180745548 Aircraft Design Project 2

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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT

ACKNOWLEDGEMENT

We take this opportunity to thank our beloved Chairperson Dr.S.Thangam

Meganathan, Rajalakshmi Engineering College, Thanalam, for providing

good Infrastructure with regards to our project and giving enthusiasm in

pursuing the studies.

We also express our thanks to our beloved Principal, Dr.G. Thanigaiaras!,

who has been a constant source of inspiration and guidance throughout our

course.

We would like to thank, Mr."ogesh K!mar Sinha, #ea o$ the e%artment,

De%artment o$ Aerona!ti&al Engineering, for allowing us to take up this

project and his timely suggestions.

We express our sense of gratitude to Mr.Karthik, 'roje&t G!ie for his help,

through provoking discussions, invigorating suggestions extended to us with

immense care, eal throughout our work.

I would like to express my gratitude to my parents for their hard work and

continuous support, which helped me in pursuing higher studies. !ppreciation is

also extended to all the faculty and students that I have had the privilege of

working with throughout my years of college at RA(ALAKS#M)

ENG)NEER)NG COLLEGE. "ast but not the least, I would like to thank my

friends for their constant support and help in all my endeavors



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)NDE*

S.NO CONTENTS 'AGE NO.

# Introduction $

% &'n (iagram #)

) *ust &'n diagram %+

Critical loading performance and final &'n

diagram%

$ -tructural design study theory approach %/

0 "oad estimation on wings )%

1 "oad estimation on fuselage

/2alancing and maneuvering loads on tail plane,

rudder and aileron loads3

3 (etailed structural layouts $

#+(esign of some components of wing and

fuselage0%

## 4aterial selection 03

#% (esign report 10

#) 5hree view diagram 13

# Conclusion /+

#$ 2ibliography /+



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NOMENCLAT+RE

!.6. ' !spect 6atio

b ' Wing -pan 7m8

C ' Chord of the !irfoil 7m8

C root ' Chord at 6oot 7m8

C tip ' Chord at 5ip 7m8

C ' 4ean !erodynamic Chord 7m8

Cd ' (rag Co'efficient

Cd,+ ' 9ero "ift (rag Co'efficient

C p ' -pecific fuel consumption 7lbs:hp:hr8

C" ' "ift Co'efficient

( ' (rag 7;8

< ' <ndurance 7hr8

e ' =swald efficiency

" ' "ift 7;8

7":(8loiter ' "ift'to'drag ratio at loiter

7":(8cruise ' "ift'to'drag ratio at cruise

4 ' 4ach number of aircraft

4ff ' 4ission fuel fraction

6 ' 6ange 7km8

6e ' 6eynolds ;umber

- ' Wing !rea 7m>8

5 ' 5hrust 7;8

&cruise ' &elocity at cruise 7m:s8

&stall ' &elocity at stall 7m:s8&t ' &elocity at touch down 7m:s8

Wcrew ' Crew weight 7kg8

Wempty ' <mpty weight of aircraft 7kg8

Wfuel ' Weight of fuel 7kg8

W payload ' Payload of aircraft 7kg8

W+ ' =verall weight of aircraft 7kg8

W:- ' Wing loading 7kg:m>8 ρ∞ ' (ensity of air 7kg:m?8



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!stringer ' Cross sectional area of stringers

! ' 5otal cross sectional area

! spar ' Cross sectional area of spar

at'-lope of the C" vs. @ curve for a horiontal tail

a'(istance of the front spar from the nose of the aircraft

bw'Width of the web

bf 'Width of the flange

Ixx ' -econd moment of area about A axis

I ' -econd moment of area about 9 axis

B ' *ust alleviation factor

n max ' 4aximum load factor

tw ' 5hickness of the web

tf ' 5hickness of the flange

5 ' 5orue

D ' *ust velocity

&cruise ' Cruise velocity

&s ' -talling velocity

' !ngle of Eaw

.



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-. )NTROD+CT)ON



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M)L)TAR" TRANS'ORT A)RCRAT

4ilitary transport aircraft or military cargo aircraft are typically fixed and rotary

wing cargo aircraft which are used to deliver troops, weapons and other military

euipment by a variety of methods to any area of military operations around the

surface of the planet, usually outside of the commercial flight routes

in uncontrolled airspace. =riginally derived from bombers, military transport

aircraft were used for delivering airborne forces during the -econd World War

and towing military gliders. -ome military transport aircraft are tasked to

perform multi'role duties such as aerial refueling and, tactical, operational and

strategic airlifts onto unprepared runways, or those constructed by engineers.

CLASS))CAT)ON O M)L)TAR" TRANS'ORTS

Fixed wing transport aircraft

5ransport Gelicopters

What is an Airli$t/

!n airlift is the organied delivery of supplies or personnel primarily

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

airlifting. 5ypically, strategic airlifting involves moving material long distances

7such as across or off the continent or theater8, whereas a tactical airlift focuses

on deploying resources and material into a specific location with high precision.

(epending on the situation, airlifted supplies can be delivered by a variety of

means. When the destination and surrounding airspace is considered secure, the

aircraft will land at an appropriate airport or airbase to have its cargo unloaded

on the ground. When landing the craft, or distributing the supplies to a certain

area from a landing one by surface transportation is not an option, the cargo

aircraft can drop them in mid'flight using parachutes attached to the supply

containers in uestion. When there is a broad area available where the intended



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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. Gowever,

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

dangerous to land in, a "ow !ltitude Parachute <xtraction -ystem drop is used.

CLASS))CAT)ON O A)RL)TS

-56!5<*IC !I6"IF5

5!C5IC!" !I6"IF5

STRATEG)C A)RL)T

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

or personnel over long distances. 5ypically, this involves airlifting the reuired

items between two airbases which are not in the same vicinity. 5his

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

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

considered strategi& airli$ters. 5his contrasts with tactical airlifters, such as

the C'#)+ Gercules, which can normally only move supplies within a

given theater of operations.

<A!4P"<H "ockheed C'$ *alaxy, !ntonov !n'#%

TACT)CAL A)RL)T

5actical airlift is a military term for the airborne transportation of supplies andeuipment within a theatre of operations 7in contrast to strategic airlift8. !ircraft

which perform this role are referred to as ta&ti&al airli$ters. 5hese are

typically turboprop aircraft, and feature short landing and take'off distances and

low'pressure tires allowing operations from small or poorly'prepared airstrips.

While they lack the speed and range of strategic airlifters 7which are

typically jet'powered8, these capabilities are invaluable within war ones.

"arger helicopters such as the CG'1 Chinook and 4il 4i'%0 can also be used



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to airlift men and euipment. Gelicopters have the advantage that they do not

reuire a landing strip and that euipment can often be suspended below the

aircraft allowing it to be delivered without landing but are highly inefficient.

5actical airlift aircraft are designed to be maneuverable, allowing low'altitude

flight to avoid detection by radar and for the airdropping of supplies. 4ost are

fitted with defensive aids systems to protect them from attack by surface'to'air

missiles.

<A!4P"<H Gercules C'#)+, "ockheed C'## -tarlifter

DES)GN O AN A)R'LANE

!irplane design is both an art and a science. Its the intellectual engineering

process of creating on paper 7or on a computer screen8 a flying machine to

meet certain specifications and reuirements established by potential

users 7or as perceived by the manufacturer8 and

pioneer innovative, new ideas and technology

5he 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 7generally proprietary8

design procedure and databases 7hand books etc8 that are a part of every

airplane manufacturer.

