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ABBREVIATION

- Density of air

- Dynamic viscosity

- Taper ratio

AR - Aspect ratio

b - Wing span

S - Wing area

Swet - Wetted area

Sref - Reference area

C - Chord of the airfoil

Croot - Chord at root

Ctip - Chord at tip

CD - Drag Co-efficient

CL - Lift Co-efficient

D - Drag

L - Lift

E - Endurance

g - Acceleration due to gravity

M - Mach number of aircraft

R - Range

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- Climb Angle

T - Thrust

Re - Reynolds number

ROC - Rate of climb

SL - Landing distance

STO - Take off distance

VCruise- Velocity at cruise

Vstall - Velocity at stall

WCrew - Crew weight

We - Empty weight of aircraft

WF - Weight of fuel

Wpayload- Payload of aircraft

W0 - Overall weight of aircraft

WL - Wing loading

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ABSTRACT

The purpose of our design project was to design a 200 seater passeng

medium range international aircraft by comparing the data and specifications

present aircrafts in this category. Performance characteristics calculations have al

been performed. Necessary graphs have also been plotted from where certa

values where deduced. The aircraft possess a low wing, tricycle landing gear andconventional tail arrangement.

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COMPARATIVE DATA SHEET

Sl no Name of a/c No of

seats

Service

ceiling (km)

Wing span

(m)

Wing area

(m2)

Aspect ratio

1 Bombardier crj100 50 12496 21.21 48.35 16:9

2 Bombardier crj200 50 12496 21.21 48.35 16:9

3 Antonov An-140 52 7600 24.505 51 16:9

4 Fokker-50 58 7620 29 70 12:1

5 Bombardier-Dash 8 50 7680 27.43 56.2 11:1

6 CASA/IPTN CN-235 44 7620 25.8 59.1 11.27:1

7 ATR-42 50 7600 24.5 54.5 11.1:1

8 XIAN MA600 60 7622 29.2 72 12:1

9 VISCOUNT V 700 60 7620 28.56 89 11:1

10 Saab 2000 58 9450 24.76 55.7 11:0

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Sl no Name of a/c Type of engine No of engine Thrust power (KW)

1 Bombardier crj100 2 Turbo prop 6830

2 Bombardier crj200 2 Turbo prop 6875

3 Antonov An-140 2 Turbo prop 1,838

4 Fokker-50 2 Turbo prop 1,864

5 Bombardier Dash 8

Q-300

2 Turbo prop 6875

6 CASA/IPTN CN-235 2 Turbo prop 1,305

7 ATR_42 2 Turbo prop 1465

8 Xian ma 600 2 Turbo prop 2148

9 Vickers viscount v700 2 Turbo prop 1484

10 Saab 2000 2 Turbo prop 3096

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s

Sl no Name of a/c Empty fuel wt (n) Maximum takeoffwt (kg)

Payload (n)

1 Bombardier crj100 13650 24041 6124

2 Bombardier crj200 13830 24041 6124

3 Antonov An-140 12810 21500 13227

4 Fokker-50 12250 20820 6080

5 Bombardier Dash 8

Q-300

19500 2720

6 CASA/IPTN CN-235 9800 15100 4000

7 ATR_42 10600 15550 4500

8 Xian ma 600 13700 21800 5500

9 Vickers viscount v700 16718 30617 6000

10 Saab 2000 13800 22800 5900

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Sl no Name of a/c Range (km) Cruise Speed

(km/hr)

Wing

loading

MAXIMUM

TAKE OFF

SPEED

1 Bombardier crj100 3000 860 497.22 786

2 Bombardier crj200 3045 860 497.22 786

3 Antonov An-140 3270 540 421.23 540

4 Fokker-50 2055 560 297.4

5 Bombardier Dash 8

Q-300

2055 500 346.97

6 CASA/IPTN CN-235 4355 450 255.49

7 ATR_42 1950 500 285.32

8 Xian ma 600 2450 514 302.77

9 Vickers viscount v700 2600 522 344.01

10 Saab 2000 2100 685 409.33

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0

2000

4000

6000

8000

10000

12000

14000

0 1000 2000 3000 4000 5000

Serviceceiling(km)

Range (km)

Service ceiling (km)

Service ceiling (km)

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0

100

200

300

400

500

600

0 1000 2000 3000 4000 5000

Wingloading

Range(km)

Wing loading VS Range

Wing loading

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0

100

200

300

400

500

600

700

800

900

1000

0 1000 2000 3000 4000 5000

CruiseSpeed(km/hr)

Range(km)

Cruise Speed (km/hr)

