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EIGHT SEATER SHORT RANGE BUSINESS JET AIRCRAFT AN AIRCRAFT DESIGN PROJECT-II REPORT

Submitted by S.VIGNESH (30609101062) P.VASANTHA PRABHU (30609101059) J.SELVA KUMAR (30609101051) N.VIGNESH (30609101061) in partial fulfillm fulfillment for the award of the degree of

BACHELOR OF ENGINEERING in AERONAUTICAL

JEPPIAAR ENGINEERING COLLEGE, CHENNAI 600 119

ANNA UNIVERSITY: CHENNAI 600 025 NOV/DEC 2012

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DEPARTMENT OF AERONAUTICAL ENGINEERING JEPPIAAR ENGINEERING COLLEGE ANNA UNIVERSITY, CHENNAI

BONAFIDE CERTIFICATE Certified that this project report ‘EIGHT SEATER SHORT RANGE BUSINESS JET AIRCRAFT’ is a bonafide work of __S.VIGNESH___________who carried out project under my supervision. Submitted for the examination held on ____6.11.2012________

PROJECT GUIDE

INTERNAL EXAMINER

HEAD OF THE DEPARTMENT

EXTERNAL EXAMINER

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ACKNOWLEDGEMENT It gives us immense pleasure in expressing our sincere gratitude to Honourable Dr.Jeppiaar, M.A., B.L., Ph.D., founder and Chairman of Jeppiaar

Engineering College for bestowing us with an opportunity to bring out this project as a successful one. We are very much grateful to our principal Dr.Susil Lal Das, M.Sc., Ph.D., for their encouragement and moral support. We are very much indebted to Mr.G.Prabakaran (HOD) Aeronautical Department for giving me his able support and encouragement. At this juncture I must emphasis the point that this AIRCRAFT DESIGN PROJECT-II would not have been possible without the highly informative and valuable guidance by our respected preceptor (Ms. Puja Sunil and Ms. Usha Bharathi), Mr. Balaraman whose vast knowledge and experience has must 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.

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ABSTRACT

The aim of this design project is to analysis an 8 Seater Short Range Executive Aircraft by a structural analysis of Shear force and Bending moment. Have to design a more strength aircraft by give the support of stringer, ribs, spar in Wing section and to give the support of stinger, bulkhead, longer in Fuselage. The flying strength of aircraft is analysis by Vn diagram. Then the Design of Miscellaneous Members of Wing Fuel Tank, Rib location and direction, Empennage Design, Auxiliary Surfaces, Wing –Fuselage Intersection, Flutter, Aileron Buzz and Buffeting. Then the necessary graphs have to be plotted for further performance calculation. Required diagrams are also drawn.

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TABLE OF CONTENTS TITLE

1

PAGE

Abstract

i

List of symbols

iv

List of figures

v

List of tables

vi

List of Graphs

vii

Introduction 1.1 Brief review of ADP - 1

1

1.2 Structural Design - Overview

2

2

V-n diagram

4

3

Structural Design of wing

4

3.1 Introduction

10

3.2 Air –Inertia Load Estimation

11

3.3 Shear force & Bending moment Distribution

13

3.4 Material Selection

14

3.5 Wing Spar and Stringer Design

18

3.6 Shear flow Distribution

33

Structural Design of Fuselage 4.1 Design of fuselage

39

4.2 Stringer Design

39

4.3 Shear flow Distribution

43

4.4 Bulkhead design

46

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5

Design of Miscellaneous Members 5.1 Wing Fuel Tank

47

5.2 Rib location and direction

48

5.3 Empennage Design

49

5.4 Auxiliary Surfaces

50

5.5 Wing –Fuselage Intersection

52

5.6 Flutter

53

5.7 Aileron Buzz

54

5.8 Buffeting

55

6

Final design- three view diagram

56

7

Conclusion

57

Bibliography

58

Website reference

59

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LIST OF SYMBOLS A

Total cross sectional area

ft2

ASPAR

Cross sectional area of spar

ft2

Afuselage

Cross sectional area of fuselage

ft2

at

Slope of the CL vs. α curve for a horizontal tail.

Deg-1

a

Distance of the front spar from the nose of the aircraft

ft

b

Distance of the rear spar from the nose of the aircraft

ft

b

Wing span

ft

bw

Width of the web

Ft

bf

Width of the flange

Ft

Cwing

Chord of the actual wing

Ft

Celliptic

Chord of the elliptic wing

Ft

C.G

Centre of gravity

Ft

g

Acceleration due to gravity

ft/s2

D

Drag

Lb

E

Youngĵs modulus

lb/ft2

FOS

Factor of safety

Ftu

Tensile ultimate strength

lb/ft2

H

Height of the C.G from the ground level

Ft

iw

Orientation of wing on fuselage

Deg

Ixx

Second moment of area about X axis

ft4

Lw

Local lift

Lb

WT.O

Takeoff weight

Lb

W/S

Wing loading

lb/ft2

y

span location

Ft

α

Angle of attack

Deg

β

Turnover angle

Deg

ρ

Density

lb/ft3

ρo

Density of air at sea level

lb/ft3

σ

Bending stress

lb/ft2

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LIST OF FIGURES

FIGURE

TITLE

PAGE NO.

1

Load on the aircraft

8

2

Typical V-n diagram

9

3

Final V-n diagram

14

4

Linear lift distribution

16

5

Elliptic Lift Distribution

17

6

Wing Separation Of Finite Section

22

7

Torque Distribution Over an Airfoil

26

8

Shear Center On The Chord

26

9

Wing Spar Arrangement

30

10

Different Spar Selection

31

11

Cross Section of Rear Spar

33

12

Cross Section of Middle Spar

35

13

Semimonocoque And Monocoque

39

14

Cross section Of Z-section

42

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LIST OF TABLES TABLE

TITLE

PAGE NO.

1

Specification of the Aircraft Design Project I

7

2 3 4

Load Factor Velocity VS Load Factor Load on Factor of safety

10 11 13

5

Span VS Linear Lift Distribution

17

6

Span VS Elliptic Lift Distribution

18

7

Span Vs Schrenk's value

18

8

Span Vs Load acting on wing

20

9 10 11 12 13

Centroid table Span VS Shear Force Span VS Bending Moment Span wise VS Shear Force Span wise VS Bending Moment0 Span wise VS Torque at Normal force

22 23 23 25 25

Spanwise VS Mean Aerodynamic Chord

28

16

Frontspar Centroid calculation table

33

17

Front Spar Bending Stress

34

18

Middle Spar Centroid Calculations

35

19

Middle Spar Bending Stress

36

20

Rear Spar Centroid Calculations

37

21

Rear Spar Bending Calculation

38

22 23

Fuselage Structure Analysis Stringer Stress Tabulation

41 44

22

Weight, Moment, Shear Force, Bending Moment

45

14 15

27

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LIST OF GRAPHS

GRAPHS

TITLE

PAGE NO.

1

Span VS Linear Lift Distribution

16

2

Span vs Elliptic lift Distribution

19

3

Span Vs Load acting on wing

20

4

Span Vs Fuel weight distribution

21

5

NET SHEAR FORCE vs BENDING MOMENT

25

6

Span wise VS Bending Moment

28

7

Span vs Net Torque Force

32

8

Stringer location in Fuselage

46

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1 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 8 seater business jet.

The specifications of the Aircraft design project-I as follows: S.No

DESIGN PARAMETER

MAGNITUDE

UNIT

1.

Cruising speed

236.11

m/s

2.

Wing span

12.84

m

3.

Aircraft Length

19.5

m

4.

Wing area

21.84

m2

5.

Height

4.8

m

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6.

Aspect ratio

7.55

(No unit)

7.

Wing loading

585.89

Kg/m2

8.

Empty weight

7,296

Kg

9.

Maximum take-off weight

1,25,568

N

10.

Pay load

1280

kg

11.

No. of engines

2

(No unit)

12.

Thrust power

32.00

Kn

13.

Range

5200

Km

14.

Service ceiling

1,3700

M

15.

Mach no.

0.715

(No unit)

16.

Thrust/weight ratio

0.25

(No unit)

17.

Gliding angle

4.23

o

18.

Seating capacity

8

(No unit)

19.

Fuselage

10.54

M

20.

Take-off distance

1,007.46

M

21.

Landing distance

710.3

M

22.

Rate of climb

1.298 × 10-3

m/s

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LOADS ON THE AIRCRAFT: The structure of an aircraft is required to support two classes of loads, first termed ground loads, includes all loads encountered by the aircraft during movement or transportation on the ground such as taxing, landing loads, towing etc, while the second is the air loads, comprises loads imposed on the structure. The two classes of loads of loads may be still classified as surface forces acting on the surface of the structure and body forces acting over the volume of the structure. Basically all air loads are the resultant of the pressure distribution over the surfaces of the skin produced by steady flight, maneuver or gust conditions. Generally these causes bending, shear, torsion in all parts of the structure in addition to local normal pressure loads imposed on the skin. Ground loads encountered in landing and taxing subject the aircraft to concentrated shock loads through the undercarriage system. The majority of the aircraft have their main undercarriage located in the wings with nose wheel or tail wheel in the vertical plane of symmetry. Clearly the position of the undercarriage should be in such a position so as to produce minimum loads on the wing structure.

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2 ESTIMATION OF V-N DIAGRAM The control of weight in aircraft design is of extreme importance. Increase in weight requires stronger structures to support them, which in turn lead to further increase in weight & so on. Excess of structural weight means lesser amounts of payload, affecting the economic viability of the aircraft. Therefore there is need to reduce aircraft’s weight to the minimum compatible with safety. Thus to ensure general minimum standards of strength & safety, airworthiness regulations lay down several factors which the primary structures of the aircraft must satisfy. These are 1. LIMIT LOAD: the maximum load that the aircraft is expected to experience in normal operation. 2. PROOF LOAD: product of the limit load and proof factor(1.0-1.25) 3. ULTIMATE LOAD : product of limit load and ultimate factor(1.0-1.5) The aircraft’s structure must withstand the proof load without detrimental distortion & should not fail until the ultimate load has been achieved. V-n Diagram: A chart of Velocity versus load factor (V-n diagram) is another way of showing limits of aircraft performance. It shows how much load factor can be safely achieved at different

airspeeds.

