Project 2 Azar ADP ASS

8/16/2019 Project 2 Azar ADP ASS

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ADP-II AMPHIBIANS AIRCRAFT

ACKNOWLEDGEMENT

Firstly, I would like to thank the almighty God for always

being by my side and providing me with strength and capability

to face all types of situations during this project tenure.

I thank our beloved CE ! "ecretary Dr. P Krishna

Kumar , #ehru Group of Institutions, Coimbatore for providing

the facilities.

I e$tend my fullest and ever owing thanks to Dr. P.

Maniiarasan , %rincipal, #ehru Institute of Engineering and

&echnology Coimbatore for the academic freedom and

inspiration.

'e also thank our professor and (ead of the )epartment,

Prof V. Sankar , and Ms. Veni Grace and sta* members of the +eronautical department of #ehru Institute of Engineering

and &echnology for leading their support to this project.

I also thank everyone who lent us support in the

completion of this project.


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ABSTRACT

Aircraft design is a separate discipline of aeronautical engineering – different from the analytical disciplines such as aerodynamics, structures,

controls, and propulsion. An aircraft designer needs to be well versed in these

subjects. Instead, the designer’s time is spent doing something called “design”,

creating the geometric description of a thing to be built.

An amphibious aircraft or amphibian is an aircraft that can ta e off

and land on either land or water. Amphibious aircraft are slower, more comple!and more e!pensive to purchase and operate than comparable landplanes but are

also more versatile. Amphibians are also engineered with retractable wheels and

floats, the class which has retractable floats which act as e!tra fuel tan s since

fuel li"uids weigh less than water of e"ual volume# these floats are removable

for e!tended land$snow operations if and when use of e!tra fuel tan s is

undesired. In addition, amphibious aircraft are useful in light transport in remote

areas, where they are re"uired to operate not only from airstrips, but also from

la es and rivers.

%o there is a need to conduct a literature survey related to what sort

of aircraft is going to be designed.

&he aim of the project is concentrated towards the design of

amphibian aircraft. &he objective of this project is to provide a better design by

manipulating the previous designs .


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INDEX

S.NO CONTENTS PAGE NO.

' Introduction (

) *+n iagram '-

- ust *+n diagram )/

01ritical loading performance and final *+ndiagram

)0

( %tructural design study –theory approach )2

3 4oad estimation on wings -)

5 4oad estimation on fuselage 00

26alancing and maneuvering loads on tail plane,rudder and aileron loads

07

7 etailed structural layouts (0

'/esign of some components of wing and

fuselage3)

'' 8aterial selection 37

') esign report 53

'- &hree view diagram 57

'0 1onclusion 2/

'( 6ibliography 2/

NOMENCLATURE:


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A.9. + Aspect 9atio

b + :ing %pan ;m<

1 + 1hord of the Airfoil ;m<

1 root + 1hord at 9oot ;m<

1 tip + 1hord at &ip ;m<

1 + 8ean Aerodynamic 1hord ;m<

1d + rag 1o+efficient

1 d,/ + =ero 4ift rag 1o+efficient

1 p + %pecific fuel consumption ;lbs$hp$hr<

1 4 + 4ift 1o+efficient + rag ;><

? + ?ndurance ;hr<

e + @swald efficiency

4 + 4ift ;><

;4$ < loiter + 4ift+to+drag ratio at loiter

;4$ umber

% + :ing Area ;m <

& + &hrust ;><* cruise + *elocity at cruise ;m$s<

* stall + *elocity at stall ;m$s<

* t + *elocity at touch down ;m$s<

: crew + 1rew weight ; g<

: empty + ?mpty weight of aircraft ; g<

: fuel + :eight of fuel ; g<

: payload + Bayload of aircraft ; g


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:/ + @verall weight of aircraft ; g<

:$% + :ing loading ; g$m <

ρ∞ + ensity of air ; g$mC<

A stringer + 1ross sectional area of stringers

A + &otal cross sectional area

A spar + 1ross sectional area of spar

a t+%lope of the 14 vs. D curve for a horiEontal tail

a+ istance of the front spar from the nose of the aircraft

bw+:idth of the web

b f +:idth of the flangeI!! + %econd moment of area about F a!is

IEE + %econd moment of area about = a!is

G + ust alleviation factor

n ma! + 8a!imum load factor

tw + &hic ness of the web

tf + &hic ness of the flange& + &or"ue

H + ust velocity

* cruise + 1ruise velocity

* s + %talling velocity

+ Angle of aw

.


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1. INTRODUCTION


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AMPHIBIANS AIRCRAFT

The basic configuration of the Be-200 amphibious aircraft isindended for fighting the forest fires using the fire extinguishantfluids. While doing this, the aircraft can fulfill the following tasks:

stop and restrain the spread of the big forest fires bde!eloping the protecting strip due to multiple drops on thefire edge"

extinguishing the small fire and fire which onl starts tode!elop"

deli!er of fire brigades and fire extingushing e#uipment tothe fire region b landing on preselected water area of aifield,and return to the base.

Many early aircra ! "e#i$n# %ere "e&el'(e" 'r !a)e*' an" lan"in$ 'n%a!er. +i!, !,e a-#ence ' rea"ily a&aila-le le&el an" '%e" iel"#/ la)e#an" e&en !,e 'cean %ere l'')e" 0('n a# i"eal c,'ice# 'r a (lace !' lan" 'r!a)e*' . T,i# all'%e" '(era!i'n in a %i"e ran$e ' ,ea"in$# !'acc' '"a!e %in" "irec!i'n. I! al#' "i" n'! re 0ire any (re(ara!i'n 'rlan"in$ 'r !a)e*' '!,er !,an a 0ic) l'') !' a)e #0re !,a! -'a!# 'r"e-ri# %ere n'! in !,e i$,! (a!,. T,i# (r'&i"e" a "eci"e" a"&an!a$e '&erlan"*-a#e" '(era!i'n# %,ere real e#!a!e ,a" !' -e (0rc,a#e" 'r ren!e"/'-#!acle# 2!ree #!0 (# an" r'c)#3 cleare"/ an" $ra## c0! !' a rea#'na-le,ei$,! 'r a ,ar"ene" ear!, 'r aca"a #0r ace (re(are". In e er$encie#/a la)e %a# al#' 're li)ely !' -e clear ' '-#!acle# !,an a ar er4# iel"!,a! i$,! -e ille" %i!, ca!!le 'r -i#ec!e" -y a ence.

Hence/ any early air(lane "e#i$ner# '(!e" 'r a #ea(lane c'n i$0ra!i'n.In !,e e&en! !,a! lan" '(era!i'n %a# #'0$,!/ an a (,i-ian "e#i$n ' ere"!,e ca(a-ili!y ' %a!er 'r lan" '(era!i'n. In ac!/ "0e !' !,e (0-lic4# lac) ' c'n i"ence in air(lane en$ine relia-ili!y/ i! %a# n'! 0n!il al '#! !,e i"*!%en!ie!, cen!0ry !,a! l'n$/ '&er%a!er (a##en$er li$,!# 2!ran#a!lan!ic/!ran#(aci ic/ Cari--ean/ e!c.3 %ere r'0!inely a!!e (!e" in any!,in$ '!,er!,an #ea(lane# 'r a (,i-ian#.

+i!, e5!en#i&e 0#e ' lan"*-a#e" aircra ! !' !ran#('r! ili!ary (er#'nnel"0rin$ +'rl" +ar II an" %i!, i (r'&e en! in en$ine relia-ili!y/ !,e lyin$(0-lic $aine" !,e c'n i"ence nee"e" 'r #0c, aircra ! !' re(lace !,eir


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%a!er*-a#e" c'0n!er(ar!#. T,i# all'%e" inlan" air('r!# !' re(lace c'a#!al#i!e# a# ('r!# ' en!ry an" e5i! 'r '&er#ea# li$,!# an" !,e lar$e a (,i-ian#an" #ea(lane# ' !,e 1678# an" 1698# %ere re!ire" r' #er&ice.