Cruise Speed (km/hr)

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10:55

11:02

11:09

11:16

11:24

11:31

11:38

11:45

11:52

12:00

12:07

0 500 1000 1500 2000 2500 3000

Aspectratio

Range (km)

Aspect ratio

Aspect ratio 16:09 16:09 16:09

12:01 11:01 11.27:1 11.1:1

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0

50

100

150

200

250

300

350

400

0 1000 2000 3000 4000 5000

THRUSTTOWEIGHT

RANGE

RANGE VS THRUST TO WEIGHT RATIO

Series1

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SL.NO SPICIFICATION DATA

1. Aspect ratio10.9

2. Cruise speed 520KM\Hr

3. Service ceiling 7800M

4. Thrust to weightratio

110

5. Range 2500KM

6. Wing loading 320

7. Max takeoff weight 21000KG

8. L/D ratio 12.2

9. Max takeoff speed

10. Endurance 6.51hrs

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OVERALL WEIGHT ESTIMATION:

The structural design , the complexity of the load distributio

through a redundant structure , and the large number of sophisticate

systems in an a/c , makes weight est a difficult and precarious task.

When the detail design drawing are complicit, the engg s cal the wt

each and every part, and add all and cal the wt. But in the beginnin

phase of design process, this cannot be accomplished because there a

neither detailed drawing of the a/c nor the detail of the various parts the a/c.

And so some approximations are made and overall weight is eventually

estimated.

Overall takeoff weight is given by,

W crew + w payload

W0= .

[1-(we/w0)-(wf/w0)]

Where:-

Wc =crew wt

Wp =payload wt

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Wf =fuel wtWe =structural wt

First approximation:

WC+wp.1

(w0)1 =

[1-(we/WO)-(wf/w0)]

Where:-

(We/w0)= [(Wfixed/w0)+(Wpower plant/w0)+(wstructural/w0)]

Given:

W structural=0.3wo

W power plant=0.06wo

W fixed equipment=0.045wo

W fuel=0.15wo

(We/w0) =0.045+0.06+0.3=0.405

(Wf/w0) =0.15

(WC+wp.1) =1000N+700N=1700N (for one passenger)

*In this we have forty persons including the pilot so,

[Wc/wp.1]=50*1700=85000N

(w0)1= 85000

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= 191011.236N

[1-0.405-0.15]

Second approximation:

Now from the graph of (we/w0) and the max takeoff wt,

We get,

According to a/c max takeoff wt Is 15000kg and (we/w0) =0.405

We know for complete one flight we can separate as

Warm up and take off

Climb

Cruise

Loiter

Desent Landing

[(wf/w0)=1.06(1-(wx/w0)]

Where,

(Wx/w0) = (w1/wo)(w2/w1)(w3/w2)(w4/w3)(w5/w4)(w6/w5)

General flight pattern of forty seated aircraft:

WHERE:-

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(W1/wo) =warm up and takeoff

(W2/w1) = climb

(W3/w2)= cruise

(W4/w3)= loiter

(W5/w4)= Desent

(W6/w5)= landing

(W1/wo)= (wi/wi-1) =0.96

(W2/w1)=0.98

(W3/w2) = (wi/wi-1) +exp^ (-RC/V (L/D))

Where

R=range=2400km

V=max takeoff speed=520km/hr

C=Sp fuel consumption

For TURBOPROP type of engine which we have to select &value from

the table

Cbhp =0.6 and p =0.8

So,

C=0.6*520/550*0.8= 0.78

(W3/W2)=e^(-(2400*0.78/520*14.2))=0.77

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(W4/W3)=EXP^ (-(EC/ (L/D))

Where

E=endurance= 3.54hr

C=sp. fuel consumption

For loiter

L/D=0.866*L/Dmax= 12.2

(W4/W3)=e^ (-3.54*0.822/12.2)=0.787

(W5/W4)= 0.918

(W6/W5)= 0.995

(Wx/Wo)=(W1/Wo)(W2/W1)(W3/W2)(W4/W3)(W5/W4)(W6/W5)=

(0.96) (0.985) (0.77) (0.787) (0.918) (0.995)

(Wx/Wo)= 0.523

(Wf/Wo)=1.06(1-Wx/Wo) = 1.06(1-0.523)= 0.505

(WO)^2=WC+Wp1/ (1-We/Wo-Wf/Wo)

=85000/ (1-0.505-0.405) =192011.36N

Hence first app of forty seater a/c by graphical method is

WO= 192011.36

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Third approximation:

(We/WO)=A WO^cKvs

A=2.36 C= -0.18

Kvs=variable sweep const=1.00 if fixed wing

From second approximation

WO = 192011.36 N

(We/WO)=A Wo^c Kvs=2.36*(5.44*10^5) ^-0.18*1=0.219

Hence overall takeoff wt:

Wo=Wc+Wp1/1-(We/Wo)-(Wf/Wo)=85000/1-0.219-0.470)N

= 194366.54N

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Aerofoil selection

Aerofoil is nothing but the cross section area of wing. The shape

necessary for any wing to produce lift force and hence the prop

selection of aerofoil becomes an important and mandatory step

design.

