HINDUSTAN COLLEGE OF ENGINEERING

AIRCRAFT DESIGN PROJECT-II

10 SEATER BUSINESS JET

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

1

ACKNOWLEDGEMENT

I would like to extent my heartfelt thanks to Mr. E. Rajakuperan (Head of Aeronautical Department) for giving me his able support and encouragement. At this juncture I must emphasis the point that this DESIGN PROJECT would not have been possible without the highly informative and valuable guidance by Miss Anindya, whose vast knowledge and experience has helped us go about this project with great ease. We have great pleasure in expressing our sincere & whole hearted gratitude to them.

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

CONTENTS2

SR NO TOPIC PAGE

1 INTRODUCTION 5

2 V-n DIAGRAM 6

3 GUST LOADS 14

4 WING STRUCTURAL LAYOUT 20

5 WING BOX CONFIGURATION 23

6 FUSELAGE STRUCTURAL ANALYSIS 29

7 WING LOADING 32

8 FUSELAGE STRESS ANALYSIS 35

9MANUVERING LOADS ON AIRCRAFT

CONTROL SURFACES41

10 MATERIAL SELECTION 47

11DESIGN OF STRUCTURAL COMPONENTS OF

THE WING50

12 FLIGHT CONTROLS 58

13 LANDING GEAR CONFIGURATION 63

14 THREE-VIEW DIAGRAM 66

15 BIBLIOGRAPHY 68

INTRODUCTION

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Aircraft Design Project-II is a continuation of Aircraft Design Project-I. As

mentioned in our earlier project, Business jet, private jet or,

colloquially, bizjet is a term describing a jet aircraft, usually of smaller size,

designed for transporting groups of up to 19 business people or wealthy

individuals. Business jets may be adapted for other roles, such as the evacuation

of casualties or express parcel deliveries, and a few may be used by public

bodies, governments or the armed forces. The more formal terms of corporate

jet, executive jet, VIP transport or business jet tend to be used by the firms that

build, sell, buy and charter these aircraft. In our Aircraft Design Project-I, we

have performed a rudimentary analysis. We have carried out a preliminary

weight estimation, power plant selection, aerofoil selection, wing selection and

aerodynamic parameter selection and analysis. Apart from the above mentioned,

we have also determined performance parameters such lift, drag, range,

endurance, thrust and power requirements.

Aircraft Design Project-II deals with a more in-depth study and analysis of

aircraft performance and structural characteristics. In the following pages we

have carried out structural analysis of fuselage and wings and the appropriate

materials have been chosen to give our aircraft adequate structural integrity. The

flight envelope of our aircraft has also been established by constructing the V-n

diagram. We have also determined the landing gear position, retraction and

other accompanying systems and mechanisms.

The study of all the above mentioned characteristics, has given us insight into

the complexity of designing a subsonic multi-role 10 seater business jet.

V-n DIAGRAM FLIGHT ENVELOP

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The velocity load factor diagram gives important guidelines to engineers in

defining the characteristics of the flight envelop. The V-n diagram is a graphical

representation of the flight conditions under different velocities and flight loads.

The V-n diagram is therefore defined as the plot showing the structural and

aerodynamic limitations of an aircraft at different velocities.

The flight operating strength of an aircraft is projected on a graph whose

horizontal scale is the airspeed and vertical speed is the load factor.

The design of an aircraft or any structural component in an aircraft is dedicated

by the loads it can withstand. Initial structural analysis is a part of conceptual

design process. Before an actual structural member can be sized and analyzed,

the loads it can withstand must be determined. To determine the loads acting on

each structural member and ultimately the loads on entire aircraft, the term load

factor is defined.

Load Factor

A load factor is the ratio of the total air load acting on the airplane to the gross

weight of the airplane. For example, a load factor of 3 means that the total loads

on an airplane’s structure is three times its gross weight. Load factors are usually

expressed in terms of “G”—that is, a load factor of 3 may be spoken of as 3 G’s,

or a load factor of 4 as 4 G’s.

It is interesting to note that in subjecting an airplane to 3 G’s in a pull-up from a

dive; one will be pressed down into the seat with a force equal to three times the

person’s weight. Thus, an idea of the magnitude of the load factor obtained in

any maneuver can be determined by considering the degree to which one is

pressed down into the seat. Since the operating speed of modern airplanes has

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increased significantly, this effect has become so pronounced that it is a primary

consideration in the design of the structure for all airplanes.

We know that the load factor (n) is given by;

n=L/W

Where, L = total lift

W = total weight

When L = W (steady level un-accelerated flight), n=1 and this is also termed as

‘1g load factor’.

There are two kinds of V-n diagrams:

1. Maneuvering V-n diagram

2. Gust V-n diagram

Maneuvering V-n diagram

The general flight envelop which shows flight characteristics for various load

factors and velocities is called as the maneuvering V-n diagram. The

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performance of an aircraft under normal flight attitudes and maneuvers is

obtained from this flight envelop.

N=L/W= (1/2V2SCL)/W---------------------- (1)

Therefore at higher speeds maximum load factor is limited by the structural

design of the airplane. Thus the aircraft load factor expresses the maneuvering

of an aircraft as a multiple of standard acceleration due to gravity. At lower

speeds the highest load factor an aircraft may experience is limited by the

maximum lift.

Max N=L/W= (1/2V2SCL, max)/W---------------------- (2)

 The figure represents the flight operating strength of an aircraft on a graph

whose vertical scale is based on load factor. It is valid only for a specific weight,

configuration and altitude and shows the maximum amount of positive or

negative lift the airplane is capable of generating at a given speed. Also shows

the safe load factor limits and the load factor that the aircraft can sustain at

various speeds.      

7

 The lines of maximum lift capability (curved lines) are the first items of

importance on the Vg diagram. The aircraft in the chart above is capable of

developing no more than +1 G at 62 mph, the wing level stall speed of the

aircraft. Since the maximum load factor varies with the square of the airspeed,

the maximum positive lift capability of this aircraft is 2 G at 92 mph, 3 G at 112

mph, 4.4 G at 137 mph, and so forth. Any load factor above this line is

unavailable aerodynamically (i.e., the aircraft cannot fly above the line of

maximum lift capability because it stalls). The same situation exists for negative

lift flight with the exception that the speed necessary to produce a given

negative load factor is higher than that to produce the same positive load

factor.   

