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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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
)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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
!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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
• (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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
0. 12n Diagram
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
)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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
¬¿¿¿¿
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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
¬¿¿¿¿¬¿¿¿
¿ 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
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/./ '#.%
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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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
;. CR)T)CAL LOAD)NG 'ERORMANCE
AND )NAL 12n D)AGRAM
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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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2
ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
Pull'up %./%
Pull'down +./%
Climb #.0//
inal 12n Diagram
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2
ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
>. STR+CT+RAL DES)GN ST+D" 4
T#EOR" A''ROAC#
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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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2
ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
?. LOAD EST)MAT)ON ON W)NGS
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
-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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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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)
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
*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 "
Lift per Len!t$ %&'"(
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
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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 "
Lift per Len!t$ %&'"(
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 "
Lift per Len!t$ %&'"(
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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$ %&'"(
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
!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.
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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 ]−
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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 %&"(
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2
ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
12 LOAD E,TIMATIO& O& F3,ELAGE
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
• 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 $or&e an =ening moment ta=!lation
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2
ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
Distan&e5m6 Loa 5N6 Shear or&e 5N6 8ening Moment 5Nm6
+ + + +
.3# '%3) '%3) '#)%#1.+#)
3.3/% '$#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 %&(
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2
ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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 %&"(
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2
ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
. 8ALANC)NG AND MANE+1ER)NG
LOADS ON TA)L 'LANE, R+DDER AND
A)LERON LOADS
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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.
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
→ 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
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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.
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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.
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
. DETA)LED STR+CT+RAL LA"O+TS
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
+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
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→ -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.
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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
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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
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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
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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
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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.
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
-H. DES)GN O SOME COM'ONENTS O
W)NG AND +SELAGE
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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
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/ , + / 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
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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
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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,
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/ 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
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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
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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
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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
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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.
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!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.
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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.
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
;. 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 %#+ )+
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
DES)GN RE'ORT
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
Design Re%ort
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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
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
Three 7ie< igram
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ADP-II HEAVY-LIFT MILITARY CARGO AIRCRAFT
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.