Aerodynamic deisgn report1

8/7/2019 Aerodynamic deisgn report1

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CONCEPTUALAERODYNAMIC DESIGN

200 SEATER PASSENGER

AIRCRAFT

Guided by, Dr. S. Elangovan Ph.D.

Submitted by,

S.Poonkodi (20033016)K.Vasanth Kumar (20033034)S.Anand (20033035)S.Rengarajan (20033019)S.Ashok (20033004)G.Prabakaran (20033017)D.Kishore (20033011)


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Collection of comparative data:

In the designers perspective it is necessary to compare the existing airplanes

that are of same type as that of our desired airplane. Their important parameters,

positive aspects to be considered and pitfalls to be overcome are taken intoconsideration.

The comparative data for our aircraft was collected from the book Janes

All the Worlds Aircrafts based upon the nearest no. of passengers. Theparameters compared were

1. Velocity (V)

2. Range (R)

3. Weight (W)

4. Wing area (S)

5.

Aspect ratio (AR)6. Wing span (b)

7. Overall length (l)

Selection of Parameters:

The comparative data for different 100 seater aircrafts were studied

and the following fundamental design parameters were selected (Graphs for

comparative data are enclosed).

Sl.

No.Parameters Values

1 Cruising velocity of the airplane V 238 m/s

2 Aspect Ratio 7.9

3 Wing Loading 5837 N/m2

4 Range 5000 km5 Span to Length ratio (b/l) 0.75


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Comparative data for 200 seater aircraft

Sl.

No.Name of the aircraft Passengers Velocity Altitude Range Weight Wing area

V

(m/s) (m) R (km) W (kg) S (m)2

1 Airbus A300-600 190 238 10668 8300 153000 219.00

2 Airbus A321-100 189 226 11278 4300 82200 123.00

3 Boeing 727-200 218 232 10668 3320 95030 157.90

4 Boeing 737-800 186 235 12500 5425 70535 135.00

5 Boeing 757-200 196 232 11000 3850 93500 184.17

6 Boeing 767-200 186 238 11887 7079 113395 185.25

7 Douglas DC-8-32 176 235 9150 7410 140614 257.40

8 Douglas DC-8-63CF 180 238 9150 3445 161025 271.90

9 Ilyushin II-86 234 232 11000 3600 208000 320.00

10 Ilyushin Il-62 186 214 12000 6700 162200 279.50

11 McDonnell MD -83 172 229 10668 4635 72575 112.30

12 McDonnell MD 81 172 229 10668 2897 63503 112.30

13 McDonnell MD -90 172 226 10668 3860 70760 112.30

14 Tupolev Tu-154M 180 238 12100 3900 102000 201.50

15 Tupolev Tu-204 202 235 11900 6600 100000 201.45


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Wing loading Vs Mach

0.00

100.00

200.00

300.00

400.00

500.00

600.00

700.00

800.00

900.00

0.70 0.72 0.74 0.76 0.78 0.80 0.82 0.84 0.86

Mach number

W

/s(kgm

-2)

Aspect ratio Vs Mach number

0.0

2.0

4.0

6.0

8.0

10.0

12.0

0.70 0.72 0.74 0.76 0.78 0.80 0.82 0.84 0.86

Mach number

A.R.


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Fineness ratio Vs Mach number

0.00

2.00

4.00

6.00

8.00

10.00

12.00

14.00

16.00

0.70 0.72 0.74 0.76 0.78 0.80 0.82 0.84 0.86

Mach number

Finenessratio

Wing Planform:


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Preliminary Weight Estimation:

The formula to be used for the approximate first weight estimation for

the design of the required aircraft is given below and the proportional weights for the

components were assumed as shown below.

