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JAWAHARLAL
INSTITUTE OF TECHNOLOGY(Approved by AICTE & Affiliated to Anna University)
COIMBATORE641 105
NAME : ___________________________________________
REG.NO : ___________________________________________
SUBJECT : ___________________________________________
COURSE : ___________________________________________
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JAWAHARLAL INSTITUTE OF TECHNOLOGY
COIMBATORE641 105
DEPARTMENT OF AERONAUTICAL ENGINEERING
Certified that this is the bonafide record work done by
. in the AIRCRAFT DESIGN LABIof this institution as prescribed by the Anna University, Coimbatore for the
........ Semester during the year 2011-2012
Staff In charge: Head of the Department
University Register No.:
Submitted for the Practical Examination of the Anna University conducted on
INTERNAL EXAMINER EXTERNAL EXAMINER
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CONTENT
S.NO NAME OF THE EXPERINMENT PAGE NO
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ABSTRACT
In the Aircraft Design Project we have decided to design a super
jumbo aircraft with a passenger seating capacity of 800 nos. The
aircraft parameters like cruise velocity, cruise altitude, wing loading
etc. And weight estimation, airfoil selection, wing selection, landing
gear selection has been made with extreme care after a several
comparison with a few same types of aircrafts. . The adequate details
have been collected to make our calculation easier and to make
design more precision. The details have been collected from various
sources which are given in the bibliography.
Even though there are huge jumbo aircrafts exist there such as A380,
B747, A340, MD-12LR which having a seat capacity around a 600 in
no. only A380 and B747 are the double deck aircrafts ever built for
civil aviation.
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SYMBOLS AND ABBBREVIATIONS
A : area
A1 : intake highlights area
Ath : throat area
APR : augmented power rating
AR : aspect ratio
AW : wetted area
a : speed of sound; acceleration : Average acceleration at 0.7 V2ac : aerodynamic centre
B :breadth, width
b :span
CR : CB root chord
CD : drag coefficient
CDi : induced drag coefficient
CDp : parasitic drag coefficient
CDpmin: minimum parasitic drag coefficient
CDw : wave drag coefficient
Cv : specific heat at constant volume
CF : overall skin friction coefficient; force coefficient
Cf : local skin friction coefficient; coefficient of friction
CL : lift coefficient
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Cl : sectional lift coefficient; rolling moment coefficient
CLi : integrated design lift coefficient
CL : lift curve slope
CL : sideslip curve slope
Cm : pitching-moment coefficient
Cn : yawing-moment coefficient
Cp : pressure coefficient; power coefficient; specific heat at constant
pressure
CT : thrust coefficient
CHT : horizontal tail volume coefficient
D : Drag
E : Endurance
e : Oswald efficiency
g : Acceleration due to gravity
G : Factor due to ground effect
JA, JT : Symbols
h : Height from ground
hOB : Obstacle height
k1 : Proportionality constant
kuc : Factor depends on flap deflection
KA , KT : Symbols
L : Lift
loiterD
L
: Lift-to-drag ratio at loiter
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cruiseD
L
: Lift-to-drag ratio at cruise
M : Mach number of aircraft
mff : Mission segment fuel fraction
N : Time between initiation of rotation and actual
R : Range
Re : Reynolds Number
R/C : Rate of climb
S : Wing Area
Sa : Approach distance
Sab : Distance require to clear an obstacle after becoming airborne
Sf : Flare distance
Sg : Ground Roll
Sref. : Reference surface area
Swet.. : Wetted surface area
T : Thrust
P : Power
Pcruise : Thrust at cruise
Ptake-off : Thrust at take-off
loiterW
P
: Thrust-to-weight ratio at loiter
cruiseW
P
: Thrust-to-weight ratio at cruise
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takeoffW
P
: Thrust-to-weight ratio at take-off
Vcruise : Velocity at cruise
Vstall : Velocity at stall
VLO : Lift off Speed
VTD : Touch down speed
Wcrew : Crew weight
Wempty : Empty weight of aircraft
Wfuel : Weight of fuel
Wpayload : Payload of aircraft
W0 : Overall weight of aircraft
S
W : Wing loading
: Density of air
: Dynamic viscosity
r : Co-efficient of rolling friction
: Tapered ratio
OB : Angle between flight path and take-off
: Turning angle
: Gliding angle
R/C : Rate of climb
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INTRODUCTION
Need for airplane design
An airplane is designed to meet the functional, operational and safety
requirements set by or acceptable to the ultimate user. The actual process of design is a
complex and long drawn out engineering task involving:
Selection of airplane type and shape
Determination of geometric parameters
Selection of power plant
Structural design and analysis of various components and Determination of airplane flight and operational characteristics.
Over the year of this century, aircraft have evolved in many directions and
the design of any modern plane is a joint project for a large body of competent engineers
and technicians, headed by a chief designer. Different groups in the project specialize in the
design of different components of the airplane, such as the wing, fuselage etc.
A new experimental plane has to meet higher performance requirements
than similar planes already in service. Hence design laboratories involved in experimental
and research work are indispensable adjuncts to a design office. These laboratories as well
as allied specialized design offices and research institutions are concerned in helping the
designer to obtain the best possible solutions for all problems pertaining to airplane design
and construction and in the development of suitable components and equipment.
Airplane design procedure is basically a method of trial and error for the design
of component units and their harmonization into a complete aircraft system. Thus each trial
aims at a closer approach to the final goal and is based on a more profound study of the
various problems involved. The three phases of aircraft design are
Conceptual design
Preliminary design
Detailed design
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Phase of aircraft design
FIG: 1
FIG: 2
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Conceptual design
Aircraft design can be broken into three major phases, as depicted in figure. Conceptual
design is the primary focus of this book. It is in conceptual design that the basic questions ofconfiguration arrangement, size and weight, and performance are answered.
