Design project IHINDUSTAN COLLEGE

OF ENGINEERING10 SEATER BUSINESS JET

THRUSTMASTER-I

By, (reg no: 30507101064) (reg no: 30507101075) (reg no: 30507101306)

Acknowledgement

I would like to extend my heartful thanks to Dr. P.K. Dash, Head of Aeronautical Department for giving me his able support and encouragement. At this juncture I must emphasis the point that this design project would not have been possible without the highly informative and valuable guidance by senior Prof. P.S. Venkatnarayanan, whose vast knowledge and experience has made us go about this project with great ease. We have great pleasure in expressing our sincere and whole hearted gratitude to them.

It is worth mentioning about my teammates, friends and colleagues of the aeronautical department for extending their kind help whenever required. I thank one and all who have directly or indirectly helped me in making yhis design project a great success.

INTRODUCTION

Business jet, private jet or, colloquially, bizjet is a term describing a jet aircraft, usually of smaller size, designed for transporting groups of business people or wealthy individuals. Business jets may be adapted for other roles, such as the evacuation of casualties or express parcel deliveries, and a few may be used by public bodies, governments or the armed forces. The more formal terms of corporate jet, executive jet, VIP transport or business jet tend to be used by the firms that build, sell, buy and charter these aircraft.

The business jet industry groups the jets into five loosely-defined classes:

Heavy jets

The most exclusive type of private jet is the heavy jet type, which is designed for the ultimate in large capacity luxury air travel. Aircraft of this class includes:

Airbus o Airbus A318 Elite

o Airbus A319CJ

Boeing

o Boeing Business Jet

Large size jets

Bombardier Aerospace o Bombardier Global 5000

o Bombardier Global Express

Super mid-size jets

The elite class of the business and private jet aircraft are the super mid-size jets that feature wide body cabin space, high altitude, speed, and ultra long range capabilities. These ultra luxurious private jets combine the long range

transatlantic capability with the speed and comfort of a wide body, high altitude aircraft. Aircraft of this class include:

Bombardier Aerospace o Bombardier Challenger 300

o Challenger 605

Mid-size jets

These aircraft are suitable for longer range travel such as transcontinental flights and for travel with larger passenger capacity requirements. Aircraft of this class includes:

Bombardier Aerospace o Learjet 60 XR

o Learjet 85

Gulfstream

o Gulfstream 150

o Gulfstream 250

Light jets

The light jets have been a staple of the business jet industry since the advent of the Lear Jet in the early 1960s. The light jets provide access to small airports and the speed to be an effective air travel tool. Aircraft of this class includes:

Bombardier Aerospace o Learjet 40

o Learjet 40 XR

o Learjet 45

Very light jets

Very light jets, also known as Micro jets or VLJs, are designed to provide air travel, for example, to the more than 5,000 small community airports in the United States. VLJs have a maximum takeoff weight of not more than 10,000lb. Aircraft of this class include:

Adam Aircraft Industries o Adam A700

DESIGN PROCESS

Airplane design is the intellectual engineering process of creating on

paper or in a computer aided screen , a flying machine to (1) meet certain specifications and requirements established by potential users (or as perceived by the manufacturer) and/or (2) pioneer innovative, new ideas and technology. An example of the former is the design of the most commercial transports, starting at least with the Douglas DC IN 1932, which was designed to meet or exceed various specifications by an airline company. An example of the latter is the design of the rocket powered Bell X-1, the airplane designed to exceed the speed of sound in the level of climbing flight. The design process is in fact an intellectual activity, but a rather a special one that is tempered by good intuition developed through experience, by attention paid to successful airplane design that have been used in the past, and by (general proprietary) design procedures and databases (handbooks, data books etc.) that are a part of every airplane manufacturer.

Conceptual Design Studies

The first activity in the project design process is the conceptual design study. In this phase conventional and novel configurations are considered to determine layouts which are technically feasible and commercially viable, at the start of the phase all options are considered and during the conceptual design phase the quantity of data generated on each design will be relatively limited. The outcome of the study is the knowledge of the feasibility of various concepts and an estimate of the likely dimension and most favoured configuration.