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The maneuverability of the aircraft is also dictated by the loads falling on the structures during the maneuvers. Both the aerodynamic and structural limitations for a given airplane are illustrated in the V-n diagram, a plot of load factor versus flight velocity. A V-n diagram is type of flight envelope for the aircraft establishing the maneuver boundaries. The BCAR (British civil airworthiness requirements) has given the basic strength and flight performance limits of various categories of the aircraft. They are listed below

Category

Positive load factor (n+)

Negative load factor(n-)

Normal

3.8

-1.5

Semi aerobatic

4.5

-2

Fully aerobatic

6

-3

Tabular column 2: LOAD FACTOR The 8 seater executive aircraft comes under the normal category. Therefore the load factor limits for the aircraft is 3.8 & -1.5. The V-n diagram for the aircraft is drawn for the two cases namely, 1. Intentional maneuver( pilot induced maneuver ) 2. Unintentional maneuver( gusts) INTENTIONAL MANEUVER: Intentional maneuvers are induced by the pilot during climb, pull up or pull down, banking the plane etc... The load factor is function of velocity. The expression relating the load factor and the velocity is given by nmax =

Where nmax is the maximum load factor, V is the speed of the aircraft, Vs is the stalling speed of the aircraft. The stalling speed of the aircraft Vs 2 =

Vs= 59.197 m/s

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For various values of V, nmax is calculated and tabulated below,

V

Nmax=

V

Nmax=

4

118.394

-0.25

29.598

3.8

115.396

-0.5

41.858

3.5

110.74

-0.75

51.256

3

102.53 -1

59.197

2.5

93.599

2

83.71

1.5

72.501

1

59.197

Tab3. Velocity VS Load Factor

The cruising speed of the aircraft is 236.11 m/s. The dive speed of the aircraft is the maximum speed of the aircraft. The dive speed is equal to the sum of the cruising speed and 60 knots. VD = 236.11 + 60 knots = 236.11 +30.56 m/s = 266.67 m/s GUSTS: The movement of air in turbulence is known as gusts. It produces changes in wing incidence, thereby subjecting the aircraft to sudden or gradual increases or decreases in lift from which normal accelerations result. These may be critical for large, high speed aircraft and may possibly cause higher loads than control initiated maneuvers. Thus in the gust analysis, the change in load factor due to the gust is calculated. The BCAR has given standard gust velocities for stall, cruise, dive speeds as 66, 50, 25 ft/s

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respectively. The small change in load factor ∆n due to the gust is calculated by assuming a sharp gust. The change in load factor ∆n = Where ρ è density at cruising altitude, a è lift slope, in radians U è gust velocity in m/s V è velocity of the aircraft in m/s W/S è wing loading in N/m2 In the above formula, gusts are assumed to be sharp but it is usually graded, hence a relief factor called gust alleviation factor K is introduced in the term. The value of the K is obtained from the book “AIRPLANE AERODYNAMICS AND PERFORMANCE” by JAN ROSKAM Where K =

,

!" µ =#$%$&

%$'(

Where ρ is the density, C is the mean aerodynamic chord, g is the acceleration due to gravity; CLα is the slope lift coefficient. The CLα (corrected for aspect ratio) is 0.0962/deg.

µ =

$ $)$ *

µ=

$+,+

,$-.$.+$.$+

µ = 50.36

K= =

$/ /

Κ = 0.796

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K = Therefore ∆n = 0

123 # !"

For STALL SPEED V= 59.197 m/s, U= 20m/s ∆n = 0

=

# 123 !"

+-/$,$.$+$$-.-+ $+,+

∆n = 1.062 For CRUISE SPEED V=236.11 m/s, U= 15m/s ∆n = 0

=

# 123 !"

+-/$,$.$+$.$/.. $+,+

∆n = 2.725 For DIVE SPEED V= 266.67 m/s, U= 7.5 m/s ∆n = 0

=

# 123 !"

+-/$,$.$+$+$///+

∆n ∆ = 1.794

$+,+

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V

1+∆ ∆n

1-∆ ∆n

59.197

2.062

-0.062

236.11

3.725

-1.725

266.67

2.794

-1.794

Tab 4.load Factor of safety

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3 STRUCTURAL ANALYSIS AND DESIGN OF WINGS WING STRUCTURAL ANALYSIS: The structural design of the wing requires a complete quantitative knowledge of the different loads it will be subjected to during its flight regime. These loads can be briefly classified as 1. Distributed loads - Loads such as aerodynamic loads, weight of the wing and weight of fuel. 2. Concentrated loads – Loads such as thrust, engine weight, landing gear weight and armament weight. LOADS ACTING ON WING: As both the wings are symmetric, let us consider the starboard wing at first. There are three primary loads acting on a wing structure in transverse direction which can cause considerable shear forces and bending moments on it. They are as follows: v Lift force (given by Schrenk’s curve) v Self-weight of the wing v Weight of the power plant v Weight of the fuel in the wing SCHRENK’S CURVE: Lift is a component of the resultant aerodynamic force acting at the centre of pressure of an aerodynamic chord, along a direction perpendicular to the direction of the relative wind. At a particular altitude and at a specific angle of attack, Lift varies along the wing span due to the variation in chord length along the span. Schrenks curve defines this lift distribution over the wing span of an aircraft. Since the wings of an aircraft are symmetrical about the longitudinal axis, the Schrenks curve for the starboard wing alone can be obtained at first. This is given by y=

45 4

where y1 è linear variation of lift along the wing semi-span y2 è equivalent elliptic lift distribution along the wing semi-span

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TO FIND y1: Lift force is found along the line joining the aerodynamic centers of chords along the wing span. Hence, the wing is rotated about the wing root so that the line joining the aerodynamic centers becomes the horizontal line. 6 a= ) 789:

a=

.,

789..; .<

)

=6.55 Lift per unit length at wing root

= CL×0.5×ρ×V2×CR = 0.23884×0.5×1.4×236.1112×2.55 = 23766.98 N/m

Lift per unit length at wing tip

= CL×0.5×ρ×V2× Ct =0.23884×0.5×1.4×236.1112×0.84 = 7829.12 N/m

6.55 m Fig.4 Linear lift distribution Area under trapezoid life distribution = 155673.719 Equation of linear lift distribution for starboard wing Y1 = -1195.289x + 23766.98 Equation of linear lift distribution for port wing we have to replace x by –x in general, Y1 = 1195.289x + 23766.98

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X

Y1

0

23766.98

1

22571.7

2

21376.42

3

20181.14

4

18985.86

5

17790.58

6

16595.3

6.55

15937.89 Tab5. Span VS Linear Lift Distribution

LINEAR LIFT DISTRIBUTION

25000 20000 15000 Series1

10000

Series2 5000 0 0

1

2

3

4

5

6

6.55

SPAN (a)

Graph.4 Span VS Linear Lift Distribution

Elliptic Lift Distribution: Twice the area under the curve or line will give the lift which will be required to overcome weight Considering an elliptic lift distribution we get =>6?

L/2 = W/2 = , Where b1 is Actual lift at root A= And a is wing semi span

=>6? ,

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Lift at Tip b = 15138.35 N/m

15138.35 N/m

6.55 m Fig 5. Elliptic Lift Distribution

Y2 = 1155.60@A BCD E F

x

Y2

0

15138.317

1

14960.852

2

14415.354

3

13457.142

4

11987.651

5

9779.05

6

6072.292

6.55

0 Tab 6: Span VS Elliptic Lift Distribution

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Construction of Schrenk”s Curve: Schrenk”s Curve is given by

Y1+Y2 =

55BGHBIDJKKBH55GGKCLABCDMF

Y1+Y2 = -597.64 x + 45489.019 + 1155.60 @A BCD E F

x

Y1+Y2

0

19452.65

1

18766.28

2

17895.89

3

16819.14

4

15486.76

5

13784.82

6

11333.796

6.55

7968.95 Tab 7:SpanVsSchrenk's value

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Graph 2: Span vs Elliptic lift Distribution

Load Estimation on wings Description: The solution methods which follow Euler’s beam bending theory (σ/y=M/I=E/R) use the bending moment values to determine the stresses developed at a particular section of the beam due to the combination of aerodynamic and structural loads in the transverse direction. Most engineering solution methods for structural mechanics problems (both exact and approximate methods) use the shear force and bending moment equations to determine the deflection and slope at a particular section of the beam. Therefore, these equations are to be obtained as analytical expressions in terms of span wise location. The bending moment produced here is about the longitudinal (x) axis. Loads acting on wing: As both the wings are symmetric, let us consider the starboard wing at first. There are three primary loads acting on a wing structure in transverse direction which can cause considerable shear forces and bending moments on it. They are as follows: ª Lift force (given by Schrenk’s curve) ª Self-weight of the wing ª Weight of the power plant ª Weight of the fuel in the wing

Self-Weight (y3): Self-weight of the wing, Wwing = 5548.06 kg ×9.81 = 54426.46 N

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Wport wing =-27213.23 N Wstar board wing = -27213.23 N / / PQM/R N O =N

y3= -290.52 (x-6.55)2 x

y

0

-12464.07

1

-8948.74

2

-6014.49

3

-3661.28

4

-1889.11

5

-697.97

6

87.88

6.55

0 Tab 8: SpanVs Load acting on wing

Graph 3: Span Vs Load acting on wing

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Power Plant: According to our design data, Our Aircraft power plant is attach to rear fuselage. So, power plant calculation won’t be calculated.

Fuel Weight Distribution: Wf = 23215.55 Consider as equation, yf = 1902.91x-12464.07

2

-8658.25

3

-6755.35

4

-4852.43

5

-2949.52

Fuel weight Distribution 0 -1000

2

3

4

5

6

-2000 -3000 -4000 -5000 -6000 -7000 -8000 -9000 -10000

Graph 4. Span Vs Fuel weight distribution

6.55

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SHEAR FORCE & BENDING MOMENT SHEAR FORCE AND BENDING MOMENT DIAGRAMS OF A WING DUE TO LOADS IN TRANSVERSE DIRECTION AT CRUISE CONDITION: The solution methods which follow Euler's beam bending theory (σ/y=M/I=E/R) use the bending moment values to determine the stresses developed at a particular section of the beam due to the combination of aerodynamic and structural loads in the transverse direction. Most engineering solution methods for structural mechanics problems (both exact and approximate methods) use the shear force and bending moment equations to determine the deflection and slope at a particular section of the beam. Therefore, these equations are to be obtained as analytical expressions in terms of span wise location. The bending moment produced here is about the longitudinal (x) axis. As both the wings are symmetric, let us consider the starboard wing at first. There are three primary loads acting on a wing structure in transverse direction which can cause considerable shear forces and bending moments on it. They are as follows: → Lift force (given by Schrenk's curve) → Self-weight of the wing → Weight of the powerplant S.No.

