design of medium range business jet transport

8/14/2019 Design of Medium range business jet transport

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RAJALAKSHMI ENGINEERINGCOLLEGE

THANDALAM 602 105

BONAFIDE CERTIFICATE

This is to certify that this is a bonafide record of work done by the student

BALAJI.M, IV year Aeronautical Engineering in the

AI ! A"T #E$I%& ' (JE!T )II Laboratory during the year *+ *-*+ .

Signatu ! "# Fa$u%t&'in'C(a g!Su)*itt!+ #" t(! , a$ti$a% E-a*inati"n (!%+ "n ..........//

Int! na% E-a*in! E-t! na% E-a*in!

UNIVERSITY REGISTER No.


2/61

ACKNO LEDGEMENT

/e would like to e0tend our heartfelt thanks to 'rof. 1ogesh 2u3ar $inha

45ead of Aeronautical #e6art3ent7 for gi8ing us his able su66ort and

encourage3ent. At this 9uncture we 3ust e36hasis the 6oint that this design

6ro9ect would not ha8e been 6ossible without the highly infor3ati8e and 8aluable

guidance of Mr. $urendra Bogadi 4Assistant 6rofessor of Aeronautical

#e6art3ent7, whose 8ast knowledge and e06erience has greatly hel6ed us in this

6ro9ect. /e ha8e great 6leasure in e06ressing our sincere and whole hearted

gratitude to the3.

It is worth 3entioning about 3y friends and colleagues of the Aeronautical

de6art3ent for e0tending their kind hel6 whene8er the necessity arose. I thank

one and all who ha8e directly or indirectly hel6ed us in 3aking this design.


3/61

NOMENCLAT RE

A. . - As6ect atio

b - /ing $6an 437

! - !hord of the Airfoil 437

! root - !hord at oot 437

! ti6 - !hord at Ti6 437

C - Mean Aerodyna3ic !hord 437

!d - #rag !o-efficient

! d,+ - :ero Lift #rag !o-efficient

!6 - $6ecific fuel consu36tion 4lbs;h6;hr7

!L - Lift !o-efficient

# - #rag 4&7

E - Endurance 4hr7

e - (swald efficiency

L - Lift 4&7

4L;#7loiter - Lift-to-drag ratio at loiter 4L;#7cruise - Lift-to-drag ratio at cruise

M - Mach nu3ber of aircraft

Mff - Mission fuel fraction

- ange 4k37

e - eynolds &u3ber

$ - /ing Area 43


4/61

/crew - !rew weight 4kg7

/e36ty - E36ty weight of aircraft 4kg7

/fuel - /eight of fuel 4kg7

/6ayload - 'ayload of aircraft 4kg7/+ - (8erall weight of aircraft 4kg7

/;$ - /ing loading 4kg;3 cur8e for a hori?ontal tail

a-#istance of the front s6ar fro3 the nose of the aircraft

bw-/idth of the web

b f -/idth of the flange

I00 - $econd 3o3ent of area about @ a0is

I?? - $econd 3o3ent of area about : a0is

2 - %ust alle8iation factor

n 3a0 - Ma0i3u3 load factor

tw - Thickness of the web

tf - Thickness of the flange

T - Tor ue

- %ust 8elocity

V cruise - !ruise 8elocity V s - $talling 8elocity


5/61

AIM OF THE ,ROJECTThe ai3 of this 6ro9ect is to design the Mediu3 range business 9et trans6ort and

to satisfy the 3ission, o6eration and lu0ury re uire3ents of the aircraft

The following 3ission re uire3ents are to be satisfied by the aircraft

. 'ro8ide 3ediu3 range 6assenger trans6ort for a grou6 of 6eo6le as a

charted flight or the like.

*. To facilitate a busy business 3an to reach his destination in ti3e withlu0ury and co3fort.

. To 6ro8ide onboard facilities to a business 6erson to carry out his or her

work e8en during the 9ourney without any interru6tion.

C. To incor6orate all 3odern flight control syste3s, structural 3aterials etc

in order to 6ro8ide a safe 9ourney.

D. Most i36ortantly to 6ro8ide fuel efficiency to 3ake our aircraft a

co33ercial success for the o6erating airlines /


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ABSTRACT The aim of the project is to do a systematic design of a medium-rangebusiness jet capable of flying passengers through regional routes incorporated

with advanced avionics, structures and systems ,with high fuel efficiency so that ityields a good commercial profit to the airlines.

INTROD CTION

1/2 OBJECTI3ES

To 3eet the functional 6ur6ose and safety re uire3ents set out or acce6table to the

agricultural use.

1/2/1 ACT AL ,ROCESS OF DESIGN

$election of aircraft ty6e and sha6e

#eter3ination of geo3etric 6ara3eters

$election of 6ower 6lant

#eter3ination of aircraft flight and o6erational characteristics.

1/2/2 DISTINCT STAGES OF AIRCRAFT DESIGN 'ro9ect "easibility $tudy

'reli3inary #esign

#esign 'ro9ect

1/2/4 ,ROJECT FEASIBILIT ST D t" !7"%7! a 8ati8#a$t" &

89!$i#i$ati"n: !o36rehensi8e 3arket sur8ey

$tudies on o6erating conditions for the air6lane to be designed


7/61

$tudies on rele8ant design re uire3ents 4s6ecified by Airworthiness Authorities7

E8aluation of si3ilar e0isting designs

$tudies on 6ossibilities of introducing new conce6ts

!ollection of data on rele8ant 6ower 6lants Laying down ' ELIMI&A 1 $'E!I"I!ATI(&$

1/2/; ,RELIMINAR DESIGNIt consists of the initial stages of design, resulting in the 6resentation of a B (!5 E

!ontaining 6reli3inary drawings and clearly stating the o6erational ca6abilities of the air6lane

being designed. This Brochure has to be A'' (VE# by the 3anufacturer and;or the

custo3er.

1/2/5 T(! 8t!98 in7"%7!+

The $tructural design in8ol8es

#eter3ination of loads acting on aircraft

V-n diagra3 for the design study

%ust and 3anoeu8rability en8elo6es

$chrenkFs !ur8e

!ritical loading 6erfor3ance and final V-n gra6h calculation

#eter3ination of loads acting on indi8idual structures

$tructural design study ) Theory a66roach


8/61

Load esti3ation of wings

Load esti3ation of fuselage.

