
0.15
0.2
0.25
0.3
0.35
0.4
0 50000 100000 150000 200000 250000 300000
MPa
y/M
To
Maximum Takeoff weight (kg)
MPay/MTo Vs Maximum Takeoff weight (kg)
MPay/MTo
20
= 75000 kg
Figure3.2 Empty mass Fraction vs. Maximum Takeoff weight
0
0.1
0.2
0.3
0.4
0.5
0.6
0 50000 100000 150000 200000 250000 300000
ME/
MTo
Maximum Takeoff weight (kg)
ME/MTo Vs Maximum Takeoff weight (kg)
ME/Mto
21
Final Values from rough weight estimate:
Mass Fraction
Payload 0.15
Fuel 0.45
Structure 0.32
Power plant 0.07
Fixed equipments 0.01
Total 1.00
Table 3.2 Values from rough weight estimate
22
4. Redefined Mass Estimation
4.1 Mission profile analysis
Profile 1: Strategic bombing mission
Analysis of Mission Profile:
Warmup and takeoff
Climb or descend
Landing
(
)
(
)
2’
3’ 5’ 4’
6’ 7’
8’
9’ 10’ 0 1
R
3 2
h
10000 km
1000 km
1000 km
9000 km
1/2 hr
Figure 4.1 Mission profile for Strategic bombing
23
Figure 4.2 Empty mass fraction vs takeoff mass -- taken from "Aircraft Design: A Conceptual Approach" by Daniel
P.Raymer
Analysis of Mission Profile
TSFC values for Bomber
Cruise Loiter
0.5 0.4
Table 4.1 TSFC values in lb/lbf-hr for bomber
Where A & c are constants from the historic data for bomber
c = -0.07; A = 0.93 (taken from Aircraft Design: A Conceptual Approach by
Daniel P.Raymer)
24
(
( ))
(
( ))
4.2: Before Refueling
Part 1: Before refueling:
Warm-up and Take-off: (0-1)
Climb: (1-2)
Cruise at 60% of maximum speed: (2-3)
For analysis (L/D) optimal = 17
Thrust Specific fuel Consumption C = 0.0001389 (kg / N-s)
0 1
R
3 2
h
10000 km
Figure 4.3 Mission profile before refueling
25
Range R2-3 = 10000 km
⁄
Descend: (3-R)
Total Mass fraction for first part of mission profile:
Fuel Mass fraction for first half of mission profile:
(
)
Thus the range of 10000 km can be interpreted as a combat radius of 5000 km.
4.3. Refueling:
Operation:
The tanker aircraft flies straight and level and extends the hose/drogue
which is allowed to trail out behind and below the tanker under normal
aerodynamic forces. The pilot of the receiver aircraft extends his probe (if
required) and uses normal flight controls to fly the refueling probe directly into
the basket. This requires a closure rate of approximately two knots (walking
26
speed) in order to establish solid probe/drogue couple and pushing the hose
several feet into the HDU. Too little closure will cause an incomplete
connection and no fuel flow (or occasionally leaking fuel). Too much closure is
dangerous because it can trigger a strong transverse oscillation in the hose,
severing the probe tip. Another significant danger is that the drogue may hit the
recipient aircraft and damage it—instances have occurred in which the drogue
has shattered the canopy of a fighter aircraft, causing great danger to its pilot.
Figure 4.4 A Tu-95MS simulating aerial refueling with an Ilyushin Il -78
The optimal approach is from behind and below (not level with) the
drogue. Because the drogue is relatively light (typically soft canvas webbing)
and subject to aerodynamic forces, it can be pushed around by the bow wave of
approaching aircraft, exacerbating engagement even in smooth air. After initial
contact, the hose and drogue is pushed forward by the receiver a certain distance
(typically, a few feet), and the hose is reeled slowly back onto its drum in the
HDU. This opens the tanker's main refueling valve allowing fuel to flow to the
drogue under the appropriate pressure (assuming the tanker crew has energized
the pump). Tension on the hose is aerodynamically balanced by a motor in the
HDU so that as the receiver aircraft moves fore and aft, the hose retracts and
extends, thus preventing bends in the hose that would cause undue side loads on
the probe. Fuel flow is typically indicated by illumination of a green light near
the HDU. If the hose is pushed in too far or not far enough, a cutoff switch will
inhibit fuel flow, which is typically accompanied by amber light.
Disengagement is commanded by the tanker pilot with a red light.