Curve/Component

Area/Structural weight (N)

Centroid

1

Y1

155673.719

5.458

2

Y2

77896.859

2.781

3

Wing

54426.46

1.637

4

Fuel

23215.55

1.31

Tab 9: Centroid table

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∑V=0 77836.859-54426.46-23215.55-VA = 0 VA = 194.85 ∑MA = 0 MA+ (54426.465×1.637)+(23215.55×1.31)-(155673.719×5.458)-(77836.859×2.781) MA – 946622.97 = 0 MA= 946622.97

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SHEAR FORCE T

1

F

FD

U -290.52 T

KGG

D

S.F1 = -298.84 x2 + 45489.019 x + 577.8 @A BCD E S - 42.903 Sin-

V A BCF V K GGF U -194.85

SPAN

Shear Force

0

-19625.79

1

-8113.17

2

1496.55

3

9203.37

4

15007.29

5

18908.31

6

20906.43

6.55

21194.28 Tab 10: Span VS Shear Force

BENDING MOMENT:

F

B.M = -199.21 x3 + 15163.006 x2 + 288.9 [x (x@A BCD E S ) + 42.903 Sin -1KGG] + 385.2 FA

(42.903- x2)1.5-290.52 ( + 21.45 x2- 2.18 x3) + 94662.97 5

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SPAN

BENDING MOMENT

0

754870.44

1

1043525.353

2

1062831.914

3

1084364.515

4

1108829.591

5

1318721.935

6

1381469.212

6.55

943585.515 Tab 11 : Span VS Bending Moment

NET SHEAR FORCE AND BENDING MOMENT DIAGRAM:

1600000

BENDING MOMENT

1400000 1200000 1000000 800000 600000 400000

SHEAR FORCE

200000 0 -8

-6

-4

-2

0

2

4

6

8

WING SPAN

Graph.5 NET SHEAR FORCE vs BENDING MOMENT

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Shear force and bending moment diagrams due to loads along chordwise direction at cruise condition: Aerodynamic center- This is a point on the chord of an airfoil section where the bending moment due to the components of resultant aerodynamic force (Lift and Drag) is constant irrespective of the angle of attack. Hence the forces are transferred to this point for obtaining constant Ma.c Shear center- This is a point on the airfoil section where if a force acts, it produces only bending and no twisting. Hence the force is transferred to this point and the torque is found. Angle of Attack (max) = 15.00,

Angle of Attack (Zero lift) = -1.0

Cruise CL = 1.40

V = 236.11 m/s

ρ = 0.23884 kg/m3

CD = 0.0025

SHEAR FORCE BENDING MOMENT: Co-efficient of force at normal direction Cn = CL cos ά + CD sin ά = 1.398 Co-efficient of force at Chordwise direction CC = CL sin ά + CD cos ά = 0.026 Therefore, Force per unit length= Cc×0.5×ρ×V2×C Force at Cr = 441.39 N/m Force at Ct = 145.39 N/m For Linear, y = 23.05 x + 145.39 à1 Shear Force: Integrate Eqn. 1 5HA

NC

W 55 GGFD V 5AG DBF

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Span wise

Shear force

0

0

1

156.915

2

336.88

3

539.895

4

765.96

5

1015.07

6

1287.24

7

1582.45

8

1900.72

9

2242.03

10

2606.40

11

2993.81

12

3404.28

12.84

3766.88 Tab 12: Span wise VS Shear Force

BENDING MOMENT: .,

X

W3.841 x3+72.69x2

Span wise

Bending moment

0

0

1

76.531

2

176.108

3

321.77

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4

536.584

5

843.575

6

1265.796

7

1826.293

8

2548.112

9

3454.299

10

4567.9

11

5911.961

12

7509.528

12.84

9064.425 Tab 13: Span wise VS Bending Moment

14000

BENDING MOMENT & SHEAR FORCE BENDING MOMENT

12000 10000 8000 6000

SHEAR FORCE

4000 2000 0 CHORDWISE

Graph 6. Span wise VS Bending Moment

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TORQUE DUE TO NORMAL FORCES AND CONSTANT PITCHING MOMENT AT CRUISE CONDITION: Aerodynamic center- This is a point on the chord of an airfoil section where the bending moment due to the components of resultant aerodynamic force (Lift and Drag) is constant irrespective of the angle of attack. Hence the forces are transferred to this point for obtaining constant Ma.c Shear center- This is a point on the airfoil section where if a force acts, it produces only bending and no twisting. Hence the force is transferred to this point and the torque is found

The lift and drag forces produce a moment on the surface of cross-section of the wing, otherwise called a torque, about the shear center. Moment about the aerodynamic center gets transferred to the shear center. The shear center on the chord under which it is locates.

Cruise condition (Normal Force) T= ½ Cn ρ V2 c × 0.034 C

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= 1325.51 C2 Where, C à chord the equation for chord can also be represented in terms of x by taking C = mx +k,

C = 0.264 x +2.55 Therefore, Torque = 30.751x3 + 8418.46 x + 892.33 x2

Span wise

Torque at Normal force

0

0

1

9541.54

2

21052.24

3

34716.62

4

50719.18

5

69244.42

6

90476.85

7

114600.98

8

141801.31

9

172262.34

10

206168.6

11

243704.57

12

285054.76

12.84

322871.54

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Tab 14:Span wise VS Torque at Normal force TORQUE DUE TO CHORDWISE FORCE: Torque per unit length T2 = FC×0 T2 = 0 TORQUE DUE TO MEAN AERODYNAMIC CHORD: Torque due to Ma.c = YZ[\ ×0.5×ρ×V2×C×C T3 = -3347.9 ×C2 T3 = -77.671 x3- 21768.04 x- 2253.806 X2 Span wise

MEAN AERODYNAMIC CHORD

0

0

1

-24099.51

2

-53172.67

3

-87685.49

4

-128104.0

5

-174894.22

6

-228522.19

7

-289453.92

8

-358155.45

9

-435092.80

10

-520732.01

11

-615539.06

12

-719980.03

12.84

-815496.45

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Tab 15: Spanwise VS Mean Aerodynamic Chord NET TORQUE: Then the different torque components are brought together in a same graph to make a comparison The net torque will be sum of all the above torques (i.e.) torques due to normal force, chordwise force, powerplant and aerodynamic moment.

400000 200000 0 1 -200000

2

3

4

5

6

7

8

9 10 11 12 13 14 Series1 Series2

-400000

Series3

-600000 -800000 -1000000

Graph.7 Span vs Net Torque Force

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LOAD ESTIMATION OF WINGS WING STRUCTURAL LAYOUT Specific Roles of Wing (Mainwing) Structure: The specified structural roles of the wing (or main plane) are: Ø 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. Ø To react against: Landing loads at attachment points Loads from pylons/stores Wing drag and thrust loads Ø To provide: Fuel tank age space Torsional rigidity to satisfy stiffness and aero-elastic requirements. Ø 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 v Form the main span wise beam v Transmit bending and torsional loads v Produce a closed-cell structure to provide resistance to torsion, shear and tension loads. In particular: v Webs – resist shear and torsional loads and help to stabilize the skin. v Flanges - resist the compressive loads caused by wing bending. SKIN v To form impermeable aerodynamics surface v Transmit aerodynamic forces to ribs & stringers v Resist shear torsion loads (with spar webs). v React axial bending loads (with stringers). STRINGERS v Increase skin panel buckling strength by dividing into smaller length sections. v React axial bending loads RIBS v Maintain the aerodynamic shape v Act along with the skin to resist the distributed aerodynamic pressure loads v Distribute concentrated loads into the structure & redistribute stress around any discontinuities v Increase the column buckling strength of the stringers through end restraint v Increase the skin panel buckling strength. SPAR DEFINITION: The maximum bending moment from previous section was found to be as 2897784.51 Nm. Therefore we define 3 Spars with front spar at 15% of chord, middle spar at 45% of chord and rear spar at 70% of chord. The position of the three spars from the leading edge of the root chord is given below as follows:

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Front spar - 15% of chord = 2.442 m Middle spar - 45% of chord = 7.326 m Rear spar - 70% of chord = 11.396 m Bending moment M = Max BM * FOS * n = 2897784.51 × 1.5 × 3.8 = 16517371.71Nm The Structural load bearing members in the wing are the Spars and Stringers. The bending moment carried by the Spars is 70% and that of Stringers is 30% of the total Bending Moment. Bending Moment taken by Spars is = 0.7 x 16517371.71 = 11562160.19 Nm The cross section of the spar chosen here is an I-section For each spar we are determining the following parameters: A. B. C. D.

Centroid Moment of Inertia Bending Moment Bending Stress

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FRONT SPAR: ª Height of the spar = 38 cm ª Breadth of the spar = 16 cm ª Thickness of the spar = 4.5 cm

Fig 11: Cross Section of Rear Spar To find out the centroid, the following calculations are made: Element

Area(A) (cm2 )

x (cm)

y (cm)

Ax (cm3 )

Ay (cm3 )

Ax2 (cm4 )

Ay2 (cm4 )

Icx (cm4 )

Icy (cm4 )

1

72

8

2.25

576

162

4608

364.5

121.5

1536

2

130.5

8

19

1044

2479.5

8352

47110.5

9145.8

220.22

3

72

8

35.75

576

2574

4608

92020.5

121.5

1536

Total

274.5

2196

5215.5

17568

139495.5

9388.87

3292.22

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Front Spar Calculations: Centroid = X =

]^_ ]^

= 8 cm;

Y=

]^` ]^

= 19 cm

I xx = Σ Icx + ΣAy2 – ΣAY2 I xx = (9388.87) + (139495.5) – (274.5)(19)2 I xx = 49789.88 cm4

I yy = Σ Icy+ ΣAx2 – ΣAX2 I yy = (3292.22) + (17568) – (274.5) (8)2 I yy = 3292.22 cm4 The FRONT SPAR carries 35 % of the BM carried by the Spars. Thus, Front spar BM = 0.35 x 1156216019 N-cm

= 404675606.7 N cm Front Spar Bending Stress: a Bending Stress, σ z = c by

POINTS

COORDINATES (y) (cm)

BENDING STRESS (N/cm2)

A

19

154425.68

B

14.5

117851.18

C

14.5

117851.18

D

-14.5

-117851.18

E

-14.5

-117851.18

F

-19

-154425.68

The bending stress at various points whose co-ordinates are determined with centroid as the origin are calculated from above formula and tabulated.