In !,e $eneral a&ia!i'n 2GA3 iel"/ #ea(lane# an" a (,i-ian# ,a&e al%ay#'cc0(ie" a # all -0! i ('r!an! nic,e in !,e ar)e!(lace/ 0#e" (ri arily

'r '(era!i'n# in!' an" '0! ' re '!e area#%,ere la)e#%ere 're (len!i 0l!,an air('r!#. T'"ay/ '#! #0c, aircra ! !en" !' -e l'a!(lane#4/ aircra !'ri$inally "e#i$ne" 'r lan" '(era!i'n !' %,ic, ,a&e -een a""e" ra!,erlar$e l'a!# !' re(lace !,e c'n&en!i'nal %,eele" 0n"ercarria$e. S0c,aircra ! are 0#0ally c'n#i"era-ly #l'%er in li$,! an" 're li i!e" in(er 'r ance !,an !,eir 'ri$inal "e#i$n# "0e !' !,e a""e" %ei$,! an" "ra$' !,e l'a!#. In a!!e (!# !' $e!-e!!er '&erall (er 'r ance/ a e% #(ecial!yaircra ! ,a&e -een "e#i$ne" a# a (,i-ian# %i!, a ,0ll 0#ela$e. H'%e&er/!,e c' (r' i#e# re 0ire" !' all'% -'!, lan" an" %a!er '(era!i'n# ,a&e#!ill re#0l!e" in a""e" %ei$,! an" c' (le5i!y/ an" a l'%er cr0i#e #(ee" !,anc'n&en!i'nal lan"*-a#e" aircra ! "e#i$n#.

In !,e 'll'%in$ #0 ary ' !,e "e#i$n (r'ce##/ e (,a#i# %ill -e (lace" 'n!,e ac!'r# 0ni 0e !' a (,i-ian aircra !. C'n#i"era!i'n ' a#(ec!# ' !,e(r'ce## !,a! are c' 'n !' all aircra ! "e#i$n# %ill -e $i&en 're c0r#'ryc'&era$e.

+,a! i# an Airli !;An airlift is the organiEed delivery of supplies or personnel primarily

via aircraft. Airlifting consists of two distinct types, strategic and tactical

airlifting. &ypically, strategic airlifting involves moving material long distances

;such as across or off the continent or theater


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receivers have control without fear of the enemy interfering with collection

and$or stealing the goods, the planes can maintain a normal flight altitude and

simply drop the supplies down and let them parachute to the ground. Jowever,

when the area is too small for this method, as with an isolated base, and$or is too

dangerous to land in, a 4ow Altitude Barachute ?!traction %ystem drop is used.

+,a! i# an Airli !;

An airlift is the organiEed delivery of supplies or personnel primarily

via aircraft. Airlifting consists of two distinct types, strategic and tactical

airlifting. &ypically, strategic airlifting involves moving material long distances

;such as across or off the continent or theater


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%&9A&? I1 AI94IK&&A1&I1A4 AI94IK&

STRATEGIC AIRLIFT

%trategic airlift is the use of cargo aircraft to transport materiel, weaponry,

or personnel over long distances. &ypically, this involves airlifting the re"uired

items between two airbases which are not in the same vicinity. &his

allows commanders to bring items into a combat theater from a point on the

other side of the planet, if necessary. Aircraft which perform this role are

considered #!ra!e$ic airli !er# . &his contrasts with tactical airlifters, such as

the 1+'-/ Jercules, which can normally only move supplies within a

given theater of operations.

?FA8B4?L 4oc heed 1+( ala!y, Antonov An+')0

TACTICAL AIRLIFT

&actical airlift is a military term for the airborne transportation of supplies and

e"uipment within a theatre of operations ;in contrast to strategic airlift


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supplies. 8ost are fitted with defensive aids systems to protect them from attac

by surface+to+air missiles.

?FA8B4?L Jercules 1+'-/, 4oc heed 1+'0' %tarlifter

DESIGN OF AN AIRPLANE

Airplane design is both an art and a science. It’s the intellectual engineering

process of creating on paper ;or on a computer screen< a flying machine to

meet certain specifications and re"uirements established by potentialusers ;or as perceived by the manufacturer< and

pioneer innovative, new ideas and technology

&he design process is indeed an intellectual activity that is rather specified one

that is tempered by good intuition developed via by attention paid to successful

airplane designs that have been used in the past, and by ;generally proprietary<

design procedure and databases;hand boo s etc< that are a part of every airplanemanufacturer.

PHASES OF AIRPLANE DESIGN

&he complete design process has gone through three distinct phases that are

carried out in se"uence. &hey are

1onceptual design

Breliminary design

etailed design


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CONCEPTUAL DESIGN

&he design process starts with a set of specifications ;re"uirementso part

of the design is ever carried out in a total vacuum unrelated to the other parts.

PRELIMINAR< DESIGN


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In the preliminary design phase, only minor changes are made to the

configuration layout ;indeed, if major changes were demanded during this

phase, the conceptual design process have been actually flawed to begin with. It

is in the preliminary design phase that serious structural and control system

analysis and design ta e place. uring this phase also, substantial wind tunnel

testing will be carried out and major computational fluid dynamics ;1K <

calculations of the computer flow fluid over the airplane configurations are

done.

Its possible that the wind tunnel tests the 1K calculations will in cover some

undesirable aerodynamic interference or some une!pected stability problems

which will promote change to the configuration layout. At the end of

preliminary design phase the airplane configuration is froEen and preciously

defined. &he drawing process called lofting is carried out which mathematically

models the precise shape of the outside s in of the airplane ma ing certain that

all sections of the aircraft property fit together

&he end of the preliminary design phase brings a major concept to commit the

manufacture of the airplane or not. &he importance of this decision point for

modern aircraft manufacturers cannot be understated, considering the

tremendous costs involved in the design and manufacture of a new airplane.

DETAIL DESIGN

&he detail design phase is literally the nuts and bolts phase of airplane design.

&he aerodynamic, propulsion, structures performance and flight control analysis

have all been finished with the preliminary design phase. &he airplane is now

simply a machine to be fabricated. &he pressure design of each individual rib,

spar and section of s in now ta e place. &he siEe of number and location of

fastness are determined. At this stage, flight simulators for the airplane are


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developed. And these are just a few of the many detailed re"uirements during

the detail design phase. At the end of this phase, the aircraft is ready to be

fabricated.

OUTLINE AIRCRAFT DESIGN PRO=ECT >:

&he structural design of the aircraft which is done in aircraft design project )

involvesL

etermination of loads acting on aircraft

• *+n diagram for the design study

• ust and maneuverability envelopes

• %chren ’s 1urve

• 1ritical loading performance and final *+n graph calculation

etermination of loads acting on individual structures

• %tructural design study – &heory approach

• 4oad estimation of wings

• 4oad estimation of fuselage.

•

8aterial %election for structural members• etailed structural layouts

• esign of some components of wings, fuselage

Para e!er# !a)en r' aircra ! "e#i$n (r'?ec! 1:

Para e!er# @al0e#

:ing loading; g$m ) < -(5.3


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8ach number /.37

&hrust to weight ratio /.'(()3

Aspect ratio 7.'

Altitude; m< 2

8a!imum 4ift coefficient '-.-

:ing span;m< -).52

:ing planform area;m ) < ''5.0

Kuel weight; g< '3'-/

?ngine weight; g< '0(/

@verall weight; g< 0-.3

1ruise speed;Gm$hr< 5'/

%talling speed;Gm$hr< ''3

%ervice speed; m< 5)/

9oot chord;m< (.(2

&ip chord;m< '.5)Muarter chord sweep angle;deg< (.5(

8ean aerodynamic chord;m< 3./-/52

&hrust per engine;G>< 5-.3

9ange; m< -3//

Bayload; g< ')///


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>. @*n Dia$ra

INTRODUCTION:

Airplanes may be subjected to a variety of loading conditions in flight. &hestructural design of the aircraft involves the estimation of the various loads on

the aircraft structure and designing the airframe to carry all these loads,

providing enough safety factors, considering the fact that the aircraft under

design is a commercial transport airplane. As it is obviously impossible to

investigate every loading condition that the aircraft may encounter, it becomes

necessary to select a few conditions such that each one of these conditions will

be critical for some structural member of the airplane.