 If the aircraft is flown at a positive load factor greater than the positive limit

load factor of 4.4, structural damage is possible. When the aircraft is operated in

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this region, objectionable permanent deformation of the primary structure may

take place and a high rate of fatigue damage is incurred. Operation above the

limit load factor must be avoided in normal operation.   

There are two other points of importance on the Vg diagram. One point is the

intersection of the positive limit load factor and the line of maximum positive

lift capability. The airspeed at this point is the minimum airspeed at which the

limit load can be developed aerodynamically. Any airspeed greater than this

provides a positive lift, which is sufficient to damage the aircraft. Conversely,

any airspeed less than this do not provide positive lift capability sufficient to

cause damage from excessive flight loads. The usual term given to this speed is

“maneuvering speed,” since consideration of subsonic aerodynamics would

predict minimum usable turn radius or maneuverability to occur at this

condition. The maneuver speed is a valuable reference point, since an aircraft

operating below this point cannot produce a damaging positive flight load. Any

combination of maneuver and gust cannot create damage due to excess air load

when the aircraft is below the maneuver speed.   

  The other point of importance on the Vg diagram is the intersection of the

negative limit load factor and line of maximum negative lift capability. Any

airspeed greater than this provides a negative lift capability sufficient to damage

the aircraft; any airspeed less than this does not provide negative lift capability

sufficient to damage the aircraft from excessive flight loads.   

 The limit airspeed (or redline speed) is a design reference point for the aircraft

—this aircraft is limited to 225 mph. If flight is attempted beyond the limit

airspeed, structural damage or structural failure may result from a variety of

phenomena.    

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The aircraft in flight is limited to a regime of airspeeds and Gs which do not exceed the limit (or redline) speed, do not exceed the limit load factor, and cannot exceed the maximum lift capability. The aircraft must be operated within this “envelope” to prevent structural damage and ensure the anticipated service lift of the aircraft is obtained. The pilot must appreciate the Vg diagram as describing the allowable combination of airspeeds and load factors for safe operation. Any maneuver, gust, or gust plus maneuver outside the structural envelope can cause structural damage and effectively shorten the service life of the aircraft.

The general V-n diagram (for maneuvering load) is calculated using the

following velocities:

1. Cruise speed

2. Maneuver speed

3. Dive speed

4. Stall speed

Stall Speed

Stall speed is the slowest speed the aircraft can travel. If the speed of the aircraft

decreases below the stall speed the aircraft will not be able to sustain steady

flight and will stall.

Since stall speed is a function of coefficient of lift

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Stall speed is given by,

VS= [(2*GW)/ (p*S*CL)] 0.5

S - wing area, square feet

GW - gross weight, pounds

p - Density of air, at sea level = 0.00238 slugs/cubic feet

Vs - stall speed, feet per second

CL - lift coefficient ,for conventional aircraft with plain flaps CL = 1.8

For our aircraft we have the following specifications

GW=20000kg=44092.45pounds

S=57.42 m2 = 618.01 ft2

VS= [(2*44092.45)/ (0.00238*618.01*1.8)] 0.5

VS = 182.5 fps

VS = 55.63 m/s.

Maneuvering speed

Maneuvering speed is the highest speed at which full deflection of the controls

about any one axis are guaranteed not to overstress the airframe. At or below

this speed, the controls may be moved to their limits. Above this speed, moving

the controls to their limits may overstress the airframe and potentially cause a

structural failure. It is normally designated as VA .

The maneuvering speed is given by

Va= Vstall * (positive limit load factor)0.5

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The positive limit load factor for our aircraft is n(+ve)=L/W(at CL max) =L/W(at max CL)

=(0647123.4/196000) = 3.3

Therefore Va = 55.63*3.30.5

Va=101m/s

Cruise speed

VC=200m/s (from design data sheet)

Dive speed

Vd= 1.25*Vc

Vd= 1.25*200

Vd = 250m/s

Thus the V-n diagram plotted based on these values is as given below:

Fig. V-n diagram for maneuvering load.

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Gust V-n Diagram

Gust loads are encountered anytime the aircraft encounters a rush of wind. Gust

loads are also encountered when the aircraft is flying in a thunderstorm or in

turbulence. These loads can be higher than maneuvering loads. Gust is very un

predictable and hence the gust V-n diagram must be given importance in order

to establish a safe flight envelop.

When an aircraft experiences a gust loads, there is generally an increase or

decrease in the angle of attack. The figure indicates the effect of upward gust of

velocity U. The angle of attack is approximately U divided by V and the change

in lift is approximately proportional to the gust velocity.

The change in the aircraft load factor due to gust is derived as follows:

Δα =tan-1(U/V)

ΔL=1/2ρV2S(CL,a Δα)

ΔL=1/2ρVSCL,a Δα

13

Thus the change in load factor is Δn= ΔL/W= ρUVCL,a

2(W/S)

Where U is the upward component of velocity due to gust loads ;V is the

direction of relative wind, ρ is the density and CL,a is the changed lift coefficient

due to gust.

Gust reduces the acceleration of the aircraft by as much as 40%.To account for

this, a gust elevation factor K has been devised and applied to measure gust data.

The gust velocity Ugust is given as:

Ugust=K*U

For subsonic K = (0.88µ)/(5.3+µ)

For supersonic K = µ1.03/(6.95+ µ1.03)

The mass ratio µ = 2(W/S)

ρcgCL,a

Where c is the mean aerodynamic chord. The mass ratio accounts for the fact

that a small light plane encounters the gust more rapidly than a large plane. For

most years, the standard vertical gust has been U=35ft/sec or 10.668m/s. This

value is a suitable gust velocity has been used in the following calculations.

Therefore for a stall speed Vstall of 55.63m/s, the change in load factor is

calculated as

Δn= ρUVstallCL,a

2(W/S)

=(1.22*55.63*0.2*10.6)/((2*196000)/57.42)

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=0.0210

ngust=n+Δn

Therefore, for the stall speed of Vstall, the gust load factor is

ngust=1+0.0210=1.0210

For design maneuvering speed of Va of 101m/s.

Δn= ρUVaCL,a

2(W/S)

=(1.22*101*0.2*10.6)/((2*196000)/57.42)

=0.038

ngust=3.3+0.038=3.338

For the design cruise speed, Vc=200m/s.

Δn= ρUVcCL,a

2(W/S)

=(1.22*200*0.2*10.6)/((2*196000)/57.42)

=0.075

ngust=3.3+0.075=3.375

For the design dive speed, Vdive=250m/s.