27
4.4. After Refueling:
Part 2: After refueling:
Cruise at 60% of maximum speed: (R-2)
Descend: (2’-3’)
Cruise: (3’-4’)
Bombing (4’)
Climb: (4’-5’)
R 1000 km
h
2’
3’ 5’ 4’
6’ 7’
8’
9’ 10’
1000 km 9000 km
Figure 4.5 After refueling mission profile
28
Cruise: (5’-6’)
Loiter: (6’-7’)
Loiter time = ½ hr
Descend: (7’-8’)
Landing: (8’-9’)
Total Mass fraction for second part of mission profile:
29
Total fuel mass fraction after refueling:
(
)
(
)
(
)
(
)
Hence (
)
is taken since the value turns out to be (
)
30
Replacing as X in excel to solve the implicit function
X ranges from 500000 to 520000 since initial mass estimate is 500000 kg
Thus mass of the aircraft is 508034.68 kg
X f(x)
500000 -9458.077
510000 2308.6163
520000 14026.793
X f(x)
501000 -8279.156
502000 -7100.742
503000 -5922.833
504000 -4745.425
505000 -3568.517
506000 -2392.106
507000 -1216.19
508000 -40.76524
509000 1134.1693
X f(x)
508000 -40.76524
508100 76.750211
X f(x)
508000 -40.76524
508010 -29.01347
508020 -17.26176
508030 -5.51009
508040 6.2415284
X f(x)
508030 -5.51009
508031 -4.334926
508032 -3.159763
508033 -1.984599
508034 -0.809437
508035 0.3657252
X f(x)
508034.6 -0.10434
508034.7 0.0131767
X f(x)
508034.67 -0.022078
508034.68 -0.010327
508034.69 0.001425
508034.7 0.0131767
508034.71 0.0249283
31
Take-off Weight of the aircraft:
4.5.Thrust Estimation
T=0.28×4983820.113
=1395469.632N
T=1395.469 KN
32
⁄
TA = 1395.469 KN
T =
T=348.86725 KN/Engine
33
5. Power Plant Selection
5.1. Comparative data of engines
Thrust Estimation:
(
(
( ⁄ )
(
)
(
)
)
)
For the chosen parameters:
T=0.28×4983820.113
=1395469.632N
T=1395.469 KN
⁄
TA = 1395.469 KN
T =
T=348.86725 KN/Engine
34
The performance data of various turbofan engines with thrust of range 330 kN
to 500 kN were collected from the following resources
www.jet-engine.net
www.wikipedia.org
S.No Name of the Engine Manufacturer Type
Length (m)
Diameter (m)
Wet weight (kg)
Dry Weight (kg)
1 Trent-900 Rolls Royce Turbofan 3 Shaft 4.55 2.94
6271
2 GP-7000 Engine Alliance Turbofan 2 Shaft 4.74 3.16 6800 6712
3 GE 90-76B GE Turbofan 2 Shaft 4.90 3.40 7540 7074
4 GE 90-92B GE Turbofan 2 Shaft 4.90 3.40 7648 7074
5 GP 7270 Engine Alliance Turbofan 2 Shaft 4.75 3.15 6800 6712
6 GE90-110B1 GE Turbofan 2 Shaft 4.90 3.40 8253 7550
7 GP 7277 Engine Alliance Turbofan 2 Shaft 4.75 3.15 6482 6033
8 PW 40477 Pratt & Whitney Turbofan 2 Shaft 4.87 3.01 6986 6598
9 GE 90-85B GE Turbofan 2 Shaft 4.90 3.40 7474 7074
10 GE 90-94B GE Turbofan 2 Shaft 4.90 3.40 8253 7550
11 GE 90-90B GE Turbofan 2 Shaft 4.90 3.40 7548 7074
12 GE 90-115B GE Turbofan 2 Shaft 4.90 3.40 8283 7550
Table 5.1. Collection of Engine Comparative Data
S.No Name of the Engine
Maximum Thrust (kN)
Overall Pressure Ratio
Thrust to Weight Ratio
Fan Diameter (m)
1 Trent-900 360 38 5.15
2 GP-7000 363 43.9 4.73 2.95
3 GE 90-76B 340 41.9 5.4 3.12
4 GE 90-92B 409 41.8 5.7 3.12
5 GP 7270 311 43.9 4.73 2.95
6 GE90-110B1 489 42.5 6.2 3.12
7 GP 7277 343 43 5.2 2.95
8 PW 40477 343 40 2.84
9 GE 90-85B 377 42 3.12
10 GE 90-94B 417 42 5.6 3.12
11 GE 90-90B 400 42 5.4 3.12
12 GE 90-115B 512 42 6.3 3.12
Table 5.2. Collection of Engine Comparative Data (Cont.)
35
5.2. Engine Selection:
From this we select Engine Alliance GP 7000
Specification of Engine
Name of the Engine GP-7000
Manufacturer Engine Alliance
Type Turbofan 2 Shaft
Length (m) 4.74
Diameter (m) 3.16
Wet weight (kg) 6800
Dry Weight (kg) 6712
Maximum Thrust (kN) 363
Overall Pressure Ratio 43.9
Thrust to Weight Ratio 4.73
Fan Diameter (m) 2.95
Table 5.3. Selected Engine Datas
5.3. Redefined Thrust to weight ratio:
Closer to initial value assumed value of 0.28
TSFC ≈ 0.8
TSFC = 0.7913 N/N - hr
= 0.7913
= 0.080662 kg/N- hr
TSFC = 0.02240627 kg/N -s
Service ceiling evaluation:
By taking service ceiling as h=15 km
√
36
√
√
Number of Engines = 4
37
6. Airfoil selection and Wing Geometry estimates
6.1. Main Parameter Selection:
Wing Loading:
6.2 Fuel volume consideration:
ρF can vary from 600 kg/m3 to 800 kg/m
3.
