KARUNYA NAGAR, COIMBATORE – 641 114

Department of Aerospace EngineeringSCHOOL OF MECHANICAL SCIENCES

2012 - 2013

09AE217- AIRCRAFT DESIGN PROJECT LAB

REGISTER NO: 09FL037

It is hereby certified that this is a bonafide report of work on

Aircraft Design Project done by Mr. Jackson J Panancherry of IV

B.Tech (Aerospace) during the period July - November 2012 and

submitted for university practical examination held on 07-11-2012.

STAFF-IN-CHARGE DIRECTOR / HOD(AERO)

INTERNAL EXAMINER EXTERNAL EXAMINER

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ACKNOWLEDGEMENT

I would like to extend my heartfelt thanks to Dr. Pradeep Kumar (Head of Aerospace

Department), for giving me his able support and encouragement. I want to highlight the fact

that my project would not have been possible without the highly informative and valuable

guidance given by Mr. Jeevanandam, whose vast knowledge and experience has made me go

about this project with great ease. I have great pleasure in expressing my sincere and whole-

hearted gratitude to him. I would also like to express the members of my team: Daniel

Jebakumar, Jackson J Pananchery and Benson M Varughese for their support and

involvement in completing the project successfully. Above all, I would like to thank God

Almighty for His ever-present Grace and Mercy.

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INDEX

SL. NO CONTENT PAGE NO.

1. Introduction to Aircraft Design 4

2. Aim of the Project 6

3. Main features of the Aircraft 7

4. Comparative study 10

5. Weight Estimation 15

6. Power plant selection 18

7. Theoretical specification 19

8. Fuselage preliminary design 20

9. Survey and incorporations 21

10. Wing design 23

11. Aircraft controls 25

12. Individual component weight estimation 34

13. Performance calculations 34

14. Three view diagram 35

15. Seating layout 39

16. Weight balance and control 41

17. Determining the CG 46

18. V-n diagram 47

19. Aerodynamic force calculations 50

20. Conclusion 55

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INTRODUCTION TO AIRCRAFT DESIGN

Aircraft design is the intellectual engineering process of creating on a paper a flying

machine to meet the specification and requirements established by potential user or as

perceived by manufacturer an pioneer innovative, new ideas and technology. From the time

that an airplane first materialize as a new thought in the mind of one or more persons to the

time that the finished products roll out of the manufacturers door, the design process has gone

through three distinct phases that carried out in sequence. These phases are in chronological

order,

Conceptual design

Preliminary design

Detail design

The Design process

Conceptual Design

The design process starts with a set of specification for new planes, or much less

frequently as the response to desire to implement some pioneering innovative new ideas and

technology. In either case, there is rather concrete goal toward which the designers are aiming.

The first step toward achieving that goal constitutes the conceptual design phase. Here with

certain assumptions the overall shape, size, weight and performance of the new design

determined. The product of conceptual design phase is a layout of the airplane configuration.

However, the conceptual design phase determines such fundamental aspects such as the

Shape of the wings

Location of the wings relative to the fuselage

Shape and location of horizontal and vertical stabilizers

Whether to use canard surface or not

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Engine size and placement

The major drivers during design process are aerodynamics, propulsion and flight performance.

Preliminary Design

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

layout. It is in the preliminary design that serious structural and control system analysis and

design take place. During this phase also substantial wind tunnel testing will be carried out and

major computational fluid dynamic calculations of the complete flow field over the airplane

configuration will be made. It is possible that the wind tunnel tests or CFD calculation will

uncover aerodynamic interference or some unexpected stability problems. At the end of the

preliminary design phase, the airplane configuration is precisely defined. The drawing process

called lofting is carried out which mathematically models the precise shape of the outside skin

of the airplane making certain that all sections of the aircraft properly fit together. The end of

the preliminary design phase brings a major decision to commit to the manufacturer of the

airplane or not. The importance of this decision point for modern aircraft manufacturers cannot

be understand, considering the tremendous costs involved in the design and manufacture of

new airplane.

Detail Design

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

aerodynamic, propulsion, structure, performance and flight control analyses have all been

finished with the preliminary design phase. For detail design, the airplane is now simply a

machine to be fabricated. The precise design of each individual rib, spar, and selection of the

skin now takes place. The size, number and location of fasteners are determined. The

manufacturing tools and jigs are designed. At the end of this, aircraft is ready to be fabricated.

OBJECTIVE OF THE PROJECT

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To design a 300 seater, twin engine sonic speed, high performance, passenger aircraft with the

following features:

To attain a max speed close to Mach 1,

Control surfaces to aid improved maneuverability,

Space conserving fuselage.

How we started

Proposals of Designing a Fighter Aircraft, a UAV, a Cargo Aircraft and a Passenger

Aircraft were put up.

Having studied the pros and cons of designing the above mentioned aircrafts at our

level, we arrived at a decision to go with a passenger aircraft for the following reasons.

Why Passenger Aircraft?

Conceived With A Novel Idea To Serve People At All Times;

To Introduce Low-cost Air Transport;

To Usher In A New Generation Of Passenger Aircrafts;

To Introduce Economically Viable Aircraft Manufacturing;

A More Business Oriented Approach

As Beginners, To Understand, Learn and Equip Ourselves In Designing More-complex

Aircrafts Like An Unstable Fighter Aircraft.

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MAIN FEATURES OF THE AIRCRAFT

1) Swept Back Wings:-

Swept back wings contribute to more lateral stability.

Swept back wings produce less lift, so in turbulent weather they are less susceptible to

abrupt changes.

They are designed with low thickness and high fineness ratio, hence less form drag.

Generally they are tapered, so less induced drag.

Capable of flying at high Mach no. as actual relative wind speed is at an angle to the

wing leading surface and therefore the wind component perpendicular to the wing

leading edge is less, and hence the wing senses less speed than actual.