The first question is can an affordable aircraft be built that meets the requirements? if not,
the customer may wish to relax the requirements.
Conceptual design is a very fluid process. New ideas and problems emerge as a design is investigated
in increasing detail. Each time the latest design is analyzed and sized, it must be redrawn to reflect
the new gross weight, fuel weight, wing size, and other changes. Early wind tunnel test often revel
problems requiring some changes to the configuration.
Preliminary design
Preliminary design can be said to begin when the major changes are over. The big
questions such as whether to use a canard or an aft tail have been resolved. The
configuration arrangement can be expected to remain about as shown on current drawing,
although minor revisions may occur. At some point late in preliminary design, even minor
changes are stopped when a decision is made to freeze the configuration.
During preliminary design the specialists in area such as structure landing gear and
control systems will design and analyze their portion of the aircraft. Testing is initiated in
areas such as aerodynamics, propulsion, structures, and control. A mockup may be
constructed at this point.
A key activity during preliminary design is lofting. Lifting is the mathematical
modeling of the outside skin of the aircraft with sufficient accuracy to insure proper fit
between its different parts, even if they are designed by different designers and possibly
fabricated in different location. Lofting originated in shipyards and was originally done with
long flexible rulers called splines. This work was done in a loft over the shipyard; hence
the name.
The ultimate objective during preliminary design is to ready the company for the detail
design stage, also called full-scale development. Thus, the end of preliminary design usually
involves a full scale development proposal. In todays environment, this can result in a
situation jokingly referred to as you-bet-your-company. The possible loss on an overrun
contrast o from lack of sales can exceed the net worth of the company! Preliminary design
must establish confidence that the airplane can be built in time and at the estimated cost.
Detailed design
Assuming a favorable decision for entering full scale development, the detail design
phase begins in which the actual pieces to be fabricated are designed. For example, during
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conceptual and preliminary design the wing box will be designed and analyzed as a whole.
During detail design, that whole will be broken down in to individual ribs, spars and skins,
each of which must be separately designed and analyzed.
Another important part of detailed is called production design. Specialist determine
how the airplane will be fabricated, starting with the smallest and simplest subassemblies
and building up to the final assembly process. Production designers frequently wish to
modify the design for ease of manufacture; that can have a major impact on performance or
weight. Compromises are inevitable, but the design must still meet the original
requirements.
It is interesting to note that in the Soviet Union, the production design is done by a
completely different design bureau than the conceptual and preliminary design, resulting in
superior reducibility at some expense in performance and weight.
During detail design, the testing effort intensifies. Actual structure of the aircraft is
fabricated and tested. Control laws for the flight control system arte tested on an iron-
bird simulator, a detailed working model of the actuator and flight control surfaces. Flight
simulator are developed and flown by both company and customer test pilot.
Detail design ends with fabrication of the aircraft. Frequently the fabrication Begins on
part of the aircraft before the entire detail-design effort is completed. Hopefully, changes to
already- fabricated pieces can be avoided. The further along a design progresses, the more
people are involved. In fact, most of the engineers who go to work for a major aerospace
company will work in preliminary on detail design.
Classification of airplanes design
Functional classification:
The airplane today is used for a multitude of activities in civil and
military fields. Civil applications include cargo transport, passenger travel, mail distribution,
and specialized uses like agricultural, ambulance and executive flying. The main types of
military airplane at the present time are fighters and bombers. Each of these types may be
further divided into various groups, such as strategic fighters, interceptors, escort fighters,
tactical bombers and strategic bombers. There are also special aircraft, such as ground
attack planes and photo-re-connaisance planes. Sometimes more than one function may be
combines so that we have multi-purpose airplanes like fighter-bombers. In addition to
these, we have airplanes for training and sport.
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Classification by power plants:
Types of engines used for power plant:
Piston engines (krishak, Dakota, super constellation)
Turbo-prop engines ( viscount,friendship,An-102)
Turbo-fan engines (HJT16, Boeing series, MIG-21)
Ramjet engines
Rockets (liquid and solid propellants) (X-15A)
Location of power plant:
Engine ( with propeller) located in fuselage nose (single engine)
(HT-2,Yak-9,A-109)
Pusher engine located in the rear fuselage (Bede XBD-2) Jet engines submerged in the wing
1. At the root(DH Comet, Tu-104,Tu-16)
2. Along the span (Canberra, U-2, YF-12A)
Jet engines in nacelles suspended under the wing (pod
mountings) (Boeing 707,DC-8,Convair 880)
Jet engines located on the rear fuselage (Trident, VC10 ,i1-62)
Jet engines located within the rear fuselage (Hf 24,
lighting,MIG-19)
Classification by configuration:
Airplanes are also classified in accordance with their shape and structural
layout, which in turn contribute to their aerodynamic, tactical and operational
characteristics. Classification by configuration is made according to:
Shape and position of the wing
Type of fuselage Location of horizontal tail surfaces
Shape and position of the wing:
Braved biplane(D.H. Tiger moth)
Braced sesquiplane (An-2)
Semi-cantilever parasol monoplane (baby ace)
Cantilever low wing monoplane (DC-3,HJT-16,I1-18,DH Comet)
Cantilever mid wing monoplane (Hunter, Canberra)
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Cantilever high wing monoplane (An-22,Brequet 941 Fokker
Friendship)
Straight wing monoplane (F-104 A)
Swept wing monoplane (HF-24, MIG-21, Lighting)
Delta monoplane with small aspect ratio (Avro-707, B-58 Hustler,Avro Vulcan)
Type of fuselage
Conventional single fuselage design ( HT-2,Boeing 707
Twin- fuselage design
Pod and boom construction (Packet, Vampire)
Types of landing gear:
Retractable landing gear (DC-9,Tu-114,SAAB-35)
Non- retractable landing gear (pushpak, An-14, Fuji KM-2)
Tail wheel landing gear (HT-2,Dakota,Cessana J85 C)
Nose wheel landing gear (Avro-748, Tu-134,F-5A)
Bicycle landing gear (Yak-25,HS-P,112)
THE DESIGN
Design is a process of usage of creativity with the knowledge of science
where we try to get the most of the best things available and to overcome the pitfalls the
previous design has. It is an iterative process to idealism toward with everyone is marching
still.