For ρF = 800 kg/m3
38
Volume of fuel accommodated in wing:
(
((
)
))
(
) (
(
))
(
)
Selecting NACA 653-418 airfoil of fineness ratio (t/c ratio) as 0.18
(
)
( )
( )
39
6.3 Takeoff Analysis:
Figure 6.1 Runway length survey for military installations
SR = 2000 m for around 68% of airbase in the world.
Assuming take off at 60% of runway length and accelerating at 20% the
gravitational attraction, where vi is initial velocity during takeoff.
40
(
)
(
)
Where can be denoted as t also since we use MAC to obtain the thickness
value
Thickness based Reynolds Number:
41
Figure 6.2 Cl vs Angle of attack curve for NACA653418 at angle of attack 0.5 deg
42
Symbol Re x/c y/c Angle of
attack
CL max
5.9×106 0.266 -0.052 18° 1.42
8.9×106 0.267 -0.047 18° 1.51
Table 6.1 Airfoil data at various Re.
At Re = 6.3×106 by interpolating we get
Location of aerodynamic center x/c = 0.2668
y/c = -0.0491
CL max = 1.48
α = 18°
6.4 Flap selection:
Flap Chosen is Triple slotted flap
Wing setting angle or incidence angle iw= 3 degree
Required Flap Deflection = 60°
Change in CL due to flap deflection:
43
Figure 6.3 Drag polar curve forNACA653418 at angle of attack 0.5 deg
44
6.5. Wing geometry:
Sweep Analysis:
For airfoil NACA 653- 418
At x/c = 0.46
(
)
Figure 6.4 Variation of local velocity with the free stream velocity
45
√
√
√
√
Hence by comparing v and aMSL and aalt it is clear that the shock wave is
formed.
In order to avoid this unwanted phenomenon we need to sweep the wing.
Critical Mach number:
(
)
If the maximum velocity reached on the upper surface is equal to the lowest
possible value of speed of sound then the velocity V∞ will be critical velocity
which corresponds to critical Mach number
Or simply M x/c= 0.46 =1
(
)
(
)
46
√
Critical Mach number for the airfoil:
Figure 6.5 Swept back wing
By using a trapezoidal and sweepback we may get
47
For 1000 km/hr or 277.77 m/s speed we get the sweep to avoid shock on upper
surface as
Where is due to taper property
(
⁄
)
Mean Aerodynamic Chord (MAC) ( ) for swept back wing:
(
)
Span wise location of MAC
On simplifying we get :
(
)
48
Where λ is taper ratio: λ can vary from 0 to 1.
By evaluating the above four equations we get :
λ cr (m) ct (m) (rad) (deg)
1 8.695 8.695 0.000 0.000
0.9 9.153 8.237 0.025 1.435
0.8 9.661 7.729 0.053 3.028
0.7 10.229 7.161 0.084 4.802
0.6 10.869 6.521 0.118 6.786
0.5 11.593 5.797 0.157 9.015
0.4 12.421 4.969 0.201 11.529
0.3 13.377 4.013 0.251 14.372
0.2 14.492 2.898 0.307 17.595
0.1 15.809 1.581 0.371 21.251
0 17.390 0.000 0.443 25.384
Table 6.2 Angle of taper for various taper ratios
Taking the value
λ = 0.5
cr = 11.593 m
ct = 5.797 m
Span wise location of MAC:
Hence:
49
Figure 6.6 Effect of aspect ratio on lift curve slope
√
a= 0.1213507 /degree for a’ = 0.15/ degree
50
7. Landing gear design
7.1. Tyre selection:
7.1.1. Load Distribution:
Typical load of aircraft while landing
Possibility of aborting mission would lead to
And during static condition
Typically main landing gear takes around 90 % of load and the Nose landing
gear takes around 10% of total load.