Swept-Back Wings

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2) Canards:

Possibility for very good stalling characteristics without elevator stops.

Sometimes a desirable layout from the packaging standpoint: Main wing carry-through

behind cabin, pusher engine installation simplified.

Synergistic use of winglets for directional stability.

In certain cases a simplified control linkage is possible.

When wing flaps are not desired (for simplicity as in ultra-lights, or competition rules as

with standard class sailplanes for example) the CLmax of a canard may exceed that of an aft-

tail airplane.

For unstable aircraft, canard designs may have a CLmax and/or drag advantage.

Control authority is larger for unstable canard aircraft at high CL than for unstable aft-tail

designs.

Canards in Valkyre

3.) V-Tail

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The V-tail is lighter, has less wetted surface area, and thus produces less drag.

In modern day, light jet general aviation aircraft such as the Cirrus Jet, Eclipse 400 or

the unmanned aerial drone Global Hawk often have the power plant placed outside the

aircraft to protect the passengers and make certification easier. In such cases V-tails are

used to avoid placing the vertical stabilizer in the exhaust of the engine, which would

disrupt the flow of the exhaust, reducing thrust and increasing wear on the stabilizer,

possibly leading to damage over time.

V-Tail in F-22 Raptor

COMPARATIVE STUDY

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Once decisions have been made on the configuration(s) to be further considered it is

necessary to size the aircraft. A three-view general arrangement scale drawing for each aircraft

configuration will be required. Little detail will be known at this stage about the aircraft

parameters (wing size, engine thrust, and aircraft weight) so some crude estimates have to be

made. This is where data from previous/existing aircraft designs will be useful. Although the

new design will be different from previous aircraft, such inconsistencies can be ignored at this

stage. Representative values from one or a small group of the specimen aircraft for wing

loading, thrust loading and aircraft take-off weight will be used. It is also possible to use a

representative wing shape and associated tail sizes. The design method that follows is an

iterative process that usually converges on a feasible configuration quickly. The initial general

arrangement drawing, produced to match existing aircraft parameters, provides the starting

point for this process. Even though designs maybe relatively crude at this stage it is important

to draw it to scale making approximations for the relative longitudinal position of the wing and

fuselage and the location of tail surfaces and landing gear.

A spreadsheet is the best way of recording numerical values for common parameters (e.g. wing

area, installed thrust, aircraft weights (or masses), etc.). A database is a good way to record other

textural data on the aircraft (e.g. when first designed and flown, how many sold and to whom, etc.). The

geometrical and technical data can be used to obtain derived parameters (e.g. wing loading, thrust to

weight ratio, empty weight fraction, etc.). Such data will be used to assist subsequent technical design

work. It is possible, using the graph plotting facilities of modern spreadsheet programs, to plot such

parameters for use in the initial sizing of the aircraft. For instance, a graph showing wing loading

against thrust loading for all our aircrafts will be useful in selecting specimen aircraft to be used in

comparison with our design.

A deep study was made of various aircrafts belonging to the same range of passenger

capacity.

Extensive data collection performed.

Comparison of various parameters was made.

Mean Of Important Parameters Was Made

A Total Of 17 Aircrafts Were Taken And Studied

19 Parameters Were Obtained, Some Of Which Were Results Of Simple Arithmetic

Operation Of Already Existing Parameter

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Contribution of Different Aircrafts:-

Aircrafts Contribution to T/W Calculation:-

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Aircrafts Contribution to F/W Ratio:-

Aircrafts Contribution to E/W Ratio:-

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Weight Distribution:-

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WEIGHT ESTIMATION

Most aircraft layouts start with the drawing of the fuselage. For many designs the

geometry of the fuselage can be easily proportioned as it houses the payload and cockpit/flight

deck. These parameters are normally specified in the project brief. They can be sized using

design data from other aircraft. The non-fuselage components (e.g. wing, tail, engines and

landing gear) are added as appropriate. With a reasonable first guess at the aircraft

configuration, the aircraft can be sized by making an initial estimate of the aircraft mass. Once

this is completed it is possible to more accurately define the aircraft shape by using the

predicted mass to fix the wing area and engine size.

Initial mass (weight) estimation

The first step is to make a more accurate prediction of the aircraft maximum (take-off )

mass/weight.

Aircraft design textbooks show that the aircraft take-off mass can be found from:

MTO = MUL

1 − (ME/MTO) − (MF/MTO)

Where,

MTO = maximum take-off mass

MUL = mass of useful load (i.e. payload, crew and operational items)

ME = Empty mass

MF = fuel mass

If aircraft operational mass is used for ME, the crew and operational items in MUL would not

be included. One of the main difficulties in the analysis at this stage is the variability of

definitions used for mass components in published data on existing aircraft. Some

manufacturers will regard the crew as part of the useful load but others will include none or

just the minimum flight crew in their definition of empty/operational mass. Such difficulties

will be only transitional in the development of your design, as the next stage requires a more

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detailed breakdown of the mass items. The three unknowns on the right-hand side of the

equation can be considered separately:-

(a) Useful load

The mass components that contribute to MUL are usually specified in the project brief and

aircraft requirement reports/statements.

(b) Empty mass ratio

The aircraft empty mass ratio (ME/MTO) will vary for different types of aircraft and for

different operational profiles. All that can be done to predict this value at the initial sizing stage

is to assume a value that is typical of the aircraft and type of operation under consideration.

The data from existing/competitor aircraft collected earlier is a good source for making this

prediction. Figure below shows how the data might be viewed.

(c) Fuel fraction

For most aircraft the fuel fraction (MF/MTO) can be crudely estimated from the fact that fuel

weight is 30 % of Maximum Take-off Weight.