Design of any system is of successful application of fundamentals of
physics. Thus the airplane design incorporates the fundamentals of aerodynamics,
structures, performance and stability & control and basic physics. These are based on
certain degree of judgment and experience. Every designer has the same technical details
but each design prevails its own individuality and the mode of the designer.
Here the preliminary design has been done of an executive Transport
Aircraft. The basic requirements are the safe, comfortable and economic transport mode
with reasonable time period of flight. Here comfort and safety are given primary
importance.
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FIG: 3
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COMPARATIVE DATA SHEET
In the designers perspective it is necessary to compare the existing
airplanes that are of same type as that of our desired airplane. Their important parameters,
positive aspects to be considered and pitfalls to be overcome are taken into consideration.
Manufacturer AIRBUS BOEING BOEING McDON.
Type
Model
A340-
600
747-
400
777-
300
/DOUG.
MD-12LR
Initial service date 2002 1988 1998 -
Engine Manufacturer R-R P&W4062 R-R R-R/GE/PW
Model / Type Trent 556 4056 Trent 895 CF6-80C2
No. of engines 4 4 2 4
SFC .54lb/lbf-hr .56lb/lbf-hr .575lb/lbf-hr .23lb/lbf-hr
Bypass ratio 36:03:01 5.0:1 38:04:01 5-5.31
Dry Weight 4835kg 4890kg 5942kg 4472.42kg
Diameter 2.5m 2.54m 3m 2.69m
Length 3.9m 4.41m 4.36m 4.26m
Static thrust (kN) 249.1 252.4 423.0 284.7
Accommodation:
Max. seats (single
class)
475 660 550 660
Two class seating 440 496 479
Three class seating 380 412 394 481
No. abreast 9 10 10 11/8
Hold volume (m) 187.74 171.00 200.50 126.40
Volume per
passenger
0.40 0.26 0.36 0.19
Mass (Weight) (kg):
Ramp 365900 397730 299600
Max. take-off 365000 396830 299370 430846
Max. landing 254000 285760 237685 291468
Zero-fuel 240000 242670 224530 273308
Max. payload 63000 61186 68570 85489
Max. fuel payload 29311
Design payload 36100 39140 45695
Design fuel load 151890 176206 197332
Operational empty 177010 181484 155960 187819
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Manufacturer AIRBUS BOEING BOEING McDON.
Type
Model
A340-
600
747-
400
777-
300
/DOUG.
MD-12LR
Weight Ratios:
Ops empty/Max. T/O 0.485 0.457 0.521 0.436Max. Payload/Max.
T/O
0.173 0.154 0.229
Max. Fuel/Max. T/O 0.423 0.407 0.452
Max. Landing/Max.
T/O
0.696 0.720 0.794 0.677
Fuel (litres):
Standard 195620 204350 171170
Optional 216850
Fuselage:
Length (m) 69.57 68.63 72.88 58.82
Height (m) 5.64 8.10 6.20 8.51
Width (m) 5.64 6.50 6.20 7.47
Finess Ratio 12.34 10.56 11.75 7.87
Wing:
Area (m) 437.30 525.00 427.80 543.00
Span (m) 61.20 62.30 60.90 64.92
MAC(m) 8.35 9.68 8.75 9.80
Aspect Ratio 8.56 7.39 8.67 7.76
Taper Ratio 0.220 0.275 0.149 0.215
Average (t/c) % 9.401/4 Chord Sweep () 31.10 37.50 31.60 35.00
High Lift Devices:
Trailing Edge Flaps
Type
S2 S3 S2/S1 S2
Flap Span/Wing Span 0.625 0.639 2.758
Area (m2) 78.7
Leading Edge Flaps
Type
Slats kruger Slats slats
Area (m) 48.1
Manufacturer AIRBUS BOEING BOEING McDON.
Type
Model
A340-
600
747-
400
777-
300
/DOUG.
MD-12LR
Vertical Tail:
Area (m) 47.65 77.10 53.23 96.10
Height (m) 9.44 10.16 9.24 12.90
Aspect Ratio 1.87 1.34 1.60 1.73
Taper Ratio 0.350 0.330 0.290 0.345
1/4 Chord Sweep () 45.00 45.00 46.00 40.00
Tail Arm (m) 27.50 30.00 31.65 24.50
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Sv/S 0.109 0.147 0.124 0.177
SvLv/Sb 0.049 0.071 0.065 0.067
Horizontal Tail:
Area (m) 93.00 136.60 101.26 113.80
Span (m) 21.50 22.08 21.35 22.55
Aspect Ratio 4.97 3.57 4.50 4.47
Taper Ratio 0.360 0.265 0.300 0.326
1/4 Chord Sweep () 30.00 32.00 35.00 35.00
Tail Arm (m) 28.60 32.50 32.95 24.67
Sh/S 0.213 0.260 0.237 0.210
ShLh/Sc 0.729 0.874 0.891 0.528
Undercarriage:
Track (m) 10.70 11.00 11.00 11.59
Wheelbase (m) 32.50 25.60 25.80 26.84
Turning radius (m) 42.80 41.00No. of wheels
(nose;main)
2;12 2;16 2;12 2;16
Main Wheel
diameter (m)
1.250 1.118
Main Wheel width
(m)
0.457
Nacelle:
Length (m) 6.10 5.64 7.30 7.27
Max. width (m) 3.05 2.90 3.20 3.10
Spanwise location 0.296/0.625 0.376/0.667 0.326 0.370/0.630
Manufacturer AIRBUS BOEING BOEING McDON.