'#ASES O A)R'LANE DES)GN

5he complete design process has gone through three distinct phases that are

carried out in seuence. 5hey are

Conceptual design



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Preliminary design

(etailed design

CONCE'T+AL DES)GN

5he design process starts with a set of specifications 7reuirements8for a new

airplane, or much less freuently as the response to the desire to implement

some pioneering, innovative new ideas and technology. In either case, there is a

rather concrete good towards which the designers are aiming. 5he first steps

towards achieving that goal constitute the conceptual design phase. Gere, within

a certain somewhat fuy latitude, the overall shape, sie, weight and

performance of the new design are determined.

5he product of the conceptual design phase is a layout on a paper or on a

computer screen8 of the airplane configuration. 2ut one has to visualie this

drawing as one with flexible lines, capable of being slightly changed during the

preliminary design phase. Gowever the conceptual design phase determines

such fundamental aspects as the shape of the wings 7swept back, swept forward

or straight8, the location of the wings related to the fuselage, the shape and

location of the horiontal and vertical tail, the use of a engine sie and

placement etc, the major drivers during the conceptual design process are

aerodynamics, propulsion and flight performance.

-tructural and context system considerations are not dealt with in any detail.

Gowever they are not totally absent. (uring the conceptual design phase the

designer is influenced by such ualitative as the increased structural loadsimposed by a high horiontal tail location trough the fuselage and the



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difficulties associated with cutouts in the wing structure if the landing gear are

to be retracted into the wing rather than the fuselage or engine nacelle. ;o part

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

'REL)M)NAR" DES)GN

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

configuration layout 7indeed, 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 take place. (uring this phase also, substantial wind tunnel

testing will be carried out and major computational fluid dynamics 7CF(8

calculations of the computer flow fluid over the airplane configurations are

done.

Its possible that the wind tunnel tests the CF( calculations will in cover some

undesirable aerodynamic interference or some unexpected stability problems

which will promote change to the configuration layout. !t the end of

preliminary design phase the airplane configuration is froen and preciously

defined. 5he drawing process called lofting is carried out which mathematically

models the precise shape of the outside skin of the airplane making certain that

all sections of the aircraft property fit together

5he end of the preliminary design phase brings a major concept to commit the

manufacture of the airplane or not. 5he 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.

DETA)L DES)GN



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5he detail design phase is literally the nuts and bolts phase of airplane design.

5he aerodynamic, propulsion, structures performance and flight control analysis

have all been finished with the preliminary design phase. 5he airplane is now

simply a machine to be fabricated. 5he pressure design of each individual rib,

spar and section of skin now take place. 5he sie of number and location of

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

developed. !nd these are just a few of the many detailed reuirements during

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

fabricated.

O+TL)NE A)RCRAT DES)GN 'RO(ECT 0

5he structural design of the aircraft which is done in aircraft design project %

involvesH

(etermination of loads acting on aircraft

• &'n diagram for the design study

• *ust and maneuverability envelopes

• -chrenks Curve

• Critical loading performance and final &'n graph calculation

(etermination of loads acting on individual structures

• -tructural design study 5heory approach

• "oad estimation of wings

• "oad estimation of fuselage.

• 4aterial -election for structural members



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• (etailed structural layouts

• (esign of some components of wings, fuselage

'arameters taken $rom air&ra$t esign %roje&t -

'arameters 1al!es

Wing loading7kg:m%8 0$.$

4ach number %.1%

5hrust to weight ratio +.#$$%0

!spect ratio /.3%!ltitude7km8 #)

4aximum "ift coefficient #./%/)

Wing span7m8 1+

Wing planform area7m%8 $3

Fuel weight7kg8 /10#3

<ngine weight7kg8 )0)+

=verall weight7kg8 )$3))#

Cruise speed7Bm:hr8 3++

-talling speed7Bm:hr8 %$+

-ervice speed7km8 #)

6oot chord7m8 #%.$

5ip chord7m8 ).#)$

Juarter chord sweep angle7deg8 )./)o

4ean aerodynamic chord7m8 $.)

5hrust per engine7B;8 #)1

6ange7km8 $++

Payload7kg8 3++++



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0. 12n Diagram



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)NTROD+CT)ON

!irplanes may be subjected to a variety of loading conditions in flight. 5he

structural 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. !s 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.

1elo&it3 4Loa a&tor 512n6 iagram

5he control of weight in aircraft design is of extreme importance. Increases in

weight reuire stronger structures to support them, which in turn lead to further

increases in weight and so on. <xcess of structural weight mean lesser amounts

of payload, thereby affecting the economic viability of the aircraft. 5he aircraft

designer is therefore constantly seeking to pare his aircrafts weight to the

minimum compatible with safety. Gowever, to ensure general minimum

standards of strength and safety, airworthiness regulations 7!v.P.31+ and

2C!68 lay down several factors which the primary structure of the aircraft

must satisfy. 5hese are the

• Limit loa, which is the maximum load that the aircraft is expected to

experience in normal operation.

• 'roo$ loa, which is the product of the limit load and the %roo$ $a&tor 7#.+'

#.%$8, and

• +ltimate loa, which is the product of the limit load and the !ltimate $a&tor

7usually #.$8. 5he aircrafts structure must withstand the proof load without

detrimental distortion and should not fail until the ultimate load has been

achieved.



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5he basic strength and fight performance limits for a particular aircraft are

selected by the airworthiness authorities and are contained in the flight envelope

or 12n iagram.

5here are two types of & n diagram for military airplanes H

&n maneuver diagram and

&n gust diagram

1 4 n MANE+1ER D)AGRAM

5he 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

359331+10000n¿



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lim ¿( pos)≥2.164

n¿

5he maximum positive limit load factor for military transport aircraft should be

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

n¿

5he maximum negative limit load factor is given by

¬¿¿

lim ¿( pos)

¿lim ¿¿¿n¿

¬¿¿¿

lim ¿¿¿n¿

¬¿¿¿

lim ¿¿¿n¿

5here are four important speeds used in the & n diagram

# g stall speed &-

(esign maneuvering speed &!

(esign cruise speed &C

(esign diving speed &(

'ositi7e - 4 g stall s%ee 1S

V S=√ 2

ρ C Nmax

W

S

C Nmax=1.1×C Lmax

C Nmax=1.1×1.138

C Nmax=1.252

V S=√ 2

1.125×1.252×654.5

V S=30.48m /s

Negati7e - 4 g stall s%ee 1Sneg



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¬¿¿¿¿

Nmax¿ ρC ¿

2¿

S¬¿=√ ¿V ¿

¬¿¿¬¿¿¿

Lmax¿¿

Nmax¿

C ¿¬¿¿¿

Nmax¿

C ¿¬¿¿¿

Nmax¿

C ¿

S¬¿=

√ 2

1.125×0.594

× 654.5

V ¿

S¬¿=44.26m / sV ¿

Design Mane!7ering s%ee 1A $or %ositi7e loa $a&tor

2nlim ¿( pos)

ρ C Nmax

W

S

V A=√ ¿

V A=√ 2×3

1.125×1.252×654.5

V A=52.80m/ s

Design Mane!7ering s%ee 18 $or negati7e loa $a&tor



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¬¿¿¿¿¬¿¿¿

¿ Nmax¿

¿lim ¿¿¿

2n¿¿

V B=√ ¿

V B=√ 2×1.2

1.125×0.594 ×654.5

V B=48.48m /s

Design Cr!ise s%ee 1C

From !ircraft (esign Project #,

&C K &cruise K 3++ km:hr

&C K %$+ m:s

Design Di7ing S%ee 1D

5he design diving speed must satisfy the following relationshipV D ≥1.25V cruise

V D=1.25× %$+

V D=312.5m /s

C!r7e OA

5he velocity along the curve =! is given by the expression

V Sn=

√ 2n

ρC Nmax

W

S

From this expression the load factor along the curve =! is given by

n= ρ C NmaxV

2

2

1

W

S

n=1.125×1.252V

2

2

1

654.5

n=1.076×10−3

V 2



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1elo&it3 m9s 'ositi7e Loa a&tor n