Aerofoil design is a measure fact of aerodynamics. The aerofoil

completely affected by the flight regime in which the a/c is intended

operate

We know that,

lift=1/2.V^2.CI.S

V stall= 0.25 Vcruise=125km/hr= 37.5 m/sec

For steady and level condition,

L=W

W=1/2.V stall^2.CI max .S

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CI max=2(W/S) 1/.Vstall^2

Where (W/S) =wing loading

CI max= (320*2/ (37.5)2*)2.830

CL max=1.28

Re= V stall. 1/r

Where L is chord length

L=S/AR=54/10.9=4.954

REYNOLDS NUMBER:

R=kinematic viscosity=2.1584*10-5

/sec

Re=0.42*37.33*4.954/2.15*10-5

=3.59*106

So from Reynolds number and CI max we can find the aerofoil NACA

4214, we got the aerofoil NACA4214 as

And for this aerofoil selection the drag co efficient

Cd=0.05

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Thrust calculation:

Thrust is nothing but a reaction force which when acts on bod

produces a motion. It is expressed in Newton. Newtons second law an

third law quantitatively defined thrust together. An aircraft propels wi

the aid of this reaction force. There are three diff independent actio

which causes this reaction to occur

Spinning blades of the propeller or rotating turbine of jet , or th

blast of the propellants in the rocket engine , force a mass of air or gtowards the rear and this mass exerts an equal and opp force on th

system and thus we have a force named , thrust

An a/c with a good propulsion sys design, integrated with a goo

structural char will automatically surprise other phases of design and w

emerge out successfully

We know that

(thrust/wt) takeoff= (thrust/weight) cruse*(Ttakeoff/Tcruise/Wtakeoff)

We know that

(thrust/weight) cruise=1/ (L/D) cruise

=1/12.2=0.0819

We will get

(Wcruise/Wtakeoff) as

(Wx/Wo) = (W1/Wo)(W2/Wo)(W3/W2)

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W1/Wo=0.96Wo

W1=0.97*(191011) =185280.67

W2/W1=0.985

W2=0.985*185280.67

=182501.46

W3/W2=0.818*182501.46

W3=149286.19

As we have to find the fraction of wt at take off to0 the wt at cruise,

Wb/W1=149286.19/185280.67=0.805

(thrust/wt) takeoff= (thrust/wt) cruise*(T take off/T cruise)*(W cruise/W

takeoff)

Thrust=2100kw

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Wing design:

Wings are most important in a/c to generate lift.

The cross section of wing is normally and mostly an aerofoil shape. Th

assures stream lining and reduces the drag.

The design and selection of wing includes chord calculatio

aerodynamic center location, and selection of various angles.

CHORD:It is the line joining from leading edge to trailing edge of an aerofoil,

case of a tapered wing the chord at the root is greater than the chord

the tip.

ROOT CHORD:

It can be determined by evoking the expression.

C root=2S/ b (1)

S=wing area

b=wing span

=chord taper ratio

for low speed subsonic a/c is 0.45

S= (W/W/S)

= (21000/320)=65.625m2

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AR=b^2/S

10.9=b2/65.625

b=26.74m

Root chord (C root)

2S/ b (1+1mda)

=2*65.625/26.74*(1+0.45)

=3.385m

C root= 3.385m

TIP CHORD:

Tip chord can be found from the below expression

=tip chord/root chord

tip chord C tip=root chord

=0.45*3.385=1.523m

ROOT MEAN CHORD:

()=2/3Croot(1^2/(1/))

2/3*3.385(1+0.452/1+0.45)=

()=1.87m

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AERODYNAMINC CENTER (A/D):

It is the point at which the aerofoil where in the coefficient of mome

does not vary with in the change in the angle of attack. Aerodynam

center always lies on the root main chord as a distance of 0.25 from th

leading edge of sub sonic a/c

(A/D)=0.25C root

=0.25*2.99

(A/D)=0.7475m

DISTANCE OF MEAN CHORD FROM THE FUSELAGE CENTER:

The distance () can be cal from the formula

=(b/6)12/1)

=26.74/6*[1+ (2*0.45)]/1+0.45

=5.839

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FUSELAGE DESIGN:

It is hallow and strong tube. It is the main part of the a/c which holds th

crew, cargo and the passenger. It indicates all the parts of the a/c an

holds them up. The fuselage is always streamlined to minimize th

amount of drag produced. The center of the gravity always lies with

the fuselage. This fuselage consists of cockpit where all the a/c contro

are controlled here

As we know

L/b=0.84

L=length of fuselage

b=span

We know

b=26.74

l=26.74*0.84

l=22.46m

As we know the fitness ratio as.