Initial Weight estimation:-

Empty mass fraction = 0.56278

Fuel mass fraction = 0.424486

Per passenger =120 kg

Total passengers = 300

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Total no. Of pilots = (2+1)

Pilot weight =120 kg

No. Of cab attendants = (16 + 4)

Weight of each Attendant = 100 kg

Therefore,

Total weight (m) = (135 * 300) + (3 * 120) + (20 * 100)

= 94490 kg

MTOW = m = 94490 = 688603lbs

1- M e - M f 1 - 0.3 – 0.56278

M to M to

Thrust calculation:-

Thrust = (T/Mto)avg x MTOW

= 0.31 X 688603

= 21415lbs

Wing Area:-

S = W/ (W/S)avg

= 688603/134.27051

S =5128.475 sq.ft

Wing span:-

b = (AR x S)0.5

= (10 x 5128.475)0.5

b = 226.461 ft

Mean Chord:-

MAC = (b/AR)

=226.461/10

MAC=22.461 ft

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POWER PLANT SELECTION

Thrust required for our aircraft = 214155 lb.

Thrust by 2GE 90 115b engine = (2 x 115,300 lb.)

= 230600 lb.

Specifications:-

General characteristicsType :  axial flow, twin-shaft, bypass turbofan engineLength : 287 in (7,290 mm)Diameter : overall: 135 in (3.429 m); fan: 128 in (3.251 m)Dry weight : 18,260 lb. (8,283 kg).

ComponentsCompressor:  axial- 1 wide chord swept fan, 4 low pressure stages, 9 high pressure stagesTurbine : axial- 6 low pressure stages, 2 high pressure stages.

PerformanceMaximum thrust : max at sea level: 115,300 lb. (514 kN).Overall pressure ratio: 42:1Thrust-to-weight ratio: approx. 6.3:1

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THEORETICAL SPECIFICATION:-

Maximum Take-off Weight 688603 lbs.

Thrust 214155lbs.

Power-plant used GE90 115B

Wing Area 5128.475ft.sq

Aspect Ratio 10

Wing Span 226.461ft

Mean Chord 22.6461ft.

Proposed Maximum speed Mach>1

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FUSELAGE PRELIMINARY DESIGN

Inspired from the shape of diamond, the hardest substance, the fuselage is designed

with a view to economically and efficiently utilize the available space. Designed with a view to

ensure space is effectively used, minimizing wastage. Fuselage is triangular in shape where a

platform is used to separate fuel storage from payload storage. The triangular shape prevents

the wastage of unoccupied spaces present in conventional aircrafts. This ensures the efficient

utilization of space, leading to material reduction promoting weight reduction and ultimately

lowering Drag

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SURVEY AND INCORPORATIONS

The following aircrafts were referred to and design features were adopted from to develop and

design our aircrafts

Wing from B2 Spirit

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V-Tail from F22 Raptor

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WING DESIGN

If the wings of large birds like the Steppe Eagle were too long, their turning circle

would be too big to fit inside the rising columns of warm air which they use to soar. The

eagle’s wings perfectly balance maximum lift with minimum length by curling feathers up at

the tips until they are almost vertical. This provides a barrier against the vortex for highly

efficient flight. If built to a conventional design, the A380’s wingspan would have been three

metres too long for the world’s airports. But thanks to small devices known as ‘winglets,’

which mimic the upward curl of the eagle’s feathers, the A380’s wings are in compliance with

airport limits by 20 cm. but still provide enough lift for the world’s largest passenger aircraft to

fly efficiently – saving fuel, lowering emissions and reducing airport congestion.

Bio-mimicry: Our wing was drew its inspiration from the Eagle

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Winglets help reduce vortices in Aircrafts

The Wing for the aircraft is adopted from B2- Spirit for a simple reason that no other swept-

back aircraft had wings with control surfaces sharing the same design concepts.

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Wing Design

NACA 653-019 at the root

NACA 653-018 at the tip

About The Airfoil:-

It belongs to the 6 Series of Airfoils.

It has the area of minimum pressure 50% of the chord back.

It maintains low drag, 0.3 above and below the lift coefficient.

It has a minimum thickness of 19% and 18% at the root and tip respectively.

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AIRCRAFT CONTROLS

PRIMARY CONTROLS SECONDARY CONTROLS

Elevons Canards

V-tail Slats

Drag split rudders

CANARDS:-

In aeronautics, canard (French for "duck") is a wing configuration of fixed-wing aircraft in

which the forward horizontal surface is smaller than the rearward one, the former sometimes

being known as the "canard" or fore- plane, while the latter is the main wing.

In contrast a conventional aircraft has a small horizontal surface or tailplane behind the main

wing.

General characteristics:-

A canard design tends to be less controllable than a conventional design because ailerons on

the main wing may be subject to turbulence from the canards that varies widely at different

angle of attack , leading to conditions of deep stall. "Incorporating roll control on the canard is

basically less efficient because of an adverse downwash influence on the main wing opposing

the canard rolling-moment input."Canards have poor stealth characteristics because they

present large, angular surfaces that tend to reflect radar signals.

Canard classes:-

Other classes include the close-coupled type and active vibration damping. Canard designs fall

into two main classes: the lifting-canard and the control- canard.

Rutan Long-Ez, with lifting-canard ahead of the cockpit

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Lifting Canard

In this configuration, the weight of the aircraft is shared between the main wing and the

canard wing. It may be described as an extreme conventional configuration with the following

features:

A small highly-loaded wing

The C.G significantly aft, at 500% chord

A relatively large lifting tail to give enough stability with this extreme aft c.g.

A lifting-canard generates an upload, in contrast to a conventional aft-tail which

generates negative lift that must be counteracted by extra lift on the main wing. This may

appear to unambiguously favour the canard. However, the downwash interaction between the

two surfaces is unfavourable for the canard, and favourable for the conventional tail, so the

difference in overall induced drag is actually not obvious, and depends on the details of the

configuration. The canard lift appears to increase the overall lift capability of the configuration.