Conceptual design begins with a set of requirements put forward by a prospective customer or by any organization. The first preference will be given to the customer’s needs. Design details include aircraft range, payload, speed, versatility etc.

The actual design usually begins with a conceptual sketch. The conceptual sketch will include approximate wing and tail geometries, fuselage shape, locations of the engines, landing gear etc.

Preliminary Design Studies

The conceptual designs created will be compared and those which are considered unfeasible or too commercial risky will be eliminated. The remainder will be compared after careful consideration of the demands laid out by the customer. It is important not to carry many options forward to the next stage as this will dissipate the available effort and slow down the detailed definition of the preferred design. However, care must be taken to avoid discarding design layouts too quickly as some may lead to evolutionary configurations which could give the aircraft a competitive advantage over aircraft from other companies.

Detailed Design Studies

The detailed design phase is started towards the end of the preliminary analysis. In this part of the design process the layout is refined to a greater level of detail. With the external shape fixed the structural frame work will be defined. In this phase there will be an increasing reluctance to make radical geometrical changes to the overall layout of the aircraft. Throught this phase the aircraft weight and performance estimates will be continuously updated as more details of the aircraft layout becomes available.

SEVEN INTELLECTUAL POINTS FOR CONCEPTUAL DESIGN:

ABSTRACT

The following design requirements and research studies are set for the project:

Design an aircraft that will transport 10 business class passengers and their baggage over a design range of 3500 km at a cruise speed of about 800 km/h.

To provide the passengers with high levels of safety and comfort.

To operate from regional airports.

To use advanced and state of the art technologies in order to reduce the operating costs.

To offer a unique and competitive service to existing scheduled operations.

To assess the development potential in the primary role of the aircraft.

To produce a commercial analysis of the aircraft project.

ALTERNATIVE ROLES

Our aircraft will have a fuselage size that is more spacious than normally

associated with a 10-seat business jet. The long range requirement will

demand a high fuel load and this will make the aircraft maximum design

weight heavier than normal for 10 seat business class jet aircraft. Both

this aspect suggests that the aircraft can be converted into a higher

capacity short range airliner. A study will be required to investigate such

variants. This type investigation may result in recommendations to

change the baseline aircraft geometry to make such developments easier

to achieve. For example, using the extra fuselage space up to six abreast

seating in the higher capacity aircraft can be made. Without such a

change, six abreast seating may not be feasible.

COMPARITIVE DATA SHEET

To have an idea about the parameters for our design process we compare the important parameters of already existing aircraft similarly related to our interest. The important parameters thus obtained are plotted with respect to the cruising speed and the approximate value of the parameters for the new aircraft is found out.

The parameters involved in the CDS are listed below:

Aspect ratio Maximum takeoff weight Empty weight Power Cruise speed Wing loading Range Length Wing span Wing area

S No NAME OF AIRCRAFT CAPACITY ALTITUDE RANGE SPEED ASPECT RATIO

(passengers) (m) (km) (km/h)

1 Falcon 20F 14+2 12800 3350 862 6.5

2 Piaggio P 180 Avanti 9+2 12500 2592 732 12.97

3 Cessna Citation I 7+2 12450 2559 916 7.3

4 Hawker 850 XP 13+2 12500 4893 830 7.003

5Aero commander 500 S 6+2 5914.8 1915 346 5.31

6 Learjet 24D 7+2 13720 2037 886 9.765

7Raython Hawker 800 XP 10+2 13458 8790 890 9.9

8 A38 8+2 4800 4059 790 7.4

9 Bombadier Learjet 40 7+2 15540 3156 865 9.9

10Bombadier Learjet 60R 9+2 15545 4330 859 12.99

11 Challenger 800 15+2 12497 5129 850 14.82

12 Bombadier Global XRS 12+2 15500 11390 860 9.6

13 Global 5000 13+2 15554 9360 950 15.3

14 Dassult Falcon 14+3 15542 11019 953 9.709

15 Learjet 24E 7+2 13724 2052 886 8.58

S No NAME OF AIRCRAFT WING LOADING WING SPAN LENGTH MAX TOW R/C WING AREA

(km/sq.m) (m) (m) (kg) (m/min) (sq.m)