+ +

#+ +.#+10

%+ +.)+

)+ +.30/

+ #.1%#0

$+ %.03

$%./+ )

C!r7e OG

5he negative load factor along the curve =* is given by the expression

¬¿¿

¿V 2

¿ Nmax¿

ρ C ¿¬¿=¿n¿

¬¿=1.125×0.594V

2

2

1

654.5

n¿

¬¿=5.10504 ×10−4

V 2

n¿

1elo&it3 m9s Negati7e Loa a&tor nneg

+ +

#+ '+.+$#+$

%+ '+.%+%

)+ '+.$3$

+ '+./#0/

$+ '#.%10%0



2


/./ '#.%



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:. G+ST 12n D)AGRAM





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μg=

2( W

S ) ρ C C Lα

where,

k g−¿ *ust alleviation factor U g−¿ (erived gust velocityV B−¿ (esign speed for maximum gust intensityV c−¿ (esign cruise velocityV D−¿ (esign diving velocityC Lα −¿ =verall lift curve slope rad'#

C −¿ Wing mean geometric chordW

S =654.5

kg

m2

ρ=1.225 kg/m3

C Lα =10.03 C =5.434 m

μg= 2×654.5

1.225×5.434×10.031=19.334

k g=0.88×19.334

5.3+19.334 =0.6926

Construction of gust load factor line for speed V B=52.80m/s 7take

U g=20.11m /s 8

lim ¿=1.746

+n¿

lim ¿=0.2543

−n¿

Construction of gust load factor line for speed V c=250m /s 7take

s

U g=15.24m/¿¿

lim ¿=3.525

+n¿

lim ¿=−1.525

−n¿

Construction of gust load factor line for speed V c=312.50m /s 7take

U g=40m /s 8

lim ¿=2.5025

+n¿

lim ¿=−0.5035

−n¿



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;. CR)T)CAL LOAD)NG 'ERORMANCE

AND )NAL 12n D)AGRAM



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CR)T)CAL LOAD)NG 'ERORANCE

5he greatest air loads on an aircraft usually comes from the generation of lift

during high'g maneuvers. <ven the fuselage is almost always structurally sied

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

fuselage. !ircraft load factor 7n8 expresses the maneuvering of an aircraft as a

standard acceleration due to gravity.

!t lower speeds the highest load factor of an aircraft may experience is limited

by the maximum lift available. !t higher speeds the maximum load factor is

limited to some arbitrary value based upon the expected use of the aircraft. 5he

maximum lift load factor euals #.+ at levels flight stall speed. 5his is the

slowest speed at which the maximum load can be reached without stalling.

5he aircraft maximum speed, or dive speed at right of the &'n diagram

represents the maximum dynamic pressure and maximum load factor is clearly

important for structural siing. !t this condition, the aircraft is at fairly low

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

approximately vertical in the body axis. 5he most common maneuvers that we

focused are,

"evel turn

Pull up

Pull down

Climb

Le7el t!rn

5he value of minimum radius of turn is given by the formula,

!min=4 k (W

S )g ρ( " W )√1−4 kC D

0( " W )5he 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 known values,

!min=688.698m

n !min=¿ #.)$#

'!ll2!% Mane!7er

!= V ∞

2

g (n−1)

-ubstituting the known values and 6 K )$++ m

n=2.82

'!ll2o<n Mane!7er

!= 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=0.82

Clim=

( "

W )−# ¿2−( 4C Do

$eA!)}0.5

¿

[( "

W )−#

]+{¿n=¿

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

-ubstituting the known values

n=1.688

Mane!7er Loa a&tor n

"evel turn #.)$#



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Pull'up %./%

Pull'down +./%

Climb #.0//

inal 12n Diagram



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>. STR+CT+RAL DES)GN ST+D" 4

T#EOR" A''ROAC#



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STR+CT+RAL DES)GN ST+D" 4 T#EOR" A''ROAC#

!ircraft loads are those forces and loadings applied to the airplanes structural

components to establish the strength level of the complete airplane. 5hese

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

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

during other operations such as catapulted take'offs, arrested landings, or

landings in water.

5he determination of design loads involves a study of the air pressures and

inertia forces during certain prescribed maneuvers, either in the air or on the

ground. -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 examination of some of the criteria.

Important consideration in determining the extent of the load analysis is the

amount of structural weight involved. ! fairly detailed analysis may be

necessary when computing operating loads on such items as movable surfaces,

doors, landing gears, etc. proper operation of the system reuires an accurate

prediction of the loads.

!ircraft loads is the science of determining the loads that an aircraft structure

must be designed to withstand. ! large part of the forces that make up design

loads are the forces resulting from the flow of air about the airplanes surfaces'the same forces that enable flight and control of the aircraft.

Loa $a&tors

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

airplane. (uring maneuvers or flight through turbulent 7gusty8 air, however,

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

the airplane structure. 5he amount of additional loads depends on the severity of





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STR+CT+RAL DES)GN CR)TER)A

5he structural criteria define the types of maneuvers, speed, useful loads,

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

5hese 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. 5he basic

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

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



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?. LOAD EST)MAT)ON ON W)NGS



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Des&ri%tion

5he solution methods which follow <ulers beam bending theory

7N:yK4:IK<:68 use the bending moment values to determine the stresses

developed at a particular section of the beam due to the combination of

aerodynamic and structural loads in the transverse direction. 4ost engineering

solution methods for structural mechanics problems 7both exact and

approximate methods8 use the shear force and bending moment euations to

determine the deflection and slope at a particular section of the beam.

5herefore, these euations are to be obtained as analytical expressions in termsof span wise location. 5he bending moment produced here is about the

longitudinal 7x8 axis.

Loas a&ting on <ing

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

5here are three primary loads acting on a wing structure in transverse direction

which can cause considerable shear forces and bending moments on it. 5hey are

as followsH

"ift force 7given by -chrenks curve8

-elf'weight of the wing

Weight of the power plant

Weight of the fuel in the wing

Shear $or&e an =ening moment iagrams !e to loas along trans7erse

ire&tion at &r!ise &onition

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

attack and sweep along the span. -chrenks curve defines this lift distribution

over the wing span of an aircraft, also called simply as "ift (istribution Curve.



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-chrenks Curve is given by

*= *

1+ *

2

2

wherey# is "inear &ariation of lift along semi wing span also named as "#

y% is <lliptic "ift (istribution along the wing span also named as "%

Linear li$t istri=!tion 5tra%e@i!m6

"ift at root

Lroo( =

ρV 2C Lcroo(

2

"root K /+#01./) ;:m

"ift at tip

L(ip= ρV

2C L c (ip

2

"tip K %++#.30 ;:m

2y representing this lift at sections of root and tip we can get the euation for

the wing.