Material $election for structural 3e3bers

#etailed structural layouts

#esign of so3e co36onents of wings, fuselage

1/2 ,ARAMETERS DERI3ED FROM AD, 1

1/2/1 DESIGN ,ARAMETERS

S/NO DESIGN DATA DESIGN 3AL ES

cruise s6eed G + k3;hr

* ole Mediu3 range business 9et

!rew *

C Length D 3

D /ing s6an G.H 3

H 5eight D. 3

E36ty weight G+++ kg

G Ma0 takeoff weight CDD+ kg

6ower 6lant *

+ /ing area CH.DG* s .3

ange DD++ k3

* "uel weight DD* .*DC kg

wing loading *. D kg;s .3

C Engine ty6e Turbo fan

D Engine Thrust .H k& 4 engine7


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1/2/2 ,ERFORMANCE ,ARAMETERS

Ta)%! 1'2 ,! #" *an$! ,a a*!t! 8

S/NO ,ERFORMANCE DATA ,ERFORMANCE 3AL ES

/ing loading /;$ *. D kg;s .3

* /ing area $ CH.DG* s .3

:ero-lift to drag coefficient ! #o +.+ G

C #rag-due to )lift coefficient 2 +.+ C

D As6ect ratio A D. *H

H Thrust 4one engine7 .H k&

AIRFOIL S,ECIFICATIONS

Thickness

Kc

!a3ber

.GKc

Trailing edge angle

+.* deg

Lower flatness

H .DKc

Leading edge radius

.DKc

Ma0 ! L

.GCD


10/61

Ma0 ! L angleD.+ deg

Ma0 L;#

HG.D *

Ma0 L;# angle

D deg

Ma0 L;# ! L

. C

$tall angle

.+ deg

:ero-lift angle

-H.+ deg

1.3 THREE VIEW DIAGRAM OF MEDIUM RANGE BUSINESS JET:


11/61

THE 3'n DIAGRAM1/ 3 n Diag a*

"light regi3e of any aircraft includes all 6er3issible co3binations of s6eeds, altitudes,

weights, centres of gra8ity, and configurations. This regi3e is sha6ed by aerodyna3ics,

6ro6ulsion, structure, and dyna3ics of aircraft. The borders of this flight regi3e are called

flight en8elo6e or 3anoeu8ring en8elo6e. The safety of hu3an onboard is guaranteed by

aircraft designer and 3anufacturer. 'ilots are always trained and warned through flight

instruction 3anual not to fly out of flight en8elo6e, since the aircraft is not stable, or not

controllable or not structurally strong enough outside the boundaries of flight en8elo6e. A

3isha6 or crash is e06ected, if an aircraft is flown outside flight en8elo6e.

The flight en8elo6e has 8arious ty6es each of which is usually the allowable 8ariations of one

flight 6ara3eter 8ersus another 6ara3eter. These en8elo6es are calculated and 6lotted by flight

3echanics engineers and e36loyed by 6ilots and flight crews. "or instance, the load 3asters of

a cargo aircraft 3ust 6ay e0tra caution to the center of gra8ity location whene8er they

distribute 8arious loads on the aircraft. There are se8eral crashes and 3isha6s that safety

board s re6ort indicated that load 3aster are res6onsible, since they de6loyed 3ore loads than

allowed, or 3is6laced the load before take-off. &ose hea8y and tail hea8y are two flight

conce6ts that 6ilots are fa3iliar and e06erienced with, and are trained to deal with the3 safely.

'ilots are using se8eral gra6hs and charts in their flight o6erations. "our i36ortant en8elo6es

are as follows


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1/ #iagra3 of 8ariations of aircraft lift coefficient 8ersus Mach nu3ber 4!L ) M7

2/ #iagra3 of 8ariations of airs6eed 8ersus altitude 4V ) h7

4/ #iagra3 of 8ariations of centre of gra8ity 8ersus aircraft weight 4@cg ) /7

;/ #iagra3 of 8ariations of airs6eed 8ersus load factor 4V ) n7

(ne of the 3ost i36ortant diagra3s is referred to as flight envelope . This en8elo6e

de3onstrates the 8ariations of airs6eed 8ersus load factor 4V ) n7. In another word, it de6icts

the aircraft li3it load factor as a function of airs6eed. (ne of the 6ri3ary reasons that this

diagra3 is highly i36ortant is that, the 3a0i3u3 load factor that is e0tracted fro3 this gra6h

is a reference nu3ber in aircraft structural design. If the 3a0i3u3 load factor is under-

calculated, the aircraft cannot withstand flight load safely. "or this reason, it is reco33ended

to structural engineers to recalculate the V-n diagra3 on their own as a safety factor.

In this section, details of the techni ue to 6lot the V- n diagra3 in introduced. The "igure

shows a ty6ical V-n diagra3 for a %A aircraft. This diagra3 is, in fact, a co3bination of two

diagra3s

1/ T(! 3'n +iag a*


13/61

N" Ai $ a#t t&9! Ma-i*u* 9"8iti7! %"a+#a$t" Ma-i*u* n!gati7! %"a+#a$t" &or3al 4non-

acrobatic7*.D ) .G - to - .D

* tility 4se3i-acrobatic C.C - .G Acrobatic H -

C 5o3ebuilt D -*D Trans6ort ) C - to -*H 5ighly 3aneu8erable H.D ) * - to -H

Bo3ber * ) C - to -*

2/2' 3 n Diag a*


14/61

The to6 cur8e is literally a 6lot of abo8e e uation the region abo8e this cur8e in the V-n

diagra3 is the stall area. $ince, no aircraft can fly continuously at a flight condition abo8e this

cur8e, so this is one of the li3its on the aircraft 3aneu8erability. Because the aircraft angle ofattack will be abo8e stall angle. Based on the e uation, as the airs6eed increases, the 3a0i3u3

load factor will increase 6ro6ortionally to V*. 5owe8er, n3a0 cannot be allowed to increase

indefinitely. It is constrained by the structural strength 4structural li3it load factor7. The to6

hori?ontal line denotes the 6ositi8e li3it load factor in the V-n diagra3.