1 Falcon 20F 354 16.3 17.15 13000 1524 41

2 Piaggio P 180 Avanti 327 14.03 14.41 5239 900 16

3 Cessna Citation I 208 14.35 13.26 5375 828 25.9

4 Hawker 850 XP 441 16.5 15.6 12701 870 34.75

5 Aero commander 500 S 492 14.95 11.22 3060 409 23.7

6 Learjet 24D 299 13.12 14.5 6136 2073 21.53

7 Raython Hawker 800 XP 390 14.56 16.39 9545 859 32

8 A38 471 16.6 12.8 3150 696 67

9 Bombadier Learjet 40 390 14.56 16.93 9545 859 28.95

10 Bombadier Learjet 60R 433 13.34 17.88 16600 660 24.6

11 Challenger 800 451 21.21 26.77 24041 756 48.35

12 Bombadier Global XRS 468 28.65 30.3 44500 793 94.9

13 Global 5000 451 28.65 29.5 41957 810.6 53.29

14 Dassult Falcon 435 26.21 23.19 31750 774.6 70.7

15 Learjet 24E 342 13.35 14.52 6136 850 24.57

DESIGN DATA

PARAMETER VALUE

CRUISE SPEED 720km/hASPECTRATIO 10.25LENGTH 20mWING SPAN 22.88mWEIGHT 16,642kgRANGE 6000kmMAX. TAKEOFF WEIGHT 20,000kgSERVICE CEILING 10200mCAPACITY 10THRUST 98KNMAXIMUM SPEED 750km/hWING LOADING 350kg/m2

WEIGHT ESTIMATION

OVERALL WEIGHT ESTIMATION

The second pivot point in our conceptual design analysis is the preliminary weight estimation of the aircraft. There are various ways to subdivide and categorize the weight components of the airplane.

1. Crew weigh Wcrew : The crew comprises of the people necessary to operate the aircraft.

2. Payload weight Wpayload : The payload is what the airplane is intended to transport i.e. passengers, baggage etc.

3. Fuel weight Wfuel : it is the weight of the fuel in the fuel tank. Since the fuel is consumed during the course of the flight, Wfuel is variable, decreasing with time during the flight.

4. Empty weight Wempty : This is the weight of everything else the structure, the engines, electronic equipments, landing gear, fixed weights etc.

The design take-off weight is the weight of the aircraft at the instant it begins its mission. It includes the weight of all the fuel on board at the beginning of the flight.

W=WSTRUCT+WFIXEDEQUI+WCREW+WPASSENGER+WPOWERPLANT+WFUEL

W fuel is approx. =0.3Wtakeoff

Wstructure is approx. =0.32 Wtakeoff

Wpower plant is approx. =0.055 Wtakeoff

Wfixed equipment is approx. =0.05 Wtakeoff

Wcrew+Wpassengers is approx. =0.275 Wtakeoff

WSTUCT =0.32 Wtakeoff

WSTUCT =6432kg

WFIXED EQUIP =0.05 Wtakeoff

WFIXED EQUIP =1005kg

WPOWERPLANT =0.055 Wtakeoff

WPOWERPLANT =1105.5kg

WFUEL =0.3 Wtakeoff

WFUEL =6030kg

WCREW+WPASSENGER =0.275 Wtakeoff

WCREW+WPASSENGER =5527.5kg

W=0.725W+(no. Of passengers)*180+(crew)*100

=0.725(20,000)+((10*1800)/9.8)+((3*1000)/9.8)

=18,142Kg

Therefore the gross weight of the aircraft is 18,142 Kg

AEROFOIL SELECTION

AIRFOIL TERMINOLOGY

The various terms related to airfoils are defined below :

The mean camber line is a line drawn midway between the upper and lower surfaces

The cord line is a straight line connecting the leading and the trailing edges of the aerofoil, at the ends of the mean caber line.