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MIDDLE SPAR: ª Height of the spar = 41.6 cm ª Breadth of the spar = 18 cm ª Thickness of the spar = 5 cm

Fig 12: Cross Section of Middle Spar

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To find out the centroid, the following calculations are made: Element

Area(A) (cm2 )

x (cm)

y (cm)

Ax (cm3 )

Ay (cm3 )

Ax2 (cm4 )

Ay2 (cm4 )

Icx (cm4 )

Icy (cm4 )

1

90

9

2.5

810

225

7290

562.5

187.5

2430

2

158

9

20.8

1422

3286.4

12798

68357.12

13147.7

329.17

3

90

9

39.1

810

3519

7290

137592.9

187.5

2430

Total

338

3042

7030.4

27378

206512.5

13522.7 5189.17

Middle Spar Calculations Centroid = X =

]^_ ]^

= 9 cm;

Y=

]^` ]^

= 20.8 cm

I xx = Σ Icx + ΣAy2 – ΣAY2 I xx = (13522.7) + (206512.5) – (338) (20.8)2 I xx = 60467.7 cm4

I yy = Σ Icy+ ΣAx2 – ΣAX2 I yy = (5189.17) + (27378) – (338) (9)2 I yy = 5189.17 cm4 The bending moment carried by the middle spar is 40% of the total bending moment carried by the spars. Middle Spar BM = 462486407.6 N-cm Bending Stress, σ z =

d e

y

P a g e | 46 [email protected]

POINTS

COORDINATES (y) (cm)

BENDING STRESS (N/cm2)

A

20.8

159088.52

B

15.8

120846.09

C

15.8

120846.09

D

-15.8

-120846.09

E

-15.8

-120846.09

F

-20.8

-159088.52

The bending stress at various points whose co co-ordinates ordinates are determined with centroid as the origin are calculated from above formula and tabulated. REAR SPAR ª Height of the spar = 17.72 cm ª Breadth of the spar = 7.6 cm ª Thickness of the spar = 2.5 cm

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To find out the centroid, the following calculations are made: Element

Area(A) (cm2 )

x (cm)

y (cm)

Ax (cm3 )

Ay (cm3 )

Ax2 (cm4 )

Ay2 (cm4 )

Icx (cm4 )

Icy (cm4 )

1

19

3.8

1.25

72.2

23.75

274.36

29.6875

9.896

91.45

2

31.8

3.8

8.86

120.84

281.748

459.19

2496.287

428.76

16.56

3

19

3.8

16.47

72.2

312.93

274.36

5153.957

9.896

91.45

Total

69.8

265.24

618.428

1007.9

7679.932

448.552

199.46

Rear Spar Calculations Centroid = X =

]^_ ]^

= 3.8 cm;

Y=

]^` ]^

= 8.86 cm

I xx = Σ Icx + ΣAy2 – ΣAY2 I xx = (448.552) + (7679.932) – (69.8) (8.86)2 I xx = 2649.184 cm4

I yy = Σ Icy+ ΣAx2 – ΣAX2 I yy = (199.46) + (1007.9) – (69.8) (3.8)2 I yy = 199.46 cm4 Rear Spar carries 25 % of the spar Bending Moments. Bending Moment = 289054004.8 N-cm d

Bending Stress, σ z = e y

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The bending stresses at various points are obtained as: Rear Spar Bending Stress POINTS

COORDINATES (y) (cm)

BENDING STRESS (N/cm2)

A

8.86

966719.74

B

6.36

693943.29

C

6.36

693943.29

D

-6.36

-693943.29

E

-6.36

-693943.29

F

-8.86

-966719.74

The bending stress at various points whose co-ordinates are determined with Centroid as the origin are calculated from above formula and tabulated.

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4 STRUCTURAL ANALYSIS AND DESIGN OF FUSELAGE FUSELAGE STRUCTURAL LAYOUT: The fuselage is the main structure, or body, of the aircraft. It provides space for personnel, cargo, controls, and most of the accessories. The power plant, wings, stabilizers, and landing gear are attached to it. There are two general types of fuselage construction construction—welded welded steel truss and monocoque designs. The welded steel truss was used in smaller Navy aircraft, and it is still being used in some helicopters. onocoque design relies largely on the strength of the skin, or covering, to carry The monocoque various loads. The monocoque design may be divided into three classes - monocoque, semimonocoque and reinforced shell.

Fig 13:Semimonocoque And Monocoque

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v The true monocoque construction uses formers, frame assemblies, and bulkheads to give shape to the fuselage. However, the skin carries the primary stresses. Since no bracing members are present, the skin must be strong enough to keep the fuselage rigid. The biggest problem in monocoque construction is maintaining enough strength while keeping the weight within limits. v Semimonocoque design overcomes the strength-to-weight problem of monocoque construction. In addition to having formers, frame assemblies, and bulkheads, the semimonocoque construction has the skin reinforced by longitudinal members. v The reinforced shell has the skin reinforced by a complete framework of structural members. Different portions of the same fuselage may belong to any one of the three classes. Most are considered to be of semimonocoque-type construction. The semimonocoque fuselage is constructed primarily of aluminum alloy, although steel and titanium are found in high-temperature areas. Primary bending loads are taken by the longerons, which usually extend across several points of support. The longerons are supplemented by other longitudinal members known as stringers. Stringers are more numerous and lightweight than longerons. The vertical structural members are referred to as bulkheads, frames, and formers. The heavier vertical members are located at intervals to allow for concentrated loads. These members are also found at points where fittings are used to attach other units, such as the wings and stabilizers. The stringers are smaller and lighter than longerons and serve as fill-ins. They have some rigidity but are chiefly used for giving shape and for attachment of skin. The strong, heavy longerons hold the bulkheads and formers. The bulkheads and formers hold the stringers. All of these join together to form a rigid fuselage framework. Stringers and longerons prevent tension and compression stresses from bending the fuselage. The skin is attached to the longerons, bulkheads, and other structural members and carries part of the load. The fuselage skin thickness varies with the load carried and the stresses sustained at particular location. There are a number of advantages in using the semimonocoque fuselage. v The bulkhead, frames, stringers, and longerons aid in the design and construction of a streamlined fuselage. They add to the strength and rigidity of the structure. v The main advantage of the semimonocoque construction is that it depends on many structural members for strength and rigidity. Because of its stressed skin construction, a semimonocoque fuselage can withstand damage and still be strong enough to hold together.

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Loads and its distribution: To find out the loads and their distribution, consider the different cases. The main components of the fuselage loading diagram are: v Weight of the fuselage v Engine weight v Weight of the horizontal and vertical stabilizers v Tail lift v Weight of crew, payload and landing gear v Systems, equipment, accessories Symmetric flight condition, steady and level flight: (Downward forces negative) Values for the different component weights are obtained from aerodynamic design calculations. S.No.

Equipment and Component

Length from Ref. point

Weight (N)

Moment

Shear Force

Bending Moment

1

Nose

0.58

3933.81

2281.609

3933.81

2281.809

2

Pilot(2)

2.02

1962.00

3963.24

5895.81

6244.84

3

Cockpit

2.62

4944.24

12953.908

10840.05

19198.757

4

Wing

6.62

54426.46

360303.16

65266.51

379501.91

5

Passenger(3)

6.63

3237.30

21463.29

68503.81

400965.2

6

Passenger(3)

7.92

3237.30

25639.41

71741.11

426604.61

7

Passenger(2)

9.85

2060.10

20291.98

73801.21

446896.59

8

Crew(2)

10.32

2060.10

21260.23

75861.31

468156.82

9

Weight of Fuselage Sheet

10.50

2000.68

21007.14

77861.99

489163.96

10

Power Plant

14.23

3937.83

56036.74

81799.92

545200.72

11

Horizontal Tail

17.10

13232.807

226280.9

95032.72

771481.607

12

Vertical tail

17.60

8971.34

157895.58

104004.06

929377.18

SHEAR FORCE AND BENDING MOMENT DIAGRAM

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To determine the shear force and bending moment diagram for the wing we assume that the wing is a cantilever beam with the root end fixed while the tail end is free. For a cantilever beam the shear force is a given by, Shear Force = Rx fQ g Bending Moment = Tabulation for the values of shear force and bending moment at various positions along the span is as follows.

Graph 9. Length from Ref. point Vs Shear Force 120000

Shear force

100000 80000 60000 40000 20000 0 0.58 2.02 2.62 6.62 6.63 7.92 9.85 10.32 10.5 14.23 17.1 17.6 Length from Ref. point

Bending moment

Graph 10. Lengtth from Ref. point Vs Bending Moment 1000000 900000 800000 700000 600000 500000 400000 300000 200000 100000 0 0.58 2.02 2.62 6.62 6.63 7.92 9.85 10.32 10.5 14.23 17.1 17.6 Leght from Ref. point

FUSELAGE STRUCTURAL ANALYSIS

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Structural analysis of fuselage like that of wing is of prime importance while designing an aircraft. As the fuselage is the one which houses the pilot, the power plant and also part of the payload its structural integrity is a matter of concern. While analyzing the fuselage structure the section must be idealized. Idealization involves the conversion of a stringer and its accompanying skin thickness into a concentrated mass known as a boom. The shear flow analysis of the fuselage simulating flight conditions is shown below. X (m) 1.005

Y (m) 0

0.985

0.26

0.88

0.48

0.72

0.72

0.48

0.88

0.26

0.985

0

1.005

-0.26

0.985

-0.48

0.88

-0.72

0.72

-0.88

0.48

-0.985

0.26

-1.005

0

-0.985

-0.26

-0.88

-0.48

-0.72

-0.72

-0.48

-0.88

-0.26

-0.985

0

-1.005

0.26

-0.985

0.48

-0.88

0.72

-0.72

0.88

-0.48

0.985

-0.26

1.005

0

P a g e | 54 [email protected]

Stringer location in Fuselage 1.5

1

0.5

0 -1.5

-1

-0.5

0

0.5

-0.5

-1

-1.5

The stringer used is of Z type. The following are its dimensions Cross sectional area of each stringer is 100mm2 Fig 14: Cross section of Z Z-section

1

1.5

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The above stringer section is uniformly used throughout the fuselage as shown above in order to provide the fuselage the required load carrying capacity. The diagram showed adjacent is of the idealized fuselage structure. The idealization process is carried out in the following way. STRESS ANALYSIS: IDEALIZATION: The boom 1 is given by

where, B1 èArea of Boom 1 tD èThickness of skin panel b è Circumferential distance between 2 stringers

By Symmetry, B1 = B9, B2 = B8, B10 = B16, B3 = B7 , B11 = B15, B4 = B6 = Bl2 = B14 ,B5 = B13 .h

B1=100+ (0.65×1.37×

/

) [2+

,..