The flight 8elocity corres6onding to the intersection between the left cur8e and to6 hori?ontal

line 4'oint A7 is referred to as corner velocity , and designated as VP 4V star7. The corner

8elocity can be obtained by sol8ing e uation for 8elocity, yielding

''''''''''''''''''''' 4

where the 8alue of n3a0 corres6onds to that at 6oint A . This s6eed so3eti3es is referred to as

maneuvering speed 4VA7, and is su33ari?ed as

'''''''''''''''''' ;

The 6oint A is then called the manoeuvre point . At this 6oint, both lift coefficient and load

factor are si3ultaneously at their highest 6ossible 8alues. The corner 8elocity is an interesting

8elocity for fighter 6ilots. At s6eeds less than VP, it is not 6ossible to structurally da3age the

aircraft due to generation of load factor less than n3a0. 5owe8er, the bank angle is not high

enough for a tight turn. In contrast, at s6eeds greater than VP, 3anoeu8rability decreases, since

the s6eed is too high.

The right hand side of the V ) n diagra3, 8ertical line B!, is a high s6eed li3it. This s6eed is

usually selected to be the di8e s6eed. At flight s6eeds higher than this li3it, the dyna3ic

6ressure 4 7 is higher than the design 8alue for the aircraft. At the s6eed abo8e di8e s6eed,

destructi8e 6heno3ena such as flutter, aileron re8ersal, and wing di8ergence, 3ay ha66en that

leads structural da3age, or failure, or disintegration. This s6eed li3it 4di8e s6eed7 is a red-line

s6eed for the aircraft it should ne8er be e0ceeded. The di8e s6eed 4V#7 is usually higher than


15/61

aircraft 3a0i3u3 s6eed 4V3a07, and the aircraft 3a0i3u3 s6eed 4V3a07 is often higher than

aircraft cruising s6eed 4V!7. "ro3 "A 'art * , the following regulations ha8e been directly

co6ied

----------- 5

The botto3 line of the V ) n diagra3, gi8en by hori?ontal line !# corres6onds with 3a0i3u3

negati8e li3it load factor that is a structural li3it when the aircraft is in a situation such as

in8erted flight. The botto3 left cur8e corres6onds to negati8e stall angle of attack. $ince 3ost

wing airfoils ha8e 6ositi8e ca3ber, their 6ositi8e stall angles are often 3uch higher than the

absolute 8alues of their negati8e stall angles. This cur8e defines the negati8e stall area.

2/4' CALC LATIONS

"ro3 table,

nQ8eOC and n-8eO -*

2/4/1 ,OSITI3E C R3E

"ro3 e u, C

Maneuvering speed 3A:

VA O . 3;s

"ro3 e u, D

Di7! 89!!+ 3 D:

V# O **. 3;s

Then,

LO+.DR .**DRCH. CDR .GCDR


16/61

Fro3 that we get

nO4 DG* .G+7;4 C* D.D7

>4/;?56 ''''''''''''''''''''' 6

In e u H,by substituting 8arious 8 u6to we get cur8e

2/4/2 NEGATI3E C R3E

"ro3 e u, C

Maneuvering speed 3A:

VAOHG.GC3;s

"ro3 e u, D

LO HCC .C &

And

nO-+. D*

In e u ,by substituting 8arious 8 u6to we get cur8e.

4Insert V-n 3aneu8er diagra37

4/ Gu8t En7!%"9!

4/1D!8$ i9ti"n

The at3os6here is a dyna3ic syste3 that enco36asses 8ariety of 6heno3ena. $o3e

of these 6heno3ena include turbulence, gust, wind shear, 9et strea3, 3ountain wa8e andther3al flow. In this section, we concentrate on only gust, since it is not 6redictable, but is

ha66ening during 3ost high altitude flights. /hen an aircraft e06eriences a gust, the

i33ediate effect is an increase or decrease in the angle of attack. . /hen an u6ward gust with

a 8elocity of Vg, hits under the nose of an aircraft with the 8elocity of V, the instantaneous

change 4increase7 in the angle of attack 4 >7


17/61

This indicates that gust will change load factor and will generate a load called gust load. The

loads e06erienced when an aircraft encounters a strong gust 3ay so3eti3es e0ceed the

3aneu8er load. Thus we 3ust 6ay attention to gust load when 6lotting V-n diagra3. As soon

as we know the gust 8elocity, we are able to deter3ine gust load. It is 8ery hard to 3easure

gust 8elocity, since it ha66ens suddenly. The design re uire3ents for gust 8elocities are

e0tracted fro3 flight test data.

eference %ust Velocity 4 ref 7 Sat sea le8el D3;s.

#esign %ust Velocity 4 ds7 S ref R 2

4/2 C"n8t u$ti"n

The increase in the load factor due to the gust can be calculated by

"or cur8e abo8e V-a0is

/here,

K -%ust Alle8iation "actor

3a0 -Ma0i3u3 deri8ed %ust Velocity

a - Lift !ur8e $lo6e for wing

"or cur8e below V-a0is

Gust Alleviation Factor (K):

Lateral Mass Ratio ():

/here,


18/61

g -Acceleration due to %ra8ity

@-Mean Aerodyna3ic !hord

O

$t -!hord at ti6$ -!hord at rootwhere,cr O . *3c t O .HH3

O +H.HG

2O+. H

"or cur8e below V-a0is

O .HGP +-CPuP8

By using the e uations and for 8arious s6eeds of 3a0 we get the following gust lines

load factors at the 8arious 6oints can be found using the for3ula using the corres6onding

8alues of 3a0


19/61

;/SCHRENK S C R3E

x elliptic

01773.7

1 17!3.!!3

"1733.0

#$

31!$0.7

$

#1!0#.7

!11 01.3

!%


20/61

!13!#.#

0"

711$"."

#$

$%"$.!$

%

%0 .$1

#$

%.3#173."3

0"

xschren&

011%33.

1113$".

#!


21/61

"10$10.