Wo = Wstruc + Wp/l + Wcrew + Wfuel + Wp/p+ Wfe

We = 0.53 WoWp/l = 200 kN (1000 N passenger weight including baggage)

Wcrew = 13 kN (1000 N / crew member)

Wfuel = 0.25 WoWp/p = 0.1 W

Wfe = 0.045 W

Wo = 0.53 Wo + 2*105 + 13*103 + 0.25 Wo + 0.1 Wo + 0.045 Wo

Wo = 955 kN

Refined Weight Estimation:

From comparative data

Treq/W = 0.22

Treq = 126 kN per engine

T = To * 1. 2

Wfuel = Ne * T* c * Range * 1.2/VcWhere

T = Thrust at an altitude of 10500 m

Ne = No. of engines = 2

= /o = 0.317c = Specific fuel consumption = 1.5*10-4 N/Ns

Vc = Velocity at cruise = 237.75 m/s

Wfuel = 240.05 kN

Engine Selection:

CFM56-5B1

T O thrust 187 kN

Weight of the engine 32.32 kN

SFC 16.09 mg/Ns

Weight of Fuel 407.36 kN


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Airfoil and Wing dimensions:

Aspect Ratio(AR) 9Area(S) 221.26 m

2

Span(b) 44.62 m

Length(L) 44.62 m

Calculation of CL, M, McrDo:

The coefficient of lift is calculated for the wing loading, cruise speed and flying

altitude,

Lift Coefficient CLreq = 2)/(2

V

SW

= 2.07

acruise = 297.78 m/s

Mcruise = Vcruise / acruise = 0.8

McrD = Mcruise + 0.04 = 0.84

McrDo = McrD + McrD(AR) - McrD(CL)= 0.84 + 0 -(-0.045)

= 0.885

Calculation of, t/c, , :

Using the graph between McrDo & t/c in Subsonic & Supersonic Airplane

Design by Gerald Corning and fixing a sweep back angle ( ) of 35, the ratio of t/cwas selected. In order to maintain the closeness of the design to the realistic airplanes

a variable taper ratio from the wing root to the tip is used.

Sweep back 25t/c 0.15

Taper Ratio ( ) 0.24

Dihedral () 7


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Airfoil selection:

From the aerofoil data book various airfoils of required t/c are taken and are

tabulated for maximum lift coefficient and minimum drag.

From the table the airfoil with optimum combination of maximum lift

coefficient and minimum drag coefficient is selected.

Airfoil CLmax CDomin L/D

652015 1.4 0.004 350

652215 1.5 0.004 375

652415 1.6 0.004 400

652415 (a = 0.5) 1.6 0.004 400

Selected airfoil : NACA652-415

Cruise CL = 0.532

Volume of fuel:

Cr=)1(

2+b

S = 7.34 m

Ct = 0.24 * Cr = 1.76 m

Cmean = 5.12 m

Volume of the fuel that can be stored in the wing

V = (0.5*Cm * t/c * Cm * b/2 * 0.75)*0.75*2 = 39.75 m3

So volume of fuel that can be stored in wings is 39.75 m3

Total Volume of the fuel to be carried

Vt =)*81.9( fuel

fuelW

= 39.75 m3

The fuel is completely accommodated in the wing.


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Flap selection:

With zero final velocity and with deceleration aided by thrust reverser, landing

speed is calculated using the following equation. Since our aircraft is a 200 seater

aircraft,

Runway length = 2500 m

Ground run= 60% of (Runway length) = 1500 m

Landing Velocity Vldg = (2*a*SL)^0.5 = 76.72 m/s

Stalling Velocity VStall= Vldg / 1.3 = 59.01m/s

Takeoff Velocity VTO = 1.2 * VStall = 70.82 m/s

Cruise CL = 2*W/(S*alt*Vcr2) = 0.532

CL max = 2*W/(S*sl*Vst2) = 2.74

Takeoff CL = 2*W/(S*sl*Vto2) = 1.9

Landing CL = 2*W/(S*sl*Vldg2) = 1.62

Average CL = CL Required = 1.76

CLreq = 0.11

Fowler Flap of20% chord with deflection of25 provides the required CL.