Type
Model
A340-
600
747-
400
777-
300
/DOUG.
MD-12LR
Loadings:
Max. power load
(kg/kN)
366.32 393.06 353.87 378.33
Max. wing load
(kg/m2)
834.67 755.87 699.79 793.45
Thrust/Weight Ratio 0.2783 0.2593 0.2881 0.269
Take-off (m):
ISA sea level 3100 3310 3080
ISA +20C SL. 3550 3600 3540
ISA 5000ft 4250 4390
ISA +20C 5000ft
Landing (m):
ISA sea level. 2240 2130 1860 2577
ISA +20C SL. 2240 2130 1860
ISA 5000ft 2410
ISA +20C 5000ft 2410Speeds (kt/Mach):
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V2 185
Vapp 144 153
Vno/Mmo 330/M0.86 365/M0.92 330/M0.87 /M0.85
Vne/Mme 365/M0.93 445/M0.97
CLmax(T/O) 1.92
CLmax(L/D @ MLM) 2.87 2.38
Max. cruise :
Speed (kt) 507
Altitude (ft) 35000
Fuel consumption
(kg/h)
11370
Long range cruise:
Speed (kt) 490
Altitude (ft) 35000
Fuel consumption
(kg/h)
9950
Manufacturer AIRBUS BOEING BOEING McDON.
Type
Model
A340-
600
747-
400
777-
300
/DOUG.
MD-12LR
Range (nm):
Max. payload 5700 6857 8000
Design range 7500 7100 5604Max. fuel (+ payload) 7800 8310
Ferry range 8800
Design Parameters:
W/SCLmax 2857.63 3117.51
W/SCLtoST 3912.54 4579.90
Fuel/pax/nm (kg) 0.0460 0.0500
Seats x Range
(seats.nm)
3300000 3521600
TABLE: 1
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DESIGNING OF SUPER JUMBO AIRCRAFT
REQUIRMENTS:
Passengers: 820(1-class)
Range: 10300 km
Pay load: 83900 Kg
Cruise Mach: 0.85
Altitude: 35000ft
CALCULATION OF CRUISE VELOSITY
Temperature at 35000ft: 216k
Velocity of sound at 35000ft.=297.88m/s
V cruise = M * velocity of sound
=0.85*297.88
=253.055
=253m/s
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MISSION PROFILE
FIG: 4
0-1Take-off
1-2Climbing
2-3Cruising
3-4Descending
4-5Loitering
5-6Descending
6-7Landing
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SEGMENT DETAILS
MISSION
SEGMENT
DESCRIPTION ALTITUDE DISTANCE TIME
0-1 GROUND
RUN
0 2750m 5min
1-2 ASCENT 0-13Km 20 Km 6 min
2-3 CRUSING 13 Km 10300 Km 9 hrs.
3-4 LOITER13 Km
5 km 30min
4-5 DECENT 13-0 Km 15 Km 5 min
5-6 LANDING
0 2050 m 2min
TABLE: 2
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ESTIMATION OF WEIGHT
The weight of the aircraft (W) is the key factor in almost aircraft performance problems. The
gross weight is distributed in the following manner:
W = Wstruc+ Wcrew+ Wpass+ Wfe+ Wpp+ Wf
Here,
Wstructureconsists of the wing, fuselage, under-carriage & the empennage and accounts
for about 32% of the gross weight, i.e., 0.32W.
Wfixed equipment includes the passenger seats, food, baggage racks, lavatories, air-
conditioning, avionics and other passenger amenities. This adds to the weight by about0.05W.
Wpowerplant is the weight of the engine and its systems. The initial assumption of engine
weight is assumed to be 0.055W which may be modified later to suit thrust requirements.
Wfuel is the weight contribution of the fuel to the total weight. It depends on the range also
includes the Reserve fuel that is used in case of an emergency. It adds to the gross weight by
a factor of 0.3W.
Wcrew+ Wpassengersaccounts for the remaining weight. i.e., 0.275W. Taking passenger &
baggage weight into consideration, a maximum of 1800N per passenger is permissible. As
for a crew member, 1000N would suffice.
WARMUP AND TAKE OFF:
W1/W0=0.97
W0=Takeoff weight
W1= Weight at the end of take off
CLIMB:
W2/W1-0.985
W1= Weight at the start of climb
W2= Weight at the end of climb
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CRUISE:
W3/W2= W2 = Weight at the start of cruise
W3= Weight at the end of cruise
RANGE: HEAD WIND CORRECTION
Gross safe range = Gross still in air range/1.5
=10300/1.5
=6866.667
Vcru = 253m/s
Vcr= 910.8km/hr.
Time=7.539 hr.
Head wind = 15m/s
= 54km/hr.
Additional distance = 7.539*54
=407.106km
TOTAL RANGE = FERRY RANGE+RANGE CORRECTION FOR THE HEAD WIND
= 10300+407.106
=10707.106km
(L/D)cru =0.866*(L/D)max
TO FIND (L/D)max
Aspect ratio = 7.5
From the Wetted area ratio chart (chart 1)
For swept back wing Swet/Sref=6
Wetted aspect ratio = b2/Swet
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Chart: 1 Wetted area ratios
Wetted aspect ratio = A/(Swet/Sref)
= 7.5/6
=1.25
From Maximum lift to drag ratio trends chart (chart-2)
(L/D)max= 17
For cruise,
(L/D) = 0.866*(L/D)max
=0.866*17
= 14.722
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Chart 2
W3/W2= =
= 0.6708
W3/W2=0.6708
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LOITER:
W3/W4
E = (1/ct)*(L/D)max*ln(W3/W4)E = 30 min = 0.5 hr.