The chord is the length of the chord line and is the characteristic dimension of the airfoil.

The maximum thickness and the location of it

are expressed as a percentage of the chord.

For the symmetrical aero foil both mean camber line and chord line pass from the centre of gravity of the aero foil and they touch at leading and trailing edge of aerofoil.

The aerodynamic centre is the chord wise length about which the pitching moment is independent of the lift coefficient and the angle of attack.

The centre of pressure is the chord wise location about which pitching moment is zero.

NACA airfoils are airfoil shapes for aircraft wings developed by national advisory committee for aircraft wings developed by the National Advisory Committee for Aeronautics “NASA”. The shape of the NASA airfoils is described using a series of digits following the word “NACA”. The parameters in the numerical code can be entered into equations to precisely generate the cross section of the airfoil and calculate its properties.

CRUISE VELOCITY= 720 km/h=200m/s

ASPECT RATIO = 10.25

CRUISE ALTITUDE=10200 m

From the graph we have W/S= 350 kg/m2

CL=2×(W/S)/ρv2stall

VSTALL=0.25Vcruise=180 km/hr

CL= 0.17

Thus CL was found to be 0.17. With this as the primary criterion for airfoil selection, we opt for the airfoil NACA 63-210.

THICKNESS TO CHORD RATIO:

For low-speed subsonic aircraft relatively high t/c values (up to 0.2) acceptable at wing root – gives good structural depth with small profile drag penalty.From the last two digit NACA aerofoil we have a thickness to cord ratio of about 10% of the chord.

This aerofoil was selected due to its apt t/c ratio of around 10 % of the chord and has a design lift coefficient as required. The details of the aerofoil are:

The 2nd digit indicates the distance of min. pressure area i.e. 30% of the chord.

The 3rd digit indicates the design lift coefficient in tenths.CL=0.2.

The final 2 digit indicates the maximum thickness in tenths of % of the chord. In this case it is 10%.

NACA 63-2100

DRAG POLAR

ON THE SELECTED AIRFOIL

CHARACTERISTICS OF AIRFOIL

WHEN KEPT IN A WIND TUNNEL

ENGINE SELECTION

The gross weight estimate of the aircraft has been done. It is known that the weight of the power plant used is roughly 5.5% of the overall weight.The gross weight of the aircraft is given by

Wo=18,142 kg

From this it is clear that the power plant weight is approximately,

Wpower plant =0.055*18142=998kg

In most of business jets, it is common practice to have twin engine and our design has the same configuration. Hence the weight of the engine is 500kg. From the design sheet it was known that that the engines should produce approximately 98 KN of thrust ,i.e.= 49 KN ≈ 50 KN each.

The table below lists few engines that meet our demands.

COMPANY ENGINE THRUST(KN) WEIGHT SFC

Rolls Royce RB 183-2 Mk 45.8 716 0.75General Electric CF 34 3A 46.8 520 0.42General Electric CF 34-3B 44.2 467 0.45Pratt & Whitney PW810 48 512 0.62

Rolls Royce AE3007A1P 54 478 0.59

From the above chart we select for Rolls Royce AE3007 A1P.

Rolls Royce AE 3007

Fuel weight validation

The choice of a suitable engine, having been made, it is now possible to estimate the amount of fuel required for a flight at the given cruising speed for the given range.

Wfuel = (no. of engines) x (thrust at altitude) x Range x SFC x 1.2 Cruise velocity

The factor of 1.2 is provided for reserve fuel.

Thrust at altitude is calculated using the relation

Tσ is the thrust at the altitude and To is the thrust at sea level.

2.10 * TT

0 alt

ρalt=density of air at 10000m and ρo is the density of air at sea level.