%

310"17.

!%

#%!00.7

11$% !.3

!%

!$"7$.#

"

77 .3

#$

$!7!0.$

3%

%7%7.0

1

%.3#"77.$

0"


22/61

;/1 D!8$ i9ti"n

Lift 8aries along the wing s6an due to the 8ariation in chord length, angle of attack andswee6 along the s6an. $chrenkFs cur8e defines this lift distribution o8er the wing s6an of anaircraft, also called si36ly as Lift #istribution !ur8e.$chrenkFs !ur8e is gi8en by

/herey is Linear Variation of lift along se3i wing s6an also na3ed as Ly* is Elli6tic Lift #istribution along the wing s6an also na3ed as L*

;/2 Lin!a Li#t Di8t i)uti"n

Lift at root

Lroot O *+ .D &;3

Lift at ti6

Lti6 O + D . D &;3

By re6resenting this lift at sections of root and ti6 we can get the e uation for the wing.

E uation of linear lift distribution for starboard wing

y O- + G . 0Q*+ .D

E uation of linear lift distribution for 6ort wing we ha8e to re6lace 0 by )0 in general,

y O +G . 0Q*+ .D

x linear

0101 %.

71 %!1$.$

"%077.$

3 $ 3!.%


23/61

#7%% .%

7#

!!%1#.0

7 !373.1$

$3".1

% "%1."

%.3#10#.

7"

;/4 E%%i9ti$ Li#t Di8t i)uti"n

Twice the area under the cur8e or line will gi8e the lift which will be re uired to o8erco3eweight!onsidering an elli6tic lift distribution we get


24/61

/here b is Actual lift at rootAnd a- is wing se3i s6an Lift at ti6

b O H . &;3

E uation of elli6tic lift distribution,

y*O .G4GG. -0*7+.D

then

y*;*O GG. 4GG. -0*7+.D

x elliptic

01773.7

117!3.!

!3

"1733.0

#$

31!$0.7

$

#1!0#.7

!11 01.3

!%

!13!#.#

0"

711$"."

#$

$%"$.!$

%

%0 .$1

#$

%.3#173."3

0"


25/61

;/; C"n8t u$ti"n "# S$( !n 8 Cu 7!

$chrenkFs !ur8e is gi8en by

yO-DC+. D0Q + D . DQ GG. 4GG. -0*7+.D

$ubstituting different 8alues for 0 we can get the lift distribution for the wing se3i s6an


26/61

/e get the following schrenkFs distribution for lift.

xschren&

011%33.

1 113$".#!

"10$10.

%

310"17.

!%

#%!00.7

11$% !.3

!%

!$"7$.#

"

7 7 .3#$

$!7!0.$

3%

%7%7.0

1

%.3#"77.$

0"

5/ LOAD ESTIMATION ON ING

5/1 D!8$ i9ti"n


27/61

The solution 3ethods which follow EulerFs bea3 bending theory 4U;yOM;IOE; 7 usethe bending 3o3ent 8alues to deter3ine the stresses de8elo6ed at a 6articular section of the

bea3 due to the co3bination of aerodyna3ic and structural loads in the trans8erse direction.Most engineering solution 3ethods for structural 3echanics 6roble3s 4both e0act anda66ro0i3ate 3ethods7 use the shear force and bending 3o3ent e uations to deter3ine thedeflection and slo6e at a 6articular section of the bea3. Therefore, these e uations are to beobtained as analytical e06ressions in ter3s of s6an wise location. The bending 3o3ent

6roduced here is about the longitudinal 407 a0is.

5/2 L"a+8 a$ting "n


28/61

/ /I&% O+. *D P CDD+P .G O +. &

/ /I&% O +. &

/ 6ortwing O G GD. 4Acting #ownwards7

/ starboard OG GD. 4Acting #ownwards7

Assu3ing 6arabolic weight distribution

/here b ) wing s6an

/hen we integrate fro3 0O+ 4root location7 to 0Ob 4ti6 location7 we get the net weight of 6ortwing.

y O-C. *40- . CD7*

$ubstituting 8arious 8alues of 0 in the abo8e e uation we get the self-weight of the wing

0 self weight

+ - D . DDG

-*GH. * G

* -***.*H G- HD.GH G

C - . +C G

D - . G DG

H -CH.+ G G

-**.HDD G


29/61

G - .CD G

-+.C + G

. CD +

5/4/2 Fu!%

Again by using general for3ula for straight line yO30 Q c we get,

yf ODD* .*D2g

0 self weight fuel dist.+ - D . DDG - +*.-*GH. * G -*HD.HCG

* -***.*H G -**G. GH- HD.GH G - . *C

C - . +C G - D .GH*D - . G DG - H.HH -CH.+ G G - . G

-**.HDD G -C*.+ HG - .CD G -C.G C


30/61

-+.C + G *.CCG. CD + CD. +

Ta)%! 5'1 L"a+8 8i*9%i#i!+ a8 9"int %"a+8

Cu 7! $"*9"n!nt A !a !n$%"8!+ 8t u$tu a%

y ;* C C*.GH C. C

y*;* HD D.** +.

/ing G GD. .D+*

"uel * . H .C*

5/4/4 R!a$ti"n #" $! an+ B!n+ing *"*!nt $a%$u%ati"n8


31/61

Then,

C C*.GHQ HD D.**-G GD. -* . H-VA O+

VAO DC* . &

Then,MA Q4 + .CCP .C7Q4**G . GP*. 7-4 * H.+ PC.+ 7-4 * G .GPC. G7 O +

MAO **H++.* &3 &ow we know V A and M A, using this we can find out shear force and Bending 3o3ent.

5/4/; SHEAR FORCE

By using the corres6onding 8alues of 0 in a66ro6riate e uations we get the 6lot of shear force.

@ bc cd da+ - DC* .

-HC+C*.D -HC *H.* - D HCC.H

-C CC*.*C - C.GD -*C .H - HDD . - GGD.