@el'ci!y L'a" Fac!'r 2@*n3 "ia$ra :

&he control of weight in aircraft design is of e!treme importance. Increases in

weight re"uire stronger structures to support them, which in turn lead to further

increases in weight and so on. ?!cess of structural weight mean lesser amounts

of payload, thereby affecting the economic viability of the aircraft. &he aircraft

designer is therefore constantly see ing to pare his aircraft’s weight to the

minimum compatible with safety. Jowever, to ensure general minimum

standards of strength and safety, airworthiness regulations ;Av.B.75/ and

61A9< lay down several factors which the primary structure of the aircraft

must satisfy. &hese are the

• Li i! l'a" , which is the ma!imum load that the aircraft is e!pected to

e!perience in normal operation.• Pr'' l'a" , which is the product of the limit load and the (r'' ac!'r ;'./+

'.)(


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&he basic strength and fight performance limits for a particular aircraft are

selected by the airworthiness authorities and are contained in the

flightenvelopeor @*n "ia$ra .

&here are two types of * – n diagram for military airplanes L

*–n maneuver diagram and*–n gust diagram

@ n MANEU@ER DIAGRAM:

&he positive design limit load factor must be selected by the designer, but must

meet the following condition

lim ¿( pos)≥ 2.1 + 24000

W +10000n¿

lim ¿( pos)≥ 2.1 + 24000

43600 +10000n¿


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lim ¿( pos)≥ 2.55n¿

&he ma!imum positive limit load factor for military transport aircraft should be

in the range ) to -. %o for our aircraft we ta elim ¿( pos)= 2. 3

n¿

&he ma!imum negative limit load factor is given by¬¿¿

lim ¿( pos)¿

lim ¿¿¿n¿

¬¿¿¿

lim ¿¿¿n¿

¬¿¿¿

lim ¿¿¿n¿

&here are four important speeds used in the * – n diagram

' – g stall speed * %esign maneuvering speed * Aesign cruise speed * 1

esign diving speed *

P'#i!i&e 1 $ #!all #(ee" @ S

V S=√ 2 ρ C Nmax W SC Nmax= 1.1 ×C LmaxC Nmax= 1.1 × 1.131C Nmax= 1.246

V S=√ 21.125 × 1.246 × 357.6V S= 22.588 m /s

Ne$a!i&e 1 $ #!all #(ee" @ Sne$


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¬¿¿¿¿

Nmax¿ ρC ¿

2¿

S¬¿= √ ¿V ¿

¬¿¿

¬¿¿¿

Lmax¿¿

Nmax¿C ¿¬¿¿¿

Nmax¿C ¿¬¿¿¿

N max¿C ¿

S¬¿=

√ 2

1.125 × 0.495 × 357.6

V ¿

S¬¿= 33.55 m /sV ¿

De#i$n Mane0&erin$ #(ee" @ A 'r ('#i!i&e l'a" ac!'r

2nlim ¿( pos)

ρ C NmaxW S

V A= √ ¿

V A=√ 2 × 2.51.125 × 1.246 × 357.6V A= 27.009 m/ s

De#i$n Mane0&erin$ #(ee" @ B 'r ne$a!i&e l'a" ac!'r


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¬¿¿¿¿

¬¿¿¿¿ Nmax¿¿

lim ¿¿¿

2 n¿¿

V B= √ ¿

V B=√ 2 × 1.021.125 × 0.495 × 357.6V B= 17.19 m /s

De#i$n Cr0i#e #(ee" @ C

Krom Aircraft esign Broject ',

* 1 N * cruise N 5'/ m$hr

* 1 N '75.)) m$s

De#i$n Di&in$ S(ee" @ D

&he design diving speed must satisfy the following relationshipV D ≥ 1.25 V cruise

V D= 1.25 × '75.))

V D= 246.525 m /s

C0r&e OA

&he velocity along the curve @A is given by the e!pression

V Sn=√ 2 n

ρ C NmaxW S

Krom this e!pression the load factor along the curve @A is given by

n= ρ C Nmax V 2

2

1

W S

n=1.125 × 1.246 V

2

2

1

357.6

n= 1.959 × 10 − 3 V 2


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@el'ci!y # P'#i!i&e L'a" Fac!'r n

/ /

'/ /.'7(-

)/ /.52-7

-/ '.53-'

-- )./--

-5 ).'0(

(/ ).'0(

C0r&e OG

&he negative load factor along the curve @ is given by the e!pression¬¿¿

¿V 2¿

Nmax¿

ρ C ¿¬¿= ¿n¿

¬¿=1.125 × 0.495 V

2

2

1

357.6

n¿

¬¿= 7.79 × 10 − 4 V 2n¿

@el'ci!y # Ne$a!i&e L'a" Fac!'r n ne$

/ /

'/ +/.'7(7

)/ +/.('')

-/ +/.2//)

-5 +/.7)

(/ +/.7)


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7. GUST @*n DIAGRAM


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Description:

ust is a sudden, brief increase in the speed of the wind. enerally, winds are

least gusty over large water surfaces and most gusty over rough land and near

high buildings. :ith respect to aircraft turbulence, a sharp change in wind speed

relative to the aircraft# a sudden increase in airspeed due to fluctuations in the

airflow, resulting in increased structural stresses upon the aircraft . S,ar(*e"$e"

$0#! ;u< is a wind gust that results in an instantaneous change in direction or

speed.

Deri&e" $0#! &el'ci!y ;H g or H ma! < is the ma!imum velocity of a sharp+edged

gust that would produce a given acceleration on a particular airplane flown inlevel flight at the design cruising speed of the aircraft and at a given air density.

As a result a )(O increase is seen in lift for a longitudinally disturbing gust.

&he effect of turbulence gust is to produce a short time change in the effective

angle of attac . &hese changes produce a variation in lift and thereby load

factor. Kor * 6 , a gust velocity of )/.''32 m$s is assumed. Kor * 1 , a gust

velocity of '(.)0 m$s at sea level is assumed. Kor * , a gust velocity of 5.)3m$sis assumed.

E ec!i&e $0#! &el'ci!y L &he vertical component of the velocity of a sharp+

edged gust that would produce a given acceleration on a particular airplane

flown in level flight at the design cruising speed of the aircraft and at a given air

density.

C'n#!r0c!i'n ' $0#! l'a" ac!'r line#

&he gust load factor lines are defined by the following e"uations

lim ¿= 1 ± (k g U g V C Lα ρ)2 W

Sn ¿

k g=0.88 μg5.3 + μg


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μg=2 (W S )

ρ Ć C Lα where,

k g− ¿ ust alleviation factor U g− ¿ erived gust velocityV B− ¿ esign speed for ma!imum gust intensityV c− ¿ esign cruise velocityV D− ¿ esign diving velocityC Lα − ¿ @verall lift curve slope rad +'

Ć − ¿ :ing mean geometric chordW S

= 357.6 kgm

2 ρ= 1.225 kg/m3

C Lα = 13.3 Ć = 0.1602 m

μg = 2 × 357.6

1.225 × 0.1602 × 13.3 = 11.334

k g=0.88 × 19.3345.3 +19.334 =

0.3926

1onstruction of gust load factor line for speed V B= 17.19 m/s ;ta e

U g = 5.0 m /s <lim ¿= 1.0446

+n¿lim ¿= 0.0543

− n¿

1onstruction of gust load factor line for speed V c= 197.22 ;ta es

U g = 6 m /¿¿lim ¿= 1.225

+n¿lim ¿=− 0.7325

− n¿

1onstruction of gust load factor line for speed V c= 247.525 m /s ;ta e

U g = 10 m /s <lim ¿= 1.8025

+n¿lim ¿=− 0.2035

− n¿


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9. CRITICAL LOADING PERFORMANCE AND FINAL @*n DIAGRAM


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CRITICAL LOADING PERFORANCE:

&he greatest air loads on an aircraft usually comes from the generation of lift

during high+g maneuvers. ?ven the fuselage is almost always structurally siEed

by the lift of the wings rather than by the pressures produced directly on the

fuselage. Aircraft load factor ;n< e!presses the maneuvering of an aircraft as a

standard acceleration due to gravity.

At lower speeds the highest load factor of an aircraft may e!perience is limited

by the ma!imum lift available. At higher speeds the ma!imum load factor is

limited to some arbitrary value based upon the e!pected use of the aircraft. &he

ma!imum lift load factor e"uals './ at levels flight stall speed. &his is theslowest speed at which the ma!imum load can be reached without stalling.

