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CHAPTER 1
Introduction
Business jet, private jet, orbizjet, is a term describing a jet aircraft, usually of smaller
size, designed for transporting groups of up to 19 individuals. Business jets may be adapted for
other roles, such as the evacuation of casualties or express parcel deliveries, and some are used
by public bodies, government officials or the armed forces. The more formal terms of corporate
jet, executive jet, VIP transport or businessjettend to be used by the firms that build, sell, buy
and charter these aircraft
Most business jet aircrafts are of low-wing design and have engines mounted at the aft end of thefuselage. Except for one three-engine and one four-engine design, all are used. Most of the
modern aircraft produced today have turbofan engines; some of these are repowered versions ofthe aircraft that originally appeared with turbojet engines. The wings of most of the aircraft have
a modest amount of sweepback, although one business jet described below has a swept forward
wing. Like any aircraft, the size and performance of business jets vary with the function forwhich the aircraft has been designed. Aircraft are available that vary in gross weight from about
11000to 65000 pounds. Cruising speeds lie in the range from 0.7 to 0.85 Mach number. Ranges
Vary from intercontinental values as low as 1150 miles. Many of the new aircrafts being
produced have at least nonstop transcontinental capability. The number of passengers that can beaccommodated, even on an aircraft of the same design, varies widely depending on the interior
cabin arrangements. Aircraft can be found with the capability of carrying from 5 to 18passengers.
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CHAPTER-2
Load Factor Requirements
Load factor is defined as the ratio of the lift of an aircraft to its weight and represents a globalmeasure of the stress ("load") to which the structure of the aircraft is subjected:
Where:
n = Load factor
L = Lift
W= Weight
Since the load factor is the ratio of two forces, it is dimensionless. However, its units aretraditionally referred to as g, because of the relation between load factor and apparentacceleration of gravity felt on board the aircraft. A load factor of one, or 1 g, represents
conditions in straight and level flight, where the lift is equal to the weight. Load factors greater
or less than one (or even negative) are the result of maneuvers or wind gusts
Load factor, n = WS CV
21 M A X2
From Aircraft Design Lab-1 Data Book
1. Takeoff weight = 10000kg
2. Air density at ground = 1.225 kg/m3
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3. Clmax = 1.71, -1.1
4. Wing span Area = 14.22 m2
Table 1- Load factor for various aircraft velocities
V n
0 0
10 0.006998
20 0.02799
30 0.062978
40 0.111961
50 0.17493960 0.251912
70 0.34288
80 0.447843
90 0.566801
100 0.699754
110 0.846703
120 1.007646
130 1.182585
140 1.371519150 1.574447
160 1.791371
170 2.02229
180 2.267204
190 2.526113
200 2.799018
210 3.085917
220 3.386811
230 3.701701240 4.030585
250 4.373465
260 4.73034
270 5.101209
280 5.486074
290 5.884934
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300 6.297789
310 6.72464
320 7.165485
330 7.620325
340 8.089161
350 8.571991
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CHAPTER-3
V-n Diagram
A chart of velocity versus load factor (or V-n diagram) is another way of showing limits
of aircraft performance. It shows how much load factor can be safely achieved at different
airspeeds.
The higher the speed, the greater is the load factorn. Compressibility has an effect on the
V-n diagram. In principle, it may be necessary to construct several V-n diagrams representing
different altitudes. This chapter explains only the role of the V-n diagram in aircraft design.
Figure 2 represents a typical V-n diagram showing varying speeds within the specified structural
load limits. The figure illustrates the variation in load factor with airspeed for maneuvers. Somepoints in a V-n diagram are of minor interest to configuration studiesfor example, at the point
V= 0 and n = 0 (e.g., at the top of the vertical ascent just before the tail slide can occur). The
points of interest are explained in the remainder of this section. Inadvertent situations may take
aircraft from within the limit-load boundaries to conditions of ultimate-load boundaries.