Tyre selection:

During landing and takeoff, the undercarriage supports the total weight of the

airplane. Undercarriage is of three types:1. Bicycle type

2. Tricycle type

3. Tricycle tail wheel type.

A tricycle wheel type needs more takeoff distance and floor is also needs to be

inclined. So we have selected a tricycle nose wheel type. Also the types of runways

are also be selected with due care since depending on this criterion, wheels and tyres

are selected.

Being an executive transport vehicle, it may land on grass fields and may even

on beaches. For tricycle nose wheel type undercarriage, the nose wheel carries 10 %

of the total load and the main undercarriage carries 90% of the total load.

For different runways, the allowable loadings are given by,

Grass 21.1 N/cm2

Grass Strip 36.9 N/cm2

Asphalt (Tar) 73.9 N/cm2

Concrete 116.1 N/cm2


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Main landing gear takes up about 90% of the total weight

Weight taken by one wheel = 109.88 kN

Tyre specifications:

Main landing gear:Track = 10.8 m

Number of wheels = 12 (6+6), 3 wheel bogeys

Load shared by each wheel = 71.69 kN

Wheel diameter = 1.13 m

Wheel width = 0.373m

Wheel contact area = 0.1019 m2

Nose landing gear:

Number of wheels = 2

Load shared by each wheel = 47.73 kN

Wheel diameter = 0.963 m

Wheel width = 0.323 m

Wheel contact area = 0.089 m2

Horizontal Tail parameters:

Volume ratio VHT = 0.8736

Tail arm lHT = 17.2 m

Tail area SHT = 42.52 m

2

Span b = 13.59 m

Root chord Cr = 4.97 m

Tip chord Ct = 1.29 m

Aspect ratio A.R = 4.345

Vertical Tail parameters:

Volume ratio VVT = 0.0805

Tail arm lVT = 16.08 m

Tail area SVT = 29.47 m2

Tail height h = 7.07 m

Root chord Cr = 6.12 m

Tip chord Ct = 2.20 m

Aspect ratio A.R = 1.7

The airfoil chosen for both horizontal tail and vertical tail is NACA-0012


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Center of gravity calculations:

An aircraft is a rigid (assumed) system comprising of many more components

with all these components to be in the air medium. To have a stable aircraft system

and easily controllable, its center of gravity should be positioned in an appropriatemanner. So the weights in the aircraft should be distributed such that it has a defined

c.g position, which is critical. Also the weight distribution should be such that on

certain situations where some components may be consumed or even removed, its c.g.

movement should be in a controllable manner so that is not compromised.

One important condition is that when fully loaded, the c.g. is at 30 % of mean

aerodynamic chord and in different situations such as landing, with or without

payload, the c.g. movement should be restricted within 25% of mean aerodynamic

chord and 35% of mean aerodynamic chord.

Location of different components from the nose:

Sl. No. Component Distance (m)

1 Instruments 1.80

2 Pilot 3.25

3 Nose Wheel 4.70

4 Crew 7.00

5 Passengers 23.00

6 Cargo Forward 12.00

7 Fuselage structure 24.30

8 Cargo Aft 33.00

9 Wing assembly 17.00

10 Toilets Front 23.00

11 Toilets Aft 36.40

12 Rear Crew 39.00

13 Galley 41.00

14 Horizontal Tail 46.00

15 Vertical Tail 44.00


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Front View


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Top View


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Side View

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Calculation of Static margin:

Static margin an important parameter that determines the stability of the

aircraft is calculated for various flight conditions and it is seen that static margin is

always positive, in other words neutral point is always ahead of center of gravity andthe aircraft is stable.