0.5 = (1/0.4)*17*ln(W3/W4)
0.5 = 42.5*ln(W3/W4)
Ln (W3/W4) = 0.0117647
0.0117647
= 1.011834
W4/W3= 1/1.011834
= 0.98813
W4/W3= 0.98813
LAND:
W5/W4 = 0.995
W4= Initial weight while landing
W3= Final weight while landing
HALTING:
Wf/Wg= 1.06(1-Wh/Wg)
Wh/Wg=W5/W4*W4/W3*W3/W2*W2/W1*W1/W0
= 0.995*0.98813*0.6708*0.985*0.97
= 0.63014
Wf/Wg = 1.06(1-0.63014)
= 0.392052
Wf/Wg = 0.392052
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WEIGHT RATIOS FOR DIFFERENT (L/D)cru VALUES
TABLE: 3
(L/D)max
W1/
W0
W2/
W1
W3/
W2
W4/
W3
W5/
W4
Wf/
Wg
(L/D)cruise
11 0.97 0.985 0.539 0.978 0.995 0.502 9.526
12 0.97 0.985 0.568 0.981 0.995 0.530 10.392
13 0.97 0.985 0.593 0.982 0.995 0.554 11.258
14 0.97 0.985 0.616 0.984 0.995 0.576 12.124
15 0.97 0.985 0.636 0.985 0.995 0.595 12.99
16 0.97 0.985 0.654 0.986 0.995 0.613 13.856
17 0.97 0.985 0.671 0.991 0.995 0.630 14.722
17.43 0.97 0.985 0.677 0.989 0.995 0.637 15.094
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EMPTY WEIGHT
We= Empty weight of the aircraft.
We/Wg= AWgC
A=1.02
C= -0.06
We/Wg = 1.02(Wg) ^-0.06
DIFFERENT WgVALUES FOR VARIOUS (L/D)
TABLE: 4
Wf/Wg (L/D)max Wpay/Wg Wg We/Wg
0.5277 11 0.0108 550000 0.4615
0.4985 12 0.0401 551000 0.4614
0.4727 13 0.0666 555000 0.4613
0.4497 14 0.0891 557000 0.4612
0.4289 15 0.1101 560000 0.4610
0.4103 16 0.1291 565000 0.4607
0.3921 17 0.1475 569000 0.4605
0.3852 17.43 0.1543 571000 0.4604
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TO FIND THE GROSS WEIGHT
Wg= Wpay + Wf + We
1=Wpay/Wg + Wf/Wg + We/Wg
Wg= Wpay/(1- Wf/WgWe/Wg)
We/Wg = 1.02(Wg)^(-0.06)
Wg = Wpay/ {1-Wf/Wg-[1.02(Wg)^(-0.06)]}
FOR
(L/D) max = 17;
Wf/Wg= 0.3921
By substituting the values in above equation we get
We/Wg= 1.02(569000) ^ (-0.06)
= 0.4601
Wg = (Wcrew+ Wpayload)/ {1-Wf/Wg-[1.02(Wg)^(-0.06)]}
= (73800+10100) / (1-0.3921-0.4601)
=569199.46 kg
Wg= 569199.46
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We/Wg= 1.02(569000) ^ (-0.06)
= 0.4601
Wpay/Wg= 83900/569199.46
= 0.1474
Wpay/Wg + We/Wg + Wf/Wg = 1
0.1474+0.461+0.3921 = 1.0005
1.0005 ~ 1
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GRAPH: 1
(L/D)max Wf/Wg
11 0.5277
12 0.4985
13 0.4727
14 0.4497
15 0.4289
16 0.41025
17 0.39205
17.43 0.38518
TABLE: 5
0
0.05
0.1
0.15
0.2
0.25
0.3
0.35
0.4
0.450.5
0.55
0.6
0 2 4 6 8 10 12 14 16 18 20
Wf/Wg
(L/D)max
(L/D)maxvs Wf/Wg Wf/Wg
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GRAPH: 2
(L/D)max Wg
11 550000
12 551000
13 555000
14 557000
15 560000
16 565000
17 569000
17.43 571000
TABLE: 6
547500
550000
552500
555000
557500
560000
562500
565000
567500
570000
572500
0 2 4 6 8 10 12 14 16 18 20
Wg
(L/D)max
(L/D)maxVsWg
Wg
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SELECTION OF WING LOADING BASED ON
LANDING SPEED/LANDING DISTANCE
Approach Velocity:
Va~ 1.3(Vs)land
VTD ~ 1.15(Vs) land
Sland(feet) = 0.3{ Va (in knots)}2
Vs = {2Wland/(S*CLmax*0*)}0.5
= (2Pland/CLmax*0)0.5
Pland = (CLmax* 0*Vs2)/2
=1.0, In the sea level unless otherwise prescribed landing altitude.
CLmax =3
Landing Distance, land(in ft.)=6725.72ft.
Approach Velocity.
Va = (Sland /.3).5
=149.7299knots.
1 Knot=1.853km/hr.
=0.5148m/s
Va=77.08m/s
Stalling Velocity.
Vstall=Va/1.3
=59.2923m/s
Pland=(3*1.225*1*59.232)/2.
=6446.30N/m3.
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Pland= 6446.30
Plandfor Vs=6446.30N/m3.
Pland for Vs+10% =7816.445N/m3.
Pland for Vs-10% =5232.4966N/m3.
W/S = Pland*(WTO/Wland).
For Stalling Velocity of 59.8m/s.
Pland =(3*1.225*1*59.82)/2
=6570.97N/m2
S =Wg/Pland
= (569000*9.81)/6570.97
= 849.476m2.