Δn= ρUVdCL,a

2(W/S)

Δn =(1.22*250*0.2*10.6)/((2*196000)/57.42)

=0.094

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ngust=3.3+0.094=3.394

Similarly for the negative angles of attack, the negative lift coefficient is

considered which in turn gives the negative load factor i.e. -1.5 and the load

factor for gust is as follows:

For Vstall,ngust = -1+0.0210= -0.979

For Va ,ngust = -1.5+0.038= -1.462

For Vc,ngust = -1.5+0.075= -1.425

For V d,ngust = -1.5+0.094= -1.406

Based on these values the V-n diagram for gust encounter is plotted as shown

below:

Fig. Gust V-n diagram

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It is assumed that the aircraft is in 1-g load factor when the aircraft experiences

gust.

Notice the shift in the V-n diagram due to gust effects. The load factor between,

dive cruise maneuver is assumed to follow a straight line. The gust line for

stall ,cruise and maneuver can be observed clearly in the above graph.

Therefore joining the points B, C, D, E, D’, and F complete the gust V-n

diagram.

The maneuvering and the gust V-n diagram are combined to determine the most

critical load factor at each speed. Since the gust loads are greater than the limit

loads, the increased limit load at all velocities has been denoted by the dotted

line.

Fig. Combined V-n diagram

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One interesting point to note for gust V-n diagram is that the load factor due to

gust increases if the aircraft is lighter. This is counter to the natural assumption

that the an aircraft is more likely to have structural failure if it is heavily loaded.

In fact the change in lift due to gust is heavily unaffected by the weight, so that

the change in wing stress is same in either case. If the aircraft is lighter the same

lift increase will cause greater vertical acceleration and hence the rest of the

aircraft experiences greater stress.Aeroelastic effect also influences load factor

due to gust.

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WING STRUCTURAL LAYOUT

SPECIFIC ROLES OF WING (MAINPLANE) STRUCTURE:

The specified structural roles of the wing (or main plane) are:

To transmit: wing lift to the root via the main span wise beam

Inertia loads from the power plants, undercarriage, etc, to the main beam.

Aerodynamic loads generated on the aerofoil, control surfaces & flaps to

the main beam.

To react against:

Landing loads at attachment points

Loads from pylons/stores

Wing drag and thrust loads

To provide:

Fuel tank age space

Torsion rigidity to satisfy stiffness and aero elastic requirements.

To fulfill these specific roles, a wing layout will conventionally compromise:

Span wise members (known as spars or booms)

Chord wise members(ribs)

A covering skin

Stringers

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Basic Functions of wing Structural Members

The structural functions of each of these types of members may be

considered independently as:

Spars:

Form the main span wise beam

Transmit bending and torsional loads

Produce a closed-cell structure to provide resistance to torsion, shear and

tension loads.

20

Skin:

To form impermeable aerodynamics surface

Transmit aerodynamic forces to ribs & stringers

Resist shear torsion loads (with spar webs).

React axial bending loads (with stringers).

Stringers:

Increase skin panel buckling strength by dividing into smaller length

sections.

React axial bending loads

Ribs:

Maintain the aerodynamic shape

Act along with the skin to resist the distributed aerodynamic pressure

loads

Distribute concentrated loads into the structure & redistribute stress

around any discontinuities

Increase the column buckling strength of the stringers through end

restraint

Increase the skin panel buckling strength.

21

WING BOX CONFIGURATIONS

Several basic configurations are in use now-a-days:

Mass boom concept Box Beam(distributed flange) concept-built-up or integral

construction Multi-Spar Single spar D-nose wing layout

Mass Boom Layout

In this design, all of the span wise bending loads are reacted against by substantial booms or flanges. A two-boom configuration is usually adopted but a single spar “D-nose” configuration is sometimes used on very lightly loaded structures. The outer skins only react against the shear loads. They form a closed-cell structure between the spars. These skins need to be stabilized against buckling due to the applied shear loads; this is done using ribs and a small number of span wise stiffeners.

Box Beam or Distributed Flange Layout:22

This method is more suitable for aircraft wings with medium to high load

intensities and differs from the mass boom concept in that the upper and lower

skins also contribute to the span wise bending resistance.

Another difference is that the concept incorporates span wise stringers

(usually “z” section) to support the highly –stressed skin panel area. The

resultant use of a large number of end-load carrying members improves the

overall structural damage tolerance.

Design Difficulties Include:

Interactions between the ribs and stringers so that each rib either has to

pass below the stringers or the load path must be broken. Some examples

of common design solutions are shown in figure

Many joints are present, leading to high structural weight, assembly times,

complexity, costs & stress concentration areas.

The concept described above is commonly known as built-up

construction method. An alternative is to use a so-called integral construction

method. This was initially developed for metal wings, to overcome the inherent

drawbacks of separately assembled skin-stringer built-up construction and is

very popular now-a-days. The concept is simple in that the skin-stringer panels

are manufactured singly from large billets of metal. Advantages of the integral

construction method over the traditional built-up method include:

Simpler construction & assembly

Reduced sealing/jointing problems

Reduced overall assembly time/costs

Improved possibility to use optimized panel tapering23

Disadvantages include:

Reduced damage tolerance so that planks are used

Difficult to use on large aircraft panels.

Fig. Basic metal-sparred wing using a honeycomb   'D' box   leading edge

Types of spars:

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In the case of a two or three spar box beam layout, the front spar should be located as far forward as possible to maximize the wing box size, though this is subject to there being:

Adequate wing depth for reacting vertical shear loads. Adequate nose space for LE devices, de-icing equipment, etc.

This generally results in the front spar being located at 12 to 18% of the chord length. For a single spar D-nose layout, the spar will usually be located at the maximum thickness position of the aerofoil section. For the standard box beam layout, the rear spar will be located as far as aft as possible, once again to maximize the wing box size but positioning will be limited by various space requirements for flaps control surfaces spoilers etc

This usually results in a location somewhere between about 55 and 70% of the chord length. If any intermediate spars are used they would tend to be spaced uniformly unless there are specific pick-up point requirements.

Ribs:25

For a typical two spar layout, the ribs are usually formed in three parts

from sheet metal by the use of presses and dies. Flanges are incorporated around

the edges so that they can be riveted to the skin and the spar webs Cut-outs are

necessary around the edges to allow for the stringers to pass through Lightening

holes are usually cut into the rib bodies to reduce the rib weight and also allow

for passage of control runs fuel electrics etc.