] + (0.65×1.37×

.h /

)[2+

,..

]

=815582.12 Similarly for boom 2 ,

B2 = 815582.12 mm2 Similarly B3 = 815582.12 mm2, B4 =815582.12 mm2. We note that stringers 5 and 13 lie on the neutral axis of the section and are therefore unstressed; the calculation of boom areas B5 and B13 does not then arise. Thus, we have B1:B16 = 815582.12 mm2

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We know that, Ixx = By2

c__? = 24.67 m4; c__g = 13.77 m4; c__R = 6.12 m4; c__i = 1.11 m4 Maximum bending moment = 2897784.51 Nm Hence the Bending moment acting on the fuselage M = Max.B.M × n× FOS =2897784.51 × 3.8×1.5 =16517371.71 Nm Ixx = 24.67 m4 The value of stress acting is given by the expression: =

./.++.+.$` ,/+

STRINGER/BOOM

Y (m)

STRESS x 10 6 (Nm-2)

1

5.5

3.68

2, 16

4.11

2.75

3, 15

2.74

1.83

4, 14

1.37

0.9

5, 13

0

0

6, 12

-1.37

-0.9

7, 11

-2.74

-1.83

8, 10

-4.11

-2.75

9

-5.5

-3.68

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5 Design of Miscellaneous Members Wing fuselage intersection The 8 seater business jet aircraft has low wing configuration, thus the entire wing structure continues in the way of airplane body. Four pin design concept: This concept is adopted as it is the most simple and straight forward method used in Horizon 4000 transport, during 1950s. The lift and moment loads can be carried between the wing and fuselage by simple shear on the four pins. The drag and thrust is taken by breather web. This design allows the wing spar and fuselage bulkheads to deflect independently of each other such that no spar moment is directly transferred to the bulkheads.

The wing-body juncture produces aerodynamic interference which in turn promotes flow separation with its attendant higher drag and unsteady buffeting. This adverse pressure gradient and consequent flow separation can be minimized using contoured surface called fillet. Engine mount An engine mount is a frame that supports the engine and holds it to the fuselage or nacelle. Usually it is made of built up sheet metal, welded steel tubing. The turbofan engine, “HONEY WELL TFE731-20” is wing mounted. A typical turbofan engine installation for a low wing aircraft configuration similar to that of this aircraft is shown below,

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The pylon has three spars (longerons) – Upper, middle and lower- and three major bulkheads, and is attached to the wing at four primary points. These are two mid mid--spar fittings, an upper link and a diagonal brace (drag strut). The attachment pins are secured with “fuse” bolts which are hollow carbon steel devices that have been heat treated to shear fail at a definite load. In the landing break way condition (wheels (wheels-up landing); anding); the sequence is designed to fail the upper and lower links so that the pylon rotates around the mid mid-spar spar and upward. The wing pylon design provides considerable load path redundancy such that an upper link can fail, partially or completely, and there ere is an alternate path path- lower diagonal brace. The below figure shows the engine mounts.

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Empennage Design Horizontal Stabilizer: The horizontal tail of the aircraft is conventional and consists of a fixed tail box. The horizontal stabilizer is usually a two spar structure consisting of a Centre structural box section and two outer sections. The stabilizer assembly is interchangeable (symmetrical airfoil section) as a unit at the fuselage attach points and the outer sections are interchangeable at the attachment to the center box. The two basic horizontal stabilizer box constructions for modern transports are 1. Box constructions with spars, closer light rib spacing (usually less than 10 inches) and surface (may be tapered skins) without stringer reinforcement. The feature of this design is the low manufacturing cost and high torsional stiffness require by the flutter analysis.

2. Box construction with spar stronger ribs and surface skins with stringer reinforcements (skin-stringer or integrally stiffened panels) is a lighter weight structure.

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Vertical Stabilizer: The structural design of the vertical stabilizer is essentially the same as for the horizontal stabilizer is essentially the same as for horizontal stabilizers. The vertical stabilizer box is a two or multi spar structure (general aviation airplanes usually use single spar design) with cover panels (with or without ribs). The root of the box is terminated at the aft fuselage conjuncture with fittings or splices.

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WING FUEL TANKS: In addition to providing the required strength and stiffness, the structural box almost always has to provide fuel space. Integral tanks, as opposed to separate internally supported types, are preferred since their use enables the maximum advantage to be taken of the available volume. Integrally machined or moulded constructions, which use a small amount of large components, are obviously an advantage since sealing is reduced to a minimum. The major problem occurs at tank end ribs, particularly in the corners of the spar web and skins, and at lower surface access panels. The corner difficulty is overcome by using special “suitcase” corner fittings. Access panels should be large enough for a person to get through so that the inside can be inspected and resealed if necessary. On shallow section wings, the access has to be in the lower surface so that the operator can work in an acceptable way even if the depth is insufficient to climb in completely. Apart from the sealing problems, lower surface access panels are in what is primarily a tension skin and so introduce stress concentrations in an area where crack propagation is a major consideration. The access panels are arranged in a span-wise line so the edge reinforcing can be continuous and minimum stress concentration due to the cut-outs. Access panels are often designed to carry only shear and pressure loads, the wing bending being reacted by the edge reinforcing members. A deep wing can avoid these problems by using upper surface access panels but this is not a preferred aerodynamic solution. 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: v 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. v The control hinge loads and the resulting shear forces and bending moments should be equalized as far as is possible. v Structural failure of a single hinge should be tolerated unless each hinge is of fail-safe design and can tolerate cracking one load path.

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These points suggest the use of a relatively large number of discrete hinges but there are difficulties associated with this solution there are the obvious loads likely to be induced in the control by the distortion under load of the main surface to which it is attached may be significant. These problems do not arise if only two hinge points are used as any span-wise distortion or misalignment can be accommodated by designing one of the hinges so that it can rotate about a vertical axis. When more than two hinges are used the „floating hinge concept cannot fully overcome the problems. However, it is possible to design the control surface so that it is flexible in bending and indeed the more hinges there are the easier this is to accomplish. One hinge must always be capable of reacting side loads in the plane of the control surface. The hinges are supported near to the aft extremities of the main surface ribs. 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. Thus the structural layout may consist of an integral root rib or pivot or stub spar arrangement to which is attached a number of shear webs fanning out towards the extremities of the surface, possibly in conjunction with full depth honeycomb. High skin shear loading is inevitable due to the need to bring the loads to the two concentrated points. Shear loads due to torsion may be limited by locating the operating point on the root rib some distance away from the pivot. 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 local geometry allows minimizing 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 majority of flaps and slats are split into span wise segments of no greater lengths than can be supported at two or three locations. As with control surfaces, the locations of the support points are established so as to minimize local deformations since the various slots are critical in determining the aerodynamic performance. Sometimes the actuation may be located at

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a different pan wise position from the support points. This is often a matter of convenience, layout clearances, and the like. 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. Figure shows a cross section of a typical slotted flap of metal construction but the same layout applies if composite materials are used. In many cases the slipstream or afflux from power plants impinges upon a flap and this is likely to require special consideration in the design. Additional stiffness is not necessarily the answer because acoustic fatigue characteristics are often worse at higher panel frequencies. However the extensive local support offered by sandwich construction, either in panel or full depth configuration, is usually beneficial. This leads naturally to the application of reinforced plastic materials. Trailing edge flaps tends to be prone to damage by debris thrown up by the landing gear and it may be desirable to use Kevlar or glass rather than carbon fibers for the lower surface, but material compatibility needs to be considered. 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, the resulting constraint stresses being allowed for. It is usual for the wing structure of large aircraft to include a production joint at the side of the fuselage and this is virtual essential for swept wings. 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. Similar remarks also apply to the attachment of the horizontal stabilizer when the incidence setting is fixed. If the surface is also used for trimming or control, some special consideration is necessary in the location of the pivot and actuation fittings. These usually require a relatively heavily loaded rib or a pair of ribs, and where possible at least one of the attachment points should be close to the rib or spar intersection. It is desirable to arrange for the lateral distance between the pivots to be as great as possible to minimize pivot loads resulting from asymmetric span-wise loading. When the controls are manually operated, it is simplest if the elevator-hinge line and pivot coincide. Fins are usually built integrally with the rear fuselage. This is mainly due to the different form of loading associated with the geometric asymmetry.

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Flutter: Flutter as the dynamic instability of an elastic body in an airstream. It is found most frequently in aircraft structures subjected to large aerodynamic loads such as wings, tail units and control surfaces. Flutter occurs at a critical or flutters speed Vf which in turn is defined as the lowest airspeed at which a given structure will oscillate with sustained simple harmonic motion. Flight at speeds below and above the flutter speed represents conditions of stable and unstable (that is divergent) structural oscillation, respectively. Generally, an elastic system having just one degree of freedom cannot be unstable unless some peculiar mechanical characteristic exists such as a negative spring force or a negative damping force. However, it is possible for systems with two or more degrees of freedom to be unstable without possessing unusual characteristics. The forces associated with each individual degree of freedom can interact, causing divergent oscillations for certain phase differences. The flutter of a wing in which the flexural and torsional modes are coupled is an important example of this type of instability. Some indication of the physical nature of wing bending–torsion-flutter may be had from an examination of aerodynamic and inertia forces during a combined bending and torsional oscillation in which the individual motions are 90 out of phase. In a pure bending or pure torsional oscillation the aerodynamic forces produced by the effective wing incidence oppose the motion; the geometric incidence in pure bending remains constant and therefore does not affect the aerodynamic damping force, while in pure torsion the geometric incidence produces aerodynamic forces which oppose the motion during one-half of the cycle but assist it during the other half so that the overall effect is nil. Thus, pure bending or pure torsional oscillations are quickly damped out. This is not the case in the combined oscillation when the maximum twist occurs at zero bending and vice versa; i.e. a 90 phase difference. The type of flutter described above, in which two distinctly different types of oscillating motion interact such that the resultant motion is divergent, is known as classical flutter. Other types of flutter, non-classical flutter, may involve only one type of motion. For example, stalling flutter of a wing occurs at a high incidence where, for particular positions of the span wise axis of twist, self-excited twisting oscillations occur which, above a critical speed, diverge. Aileron Buzz: Another non-classical form of flutter, aileron buzz, occurs at high subsonic speeds and is associated with the shock wave on the wing forward of the aileron. If the aileron oscillates downwards the flow over the upper surface of the wing accelerates, intensifying the shock and

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resulting in a reduction in pressure in the boundary layer behind the shock. The aileron, therefore, tends to be sucked back to its neutral position. When the aileron raises the shock intensity reduces and the pressure in the boundary layer increases, tending to push the aileron back to its neutral position. At low frequencies these pressure changes are approximately 180 out of phase with the aileron deflection and therefore become aerodynamic damping forces. At higher frequencies a component of pressure appears in phase with the aileron velocity which excites the oscillation. If this is greater than all other damping actions on the aileron a high frequency oscillation results in which only one type of motion, rotation of the aileron about its hinge, is present, i.e. aileron buzz. Aileron buzz may be prevented by employing control jacks of sufficient stiffness to ensure that the natural frequency of aileron rotation is high. Buffeting: Buffeting is produced most commonly in a tail plane by eddies caused by poor airflow In the wing wake striking the tail plane at a frequency equal to its natural frequency; a resonant oscillation having one degree of freedom could then occur. The problem may be alleviated by proper positioning of the tail plane and clean aerodynamic design.

