sagar final adp
TRANSCRIPT
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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