S = 849.476m2
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GRAPH: 3
W/S Vs
4593.75 50
6615 60
9003.75 70
11760 80
14883.75 90
18375 100
TABLE : 7
0
10
20
30
40
50
60
70
80
90
100
110
0 2500 5000 7500 10000 12500 15000 17500 20000
Vs
W/S
W/S vs Vs
Vs
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SELECTION OF WING LOADING BASED ON
MAXIMUM SPEED
For a High Subsonic Aircraft
Mmax = Mcric+0.04
=0.9+0.04
=0.94
So Maximum Velocity Vmax= 278.18m/s.
t`=Tvmax/W
t`=(.5**Vmax2*s*Cd)/W
0.5Vmax2=qmax ; W/S=P
t`=(Cdqmax)/P
CD0=Cfe*Swet/S
Estimation Of Wetted Area:
CD0(approximate)=0.015
Log10Swet=C+d*log10WTO
WTOin lbs.& Swetin ft.2
Cfe= 0.0030
From reference 2 Table 3.5,Page 122
C=0.0199
d=0.778
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Log10Swet= 0.0199+0.778*log10 (1254430.18*0.97)
=3.0625
Swet = 56780.57ft2
=5223.813m2
Cd0= Cfe*(Swet/S)
=0.003*(6)
=0.018~0.015
Determination Of Drag:-
Drag is the resolved component of the complete aerodynamic force which is parallel to the
flight direction (or relative oncoming airflow). It must always act to oppose the direction of
motion.
It is the undesirable component of the aerodynamic force while lift is the desirable
component.
There are only two sources of aerodynamic force on a body moving through a fluid- thepressure distribution and the shear stress distribution acting over the body surface.
Therefore there are only two general types of drag:
Pressure Drag: due to a net imbalance of surface pressure acting in the drag direction.
Friction Drag: due to the net effect of the shear stress acting in the drag direction.
Amount of drag generated depends on the Planform area (S), air density (), flight speed (V),
drag coefficient (CD) CDis a measure of aerodynamic efficiency and mainly depends upon
the Section shape, Planform geometry, angle of attack (), compressibility effects (Mach
number), and viscous effects (Reynolds number).
Cd=Cd0+KCl2.
Cd0Parasite Drag Coefficient.
D0= D0wing+D0Fus+DoNac+D0HT+D0VT+D0ETC
CD0=D0/ (0.5V2
S)
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Page | 40 ADL
K = 1/ (*A*)
e = 0.85
A = 7.5
K = 1/(*0.85*7.5)
K = 0.04993
Cd= Cd0+KCL2
= 0.4671
Estimation of drag polar
Configuration Cd0
Clean - 0.8 to 0.85
Take off flaps 0.01 to 0.02 0.75 to 0.8
Landing flaps 0.05 to 0.075 0.7 to 0.75
Landing gear 0.015 to 0.025 No effect
TABLE : 8
1. Clean configuration
CD clean = KCL2
= 0.04993*32
= 0.4494
2. Take- off flaps (gear up)
CD = CD0 + KCL2
e (1) = 0.8
CD0(1) = (0.02+0.025)
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Page | 41 ADL
Cd= 0.5225
2. Take- off flaps (gear up)
CD = CD0 + KCL2
e (2) = 0.8
CD0(2) = 0.07
Cd= 0.5475
4. Landing flaps (gear up)
CD0(3) = 0.095
e(3) = 0.75
Cd= 0.6043
5. Landing flaps (down up)
CD0(4) = 0.1
e(4) = 0.75
Cd= 0.6093
Break up Drag Polar
CD= F1 + F2(w/s) + F3 (w/s)2
F1= sum of the CD values of wing, stabilizers area.
F1 = Cfe* (Swet/s)wing*(1 + Sstabilizer/S)
= 0.003*(Swet/s)wing*(1 + Sstabilizer/S)
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Page | 42 ADL
(Swet/s)wing= 6.0
( Sstabilizer/S) = 0.26
F1= 0.003*3*(1+0.26)
F1 = 0.02275
F2= (CD-F1)/ (w/s)
= 0.4671-0.02275/6446.30
F2 = 6.893*10-5
F3= (*A**(0.5**V2
max)2
)-1
; {K/q2}
At 13km the density is 0.266kg/m3
F3= (*7.5*0.85*(0.5*0.266*278.182)2)-1
F3= 4.9337*10-10
dt`/dP = 0;{Recall thrust loading equation}
qmax*(-F1/P2+0+F3) = 0
Pvmax=
= Pvmax= 6791.0220
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Page | 43 ADL
GRAPH : 4
TABLE : 9
0
50
100
150
200
250
300
350
400
0 2000 4000 6000 8000 10000 12000
Vmax
W/S
W/S vs Vmax
Vmax
W/S Vmax
877.5039 100
1974.38 150
3510.016 200
5484.399 250
7897.535 300
10749.42 350
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Page | 44 ADL
SELECTION OF WING LOADING BASED ON
ABSOLUTE CELING
t `Hmax=Dmin/W
= 1/ (W/D)max
CL/ (L/D)max= (CD0/k)0.5
(CD)(L/D) max= 2CDo
t `Hmax= 1/(L/D)max
=
==
q Hmax= 0.5*Hmax*Vmax2
t `H= Treq/W
= qHmax*( )P =
* +
CDo= 0.018; K = 0.04993; Hmax= 0.266
P = P = 6179.579
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Page | 45 ADL
For different values, we get the results as shown in the below table.