Figure 1- Typical V-ndiagram showing load and speed limits
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Table 2- V Vs n
V n negative n
0 0 0
10 0.0069975 -0.004453
20 0.0279902 -0.017812
30 0.0629779 -0.040077
40 0.1119607 -0.071248
50 0.1749386 -0.111325
60 0.2519116 -0.160307
70 0.3428796 -0.218196
80 0.4478428 -0.284991
90 0.5668011 -0.360692
100 0.6997544 -0.445298
110 0.8467028 -0.538811
120 1.0076463 -0.641229
130 1.1825849 -0.752554
140 1.3715186 -0.872785
150 1.5744474 -1.001921
160 1.7913712 -1.139964
170 2.0222902 -1.286912
180 2.2672042 -1.442766
190 2.5261133 -1.607527
200 2.7990175 -1.781193
210 3.0859168 -1.963765
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220 3.3868112 -2.155244
230 3.7017007 -2.355628
240 4.0305853 -2.564918
250 4.3734649 -2.783114
260 4.7303396 -3.010216
270 5.1012095 -3.246224
280 5.4860744 -3.491138
290 5.8849344 -3.744958
300 6.2977895 -4.007684
310 6.7246396 -4.279316
320 7.1654849 -4.559854
330 7.6203253 -4.849298
340 8.0891607 -5.147648
350 8.5719912 -5.454904
Cruise Velocity Vc = 300 m/s
Dive Velocity, VD = 1.5 * Vc
= 1.5 * 300
VD = 450m/s
Using all these data, the final V-n Diagram is given in the following figure.
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Figure 2- V-n Diagram
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CHAPTER 4
Other Load Calculations
4.1 Interception
An interceptor aircraft (or simply interceptor) is a type offighter aircraft designed
specifically to prevent missions of enemy aircraft, particularly bombers and reconnaissance
aircraft, and destroy them relying usually on great speed and powerful armament.
The load acting on the interceptor is given by the following formula
Table 3- V Vs n for Interception at various densities
Velocity(m/s) Density(Kg/m ) n
50 1.225 0
50 0.5489 0.1218558
50 0.3652 0.0810744
100 1.225 1.0878
100 0.5489 0.4874232
100 0.3652 0.3242976
150 1.225 2.44755150 0.5489 1.0967022
150 0.3652 0.7296696
200 1.225 4.3512
200 0.5489 1.9496928
200 0.3652 1.2971904
http://en.wikipedia.org/wiki/Fighter_aircrafthttp://en.wikipedia.org/wiki/Bomber_aircrafthttp://en.wikipedia.org/wiki/Reconnaissance_aircrafthttp://en.wikipedia.org/wiki/Reconnaissance_aircrafthttp://en.wikipedia.org/wiki/Reconnaissance_aircrafthttp://en.wikipedia.org/wiki/Reconnaissance_aircrafthttp://en.wikipedia.org/wiki/Bomber_aircrafthttp://en.wikipedia.org/wiki/Fighter_aircraft7/29/2019 Aircraft Design Project.pdf
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250 1.225 6.79875
250 0.5489 3.046395
250 0.3652 2.02686
300 1.225 9.7902
300 0.5489 4.3868088
300 0.3652 2.9186784
350 1.225 13.32555
350 0.5489 5.9709342
350 0.3652 3.9726456
Figure 3- V vs n for Interception
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4.2 Instantaneous Turn Rate
Instantaneous turn rate describes maximum g turns which cause a loss in energy, either in
the form of speed or altitude. These turns are unsustainable, although to some degree the energy
loss may be compensated for by increasing thrust, known as applying "excess specific power."
This usually occurs during hard turns or even harder breaks.
The immediate bank angle an aircraft can achieve before drag seriously bleeds off
airspeed is known as its instantaneous turn performance. An aircraft with a small, highly loaded
wing may have superior instantaneous turn performance, but poor sustained turn performance: it
reacts quickly to control input, but its ability to sustain a tight turn is limited.