Load taken by wheels in nose landing gear = 0.1× 4983820.113
= 498.382 kN
Load taken by wheels in main landing gear = 0.9× 4983820.113
= 4485.438 kN
Nose Landing
Gear
Main Landing
Gear
Number of wheels 4 20
Total load supported (kN) 498.382 4485.438
Load taken by each
wheel (kN) 124.5955 224.2719
Tyre Pressure (psi) 200 200
Tyre Pressure (bar) 14.28 14.28
Table 7.1 Load Distribution
The load is taken by the tyre due to internal pressure,
51
Figure 7.1 Typical tyre pressures - taken from Aircraft design: Conceptual Approach by Daniel P. Raymer
7.1.2.Tyre Selection For Nose wheel:
Wheel diameter = AWwB (A=1.63, B=0.315)
dw =1.63 (27876.30)0.315
dw = 40.96 in ( 1.0403 m)
Wheel width = AWwB (A=0.1043, B=0.480)
ww =0.1043 (27876.30)0.480
ww = 14.190 in (0.3604m)
Figure 7.2 Emprical relations and constants for tyre selection - taken from Aircraft design: Conceptual Approach by
Daniel P. Raymer
52
Contact area:
Figure 7.3 Tyre contact area
From figure the contact area will be neither rectangular nor elliptic but a
combination of both.
√ (
)
(
√
)
(
√ )
Hence Rt rolling radius for nose landing gear assembly has reduced by 6.667%
of the wheel radius
53
7.1.3. Tyre Selection for Main landing gear:
Wheel diameter = AWwB (A=1.63, B=0.315)
dw =1.63 (50357.2)0.315
dw = 49.36 in ( 1.253 m)
Wheel width = AWwB (A=0.1043, B=0.480)
ww =0.1043 (50357.2)0.480
ww = 18.848 in (0.4787m)
Contact area:
√ (
)
(
√
)
(
√ )
7.2. Runway Loading:
Runway loading estimates for both Main and nose landing wheel:
54
For a rigid runway:
Nominal working stress on Concrete pavement: 400 psi or 2.75 MN/m2
Concrete Elastic modulus E= 27.5 GPa
55
8. Dimensional estimates
8.1. Basic Dimensions:
Span to height ratio
Span to length ratio:
Where length is the length of the fuselage
Total length m
56
8.2. Configuration of tail:
Figure 8.1 T tail configuration
8.2.1. Horizontal stabilizer:
Airfoil used: NACA 0012
Horizontal stabilizer sizing: 15% of wing area
A.Rh =4.5
57
Sweep analysis for the horizontal tail:
Tail will not be affected by downwash since we use T tail
Horizontal tail geometry:
Sweep Analysis:
For airfoil NACA 0012
Figure 8.2 Local velocity vs free stream velocity for NACA 0012 airfoil
58
Figure 8.3 CL vs angle of attack curve for NACA 0012
59
Figure 8.4 Drag polar curve for NACA 0012
60
At x/c = 0.125
(
)
M x/c= 0.125 =1
(
)
√
Critical Mach number for the airfoil:
By using a trapezoidal and sweepback we may get
61
For 1000 km/hr or 277.77 m/s speed we get the sweep to avoid shock on upper
surface as
Where is due to taper property
(
⁄ )
For horizontal tail:
By evaluating for λ which varies from 0 to 1 we get:
λh crh cth h (rad) h
1 4.60 4.60 0.00 0.00
0.9 4.84 4.36 0.01 0.76
0.8 5.11 4.09 0.03 1.60
0.7 5.41 3.79 0.04 2.55
0.6 5.75 3.45 0.06 3.60
0.5 6.13 3.07 0.08 4.80
0.4 6.57 2.63 0.11 6.16
0.3 7.08 2.12 0.13 7.72
0.2 7.67 1.53 0.17 9.53
0.1 8.37 0.84 0.20 11.64
0 9.20 0.00 0.25 14.13
Table 8.1 Variation of taper angle of horizontal tail for various taper ratio
Taking λh as 0.4
crh = 6.5729 m
cth = 5.797 m
Span wise location of MAC:
62
Hence:
8.2.2. Vertical Stabilizer Geometry:
Vertical stabilizer sizing: 9% of wing area
Airfoil used: NACA 0012
A.Rv =0.9
63
λv crv ctv v (rad) v
1 7.981454716 7.981454716 0 0
0.9 8.40153128 7.561378152 0.02299242 1.317407458
0.8 8.868283017 7.094626414 0.048510003 2.779500429
0.7 9.389946724 6.572962707 0.076953376 4.40923372
0.6 9.976818394 5.986091037 0.108800855 6.23401364
0.5 10.64193962 5.32096981 0.144623174 8.286541884
0.4 11.40207817 4.560831266 0.185098527 10.60567718
0.3 12.2791611 3.68374833 0.231024637 13.23712703
0.2 13.30242453 2.660484905 0.283320456 16.23354517
0.1 14.51173585 1.451173585 0.343001791 19.6531346
0 15.96290943 0 0.411098972 23.55493075
Table 8.2 Variation of taper angle of vertical tail for various taper ratio
Taking λt as 0.7
crt = 9.389 m
ctt = 6.5729 m = chr
Span wise location of MAC:
Hence:
64
9. Preparation of Layout
Configuration: Anhedral high wing, T-Tail configuration.
Nose radius rn= 3m
9.1. Wing Location and CG Estimation:
Reference is taken from nose:
Where X is the location of wing root L.E. from the nose fuselage and Xfinal is the
location of cg from L.E at root.