However, pitch stability flight safety requirements dictate that the canard must stall before the

main wing, so the main wing can never reach its maximum lift capability. Hence, the main

wing must then be larger than on the conventional configuration, which increases its weight

and profile drag.

Pitch stability requires that the lift slope of the canard wing is lower than the lift slope

of the main wing: to achieve stability, the change in lift coefficient with angle of attack should

be less than that for the main plane. The first powered airplane to fly, the Wright Flyer, a

lifting-canard, was pitch unstable. Following the first flight, the Wright Flyers had some ballast

added to the nose. The most common way in which pitch stability can be achieved is to

increase the wing loading of the canard. This tends to increase the lift induced drag of the

foreplane, which may be given a high aspect ratio in order to limit drag. A canard airfoil has

commonly a greater airfoil camber than the wing.

With a lifting-canard, the main wing must be located further aft of the center of gravity

than with a conventional aft tail, and this increases the nose-pitching moment caused by the

deflection of trailing-edge flaps. Highly loaded canards do not have sufficient extra lift to

balance this moment, so lifting-canard aircraft cannot readily be designed with powerful

trailing-edge flaps. NASA has investigated the use of a stowable canard for use at low speed

that is withdrawn from the airstream at high speeds in order to avoid the Wave drag penalty of

a canard design.

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Control-Canard

A deflected control-canard on an RAF Typhoon F2

In the later control-canard, most of the weight of the aircraft is carried by the main wing

and the canard wing is used primarily for longitudinal control during manoeuvrings. Thus, a

control-canard mostly operates only as a control surface and is usually at zero angle of attack,

carrying no aircraft weight in normal flight. One benefit obtainable from a control-canard is

avoidance of pitch-up. An all-moving canard capable of a significant nose-down deflection will

protect against pitch-up. As a result, the aspect ratio and wing-sweep of the main wing can be

optimized without having to guard against pitch-up.

Close-Coupled Canard:

Saab 37 Viggen of the Swedish Air Force

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In the close-coupled canard, the fore-plane is located just above and forward of the

main wing. At high angles of attack the canard surface directs airflow downwards over the

wing, reducing turbulence which results in reduced drag and increased lift. The canard

foreplane may be fixed as on the IAI Kfir, or have landing flaps as on the Saab Viggen, or it

may be moveable and also act as a control-canard during normal flight as on the Dassault

Rafale. A close-coupled canard is very useful for a supersonic delta wing design which gains

lift in both transonic flight (such as for super cruise) and also in low speed flight (such as take

offs and landings). A moustache is a small, high aspect ratio fore-plane of close-coupled

configuration. The surface is typically retractable at high speed and is deployed only for low-

speed flight.

Advantages:

Possibility for very good stalling characteristics without elevator stops.

Sometimes a desirable layout from the packaging standpoint: Main wing carry-through behind

cabin, pusher engine installation simplified.

Synergistic use of winglets for directional stability.

In certain cases a simplified control linkage is possible.

When wing flaps are not desired (for simplicity as in ultra-lights, or competition rules as with

standard class sailplanes for example) the CLmax of a canard may exceed that of an aft-tail

airplane.

For unstable aircraft, canard designs may have a CLmax and/or drag advantage.

Control authority is larger for unstable canard aircraft at high CL than for unstable aft-tail

designs.

Disadvantages:

Fuel center of gravity lies farther behind aircraft c.g. than in conventional designs. This means

that a large c.g. range is produced or that the fuel must be held elsewhere (e.g. strakes near the

wing root.)

CLmax problems with flaps or margin on the entire wing: Flaps produce a larger pitching

moment about the c.g. on a canard aircraft. This results in the need for both large canard

aerodynamic incidence change and high maximum canard lift coefficient. Note that since the

value of a S is usually larger for canard designs, Cm0 has a greater impact on L than it does on

aft-swept designs.

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Induced drag / CLmax incompatibility: Canard designs can achieve equal or better CLmax values

than conventional designs, and similar values of span efficiency. However, the configurations

with high CLmax values have terrible values of e and those with respectable e’s have low

maximum lift coefficients.

Directional stability: The distance from the aircraft c.g. to the most aft part of the airplane is

usually smaller on canard aircraft. This poses a problem for locating a vertical stabilizer and

may result in very large vertical surfaces. (Note, however, that winglets may be used to

advantage in this case.)

Wing twist distribution is strange and CL dependent: The wing additional load distribution is

distorted by the canard wake.

Power effects on canard - deep stall: Accidents have been associated with tractor canard

configurations for which the propeller slipstream has prevented canard stall before wing stall.

The result is a possible deep-stall problem.

Finally, and perhaps most importantly, canard sizing is much more critical than aft tail sizing.

By choosing a canard which is somewhat too big or too small the aircraft performance can be

severely affected. It is easy to make a very bad canard design.

In a normal configured aircraft when the airplane stalls it is the main wing that stalls and

the nose drops until sufficient speed can be regained for the elevator to once more be

effective and the wing have sufficient lift to return to level flight. In a canard aircraft, it

is the front wing that stalls, and the main wing keeps flying. The nose only drops a few

degrees before normal flight is achieved. You can hold the stick all the way back in a

canard and it will do a series of ups and down until you get sea sick, but the plane never

enters into a deep stall, it recovers itself.

Yes, a canard is a far safer aircraft to fly due to this self-recovery action of the canard.

The downside is that most canards have a very long take off run before they can rotate

and the approach speed is much higher than in a normal configured aircraft. Many

canards do not have flaps as a safety measure due to the some poor flight characteristics

when they are deployed. Most of the small experimental canards, which were very

popular several decades ago, had their vertical stabilizers and rudders located on the

wing tips, elevator action was built into the canard wing in front

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The Piaggio 180 was one of several canards produced for the business pilot. The

Beechcraft Starship was another. The P-180 only used a small nose mounted

Stabilizing canard.