L/d=4.5

d=22.46/4.5

d=4.991m

Length of the empennage of the nose

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=1/10 of length of the fuselage

=1/10(22.46)

=2.246m

DIAMETER OF THE NOSE CONE AND EMPENNAGE:

Lnose/dia nose=1.7

Dia nose=2.246/1.7=1.321m

Length emp/dia emp=1.8

Dia emp=2.246/1.8=1.247m

Wt area:

S reference =C root * dia fuselage

= 3.385 *4.991=16.89m

We know

S=S ref+ S wet

We have seen:

S=65.625m

S wet=S-Sref

=65.625-16.89=48.735m

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Drag and lift estimation:

Like thrust and wt, drag too is a force the thrust produced by the pow

plant, has to overpower those drag force and only by that way the a

can be suspends in the air. Thrust is an aerodynamic force to th

direction of motion.

The drag can be est by the formula

D= (1/2)Vstall^2SCd

Cd is coefficient of drag

CD=CDo=[(CLmax) ^2/(eAR)+

The right hand side has two terms in the eq.the first is parasite drag ter

and the second is the induced drag term.

CDo=CDe*(S wet/S ref)CDe is the effective friction co efficient and its value is 0.003

S wet/S ref=48.735/16.89=2.885

CDo=0.01*2.885=0.02885

CD=CDo+ (CLmax/e AR)

=0.02885+ (1.28/3.14*0.9*10.9)=0.07038

Drag= (1/2) Vstall^2S*CD

=0.5*0.42*37.252*65.625*0.07038=1363.9N

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lift:

The term lift is generally associated with an a/c although lift is als

generated by rotor on helicopter, sails, and keels on the sailboats, an

wind turbines. While the common meaning of lift suggests that

opposes gravity, aerodynamic lift can be in any direction.

We very well know that lift L is given by

L= (1/2)Vstall^2sCLmax

=0.5*0.42*37.332*65.625*1.28

=24581.84N

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Performance:

The evaluation of performance of an aircraft involves of quantiti

such as take off distance, landing distance, rate of climb, climb ang

bank angle, and minimum load factor for turn angle, turn radius, angul

velocity, and endurance. The performance is evaluated with the help

simple parameters like weight, lift and drag coefficients, and engin

thrust characteristics. The formulas used are in the perspective

preliminary design stage of an aircraft. Accurate results require muc

more numerical data.

Take off distance Sto

The distance or the length of the runway required for an aircraft to be a

born success fully is given by the formulaSto = * 1.44 (Wtk)^2+ / (gsCLmax)+ * (T-D) + r (Wo - L) ]

r is the friction coefficient of runway and is usually 0.02

is the density at sea level

T, W, Wtk, WO are in new ton

Sto = [ 1.44*2.8 * 1020 ] / (9.81 * 1.225 * 500.18 * 0.5867 )

[258200 - 10520 + 0.02 *50000 ]

= 1400m

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Hence Sto = 1.4 km

Landing distance SL

The length of the run way required for an air craft to come to a halt afte

it touches is given by

SL = *1.67 ( wl ) ^ 2+ / (gsCLmax ) * (T-D) + R (Wo L )]

Landing weight WL is calculated from the expression below

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WL = WL 30% (WO)

= 1.12 km

Hence , SL = 2.8 km

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Rate of climb Roc and climb angle

Rate of climb is nothing but the rate at which an air craft climbs up whe

it is in pitch up position. The angle at which it climbs in to the

atmosphere is called climb angle

They are given by the following formulaes

Roc = Vcruise * (T-D) / WO

=[520*(1700000-1275000) ] / 1729650

=125.8 m / s

Hence, Roc = 125.8 m / s

Roc = Vcruise sin

Sin =RoC / Vcruise

= sin^-1 (125.8 / 520)

=140

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Bank angle:

The angle between the lateral axis and the horizontal surface of the

earth is called bank angle

It is given by

= cos^ -1 (W/L)

= COS^-1 (0.61 / 1.53)

=66o42

Load factor for turn (nm)

Minimum Load factor required for turn is given by,

Nm = [T/W] * [L/D] where [T/W] is 0.25 from table

=0.25*12.2

=3.05

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Turn angle

The angle with which the aircrafts makes a turn with respect to the

horizontal axis of the earth is called turn angle

Tan = (3.05^2 1 )^0.