Balance Diagram:

Configuration XC.G(m) Static Margin

Full payload - Full fuel 23.646 0.1658

No payload - Full fuel 23.838 0.1282

Half payload - Full fuel 23.730 0.1494

Full payload - Reserve fuel 23.687 0.1577

No payload - Reserve fuel 23.974 0.1017

Half payload - Reserve fuel 23.806 0.1345

Full payload Half fuel 23.669 0.1613

No payload - Half fuel 23.912 0.1138

Half payload - Half fuel 23.772 0.1412


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Drag Estimation:

In the design of an aircraft, it is a crucial step to determine the drag of the

aircraft since it directly affects the power required and the performance is sensitive to

the drag of the aircraft. Drag due to all components is to be considered.

All parts of the aircraft contribute towards drag, which should be carefully

reduced by proper selection and design. From experience and experimental work

some approximation has been done for the calculation of fuselage drag and other

parts.

The various components of drag are

1. Parasite drag

2.

Induced drag3. Interference drag

4. Drag due to compressibility correction

CDtotal = CDo+ CDi+ CDint+ CDcompThe zero lift drag of the aircraft can be estimated from the formula,

wing

DDo

S

SCC

=

Sl.

No.Component S (m

2) CD CD * S (m

2)

1 Fuselage 16.75 0.0776 1.3003

2 Wing 166.53 0.0060 0.9992

3 Horizontal Tail 45.52 0.0063 0.2868

4 Vertical Tail 29.44 0.0059 0.1737

5 Power plant 2.52 0.0231 0.0583

6 Under carriage

a Main Wheel 0.31 0.6 0.186

b Nose Wheel 0.37 0.6 0.222

7 Flaps

a Take off (45o) 30 0.1 3

b Landing (60 o) 30 0.15 4.5


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CDo at various conditions

1. Cruise 0.0108

2. Take off 0.0214

3. Landing 0.0265

eARK

1=

Wing efficiency factor e = 0.9

Aspect ratio A.R =7.9

K = 0.0474

The Drag polar for the 3 conditions becomes

1. Cruise 0.0108 + 0.0474CL2

2. Take off 0.0214 + 0.0474CL2

3. Landing 0.0265 + 0.0474CL2

Variation of CLVs CD:

Take off:-

Sl. Velocity CL M KCL2

CD = CDo+ KCL2

1 0 0.00 0.00 0.00000 0.0214

2 100 0.95 0.30 0.0431 0.0644

3 120 0.66 0.36 0.0208 0.0421

4 140 0.48 0.42 0.0112 0.0326

5 160 0.37 0.48 0.0066 0.0279

6 180 0.29 0.55 0.0041 0.0255

7 200 0.24 0.61 0.0027 0.02418 220 0.20 0.67 0.0018 0.0232

9 240 0.17 0.73 0.0013 0.0227


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Cruise:-


CD = CDo+ KCL2

1 0 0.00 0.00 0.00000 0.01082 100 3.01 0.27 0.4291 0.4400

3 120 2.09 0.40 0.2070 0.2178

4 140 1.53 0.47 0.1117 0.1225

5 160 1.17 0.54 0.0655 0.0763

6 180 0.93 0.61 0.0409 0.0517

7 200 0.75 0.67 0.0268 0.0376

8 220 0.62 0.74 0.0183 0.0291

9 240 0.52 0.81 0.0129 0.0237

Landing:-


CD = CDo+ KCL2

1 0 0.00 0.00 0.0000 0.0265

2 100 0.95 0.30 0.0431 0.0696

3 120 0.66 0.36 0.0208 0.0473

4 140 0.49 0.42 0.0112 0.0377

5 160 0.37 0.48 0.0066 0.0331

6 180 0.29 0.55 0.0041 0.0306

7 200 0.24 0.61 0.0027 0.0292

8 220 0.19 0.67 0.0018 0.0283

9 240 0.16 0.73 0.0013 0.0278

CL

Vs CD

0.0000

0.2000

0.4000

0.6000

0.8000

1.0000

1.2000

0.0000 0.0200 0.0400 0.0600 0.0800

CD

CL

Take off

Landing

Cruise


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Performance curves:

The curves drawn between thrust to velocity and power to velocity are the

most important performance indicator graphs, as they give an indication of the thrust

required, thrust available, power required and power available. The rate of climb vsaltitude graph can be used to find the absolute and service ceilings.