This Grumman canard was a research project testing the swept forward wing

configuration. Note the small canard auxiliary control surface on the nose.

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ELEVONS

Planes are built with a series of complex mechanisms that allow the pilot to maintain

control during flight. Every plane found in the air today requires specific controls in order to

maintain flight and properly navigate the crystal blue skies. Even from the very first examples

of plane engineering there has been a need for elevators and stabilizers to control a plane

during flight. Now with the coming age of technology, where jets crack the sound barrier as

though it was just a typical joy ride, there is a new breed of aircraft controls that have emerged

known as the "ELEVON"!

Elevons are specific surfaces that utilize airflow to allow directional changes such as altitude

control (pitch), lateral axis control (roll), and vertical axis control (Yaw).

They make use of a combination of both traditional stabilizer and aileron flap controls, which

have been in use since the dawn of flight.

Almost all planes have a fixed vertical and horizontal line of axis and a common type of tail

section. These plane types utilize such mechanisms like the elevators (Pitch control) found on

the rear stabilizer wing, ailerons (Roll control), which are flaps found on both trailing wing

edges, and a rear rudder fin that controls steering (Yaw).

There are though a few planes that fall into the Delta-winged class, which have either no tail

section at all, or the tail section has barely any functionality, save the stability factor. These

arrow or Y shaped planes require a new way to implement the control axis, which is where the

elevons come into play.

Similar to a typical aileron the Elevon is a flap that is fixed to the trailing edge of the wing,

however there are additional flaps that can control the same properties that are determined by

the more typical stabilizers in the aft section of non-delta-wing planes.

The Concord for example has a total of six elevons that control pitch and roll control. The

design of this plane type utilizes an outer elevon that is found at the outer edge tip of the wing.

The second elevon is known as the middle elevon, which is right next to the outer elevon.

Lastly there is the inner elevon that is closest to the body of the plane. Not every plane utilizes

this particular configuration since some planes are completely without a tail section.

Many fighter jets or stealth planes used in the military today utilize elevons for flight control,

such as the Nighthawk F-117, SR-17 Blackbird and the infamous alien like B-2 Spirit, which

actually utilizes elevons and special thrusters that require complex computer systems to

implement Roll and Yaw control.

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V-TAIL

In aircraft, a V-tail (sometimes called a Butterfly tail) is an unconventional

arrangement of the tail control surfaces that replaces the traditional fin and horizontal surfaces

with two surfaces set in a V-shaped configuration when viewed from the front or rear of the

aircraft. The rear of each surface is hinged, and these movable sections, sometimes

called ruddervators, combine the tasks of the elevators and rudder.

Advantages:-

Ideally, with fewer surfaces than a conventional three-aerofoil tail or a T-tail, the V-tail is

lighter, has less wetted surface area, and thus produces less drag.

In modern day, light jet general aviation aircraft such as the Cirrus Jet, Eclipse 400 or the

unmanned aerial drone Global Hawk often have the power plant placed outside the aircraft to

protect the passengers and make certification easier.

In such cases V-tails are used to avoid placing the vertical stabilizer in the exhaust of the

engine, which would disrupt the flow of the exhaust, reducing thrust and increasing wear on

the stabilizer, possibly leading to damage over time.

Disadvantages:-

Combining the pitch and yaw controls is difficult and requires a more complex control system.

The V-tail arrangement also places greater stress on the rear fuselage when pitching and

yawing.

Drag Split Rudder:-

Located in the wing tips at 18 degrees angle with the aircraft centreline. They were designed

for pilots to be used during high speed manoeuvres as air –brakes.

SECONDARY CONTROLS

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Slats:-

Aerodynamic surfaces on the leading edge of the wings which allow the wing to

operate at a higher angle of attack resulting in higher coefficient of lift. By deploying slats an

aircraft can fly at slower speeds, or take off and land in shorter distances.They are usually used

while landing or performing manoeuvres which take the aircraft close to the stall, but are

usually retracted in normal flight to minimize drag.

Spoilers:-

On some airplanes, high-drag devices called spoilers are deployed from the wings to

spoil the smooth airflow, reducing lift and increasing drag. Spoilers are used for roll control on

some aircraft, one of the advantages being the elimination of adverse yaw. To turn right, for

example, the spoiler on the right wing is raised, destroying some of the lift and creating more

drag on the right. The right wing drops, and the airplane banks and yaws to the right.

Deploying spoilers on both wings at the same time allows the aircraft to descend without

gaining speed. Spoilers are also deployed to help shorten ground roll after landing. By

destroying lift, they transfer weight to the wheels, improving braking effectiveness.

INDIVIDUAL COMPONENT WEIGHT ESTIMATION:

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Wing Structure

MTOM = 688603kg

Wing area = 476.450 sq. m

Wing aspect ratio = 10

Wing taper ratio = 0.3

Wing av. thickness = 15%

From standard results, Mwing = 10.4%MTOW

=71614.712 Kg

(Inclusive of slats, elevons and spoilers)

Tail Structure

Tail structure inclusive of V-Tail and Canards= MP = 1.9% MTOM=13083.457 kg

Fuselage Structure

11.5% MTOM= 79189.345kg

Nacelle Structure

2.1% MTOM= 14460.663kg

Landing Gear

4.45% MTOM= 30642 kg

PERFORMANCE CALCULATIONS

Range:-Range = (V/c)(L/D) loge(M1/M2)

Where,

V = cruise speed = M0.85∗ = 255m/s = 485 kts

c = assumed engine fuel consumption = 0.55 N/hr.