Xfinal = 0.35 (Xcr -Xct)
Xfinal =11.59 m
Substituting the values from condition 1 in the above equation:
Fence the wing root L.E. has to be fixed at 14.645 m from nose.
That is from nose cg lies at 14.645 +11.69 = 26.223 m
65
Figure 9.1 Wing Details for cg estimates
66
Condition 1 Full Payload and Full Fuel
Fuselage alone analysis
S.No Components
Distance from reference line (m) Mass (kg) Weight (N) Moment (Nm)
1 Crew 3.043 270 2648.7 8059.9941
2 Nose landing gear 6.086 3600 35316 214933.176
3 Payload bay 1 12.172 37500 367875 4477774.5
4 Fixed equipments 18.641 1149.19 11273.5539 210150.3182
5 Excess mass 22.641 21600 211896 4797537.336
6 Fuselage mass 22.641 43200 423792 9595074.672
7 Fuel in fuselage 22.641 95280.42 934700.9202 21162563.53
8
Main landing gear assembly 1 22.641 10800 105948 2398768.668
9
Main landing gear assembly 2 30.3211 7200 70632 2141639.935
10 Payload bay 2 30.321 37500 367875 11154337.88
11 Horizontal stabilizer 45.367 6400 62784 2848321.728
12 Vertical Stabilizer 48.501 3200 31392 1522543.392
Total 267699.61 2626133.174 60531705.13
Cg from Nose 23.04974695
Wing alone analysis
S.No Components
Distance from reference line (m) Mass (kg) Weight (N) Moment (Nm)
1 Wing structure 16.35 64000 627840 10265184
2 Wing fuel 1 11.85 74567.285 731505.0659 8668335.03
3 Wing fuel 2 18.428 74567.285 731505.0659 13480175.35
4 Power plant 1 9.57 13600 133416 1276791.12
5 Power plant 2 15 13600 133416 2001240
Total 240334.57 2357682.132 35691725.5
Cg from L.E 15.13848072
Entire aircraft
cg from L.E. Root 11.60343233 Grand total 508034.18 4983815.306 96223430.63
Table 9.1 Cg estimate for fully loaded condition
Similarly for different cases i.e. conditions of loading must be evaluated by fixing the wing at
the location x. For this case 1 the cg lies at 27.085 m
67
9.2. Three views of Aircraft:
Figure 9.2 Top View
68
Figure 9.3 Front View
69
Figure 9.4 S ide view
70
Condition 2 Full Payload and Reserve Fuel
Fuselage alone analysis
S.No Components
Distance from reference line (m) Mass (kg) Weight (N) Moment (Nm)
1 Crew 3.043 270 2648.7 8059.9941
2 Nose landing gear 6.086 3600 35316 214933.176
3 Payload bay 1 12.172 0 0 0
4 Fixed equipments 18.641 1149.19 11273.5539 210150.3182
5 Excess mass 22.641 21600 211896 4797537.336
6 Fuselage mass 22.641 43200 423792 9595074.672
7 Fuel in fuselage 22.641 19056.084 186940.184 4232512.707
8
Main landing gear assembly 1 22.641 10800 105948 2398768.668
9
Main landing gear assembly 2 30.3211 7200 70632 2141639.935
10 Payload bay 2 30.321 0 0 0
11 Horizontal stabilizer 45.367 6400 62784 2848321.728
12 Vertical Stabilizer 48.501 3200 31392 1522543.392
Total 116475.274 1142622.438 27969541.93
Cg from Nose 24.47837623
Wing alone analysis
S.No Components
Distance from reference line (m) Mass (kg) Weight (N) Moment (Nm)
1 Wing structure 16.35 64000 627840 10265184
2 Wing fuel 1 11.85 14913.457 146301.0132 1733667.006
3 Wing fuel 2 18.428 14913.457 146301.0132 2696035.071
4 Power plant 1 9.57 13600 133416 1276791.12
5 Power plant 2 15 13600 133416 2001240
Total 121026.914 1187274.026 17972917.2
Cg from L.E 15.13796882
Entire aircraft cg from L.E. Root 12.54876323 Grand total 237502.188 2329896.464 45942459.12
Table 9.2 Cg estimate for full payload and reserve fuel
71
Condition 3 Half Payload and Full Fuel
Fuselage alone analysis
S.No Components
Distance from reference line (m) Mass (kg) Weight (N) Moment (Nm)
1 Crew 3.043 270 2648.7 8059.9941
2 Nose landing gear 6.086 3600 35316 214933.176
3 Payload bay 1 12.172 18750 183937.5 2238887.25
4 Fixed equipments 18.641 1149.19 11273.5539 210150.3182
5 Excess mass 22.641 21600 211896 4797537.336
6 Fuselage mass 22.641 43200 423792 9595074.672