<uation of linear lift distribution for starboard wing

y# K '#1#1.//1x O /+#01./)

<uation of linear lift distribution for port wing we have to replace x by x in

general,

y# K #1#1.//1x O /+#01./)

For the -chrenks curve we only consider half of the linear distribution of lift

and hence we derive y#:%

*1

2 K '/$/.3#x O++/).3#$



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0 5 10 15 20 25 30 35 40

0

10000

20000

30000

40000

50000

60000

70000

80000

90000

Linear variation of ift aon! se"i #in! span

Haf #in! span "

Lift per Len!t$ %&'"(

Elli%ti& Li$t Distri=!tion

5wice the area under the curve or line will give the lift which will be reuired to

overcome weight

Considering an elliptic lift distribution we get

L

2=

W

2 =

$a&

4

A=$a&

4

Where

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

"ift at tip,

&=4W

2 $a=64117.38 N /¿ m

<uation of elliptic lift,

*2=

√

&2(1−

x2

a

2)



2


*2=1831.925 √ (1225− x

2)

*2

2 =915.962 √ (1225− x

2)

0 5 10 15 20 25 30 35 40

0

10000

20000

30000

40000

50000

60000

70000

Eiptica ift )istri*+tion aon! se"i #in! span

Haf #in! span "


Constr!&tion o$ S&hrenks C!r7e

-chrenks Curve is given by,

*= *

1+ *

2

2

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

-ubstituting different values for x we can get the lift distribution for the wing

semi span



2


0 5 10 15 20 25 30 35 40

0

10000

20000

30000

40000

50000

60000

70000

80000

,c$ren.s c+rve for se"i #in! span

Haf #in! span "


S&hrenks &!r7e

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

0

10000

20000

30000

40000

50000

60000

70000

80000

,c$ren.s c+rve for f+ #in! span

/in! span "




2


Sel$2Weight o$ <ing 53:6

-elf'weight of the wing,

W W+N#

W ¿

=0.25

WWI;*K +.%$)$3))#3./#

WWI;*K //#%$3 ;

W portwing K ' +0%3 ; 7!cting (ownwards8

WstarboardK ' +0%3 ; 7!cting (ownwards8

!ssuming parabolic weight distribution

*3=k

( x−

&

2 )2

where b wing span

When we integrate from xK+ 7root location8 to xKb 7tip location8 we get the net

weight of port wing.

−440629=∫0

35

k ( x−&

2 )2

'x

kK '#%.))%$ *

3=−12.3325 ( x−35 )2

0 5 10 15 20 25 30 35 40

-16000

-14000

-12000

-10000

-8000

-6000

-4000

-2000

0

,ef #ei!$t of #in!

Haf #in! span "

/in! #ei!$t per +nit Len!t$ %&'"(



2


!el <eight in the <ing

5his design has fuel in the wing so we have to consider the weight of the fuel in

one wing.

W ,ue%-ing

2 =

104780.91

2 kg=52390.45 kg

W ,ue%-ing=513950.41 N

!gain by using general formula for straight line yK mx O 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+e #ei!$t in #in!

Haf #in! span "

F+e #ei!$t per Len!t$ %&'"(

'o<er %lant <eight

Power plant is assumed to be a point load,

WppK)0)+ kg K )$0#+.) ;

!cting at xK / m and xK %+ m from the root.



2


0 5 10 15 20 25 30 35 40

-60000

-40000

-20000

0

20000

40000

60000

80000

100000

Overa Loa) )istri*+tion on #in!

Haf #in! span "

Force per Len!t$ %&'"(

Loas sim%li$ie as %oint loas

C!r7e 9

&om%onent

Area en&lose 9 str!&t!ral

<eight 5N8

Centroi

5$rom <ing root6

y#:% #1$)01#.)%$ # m

y%:% #10%$#/.$)1 #./ m

Wing +0%3 /.1$ m

Fuel $#)3$+.# #+.$+# m

Power plant )$0#+.) /m, %+m



2


Rea&tion $or&e an 8ening moment &al&!lations

5he wing is fixed at one end and free at other end.

∑V =0 ,

5hen,

&!'#1$)01#.)%$'#10%$#/.$)1O+0%3O$#)3$+.#O)$0#+.)O)$0#+K+

&!K %3+)3+.#$% ;

∑ / =0 ,

5hen,

4! ' 7#1$)01#.)%$#8 ' 7#10%$#/.$)1#./8 O 7+0%3/.1$8 O

7$#)3$+.##+.$+#8 O 7)$0#+.)/8 O 7)$0#+.)%+8 K +

4! K 261419642.5 ;:m

;ow we know &! and 4!, using this we can find out shear force and 2ending

moment.

S#EAR ORCE

S0 BC =∫( * 1+ *

1

2 − *

3)'x−V A

S0 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− x

2+1225sin−1

( x35 ) +12.3325 [

x3

3 −35 x

2+1225 x ]−



2


S0 CD=S0 BC +∫ *, 'x

S0 CD=S0 BC +∫(1185.185 x . 39775.92)'x

S0 CD=S0 BC +(592.592 x2−39775.92 x)

S0 D1=S0 CD−35610.3

S0 10 =S0 D1−35610.3

S0 0A=S0 10 −(592.592 x2−39775.92 x) O $#)3$+.#

2y using the corresponding values of x in appropriate euations we get the plot

of shear force.

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

-1500000

-1000000

-500000

0

500000

1000000

1500000

2000000

,$ear force )ia!ra"

Location in #in! "

,$ear force %&(

8END)NG MOMENT

B/ BC =∬( *1+ *

2

2 + *

3−V A)'x

2+ / A

B/ BC =−143.156 x3+20041.96 x

2+457.98 x [ x √ 1225− x2+1225sin−1( x

35)]+305.32 [1225− x2 ]1.5−12.3325

B/ CD=∬( * 1+ *

2

2 + *

3+ *, −V A)'x

2+ / A

B/ CD=B/ BC +∬ *, 'x2



2


K 242C O #31.$)x)'#3//1.30x%

24(< K 24C( ')$0#+.)x

24<F K 24(< ')$0#+.)x

24F! K 24<F Q#31.$)x)'#3//1.30x%RO $#)3$+.#x

2y substituting the values of x for the above euations 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

0en)in! "o"ent )ia!ra"

Location in #in! "

0en)in! "o"ent %&"(



2


12 LOAD E,TIMATIO& O& F3,ELAGE



2


LOAD EST)MAT)ON ON +SELAGE

Fuselage 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, horiontal tail, vertical tail, landing gear

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

accommodating practically all the euipments, accessories and systems in

addition to carrying the payload. 2ecause of large amount of euipment inside

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

fuselage for access and inspection purposes. 5hese cutouts and discontinuities

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

efficient in design. !s a common member to which other components are

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

hollow beam. 5he reactions produced by the wing, tail or landing gear may be

considered as concentrated loads at the respective attachment points. 5he

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

of the fuselage structure and the various components inside the fuselage. 5hese

reaction forces are distributed all along the length of the fuselage, though need

not be uniformly .Dnlike 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. 5he main 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 form of shear only. 5he structural design of both wing andfuselage begin with shear force and bending moment diagrams for the

respective members

5o find out the loads and their distribution, consider the different cases. 5he

main components of the fuselage loading diagram areH

• Weight of the fuselage

• <ngine weight

• Weight of the horiontal and vertical stabiliers



2


• 5ail lift

• Weight of crew, payload and landing gear

• -ystems, euipment, accessories

-ymmetric flight condition, steady and level flightH 7(ownward forces negative8&alues for the different component weights are obtained from aerodynamic

design calculations.