Rib construction and configuration:

The ribs should be ideally spaced to ensure adequate overall buckling

support to spar flanges .In reality however their positioning is also influenced by

Facilitating attachment points for control surfaces, flaps, slats, spoiler

hinges, power plants, stores, undercarriage attachments etc

Positions of fuel tank ends, requiring closing ribs

A structural need to avoid local shear or compression buckling.

26

Rib Alignment Possibilities:

There are several different possibilities regarding the alignment of

the ribs on swept-wing aircraft

(a) Is a hybrid design in which one or more inner ribs are aligned

with the main axis while the remainder is aligned

perpendicularly to the rear spar

(b) Is usually the preferred option but presents several structural

problems in the root region

(c) Gives good torsional stiffness characteristics but results in

heavy ribs and complex connection

27

FUSELAGE STRUCTURAL LAYOUT

The fundamental purpose of the fuselage structure is to provide an envelope to

support the payload, crew, equipment, systems and (possibly) the power plant.

Furthermore, it must react against the in-flight maneuver, pressurization and

gust loads; also the landing gear and possibly any power plant loads. Finally, it

must be able to transmit control and trimming loads from the stability and

control surfaces throughout the rest of the structure.

Fuselage Layout Concepts

There are two main categories of layout concept in common use:

• Mass boom and longeron layout

• Semi-monocoque layout

Mass Boom & Longeron Layout

This is fundamentally very similar to the mass-boom wing-box concept. It is

used when the overall structural loading is relatively low or when there are

extensive cut-outs in the shell. The concept comprises four or more continuous

heavy booms (longerons), reacting against any direct stresses caused by applied

vertical and lateral bending loads. Frames or solid section bulkheads are used at

positions where there is distinct direction changes and possibly elsewhere along

the lengths of the longeron members. The outer shell helps to support the

longerons against the applied compression loads and also helps in the shear

carrying. Floors are needed where there are substantial cut-outs and the skin is

stabilized against buckling by the use of frames and bulkheads.

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Mass boom & longeron fuselage layout

Semi Monocoque Layout

This is the most common layout, especially for transport types of aircraft, with a

relatively small number and size of cut-outs in use. The skin carries most of the

loading with the skin thickness determined by pressurization, shear loading &

fatigue considerations.

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Fig. Semi Monocoque fuselage layout

Longitudinal stringers provide skin stabilization and also contribute to the

overall load carrying capacity. Increased stringer cross-section sizes and skin

thicknesses are often used around edges of cut-outs. Less integral machining is

possible than on an equivalent wing structure. Frames are used to stabilize the

resultant skin-stringer elements and also to transmit shear loads into the

structure. They may also help to react against any pressurization loads present.

They are usually manufactured as pressings with reinforced edges. Their spacing

(pitch) is usually determined by damage tolerance considerations, i.e. crack-

stopping requirements. The frames are usually in direct contact with the skin;

stringers pass through them and are seated into place.

30

WING LOADING

In aerodynamics, wing loading is the loaded weight of the aircraft divided by the

area of the wing. The faster an aircraft flies, the more lift is produced by each

unit area of wing, so a smaller wing can carry the same weight in level flight,

operating at a higher wing loading. Correspondingly, the landing and take-off

speeds will be higher. The high wing loading also decreases maneuverability.

We know that lift, L=CL*1/2*ρV2*S=weight

Thus W= CL*1/2*ρV2*S or W/S= CL*1/2*ρV2=wing loading

From this we can see that if wing loading increases in a constant speed

maneuver then CL, the angle of attack muincrease. Conversely if CL is increased

during a constant maneuvre, the lift and consequently the wing loading must

increase.

Most general aviation aircraft have a designed wing loading between 500 and

1000 N/m2.Aircraft designed with higher wing loading are more maneuverable

but have higher minimum speed than aircraft with lower wing loading. Wing

loading is normally stated in pounds per square foot. In most airplane designs,

wing loading is determined by considerations of Vstall and landing distance.

However, W/S also plays a major role in maximum velocity

We have,

W/S = 0.5*ρ0*Vstall2× (CL) max

W/S = 0.5×1.225×55.62× (1.17)

W/S = 2215.30kg/m2

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Let us examine the constraint imposed by the specified landing distance. The

landing distance is the sum of the approach distance Sa, the flare distance Sf,

and the ground roll Sg. The approach angle θa requires knowledge of L/D and

T/W. Since we have not made estimates of either quality yet we assume, based

on the thumb rule, i.e. θ<=3°, for small passenger aircraft we take θa=3o.

R = Vf2/0.2g= (1.23 Vstall) 2/ (0.2×9.8)

R = 2377.16m

The flare height hf is given by,

hf= R (1-cosθa) = 2377.16(1-cos3o) => hf = 1.4m

The approach distance required to clear a 50 feet obstacle is given by

Sa= (50-hf)/tanθa = (50-1.4)/tan2°

Sa = 892.99m

The flare distance Sf is given by

Sf = Rsinθa =2377.16×sin3°

Sf = 216.65m

In the equation of Sg let us assume that the lift has been intentionally made

small by retracting the flaps combined with a small angle of attack due to the

rather level orientation of the airplane relative to the ground. Furthermore,

assuming no provision for thrust reversal and ignoring the drag compared to the

friction force between the tires and the ground we have,

Sg = jN(2×W)/(ρ0×S×CLmax)).5 +(j2(W/S))/(g×ρ× CLmax×µ)

32

As stated above j=1.15 for commercial airplanes. Also, N is the time increment

for free roll immediately after touchdown, before the brakes are applied. By

assuming N=3s and µ=0.4 we get,

Sg = 1.15*3(2×W/S)/(1.225×1.17))0.5 +(1.152(W/S))/(9.81×1.225×1.17*0.4)

Sg = 4.075(W/S)0.5 + 0.235(W/S)

Since the allowable landing distance is specified in the requirement as 2100m

and we have previously determined Sa and Sf, the allowable value for Sg is

Sg = 2100-892.99-124.4

Sg = 1082.61

Therefore we have

4.075(W/S)0.5 + 0.235(W/S)=1082.61, solving for (W/S) we get,

W/S=357kg/m 2

33

FUSELAGE STRESS ANALYSIS

The fuselage has a circular cross-section as shown in the above figure. The

cross-sectional area of each stringer is 100mm2 and the vertical distances given

in the figure are measured from the mid-line of the section wall at the

corresponding stringer position.

The fuselage is subjected to a bending moment of 200 kN-m applied in the

vertical plane of symmetry. We will now be determining the direct stress

distribution at each stringer.