&he aircraft ma!imum speed, or dive speed at right of the *+n diagram

represents the ma!imum dynamic pressure and ma!imum load factor is clearly

important for structural siEing. At this condition, the aircraft is at fairly low

angle of attac because of the high dynamic pressure, so the load is

appro!imately vertical in the body a!is. &he most common maneuvers that wefocused are,

4evel turn

Bull up

Bull down

1limb

Le&el !0rn:

&he value of minimum radius of turn is given by the formula,

! min=4 k (W S )

g ρ( " W )√1 − 4 k C D 0( " W )&he load factor at minimum radius of turn is given by,


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" /W ¿2¿¿

2 −4 k C Do

¿

n ! min= √ ¿%ubstituting the nown values,

! min= 603.60 8 m

n ! min= ¿ './('

P0ll*0( Mane0&er:

! = V ∞2

g (n− 1 )

%ubstituting the nown values and 9 N -(// m

n= 2.54

P0ll*"'%n Mane0&er:

! = V ∞2

g (n− 1 )

%ince the radius for pull down is same as that of the pull up maneuver, the load

factor for pull down maneuver is found to be,

n= 0.63

Cli -:

( " W )− # ¿2 −( 4 C Do$eA!

)}0.5

¿

[( " W )− # ]+{¿n= ¿

C%im& gra'ien(# = sin ) = sin 5 = 0.87155

%ubstituting the nown values

n= 1. 458

Mane0&er L'a" Fac!'r n

4evel turn './'(


31/84

23


Bull+up ).(0

Bull+down /.3-

1limb '.0(2

Final @*n Dia$ra :


32/84

23


. STRUCTURAL DESIGN STUD< THEOR< APPROACH


33/84

23


STRUCTURAL DESIGN STUD< THEOR< APPROACH

Aircraft loads are those forces and loadings applied to the airplanes structural

components to establish the strength level of the complete airplane. &hese

loadings may be caused by air pressure, inertia forces, or ground reactions

during landing. In more specialiEed cases, design loadings may be imposed

during other operations such as catapulted ta e+offs, arrested landings, or

landings in water.

&he determination of design loads involves a study of the air pressures and

inertia forces during certain prescribed maneuvers, either in the air or on theground. %ince the primary objective is an airplane with a satisfactory strength

level, the means by which this result is obtained is sometimes unimportant.

%ome of the prescribed maneuvers are therefore arbitrary and empirical which is

indicated by a careful e!amination of some of the criteria.

Important consideration in determining the e!tent of the load analysis is the

amount of structural weight involved. A fairly detailed analysis may benecessary when computing operating loads on such items as movable surfaces,

doors, landing gears, etc. proper operation of the system re"uires an accurate

prediction of the loads.

Aircraft loads is the science of determining the loads that an aircraft structure

must be designed to withstand. A large part of the forces that ma e up design

loads are the forces resulting from the flow of air about the airplane’s surfaces+the same forces that enable flight and control of the aircraft.

L'a" ac!'r#

In normal straight and level flight the wing lift supports the weight of the

airplane. uring maneuvers or flight through turbulent ;gusty< air, however,

additional loads are imposed which will increase or decrease the net loads on

the airplane structure. &he amount of additional loads depends on the severity of


34/84

23


the maneuvers or the turbulence, and its magnitude is measured in terms of load

factor.

&he ma!imum maneuvering load factor to which an airplane is designed

depends on its intended usage. Kighters, which are e!pected to e!ecute violent

maneuvers, are designed to withstand loads commensurate with the

accelerations a pilot can physically withstand. 4ong range, heavily loaded

bombers, on the other hand, are designed to low load factors and must be

handled accordingly.

Kor a typical two spar layout, the ribs are usually formed in three parts from

sheet metal by the use of presses and dies. Klanges are incorporated around theedges so that they can be riveted to the s in and the spar webs 1ut+outs are

necessary around the edges to allow for the stringers to pass through 4ightening

holes are usually cut into the rib bodies to reduce the rib weight and also allow

for passage of control runs fuel electrics etc .


35/84

23


STRUCTURAL DESIGN CRITERIA

&he structural criteria define the types of maneuvers, speed, useful loads,

and gross weights which are to be considered for structural design analysis.

&hese are items which are under the control of the airplane operator. In addition,

the structural criteria must consider such items as inadvertent maneuvers, effects

of turbulent air, and severity of ground contact during landing. &he basic

structural design criteria, from which the loadings are determined, are based

largely on the type of the airplane and its intended use.


36/84

23


. LOAD ESTIMATION ON +INGS


37/84

23


De#cri(!i'n

&he solution methods which follow ?uler’s beam bending theory

;P$yN8$IN?$9< usethe bending moment values to determine the stressesdeveloped at a particular section of thebeam due to the combination of

aerodynamic and structural loads in the transverse direction.8ost engineering

solution methods for structural mechanics problems ;both e!act andappro!imate

methods< use the shear force and bending moment e"uations to determine

thedeflection and slope at a particular section of the beam. &herefore, these

e"uations are to beobtained as analytical e!pressions in terms of span wiselocation. &he bending momentproduced here is about the longitudinal ;!< a!is.

L'a"# ac!in$ 'n %in$

As both the wings are symmetric, let us consider the starboard wing at first.

&here arethree primary loads acting on a wing structure in transverse direction

which can causeconsiderable shear forces and bending moments on it. &hey are

as followsL

4ift force ;given by %chren ’s curve<%elf+weight of the wing:eight of the power plant

:eight of the fuel in the wing

S,ear 'rce an" -en"in$ ' en! "ia$ra # "0e !' l'a"# al'n$ !ran#&er#e

"irec!i'n a! cr0i#e c'n"i!i'n

4ift varies along the wing span due to the variation in chord length, angle of

attac and sweep along the span. %chren ’s curve defines this lift distribution

over the wing span of an aircraft, also called simply as 4ift istribution 1urve.

%chren ’s 1urve is given by


38/84

23


*= *1+ *2

2

where

y ' is 4inear *ariation of lift along semi wing span also named as 4'y) is ?lliptic 4ift istribution along the wing span also named as 4)

Linear li ! "i#!ri-0!i'n 2!ra(e i0 3:

4ift at root

Lroo( = ρV

2C Lc roo( 2

4root N 2/'35.2- >$m

4ift at tip

L(ip= ρV

2C Lc (ip2

4tip N )//0'.73 >$m

6y representing this lift at sections of root and tip we can get the e"uation for

the wing.

?"uation of linear lift distribution for starboard wing

y ' N +'5'5.225! Q 2/'35.2-

?"uation of linear lift distribution for port wing we have to replace ! by –! in

general,

y ' N '5'5.225! Q 2/'35.2-

Kor the %chren ’s curve we only consider half of the linear distribution of lift

and hence we derive y'$)

*12

N +2(2.70'! Q0//2-.7'(


39/84

23


0 5 10 15 20 25 30 35 40

0

10000

20000

30000

40000

50000

60000

70000

80000

90000

Linear variation of lift along se i !ing span

Half !ing span

Lift per Lengt" #N$ %

Elli(!ic Li ! Di#!ri-0!i'n:

&wice the area under the curve or line will give the lift which will be re"uired toovercome weight

1onsidering an elliptic lift distribution we get

L2

= W 2 = $a&

4

A= $a&4

:here

b is actual lift at root and a is wing semi span

4ift at tip,

&=4 W 2 $a

= 64117.38 N /¿ m

?"uation of elliptic lift,

*2 =

√&

2 (1 − x2

a2 )


40/84

23


*2 = 1831.925 √ (1225 − x2 )

*22 = 915.962 √ (1225 − x2)

0 5 10 15 20 25 30 35 40

0

10000

20000

30000

40000

50000

60000

70000

&lliptical lift 'istri()tion along se i !ing span

Half !ing span


C'n#!r0c!i'n ' Sc,ren)4# C0r&e:

%chren ’s 1urve is given by,

*= *1+ *2

2

*=− 858.941 x+40083.915 +915.962 √ (1225 − x2 )

%ubstituting different values for ! we can get the lift distribution for the wing

semi span


41/84

23


0 5 10 15 20 25 30 35 400

10000

20000

30000

40000

50000

60000

70000

80000

Sc"ren*+s c)rve for se i !ing span

Half !ing span


Sc,ren)4# c0r&e:

-40 -30 -20 -10 0 10 20 30 40

0

10000

20000

30000

40000

50000

60000

70000

80000Sc"ren*+s c)rve for f)ll !ing span

,ing span



42/84

23


Sel *+ei$,! ' %in$ 2y 73:

%elf+weight of the wing,

W W+N#

W ¿= 0.25

: :I> N /.)(R-(7--'R7.2'

: :I> N 22')(7 >

: portwing N +00/3)7 > ;Acting ownwards<

: starboard N + 00/3)7 >;Acting ownwards<

Assuming parabolic weight distribution

*3 = k ( x−&2 )

2

where b – wing span

:hen we integrate from !N/ ;root location< to !Nb ;tip location< we get the net

weight of port wing.