- * H.GH -*DC*. G

G CH . C+C . D

. CD H*H+.HH+ - DC* .- -HC+C*.D -HC *H.-* -D HCC.H- -C CC*.*-C - C.G


32/61

-D -*C .-H - HDD . - GGD.- - * H.GH -*DC*. G-G CH . C- +C . D- . CD H*H+.HH

5/4/5 B!n+ing *"*!nt

By substituting the 8alues of 0 for the abo8e e uations of bending 3o3ents obtained we canget a continuous bending 3o3ent cur8e for the 6ort wing.

0 bc cd da+ C*HG .


33/61

DH*D+.C

DH +D.*

* * * DG.

* HCC .

C GH GD C C++.

HH +DG*D.

CH .

CDC. C++DD.DDG D C.DGC

-*CC .G. CD -C+ CH.H

+ C*HG .

- DH*D+.C

DH +D.*

-* * * DG.

- * HCC .

-C GH G-D C C++.

H-H +DG*D.

CH .

- CDC. C++DD.DD-G D C.DGC- -*CC .G- . CD -C+ CH.H


34/61

5/; S(!a #" $! an+ )!n+ing *"*!nt +iag a*8 +u! t" %"a+8 a%"ng $(" +


35/61

" OD C.* &;3!hordwise force at ti6

" TO* . &;34sa3e chord throught out the wing7

By using y O 30 Qc again we get the e uation as1O- . 0QD C.*

0 load-chord

+ D C.*

DH*.C

* D +.HH

C G.G

C CH . *

D C D. D

H C+ .DG

.G

G C+.+C

+G.*

. CD * . + C


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The abo8e e uation gi8es the 6rofile of load acting chordwise, by integrating this abo8ee uation we get a co36onent of $hear force and again by integrating the sa3e we get theco36onent of Bending Mo3ent

y d0 O - D.GG0 *QD C.*

0 shear

+ - DC* .

- CGC .C

* - C* H.GD

- G*.+D

C - * .+

D - *GC .

H - *C*G.*

- *+C+.CD

G - HGC.CD

- H+.*

. CD - *DD. D *


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y d0 * O -D.* 0 Q* . 0*

To find fi0ing 3o3ent and the reaction force,

, Then,VAO DC* .

,

Then,MAO **H++.*

Tor ue due to nor3al forces and constant 6itching 3o3ent at cruise condition


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D!t! *inati"n "# 7a i"u8 $"*9"n!nt8 "# t" u!

D!t! *inati"n "# 7a i"u8 $"*9"n!nt8 $au8ing t" u!

5/5 T" u! at $ iti$a% #%ig(t $"n+iti"n

5/5/1 T" u! +u! t" n" *a% #" $!

T O H*.Hc*

/here

c - chord

the e uation for chord can also be re6resented in ter3s of 0 by taking cO 30 Qk,


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cO-+. H0Q . *

Therefore tor ue

T O .H 0Q .GD0- C. 0*

0 tor ue-nf

+ +

H G.H*

* * .CH

CDHH.DC

C D DD.GG

D H D.D

H H D.C*

GCHD.HH

G H.*C

HD . G

. CD G*+. GC


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5/5/2 T" u! +u! t" $(" +


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- GH.D

G - +DD.GC

- +C C.G

. CD - +H +H. D

0 tor ue-su3

+ +

- H H+.C

* - CD.

-CD+DD.HD

C -DH GG.GC

D -H +CC.HD

H - D **.*H

-G D*+.GD

G -G .H


42/61


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6/ F SELAGE DESIGN

6/1 DESCRI,TION

To find out the loads and their distribution, consider the different cases. The 3ain co36onents

of the fuselage loading diagra3 are /eight of the fuselage

Engine weight

/eight of the hori?ontal and 8ertical stabili?ers

Tail lift

/eight of crew, 6ayload and landing gear

$yste3s, e ui63ent, accessories

$y33etric flight condition, steady and le8el flight 4#ownward forces negati8e7 Values for thedifferent co36onent weights are obtained fro3 aerodyna3ic design calculation Ta)%! 6'1L"a+8 a$ting "n Fu8!%ag!

C"*9"n!ntDi8tan$! # "* !#! !n$! %in!

*:*a88

g:

N:*"*!nt N*:

Engine .* * GG* CHC.*

G *C+ . &ose landing gear *. **+ * DG.* C H .GH

Load due to wing and fuel H.H + . +

CH G.

C* ++ . Dcrew C + GH . CDD

fuselage 3ass H.D C +C+*G .H

*H GG*.G5ori?ontal stabili?er *. G+ HD.G * . CMain landing gear H.H HG+ HH +.G CC+* .Vertical stabili?er *. D D. D CC.+*


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Ta)%! 6'2 S(!a #" $! an+ )!n+ing *"*!nt ta)u%ati"n

Distance from reference line(m)

mass(kg)

weight(N) SF (N) moment(Nm)

0 0 0 0 02.3 220 -2158.2 -2158.2 - !"3.8"

1!0 -18"3.! - 022.1 -# 55."".3 "!0!.10 "###8.31 "3#5".21 2#003.".5 10# - 028!.# 23 "".51 -2"1883"." "80 -""#0.8 1"#5!.#1 - 02#.3

11.2 2188 -21 " .3 - ""8.5! -2 0 0012.3 180 -1#"5.8 -" 3 .3! -21#1!.312.# #5 -#35.#5 -#1#0.1 -!3 .03

15 0 0 0 0

SF (N)

-20000

-10000

0

10000

20000

30000

0000

50000

"0000

#0000

0 2 " 8 10 12 1 1"

SF (N)


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moment(Nm)

-300000

-200000

-100000

0

100000

200000

300000

00000

500000

0 2 " 8 10 12 1 1"

moment(Nm)


46/61

/ MATERIAL SELECTION


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/1 DESCRI,TION

Aircraft structures are basically unidirectional. This 3eans that one di3ension, the

length, is 3uch larger than the others - width or height. "or e0a36le, the s6an of the wing and

tail s6ars is 3uch longer than their width and de6th the ribs ha8e a 3uch larger chord length

than height and;or width a whole wing has a s6an that is larger than its chords or thickness

and the fuselage is 3uch longer than it is wide or high. E8en a 6ro6eller has a dia3eter 3uch

larger than its blade width and thickness, etc.... "or this si36le reason, a designer chooses to

use unidirectional 3aterial when designing for an efficient strength to weight structure.

nidirectional 3aterials are basically co36osed of thin, relati8ely fle0ible, long fiberswhich are 8ery strong in tension 4like a thread, a ro6e, a stranded steel wire cable, etc.7

An aircraft structure is also 8ery close to a symmetrical structure. Those 3ean the u6 and down

loads are al3ost e ual to each other. The tail loads 3ay be down or u6 de6ending on the 6ilot

raising or di66ing the nose of the aircraft by 6ulling or 6ushing the 6itch control the rudder

3ay be deflected to the right as well as to the left 4side loads on the fuselage7. The gusts hitting

the wing 3ay be 6ositi8e or negati8e, gi8ing the u6 or down loads which the occu6ant

e06eriences by being 6ushed down in the seat or hanging in the belt.