The section is first idealized. As an approximation we shall assume that the skin

between adjacent stringers is flat so that we may use the following equations to

determine the boom areas.

From the symmetry,

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Substituting,

i.e.,

Similarly . We note that stringers 5

and 13 lie on the neutral axis of the section and are therefore unstressed; the

calculation of the boom areas B5 and B13 does not arise.

Stinger/ Boom y

1 900 51.93

2,16 831.5 47.97755

3,15 636.61 36.7324

4,14 344.41 19.87246

5,13 0 0

6,12 -344.41 -19.8725

7,11 -636.61 -36.7324

8,10 -831.5 -47.9776

9 -900 -51.93

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For this section Ixy=0 and My=0

We know,

Where,

Solving the above equation, we obtain the direct stress distribution on the

fuselage which is shown in the above table.

SHEAR FLOW ANALYSIS

Initially the value of Sy is to be found, Sx = 0. To find Sy the circulation acing on

the cylinder (fuselage) is to be determined.

We know,

Hence the resultant load acting Sy = 82.131 kN

36

The shear flow is given by the expression:

Substituting the required values we get,

To determine the shear flow for the closed section we assume that the panel 12

is cut. Now the shear flow for the open section is determined by the following

formulae;

Rearranging the above equation we get,

Where,

Ap=Area of each panel ( in this case it is uniform)

AT=Total area

Solving the above equation we get the shear flow for the open section;

37

The following tabular column contains the shear flow values over the fuselage.

SKIN

PANEL

Stinger/

BoomBr(mm2) yr(mm)

qb,o(N/

mm)

qb(N/

mm)

1 2 - - - 0 3.947

2 3 2 534.72 831.5 -10.5357 -6.58871

3 4 3 534.72 636.61 -8.06631 -4.11931

4 5 4 534.72 344.41 -4.36392 -0.41692

5 6 5 0 0 0 3.947

6 7 6 534.72 -344.41 4.363924 8.310924

7 8 7 534.72 -636.61 8.06631 12.01331

8 9 8 534.72 -831.5 10.53571 14.48271

1 16 1 534.72 900 -11.4037 -7.45665

16 15 16 534.72 831.5 -10.5357 -6.58871

15 14 15 534.72 636.61 -8.06631 -4.11931

14 13 14 534.72 344.41 -4.36392 -0.41692

13 12 13 0 0 0 3.947

12 11 12 534.72 -344.41 4.363924 8.310924

11 10 11 534.72 -636.61 8.06631 12.01331

10 9 10 534.72 -831.5 10.53571 14.48271

38

SHEAR FLOW DIAGRAM

Therefore the shear flow diagram for the fuselage is given as follows,

39

MANEUVERING LOADS ON AIRCRAFT

CONTROL SURFACES

Aircraft load estimation combines aerodynamics, structures, and weights. Load

estimation remains a critical area because an error or faulty assumption will

make the aircraft too heavy or will result in structural failure when real loads are

encountered in flight.

Loads acting on the aircraft can be classified according to the following load

categories:

Air loads

Manoeuvre

Gust

Control deflection

Component interaction

Buffet

Landing

Vertical load factor

Spin up

Spring back

Crabbed

One wheel

Arrested

Braking

40

Inertia loads

Acceleration

Rotation

Dynamic

Vibration

Flutter

Power plants loads

Thrust

Torque

Gyroscope

Vibration

Duct pressure

Take off loads

Catapult

Aborted

Taxi

Bumps

Turning

Other loads

Towing

Jacking

Pressurization

Bird strike

Crash

Limit load41

The largest load the aircraft is expected to encounter without any

permanent deformation is known as limit load or applied load.

Design load

To provide a margin of safety, the aircraft structure is always designed

to withstand higher load than the limit load. The highest load the aircraft is

designed to withstand without breaking is the design or ultimate load.

Load sources

There are generally two cases of the load sources

1. Maneuverability cases

2. Environmental cases

Maneuverability cases

In this the loads which act on the aircraft is due to the pilot’s action.

E.g.: pull up, pull down etc.

Environmental cases

In this the loads are imposed by the environment on the aircraft where it

operates.

E.g.: turbulence loads, kinetic heating loads, bird strike etc.

Load factors42

Any force applied to an airplane to deflect its flight from a straight line

produces a stress on the structure; the amount of this force is termed as load

factor.

A load factor is the ratio of the total air load acting on the airplane to the

gross weight of the airplane.

n=L / W

For e.g., a load factor of 3 means that the total load on an airplane’s structure is

three times the gross weight.

Category limit load

Normal 3.8 to -1.25

Utility 4.4 to -1.76

Acrobatic 6.6 to -3.0

Maneuver loads

The greatest air loads on an airplane usually come from the generation of

lift during high-g maneuvers. Aircraft load factor (n) expresses the maneuvering

of an aircraft as a multiple of the standard acceleration due to gravity.

Maneuvering loads on elevator

Operation of the control surfaces produces air loads in several ways. The

greatest impact is in the effect of the elevator on angle of attack and hence the

load factor.

Deflection of control surfaces produces additional loads directly upon the wing.

Maneuver speed or pull up speed (Vp), is the maximum speed at which the pilot

can fully deflect the controls without damaging either the airframe or the control

themselves.

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The figure shows the loading distribution of a horizontal tail consisting of a

fixed stabilizer and a moving elevator. Under some combinations of angle of

attack and elevator position the stabilizer and elevator will actually have loads in

the opposite directions.

For design purposes, the elevator load is assumed to equal 40% of the total

required tail load but in the opposite direction. The distributed load shown on

the stabilizer must then be equal 140% of the tail load. The smoothest pull up

possible, with a moderate load factor, will deliver the greatest gain in the

altitude and will result in better overall performance.

The normal stall entered from straight level flight or an un-

accelerated straight climb, will not produce added load factors beyond the IG of

straight and level flight. In this event recovery is affected by snapping the

elevator control forward, negative load factors, those which impose a down load

on the wings. A recovery from stall is made by dividing only to cruising or

design maneuvering airspeed, with a gradual pull up as soon as the airspeed is

safely above stalling, can be affected with load factor not to exceed 2 or 2.5.

Maneuvering loads on ailerons

In the level turning flight, the lift of the wing is canted so that

the horizontal component of the lift exerts the centripetal force required to turn

the total lift on the wing is ‘n’ times the aircraft weight W.

n - Load factor

Turn rate (ψ) =g*(n^2) ^0.5/V

=68.76 o /second.

