Aircraft Design Project.pdf

7/29/2019 Aircraft Design Project.pdf

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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_aircraft


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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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