Ta=le - Loas a&ting on !selage

Conition !ll 'a3loa

!selage alone anal3sis

S.N

oCom%onents

Distan&e $rom

re$eren&e line 5m6Mass 5kg6 Weight 5N6 Moment 5Nm6

# Crew .3# )++ %3)#)%#1.+#)

% ;ose "anding *ear 3.3/% $)++ $#33)$#/33.#%0

) Pay "oad 2ay # #1.30$ $+++ #$+13)+03.%$

Fixed <uipment %1.$#) #1++ #0011$//).)+#

$ Fuselage mass )).#1 1#/ 1)++#%)3$1/+.#

0 4ain "anding *ear # )).#1 #$3++ #$$313$%#%)$+.%)

1 4ain "anding *ear % .1$) #+#++ 33+/#)#1#.33)

/ Payload bay % .1$) $+++ #$+#31$0%##./$

3 Goriontal stabilier 00.30 3$++ 3)#3$0%+))1.%

#+ &ertical -tabilier 1#.$/0 /++ 1+//))1+/#.$0/

5otal %#%+#/ %+13/30.$/ 1%))#)/1.0

c.g. from nose K :;.BB? m

Ta=le 0 Shear $or&e an =ening moment ta=!lation



2


Distan&e5m6 Loa 5N6 Shear or&e 5N6 8ening Moment 5Nm6

+ + + +

.3# '%3) '%3) '#)%#1.+#)

3.3/% '$#33) '$3)0 '$+$111.##)

#1.30$ '#$+ '30)/0 '/)0%0.)0)

%1.$#) '#0011 '#0011 '//3$%0+.00

)).#1 '1)++# '#%)#+ '))%3#++.10

)).#1 '#$$313 '#)33+/) ')/$+))3#.+#

).110 %+13/30 0/+/#) ))/%1330.0)

.1$) '33+/# $/#1)% %3)3)/%.0

.1$) '#$+ #+%/% 30)10#%.13

00.30 '3)#3$ 1+// ))1+/#.$0/

1#.$/0 '1+// + +

Shear $or&e on the $!selage 5$ree2$ree =eam <ith one rea&tion at its &.g.6 at

$!ll3 loae &onition

0 10 20 30 40 50 60 70 80

-2000000

-1500000

-1000000

-500000

0

500000

1000000

,$ear force )ia!ra"

Distance fro" nose %"(

,$ear force %&(



2


8ening moment on the $!selage 5$ree2$ree =eam <ith one rea&tion at its

&.g.6 at $!ll3 loae &onition

0 10 20 30 40 50 60 70 80

-50000000

-40000000

-30000000

-20000000

-10000000

0

10000000

20000000

30000000

40000000

0en)in! "o"ent )ia!ra"

Distance fro" nose %"(

0en)in! "o"ent %&"(



2


. 8ALANC)NG AND MANE+1ER)NG

LOADS ON TA)L 'LANE, R+DDER AND

A)LERON LOADS



2


Mane!7ering loas

<ach horiontal 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 conditionsH

! sudden movement of the pitching control, at the speed &!, to the

maximum aft movement, and the maximum forward movement, as limited

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

! sudden aft movement of the pitching control at speeds above &!, followed

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

combinations of normal and angular acceleration. !t speeds up to &!, the

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

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

With the airplane in unaccelerated flight at ero yaw, it is assumed that the

rudder control is suddenly displaced to the maximum deflection, as limited

by the control stops or by limit pilot forces.

With 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 eual

to #.$ times the static sideslip angle may be assumed.

! yaw angle of #$ degrees with the rudder control maintained in the neutral

position 7except as limited by pilot strength8

5he airplane must be yawed to the largest attainable steady state sideslip

angle, with the rudder at maximum deflection caused by any one of the

followingH

→ Control surface stops

→ 4aximum available booster effort

→ 4aximum pilot rudder force

→ 5he rudder must be suddenly displaced from the maximum deflection to

the neutral position.



2


→ 5he yaw angles may be reduced if the yaw angle chosen for a particular

speed cannot be exceeded inH

→ -teady slip conditions

→ Dncoordinated rolls from steep banks or

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

5he ailerons must be designed for the loads to which they are subjectedH

In the neutral position during symmetrical flight conditionsL and

2y the following deflections 7except as limited by pilot effort8, during

unsymmetrical flight conditions

-udden maximum displacement of the aileron control at &!. -uitable

allowance may be made for control system deflections.

-ufficient deflection at &C, where &C is more than &!, to produce a rate of

roll not less than obtained

-ufficient deflection at &C to produce a rate of roll not less than one'third of

that obtained

5a6S3mmetri& mane!7ering &onitionsH

Where sudden displacement of a control is specified, the assumed rate of

control surface displacement may not be less than the rate that could be applied

by the pilot through the control system. In determining elevator angles and

chord wise load distribution in the maneuvering conditions, the effect of

corresponding pitching velocities must be taken into account. 5he in'trim andout'of'trim flight conditions must be considered.

5=6Mane!7ering =alan&e &onitions

!ssuming the airplane to be in euilibrium with ero pitching acceleration, the

maneuvering conditions on the maneuvering envelope must be investigated.

5&6'it&h mane!7er &onitionsH

5he movement of the pitch control surfaces may be adjusted to take into

account limitations imposed by the maximum pilot effort, control system stops



2


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

system 7for example, stalling torue or maximum rate obtainable by a power

control system.

Maim!m %it&h &ontrol is%la&ement at &!H

5he airplane is assumed to be flying in steady level flight and the cockpit pitch

control is suddenly moved to obtain extreme nose up pitching acceleration. In

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

!irplane loads that occur subseuent to the time when normal acceleration at

the c.g. exceeds the positive limit maneuvering load or the resulting tail plane

normal load reaches its maximum, whichever occurs first, need not be

considered.

S%e&i$ie &ontrol is%la&ement

! checked maneuver, based on a rational pitching control motion vs. time

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

exceeded. Dnless lesser values cannot be exceeded, the airplane response must

result in pitching accelerations not less than the followingH

! positive pitching acceleration 7nose up8 is assumed to be reached

concurrently with the airplane load factor of #.+. 5he positive

acceleration must be eual to at least )3n7n'#8:v, 7rad:sec 8

Where Sn is the positive load factor at the speed under considerationL and & is

the airplane euivalent speed in knots.

! negative pitching acceleration 7nose down8 is assumed to be reached oncurrently with the positive maneuvering load factor. 5his negative

pitching acceleration must be eual to at least '%0n7n'#8:v, 7rad:sec 8

Where Sn is the positive load factor at the speed under considerationL and & is

the airplane euivalent speed in knots.

8alan&ing loas

! horiontal surface balancing load is a load necessary to maintaineuilibrium in any specified flight condition with no pitching acceleration.



2


Goriontal 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 reuired to balance the rudder because it will not deflect due to

gravity.

!ileron will defect in vice versa direction so it is doesnt reuire balancing

load.



2


. DETA)LED STR+CT+RAL LA"O+TS



2


+NCT)ON O T#E STR+CT+RE

5he primary functions of an aircrafts structure can be basically broken down

into the followingH

5o transmit and resist applied loads.

5o provide and maintain aerodynamic shape.

5o protect its crew, passenger, payload, systems, etc.

For the vast majority of aircraft, this leads to use of a semi'monocoue design

7i.e. a thin, stressed outer shell with additional stiffening members8 for the wing,

fuselage T empennage. 5hese notes will discuss the structural layout

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

W)NG STR+CT+RAL LA"O+T

S%e&i$i& Roles o$ Wing 5Main <ing6 Str!&t!re

5he specified structural roles of the wing 7or main plane8 areH

5o transmitH

→ wing lift to the root via the main span wise beam

→ Inertia loads from the power plants, undercarriage, etc., to the main beam.