− 440629 =∫0

35

k ( x− &2 )2

'x

N +').--)( *3 =− 12.3325 ( x− 35 )

2

0 5 10 15 20 25 30 35 40

-16000

-14000

-12000

-10000

-8000

-6000

-4000

-2000

0

Self !eig"t of !ing

Half !ing span

,ing !eig"t per )nit Lengt" #N$ %


43/84

23


F0el %ei$,! in !,e %in$:

&his design has fuel in the wing so we have to consider the weight of the fuel in

one wing.

W ,ue%-ing2

= 104780.912

kg= 52390.45 kg

W ,ue%-ing= 513950.41 N

Again by using general formula for straight line yNm! Q c we get,

*, = 1185.185 x . 39775.92

0 5 10 15 20 25 30 35

-40000

-35000

-30000

-25000

-20000

-15000

-10000

-5000

0

F)el !eig"t in !ing

Half !ing span

F)el !eig"t per Lengt" #N$ %

P'%er (lan! %ei$,!:

Bower plant is assumed to be a point load,

:ppN-3-/ g N -(3'/.- >

Acting at !N 2 m and !N )/ m from the root.


44/84

23


0 5 10 15 20 25 30 35 40

-60000

-40000

-20000

0

20000

40000

60000

80000

100000

verall Loa' 'istri()tion on !ing

Half !ing span

Force per Lengt" #N$ %

L'a"# #i (li ie" a# ('in! l'a"#:

C0r&e c' ('nen!

Area encl'#e" #!r0c!0ral%ei$,! 2N <

Cen!r'i"2 r' %in$ r''!3

y ' $) '5(-35'.-)( '0 m

y) $) '53)('2.(-5 '0.200 m

:ing 00/3)7 2.5( m

Kuel ('-7(/.0' '/.(/' m

Bower plant -(3'/.- 2m, )/m


45/84

23


Reac!i'n 'rce an" Ben"in$ ' en! calc0la!i'n#:

&he wing is fi!ed at one end and free at other end.

∑ V = 0 ,&hen,

* A+'5(-35'.-)(+'53)('2.(-5Q00/3)7Q('-7(/.0'Q-(3'/.-Q-(3'/N/

* AN )07/-7/.'() >

∑ / = 0 ,&hen,

8 A+ ;'5(-35'.-)(R'0< + ;'53)('2.(-5R'0.200< Q ;00/3)7R2.5(< Q

;('-7(/.0'R'/.(/'< Q ;-(3'/.-R2< Q ;-(3'/.-R)/< N /

8 A N 261419642.5 >$m

>ow we now * A and 8 A, using this we can find out shear force and 6ending

moment.

SHEAR FORCE

S0 BC =∫ ( *1+ *12 − *3)'x − V AS0 BC =∫ (− 858.941 x+40083.915 +915.962 √ (1225 − x2)+12.3325 ( x− 35 )2 )'x − 2490390.152

S0 BC =− 429.4705 x2 +40083.915 x+915.962 x√ 1225 − x2 +1225sin − 1( x35 ) +12.3325 [ x

3

3 − 35 x2 +1225 x]−


46/84

23


S0 CD= S0 BC +∫ *, 'xS0 CD= S0 BC +∫ (1185.185 x . 39775.92 ) 'xS0 CD= S0 BC +(592.592 x

2 − 39775.92 x)

S0 D1= S0 CD− 35610.3

S0 10 = S0 D1 − 35610.3

S0 0A= S0 10 −( 592.592 x2 − 39775.92 x) Q('-7(/.0'

6y using the corresponding values of ! in appropriate e"uations we get the plot

of shear force.

-40 -30 -20 -10 0 10 20 30 40

-1500000

-1000000

-500000

0

500000

1000000

1500000

2000000

S"ear force 'iagra

Location in !ing

S"ear force #N%

BENDING MOMENT:

B/ BC =∬( *1 + *22 + *3 − V A)'x 2 + / AB/ BC =− 143.156 x

3 +20041.96 x2 +457.98 x[ x√ 1225 − x2 +1225sin − 1( x35 )]+305.32 [1225 − x2 ]1.5 − 12.3325B/ CD=∬( *1 + *22 + *3 + *, − V A)'x 2 + / AB/ CD = B/ BC +∬ *, 'x

2


47/84

23


N 68 61 Q '75.(-! - +'7225.73! )

68 ? N 68 1 +-(3'/.-!

68 ?K N 68 ? +-(3'/.-!

68 KAN 68 ?K – S'75.(-! - +'7225.73! ) TQ ('-7(/.0'!

6y substituting the values of ! for the above e"uations of bending moments

obtained we can get a continuous bending moment curve for the port wing.

-40 -30 -20 -10 0 10 20 30 40

0

100000000

200000000

300000000

400000000

500000000

600000000

700000000

800000000

Ben'ing o ent 'iagra

Location in !ing

Ben'ing o ent #N %

./ L AD &STIMATI N N F0S&LA1&


48/84

23



49/84

23


LOAD ESTIMATION ON FUSELAGE

Kuselage contributes very little to lift and produces more drag but it is an

important structural member$component. It is the connecting member to all load

producing components such as wing, horiEontal tail, vertical tail, landing gear

etc. and thus redistributes the load. It also serves the purpose of housing

oraccommodating practically all the e"uipments, accessories and systems in

addition to carrying the payload. 6ecause of large amount of e"uipment inside

the fuselage, it is necessary to provide sufficient number of cutouts in the

fuselage for access and inspection purposes. &hese cutouts and discontinuities

result in fuselage design being more complicated, less precise and often lessefficient in design. As a common member to which other components are

attached, thereby transmitting the loads, fuselage can be considered as a long

hollow beam. &he reactions produced by the wing, tail or landing gear may

beconsidered as concentrated loads at the respective attachment points. &he

balancing reactions are provided by the inertia forces contributed by the weight

of the fuselage structure and the various components inside the fuselage. &hesereaction forces are distributed all along the length of the fuselage, though need

not beuniformly.Hnli e the wing, which is subjected to mainly unsymmetrical

load, the fuselage is much simpler for structural analysis due to its symmetrical

cross+section and symmetrical loading. &he main load in the case of fuselage

is the shear load because the load acting on the wing is transferred to the

fuselage s in in the form of shearonly. &he structural design of both wing andfuselage begin with shear force and bending moment diagramsfor the respective

members

&o find out the loads and their distribution, consider the different cases. &he

main components of the fuselage loading diagram areL

• :eight of the fuselage• ?ngine weight• :eight of the horiEontal and vertical stabiliEers


50/84

23


• &ail lift• :eight of crew, payload and landing gear• %ystems, e"uipment, accessories

%ymmetric flight condition, steady and level flightL ; ownward forces negative<*alues for the different component weights are obtained from aerodynamic

design calculations .

Ta-le 1:L'a"# ac!in$ 'n F0#ela$e

C'n"i!i'n F0ll Payl'a"

F0#ela$e al'ne analy#i#

S.N' C' ('nen!#

Di#!ance r're erence line2 3

Ma##2)$3 +ei$,! 2N3

M' en!2N 3

' 1rew 0.07' -// )70- '-)'5./'-

) >ose 4andingear 7.72) (-// ('77- ('2770.')3

- Bay 4oad 6ay ' '5.73( 0(/// 00'0(/ 57-/307.)(

0 Ki!ed ?"uipment )5.('- '5// '3355 0(22-0.-/'

( Kuselage mass --.0'5 500'2 5-//0' )0-7(52/.'