Atmospheric Chart

Altitude(feet)

Pressure(in. Hg)

Temp.(F°)

Density(%)

sea level 29.92 59.0 100 2,000 27.82 51.9 94.3 4,000 25.84 44.7 88.8 6,000 23.98 37.6 83.6 8,000 22.22 30.5 78.6 10,000 20.57 23.3 73.8 12,000 19.02 16.2 69.3 14,000 17.57 9.1 65.0 16,000 16.21 1.9 60.9

Therefore from the above chart, at 10,000 m the density decreases by 26.2%.Therefore the density becomes 0.85 kg/m3 .

σ=0.85/1.225=0.75

Tσ=98*0.751.2=34.69

Wfuel = (no. of engines) x (thrust at altitude) x Range x SFC x 1.2 Cruise velocity

Wfuel = 2* 34.69*1000* 6000*0.8*1.2 720

Wfuel=5600kg

WING DESIGN

The design of wing involves building a wing with apt shape and the structure to improve aerodynamic efficiency. The wing is basically extrusion of the 2D airfoil with a sweep. For a subsonic aircraft it is best to have the root airfoil 20-60% thicker than the tip airfoil as it reduces weight and gives more volume for fuel and landing gear.

Wing area(S)=(weight/wing loading)

=20100/350

S =57.42m2

A.R=b2/s =10.25(from design data sheet)

From this we find the wing span(b)=24.26m

Croot =(2S/b(1+λ))

Taper ratio(λ)=Ctip/Croot=0.3 for jet aircrafts

Croot = 3.7m

Ctip = 1.11m`

DETERMAINING THE MAC:

Cmean= [0.66*Croot*(1+λ+λ2/(1+λ))]

Cmean= 2.61m

Distance of mean chord from aircraft centre line

= [b*(1+2λ)/6(1+λ)] = 4.97m

Wing sweep has stability. A swept wing has natural dihedral effect.

Sweep back angle at leading edge[ʌLE] =300

LOCATION OF THE WING:

As in the case of business jets the wing is of low wing configuration. The primary advantage is that it helps in stowage of landing gear.

HIGH LIFT DEVICES:

Flaps are high devices which, in effect ,increase the camber of the wing and, in some cases, as with the fowler flap, also increase the effective wing area. Their use permits better takeoff performance and permits steeper approach angles and lower approach and landing speeds.

When deflected, flaps increase the upper chamber of the wing, the negative pressure on the wing. At the same time, they allow a build up of pressure below the wing.

During takeoff, flap settings of 100 to 200 are used to better take off performance and a better angle of climb, especially valuable when climbing out over obstacles. Flaps do indeed increase drag. The greater the flap deflection, the greater the drag. At a point of about half of their

full travel, the increase drag surpasses the increase lift and flaps become airbrakes. Most flaps can be extended to 400 from the chord of the wing. At settings between 200 to 400,the essential function of the flaps is to increase the landing capabilities, by steepening the glide without increasing the glide speed. In an approach over obstacles, the use of flaps permits the pilot to touch down much nearer the threshold of the runway. Flaps also permit a slow landing speed and act as airbrakes when the airplane is rolling to a stop after landing, thus reducing need for excessive breaking action. As a result, there is less wear on the under carriage, wheels and tyres. Lower landing speeds also reduce the possibility of ground looping during the landing roll

Slotted flaps, on the other hand, including such types as fowler and a zap produces lift in excess of drag and their partial use is therefore recommended for takeoff.

From the stand point of aerodynamic efficiency, the fowler flap is generally considered to offer the most advantages and the fewest disadvantages, especially on larger airplanes, while double slotted flaps have won wide approval for smaller types.

Changes in flap setting affect the trim of an airplane. As flaps are lowered the centre of pressure moves rearward creating a nose down pitching moment.

However, in some airplanes ,the change in air flow over the tail plane as flaps are lowered, is such that the total moment created is nose up and it becomes necessary to trim the airplane nose down.

The airplane is apt loose considerable height when the flaps are raised. At the low altitudes, therefore the flaps should be raised cautiously.

Fig. Location of flaps and ailerons in the wing

FUSELAGE CABIN LAYOUTThe cabin layout of the business jet aircraft is designed for comfort rather than seating capacity. The cabin may also have a separate section for the meeting and the refreshment centre.