→ !erodynamic loads generated on the aerofoil, control surfaces T flaps to

the main beam.

5o react againstH

→ "anding loads at attachment points

→ "oads from pylons:stores

→ Wing drag and thrust loads

5o provideH

→

Fuel tank age space→ 5orsional rigidity to satisfy stiffness and aero'elastic reuirements.

5o fulfill these specific roles, a wing layout will conventionally compromiseH

→ -pan wise members 7known as spars or booms8

→ Chord wise members7ribs8

→ ! covering skin



2


→ -tringers

8asi& !n&tions o$ Wing Str!&t!ral Mem=ers

5he structural functions of each of these types of members may be

considered independently asH

S'ARS

Form the main span wise beam

5ransmit bending and torsional loads

Produce a closed'cell structure to provide resistance to torsion, shear and

tension loads.

)n %arti&!lar

• Webs resist shear and torsional loads and help to stabilie the skin.

• Flanges ' resist the compressive loads caused by wing bending.

SK)N

5o form impermeable aerodynamics surface

5ransmit aerodynamic forces to ribs T stringers

6esist shear torsion loads 7with spar webs8.

6eact axial bending loads 7with stringers8.



2


STR)NGERS

Increase skin panel buckling strength by dividing into smaller length

sections.

6eact axial bending loads

R)8S

4aintain the aerodynamic shape

!ct along with the skin to resist the distributed aerodynamic pressure

loads

(istribute concentrated loads into the structure T redistribute stress

around any discontinuities Increase the column buckling strength of the stringers through end

restraint

Increase the skin panel buckling strength.

S'ARS

5hese usually comprise thin aluminium alloy webs and flanges, sometimes with

separate vertical stiffeners riveted on to the webs.

T3%es o$ s%arsH

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

located as far forward as possible to maximie the wing box sie, though this is

subject to there beingH

!deuate wing depth for reacting vertical shear loads.

!deuate nose space for "< devices, de'icing euipment, etc.

5his generally results in the front spar being located at #%M to #/M of the chord

length. For a single spar ('nose layout, the spar will usually located at the

maximum thickness position of the aerofoil section 7typically between )+M T

+M along the chord length8. For the standard box beam layout, the rear spar

will be located as for aft as possible, once again to maximie the wing box sie,

but positioning will be limited by various space reuirements for flaps, control

surfaces, spoilers etc. 5his usually results in a location somewhere between



2


about $$Mand 1+M of the chord length. If any intermediate spars are used, they

would tend to be spaced uniformly unless there are specific pick'up point

reuirements.

R)8S

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

sheet metal by the use of presses Tdies. Flanges are incorporated around the

edges so that they can be riveted to the skin and the spar webs. Cut'out are

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

"ightening holes are usually cut into the rib bodies to reduce the rib weight and

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

6ib bulkheads do not include any lightening holes and are used at fuel tank

ends, wing crank locations and attachment support areas. 5he rib should be

ideally spaced to ensure adeuate overall buckling support to spar flanges. In

reality, however, their positioning is also influenced byH

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

hinges, power plants, stores, undercarriage attachment etc.

Positioning of fuel tank ends, reuiring closing ribs.

! structural need to avoid local shear or compression bucklingL 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 axis 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 complex connections.

SK)N

5he skin tends to be riveted to the rib flanges and stringers, using countersunk

rivets to reduce drag. It is usually pre'formed at the leading edges, where the

curvature is large due to aerodynamic considerations.

+SELAGE STR+CT+RE



2


5he fundamental purpose of the fuselage structure is to provide an envelope to

support the payload, crew, euipment, systems and 7possibly8 the power'plant.

Furthermore, it must react against the in'flight manoeuvre, pressurisation and

gust loadsL 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 surfaces throughout the rest of the structure

Fuselage 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, horiontal tail, vertical tail, landing gear

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

accommodating practically all euipment, accessories and systems in addition

to carrying the payload. 2ecause of large amount of euipment inside the

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

for access and inspection purposes. 5hese cutouts and discontinuities result in

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

design.

!s a common member to which other components are attached, thereby

transmitting the loads, fuselage can be considered as a long hollow beam. 5he

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

concentrated loads at the respective attachment points. 5he balancing reactions

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

structure and the various components inside the fuselage. 5hese reaction forcesare distributed all along the length of the fuselage, though need not be

uniformly. Dnlike 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. 5he main 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 form of shear only. 5he structural design of both wing and fuselage begin

with shear force and bending moment diagrams for the respective members. 5he



2


maximum bending stress produced in each of them is checked to be less than

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

!selage La3o!t Con&e%ts

5here are two main categories of layout concept in common useL

4ass boom and longeron layout

-emi'monocoue layout

Mass 8oom F Longeron la3o!t

5his is fundamentally very similar to the mass'boom wing'box concept

discussed in previous section. It is used when the overall structural loading is

relatively low or when there are extensive cut'outs in the shell. 5he conceptcomprises four or more continuous heavy booms 7longeron8, reacting against

any direct stresses caused by applied vertical and lateral bending loads. Frames

or solid section

Semi2Mono&o!e la3o!t

5he semi'monocoue is the most often used construction for modern, high'

performance aircraft. Semi2mono&o!e literally means half a single shell. Gere,internal braces as well as the skin itself carry the stress. 5he vertical structural

members are referred to as bulkheads, frames, and formers. 5he heavier

vertical members are located at intervals to allow for concentrated loads. 5hese

members are also found at points where fittings are used to attach other units,

such as the wings and stabiliers.

Primary bending loads are taken by the longerons, which usually extend across

several points of support. 5he longerons are supplemented by other longitudinal

members known as stringers. -tringers are more numerous and lightweight than

longerons. 5he stringers are smaller and lighter than longerons and serve as fill'

ins. 5hey have some rigidity but are chiefly used for giving shape and for

attachment of skin.

5he strong, heavy longerons hold the bulkheads and formers. 5he bulkheads

and formers hold the stringers. !ll of these join together to form a rigid fuselage



2


framework. -tringers and longerons prevent tension and compression stresses

from bending the fuselage. 5he skin is attached to the longerons, bulkheads, and

other structural members and carries part of the load. 5he fuselage skin

thickness varies with the load carried and the stresses sustained at particular

location.



2


-H. DES)GN O SOME COM'ONENTS O

W)NG AND +SELAGE



2


DES)GN O W)NG COM'ONENT −¿ S'AR

Wing is the major lift producing surface. 5herefore, the analysis has to be very

accurate. 5he structural analysis of the wing by defining the primary load

carrying 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. 5here are two spars, one located at #$'

%+M of the chord known as the front spar, the other located at 0+'1+M of the

chord known as the rear spar. -ome of the functions of the spar includeH

5hey form the boundary to the fuel tank located in the wing.• 5he spar flange takes up the bending loads whereas the web carries the

shear loads.

• 5he rear spar provides a means of attaching the control surfaces on the

wing.

Considering these functions, the locations of the front and rear spar are fixed at

+.#1c and +.0$c respectively. 5he spar design for the wing root has been taken

because the maximum bending moment and shear force are at the root. It is

assumed that the flanges take up all the bending and the web takes all the shear

effect. 5he maximum bending moment for high angle of attack condition is

261419642.5 ;m. 5he ratio in which the spars take up the bending moment is

given as

/ ,

/ r=

2, 2

2r

2

Where

h# ' height of front spar

h% ' height of rear spar

/ ,

/ r=1.6186

2

1.47952

/ , / r

=1.197



2


/ , + / r=261419642.5

5herefore,

4f K ##/3/3)0/.$ ;m

4r K 33+0)%%./$ ;m

5he yield tensile stress Ny for !l !lloy 7!l 1+1$8 is ;>>.H>:?0 M'a. 5he area

of the flanges is determined using the relation

3 *= /

A4

where 4 is bending moment taken up by each spar,

! is the flange area of each spar,

is the centroid distance of the area K h:%.