Table 4- Load factor at different velocities at different ITR
Velocity(m/s) Turn rate(deg/s)
Turn rate(Rad/sec) n
0 1
50 2 0.03488 1.015698843
50 4 0.06976 1.06140311
50 6 0.10464 1.133488979
50 8 0.13952 1.22731668650 10 0.1744 1.338321155
50 12 0.20928 1.462596684
50 14 0.24416 1.597048179
50 16 0.27904 1.739317393
50 18 0.31392 1.887637514
150 2 0.03488 1.133488979
150 4 0.06976 1.462596684
150 6 0.10464 1.887637514
150 8 0.13952 2.357277293
150 10 0.1744 2.849549372150 12 0.20928 3.354504663
150 14 0.24416 3.867178038
150 16 0.27904 4.384863162
150 18 0.31392 4.905973752
250 2 0.03488 1.338321155
250 4 0.06976 2.040689603
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250 6 0.10464 2.849549372
250 8 0.13952 3.695626634
250 10 0.1744 4.558243943
250 12 0.20928 5.429523597
250 14 0.24416 6.305876003
250 16 0.27904 7.185445349
250 18 0.31392 8.06717947
350 2 0.03488 1.597048179
350 4 0.06976 2.683701091
350 6 0.10464 3.867178038
350 8 0.13952 5.080256508
350 10 0.1744 6.305876003
350 12 0.20928 7.537921724
350 14 0.24416 8.773686879
350 16 0.27904 10.01179428
350 18 0.31392 11.25147074
400 2 0.03488 1.739317393
400 4 0.06976 3.016769792
400 6 0.10464 4.384863162
400 8 0.13952 5.779584753
400 10 0.1744 7.185445349
400 12 0.20928 8.596982018
400 14 0.24416 10.01179428
400 16 0.27904 11.4286657
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Figure 4- V Vs n at different ITR
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4.3 Sustained Turn Rate
Sustained Turn Rate is where a plane maximizes its smallest turn radius, g - load, and
speed to acquire the best possible turn rate and continuously sustains the turn for long periods of
time, without giving up alt, speed, or degrees of turn.The maximum rate of turn possible for a
given aircraft design is limited by its wing size and available engine power: the maximum turn
the aircraft can achieve and hold is itssustained turn performance.
Table 5- Velocity and load factor for Sustained turn rate at 14 deg/sec
Velocity
(m/s) Turn rate (deg/s) Turn rate (Rad/sec) n
0 14 0.2443 1
50 14 0.24431.5976
100 14 0.24432.6850
150 14 0.2443 3.8692
200 14 0.24435.0830
250 14 0.24436.3094
300 14 0.24437.5421
350 14 0.2443 8.7786
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Figure 5- V Vs n at STR 14 deg/sec
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CHAPTER 5
Wing Load Distribution
The loads on the wing are made up of aerodynamic lift and drag forces, as well as
concentrated or distributed weight of mounted engines, stored fuel, weapons, structural elements
etc.
The following formulae were used to compute the Lift distribution on the wing.
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The following formulae can be used to compute the shear force distribution and the
moment distribution on the wing.
Table 6- Lift, Shear Force and Bending Moment acting on the Wing
Y a=Y/(b/2) c(y) L(Elip) L(trap) L(bar) V.Lift M Lift
0 0 2.19 961.0806 1371.93 1166.505 2306.739 10932.59
0.368 0.055547 2.326988 959.5967 1320.871 1140.234 2252.706 10083.71
0.736 0.111094 2.316159 955.1314 1269.813 1112.472 2195.67 9254.717
1.104 0.166642 2.297999 947.6423 1218.754 1083.198 2135.575 8446.711
1.472 0.222189 2.27233 937.0571 1167.696 1052.376 2072.33 7660.819
1.84 0.277736 2.238895 923.2692 1116.637 1019.953 2005.809 6898.202
2.208 0.333283 2.197339 906.1326 1065.579 985.8555 1935.842 6160.064
2.576 0.38883 2.147191 885.4526 1014.52 949.9863 1856.917 5447.6752.994 0.451925 2.079014 857.3379 956.5241 906.931 1779.313 4764.329