The design parameters of the cabin layout are:

Seat pitch =1.02m

Seat width =0.74m

Aisle width =0.6m

Cabin diameter =2.16m

By standards described by the Raymer, the structural thickness is given by: t=0.002dfus+0.0254m=0.0686m

The external diameter is thus =2.1+(0.0686*2)=2.2972m

The total length of the fuselage = 23 m.

As per raymer suggestion the two crew aircraft should have a minimum length of 100 inches nose. So we choose the length of the nose to be 20% of the total length. i.e. 0.2(23)= 4.5m.

The length of the tail is taken as about 25% of the length which is equal to 5.6m. The main reason for choosing such a high value for the tail is explained below.

Having a highly inclined tail may increase the angle of attack of the wing during the takeoff in or near to the ground. The high sweep angle of the tail may raise the rear fuselage level giving enough angle to the wing.

DETERMINATION OF CENTRE OF GRAVITY

Considering forces to act at the centre of various sections and taking moments about the nose C.G can be determined.

Hence,

C.G. = (160*2.25)+(1240*10.95)+(85.75*10.9)+(1105*18)160+1240+85.75+1105

C.G.=13.41m

LANDING GEAR DESIGN

The undercarriage or the landing gear is the structure(usually wheels)that supports an aircrafton the ground and allows an aircraft on the ground and allow it to taxi.

Types of landing gear:

Wheeled under carriage comes in two types:

Conventional or tail drageer under carriage ,where there are two main wheels towards the front of the aircraft and a single ,much smaller ,wheel or skid at the rear.

Tricycle under carriage ,where there are two main wheels under the wings and a third smaller wheel in the nose. The tail dragger arrangement was common during the early propeller era,as it allows more clearence for the propeller clearence .

Most of the modern aircraft have tricycle arrangement.tail draggers are considered harder to land and take off(because the arrangement is unstable a small deviation from the straight line travel is naturally amplified by the greter drag on the main wheel which has moved further away from the plane’s centre of gravity due to the deviation), and usually require special pilot training.

Sometimes a small tail wheel or skid is added to the aircraft with tricycle undercrriage,in case of tail strikes during the take off. The conocorde,for instance,had arectractable tail bumper wheel. The Boeing 727 also had a rectractable tail bumper. Some aircraft with rectractable conventional landing gear having a fixed tail wheel,which generate minimum drag( since most of the airflow past the tail wheel has been blanketed by the fuselage) and even improve yaw stability in some cases.

Rectractable landing gear:To decrease drag during the flight some under carriages retract into the wings and /or concealed behind doors;this is called rectractable landing gears.

A design for retractable landing gear was first seen in 1876 in plans for an amphibious monoplane designed by Frenchmen Alphonse and Paul Gauchot. Aircraft with at least partially retractable landing gear did not appear until 1917, and it was not until the 1920s and early 1930s that such aircraft became common.

By then, aircraft performance was improved to the point where the aerodynamic advantage of a retractable carriage justified the added complexity and weight. An aerodynamic approach of reducing the aerodynamic penalty imposed by fixed undercarriage is to attach aerodynamic fairings (often called spats ) on the undercarriage, with only the bottom of the wheels exposed.

Pilots conforming that their landing gear is down refer to “three green” or “three in the green”, a reference to the electrical indicator lights from the nose wheel and two main landing gears.

Amber light indicates that the gears are in the up-locked position; red lights indicate that the landing gear is in transit (neither down or locked nor fully retractable).

Landing gear placement:

The sizes of the tyres depend on the load distribution between the main wheels and the nose wheel. The load carried by each tyre is the equal and opposite reaction force exerted by the ground on them. FN and FM

are the force acting on the nose and the main landing gear respectively.

The takeoff weight acts through the centre of gravity. The distance between the line of action of C.G. and the FN is given by X1.

The distance between the line of action of C.G. and the FM is given by X2.The distance between FN and FM is 9.5m.