Dsing the available values,

!rea of front spar,

!f K +.)%)+33 m%

!rea of rear spar,

!r K +.%3$)+ m%

!ssumptionsH

5 sections are chosen for top and bottom flanges of front and rear spars. 2oth

the flanges are connected by a vertical stiffener through spot welding and

( ,

( -=1

From the buckling euation,

0 cr=0.388 1 (( -&- )

2

the thickness to width ratio of web( -

&-

is found to be ).3$3#. !lso from

U!nalysis and design of flight vehicle structures by 26DG;V, the flange to web

width ratio of the 5 section&,

&-=1.8



2


2y euating all the three values of the ratio in area of the section euation, the

dimensions of the spar can be found.

Dimensions $or $ront s%ar

bflangeK +.%0)) m

tflange K twebK+.%)0/$ m

bweb K +.3)11# m

Dimensions $or rear s%ar

bflangeK +.+1$3 m

tflange K twebK +.%%0 m

bweb K +./303 m

DES)GN O +SELAGE COM'ONENT −¿ STR)NGER

5he circumference of the fuselage is ).#+% m. 5o find the area of one stringer,

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

stringers in the fuselage is #0. 5he stringers are eually spaced around the

circumference of the fuselage.

Stringer S%a&ing



2


5he stringers are symmetrically spaced on the fuselage with the spacing

calculate as shown below,

Circumference of the fuselage K $D=$ ∗2∗6.86=43.102m

5otal number of stringers K #0

5herefore the stringers are spaced at the interval of K 43.102

16 =2.6939m

Stringer area &al&!lation

5he stress induced in the each stringer is calculated with the area keeping

constant in the stress term. 5hen the maximum stress 7i.e. one which has larger

numerator8 is euated with the yield strength of the material. From this area of

one stringer is calculated.

5he direct stress in each stringer produced by bending moments / x and

/ * is given by the euationH

3 = / 5

+ 55

4+ / 6

+ 66

x N /m2

Where

/ 5 =33827996.63 N

/ 6 =(12 ρ V 2

S( a ( 7 )× x

ρ is density K#.%%$ kg:m)

& is cruise velocity K %$+ m:s

-t is the tail area K 1#.0$1 m%

at is the slope of the lift curve K +.+0/# :deg

7 is the angle of yaw for asymmetric flight

7 =0.7nmax+457.2

V D

7 =3.563'eg

x is the distance between the aircraft c.g position and horiontal tail c.g

positionK :;.BB? m

5hen,



2


/ 6 =23146604.65 Nm

+ 55 = + 66 = As(inger D2

Where A s(inger is the stringer area, ( is the diameter of the fuselage K #).1% m.

/ x and

/ * reach their maximum only from the stringers # to . 5hus the

stresses are high only on these stringers. Calculating stress for stringer # to .

A K +, 9 K 0./0

3 2=

/ 5

+ 55

6 + / 6

+ 66

5

5hen,

3 1=

1232798.71

A s(ing er

N

m2

A K %.0%3, 9 K 0.)

3 2=

/ 5

+ 55

6 + / 6

+ 66

5

5hen,

3 2=

1462623.57

A s(inger

N

m2

A K ./$1/, 9 K ./$1/

3 3=

/ 5

+ 55

6 + / 6

+ 66

5

5hen,

3 3=1470322.836

As(inger

N

mm2

A K 0.))1/, 9 K %.0%$

3 4=

/ 5

+ 55

6 + / 6

+ 66

5



2


5hen, 3 4=

1251057.39

As(inger

N

mm2

5he allowable stress in the stringer is $$.+$)30% 4Pa for !l !lloy 7!l 1+1$8.

5he maximum direct stress in the stringer % is

3 2=

1462623.57

A s(inger

N

m2

5herefore the reuired stringer area of cross section is then given by

1462623.57

A s(inger

=455.053962×106

A s(inger=3.214×10−3

m2

5hus one stringer area is 3.214×10−3

m2 . 5he stringer chosen is 9 section. 5he

dimensions of the stringers are obtained from the !;!"E-I- !;( (<-I*;

=F 5G< F"I*G5 &<GIC"< -56DC5D6<- by 26DG;.

The imensions are,

twK tf K $.311 ×10−3

m

bflange K+.#)$ m

bweb K +.%03 m





2


DESCR)'T)ON

!ircraft structures are basically unidirectional. 5his means that one dimension,

the length, is much larger than the others ' width or height. For example, the

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

ribs have a much larger chord length than height and:or widthL a whole wing has

a span that is larger than its chords or thicknessL 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 thickness, etc.... For this simple reason, a designer chooses to

use unidirectional material when designing for an efficient strength to weight

structure.

Dnidirectional materials are basically composed of thin, relatively flexible, long

fibers which are very strong in tension 7like a thread, a rope, a stranded steel

wire cable, etc.8. !n aircraft structure is also very close to a symmetrical

structure. 5hose mean the up and down loads are almost eual to each other.

5he tail loads may be down or up depending on the pilot raising or dipping the

nose of the aircraft by pulling or pushing the pitch controlL the rudder may be

deflected to the right as well as to the left 7side loads on the fuselage8. 5he gusts

hitting the wing may be positive or negative, giving the up or down loads which

the occupant experiences by being pushed down in the seat or hanging in the

belt.

2ecause of these factors, the designer has to use a structural material that can

withstand both tension and compression. Dnidirectional fibers may be excellentin tension, but due to their small cross section, they have very little inertia 7we

will explain inertia another time8 and cannot take much compression. 5hey will

escape the load by bucking away. !s in the illustration, you cannot load a string,

or wire, or chain in compression.

In order to make thin fibers strong in compression, they are glued together

with some kind of an embedding. In this way we can take advantage of their

tension strength and are no longer penalied by their individual compression



2


weakness because, as a whole, they become compression resistant as they help

each other to not buckle away. 5he embedding is usually a lighter, softer resin

holding the fibers together and enabling them to take the reuired compression

loads. 5his is a very good structural material.

WOOD

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

material. 5hey have to be tall and straight and their wood must be strong and

light. 5he dark bands 7late wood8 contain many fibers, whereas the light bands

7early wood8 contain much more resin. 5hus the wider the dark bands, thestronger and heavier the wood. If the dark bands are very narrow and the light

bands uite wide, the wood is light but not very strong. 5o 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 7length and width are large

with respect to thickness8. Plywood is often used for such structures. -everal

thin boards 7foils8 are glued together so that the fibers of the various layers cross

over at different angles 7usually 3+ degrees today years back you could get them

at )+ and $ degrees as well8. Plywood makes excellent shear webs if the

designer knows how to use plywood efficiently. 7We will learn the basis of

stress analysis sometime later.8

5oday good aircraft wood is very hard to come by. Instead of using one good

board for our spars, we have to use laminations because large pieces of wood

are practically unavailable, and we no longer can trust the wood uality. From

an availability point of view, we simply need a substitute for what nature has

supplied us with until now.