(L/D) assumed to be = 17 in cruise

M1 = start mass = MTOM = 688603kg

M2 = end mass = ZFM = 245142.668 kg

Range = 12727.436 km

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THREE VIEW DIAGRAM

Side-View

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Top-View

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Front-View

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Three-View Diagram

SEATING LAYOUT

Fuselage Structure:

The fuselage width is set by the number of seats abreast, the seat width and the aisle width.

The depth is set to accommodate the cargo containers below the floor and the headroom above

the aisle.

It is desirable to split the cabin into at least two separate sections.

This makes the in-flight servicing easier and allows more options for the airline to segregate

different classes.

For the exclusive executive layout, this division will allow a quieter environment within the

cabin. A service module (catering or toilets) is positioned at this location.

External service doors and hatches are positioned here and these can act as emergency exits.

The provision of service modules and the ‘wasted’ space adjacent to the doors will add about 4

meters to the cabin length.

For our 300 seater aircraft, we have 100 Executive class seats and 200 charter seats.

Seating Plan for Executive Class

Seating pattern : 1-2-1

Number of Executive class seats = 100

Typical maximum first-class seat width= 0.7m

Aisle width =0.6m

Seat pitch value: 0.88m

Number of seats in a row =4

Total number of rows =25

Total width of the cabin =(4 × 0.7) + 0.6=3.4m

Total length of the Executive class = 22m

Seating Plan for Charter Class

Seating pattern : 2-3-2

Number of Charter class seats = 200

Typical maximum Charter class seat width= 0.405m

Aisle width =0.6m

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Seat pitch value: 0.8m

Number of seats in a row =7

Total number of rows = 31

Total width of the cabin =(7 × 0.405) + 0.56=3.4m

Total length of the Charter class =24.8 m

Cabin Dimension:-

Adding 0.2m for the pressure cabin structure :

Total fuselage external diameter equal to 3.60m

Total Length of the cabin= cabin length + the non-cabin length + service modules and

the ‘wasted’ space For our aircraft= (22+24.8+15+4) = 65.8m

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WEIGHT BALANCE AND CONTROL

There are many factors that lead to efficient and safe operation of aircraft. Among these

vital factors is proper weight and balance control. The weight and balance system commonly

employed among aircraft consists of three equally important elements: the weighing of the

aircraft, the maintaining of the weight and balance records, and the proper loading of the

aircraft. An inaccuracy in any one of these elements nullifies the purpose of the whole system.

The final loading calculations will be meaningless if either the aircraft has been improperly

weighed or the records contain an error. Improper loading cuts down the efficiency of an

aircraft from the standpoint of altitude, maneuverability, rate of climb, and speed. It may even

be the cause of failure to complete the flight, or for that matter, failure to start the flight.

Because of abnormal stresses placed upon the structure of an improperly loaded aircraft, or

because of changed flying characteristics of the aircraft, loss of life and destruction of valuable

equipment may result. The responsibility for proper weight and balance control begins with the

engineers and designers, and extends to the aircraft mechanics that maintain the aircraft and the

pilots who operate them.

Modern aircraft are engineered utilizing state-of-the-art technology and materials to

achieve maximum reliability and performance for the intended category. As much care and

expertise must be exercised in operating and maintaining these efficient aircraft as was taken in

their design and manufacturing. The designers of an aircraft have set the maximum weight,

based on the amount of lift the wings or rotors can provide under the operation conditions for

which the aircraft is designed. The structural strength of the aircraft also limits the maximum

weight the aircraft can safely carry. The ideal location of the center of gravity (CG) was very

carefully determined by the designers, and the maximum deviation allowed from this specific

location has been calculated. The manufacturer provides the aircraft operator with the empty

weight of the aircraft and the location of its empty weight center of gravity (EWCG) at the time

the certified aircraft leaves the factory.

Weight Control

Weight is a major factor in airplane construction and operation, and it demands respect from all

pilots and particular diligence by all A&P mechanics and repairmen. Excessive weight reduces

the efficiency of an aircraft and the safety margin available if an emergency condition should

arise. When an aircraft is designed, it is made as light as the required structural strength will

allow, and the wings or rotors are designed to support the maximum allowable weight. When

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the weight of an aircraft is increased, the wings or rotors must produce additional lift and the

structure must support not only the additional static loads, but also the dynamic loads imposed

by flight maneuvers. Severe uncoordinated maneuvers or flight into turbulence can impose

dynamic loads on the structure great enough to cause failure. In accordance with Title 14 of the

Code of Federal Regulations (14 CFR) part 23, the structure of a normal category airplane

must be strong enough to sustain a load factor of 3.8 times its weight. That is, every pound of

weight added to an aircraft requires that the structure be strong enough to support an additional

3.8 pounds. An aircraft operated in the utility category must sustain a load factor of 4.4, and

acrobatic category aircraft must be strong enough to withstand 6.0 times their weight.

The lift produced by a wing is determined by its airfoil shape, angle of attack, speed

through the air, and the air density. When an aircraft takes off from an airport with a high

density altitude, it must accelerate to a speed faster than would be required at sea level to

produce enough lift to allow takeoff; therefore, a longer takeoff run is necessary. The distance

needed may be longer than the available runway. When operating from a high-density altitude

airport, the Pilot’s Operating Handbook (POH) or Airplane Flight Manual (AFM) must be

consulted to determine the maximum weight allowed for the aircraft under the conditions of

altitude, temperature, wind, and runway conditions.

Effects of Weight:-

Most modern aircraft are so designed that if all seatsare occupied, all baggage allowed by the

baggage compartment is carried, and all of the fuel tanks are full, the aircraft will be grossly

overloaded. This type of design requires the pilot to give great consideration to the

requirements of the trip. If maximum range is required, occupants or baggage must be left

behind, or if the maximum load must be carried, the range, dictated by the amount of fuel on

board, must be reduced. Some of the problems caused by overloading an aircraft are:

• The aircraft will need a higher takeoff speed, which results in a longer takeoff run.