=tan ^-1 [3.05^2 - 1] ^ 0.5

=9011

1

Turn Radius R

The radius which the air craft covers a turn is turn radius and is given by

R = Vcruise^2 / [ g* ( nm ^2 1 ) ^0.5 ]

=520^2 / [9.81* (3.05^2 -1) ^ 0.5]

= 9566.06 m

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Endurance E

Endurance is nothing but the

Amount of time spent by the aircraft in air , it is calculated from the

expression

E = ln (WL/Wo)*[(L/D)]

Where (L/D) and C for endurance is 1.6 & 0.4 respectively

= - ln (1.12 / 1.91) * [12.2]

= 6.51 Hrs

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V-N DIAGRAM

The V-n diagram is a graph portraying load factor vs velocity f

given airplane, along with the constrains on both n and v due

structural limitations. The V-n diagram illustrates some particular

important aspects of overall airplane performance.

Load factor aids us in fixing boundaries to an aircraft within which th

aircraft is free to perform and operate. Load factor is dependent o

gravity and hence depending on that we have corresponding on that whave corresponding velocities and eventually V-n plot.

For our calculation, we consider load factors direct proportionality to

the square of velocity. Load factor is given by

n= V2/Vstall

2

Positive load factors indicate that the aircraft is ascending up.

When n=1, V=37.5 m/s

When n=2, V=53.03 m/s

When n=3, V=64.95 m/s

When n=4, V=75 m/s

Negative load factors are experienced by aircraft when it descends dow

When n=-1, V=-37.5m/s

When n=-2, V=-53.03 m/s

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When the velocity is 0, load factor is also zero.

Load factor n=1 gives an initial boundary limit and a dive speed of 200.2

m/s gives final boundary limit.

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The area exposed by continues lines in the plot is the regime in which

the aircraft is bound to perform an operate. The first vertical line

crossing X-axis at 33.33m/s sets the boundary of minimum speed Vmin

The second vertical line crossing X-axis at 200.25m/s sets the boundary

of maximum speed Vmax.

Thus the aircraft can operate between velocities of 37.3m/s an

200.25m/s.

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SUMMARY OF DESIGN FEATURES

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CONCLUSION

Early aircrafts were developed in response to very simple requirement

Today, complex set of requirement like specification of airplane performanc

safety, reliability and maintainability, and other are included. Because t

companies are continuing to try improve on the strategy. In the early days

airplane design, people did not do much computation. The design teams tende

to be small. Modern design projects are so complex that the problem has

examine advertisement for aircrafts; the definition of the best aircraft is vesimple. Aircraft Company sells the fastest, most efficient, quietest, most

expansive airplanes with the shortest field length. Unfortunately such an airpla

cannot exist. As professor Bryson, the father of scientific climatelogy puts it, yo

can only manke one thing best at time. The most expansive airplane would sure

not be the fastest; the most efficient would not be the most comfortable.

Airframe manufacturers are continuously creating innovative design, maki

greater use of new lightweight materials and increasing their focus on passeng

comfort. Achieving a perfect balance between these competing requiremen

represents a tremendous challenge for the design and engineering of t

airframe.

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REFERENCE

The following are the list of the books and websites which helped us to make o

dream a reality and brought our aircraft to life.

Aircraft design-conceptual approach-by Raimer

Introduction to flight-by John D. Anderson

Aerodynamics for engineers- by Arthur and carruther

Aircraft design projects for engineering students B Lloyd and R.Jenkinson

Websites

www.google.com

www.wikipedia.com

www.airliners.com

www.airtoaircombat.com

www.ebookee.com www.nasa.gov.in

http://www.google.com/http://www.google.com/http://www.wikipedia.com/http://www.wikipedia.com/http://www.airliners.com/http://www.airliners.com/http://www.airtoaircombat.com/http://www.airtoaircombat.com/http://www.ebookee.com/http://www.ebookee.com/http://www.nasa.gov.in/http://www.nasa.gov.in/http://www.nasa.gov.in/http://www.ebookee.com/http://www.airtoaircombat.com/http://www.airliners.com/http://www.wikipedia.com/http://www.google.com/

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