AL+M)N)+M ALLO"S

-o, since wood may not be as available as it was before, we look at another

material which is strong, light and easily available at a reasonable price 7thereXs



2


no point in discussing 5itanium ' itXs simply too expensive8. !luminium alloys

are certainly one answer. We will discuss the properties of those alloys which

are used in light plane construction in more detail later. For the time being we

will look at aluminium as a construction material.

E*TR+DED AL+M)N)+M ALLO"S

(ue to the manufacturing process for aluminium we get a unidirectional

material uite a bit stronger in the lengthwise direction than across. !nd even

better, it is not only strong in tension but also in compression. Comparing

extrusions to wood, the tension and compression characteristics are practicallythe same for aluminium alloys so that the linear stress analysis applies. Wood,

on the other hand, has a tensile strength about twice as great as its compression

strengthL accordingly, special stress analysis methods must be used and a good

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

avoidedY

!luminium alloys, in thin sheets 7.+#0 to .#%$ of an inch8 provide an excellent

two dimensional material used extensively as shear webs ' with or without

stiffeners ' and also as tension:compression members when suitably formed

7bent8.It is worthwhile to remember that aluminium is an artificial metal. 5here

is no aluminium ore in nature. !luminium is manufactured by applying electric

power to bauxite 7aluminium oxide8 to obtain the metal, which is then mixed

with various strength'giving additives. 7In a later article, we will see which

additives are used, and why and how we can increase aluminiumXs strength by

cold work hardening or by tempering.8 !ll the commonly used aluminium

alloys are available from the shelf of dealers. When reuested with the

purchase, you can obtain a mill test report that guarantees the chemical and

physical properties as tested to accepted specifications.



2


!s 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

5he next material to be considered for aircraft structure will thus be steel, which

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

!part from mild steel which is used for brackets needing little strength, we are

mainly using a chrome'molybdenum alloy called !I-I #)=; or #+. 5he

common raw materials available are tubes and sheet metal. -teel, due to its highdensity, is not used as shear webs like aluminium sheets or plywood. Where we

would need, say.#++ plywood, a .+)% inch aluminium sheet would be reuired,

but only a .+#+ steel sheet would be reuired, which is just too thin to handle

with any hope of a nice finish. 5hat 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 7light weight8 to give it the reuired aerodynamic

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

techniuesH steel work and fabric covering. .

COM'OS)TE MATER)ALS

5he designer of composite aircraft simply uses fibers in the desired direction

exactly where and in the amount reuired. 5he fibers are embedded in resin to

hold them in place and provide the reuired support against buckling. Instead of

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

designer uses cloth where the fibers are laid in two directions .7the woven thread

and weft8 also embedded in resin. 5his has the advantage of freedom of shape in

double curvature as reuired by optimum aerodynamic shapes and for very

appealing look 7importance of aesthetics8.



2


5odayXs fibers 7glass, nylon, Bevlar, carbon, whiskers or single crystal fibers of

various chemical compositions8 are very strong, thus the structure becomes very

light. 5he drawback is very little stiffness. 5he structure needs stiffening which

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

sandwich structureH two layers of thin uni' or bi'directional fibers are held apart

by a lightweight core 7foam or honeycomb8. 5his allows the designer to

achieve the reuired inertia or stiffness.

From an engineering standpoint, this method is very attractive and supported by

many authorities because it allows new developments which are reuired in

case of war. 2ut this method also has its drawbacks for homebuildingH ! mold is

needed, and very strict uality control is a must for the right amount of fibers

and resin and for good adhesion between both to prevent too dry or wet a

structure. !lso the curing of the resin is uite sensitive to temperature, humidity

and pressure. Finally, the resins are active chemicals which will not only

produce the well'known allergies but also the chemicals that attack our body

7especially the eyes and lungs8 and they have the unfortunate property of being

cumulatively damaging and the result 7in particular deterioration of the eye8

shows up only years after initial contact.

!nother disadvantage of the resins is their limited shelf life, i.e., if the resin is

not used within the specified time lapse after manufacturing, the results may be

unsatisfactory and unsafe.

#EA1" A)RCRAT RAW MATER)ALS

-. Magnesi!m !n expensive material. Castings are the only readily available

forms. -pecial precaution must be taken when machining magnesium because

this metal burns when hot.

0. Titani!m ! very expensive material. &ery tough material and difficult to

machine.

:. Car=on i=ers -till very expensive materials.



2


;. Ke7lar i=ers &ery expensive and also critical to work with because it is

hard to soak in the resin.

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

structure. 5he selection of the best material depends upon the application.

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

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

repairability, cost and availability. 5he 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.

For most aerospace materials, creep is a problem only at the elevated

temperature. Gowever some titanium plastics and composites will exhibit creep

at room temperatures.

5aking all the above factors into considerations, the following aluminium alloys

which have excellent strength to weight ratio and are abundant in nature are

considered.

S.No Al!mini!m Allo3"iel strength

M'a

+ltimate strength

M'a

# !l %+%' 5)$ %/+ 1+

% !l %+%' 5) %10 %1

) !l 1+1$' 50 10 $)/

!l1+1$' 50$# 0% $)/

$ !l 0+0#'+ $$ ##%

0 !l0+0'5 ##+ %+1

1 !l0+0#'50 %# %3+

/ !!0+/%50 %#+ )+



2


DES)GN RE'ORT



2


Design Re%ort



2

ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT'arameters 1al!es

-pan 1+ m

Planform area $3 m%

!spect ratio /.3%

<mpty weight #0++++ kg

4aximum takeoff weight )$3))# kg

=swald efficiency factor +.1/$%

Chord at root #%.$ m

Chord at tip ).#)$ m

5aper ratio +.%$

-weepback angle )./)

Wing loading 0$ kg:m%

Power delivered by motor $ hp

5hrust'to'weight ratio +.#$/

6ate of climb $.%% m:s

<ndurance ) hours

6ange $++ km

-tall speed 03. m:s

"anding distance #$+ m

5akeoff distance %%11.0 m

4aximum Ove "oad factor )

4aximum ve "oad factor #.%

(esign dive speed )#%.$ m:s

"ift coefficient7flaps down8 #.#)/

4inimum radius of turn 0//.03/ m

4aximum bending moment %0##30%.$ ;m

Front spar bending moment ##/3/3)0/.$ ;m

6ear spar bending moment 33+0)%%./$ ;m

2ending moment in fuselage ))/%1330.0) ;m



2


Three 7ie< igram




CONCL+S)ON

5he structural design of the Geavy'lift military cargo aircraft which is a

continuation of the aerodynamic design part carried out last semester is

completed satisfactorily. 5he aeroplane has gone through many design

modifications since its early conceptual designs expected, among these was a

growth in weight.

5o ensure continued growth in payload and the reduced cost of cargo

operations, improvements in methods, euipment and terminal facilities will be

reuired in order to reduce cargo handling costs and aircraft ground time and to

provide improved service for the shippers.

We have enough hard work for this design project. ! design never gets

completed in a flutter sense but it is one step further towards ideal system. 2ut

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

implications when applied to an aircraft design.

8)8L)OGRA'#"

1. Raymer, D.P, Aircraft Design - a Conceptual Approach , AIAA educational

series second edition 1992.

2. T.H.G.Megson , Aircraft Structures for engineering students, 4t !dition

!lse"ier #td $%A 2&&'.

(. !.).*run , Analysis and design of flight vehicle structures,1 st !dition, tri+

state oset com-any,$%A,19'(.

4. Miceal un+/ung 0iu , Airframe structural design, 2nd !dition, Hong ong

onmilit Press #td, Hong ong, 2&&1.

Documents

180745548 Aircraft Design Project 2