7 Fuel in fuselage 22.641 95280.42 934700.9202 21162563.53
8
Main landing gear assembly 1 22.641 10800 105948 2398768.668
9
Main landing gear assembly 2 30.3211 7200 70632 2141639.935
10 Payload bay 2 30.321 18750 183937.5 5577168.938
11 Horizontal stabilizer 45.367 6400 62784 2848321.728
12 Vertical Stabilizer 48.501 3200 31392 1522543.392
Total 230199.61 2258258.174 52715648.94
Cg from Nose 23.34349967
Wing alone analysis
S.No Components
Distance from reference line (m) Mass (kg) Weight (N) Moment (Nm)
1 Wing structure 16.35 64000 627840 10265184
2 Wing fuel 1 11.85 74567.285 731505.0659 8668335.03
3 Wing fuel 2 18.428 74567.285 731505.0659 13480175.35
4 Power plant 1 9.57 13600 133416 1276791.12
5 Power plant 2 15 13600 133416 2001240
Total 240334.57 2357682.132 35691725.5
Cg from L.E 15.13848072
Entire aircraft
cg from L.E. Root 12.00007718 Grand total 470534.18 4615940.306 88407374.45
Table 9.3 Cg estimate for half payload and full fuel
72
Condition 4 Half Payload and Reserve Fuel
Fuselage alone analysis
S.No Components
Distance from reference line (m) Mass (kg) Weight (N) Moment (Nm)
1 Crew 3.043 270 2648.7 8059.9941
2 Nose landing gear 6.086 3600 35316 214933.176
3 Payload bay 1 12.172 18750 183937.5 2238887.25
4 Fixed equipments 18.641 1149.19 11273.5539 210150.3182
5 Excess mass 22.641 21600 211896 4797537.336
6 Fuselage mass 22.641 43200 423792 9595074.672
7 Fuel in fuselage 22.641 19056.084 186940.184 4232512.707
8
Main landing gear assembly 1 22.641 10800 105948 2398768.668
9
Main landing gear assembly 2 30.3211 7200 70632 2141639.935
10 Payload bay 2 30.321 18750 183937.5 5577168.938
11 Horizontal stabilizer 45.367 6400 62784 2848321.728
12 Vertical Stabilizer 48.501 3200 31392 1522543.392
Total 153975.274 1510497.438 35785598.11
Cg from Nose 23.69126701
Wing alone analysis
S.No Components
Distance from reference line (m) Mass (kg) Weight (N) Moment (Nm)
1 Wing structure 16.35 64000 627840 10265184
2 Wing fuel 1 11.85 14913.457 146301.0132 1733667.006
3 Wing fuel 2 18.428 14913.457 146301.0132 2696035.071
4 Power plant 1 9.57 13600 133416 1276791.12
5 Power plant 2 15 13600 133416 2001240
Total 121026.914 1187274.026 17972917.2
Cg from L.E 15.13796882
Entire aircraft
cg from L.E. Root 11.74118831 Grand total 275002.188 2697771.464 53758515.31
Table 9.4 Cg estimate for half payload and reserve fuel
73
Condition 5 No Payload and Reserve Fuel
Fuselage alone analysis
S.No Components
Distance from reference line (m) Mass (kg) Weight (N) Moment (Nm)
1 Crew 3.043 270 2648.7 8059.9941
2 Nose landing gear 6.086 3600 35316 214933.176
3 Payload bay 1 12.172 0 0 0
4 Fixed equipments 18.641 1149.19 11273.5539 210150.3182
5 Excess mass 22.641 21600 211896 4797537.336
6 Fuselage mass 22.641 43200 423792 9595074.672
7 Fuel in fuselage 22.641 19056.084 186940.184 4232512.707
8
Main landing gear assembly 1 22.641 10800 105948 2398768.668
9
Main landing gear assembly 2 30.3211 7200 70632 2141639.935
10 Payload bay 2 30.321 0 0 0
11 Horizontal stabilizer 45.367 6400 62784 2848321.728
12 Vertical Stabilizer 48.501 3200 31392 1522543.392
Total 116475.274 1142622.438 27969541.93
Cg from Nose 24.47837623
Wing alone analysis
S.No Components
Distance from reference line (m) Mass (kg) Weight (N) Moment (Nm)
1 Wing structure 16.35 64000 627840 10265184
2 Wing fuel 1 11.85 14913.457 146301.0132 1733667.006
3 Wing fuel 2 18.428 14913.457 146301.0132 2696035.071
4 Power plant 1 9.57 13600 133416 1276791.12
5 Power plant 2 15 13600 133416 2001240
Total 121026.914 1187274.026 17972917.2
Cg from L.E 15.13796882
Entire aircraft
cg from L.E. Root 12.54876323 Grand total 237502.188 2329896.464 45942459.12
Table 9.5 Cg estimate for No payload and reserve fuel
74