Instantaneous turn rate

44

If the aircraft is allowed to slow down during the turn which is

known as instantaneous turn, the load factor ‘n’ will be limited only by the

maximum lift coefficient or structural strength of the aircraft.

Sustained turn rate

In a sustained turn rate, the aircraft is not permitted to slow down

or lose altitude during the turn. In a sustained turn the thrust must equal the drag

and the lift must equal load factor ‘n’ times the weight. Thus the maximum load

factor for sustained turn can be expressed as the product of the thrust to weight

and lift to drag ratios, assuming that the thrust axis is approximately aligned

with the flight directions.

Maneuvering loads on rudder

In flight yaw control is provided by the rudder and the directional

stability by vertical stabilizer. The vertical stabilizer and the rudder must be

capable of generating sufficient yawing moments to maintain directional control

of the aircraft. The rudder deflection, necessary to achieve these yawing

moments and the resulting sideslip angles place significant aerodynamic loads

on the rudder and on the vertical stabilizer.

Both are designed to sustain in several lateral loading conditions

leading to the required level of structural strength.

With the aircraft in un-accelerated and stabilized straight flight, the rudder is

suddenly displaced to the maximum available deflection at the current airspeed.

MATERIAL SELECTION45

The next important task is to select the various materials required to

fabricate the entire aircraft such as the skin, fuselage, wings, control surfaces

etc; without any of its components failing due to higher rates of tensile and

compressive loads, stresses and strains our aircraft is inhibited to during its

flight. Failure of material will lead to damage which results to loss of life and

expensive component. Hence the material selected should be of very high

strength in compliance with lower costs and shouldn’t tend to increase the

overall weight of the aircraft.

The aircraft being designed features the lighter-weight construction. Its materials

(by weight) are: 50% composite, 20% aluminum, 15% titanium, 10% steel, 5%

of other materials. Composite materials are significantly lighter and stronger

than traditional aircraft materials, making our aircraft lighter for its capabilities.

The aircraft will be 80% composite by volume. It contains approximately 35

tons of carbon fiber reinforced plastic, made with 23 tons of carbon fiber.

Composites are used on fuselage, wings, tail, doors, and interior. Aluminum is

used on wing and tail leading edges, titanium used mainly on engines with steel

used in various places.

COMPOSITES

Composites are the most important materials to be adapted for aviation since the

use of aluminum in the 1920s. Composites are materials that are combinations

of two or more organic or inorganic components. One material serves as a

"matrix," which is the material that holds everything together, while the other

material serves as reinforcement, in the form of fibers embedded in the matrix.

Until recently, the most common matrix materials were "thermosetting"

46

materials such as epoxy, bismaleimide, or polyimide. The reinforcing materials

can be glass fiber, boron fiber, carbon fiber, or other more exotic mixtures.

Our aircraft uses all-composite fuselage and the remaining control surfaces uses

50% composite (mostly carbon fiber reinforced plastic).This makes our aircraft

lighter compared to other aircraft in this range. Each fuselage barrel will be

manufactured in one piece, and the barrel sections joined end to end to form the

fuselage. This will eliminate the need for about 50,000 fasteners used in

conventional airplane building. The composite is also stronger, allowing a

higher cabin pressure during flight compared to aluminum. It was also added

that carbon fiber, unlike metal, does not visibly show cracks and fatigue. They

have also stated that special defect detection procedures will be put in place to

detect any potential hidden damage. Another concern arises from the risk of

lightning strikes. The aircrafts fuselage composite could have as much as 1,000

times the electrical resistance of aluminum, increasing the risk of damage during

lightning strike.

METALS

Composites aren’t the only materials integrated in our aircraft. While

composites represent 50 percent by weight (80 percent by volume) of the

structure, other materials represented are aluminum (20 percent); titanium (15

percent); steel (10 percent) and others (5 percent). Most notable among the

“other” is the widespread use of plastic heat sinks in aircraft structures. Plastics

that are highly loaded with heat-removing materials such as carbon or ceramics

which have been around for a while, but have not yet penetrated the aircraft

market. Their great advantage is their ability to be molded into net shapes. The

economics for plastics can be favorable depending on total tooling and finishing

47

costs. They can be designed with additional surface areas as fins and ribs to

improve convective heat transfer.

Costs and properties can be balanced depending on which engineering

thermoplastics are used. For example, nylon can improve economics while

liquid crystal polymer can improve properties. They are typically loaded 30 to

40 percent with thermally conductive materials.

Other new materials highlighted on our design aircraft are:

Titanium: This aircraft will be using of a new advanced alloy from titanium

which is new in the aircraft industry. The new grade, designated 5553 (Ti-5Al-

5V-5Mo-3Cr), supersedes another high-strength alloy, 1023 (Ti-10V-2Fe-3Al).

Typically, titanium has been used in engine applications for rotors, compressor

blades, hydraulic system components and nacelles.

Aluminum: New technologies are emerging for extrusions in plates in

aluminum-lithium alloys that find its application in our aircraft. It’s well known

that aluminum-lithium alloys have lower density, good and often higher strength

than conventional aluminum alloys, and provide higher modulus, and therefore,

enable weight savings.

Thermally conductive plastics offer significant improvements over conventional

plastics.

48

DESIGN OF COMPONENTS OF THE WING

FUEL TANKS

Aircraft typically use three types of fuel tanks: integral, rigid removable, and

bladder.

Integral tanks are areas inside the aircraft structure that have been sealed

to allow fuel storage. Since these tanks are part of the aircraft structure,

they cannot be removed for service or inspection. Inspection panels must

be provided to allow internal inspection, repair, and overall servicing of

the tank. Most large transport aircraft use this system, storing fuel in the

wings and/or tail of the airplane.

Rigid removable tanks are installed in a compartment designed to

accommodate the tank. They are typically of metal construction, and may

be removed for inspection, replacement, or repair. The aircraft does not

rely on the tank for structural integrity.

Bladder tanks are reinforced rubberized bags installed in a section of

aircraft structure designed to accommodate the weight of the fuel. The

bladder is rolled up and installed into the compartment through the fuel

filler neck or access panel, and is secured by means of metal buttons or

snaps inside the compartment. Many high-performance light aircraft and

some smaller turboprops use bladder tanks.

Pertaining to the initial design carried out to the aircraft, all commercial aircrafts

follow the integral type tank for safety and easier access of fuel to the engine.

49

RIB LOCATION AND DIRECTION

The span-wise location of ribs is of some consequence.