3 8ain 4andingear ' --.0'5 '(7// '((757 ()')-(/.)0-

5 8ain 4andingear ) 00.5(- '/'// 77/2' 00-0'5'.77-

2 Bayload bay ) 00.5(- 0(/// 00'0(/ '75(3)''.2(

7 JoriEontalstabiliEer 33.73 7(// 7-'7( 3)0/--5.)

'/ *ertical %tabiliEer 5'.(23 02// 05/22 --5/20'.(32

&otal )')/'2 )/57273.(2 5)--'-25.30

c.g. from nose N 79.


51/84

23


Ta-le >: S,ear 'rce an" -en"in$ ' en! !a-0la!i'n

Di#!ance2 3 L'a" 2N3 S,ear F'rce 2N3 Ben"in$ M' en!2N 3

/ / / /

0.07' +)70- +)70- +'-)'5./'-

7.72) +('77- +(07-3 +(/(555.''-

'5.73( +00'0(/ +073-23 +20-30)3.-3-

)5.('- +'3355 +'3355 +227()3/.330

--.0'5 +5-//0' +')0-'/0 +--)7'/0/.53

--.0'5 +'((757 +'-77/2- +-2(/--7'./'

-0.553 )/57273 32/2'- --2)5773.3-00.5(- +77/2' (2'5-) )7-7-2)0.30

00.5(- +00'0(/ '0/)2) 73-53').57

33.73 +7-'7( 05/22 --5/20'.(32

5'.(23 +05/22 / /

S,ear 'rce 'n !,e 0#ela$e 2 ree* ree -ea %i!, 'ne reac!i'n a! i!# c.$.3 a!0lly l'a"e" c'n"i!i'n:

0 10 20 30 40 50 60 70 80

-2000000

-1500000

-1000000

-500000

0

500000

1000000

S"ear force 'iagra

Distance fro nose # %

S"ear force #N%


52/84

23


Ben"in$ ' en! 'n !,e 0#ela$e 2 ree* ree -ea %i!, 'ne reac!i'n a! i!#c.$.3 a! 0lly l'a"e" c'n"i!i'n:

0 10 20 30 40 50 60 70 80

-50000000

-40000000

-30000000

-20000000

-10000000

0

10000000

20000000

30000000

40000000Ben'ing o ent 'iagra

Distance fro nose # %

Ben'ing o ent #N %


53/84

23


. BALANCING AND MANEU@ERING LOADS ON TAIL PLANE/RUDDER AND

AILERON LOADS


54/84

23


Mane0&erin$ l'a"#:

?ach horiEontal surface and its supporting structure, and the main wing of a

canard or tandem wing configuration, if that surface has pitch control, must be

designed for the maneuvering loads imposed by the following conditionsL

A sudden movement of the pitching control, at the speed * A, to the

ma!imum aft movement, and the ma!imum forward movement, as limited

by the control stops, or pilot effort, whichever is critical.

A sudden aft movement of the pitching control at speeds above * A, followed

by a forward movement of the pitching control resulting in the following

combinations of normal and angular acceleration. At speeds up to * A, the

vertical surfaces must be designed to withstand the following conditions. In

computing the loads, the yawing velocity may be assumed to be EeroL

:ith the airplane in unaccelerated flight at Eero yaw, it is assumed that the

rudder control is suddenly displaced to the ma!imum deflection, as limited

by the control stops or by limit pilot forces.

:ith the rudder deflected, it is assumed that the airplane yaws to the over

swing sideslip angle. In lieu of a rational analysis, an over swing angle e"ual

to '.( times the static sideslip angle may be assumed.

A yaw angle of '( degrees with the rudder control maintained in the neutral

position ;e!cept as limited by pilot strength<

&he airplane must be yawed to the largest attainable steady state sideslip

angle, with the rudder at ma!imum deflection caused by any one of the

followingL

→ 1ontrol surface stops

→ 8a!imum available booster effort

→ 8a!imum pilot rudder force

→ &he rudder must be suddenly displaced from the ma!imum deflection to

the neutral position.


55/84

23


→ &he yaw angles may be reduced if the yaw angle chosen for a particular

speed cannot be e!ceeded inL

→ %teady slip conditions

→ Hncoordinated rolls from steep ban s or

→ %udden failure of the critical engine with delayed corrective action.

&he ailerons must be designed for the loads to which they are subjectedL

In the neutral position during symmetrical flight conditions# and

6y the following deflections ;e!cept as limited by pilot effort


56/84

23


and any indirect effect imposed by limitations in the output side of the control

system ;for e!ample, stalling tor"ue or ma!imum rate obtainable by a power

control system.

Ma5i 0 (i!c, c'n!r'l "i#(lace en! a! * AL

&he airplane is assumed to be flying in steady level flight and the coc pit pitch

control is suddenly moved to obtain e!treme nose up pitching acceleration. In

defining the tail load, the response of the airplane must be ta en into account.

Airplane loads that occur subse"uent to the time when normal acceleration at

the c.g. e!ceeds the positive limit maneuvering load or the resulting tail plane

normal load reaches its ma!imum, whichever occurs first, need not beconsidered.

S(eci ie" c'n!r'l "i#(lace en!:

A chec ed maneuver, based on a rational pitching control motion vs. time

profile, must be established in which the design limit load factor will not be

e!ceeded. Hnless lesser values cannot be e!ceeded, the airplane response must

result in pitching accelerations not less than the followingLA positive pitching acceleration ;nose up< is assumed to be reached

concurrently with the airplane load factor of './. &he positive

acceleration must be e"ual to at least -7n;n+'


57/84

23


JoriEontal balancing surfaces must be designed for the balancing loads

occurring at any point on the limit maneuvering envelope and in the flap

conditions

It is not re"uired to balance the rudder because it will not deflect due to

gravity.

Aileron will defect in vice versa direction so it is doesn’t re"uire balancing

load.


58/84

23


6. DETAILED STRUCTURAL LA


59/84

23


FUNCTION OF THE STRUCTURE:

&he primary functions of an aircraft’s structure can be basically bro en down

into the followingL

&o transmit and resist applied loads.&o provide and maintain aerodynamic shape.&o protect its crew, passenger, payload, systems, etc.

Kor the vast majority of aircraft, this leads to use of a semi+monoco"ue design

;i.e. a thin, stressed outershell with additional stiffening members< for the wing,

fuselage V empennage. &hese notes will discussthe structural layout

possibilities for each of these main areas, i.e. wing, fuselage V empennage.

+ING STRUCTURAL LA


60/84

23


→ %tringers

Ba#ic F0nc!i'n# ' +in$ S!r0c!0ral Me -er#

&he structural functions of each of these types of members may be

considered independently asL

SPARS

Korm the main span wise beam&ransmit bending and torsional loadsBroduce a closed+cell structure to provide resistance to torsion, shear and

tension loads.

In (ar!ic0lar:• :ebs – resist shear and torsional loads and help to stabiliEe the s in.• Klanges + resist the compressive loads caused by wing bending.

S IN

&o form impermeable aerodynamics surface&ransmit aerodynamic forces to ribs V stringers9esist shear torsion loads ;with spar webs


61/84

23


STRINGERS

Increase s in panel buc ling strength by dividing into smaller length

sections.

9eact a!ial bending loads

RIBS

8aintain the aerodynamic shapeAct along with the s in to resist the distributed aerodynamic pressure

loadsistribute concentrated loads into the structure V redistribute stress

around any discontinuitiesIncrease the column buc ling strength of the stringers through end

restraintIncrease the s in panel buc ling strength .

SPARS

&hese usually comprise thin aluminium alloy webs and flanges, sometimes with

separate verticalstiffeners riveted on to the webs.

Ty(e# ' #(ar# L

In the case of a two or three spar bo! beam layout, the front spar should be

located as far forwardas possible to ma!imiEe the wing bo! siEe, though this is

subject to there beingL

Ade"uate wing depth for reacting vertical shear loads.Ade"uate nose space for 4? devices, de+icing e"uipment, etc.