GRAPH : 5
W/S VHmax
789.55 100
1796.738 150
3194.22 200
4990.938 250
6179.508 278.18
7186.95 300
9782.238 350
TABLE: 10
0
50
100
150
200
250
300
350
400
0 2000 4000 6000 8000 10000 12000
VHmax
W/S
W/S vs VHmax
VHmax
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SELECTION OF WING LOADING BASED ON
RATE OF CLIMB
Vc = V (T D)/W
Vc= V (t `R/C)r= (Vc/V) + 0.5*0*V
2/P*CD0
CD= F1+F2P+F3P2
(t `R/C)r= (Vc/V) + q[F1/P+F2+F3P]
Dt`R/C/dP = 0
PR/C= (F1/F3)0.5
= q (F1A)0.5
(t `R/C)r= (Vc/V) + q[ ]W/S =
V(R/C) max= ( ) L/D = 17; T/W = 0.222; mean= (1.225+0.266)/2 = 0.7455
(VR/C)2
= {0.222/ (3*0.7455 *0.018)}*W/S*
[
]
= 11.58W/S
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GRAPH: 6
W/S V(R/C)max
5397.236 250
5837.651 260
6295.337 270
6770.294 280
7262.522 290
7772.02 300
8298.79 310
8842.83 320
TABLE: 11
0
50
100
150
200
250
300
350
0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000
V(R/C)max
W/S
W/S vs V(R/C)max
V(R/C)max
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COMPARATIVE GRAPH
GRAPH: 7
0
50
100
150
200
250
300
350
400
0 2000 4000 6000 8000 10000 12000 14000 16000 18000 20000
V(R/C)max
W/S
W/S vs V(R/C)max
Vs
Vmax
VHmax
V(R/C)max
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COMPARATIVE TABLE
W/S Vs W/S Vmax W/S VHmax W/S V(R/C)max
4593.75 50 877.5039 100 789.55 100 5397.236 250
6615 60 1974.38 150 1796.738 150 5837.651 260
9003.75 70 3510.016 200 3194.22 200 6295.337 270
11760 80 5484.399 250 4990.938 250 6770.294 280
14883.75 90 7897.535 300 6179.508 278.18 7262.522 290
18375 100 10749.42 350 7186.95 300 7772.02 300
TABLE: 12&13
S. No. Design Parameter Value Unit
1. Aspect Ratio 7.5 (no unit)
2.
Wing Span 79.75 m
3.
Height 24.45 m
4. Length 72.73 m
5. Wing Area 849.47 m2
6. Max Speed 1001.448 km/hr
7. Cruise Speed 910.8 km/hr
8. Range 10300 km
9. Service Ceiling 43,028 ft
10. Rate of Climb 55.55 m/s
11. Max Take-Off Weight 569000 kg
12. Empty Weight 262000 kg
13. Payload 83900 kg
14. Crew Members 2 (no unit)
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Page | 50 ADL
FUSELAGE DESIGN
INTRODUCTION
The fuselage (from Frenchfusel "spindle-shaped") is an aircraft's main body
section that holds crew and passengers or cargo. In single-engine aircraft it will
usually contain an engine, although in some amphibious aircraft the single
engine is mounted on a pylon attached to the fuselage which in turn is used as
a floating hull. The fuselage also serves to position control and stabilization
surfaces in specific relationships to lifting surfaces, required for aircraft
stability and maneuverability.
Common practice to modularise layout:
Crew compartment, power plant system, payload configuration, fuel
volume, landing gear stowage, wing carry-through structure,
empennage, etc.
Or simply into front, centre and rear fuselage section designs.
Functions of fuselage:
Provision of volume for payload.
Provide overall structural integrity.
Possible mounting of landing gear and power plant.
Once fundamental configuration is established, fuselage layout proceeds
almost independently of other design aspects.
PRIMARY CONSIDERATIONS
Most of the fuselage volume is occupied by the payload, except for:
Single and two-seat light aircraft. Trainer and light strike aircraft.
Combat aircraft with weapons carried on outer fuselage & wing.
High performance combat aircraft.
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Payload includes:
Passengers and associated baggage.
Freight.
Internal weapons (guns, free-fall bombs, bay-housed guided weapons).
Crew (significant for anti-sub and early-warning aircraft).
Avionics equipment.
Flight test instrumentation (experimental aircraft).
Fuel (often interchangeable with other payload items on a mass basis).
Pressurisation: If required, has a major impact upon overall shape.
Overall effect depends on level of pressurisation required.
Low Differential Pressurisation:
Defined as no greater than 0.27 bar (4 psi).
Mainly applicable to fighters where crew are also equipped with
pressure suits.
Cockpit pressurisation primarily provides survivable environment in case
of suit failure at high altitude.
Also used on some general aviation aircraft to improve passenger
comfort at moderate altitude.
Pressure compartment has to avoid use of flat surfaces.
Normal (High) Differential Pressurisation:
Usual requirement is for effective altitude to be no more than 11 km
(32000 ft) ISA for passenger transports.
Implied pressure differentials are:
o 0.37 bar (5.5 psi) for aircraft at 7.6 km (25,000 ft).
o 0.58 bar (8.5 psi) for aircraft at 13.1 km (43,000 ft).
o 0.65 bar (9.4 psi) for aircraft at 19.8 km (65,000 ft).
High pressure differential required across most of fuselage for passenger
transports so often over-riding fuselage structural design requirement.
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Tail Shape
Smooth change in section required, from maximum section area to
ideally zero.
Minimisation of base area especially important for transonic/supersonicaircraft.
Important parameter for determining tail upsweep angle is ground
clearance required for take-off and landing rotation.
Typically 12oto 15
o.
FIG: 5
Typical tail section lengths are:
o 2.5 to 3.0 x diameter (subsonic)
o 6 to 7 x diameter (supersonic)
Centre Fuselage & Overall Length - Subsonic Aircraft
Theoretically minimum drag for streamlined body with fineness ratio(length/diameter) of 3.
In reality, typical value is around 10, due to:
o Need to utilise internal volume efficiently.
o Requirement for sufficiently large moment arm for
stability/control purposes.
o Suitable placement of overall CG.
Wing Location - Aerodynamics Considerations Mid-wing position gives lowest interference drag, especially well for
supersonic aircraft.