Because of these factors, the designer has to use a structural 3aterial that can withstand

both tension and co36ression. nidirectional fibers 3ay be e0cellent in tension, but due to

their s3all cross section, they ha8e 8ery little inertia 4we will e06lain inertia another ti3e7 and

cannot take 3uch co36ression. They will esca6e the load by bucking away. As in the

illustration, you cannot load a string, or wire, or chain in co36ression.


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In order to 3ake thin fibers strong in co36ression, they are Nglued togetherN with so3e

kind of an Ne3beddingN. In this way we can take ad8antage of their tension strength and are no

longer 6enali?ed by their indi8idual co36ression weakness because, as a whole, they beco3e

co36ression resistant as they hel6 each other to not buckle away. The e3bedding is usually a

lighter, softer NresinN holding the fibers together and enabling the3 to take the re uired

co36ression loads. This is a 8ery good structural 3aterial.

/2 OOD5istorically, wood has been used as the first unidirectional structural raw 3aterial.

They ha8e to be tall and straight and their wood 3ust be strong and light. The dark bands 4late

wood7 contain 3any fibers, whereas the light bands 4early wood7 contain 3uch 3ore NresinN.

Thus the wider the dark bands, the stronger and hea8ier the wood. If the dark bands are 8ery

narrow and the light bands uite wide, the wood is light but not 8ery strong. To get the 3ost

efficient strength to weight ratio for wood we need a definite nu3bers of bands 6er inch.

$o3e of our aircraft structures are two-di3ensional 4length and width are large with

res6ect to thickness7. 'lywood is often used for such structures. $e8eral thin boards 4foils7 are

glued together so that the fibers of the 8arious layers cross o8er at different angles 4usually +

degrees today years back you could get the3 at + and CD degrees as well7. 'lywood 3akes

e0cellent Nshear websN if the designer knows how to use 6lywood efficiently. 4/e will learn

the basis of stress analysis so3eti3e later.7

Today good aircraft wood is 8ery hard to co3e by. Instead of using one good board for

our s6ars, we ha8e to use la3inations because large 6ieces of wood are 6ractically una8ailable,

and we no longer can trust the wood uality. "ro3 an a8ailability 6oint of 8iew, we si36ly

need a substitute for what nature has su66lied us with until now.

/4 AL MIN M ALLO S

$o, since wood 3ay not be as a8ailable as it was before, we look at another 3aterial

which is strong, light and easily available at a reasonable price 4there s no 6oint in discussing


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Titaniu3 - it s si36ly too e06ensi8e7. Aluminum alloys are certainly one answer. /e will

discuss the 6ro6erties of those alloys which are used in light 6lane construction in 3ore detail

later. "or the ti3e being we will look at alu3inu3 as a construction 3aterial.

E-t u+!+ A%u*inu* A%%"&8

#ue to the 3anufacturing 6rocess for alu3inu3 we get a unidirectional 3aterial uite a

bit stronger in the lengthwise direction than across. And e8en better, it is not only strong in

tension but also in co36ression. !o36aring e0trusions to wood, the tension and co36ression

characteristics are 6ractically the sa3e for alu3inu3 alloys so that the linear stress analysis

a66lies. /ood, on the other hand, has a tensile strength about twice as great as its co36ression

strength accordingly, s6ecial stress analysis 3ethods 3ust be used and a good understanding

of wood under stress is essential if stress concentrations are to be a8oidedX

Alu3inu3 alloys, in thin sheets 4.+ H to . *D of an inch7 6ro8ide an e0cellent twodi3ensional 3aterial used e0tensi8ely as shear webs - with or without stiffeners - and also as

tension;co36ression 3e3bers when suitably for3ed 4bent7.

It is worthwhile to re3e3ber that alu3inu3 is an artificial 3etal. There is no alu3inu3

ore in nature. Alu3inu3 is 3anufactured by a66lying electric 6ower to bau0ite 4alu3inu3

o0ide7 to obtain the 3etal, which is then 3i0ed with 8arious strength-gi8ing additi8es. 4In a

later article, we will see which additi8es are used, and why and how we can increase

alu3inu3 s strength by cold work hardening or by te36ering.7 All the co33only used

alu3inu3 alloys are a8ailable fro3 the shelf of dealers. /hen re uested with the 6urchase,

you can obtain a N3ill test re6ortN that guarantees the che3ical and 6hysical 6ro6erties as

tested to acce6ted s6ecifications.

As a rule of thu3b, alu3inu3 is three ti3es hea8ier, but also three ti3es stronger than

wood. $teel is again three ti3es hea8ier and stronger than alu3inu3.

/; STEEL

The ne0t 3aterial to be considered for aircraft structure will thus be steel, which has the

sa3e weight-to-strength ratio of wood or alu3inu3.

A6art fro3 3ild steel which is used for brackets needing little strength, we are 3ainly using a

chro3e-3olybdenu3 alloy called AI$I C (& or C C+. The co33on raw 3aterials a8ailable

are tubes and sheet 3etal. $teel, due to its high density, is not used as shear webs like

alu3inu3 sheets or 6lywood. /here we would need, say. ++N 6lywood, a .+ * inch alu3inu3


50/61

sheet would be re uired, but only a .+ + steel sheet would be re uired, which is 9ust too thin to

handle with any ho6e of a nice finish. That is why a steel fuselage uses tubes also as diagonals

to carry the shear in co36ression or tension and the whole structure is then co8ered with fabric

4light weight7 to gi8e it the re uired aerodyna3ic sha6e or desired look. It 3ust be noted that

this 3ethod in8ol8es two techni ues steel work and fabric co8ering.