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7 CONCLUSION In conclusion, the series of short range aircrafts incorporated many unique design of future that was never seen on an operational aircraft. The design of these aircrafts points the way for the design of future of very high mach airplanes. The airplane has gone through many design modifications since its early conceptual designs expected, among these was a growth in weight. The document to provide information on the trends in various aircraft characteristics that may influence general long-term airport planning and design. These are strong indications that future trends could see the coexistence of very high capacity aircraft modules of similar capacities for the long range/very long range operations. Cargo payloads, which include mail, express and freight, are increasing in size and weight as larger aircraft service with the airlines, To ensure continued growth in payload and the profitability of cargo operations, improvements in methods, equipment and terminal facilities will be required in order to reduce cargo handling costs and aircraft ground time and to provide improved service for the shippers. We have enough hard work for this design project. A design never gets completed in a flutter sense but it is one step further towards ideal system. But during the design of this aircraft, we learnt a lot about aeronautics and its implications when applied to an aircraft design.

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BIBLIOGRAPHY

Ø “AIRPLANE AERODYNAMICS AND PERFORMANCE” by JAN ROSKAM Ø E.F. Bruhn, “Analysis and Design of Flight Vehicle Structures”, Tristate Offset Co., 1980. Ø Lloyd R. Jenkinson and James F.Marchman III., “Aircraft Design project”, ButterworthHeinemann., Burlington, 2003. Ø Anderson, John D., Jr:Aircraft Performance and Design, McGraw-Hill, Boston, 1999. Ø Megson, T.M.G; Aircraft Structures for Engineering Students, Edward Arnold, 1989 Ø Peery, D.J. and Azar, J.J., Aircraft Structures, 2nd Edition, McGraw-Hill, New York, 1993. Ø McCornic, B.W, “Aerodynamics, Aeronautics & Flight Mechanics”. John Wiley, 1995. Ø G. Corning, “Supersonic & Subsonic Airplane Design”, II Edition, Edwards Brothers Inc., Michigan, 1953. Ø Ira h. Abbott, Albert e. Von Doenhoff,and Louis S. Stivers, Jr,”Summary of Airfoil Data”, National advisory committee for aeronautics, 1947. Ø John T.Lowry., “Performance of Light Aircraft”, American institute of aeronautics and astronautics, Washington.D.C, 1935. Ø Dr.lng.S.F.Hoerner,”Fluid –Dynamic Drag”,Hoerner fluid dynamics., 1992. Ø J.B. Russell,” Performance and Stability of Aircraft”. ISBN 0-340-63170-8. Arnold 1996. Ø Mark D. Ardema, Mark C. Chambers, Anthony P. Patron, Andrew S. Hahn, Hirokazu Miura, and Mark D. Moore, “Analytical Fuselage and Wing Weight Estimation of Transport Aircraft”,1996.

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WEBSITE REFERENCE ª http://www.worldofkrauss.com/ ª http://faculty.dwc.edu/sadraey/V-n%20diagram.pdf

ª http://www.aerostudents.com/files/aircraftStressAnalysisAndStructura lDesign/reader.pdf ª http://www.cta-dlr2009.ita.br/Proceedings/PDF/60272.pdf ª http://www.emteq.com/aircraft-structural-analysis-modifications.php ª http://www.biznet.org.au/member.asp?id=1094&pid=184 ª http://www.cp.berkeley.edu/cds_ucb/UCB-05100.pdf ª http://www.aer.ita.br/~bmattos/download/fuselagem-design.pdf ª http://adg.stanford.edu/aa241/fuselayout/sstfuse.html ª http://www.mat.ethz.ch/news_events/archive/materialsday/matday01/ pdf/TempusMD.pdf ª http://www.free-online-private-pilot-ground-school.com/aircraftstructure.html ª http://www.zenithair.com/stolch701/7-design-fuselage.html ª http://www.nusil.com/products/engineering/aircraft/documents/Aircra ft%20Selection%20Guide.pdf ª http://www.scielo.oces.mctes.pt/pdf/ctm/v20n3-4/v20n3-4a11.pdf ª http://www.faa.gov/library/manuals/aircraft/amt_handbook/media/FA A-8083-30_Ch05.pdf ª http://www.ppart.de/aerodynamics/profiles/NACA5.html ª http://www.desktop.aero/appliedaero/airfoils1/airfoilgeometry.html ª http://www.mh-aerotools.de/airfoils/jf_users_manual.htm

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EIGHT SEATER SHORT RANGE BUSINESS JET AIRCRAFT AN AIRCRAFT DESIGN PROJECT-II REPORT

Submitted by S.VIGNESH (30609101062) P.VASANTHA PRABHU (30609101059) J.SELVA KUMAR (30609101051) N.VIGNESH (30609101061) in partial fulfillm fulfillment for the award of the degree of

BACHELOR OF ENGINEERING in AERONAUTICAL

JEPPIAAR ENGINEERING COLLEGE, CHENNAI 600 119

ANNA UNIVERSITY: CHENNAI 600 025 NOV/DEC 2012

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DEPARTMENT OF AERONAUTICAL ENGINEERING JEPPIAAR ENGINEERING COLLEGE ANNA UNIVERSITY, CHENNAI

BONAFIDE CERTIFICATE Certified that this project report ‘EIGHT SEATER SHORT RANGE BUSINESS JET AIRCRAFT’ is a bonafide work of __S.VIGNESH___________who carried out project under my supervision. Submitted for the examination held on ____6.11.2012________

PROJECT GUIDE

INTERNAL EXAMINER

HEAD OF THE DEPARTMENT

EXTERNAL EXAMINER

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ACKNOWLEDGEMENT It gives us immense pleasure in expressing our sincere gratitude to Honourable Dr.Jeppiaar, M.A., B.L., Ph.D., founder and Chairman of Jeppiaar

Engineering College for bestowing us with an opportunity to bring out this project as a successful one. We are very much grateful to our principal Dr.Susil Lal Das, M.Sc., Ph.D., for their encouragement and moral support. We are very much indebted to Mr.G.Prabakaran (HOD) Aeronautical Department for giving me his able support and encouragement. At this juncture I must emphasis the point that this AIRCRAFT DESIGN PROJECT-II would not have been possible without the highly informative and valuable guidance by our respected preceptor (Ms. Puja Sunil and Ms. Usha Bharathi), Mr. Balaraman whose vast knowledge and experience has must 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.

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ABSTRACT

The aim of this design project is to analysis an 8 Seater Short Range Executive Aircraft by a structural analysis of Shear force and Bending moment. Have to design a more strength aircraft by give the support of stringer, ribs, spar in Wing section and to give the support of stinger, bulkhead, longer in Fuselage. The flying strength of aircraft is analysis by Vn diagram. Then the Design of Miscellaneous Members of Wing Fuel Tank, Rib location and direction, Empennage Design, Auxiliary Surfaces, Wing –Fuselage Intersection, Flutter, Aileron Buzz and Buffeting. Then the necessary graphs have to be plotted for further performance calculation. Required diagrams are also drawn.

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TABLE OF CONTENTS TITLE

1

PAGE

Abstract

i

List of symbols

iv

List of figures

v

List of tables

vi

List of Graphs

vii

Introduction 1.1 Brief review of ADP - 1

1

1.2 Structural Design - Overview

2

2

V-n diagram

4

3

Structural Design of wing

4

3.1 Introduction

10

3.2 Air –Inertia Load Estimation

11

3.3 Shear force & Bending moment Distribution

13

3.4 Material Selection

14

3.5 Wing Spar and Stringer Design

18

3.6 Shear flow Distribution

33

Structural Design of Fuselage 4.1 Design of fuselage

39

4.2 Stringer Design

39

4.3 Shear flow Distribution

43

4.4 Bulkhead design

46

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5

Design of Miscellaneous Members 5.1 Wing Fuel Tank

47

5.2 Rib location and direction

48

5.3 Empennage Design

49

5.4 Auxiliary Surfaces

50

5.5 Wing –Fuselage Intersection

52

5.6 Flutter

53

5.7 Aileron Buzz

54

5.8 Buffeting

55

6

Final design- three view diagram

56

7

Conclusion

57

Bibliography

58

Website reference

59

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LIST OF SYMBOLS A

Total cross sectional area

ft2

ASPAR

Cross sectional area of spar

ft2

Afuselage

Cross sectional area of fuselage

ft2

at

Slope of the CL vs. α curve for a horizontal tail.

Deg-1

a

Distance of the front spar from the nose of the aircraft

ft

b

Distance of the rear spar from the nose of the aircraft

ft

b

Wing span

ft

bw

Width of the web

Ft

bf

Width of the flange

Ft

Cwing

Chord of the actual wing

Ft

Celliptic

Chord of the elliptic wing

Ft

C.G

Centre of gravity

Ft

g

Acceleration due to gravity

ft/s2

D

Drag

Lb

E

Youngĵs modulus

lb/ft2

FOS

Factor of safety

Ftu

Tensile ultimate strength

lb/ft2

H

Height of the C.G from the ground level

Ft

iw

Orientation of wing on fuselage

Deg

Ixx

Second moment of area about X axis

ft4

Lw

Local lift

Lb

WT.O

Takeoff weight

Lb

W/S

Wing loading

lb/ft2

y

span location

Ft

α

Angle of attack

Deg

β

Turnover angle

Deg

ρ

Density

lb/ft3

ρo

Density of air at sea level

lb/ft3

σ

Bending stress

lb/ft2

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LIST OF FIGURES

FIGURE

TITLE

PAGE NO.