The performance curves were drawn using the following criterion.

Weights:-

Full payload + 25% of fuel W1 = 775 kN



Altitudes:-

Z1 = 0 (S.L) Z2 = 5300 m Z3 = 10500 m

1 = 1.225 kgm-3 2 = 0.71 kgm-3 3 = 0.388 kgm-3

Rate of Climb Vs Altitude

0

5

10

15

20

25

30

35

0 5 10 15

Altitude (km)

R/C(m/s)

R/C

R/C ceil

Absolute ceiling 13,000 m

Service ceiling 12,800 m


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Thrust Vs Velocity

0

50000100000

150000

200000

250000

300000

350000

400000

0 200 400 600

Velocity (m/s)

T

hrust(N)

S.L+ W1

S.L W2

S.L W3

5300 W1

5300 W2

5300 W3

10500m W1

10500m W2

10500m W3

S.L Tav

5300m Tav

10500m Tav

Power Vs Velocity

0

20000000

40000000

60000000

80000000

100000000

120000000

140000000

160000000

180000000

0 200 400 600

Velocity (m/s)

Pow

er(W)

S.L W1

S.L W2

S.L W3

5300m W1

5300m W2

5300m W3

10500m W1

10500m W2

10500m W3

S.L Pav

5300m Pav

10500m Pav


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Take-Off distance:

Take Off is the first step for an aircraft to be really called as an aircraft.

Therefore the estimation of take-off distance is a major criterion, especially in the

case of a passenger aircraft. The following section classifies the take-off distance into

three stages namely, ground run transition and climb.

})]([{

W1.21

max

2

avgrL

grLWDTSCg

S+

=

gS tr

15.0

sinV2

lo =

Sground run = 1438.78 m

Stransition = 3.35 mSclimb = 155.63 m

Stotal = 1.59 km

Landing Distance:

})(5.1{

W1.69

max

2

avgrL

grLWDSCg

S+

=

gS

ap

f082.0

sinV2

ap =

ap

f

tan

h-h

=

apS

Landing is the crucial stage in an aircrafts flight envelope. The shorter the

landing distance, the better the flying quality. In order to achieve this, i.e., to

minimize the landing distance, we have planned to used thrust reversal and full flapdeflection during landing.

Sapproach = 286.66mSflare = 5.16 m

Sground run = 1690 m

Stotal = 1.98 km


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Minimum Turn Radius:

Rmin =

2)/(

4

1)/(

)/(4

WT

KC

WTg

SWK

Do

Rmin = 476.793 m

tmin =

tang

Vstall

tmin = 50.76 s

Stability and control:

An aircraft is actually a huge mass being suspended in the air medium. Hence

it needs to be perfectly stable and controllable. In this section stability characteristics

of this design are studies and its critical parameters are calculated. For better stability

and control some parameters are modified in an interactive process and satisfactory

results are arrived at.

Stability is studied under the following heads:

1) Longitudinal Stability

2) Lateral Stability

3) Directional Stability

For the stability calculations, the body axes are taken into consideration. The

pitch, roll and yaw axes are the x, y and z axes respectively. Since this is a

preliminary design aero elasticity is not considered and the aircraft is considered to be

a rigid body.


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Longitudinal Stability:

CMc.g Vs (Fixed)

-2

-1.5

-1

-0.5

0

0.5

1

1.5

2

-15 -10 -5 0 5 10 15

(o)

CMc.g

= -8

= -6

= -4

= -2

= 0

= 2

= 4

= 6

= 8

CMc.g Vs (Free)

-2

-1.5

-1

-0.5

0

0.5

1

1.5

2

-10 -5 0 5 10

(o )

CMc.g

= -8

= -6

= 4

= -2

= 0

= 2

= 4

= 6

= 8


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Directional Static Stability:

The directional stability defines the stability of the aircraft about the yaw axis.