/e will be discussing tubes and welded steel structures in 3ore detail later and go now to

Nartificial woodN or co36osite structures.

/5 COM,OSITE MATERIALS

The designer of co36osite aircraft si36ly uses fibers in the desired direction e0actlywhere and in the a3ount re uired. The fibers are e3bedded in resin to hold the3 in 6lace and

6ro8ide the re uired su66ort against buckling. Instead of 6lywood or sheet 3etal which allows

single cur8ature only, the co36osite designer uses cloth where the fibers are laid in two

directions .4the wo8en thread and weft7 also e3bedded in resin. This has the ad8antage of

freedo3 of sha6e in double cur8ature as re uired by o6ti3u3 aerodyna3ic sha6es and for 8ery

a66ealing look 4i36ortance of esthetics7.

Today s fibers 4glass, nylon, 2e8lar, carbon, whiskers or single crystal fibers of 8arious

che3ical co36ositions7 are 8ery strong, thus the structure beco3es 8ery light. The drawback is

8ery little stiffness. The structure needs stiffening which is achie8ed either by the usual discreet

stiffeners, -or 3ore elegantly with a sandwich structure two layers of thin uni- or bi-directional

fibers are held a6art by a lightweight core 4foa3 or Nhoneyco3bN7. This allows the designer to

achie8e the re uired inertia or stiffness.

"ro3 an engineering stand6oint, this 3ethod is 8ery attracti8e and su66orted by 3any

authorities because it allows new de8elo63ents which are re uired in case of war. But this

3ethod also has its drawbacks for ho3ebuilding A 3old is needed, and 8ery strict uality

control is a 3ust for the right a3ount of fibers and resin and for good adhesion between both to

6re8ent too NdryN or NwetN a structure. Also the curing of the resin is uite sensiti8e to

te36erature, hu3idity and 6ressure. "inally, the resins are acti8e che3icals which will not only

6roduce the well-known allergies but also the che3icals that attack our body 4es6ecially the


51/61

eyes and lungs7 and they ha8e the unfortunate 6ro6erty of being cu3ulati8ely da3aging and

the result 4in 6articular deterioration of the eye7 shows u6 only years after initial contact.

Another disad8antage of the resins is their li3ited shelf life, i.e., if the resin is not used

within the s6ecified ti3e la6se after 3anufacturing, the results 3ay be unsatisfactory and

unsafe.

/6 HEA3 AIRCRAFT RA MATERIALS

The focus of our article is our Table which gi8es t&9i$a% 7a%u!8for a 8ariety of raw

3aterials.

!olu3n lists the standard 3aterials which are easily a8ailable at a reasonable cost. $o3e of

the 3aterials that fall along the borderline between 6ractical and i36ractical are

1/ Magn!8iu* An e06ensi8e 3aterial. !astings are the only readily a8ailable for3s. $6ecial 6recaution 3ust be taken when 3achining 3agnesiu3 because this 3etal burns when hot.

2/ Titaniu* A 8ery e06ensi8e 3aterial. Very tough and difficult to 3achine.

4/ Ca )"n Fi)! 8 $till 8ery e06ensi8e 3aterials.

;/ K!7%a Fi)! 8 Very e06ensi8e and also critical to work with because it is hard to NsoakN in

the resin.

A nu3ber of 6ro6erties are i36ortant to the selection of 3aterials for an aircraft structure. The

selection of the best 3aterial de6ends u6on the a66lication. "actors to be considered include

yield and ulti3ate strength, stiffness, density, fracture toughness, fatigue, crack resistance,

te36erature li3its, 6roducibility, re6airability, cost and a8ailability. The gust loads, landing

i36act and 8ibrations of the engine and 6ro6eller cause fatigue failure which is the single 3ost

co33on cause of aircraft 3aterial failure.

"or 3ost aeros6ace 3aterials, cree6 is a 6roble3 only at the ele8ated te36erature. 5owe8er

so3e titaniu3 6lastics and co36osites will e0hibit cree6 at roo3 te36eratures.

Taking all the abo8e factors into considerations, the alu3inu3 alloy which has e0cellent

strength to weight ratio and abundant in nature.

The following list of alu3iniu3 alloys is considered.

S/N" A%u*iniu* A%%"& i!%+ 8t !ngt(M,a

%ti*at! 8t !ngt(M,a

Al *+*C- T D *G+ C +

* Al *+*C- T * H C*


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Al + D- TH C H D G

C Al + D- THD CH* D G

D Al H+H -+ DD *

6 A%606;'T; 110 20

AlH+H -TH *C * +

G AAH+G*TH * + C+

Ta)%! '1 Mat! ia% 9 "9! t& ta)%!


53/61

/1 S9a +!8ign

$6ars are 3e3bers which are basically used to carry the bending and shear loads acting

on the wing during flight. There are two s6ars, one located at D-*+K of the chord known as

the front s6ar, the other located at H+- +K of the chord known as the rear s6ar. $o3e of the

functions of the s6ar include

They for3 the boundary to the fuel tank located in the wing.

The s6ar flange takes u6 the bending loads whereas the web carries the shear loads.

The rear s6ar 6ro8ides a 3eans of attaching the control surfaces on the wing.

!onsidering these functions, the locations of the front and rear s6ar are fi0ed at +. c and +.HDcres6ecti8ely. The sy33etric airfoil4&A!A CC D7 is drawn to scale using any design softwareand the chord thickness at the front and rear s6ar locations are found to be +.* 3 and . 3res6ecti8ely.