1

Load on the aircraft

8

2

Typical V-n diagram

9

3

Final V-n diagram

14

4

Linear lift distribution

16

5

Elliptic Lift Distribution

17

6

Wing Separation Of Finite Section

22

7

Torque Distribution Over an Airfoil

26

8

Shear Center On The Chord

26

9

Wing Spar Arrangement

30

10

Different Spar Selection

31

11

Cross Section of Rear Spar

33

12

Cross Section of Middle Spar

35

13

Semimonocoque And Monocoque

39

14

Cross section Of Z-section

42

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LIST OF TABLES TABLE

TITLE

PAGE NO.

1

Specification of the Aircraft Design Project I

7

2 3 4

Load Factor Velocity VS Load Factor Load on Factor of safety

10 11 13

5

Span VS Linear Lift Distribution

17

6

Span VS Elliptic Lift Distribution

18

7

Span Vs Schrenk's value

18

8

Span Vs Load acting on wing

20

9 10 11 12 13

Centroid table Span VS Shear Force Span VS Bending Moment Span wise VS Shear Force Span wise VS Bending Moment0 Span wise VS Torque at Normal force

22 23 23 25 25

Spanwise VS Mean Aerodynamic Chord

28

16

Frontspar Centroid calculation table

33

17

Front Spar Bending Stress

34

18

Middle Spar Centroid Calculations

35

19

Middle Spar Bending Stress

36

20

Rear Spar Centroid Calculations

37

21

Rear Spar Bending Calculation

38

22 23

Fuselage Structure Analysis Stringer Stress Tabulation

41 44

22

Weight, Moment, Shear Force, Bending Moment

45

14 15

27

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LIST OF GRAPHS

GRAPHS

TITLE

PAGE NO.

1

Span VS Linear Lift Distribution

16

2

Span vs Elliptic lift Distribution

19

3

Span Vs Load acting on wing

20

4

Span Vs Fuel weight distribution

21

5

NET SHEAR FORCE vs BENDING MOMENT

25

6

Span wise VS Bending Moment

28

7

Span vs Net Torque Force

32

8

Stringer location in Fuselage

46

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1 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 8 seater business jet.

The specifications of the Aircraft design project-I as follows: S.No

DESIGN PARAMETER

MAGNITUDE

UNIT

1.

Cruising speed

236.11

m/s

2.

Wing span

12.84

m

3.

Aircraft Length

19.5

m

4.

Wing area

21.84

m2

5.

Height

4.8

m

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6.

Aspect ratio

7.55

(No unit)

7.

Wing loading

585.89

Kg/m2

8.

Empty weight

7,296

Kg

9.

Maximum take-off weight

1,25,568

N

10.

Pay load

1280

kg

11.

No. of engines

2

(No unit)

12.

Thrust power

32.00

Kn

13.

Range

5200

Km

14.

Service ceiling

1,3700

M

15.

Mach no.

0.715

(No unit)

16.

Thrust/weight ratio

0.25

(No unit)

17.

Gliding angle

4.23

o

18.

Seating capacity

8

(No unit)

19.

Fuselage

10.54

M

20.

Take-off distance

1,007.46

M

21.

Landing distance

710.3

M

22.

Rate of climb

1.298 × 10-3

m/s

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LOADS ON THE AIRCRAFT: The structure of an aircraft is required to support two classes of loads, first termed ground loads, includes all loads encountered by the aircraft during movement or transportation on the ground such as taxing, landing loads, towing etc, while the second is the air loads, comprises loads imposed on the structure. The two classes of loads of loads may be still classified as surface forces acting on the surface of the structure and body forces acting over the volume of the structure. Basically all air loads are the resultant of the pressure distribution over the surfaces of the skin produced by steady flight, maneuver or gust conditions. Generally these causes bending, shear, torsion in all parts of the structure in addition to local normal pressure loads imposed on the skin. Ground loads encountered in landing and taxing subject the aircraft to concentrated shock loads through the undercarriage system. The majority of the aircraft have their main undercarriage located in the wings with nose wheel or tail wheel in the vertical plane of symmetry. Clearly the position of the undercarriage should be in such a position so as to produce minimum loads on the wing structure.

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2 ESTIMATION OF V-N DIAGRAM The control of weight in aircraft design is of extreme importance. Increase in weight requires stronger structures to support them, which in turn lead to further increase in weight & so on. Excess of structural weight means lesser amounts of payload, affecting the economic viability of the aircraft. Therefore there is need to reduce aircraft’s weight to the minimum compatible with safety. Thus to ensure general minimum standards of strength & safety, airworthiness regulations lay down several factors which the primary structures of the aircraft must satisfy. These are 1. LIMIT LOAD: the maximum load that the aircraft is expected to experience in normal operation. 2. PROOF LOAD: product of the limit load and proof factor(1.0-1.25) 3. ULTIMATE LOAD : product of limit load and ultimate factor(1.0-1.5) The aircraft’s structure must withstand the proof load without detrimental distortion & should not fail until the ultimate load has been achieved. V-n Diagram: A chart of Velocity versus load factor (V-n diagram) is another way of showing limits of aircraft performance. It shows how much load factor can be safely achieved at different

airspeeds.

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The maneuverability of the aircraft is also dictated by the loads falling on the structures during the maneuvers. Both the aerodynamic and structural limitations for a given airplane are illustrated in the V-n diagram, a plot of load factor versus flight velocity. A V-n diagram is type of flight envelope for the aircraft establishing the maneuver boundaries. The BCAR (British civil airworthiness requirements) has given the basic strength and flight performance limits of various categories of the aircraft. They are listed below

Category

Positive load factor (n+)

Negative load factor(n-)

Normal

3.8

-1.5

Semi aerobatic

4.5

-2

Fully aerobatic

6

-3

Tabular column 2: LOAD FACTOR The 8 seater executive aircraft comes under the normal category. Therefore the load factor limits for the aircraft is 3.8 & -1.5. The V-n diagram for the aircraft is drawn for the two cases namely, 1. Intentional maneuver( pilot induced maneuver ) 2. Unintentional maneuver( gusts) INTENTIONAL MANEUVER: Intentional maneuvers are induced by the pilot during climb, pull up or pull down, banking the plane etc... The load factor is function of velocity. The expression relating the load factor and the velocity is given by nmax =

Where nmax is the maximum load factor, V is the speed of the aircraft, Vs is the stalling speed of the aircraft. The stalling speed of the aircraft Vs 2 =

Vs= 59.197 m/s

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For various values of V, nmax is calculated and tabulated below,

V

Nmax=

V

Nmax=

4

118.394

-0.25

29.598

3.8

115.396

-0.5

41.858

3.5

110.74

-0.75

51.256

3

102.53 -1

59.197

2.5

93.599

2

83.71

1.5

72.501

1

59.197

Tab3. Velocity VS Load Factor

The cruising speed of the aircraft is 236.11 m/s. The dive speed of the aircraft is the maximum speed of the aircraft. The dive speed is equal to the sum of the cruising speed and 60 knots. VD = 236.11 + 60 knots = 236.11 +30.56 m/s = 266.67 m/s GUSTS: The movement of air in turbulence is known as gusts. It produces changes in wing incidence, thereby subjecting the aircraft to sudden or gradual increases or decreases in lift from which normal accelerations result. These may be critical for large, high speed aircraft and may possibly cause higher loads than control initiated maneuvers. Thus in the gust analysis, the change in load factor due to the gust is calculated. The BCAR has given standard gust velocities for stall, cruise, dive speeds as 66, 50, 25 ft/s

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respectively. The small change in load factor ∆n due to the gust is calculated by assuming a sharp gust. The change in load factor ∆n = Where ρ è density at cruising altitude, a è lift slope, in radians U è gust velocity in m/s V è velocity of the aircraft in m/s W/S è wing loading in N/m2 In the above formula, gusts are assumed to be sharp but it is usually graded, hence a relief factor called gust alleviation factor K is introduced in the term. The value of the K is obtained from the book “AIRPLANE AERODYNAMICS AND PERFORMANCE” by JAN ROSKAM Where K =

,

!" µ =#$%$&

%$'(

Where ρ is the density, C is the mean aerodynamic chord, g is the acceleration due to gravity; CLα is the slope lift coefficient. The CLα (corrected for aspect ratio) is 0.0962/deg.

µ =

$ $)$ *

µ=

$+,+

,$-.$.+$.$+

µ = 50.36

K= =

$/ /

Κ = 0.796

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K = Therefore ∆n = 0

123 # !"

For STALL SPEED V= 59.197 m/s, U= 20m/s ∆n = 0

=

# 123 !"

+-/$,$.$+$$-.-+ $+,+

∆n = 1.062 For CRUISE SPEED V=236.11 m/s, U= 15m/s ∆n = 0

=

# 123 !"

+-/$,$.$+$.$/.. $+,+

∆n = 2.725 For DIVE SPEED V= 266.67 m/s, U= 7.5 m/s ∆n = 0

=

# 123 !"

+-/$,$.$+$+$///+

∆n ∆ = 1.794

$+,+

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V

1+∆ ∆n

1-∆ ∆n

59.197

2.062

-0.062

236.11

3.725

-1.725

266.67

2.794

-1.794

Tab 4.load Factor of safety

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3 STRUCTURAL ANALYSIS AND DESIGN OF WINGS WING STRUCTURAL ANALYSIS: The structural design of the wing requires a complete quantitative knowledge of the different loads it will be subjected to during its flight regime. These loads can be briefly classified as 1. Distributed loads - Loads such as aerodynamic loads, weight of the wing and weight of fuel. 2. Concentrated loads – Loads such as thrust, engine weight, landing gear weight and armament weight. LOADS ACTING ON WING: As both the wings are symmetric, let us consider the starboard wing at first. There are three primary loads acting on a wing structure in transverse direction which can cause considerable shear forces and bending moments on it. They are as follows: v Lift force (given by Schrenk’s curve) v Self-weight of the wing v Weight of the power plant v Weight of the fuel in the wing SCHRENK’S CURVE: Lift is a component of the resultant aerodynamic force acting at the centre of pressure of an aerodynamic chord, along a direction perpendicular to the direction of the relative wind. At a particular altitude and at a specific angle of attack, Lift varies along the wing span due to the variation in chord length along the span. Schrenks curve defines this lift distribution over the wing span of an aircraft. Since the wings of an aircraft are symmetrical about the longitudinal axis, the Schrenks curve for the starboard wing alone can be obtained at first. This is given by y=

45 4

where y1 è linear variation of lift along the wing semi-span y2 è equivalent elliptic lift distribution along the wing semi-span

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TO FIND y1: Lift force is found along the line joining the aerodynamic centers of chords along the wing span. Hence, the wing is rotated about the wing root so that the line joining the aerodynamic centers becomes the horizontal line. 6 a= ) 789:

a=

.,

789..; .<

)

=6.55 Lift per unit length at wing root

= CL×0.5×ρ×V2×CR = 0.23884×0.5×1.4×236.1112×2.55 = 23766.98 N/m

Lift per unit length at wing tip

= CL×0.5×ρ×V2× Ct =0.23884×0.5×1.4×236.1112×0.84 = 7829.12 N/m

6.55 m Fig.4 Linear lift distribution Area under trapezoid life distribution = 155673.719 Equation of linear lift distribution for starboard wing Y1 = -1195.289x + 23766.98 Equation of linear lift distribution for port wing we have to replace x by –x in general, Y1 = 1195.289x + 23766.98

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X

Y1

0

23766.98

1

22571.7

2

21376.42

3

20181.14

4

18985.86

5

17790.58

6

16595.3

6.55

15937.89 Tab5. Span VS Linear Lift Distribution

LINEAR LIFT DISTRIBUTION

25000 20000 15000 Series1

10000

Series2 5000 0 0

1

2

3

4

5

6

6.55

SPAN (a)

Graph.4 Span VS Linear Lift Distribution

Elliptic Lift Distribution: Twice the area under the curve or line will give the lift which will be required to overcome weight Considering an elliptic lift distribution we get =>6?