• Both the rate and angle of climb will be reduced.

• The service ceiling will be lowered.

• The cruising speed will be reduced.

• The cruising range will be shortened.

• Maneuverability will be decreased.

• A longer landing roll will be required because the landing speed will be higher.

• Excessive loads will be imposed on the structure, especially the landing gear.

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Weight Changes:-

The maximum allowable weight for an aircraft is determined by design considerations.

However, the maximum operational weight may be less than the maximum allowable weight

due to such considerations as high-density altitude or high-drag field conditions caused by wet

grass or water on the runway. The maximum operational weight may also be limited by the

departure or arrival airport’s runway length. One important preflight consideration is the

distribution

of the load in the aircraft. Loading the aircraft so the gross weight is less than the maximum

allowable is not enough. This weight must be distributed to keep the CG within the limits.

If the CG is too far forward, a heavy passenger can be moved to one of the rear seats or

baggage can be shifted from a forward baggage compartment to a rear compartment. If the CG

is too far aft, passenger weight or baggage can be shifted forward. The fuel load should be

balanced laterally: the pilot should pay special attention to the POH or AFM regarding the

operation of the fuel system, in order to keep the aircraft balanced in flight.

Balance and Control:-

Balance control refers to the location of the CG of an aircraft. This is of primary importance to

aircraft stability, which determines safety in flight. The CG is the point at which the total

weight of the aircraft is assumed to be concentrated, and the CG must be located within

specific limits for safe flight. Both lateral and longitudinal balance are important, but the prime

concern is longitudinal balance; that is, the location of the CG along the longitudinal or

lengthwise axis. An airplane is designed to have stability that allows it to be trimmed so it will

maintain straight and level flight with hands off the controls. Longitudinal stability

is maintained by ensuring the CG is slightly ahead of the center of lift. This

produces a fixed nose-down force independent of the airspeed. This is

balanced by a variable nose-up force, which is produced by a downward

aerodynamic force on the horizontal tail surfaces that varies directly with

the airspeed.

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If a rising air current should cause the nose to pitch up, the airplane will

slow down and the downward force on the tail will decrease. The weight

concentrated at the CG will pull the nose back down. If the nose should

drop in flight, the airspeed will increase and the increased downward tail

load will bring the nose back up to level flight. As long as the CG is

maintained within the allowable limits for its weight, the airplane will have

adequate longitudinal stability and control. If the CG is too far aft, it will be

too near the center of lift and the airplane will be unstable, and difficult to

recover from a stall.

If the CG is too far forward, the downward tail load will have to be increased to maintain level

flight. This increased tail load has the same effect as carrying additional weight; the aircraft

will have to fly at a higher angle of attack, and drag will increase. A more serious problem

caused by the CG being too far forward is the lack of sufficient elevator authority. At slow

takeoff speeds, the elevator might not produce enough nose-up force to rotate and on landing

there may not be enough elevator force to flare the airplane. Both takeoff and landing runs will

be lengthened if the CG is too far forward.

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The basic aircraft design assumes that lateral symmetry exists. For each item of weight added

to the left of the centerline of the aircraft (also known as buttock line zero, or BL-0), there is

generally an equal weight at a corresponding location on the right. The lateral balance can be

upset by uneven fuel loading or burnoff. The position of the lateral CG is not normally

computed for an airplane, but the pilot must be aware of the adverse effects that will result

from a laterally unbalanced condition. This is corrected by using the aileron trim tab until

enough fuel has been used from the tank on the heavy side to balance the airplane. The

deflected trim tab deflects the aileron to produce additional lift on the heavy side, but it also

produces additional drag, and the airplane flies inefficiently

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DETERMINING THE CG

One of the easiest ways to understand weight and balance is to consider a board with

weights placed at various locations. We can determine the CG of the board and observe the

way the CG changes as the weights are moved. The CG of a board may be determined by using

these four steps:

1. Measure the arm of each weight in inches from the datum.

2. Multiply each arm by its weight in pounds to determine the moment in pound-inches of each

weight.

3. Determine the total of all weights and of all the moments. Disregard the weight of the board.

4. Divide the total moment by the total weight to determine the CG in inches from the datum.

ITEM WEIGHT (lbs) ARM (inches) MOMENT

Wings 157883.4141 1433.07 397831.17

Fuselage 174582.621 1295.236 226125696

Engine 36521.18 1476.378 53919067

Tail 28843.959 2362.2 68135200

Total 397831.1741 568149771.3

CG 1428.1178

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Location of the Center of Gravity

V-n DIAGRAM

Flight regime of any aircraft includes all permissible combinations of speeds, altitudes,

weights, centers of gravity, and configurations. This regime is shaped by aerodynamics,

propulsion, structure, and dynamics of aircraft. The borders of this flight regime are called

flight envelope or maneuvering envelope. The safety of human onboard is guaranteed by

aircraft designer and manufacturer. Pilots are always trained and warned through flight

instruction manual not to fly out of flight envelope, since the aircraft is not stable, or not

controllable or not structurally strong enough outside the boundaries of flight envelope. A

mishap or crash is expected, if an aircraft is flown outside flight envelope.

The flight envelope has various types; each of which is usually the allowable variations

of one flight parameter versus another parameter. These envelopes are calculated and plotted

by flight mechanics engineers and employed by pilots and flight crews. For instance, the load

masters of a cargo aircraft must pay extra caution to the center of gravity location whenever

they distribute various loads on the aircraft. There are several crashes and mishaps that safety

board's report indicated that load master are responsible, since they deployed more loads than

allowed, or misplaced the load before take-off. Nose heavy and tail heavy are two flight

concepts that pilots are familiar and experienced with, and are trained to deal with them safely.

Load Factor

The load to the aircraft on the ground is naturally produced by the gravity (i.e. 1 times g).