Ideally, the rib spacing should be determined to ensure adequate overall

buckling support to the distributed flanges. This requirement may be

considered to give a maximum pitch of the ribs. In practice other

considerations are likely to determine the actual rib locations such as:

a) Hinge positions for control surfaces and attachment/operating points

for flaps, slats, and spoilers.

b) Attachment locations of power plants, stores and landing gear

structure.

c) A need to prevent or postpone skin local shear or compression

buckling, as opposed to overall buckling.

50

d) Ends of integral fuel tanks where a closing rib is required. When the

wing is upswept, it is usual for the ribs to be arranged in the flight

direction and thereby define the aerofoil section.

Ribs placed at right angles to the rear spar are usually he most

satisfactory in facilitating hinge pick-ups, but they do cause layout problems

in the root regions. There is always the possibility of special exceptions, such

as power plant or store mounting ribs, where it may be preferable to locate

them in the flight direction.

51

FIXED SECONDARY STRUCTURE

A fixed leading edge is often stiffened by a large number of

closely pitched ribs, span-wise members. Considering design of the skin

attachment it is possible to arrange for little span-wise end load to be diffused

into the leading edge and buckling of the relatively light structure is avoided.

This may imply short spam-wise sections. The presence of thermal de-icing,

high-lift devices or other installations in the leading edge also has a

considerable influence upon the detail design. Bird strike considerations are

likely to be important.

Installations also affect the trailing edge structure where

much depends upon the type of flaps, flap gear, controls and systems. It is

always aerodynamically advantageous to keep the upper surfaces as complete

and smooth as is possible. Often spoilers can be incorporated in the region

above flaps or hinged doors provided for ease of access.

HORIZONTAL STABILISER

52

When the horizontal stabilizer is constructed as a single

component across the centreline of the aircraft, the basic structural

requirements are very similar to those of a wing. Here for our aircraft we

have, the basic structural requirements are very similar to those of a wing.

VERTICAL STABILISER

Conventional tail

In conventional tail the vertical stabilizer is exactly vertical.

The vertical stabilizer is mounted exactly vertically, and the horizontal

stabilizer is directly mounted to the empennage (the rear fuselage). This is the

most common vertical stabilizer configuration.

T-tail

A T-tail has the horizontal stabilizer mounted at the top of the vertical stabilizer.

It is commonly seen on rear-engine aircraft

The vertical stabilizer presents a set of issues which are different from those

of the main plane or horizontal stabilizer. Relevant matters are:

It is not unusual to build the vertical stabilizer integrally with the rear fuselage.

The spars are extended to form fuselage frames or bulkheads. A ‘root’ rib is

made to coincide with the upper surface of the fuselage and is used to transmit

the fin root skin shears directly into the fuselage skin. Fin span-wise bending

results in fuselage torsion. Sometimes on smaller aircraft the fin is designed as a

separate component which may readily be detached. The fin attachment lugs are

arranged in both lateral and fore and aft directions so that in addition to vertical

loads they react side and drag loads.

53

There is a special situation when the horizontal stabilizer is attached

at some location across the height of the fin. The horizontal stabilizer

transmits substantial loads to the fin, usually of the same order of magnitude

as the loads on the fin itself. A particular hg loading results from the reaction

of horizontal stabilizer asymmetrical lift case, which always adds to fin

lateral air-loads

AUXILIARY SURFACES

The structural layout of the auxiliary lifting surfaces is generally

similar to that of the wing but there are differences, in part due to the smaller

size and in part due to the need to provide hinges or supports. The latter

implies that each auxiliary surface is a well-defined.

HINGED CONTROL SURFACES

Conventional training edge control surfaces are almost invariably

supported by a number of discrete hinges, although continuous, piano type,

hinges may be used for secondary tabs. To some degree the number and location

of the discrete hinges depends upon the length of the control. The major points

to be considered are:

a) The bending distortion of the control relative to the fixed surface

must be limited so that the nose of the control does mot fouls the

fixed shroud.

b) The control hinge loads and the resulting shear forces and bending

moments should be equalized as far as is possible.

c) Structural failure of a single hinge should be tolerated unless each

hinge is of fail-safe design and can tolerate cracking one load path.

54

PIVOTED CONTROL SURFACES

In certain high-performance aircraft, the whole of a stabilizing or

control surface on one side of the aircraft may be pivot about a point on its root

chord. Clearly in this case, the structural considerations are dominated by the

need to react all the forces and moments at the pivot and operating points.

Some designs incorporate the pivot into the moving surface with the

support bearings on the fuselage, while on others the pivot is attached to the

fuselage and the bearings are in the surface. The bearings should be as far apart

as the local geometry allows to minimize loads resulting from the reaction of the

surface bending moment.

HIGH LIFT SYSTEMS

There is a wide variety of leading and trailing edge high-lift systems.

Some types are simply hinged to the wing, but many require some degree of

chord-wise extension. This can be achieved by utilizing a linkage, a mechanism,

a pivot located outside the aerofoil contour or, perhaps most commonly, by

some form of track. Trailing edge flaps may consist of two or more separate

chord-wise segments, or slats, to give a slotted surface and these often move on

tracts attached to the main wing structure.

The structural design of flaps is similar to that of control surfaces

but it s simpler as there is no requirement for mass balance, the operating

mechanisms normally being irreversible. On large trailing edge flap

components, there is often more than one spar member. Especially when this

assists in reacting the support or operating loading. There may be a bending

stiffness problem in the case of relatively small chord slat segments and full

depth honey combs can be used to deal with this.

55

ATTACHMENT OF LIFTING SURFACES

The joint of the fuselage with the wing is subjected to heavy load inputs and

there is a potential for considerable relative distortion. This distortion is usually

accepted and the wing centre box is built completely into the fuselage.

It is sometimes possible to arrange the wing pick-ups as pivots on the

neutral axis or set them on swinging links. In this case, the relative motion is

allowed to take place and there are no induced stresses. Structural assembly of

the wing to the fuselage is relatively simple.

Fins are usually built integrally with the rear fuselage. This is mainly

due to the different form of loading associated with the geometric asymmetry.

56

FLIGHT CONTROLS

Aircraft flight control surfaces allow a pilot to adjust and control the aircraft's

flight attitude.

Development of an effective set of flight controls was a critical advance in the

development of aircraft.

Main control surfaces

The main control surfaces of an  aircraft are attached to the airframe on hinges

or tracks so they may move and thus deflect the air stream passing over them.

This redirection of the air stream generates an unbalanced force to rotate the

plane about in the required direction.