Force diagram for obtaining load distribution among tyres.

Taking the moments at the respective points, we get

FM= WoX1/X3

FN= WoX2/X3

We should note that FM is the sum of the reaction forces at the two main landing gears. The takeoff weight is 20,000kg.

FM= 20000*(7.9/9.5)=16631.5kg

FN= 20000*(1.6/9.5)=3368.42kg

Hence the load on the nose landing gear is 3368.42 kg.And the load on the main landing gear is 8315.75 kg each.

RAYMER’s relations derived empirically for the diameter and width of tire are based on the load on the wheel.

Wheel diameter or width of the wheel= AWB

A BWheel diameter(in) 1.51 0.349

Wheel width(in) 0.715 0.312

Main wheels:

Wheel dia. =A(FM/2)B Wheel width= A(FM/2)B

=1.51(8315.75/2)0.349 =0.715(8315.75/2)0.312

=27.66in=0.702m =9.62in=0.22m

Nose Wheels:

Wheel dia. =A(FN/2)B Wheel width= A(FM/2)B

=1.51(3368.42/2)0.349 =0.715(3368.42/2)0.312

=20.18in=0.51m =7.26in=0.18m

BRAKING SYSTEM AND TYRES:

The bias ply tyre,shown in the first picture consists of casing plies running diagonally at approximate right angles to one another. The number of plies and the angles at which they are laid dictate strength and load capacity . the latest high performance bias ply aircraft tyres inter thread reinforcing fabric(ITF). This provide additional high speed stability reduces thread distortion

Under load,protects the casing plies from damage and can act as wear indicators or rethredable tyres. Bias tyres are currently the most popular tyres fitted to the world’s fleet.

The braking system used has an anti locking braking system,which stop the wheels from locking when fully applied ,allowing greater deceleration and control during breaking,particularly in wet conditions.the brakes,developed by Dunlop,were carbon based and could bring to a halt within a mile, and with proper cooling time to dissipate heat later.

PERFORMANCE ANALYSIS

TAKE OFF PERFORMANCE:

Total take off distance=Sg+Sa

Ground roll distance= 1.44*w^2{g*ρ*s*CLMAX(T-(D+µ(W-L)))0.7VLO}

Lift of velocity(VLO) =1.2*VSTALL

=220 km/h

0.7 VLO=50.17m/s

The factor due to ground effect is given by,

Φ=16[h/b2]2/1+[16h/b]2

h=1.92m

b=1.85m

Φ=0.04

For the dry surface,

µ=0.04

D=3069.7N

Sg=1.38km is the min. distance covered by the aircraft in the ground before takeoff.

CALCULATION OF DISTANCE WHILE AIBORNE TO CLEAR AN OBSTACLE

Sa=R*sinΦOB

R=V2stall*6.96/g g-acceleration due to gravity

= 25,638.3m

Obstacle height is given by,

hOB=10.668m(for jet)

cosΦOB=(1-( hOB/R))

Flight path angle,

ΦOB=cos-1(1--( hOB/R))

=1.74470

Airborne distance,

Sa=R*sinΦOB=25,638.3*sin1.74470

=780.58m

Total take off ditance=Sa+Sg

=1.38+0.780

=2.2 km

Therefore the aircraft requires a runway of 2.2 km to takeoff.

CLIMB:

R/C=Vstall(T-D)/W

=20.833m/s

R/C= Vstall sinΦ

Φ= sin-1(R/C/ Vstall)

Φ= 6.640

3D VIEW DIAGRAMS OF THE AIRCRAFT

TOP VIEW

SIDE VIEW

FRONT VIEW

BIBLIOGRAPHY

1.AIRCRAFT DESIGN: A CONCEPTUAL APPROACH, DANIEL P. RAYMER, AIAA EDUCATION SERIES.

2.INTRODUCTION TO FLIGHT.J.D. ANDERSON

3.DESIGN OF AEROPLANE BY DARROL STINTON

4.www.aerospaceweb.org

5.www.wikipedia.org

6.www.airliners.net