But, there are other sources of load to the aircraft during flight; one of which is the acceleration

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load. This load is usually normalized through load factor (i.e. "n" times g). In another word, aircraft

load is expressed as a multiple of the standard acceleration due to gravity (g = 9.81 m/sec2 = 32.17

ft/sec2).

n = L/W

In some instances of flight such as turn and pull-up, the aircraft must generate a lift

force such that it is more than weight. In some instances; especially for missiles; this load

factor may get as high as 30. Hence, the structure must carry this huge load safely. The aircraft

structure must be strong enough to carry other loads including acceleration load such that

aircraft is able to perform its mission safely. On the other hand, if the load is more than

allowable design value, the structure will lose its integrity and may disintegrate during flight.

Load factor is usually positive, but in some instances; including pull-down, or when

encountering a gust; it may become negative. In general, the absolute value of maximum

negative load factor must not exceed 0.4 times maximum positive load factor. Past experiences

forced Federal Aviation Administration to regulate load factor on aircraft.

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V-n Diagram calculation

VD=1.55 *580.571

=899.8862m/s = 1745.779 keas

VS= 291.66 m/s = 566.987 KEAS

N=L/W = 1.17*10-5 V2

V=716.114m/s =1392.125 KEAS

For lower curve,

Vsi=245.756m/s =476.76 KEAS

-n= -L/W = 1.655*10-5*V2

LOAD FACTOR = -3

-3 = 1.655*10-5*V2

V = 425.62 m/s = 825.755 KEAS

O (0, 0)

A (1, 566.987)

B (6, 1392.125)

F (6, 1745.779)

G (-3, 1745.779)

J (-3,825.755)

K (-1,476.76)

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AERODYNAMIC FORCE CALCULATIONS

Whatever airfoil you finally select MUST be rigorously tested in a real wind tunnel

before being used in a real-world application. Accurate prediction of the maximum lift

coefficient is numerically and practically difficult.

A software-DesignFOIL was used for simulation purposes.

DesignFOIL contains analysis tools that should be used for preliminary airfoil analysis

only; i.e. for comparison of different airfoils in an effort to narrow down your choices.

DesignFOIL uses a unique method based on trend analysis of real wind tunnel data that

combines aerodynamic coefficients with both flow parameters and geometric ratios.

This method has been "calibrated" with experimental wind tunnel data

AOA Cl Cd Cm

-5 -0.616 0.0081 0.009

-4 -0.493 0.0078 0.008

-3 -0.37 0.0073 0.006

-2 -0.247 0.0054 0.004

-1 -0.123 0.0051 0.002

0 0 0.0053 0

1 0.123 0.0051 -0.002

2 0.247 0.0054 -0.004

3 0.37 0.0073 -0.006

4 0.493 0.0078 -0.008

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5 0.616 0.0081 -0.009

6 0.739 0.0086 -0.011

7 0.852 0.0095 -0.013

8 0.948 0.0103 -0.015

9 1.028 0.011 -0.017

10 1.093 0.012 -0.019

11 1.143 0.014 -0.02

12 1.178 0.0154 -0.022

13 1.2 0.017 -0.024

14 1.208 0.0187 -0.026

15 1.203 0.0207 -0.027

16 1.184 0.0228 -0.029

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Simulation to obtain Boundary Layer thickness

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Simulation to analyze the pressure distribution on the airfoil

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CL vs AoA

Lift and Drag can be found out using the following formulae:

Lift = CL x ((denity*velocity2)/2) x Area

Drag = CD x ((denity*velocity2)/2) x Area

For this calculation, cruise velocity (Mach = 0.85) was taken at sea level conditions

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AOA Cl Cd Cm Lift Drag

-5 -0.616 0.0081 0.009 -15014146.6 197426.3

-4 -0.493 0.0078 0.008 -12016192 190114.2

-3 -0.37 0.0073 0.006 -9018237.39 177927.4

-2 -0.247 0.0054 0.004 -6020282.8 131617.5

-1 -0.123 0.0051 0.002 -2997954.59 124305.4

0 0 0.0053 0 0 129180.2

1 0.123 0.0051 -0.002 2997954.593 124305.4

2 0.247 0.0054 -0.004 6020282.801 131617.5

3 0.37 0.0073 -0.006 9018237.394 177927.4

4 0.493 0.0078 -0.008 12016191.99 190114.2

5 0.616 0.0081 -0.009 15014146.58 197426.3

6 0.739 0.0086 -0.011 18012101.17 209613.1

7 0.852 0.0095 -0.013 20766319.62 231549.3

8 0.948 0.0103 -0.015 23106186.62 251048.2

9 1.028 0.011 -0.017 25056075.79 268109.8

10 1.093 0.012 -0.019 26640360.73 292483.4

11 1.143 0.014 -0.02 27859041.46 341230.6

12 1.178 0.0154 -0.022 28712117.97 375353.7

13 1.2 0.017 -0.024 29248337.49 414351.4

14 1.208 0.0187 -0.026 29443326.41 455786.6

15 1.203 0.0207 -0.027 29321458.34 504533.8

16 1.184 0.0228 -0.029 28858359.66 555718.4

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CONCLUSION

Design is a fine blend of science, creativity, presence of mind and the application of

each one of them at the appropriate time. Design of anything needs experience and an

optimistic progress towards the ideal system. The scientific society always looks for the best

product design. This involves the strong fundamentals in science and mathematics and their

skillful applications, which is a tough job endowed upon the designer.

A lot of work was put into this design project. A design never gets completed in a

flutter, but it is always a step further towards ideal system. During the course of this project,

we were exposed to industrial areas of Aircraft Design, Analysis, Modeling, Weight Balancing,

and other finer areas of Aircraft Design. An idea of how aircrafts are conceived and developed

was obtained.

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