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/2 G!"*!t i$ +i*!n8i"n8

The s6ar design for the wing root has been taken because the 3a0i3u3 bending3o3ent and shear force are at the root. It is assu3ed that the flanges take u6 all the bendingand the web takes all the shear effect. The 3a0i3u3 bending 3o3ent for high angle of attackcondition is C* ++ . D &3. the ratio in which the s6ars take u6 the bending 3o3ent is gi8enas

/hereh - height of front s6ar h* - height of rear s6ar

' f (' r)#"7003.3

The yield tensile stress U y for Al Alloy4 A + D7 is ;62M,a . The area of the flanges isdeter3ined using the relation

where M is bending 3o3ent taken u6 by each s6ar,

A is the flange area of each s6ar,? is the centroid distance of the area O h;*.

sing the a8ailable 8alues,Area of front s6ar,

Af O +.+*3 *

Area of rear s6ar,Ar O +.+*C 3 *

/4 ASS M,TIONST sections are chosen for to6 and botto3 flanges of front and rear s6ars.

Both the flanges are connected by a 8ertical stiffener through s6ot welding"ro3 the buckling e uation,

the thickness to width ratio of web is found to be +.+C D. Also fro3 YA&AL1$I$ A#E$I%& (" "LI%5T VE5I!LE $T !T E$ by B 5&Z, the flange to web width ratio

of the T section .


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By e uating all the three 8alues of the ratio in area of the section e uation, the di3ensions ofthe s6ar can be found.

/; S,ECIFICATION FOR FRONT S,ARtf O+.+* 3

b flange O+.DH 3 bwebO+. D 3/5 S,ECIFICATION FOR REAR S,AR

tf O+.+ DH 3 b flange O+.C 3 bwebO+. G G 3

?/ DETAILED DESIGN OF F SELAGE

?/1 D!8$ i9ti"n

"uselage contributes 8ery little to lift and 6roduces 3ore drag but it is an i36ortantstructural 3e3ber;co36onent. It is the connecting 3e3ber to all load 6roducing co36onentssuch as wing, hori?ontal tail, 8ertical tail, landing gear etc. and thus redistributes the load. Italso ser8es the 6ur6ose of housing or acco33odating 6ractically all e ui63ent, accessoriesand syste3s in addition to carrying the 6ayload. Because of large a3ount of e ui63ent insidethe fuselage, it is necessary to 6ro8ide sufficient nu3ber of cutouts in the fuselage for accessand ins6ection 6ur6oses. These cutouts and discontinuities result in fuselage design being 3oreco36licated, less 6recise and often less efficient in design. As a co33on 3e3ber to whichother co36onents are attached, thereby trans3itting the loads, fuselage can be considered as along hollow bea3. The reactions 6roduced by the wing, tail or landing gear 3ay be consideredas concentrated loads at the res6ecti8e attach3ent 6oints. The balancing reactions are 6ro8ided


56/61

by the inertia forces contributed by the weight of the fuselage structure and the 8ariousco36onents inside the fuselage. These reaction forces are distributed all along the length of thefuselage, though need not be unifor3ly. nlike the wing, which is sub9ected to 3ainlyunsy33etrical load, the fuselage is 3uch si36ler for structural analysis due to its sy33etricalcross-section and sy33etrical loading. The 3ain load in the case of fuselage is the shear load

because the load acting on the wing is transferred to the fuselage skin in the for3 of shear only.The structural design of both wing and fuselage begin with shear force and bending 3o3entdiagra3s for the res6ecti8e 3e3bers. The 3a0i3u3 bending stress 6roduced in each of the3is checked to be less than the yield stress of the 3aterial chosen for the res6ecti8e 3e3ber.

.

?/2 St ing! S9a$ing

The stringers are sy33etrically s6aced on the fuselage with the s6acing calculate asshown below,

!ircu3ference of the fuselage OTotal nu3ber of stringers O H

Therefore the stringers are s6aced at the inter8al of O

?/4 St ing! a !a $a%$u%ati"nThe stress induced in the each stringer is calculated with the area kee6ing constant in

the stress ter3. Then the 3a0i3u3 stress 4i.e. one which has larger nu3erator7 is e uated withthe yield strength of the 3aterial. "ro3 this area of one stringer is calculated.The direct stress in each stringer 6roduced by bending 3o3ents and is gi8en by thee uation

/here' x)#"7003.3 *m

is density O .**D kg;3V is cruise 8elocity O* + 3;s$ t is the tail areaO .* 3 *

a t is the slo6e of the lift cur8e O+.+ ;deg is the angle of yaw for asy33etric flight

+)7.07 deg

0 is the distance between the aircraft c.g 6osition and hori?ontal tail c.g 6osition in O D.G 3Then,


57/61

' )1!!"13 .71$ *m

/here is the stringer area , # is the dia3eter of the fuselage in O +.D 3.

and reach their 3a0i3u3 only fro3 the stringers to . Thus the stresses are high onlyon these stringers. !alculating, the stress for stringer to .

@O+,:O+.

Then,- 1)1!%0%33."!! / T *2 4 * m "

$i3ilarly it is calculated for other @,: 8alues

The allowable stress in the stringer is CH*M'a. The 3a0i3u3 direct stress in the stringers is

/ atinger )3!.!! mm "

Thus one stringer area is . The stringer chosen is : section. The di3ensions of thestringers are obtained fro3 the A&AL1$I$ A #E$I%& (" T5E "LI%5T VE5I!LE$T !T E$ by B 5&.

T(! +i*!n8i"n8 a !=

twOtf OC33 b flange O +33hwebOH 33

? a:/S(!a F%"< Di8t i)uti"n

The shear flow in the fuselage section is calculated by considering the fuselage as closed section. The shear flow

is calculated and hence the 3a0i3u3 shear stress is found. The shear stress thus for3ed 3ust be less than theallowable shear stress.

A cut is 3ade in the fuselage section and the shear flow for o6en cell is calculated. The shear flow for the o6en

section is found out. Then by considering 3o3ent e uilibriu3, unbalanced shear flow is found. This shear flow is

then added to the shear flow obtained earlier. The 3a0i3u3 shear flow 8alue di8ided by its thickness yields shear

stress this should be less than the allowable stress.

I@@O I11 O Astringer # *


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The shear flow e uation for the fuselage section is gi8en by s O + ) [ : dA &;3

T(! 8t ing! 8 an+ t(!i %"$ati"n a ! gi7!n )!%"

design of medium range business jet transport

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