Condition 6 Full Payload and Half Fuel
Fuselage alone analysis
S.No Components
Distance from reference line (m) Mass (kg) Weight (N) Moment (Nm)
1 Crew 3.043 270 2648.7 8059.9941
2 Nose landing gear 6.086 3600 35316 214933.176
3 Payload bay 1 12.172 37500 367875 4477774.5
4 Fixed equipments 18.641 1149.19 11273.5539 210150.3182
5 Excess mass 22.641 21600 211896 4797537.336
6 Fuselage mass 22.641 43200 423792 9595074.672
7 Fuel in fuselage 22.641 47640.21 467350.4601 10581281.77
8
Main landing gear assembly 1 22.641 10800 105948 2398768.668
9
Main landing gear assembly 2 30.3211 7200 70632 2141639.935
10 Payload bay 2 30.321 37500 367875 11154337.88
11 Horizontal stabilizer 45.367 6400 62784 2848321.728
12 Vertical Stabilizer 48.501 3200 31392 1522543.392
Total 220059.4 2158782.714 49950423.36
Cg from Nose 23.13823575
Wing alone analysis
S.No Components
Distance from reference line (m) Mass (kg) Weight (N) Moment (Nm)
1 Wing structure 16.35 64000 627840 10265184
2 Wing fuel 1 11.85 37283.6425 365752.5329 4334167.515
3 Wing fuel 2 18.428 37283.6425 365752.5329 6740087.677
4 Power plant 1 9.57 13600 133416 1276791.12
5 Power plant 2 15 13600 133416 2001240
Total 165767.285 1626177.066 24617470.31
Cg from L.E 15.13824714
Entire aircraft
cg from L.E. Root 11.36246959 Grand total 385826.685 3784959.78 74567893.67
Table 9.6 Cg estimate for full payload and half fuel
75
Condition 7 Half Payload and Half Fuel
Fuselage alone analysis
S.No Components
Distance from reference line (m) Mass (kg) Weight (N) Moment (Nm)
1 Crew 3.043 270 2648.7 8059.9941
2 Nose landing gear 6.086 3600 35316 214933.176
3 Payload bay 1 12.172 18750 183937.5 2238887.25
4 Fixed equipments 18.641 1149.19 11273.5539 210150.3182
5 Excess mass 22.641 21600 211896 4797537.336
6 Fuselage mass 22.641 43200 423792 9595074.672
7 Fuel in fuselage 22.641 47640.21 467350.4601 10581281.77
8
Main landing gear assembly 1 22.641 10800 105948 2398768.668
9
Main landing gear assembly 2 30.3211 7200 70632 2141639.935
10 Payload bay 2 30.321 18750 183937.5 5577168.938
11 Horizontal stabilizer 45.367 6400 62784 2848321.728
12 Vertical Stabilizer 48.501 3200 31392 1522543.392
Total 182559.4 1790907.714 42134367.17
Cg from Nose 23.52682209
Wing alone analysis
S.No Components
Distance from reference line (m) Mass (kg) Weight (N) Moment (Nm)
1 Wing structure 16.35 64000 627840 10265184
2 Wing fuel 1 11.85 37283.6425 365752.5329 4334167.515
3 Wing fuel 2 18.428 37283.6425 365752.5329 6740087.677
4 Power plant 1 9.57 13600 133416 1276791.12
5 Power plant 2 15 13600 133416 2001240
Total 165767.285 1626177.066 24617470.31
Cg from L.E 15.13824714
Entire aircraft cg from L.E. Root 11.87233251 Grand total 348326.685 3417084.78 66751837.49
Table 9.7 Cg estimate for half payload and half fuel
76
9.3. The variation of cg location is shown below
S.No Details CG % of MAC Variation in percentage
1 Full payload + Full fuel 0.3500 0.0000
2 No payload + Full fuel 0.3794 7.7421
3 Full payload + Reserve fuel 0.3819 8.3523
4 Half payload + Full fuel 0.3652 4.1671
5 Half payload + Reserve fuel 0.3573 2.0531
6 No payload + Reserve fuel 0.3819 8.3523
7 Full payload + Half fuel 0.3819 8.3523
8 Half payload + Half fuel 0.3613 3.1366
Table 9.8 Cg estimate for various conditions
77
10. Drag Estimation
Drag Equation for Entire Aircraft:
Where
10.1. Component Drag Estimates:
Wetted surface area:
Fuselage:
Engine:
Horizontal Stablizer:
78
Vertical Stablizer:
Nose Landing Gear:
Main Landing Gear 1:
Main Landing Gear 2:
Flap:
79
10.2. Total Drag Estimate:
S.No Components
Wetted Surface
area
Permanent components
1 Fuselage 56 0.03 1.68
2 Engine 31.3706 0.03 0.941118
3 Horizontal Stablizer 95.256 0.0052 0.495331