&his generally results in the front spar being located at ')O to '2O of the chord

length. Kor a single spar +nose layout, the spar will usually located at the

ma!imum thic ness position of the aerofoil section;typically between -/O

V0/O along the chord length


62/84

23


between about ((Oand 5/O of the chord length. If any intermediate spars

areused, they would tend to be spaced uniformly unless there are specific pic +

up point re"uirements.

RIBS

Kor a typical two spar layout, the ribs are usually formed in three parts from

sheet metal by the use ofpresses Vdies. Klanges are incorporated around the

edges so that they can be riveted to the s in and thespar webs. 1ut+out are

necessary around the edges to allow for the stringers to pass through.

4ighteningholes are usually cut into the rib bodies to reduce the rib weight and

also to allow for the passage ofcontrol runs, fuel, electrics, etc.

9ib bul heads do not include any lightening holes and are used at fuel tan

ends, wing cran locationsand attachment support areas. &he rib should be

ideally spaced to ensure ade"uate overall buc lingsupport to spar flanges. In

reality, however, their positioning is also influenced byL

Kacilitating attachment points for control surfaces, flaps, slats, spoiler

hinges, power plants,stores, undercarriage attachment etc.Bositioning of fuel tan ends, re"uiring closing ribs.A structural need to avoid local shear or compression buc ling# there are

several different possibilities regarding the alignment of the ribs on swept+

wing aircraft is a hybrid design in which one or more inner ribs are aligned

with the main a!is while the remainders are aligned perpendicularly to the

rear spar and usually the preferred option but presents several structural problems in the root region also ives good torsional stiffness characteristics

but results in heavy ribs and comple! connections.

S IN

&he s in tends to be riveted to the rib flanges and stringers, using countersun

rivets to reduce drag. It isusually pre+formed at the leading edges, where the

curvature is large due to aerodynamicconsiderations.

FUSELAGE STRUCTURE


63/84

23


&he fundamental purpose of the fuselage structure is to provide an envelope to

support the payload,crew, e"uipment, systems and ;possibly< the power+plant.

Kurthermore, it must react against the in+flightmanoeuvre, pressurisation and

gust loads# also the landing gear and possibly any power+plant loads. It must be

also be able to transmit control and trimming loads from the stability and

control surfacesthroughout the rest of the structure

Kuselage contributes very little to lift and produces more drag but it is an

important structural member$component. It is the connecting member to all load

producing components such as wing, horiEontal tail, vertical tail, landing gear

etc. and thus redistributes the load. It also serves the purpose of housing or accommodating practically all e"uipment, accessories and systems in addition

to carrying the payload. 6ecause of large amount of e"uipment inside the

fuselage, it is necessary to provide sufficient number of cutouts in the fuselage

for access and inspection purposes. &hese cutouts and discontinuities result in

fuselage design being more complicated, less precise and often less efficient in

design.As a common member to which other components are attached, thereby

transmitting the loads, fuselage can be considered as a long hollow beam. &he

reactions produced by the wing, tail or landing gear may be considered as

concentrated loads at the respective attachment points. &he balancing reactions

are provided by the inertia forces contributed by the weight of the fuselage

structure and the various components inside the fuselage. &hese reaction forcesare distributed all along the length of the fuselage, though need not be

uniformly. Hnli e the wing, which is subjected to mainly unsymmetrical load,

the fuselage is much simpler for structural analysis due to its symmetrical cross+

section and symmetrical loading. &he main load in the case of fuselage is the

shear load because the load acting on the wing is transferred to the fuselage s in

in the form of shear only. &he structural design of both wing and fuselage begin

with shear force and bending moment diagrams for the respective members. &he


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ma!imum bending stress produced in each of them is chec ed to be less than

the yield stress of the material chosen for the respective member..

F0#ela$e Lay'0! C'nce(!#

&here are two main categories of layout concept in common use#

8ass boom and longeron layout%emi+monoco"ue layout

Ma## B'' L'n$er'n lay'0!

&his is fundamentally very similar to the mass+boom wing+bo! concept

discussed in previoussection. It is used when the overall structural loading is

relatively low or when there are e!tensive cut+outsin the shell. &he conceptcomprises four or more continuous heavy booms ;longeron


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framewor . %tringers and longerons preventtension and compression stresses

from bending thefuselage.&he s in is attached to the longerons, bul heads,and

other structural members and carries part of theload. &he fuselage s in thic ness

varies with the loadcarried and the stresses sustained at particular location.


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18. DESIGN OF SOME COMPONENTS OF +ING AND FUSELAGE


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DESIGN OF +ING COMPONENT −¿ SPAR:

:ing is the major lift producing surface. &herefore, the analysis has to be very

accurate. &he structural analysis of the wing by defining the primary loadcarrying member %pars is done below.

%pars are members which are basically used to carry the bending and shear

loads acting on the wing during flight. &here are two spars, one located at '(+

)/O of the chord nown as the front spar, the other located at 3/+5/O of the

chord nown as the rear spar. %ome of the functions of the spar includeL

&hey form the boundary to the fuel tan located in the wing.• &he spar flange ta es up the bending loads whereas the web carries the

shear loads.• &he rear spar provides a means of attaching the control surfaces on the

wing.

1onsidering these functions, the locations of the front and rear spar are fi!ed at

/.'5c and /.3(c respectively.&he spar design for the wing root has been ta en

because the ma!imum bending moment and shear force are at the root. It is

assumed that the flanges ta e up all the bending and the web ta es all the shear

effect. &he ma!imum bending moment for high angle of attac condition is

261419642.5 >m. &he ratio in which the spars ta e up the bending moment is

given as

/ ,

/ r= 2,

2

2r2

:here

h ' + height of front spar

h) + height of rear spar

/ , / r

=1.6186 2

1.4795 2

/ , / r

= 1.197


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/ , + / r = 261419642.5

&herefore,

8 f N ''2727-32.( >m

8 r N770/3-)).2( >m

&he yield tensile stress P y for Al Alloy ;Al 5/5(< is 9 .8 76 > MPa . &he area

of the flanges is determined using the relation

3 *= / A4

where 8 is bending moment ta en up by each spar,

A is the flange area of each spar,

E is the centroid distance of the area N h$).

Hsing the available values,

Area of front spar,

A f N /.-)-/77 m )

Area of rear spar,

A r N /.)7(-/ m )

AssumptionsL

& sections are chosen for top and bottom flanges of front and rear spars.6oth the

flanges are connected by a vertical stiffener through spot welding and

( , ( -

= 1

Krom the buc ling e"uation,

0 cr = 0.388 1 (( -

&- )2

the thic ness to width ratio of web( -&-

is found to be -.7(7'. Also from

“Analysis and design of flight vehicle structures by 69HJ>”, the flange to web

width ratio of the & section&, &-

= 1.8


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6y e"uating all the three values of the ratio in area of the section e"uation, the

dimensions of the spar can be found.

Di en#i'n# 'r r'n! #(ar:

b flange N /.0)3-- m

tflange NtwebN/.)-32( m

bweb N /.7-55' m

Di en#i'n# 'r rear #(ar:

b flange N /.0/5(7 m

tflange NtwebN /.))300 m

bweb N /.27307 m

DESIGN OF FUSELAGE COMPONENT −¿ STRINGER

&he circumference of the fuselage is 0-.'/) m. &o find the area of one stringer,

number of stringers per "uadrant is assumed to be 0. i.e. the total number of

stringers in the fuselage is '3. &he stringers are e"ually spaced around the

circumference of the fuselage.

S!rin$er S(acin$:


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&he stringers are symmetrically spaced on the fuselage with the spacing

calculate as shown below,

1ircumference of the fuselage N $D = $ ∗2∗6.86 = 43.102 m

&otal number of stringers N '3

&herefore the stringers are spaced at the interval of N 43.102

16 = 2.6939 m

S!rin$er area calc0la!i'n:

&he stress induced in the each stringer is calculated with the area eeping

constant in the stress term. &hen the ma!imum stress ;i.e. one which has larger

numerator< is e"uated with the yield strength of the material. Krom this area of

one stringer is calculated.