3.312 0.499925 2.018448 832.3621 912.4028 872.3824 1702.641 4109.542
3.68 0.555472 1.937965 799.1726 861.3442 830.2584 1615.805 3482.97
4.048 0.611019 1.844929 760.8068 810.2857 785.5462 1523.377 2888.354
4.416 0.666566 1.737326 716.4337 759.2271 737.8304 1424.34 2327.751
4.784 0.722113 1.612241 664.8514 708.1685 686.51 1317.173 1803.594
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L Flap V Flap M Flap W Fuel V Fuel M Fuel W
Structure
V
Structure
M
Structure
350 3850 8500.8 -900 -9900 -14850 -108 -1026 -1714.5
350 3500 7084 -900 -9000 -12375 -102 -918 -1458
350 3150 5796 -900 -8100 -10125 -96 -816 -1228.5
350 2800 4636.8 -900 -7200 -8100 -90 -720 -1024.5
350 2450 3606.4 -900 -6300 -6300 -84 -630 -844.5
350 2100 2704.8 -900 -5400 -4725 -78 -546 -687
350 1750 1932 -900 -4500 -3375 -72 -468 -550.5
350 1400 1288 -900 -3600 -2250 -66 -396 -433.5
350 1050 772.8 -900 -2700 -1350 -60 -330 -334.5
350 700 386.4 -900 -1800 -675 -54 -270 -252
350 350 128.8 -900 -900 -225 -48 -216 -184.5
0 0 0 0 0 0 -42 -168 -130.5
0 0 0 0 0 0 -36 -126 -88.5
0 0 0 0 0 0 -30 -90 -57
0 0 0 0 0 0 -24 -60 -34.5
0 0 0 0 0 0 -18 -36 -19.5
0 0 0 0 0 0 -12 -18 -10.5
0 0 0 0 0 0 -6 -6 -6
0 0 0 0 0 0 0 0 0
3850 -9900 -1026
5.152 0.77766 1.465204 604.2166 657.1099 630.6633 1199.409 1318.874
5.52 0.833208 1.288722 531.4395 606.0514 568.7454 1066.513 877.4917
5.888 0.888755 1.068301 440.5431 554.9928 497.7679 907.8715 485.0148
6.256 0.944302 0.76695 316.2728 503.9342 410.1035 410.1035 150.9181
6.6249 1 0 0 0 0 0 0
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W
engine
V
Engine
M
Engine
Total V Total M
0 -317.5 -
1285.24
-5086.76 1583.653
0 -317.5 -1168.4 -4482.79 2166.3130 -317.5 -
1051.56
-3887.83 2645.657
0 -317.5 -934.72 -3301.93 3024.291
0 -317.5 -817.88 -2725.17 3304.839
0 -317.5 -701.04 -2157.69 3489.962
0 -317.5 -584.2 -1599.66 3582.364
0 -317.5 -467.36 -1056.58 3584.815
-63.5 -317.5 -350.52 -518.187 3502.109
-63.5 -254 -233.68 78.64085 3335.262
-63.5 -190.5 -140.208
659.3046 3062.062
-63.5 -127 -70.104 1228.377 2687.75
-63.5 -63.5 -23.368 1234.84 2215.883
0 0 0 1227.173 1746.594
0 0 0 1139.409 1284.374
0 0 0 1030.513 857.9917
0 0 0 889.8715 474.5148
0 0 0 404.1035 144.9181
0 0 0 0 0
42693.35
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Figure 6- Y Vs Lift distribution on wing
Figure 7- Y Vs Shear Force distribution on Wing
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Figure 8- Y Vs Bending Moment distribution on Wing
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Similar to the loads acting on a Wing, the Shear force and the Bending moment can be computed
from the following formulae.
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Table 7- Loads, Shear force and Bending Moment distribution on the Fuselage
Length x/L W Pay V Pay M Pay
0 0 0 0 1650
0.7 0.05 0 0 1650
1.4 0.1 0 0 1650
2.1 0.15 -60 60 1650
2.8 0.2 -60 120 1620
3.5 0.25 -60 180 1560
4.2 0.3 -60 240 1470
4.9 0.35 -60 300 1350
5.6 0.4 -60 360 1200
6.3 0.45 -60 420 1020
7 0.5 -60 480 8107.7 0.55 -60 540 570
8.4 0.6 -60 600 300
9.1 0.65 0 0 0
9.8 0.7 0 0 0
10.5 0.75 0 0 0
11.2 0.8 0 0 0
11.9 0.85 0 0 0
12.6 0.9 0 0 0
13.3 0.95 0 0 0
14 1 0 0 0-600
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W Struct V Struct M Struct W Engine V Engine M Engine W Wing V Wing M Wing
-125 0 3762.5 0 0 -1359.13 0 0 350
-125 125 3762.5 0 0 -1359.13 0 0 350
-125 250 3675 0 0 -1359.13 0 0 350
-125 375 3500 0 0 -1359.13 0 0 350
-125 500 3237.5 0 0 -1359.13 0 0 350
-125 625 2887.5 0 0 -1359.13 0 0 350