Rudder

The rudder is typically mounted on the trailing edge of the fin, part of

the empennage. When the pilot pushes the left pedal, the rudder deflects left.

Pushing the right pedal causes the rudder to deflect right. Deflecting the rudder

right pushes the tail left and causes the nose to yaw to the right. Centering the

rudder pedals returns the rudder to neutral and stops the yaw.

57

Ailerons

Ailerons are mounted on the trailing edge of each wing near the wingtips, and

move in opposite directions. When the pilot moves the stick left, or turns the

wheel counter-clockwise, the left aileron goes up and the right aileron goes

down. A raised aileron reduces lift on that wing and a lowered one increases lift,

so moving the stick left causes the left wing to drop and the right wing to rise.

This causes the aircraft to roll to the left and begin to turn to the left. Centering

the stick returns the ailerons to neutral maintaining the bank angle. The aircraft

will continue to turn until opposite aileron motion returns the bank angle to zero

to fly straight.

Elevator

An elevator is mounted on the trailing edge of the horizontal stabilizer on each

side of the fin in the tail. They move up and down together. When the pilot pulls

the stick backward, the elevators go up. Pushing the stick forward causes the

elevators to go down. Raised elevators push down on the tail and cause the nose

to pitch up. This makes the wings fly at a higher angle of attack which generates 58

more lift and more drag. Centering the stick returns the elevators to neutral and

stops the change of pitch. Many aircraft use a stabilator — a moveable

horizontal stabilizer — in place of an elevator. Some aircraft, such use a servo

tab within the elevator surface to aerodynamically move the main surface into

position. The direction of travel of the control tab will thus be in a direction

opposite to the main control surface

Secondary control surfaces

Slats

Slats perform the same function as flaps (that is, they temporarily alter the shape

of the wing to increase lift), but they are attached to the front of the wing instead

of the rear. They are also deployed on takeoff and landing.

As our aircraft has a primary role of short distance takeoff, we have

used slats to enable our aircraft perform both low speed and high speed. Slats are

aerodynamic surfaces on the leading edge of the wings of fixed-wing aircraft

which, when deployed, allow the wing to operate at a higher angle of attack. A

higher coefficient of lift is produced as a product of angle of attack and speed, so

by deploying slats an aircraft can fly more slowly or take off and land in a shorter

distance. They are usually used while landing or performing maneuvers which

take the aircraft close to the stall, but are usually retracted in normal flight to

minimize drag. We have chosen the pilot controllable ventilated powered slat

configuration for our aircraft.

59

Spoilers

Spoilers are plates on the top surface of a wing which can be extended upward into the airflow and disturb the linear airflow. By doing so, the spoiler creates a carefully controlled stall over the portion of the wing behind it, greatly reducing the lift of that wing section.

60

Due to the high landing speeds of our aircraft, we have fitted spoilers on to our aircraft. Thrust

reversers are not practically viable due to their high weight and space requirements. Thus spoilers are

used to slow down the aircraft while landing. A spoiler is a device intended to reduce lift in an aircraft.

Flaps

Flaps are mounted on the trailing edge of each wing on the inboard section of

each wing (near the wing roots). They are deflected down to increase the

effective curvature of the wing. Flaps raise the Maximum Lift Coefficient of the

aircraft and therefore reduce its stalling speed. They are used during low speed,

high angle of attack flight including take-off and descent for landing. Some

aircraft are equipped with "flapperons", which are more commonly called

"inboard ailerons. These devices function primarily as ailerons, but on some

aircraft, will "droop" when the flaps are deployed, thus acting as both a flap and

a roll-control inboard aileron.

61

LANDING GEAR CONFIGURATION

Retractable landing gearTo decrease drag in flight some undercarriages retract into the wings and/or fuselage with wheels flush against the surface or concealed behind doors; this is called retractable gear. Our aircraft is designed to use retractable landing gear.

Fig. nose landing gear

62

Fig. Main landing gear

63

Positioning of under carriage

Tricycle gear describes an aircraft undercarriage, or landing gear, arranged in a tricycle fashion. The tricycle arrangement has one wheel in the front, called the nose wheel, and two or more main wheels slightly aft of the center of gravity. Because of the ease of operating tricycle gear aircraft on the ground, the configuration is the most widely used on aircraft.

Tricycle gear aircraft are easier to land because the attitude required to land on the main gear is the same as that required in the flare, and they are less vulnerable to crosswinds. As a result, the majority of modern aircraft are fitted with tricycle gear. Almost all jet-powered aircraft have been fitted with tricycle landing gear, to avoid the blast of hot, high-speed gases causing damage to the ground surface, in particular runways and taxiways.

Taking these factors into consideration we have incorporated tricycle landing gear pattern.

64

Differential braking

Differential braking depends on asymmetric application of the brakes on the main gear wheels to turn the aircraft. For this, the aircraft must be equipped with separate controls for the right and left brakes (usually on the rudder pedals). The nose or tail wheel usually is not equipped with brakes. Differential braking requires considerable skill. In aircraft with several methods of steering that include differential braking, differential braking may be avoided because of the wear it puts on the braking mechanisms. Differential braking has the advantage of being largely independent of any movement or skidding of the nose or tail wheel. Our aircraft has incorporated differential braking.

Tiller steering

A tiller in an aircraft is a small wheel or lever, sometimes accessible to one pilot and sometimes duplicated for both pilots, that controls the steering of the aircraft while it is on the ground. The tiller may be designed to work in combination with other controls such as the rudder or yoke. In large airliners, for example, the tiller is often used as the sole means of steering during taxi, and then the rudder is used to steer during take-off and landing, so that both aerodynamic control surfaces and the landing gear can be controlled simultaneously when the aircraft is moving at aerodynamic rates of speed. Tiller steering is incorporated in our aircraft for easy taxiing.

65

3 VIEW DIAGRAMS OF THE AIRCRAFT

TOP VIEW

66

SIDE VIEW

FRONT VIEW

67

BIBLIOGRAPHY

REFERENCES

Airplane Design by Dr.Jan Roskam.(3rd edition)

Aircraft Design: conceptual approach by Daniel P.Raymer

Introduction to flight by John D. Anderson.

Aircraft structures by T.H.G.Megson (3rd edition).

Aircraft performance and design by John D.Anderson

WEB REFERENCES:

www.aerospaceweb.org

www.continentalaerospacestechnology.org

www.wikipedia.org

www.airliners.net

www.aiee.com

www.bombardier.com

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