adp 2 final

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.

2

CONTENTS 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

9

MANUVERING LOADS ON AIRCRAFT CONTROL SURFACES

41

10

MATERIAL SELECTION

47

11

DESIGN OF STRUCTURAL COMPONENTS OF THE WING

50

12

FLIGHT CONTROLS

58

13

LANDING GEAR CONFIGURATION

63

14

THREE-VIEW DIAGRAM

66

15

BIBLIOGRAPHY

68 3

INTRODUCTION 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.

4

V-n DIAGRAM FLIGHT ENVELOP 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 5

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 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

6

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 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 8

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. 9

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

10

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

11

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. 12

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) 14

=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 15

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 16

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

17

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.

18

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 19

• Stringers

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

21

• 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.

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.

22

Box Beam or Distributed Flange Layout:

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 23

• 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 tapering

Disadvantages include: • Reduced damage tolerance so that planks are used • Difficult to use on large aircraft panels.

24

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

Types of spars: 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.

25

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: 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 26

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.

27

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

28

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 29

carrying. Floors are needed where there are substantial cut-outs and the skin is stabilized against buckling by the use of frames and bulkheads.

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.

30

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. crackstopping requirements. The frames are usually in direct contact with the skin; stringers pass through them and are seated into place.

31

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 32

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 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)

33

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/m2

34

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, B1 = B9 , B2 = B8 = B10 = B16, B 3 = B 7 = B 11 =B 15, B 4 =B 6 =B 12 =B 14, B5 = B13

35

B1 =

tD b  σ2  t Db  σ1   2 +  , B2 = 2 +  6  σ1  6  σ2 

Substituting, B1 = 100 +

i.e.,

1× 353.43  σ 2  1 × 353.43  σ 16  2+  + 2 +  6 σ1  6 σ1   

B1 = 100 +

1× 353.43  900  2 + × 2 6 831.5  

B1 = 534.72mm 2

Similarly B2 = 534.72mm 2, B3 = 534.72mm 2, B 4 = 534.72mm 2 . 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

σ

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 36

For this section Ixy=0 and My=0 We know, σs =

Mxy I xx

Where, 2 I xx = 2 × 534.72× 900 2+ 4× 534.72× 831.5 2+ 4× 534.72× 636.612+ 4× 534. 72 × 344.41

⇒ I xx = 3.466× 10 9 mm 4

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, Γ = 2π ( −V0 + 2V∞ ) = 2π ( −261.02 + 2 × 222.22) Γ = 1152.46

L = 54.754kN Net Lift onthe fuselage = 54.754 ×1.5 = 82.131kN

Hence the resultant load acting Sy = 82.131 kN

37

The shear flow is given by the expression: qs =

−S y I xx

n

∑B y r =1

r

r

+ q s,0

Substituting the required values we get, n

qs = −2.3696 × 10 −5 ∑ B r y r r =1

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; S xη 0 − S y ξ0 = Ñ ∫ qb p

ds +2 Aqs ,0 t

Rearranging the above equation we get, M b = 2 × A p × (q12 + q 23 + q 34 + .......) + 2× A T× q s ,0

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;

qs ,0 = 3.947 N mm

38

The following tabular column contains the shear flow values over the fuselage. SKIN PANEL Stinger/ Boom Br(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

SHEAR FLOW DIAGRAM 39

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

40

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 41

• Arrested • Braking

Inertia loads • Acceleration • Rotation • Dynamic • Vibration • Flutter Power plants loads • Thrust • Torque • Gyroscope • Vibration • Duct pressure Take off loads • Catapult • Aborted Taxi • Bumps • Turning 42

Other loads • Towing • Jacking • Pressurization • Bird strike • Crash Limit load 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

43

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 factors 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

44

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. 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 unaccelerated 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 45

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 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 46

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 SELECTION 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 47

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" 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. 48

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 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-5Al5V-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 49

than conventional aluminum alloys, and provide higher modulus, and therefore, enable weight savings. Thermally conductive plastics offer significant improvements over conventional plastics.

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 50

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.

RIB LOCATION AND DIRECTION 51

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. 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.

52

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. 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.

53

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

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.

54

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. 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 55

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.

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 56

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. 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 57

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.

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. 58

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.

59

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 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

60

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.

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 61

carefully controlled stall over the portion of the wing behind it, greatly reducing the lift of that wing section.

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.

62

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.

LANDING GEAR CONFIGURATION Retractable landing gear To 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.

63

Fig. nose landing gear

64

Fig. Main landing gear

65

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.

66

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.

67

3 VIEW DIAGRAMS OF THE AIRCRAFT TOP VIEW

68

SIDE VIEW

FRONT VIEW

69

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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