&he direct stress in each stringer produced by bending moments / x and

/ * is given by the e"uationL

3 = / 5 + 55

4+ / 6 + 66

x N /m2

:here

/ 5 = 33827996.63 N

/ 6 =(12 ρV 2 S( a ( 7 )× x ρ is density N'.))( g$m -

* is cruise velocity N )(/ m$s

%t is the tail area N 5'.3(5 m )

at is the slope of the lift curve N /./32' $deg

7 is the angle of yaw for asymmetric flight

7 = 0.7 nmax+457.2

V D

7 = 3.563 'eg

! is the distance between the aircraft c.g position and horiEontal tail c.g

positionN 79.

&hen,


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/ 6 = 23146604.65 Nm + 55 = + 66 = As(inger D

2

:here As(inger is the stringer area, is the diameter of the fuselage N '-.5) m.

/ x and

/ * reach their ma!imum only from the stringers ' to 0. &hus the

stresses are high only on these stringers. 1alculating stress for stringer ' to 0.

F N /, = N 3.23

3 2 = / 5 + 55

6 + / 6 + 66

5

&hen,

3 1 =1232798.71

As(inger N m

2

F N ).3)7, = N 3.-0

3 2 = / 5 + 55

6 + / 6 + 66

5

&hen,

3 2 =1462623.57

As(inger N m

2

F N 0.2(52, = N 0.2(52

3 3 = / 5 + 55

6 + / 6 + 66

5

&hen,

3 3 =1470322.836

As(inger N

mm2

F N 3.--52, = N ).3)(

3 4 = / 5 + 55

6 + / 6 + 66

5


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&hen, 3 4 =1251057.39

As(inger N

mm2

&he allowable stress in the stringer is0((./(-73) 8Ba for Al Alloy ;Al 5/5(


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11. MATERIAL SELECTION


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DESCRIPTION

Aircraft structures are basically unidirectional. &his means that one dimension,

the length, is much larger than the others + width or height. Kor e!ample, the

span of the wing and tail spars is much longer than their width and depth# the

ribs have a much larger chord length than height and$or width# a whole wing has

a span that is larger than its chords or thic ness# and the fuselage is much longer

than it is wide or high. ?ven a propeller has a diameter much larger than its

blade width and thic ness, etc.... Kor this simple reason, a designer chooses to

use unidirectional material when designing for an efficient strength to weight

structure.Hnidirectional materials are basically composed of thin, relatively fle!ible, long

fibers which are very strong in tension ;li e a thread, a rope, a stranded steel

wire cable, etc.


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wea ness because, as a whole, they become compression resistant as they help

each other to not buc le away. &he embedding is usually a lighter, softer WresinW

holding the fibers together and enabling them to ta e the re"uired compression

loads. &his is a very good structural material.

+OOD

Jistorically, wood has been used as the first unidirectional structural raw

material. &hey have to be tall and straight and their wood must be strong and

light. &he dar bands ;late wood< contain many fibers, whereas the light bands

;early wood< contain much more WresinW. &hus the wider the dar bands, thestronger and heavier the wood. If the dar bands are very narrow and the light

bands "uite wide, the wood is light but not very strong. &o get the most efficient

strength to weight ratio for wood we need a definite numbers of bands per inch.

%ome of our aircraft structures are two+dimensional ;length and width are large

with respect to thic ness


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no point in discussing &itanium + itXs simply too e!pensive


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As a rule of thumb, aluminium is three times heavier, but also three times

stronger than wood. %teel is again three times heavier and stronger than

aluminium.

STEEL

&he ne!t material to be considered for aircraft structure will thus be steel, which

has the same weight+to+strength ratio of wood or aluminium.

Apart from mild steel which is used for brac ets needing little strength, we are

mainly using a chrome+molybdenum alloy called AI%I 0'-@> or 0'0/. &he

common raw materials available are tubes and sheet metal. %teel, due to its highdensity, is not used as shear webs li e aluminium sheets or plywood. :here we

would need, say.'//W plywood, a ./-) inch aluminium sheet would be re"uired,

but only a ./'/ steel sheet would be re"uired, which is just too thin to handle

with any hope of a nice finish. &hat is why a steel fuselage uses tubes also as

diagonals to carry the shear in compression or tension and the whole structure is

then covered with fabric ;light weight< to give it the re"uired aerodynamic

shape or desired loo . It must be noted that this method involves two

techni"uesL steel wor and fabric covering. .

COMPOSITE MATERIALS

&he designer of composite aircraft simply uses fibers in the desired direction

e!actly where and in the amount re"uired. &he fibers are embedded in resin to

hold them in place and provide the re"uired support against buc ling. Instead of

plywood or sheet metal which allows single curvature only, the composite

designer uses cloth where the fibers are laid in two directions .;the woven thread

and weft< also embedded in resin. &his has the advantage of freedom of shape in

double curvature as re"uired by optimum aerodynamic shapes and for very

appealing loo ;importance of aesthetics


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&odayXs fibers ;glass, nylon, Gevlar, carbon, whis ers or single crystal fibers of

various chemical compositions< are very strong, thus the structure becomes very

light. &he drawbac is very little stiffness. &he structure needs stiffening which

is achieved either by the usual discreet stiffeners, +or more elegantly with a

sandwich structureL two layers of thin uni+ or bi+directional fibers are held apart

by a lightweight core ;foam or WhoneycombW. Ti!ani0 : A very e!pensive material. *ery tough material and difficult to

machine.

7. Car-'n Fi-er#: %till very e!pensive materials.


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9. e&lar Fi-er#: *ery e!pensive and also critical to wor with because it is

hard to Wsoa W in the resin.

A number of properties are important to the selection of materials for an aircraft

structure. &he selection of the best material depends upon the application.

Kactors to be considered include yield and ultimate strength, stiffness, density,

fracture toughness, fatigue, crac resistance, temperature limits, producibility,

repairability, cost and availability. &he gust loads, landing impact and vibrations

of the engine and propeller cause fatigue failure which is the single most

common cause of aircraft material failure.

Kor most aerospace materials, creep is a problem only at the elevatedtemperature. Jowever some titanium plastics and composites will e!hibit creep

at room temperatures.

&a ing all the above factors into considerations, the following aluminium alloys

which have e!cellent strength to weight ratio and are abundant in nature are

considered.

S.N' Al0 ini0 All'y


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DESIGN REPORT


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De#i$n Re('r!:


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ADP-II AMPHIBIANS AIRCRAFTPara e!er# @al0e#

%pan 5/ m

Blanform area (07 m )

Aspect ratio 2.7)

?mpty weight '3//// g

8a!imum ta eoff weight -(7--' g

@swald efficiency factor /.52()

1hord at root ').(0 m

1hord at tip -.'-( m

&aper ratio /.)(

%weepbac angle -.20-

:ing loading 3(0 g$m)

Bower delivered by motor (0 hp

&hrust+to+weight ratio /.'(2

9ate of climb (.)) m$s

?ndurance - hours

9ange 0(// m

%tall speed 37.00 m$s

4anding distance '0(/ m

&a eoff distance ))55.03 m

8a!imum Qve 4oad factor -

8a!imum –ve 4oad factor '.)

esign dive speed -').( m$s

4ift coefficient;flaps down< '.'-2

8inimum radius of turn 322.372 m

8a!imum bending moment )3'0'730).( >m

Kront spar bending moment ''2727-32.( >m

9ear spar bending moment 770/3-)).2( >m

6ending moment in fuselage --2)5773.3- >m


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T,ree &ie% "i$ra :


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CONCLUSION:

&he structural design of the Jeavy+lift military cargo aircraft which is a

continuation of the aerodynamic design part carried out last semester is

completed satisfactorily. &he aeroplane has gone through many design

modifications since its early conceptual designs e!pected, among these was a

growth in weight.

&o ensure continued growth in payload and the reduced cost of cargo

operations, improvements in methods, e"uipment and terminal facilities will be

re"uired in order to reduce cargo handling costs and aircraft ground time and to

provide improved service for the shippers.:e have enough hard wor for this design project. A design never gets

completed in a flutter sense but it is one step further towards ideal system. 6ut

during the design of this aircraft, we learnt a lot about aeronautics and its

implications when applied to an aircraft design.

BIBLIOGRAPH

Project 2 Azar ADP ASS

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