4 Vertical Stablizer 57.154 0.0052 0.297201
3.41365 0.005375
Temporary Components
a Nose Landing Gear 1.4996 0.12 0.179952
b Main Landing Gear 1 7.197 0.12 0.86364
c Main Landing Gear 2 4.7984 0.12 0.575808
Landing Gear total 1.6194 0.00255
i Flap at 45° 75.645 0.016 1.21032 0.001906
ii Flap at 60° 75.645 0.02 1.5129 0.002382
Table 10.1 Coefficient of Drag for different parts of aircraft
Takeoff Performance:
Landing Performance:
Cruise Performance:
80
10.3. Drag Polar
Drag Polar Analysis:
Takeoff Landing Cruise
CL KCL2 CD Takeoff CD Landing CD Cruise (L/D)cruise
-0.5 0.011841886 0.027172886 0.027648886 0.022716886 -22.0100592
-0.4 0.007578807 0.022909807 0.023385807 0.018453807 -21.6757443
-0.3 0.004263079 0.019594079 0.020070079 0.015138079 -19.8175741
-0.2 0.001894702 0.017225702 0.017701702 0.012769702 -15.662073
-0.1 0.000473675 0.015804675 0.016280675 0.011348675 -8.81160102
0 0 0.015331 0.015807 0.010875 0
0.1 0.000473675 0.015804675 0.016280675 0.011348675 8.81160102
0.2 0.001894702 0.017225702 0.017701702 0.012769702 15.662073
0.3 0.004263079 0.019594079 0.020070079 0.015138079 19.8175741
0.4 0.007578807 0.022909807 0.023385807 0.018453807 21.6757443
0.5 0.011841886 0.027172886 0.027648886 0.022716886 22.0100592
0.6 0.017052315 0.032383315 0.032859315 0.027927315 21.4843421
0.7 0.023210096 0.038541096 0.039017096 0.034085096 20.5368352
0.8 0.030315228 0.045646228 0.046122228 0.041190228 19.4220826
0.9 0.03836771 0.05369871 0.05417471 0.04924271 18.2768171
1 0.047367543 0.062698543 0.063174543 0.058242543 17.16958
1.1 0.057314727 0.072645727 0.073121727 0.068189727 16.1314621
1.2 0.068209262 0.083540262 0.084016262 0.079084262 15.1736891
Table 10.2 Coefficient of Drag for different flying conditions
81
Drag Polar Curve:
Figure 10.1 Drag Polar curve for Entire Aircraft during takeoff, cruise and Landing
10.4. Lift to Drag Ratio:
Figure 10.2 L/D vs. CL for Entire aircraft at cruise
(L/D)max= 22
-0.6
-0.4
-0.2
0
0.2
0.4
0.6
0.8
1
1.2
1.4
0 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09
CL
CD
Drag Polar
Takeoff
Cruise
Landing
0
5
10
15
20
25
0 0.5 1 1.5 2 2.5 3
L/D
CL
L/D
L/D
82
11. Performance Calculations:
11.1. Thrust required and Thrust available analysis:
W1= 25% of Fuel and 100 % of Payload
W1= 3185533.292 N
W2= 50% of Fuel and 100 % of Payload
W2= 3784962.23 N
W3= 75% of Fuel and 100 % of Payload
W3= 4384391.173 N
Figure 11.1 Thrust scenarios at Sea level for different weights
-500
0
500
1000
1500
2000
2500
0 200 400 600
Thru
st i
n N
Tho
usa
nd
s
Velocity
Thrust values at sea level
Thrust Available
Tr at w1
Tr at w2
Tr at w3
83
Figure 11.2 Thrust scenarios at 11 km altitude for different weights
Figure 11.3 Thrust scenarios at 25 km for different weights
0
100
200
300
400
500
600
700
800
900
0 100 200 300 400 500
Thru
st i
n N
Tho
usa
nd
s
Velocity
Thrust values at 11 km
Thrust avaliable
Tr at w1
tr at w2
Tr at w3
0
500
1000
1500
2000
2500
3000
0 100 200 300 400 500
Thru
st i
n N
Tho
usa
nd
s
Velocity
Thrust values at 25 km
Thrust avaliable
Tr at w1
tr at w2
Tr at w3
84
Reference:
Websites:
www.airminded.org
www.aviationexplorers.com www.flightglobal.com
www.jet-engine.net
www.wikipedia.org
www.worldaircraftsearch.com
www.worldofkrauss.com
Books:
1. Ajoy Kumar Kundu (2010) ―Aircraft Design – Cambridge Aerospace Series‖ 2. Daniel. P. Raymer (1989) ―Aircraft Design: A Conceptual Approach‖ 3. Lloyd Jekinson & Jim Marchman (2003) ―Aircraft Design Projects For Engineering Students‖ 4. FAA ―Aircraft Weight and Balance Handbook‖ 5. Jan Roskam (1985) ―Airplane Design-Part 1: Preliminary Design and Sizing and Part 4:
Layout design of Landing Gear and system‖
6. Irs.H. Abbott & Albert E Von Doenhoff (1949) ―Theory of Wing Sections‖