-125 750 2450 0 0 -1359.13 -20 20 350
-125 875 1925 0 0 -1359.13 -20 40 336
-125 1000 1312.5 0 0 -1359.13 -20 60 308
-125 1125 612.5 0 0 -1359.13 -20 80 266
-125 1250 -175 0 0 -1359.13 -20 100 210
-125 1375 -1050 0 0 -1359.13 -20 120 140
-125 1500 -2012.5 0 0 -1359.13 -20 140 56
-125 -1000 -3062.5 0 -373.56 -1359.13 -20 -40 -42
-125 -875 -2362.5 0 -373.56 -1097.64 -20 -20 -14-125 -750 -1750 -67.33 -373.56 -836.15 0 0 0
-125 -625 -1225 -67.33 -306.23 -574.658 0 0 0
-125 -500 -787.5 -67.33 -238.9 -360.297 0 0 0
-125 -375 -437.5 -67.33 -171.57 -193.067 0 0 0
-125 -250 -175 -67.33 -104.24 -72.968 0 0 0
-125 -125 0 -67.33 -36.91 0 0 0 0
-2625 -403.98 -180
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Figure 10- X Vs V of Fuselage
W Rudder V Rudder M Rudder L Tail V L tail M L tail Total V Total M
0 0 -4975 0 0 -33600 0 -34171.6
0 0 -4975 0 0 -33600 125 -34171.6
0 0 -4975 0 0 -33600 250 -34259.1
0 0 -4975 0 0 -33600 435 -34434.1
0 0 -4975 0 0 -33600 620 -34726.6
0 0 -4975 0 0 -33600 805 -35136.6
0 0 -4975 0 0 -33600 1010 -35664.1
0 0 -4975 0 0 -33600 1215 -36323.1
0 0 -4975 0 0 -33600 1420 -37113.6
0 0 -4975 0 0 -33600 1625 -38035.6
0 0 -4975 0 0 -33600 1830 -39089.1
0 0 -4975 0 0 -33600 2035 -40274.1
0 0 -4975 0 0 -33600 2240 -41590.6
0 -1250 -4975 0 -8000 -33600 -10663.6 -43038.6
0 -1250 -4100 0 -8000 -28000 -10518.6 -35574.10 -1250 -3225 0 -8000 -22400 -10373.6 -28211.2
-250 -1250 -2350 0 -8000 -16800 -10181.2 -20949.7
-250 -1000 -1475 0 -8000 -11200 -9738.9 -13822.8
-250 -750 -775 -8000 -8000 -5600 -9296.57 -7005.57
-250 -500 -250 0 0 0 -854.24 -497.968
-250 -250 0 0 0 0 -411.91 0
-1250 -8000
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Figure 11- X Vs M of Fuselage
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CHAPTER 7
Detailed Layout of the Aircraft
The main conclusion from the constraint analysis and aircraft performance
estimations is that the aircraft landing requirements are too tight and should be renegotiated
with the customers. To provide evidence on the effects of the landing constraints, the revised
baseline layout will ignore them. The new design can be analyzed to show what landing
characteristics are feasible.
With the above strategy in mind the design point for the aircraft will be moved closer
to the intersection of the take-off and climb constraint lines, i.e.:
(T/W) = 0.20 and (W/S)= 362.35 kg/m2 (80 lb/sq. ft)
Anticipating the need to increase aircraft mass to allow more fuel to be carried, the
maximum take-off mass is increased to 9850 kg (and the structural design mass
increased to 10100 kg). Using the new values for(T/W) and (W/S)the new thrust
become:
T = 0.20 9850 = 1970 kg (SSL)
For an aspect ratio (AR) of 5, the new area gives a wing span (b) = 8.56 m and a mean
chord = 1.71 m. For an aspect ratio of 4.5 the wing geometry becomes b = 8.12 m and
mean chord = 1.80 m. Rounding these figures for convenience of the layout drawing
gives:
cmean = 1.75 m (75 ft) and b = 8.5 m (28 ft)
gives,AR = 4.86 and S = 14.87 sq. m/160 sq. ftThis geometry will be used in the new layout.
Also, since the tip chord on the previous layout seemed small, the taper ratio will
be increased to 0.33.
Hence Cmean = (Ctip + Croot )/2 = 1.75 m (assumed)
With, (Ctip/Croot) = 0.33
This gives Croot = 2.63m/8.6 ft, Ctip = 0.87m/2.8 ft
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Figure 12- Final Aircraft Layout
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CHAPTER-8
Conclusion
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