Top-mounted wing minimises trailing vortex drag, especially good for
low-speed aircraft.
Low wing gives improved landing gear stowage & more usable flap area.
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Page | 54 ADL
From the above given locations of wings, the one chosen is the Low wing
configuration which gives improved landing gear stowage & more usable flap
area.
Empennage Layout
Vertical Surface
Single, central fin most common arrangement, positioned as far aft as
possible.
Horizontal Surface
Efficiency affected by wing downwash, thus vertical location relative to
wing important.
Usually mounted higher than wing except on high wing design or with
small moment armlow tail can give ground clearance problems.
Avionics & APU
Including navigation, communications and flight control/management
equipment.
Provision necessary for adequate volume in correct location with ease of
access.
Location of radar, aerials, etc also important
o
Sensors often have to face forward/down in aircraft nose.
o
Long range search & early warning scanners sometimes located on
fuselage.
Auxiliary power unit (APU) commonly located at extreme rear of
fuselage on transport aircraft.
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2D VIEW OF FUSELAGE
FIG: 9
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Page | 57 ADL
SELECTION OF AIRFOIL
The aircraft which is to be designed having a High Subsonic cruise speed say
Mach 0.85 which belongs to transonic speed, so that to avoid profile drag
SUPERCRITICAL AIRFOILS are chosen.
From the aerofoil data book various airfoils of required t/c are taken and are
tabulated for maximum lift coefficient and minimum drag.
TABLE: 14
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Page | 59 ADL
FIG: 12 Plots of Clvs CDand Clvs from XFLR5.( NASA SC(2)0610)
FIG: 13 The airfoil NASA SC(2)0606 created using JAVAFOIL software by entering the co-ordinates.
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FIG: 14 The airfoil NASA SC(2)0606 imported to XFLR5 An Airfoil Testing software.
FIG: 15 Plots of Clvs CDand Clvs from XFLR5.( NASA SC(2)0606)
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COMBINED PLOT FOR ROOT AND TIP AIRFOILS.
FIG: 16 Plots of Clvs CDand Clvs from XFLR5.( NASA SC(2)0610 & NASA SC(2)0606)
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Airfoil for Horizontal Tail Plane.
Airfoil used in Horizontal Tail Plane is NASA SC(2)0710.
FIG: 17The airfoil NASA SC(2)0710 created using JAVAFOIL software by entering the co-ordinates.
FIG: 18 The airfoil NASA SC(2)0710 imported to XFLR5 An Airfoil Testing software.
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FIG: 19Plots of Clvs CDand Clvs from XFLR5.( NASA SC(2)0710)
Airfoil for Vertical Tail Plane.
Airfoil used in Vertical Tail Plane is NASA SC(2)0010.
FIG: 20The airfoil NASA SC(2)0010 created using JAVAFOIL software by entering the co-ordinates.
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FIG: 21 The airfoil NASA SC(2)0010 imported to XFLR5 An Airfoil Testing software.
FIG: 22Plots of Clvs CDand Clvs from XFLR5.( NASA SC(2)0010)
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2-D VIEW OF THE WING
FIG: 23
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VERTICAL TAIL
VERTICAL TAIL = 30.5*(t/c) wing
ASPECT RATIO =1.39
TAPER RATIO = 0.424
t/c = 8
VERTICAL TAIL =1.73*8
= 13.66 m
LANDING GEAR SELECTION
In aviation, the undercarriage or landing gear is the structure (usually
wheels) that supports an aircraft and allows it to move across thesurface of the earth when it is not in flying. More importance is to be
given as it carries the entire load on the ground. Landing gear usually
includes wheels equipped with shock absorbers for solid ground, but
some aircraft are equipped with skis for snow or floats water, and
skids or pontoons (helicopter)
FUNCTIONS OF LANDING GEAR carry aircraft max gross weight to take off runway
withstand braking during aborted take off
retract into compact landing gear bay
Damp touchdown at maximum weight.
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TYPES OF GEAR ARRANGEMENTS
Wheeled undercarriage comes in two types: conventional or tail
dragger undercarriage, where there are two main wheels towards
the front of the aircraft and a single, much smaller, wheel or skid atrear; tricycle undercarriage where there are two main wheels under
the wings and a third smaller wheel in the nose. most modern
aircraft have tricycle undercarriage. Sometimes a small tail wheel or
skid is added to aircraft with tricycle undercarriage arrangements.
RETRACTABLE GEAR
To decrease drag in flight some undercarriages retract into the wingsand/or fuselage with wheels flush against or concealed behind doors,
this is called retractable gear. It was in late 1920s and 1930s that
such retractable landing gear became common. This type of gear
arrangement increased the performance of aircraft by reducing the
drag.
FIG: 24
STEERING OF LANDING GEAR
The steering mechanism used on the ground with wheeled landinggear varies by aircraft, but there are several types of steering.
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RUDDER STEERING
DIRECT STEERING
TILLER STEERING
Maximum Takeoff Weight of the aircraft (from Weight Estimation) =
272.655t = 2672kN
TYRE SIZING
During landing and takeoff, the undercarriage supports the total
weight of the airplane. Undercarriage is of three types
Bicycle type
Tricycle type
Tricycle tail wheel type
FIG: 25
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ISOMETRIC VIEW DIAGRAM OF AIRCRAFT
FIG: 29
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Conclusion
The aircraft is designed and the parameters like cruise
velocity, wing loading, span etc... have been selected for ouraircraft. The weight estimation had been done to estimate
the weight of our aircraft. The wings, airfoil, landing gear
have been selected for our aircraft. The performance
calculations were also made to estimate the performance.
The aircraft parameters are in the optimum range and design
characteristics have been found to be satisfactory.
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REFERENCES
1. Aircraft Design: A Conceptual Approach 2NDEdition
Daniel P. Raymer
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