Aircraft Design Lab – i - Super Jumbo Aircraft

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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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    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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    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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    (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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    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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    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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    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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    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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    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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    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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    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

    President, Conceptual Research Corporation

    2. Airplane Design: Preliminary Sizing of Airplanes 3RD

    Edition

    Dr. Jan Roskam

    3. Lecture Notes on Aircraft Design, Department of Aerospace Engineering

    I.I.T Madras, 2007Tulapurkara.E.G

    4. The Design of the Airplane

    Darrol Stinton

    5.Aircraft Design, Cambridge Aerospace Series

    Ajoy Kumar Kundu

    6. Aircraft Performance and Design 2ND

    Edition

    John D. Anderson, Jr.

    7.Janes All the Worlds Aircraft 1999-2000,Janes information group ltd., Surrey,UK, 1999.

    Jackson, P. (Editor)

    8. Pilots Handbook of Aeronautical Knowledge 2ND

    Edition

    Federal Aviation Administration

    9. Wikipedia

    www.wikipedia.org

    10. Airfoil Investigation Database