It is the tendency of the aircraft to develop restoring moment when a disturbing isgiven in the form of a sideslip. It takes into account the rudder efficiency and the

destabilizing contributions from the wing, fuselage etc., and here the rudder power are

also calculated.

1) Contribution of wing (sweep back angle)

(Cn)wing = -0.0006 1/2 (= 250)

= -0.003

2) Contribution of wing(position)

Since our aircraft is a transport type aircraft, low wing configuration is taken

which does not contribute anything to directional stability.

(Cn)pos = 0

3) Contribution of fuselage

(Cn)fuselage =

2/1

1

2

2/1

2

1

3.5796.0

WW

hh

bL

SSK f

w

s

K= 0.1

SS = 154.425 m2

SW = 221.26 m2

LF = 44.62 m

b = 44.62 m

h1= h2= 3.52 mW1=W2= 5.73 m

Substituting these values,

(Cn)fuselage = 0.0012

4) Contribution of wing fuselage interaction

The contribution is due to the distributed flow over the rudder due to the

interaction of flow due to the fuselage and the wing. The value is calculated as

2Cn = -0.0003 per degree (from graph)


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5) Contribution of vertical tail

The stabilizing contribution of vertical tail is expressed as follows,

(Cn)v = vv

w

v

vb

l

S

Sa = 0.0064

The stability measure of an airplane is the sum of all the individual components.

(Cn) = (Cn)wing +(Cn)fuselage +(Cn)pos +2(Cn) +(Cn)v

Cn = 0.0043

CN Vs (Fixed)

-1.000

-0.500

0.000

0.500

1.000

-20.000 -10.000 0.000 10.000 20.000

(o)

CN

= -20

= -10

= 0

= 10

= 20

CNVs (Free)

-1.500

-1.000

-0.500

0.000

0.500

1.000

1.500

-20 -10 0 10 20

(o)

CN

= -20

= -10

= 0

= 10

= 20


27/29

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Lateral Static Stability:

The roll stability of the aircraft is its tendency to develop a restoring moment,

when disturbed from its equilibrium state of bank angle position. It is affected by the

influence of wings, its position, and fuselage and tail contributions.

1) Wing fuselage interference effect

For low wing configuration, (Cl) = -0.0008

2)Wing vertical tail interference effect

For low wing configuration (Cl) = 0.00016

3) Wing contribution

Dihedral effect:

(Cl) dihedral = Cl=0 + 0.0002 = 0.0008

Wing tip effect:

(Cl) wing tips = 0.0002

4) V.T. contribution

(Cl) V.T. = vv

w

v

vb

l

S

Sa = 0.00161

Cl = 0.00299The design criterion in common use for evaluating lateral control effectiveness

is the non-dimensional parameter pb/2V, with p, the rate of roll in rad/sec, b, the wing

span, and v, the true speed. For geometrically similar planes and lateral control

arrangements, this parameter is a constant and that for all airplanes the pb/2V that canbe produced by full lateral control deflection is a measure of relative lateral control

power available. The term pb/2V is actually the helix angle made by the wing tip

during the roll maneuver.

[ ]

++

=

3

)(3)1)((

3.57

22

1

2

2

3

1

3

2 kkkk

b

VP a


28/29

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Where,

= Aileron effectiveness =0.4(from graph),

= Maximum aileron deflection (30),

a

V = Rolling velocity, = Taper ratio of the wing,

k1 = Distance from the axis of the fuselage to

the starting point of the aileron,

k2 = Distance from the axis of the fuselage to

the ending point of the aileron.

Aileroron Power

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0 50 100 150 200 250 300 350

Velocity (ms-1

)

P(rads-1)

Helix Angle

0

0.005

0.01

0.015

0.02

0.025

0.03

0.035

0.04

0 50 100 150 200 250 300 350

Velocity (ms -1)

Pb/2V


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The V-n diagram:

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Aerodynamic deisgn report1