L/2 = W/2 = , Where b1 is Actual lift at root A= And a is wing semi span

=>6? ,

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Lift at Tip b = 15138.35 N/m

15138.35 N/m

6.55 m Fig 5. Elliptic Lift Distribution

Y2 = 1155.60@A BCD E F

x

Y2

0

15138.317

1

14960.852

2

14415.354

3

13457.142

4

11987.651

5

9779.05

6

6072.292

6.55

0 Tab 6: Span VS Elliptic Lift Distribution

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Construction of Schrenk”s Curve: Schrenk”s Curve is given by

Y1+Y2 =

55BGHBIDJKKBH55GGKCLABCDMF

Y1+Y2 = -597.64 x + 45489.019 + 1155.60 @A BCD E F

x

Y1+Y2

0

19452.65

1

18766.28

2

17895.89

3

16819.14

4

15486.76

5

13784.82

6

11333.796

6.55

7968.95 Tab 7:SpanVsSchrenk's value

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Graph 2: Span vs Elliptic lift Distribution

Load Estimation on wings Description: The solution methods which follow Euler’s beam bending theory (σ/y=M/I=E/R) use the bending moment values to determine the stresses developed at a particular section of the beam due to the combination of aerodynamic and structural loads in the transverse direction. Most engineering solution methods for structural mechanics problems (both exact and approximate methods) use the shear force and bending moment equations to determine the deflection and slope at a particular section of the beam. Therefore, these equations are to be obtained as analytical expressions in terms of span wise location. The bending moment produced here is about the longitudinal (x) axis. Loads acting on wing: As both the wings are symmetric, let us consider the starboard wing at first. There are three primary loads acting on a wing structure in transverse direction which can cause considerable shear forces and bending moments on it. They are as follows: ª Lift force (given by Schrenk’s curve) ª Self-weight of the wing ª Weight of the power plant ª Weight of the fuel in the wing

Self-Weight (y3): Self-weight of the wing, Wwing = 5548.06 kg ×9.81 = 54426.46 N

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Wport wing =-27213.23 N Wstar board wing = -27213.23 N / / PQM/R N O =N

y3= -290.52 (x-6.55)2 x

y

0

-12464.07

1

-8948.74

2

-6014.49

3

-3661.28

4

-1889.11

5

-697.97

6

87.88

6.55

0 Tab 8: SpanVs Load acting on wing

Graph 3: Span Vs Load acting on wing

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Power Plant: According to our design data, Our Aircraft power plant is attach to rear fuselage. So, power plant calculation won’t be calculated.

Fuel Weight Distribution: Wf = 23215.55 Consider as equation, yf = 1902.91x-12464.07

2

-8658.25

3

-6755.35

4

-4852.43

5

-2949.52

Fuel weight Distribution 0 -1000

2

3

4

5

6

-2000 -3000 -4000 -5000 -6000 -7000 -8000 -9000 -10000

Graph 4. Span Vs Fuel weight distribution

6.55

P a g e | 28 [email protected]

SHEAR FORCE & BENDING MOMENT SHEAR FORCE AND BENDING MOMENT DIAGRAMS OF A WING DUE TO LOADS IN TRANSVERSE DIRECTION AT CRUISE CONDITION: The solution methods which follow Euler's beam bending theory (σ/y=M/I=E/R) use the bending moment values to determine the stresses developed at a particular section of the beam due to the combination of aerodynamic and structural loads in the transverse direction. Most engineering solution methods for structural mechanics problems (both exact and approximate methods) use the shear force and bending moment equations to determine the deflection and slope at a particular section of the beam. Therefore, these equations are to be obtained as analytical expressions in terms of span wise location. The bending moment produced here is about the longitudinal (x) axis. As both the wings are symmetric, let us consider the starboard wing at first. There are three primary loads acting on a wing structure in transverse direction which can cause considerable shear forces and bending moments on it. They are as follows: → Lift force (given by Schrenk's curve) → Self-weight of the wing → Weight of the powerplant S.No.

Curve/Component

Area/Structural weight (N)

Centroid

1

Y1

155673.719

5.458

2

Y2

77896.859

2.781

3

Wing

54426.46

1.637

4

Fuel

23215.55

1.31

Tab 9: Centroid table

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∑V=0 77836.859-54426.46-23215.55-VA = 0 VA = 194.85 ∑MA = 0 MA+ (54426.465×1.637)+(23215.55×1.31)-(155673.719×5.458)-(77836.859×2.781) MA – 946622.97 = 0 MA= 946622.97

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SHEAR FORCE T

1

F

FD

U -290.52 T

KGG

D

S.F1 = -298.84 x2 + 45489.019 x + 577.8 @A BCD E S - 42.903 Sin-

V A BCF V K GGF U -194.85

SPAN

Shear Force

0

-19625.79

1

-8113.17

2

1496.55

3

9203.37

4

15007.29

5

18908.31

6

20906.43

6.55

21194.28 Tab 10: Span VS Shear Force

BENDING MOMENT:

F

B.M = -199.21 x3 + 15163.006 x2 + 288.9 [x (x@A BCD E S ) + 42.903 Sin -1KGG] + 385.2 FA

(42.903- x2)1.5-290.52 ( + 21.45 x2- 2.18 x3) + 94662.97 5

P a g e | 31 [email protected]

SPAN

BENDING MOMENT

0

754870.44

1

1043525.353

2

1062831.914

3

1084364.515

4

1108829.591

5

1318721.935

6

1381469.212

6.55

943585.515 Tab 11 : Span VS Bending Moment

NET SHEAR FORCE AND BENDING MOMENT DIAGRAM:

1600000

BENDING MOMENT

1400000 1200000 1000000 800000 600000 400000

SHEAR FORCE

200000 0 -8

-6

-4

-2

0

2

4

6

8

WING SPAN

Graph.5 NET SHEAR FORCE vs BENDING MOMENT

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Shear force and bending moment diagrams due to loads along chordwise direction at cruise condition: Aerodynamic center- This is a point on the chord of an airfoil section where the bending moment due to the components of resultant aerodynamic force (Lift and Drag) is constant irrespective of the angle of attack. Hence the forces are transferred to this point for obtaining constant Ma.c Shear center- This is a point on the airfoil section where if a force acts, it produces only bending and no twisting. Hence the force is transferred to this point and the torque is found. Angle of Attack (max) = 15.00,

Angle of Attack (Zero lift) = -1.0

Cruise CL = 1.40

V = 236.11 m/s

ρ = 0.23884 kg/m3

CD = 0.0025

SHEAR FORCE BENDING MOMENT: Co-efficient of force at normal direction Cn = CL cos ά + CD sin ά = 1.398 Co-efficient of force at Chordwise direction CC = CL sin ά + CD cos ά = 0.026 Therefore, Force per unit length= Cc×0.5×ρ×V2×C Force at Cr = 441.39 N/m Force at Ct = 145.39 N/m For Linear, y = 23.05 x + 145.39 à1 Shear Force: Integrate Eqn. 1 5HA

NC

W 55 GGFD V 5AG DBF

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Span wise

Shear force

0

0

1

156.915

2

336.88

3

539.895

4

765.96

5

1015.07

6

1287.24

7

1582.45

8

1900.72

9

2242.03

10

2606.40

11

2993.81

12

3404.28

12.84

3766.88 Tab 12: Span wise VS Shear Force

BENDING MOMENT: .,

X

W3.841 x3+72.69x2

Span wise

Bending moment

0

0

1

76.531

2

176.108

3

321.77

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4

536.584

5

843.575

6

1265.796

7

1826.293

8

2548.112

9

3454.299

10

4567.9

11

5911.961

12

7509.528

12.84

9064.425 Tab 13: Span wise VS Bending Moment

14000

BENDING MOMENT & SHEAR FORCE BENDING MOMENT

12000 10000 8000 6000

SHEAR FORCE

4000 2000 0 CHORDWISE

Graph 6. Span wise VS Bending Moment

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TORQUE DUE TO NORMAL FORCES AND CONSTANT PITCHING MOMENT AT CRUISE CONDITION: Aerodynamic center- This is a point on the chord of an airfoil section where the bending moment due to the components of resultant aerodynamic force (Lift and Drag) is constant irrespective of the angle of attack. Hence the forces are transferred to this point for obtaining constant Ma.c Shear center- This is a point on the airfoil section where if a force acts, it produces only bending and no twisting. Hence the force is transferred to this point and the torque is found

The lift and drag forces produce a moment on the surface of cross-section of the wing, otherwise called a torque, about the shear center. Moment about the aerodynamic center gets transferred to the shear center. The shear center on the chord under which it is locates.

Cruise condition (Normal Force) T= ½ Cn ρ V2 c × 0.034 C

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= 1325.51 C2 Where, C à chord the equation for chord can also be represented in terms of x by taking C = mx +k,

C = 0.264 x +2.55 Therefore, Torque = 30.751x3 + 8418.46 x + 892.33 x2

Span wise

Torque at Normal force

0

0

1

9541.54

2

21052.24

3

34716.62

4

50719.18

5

69244.42

6

90476.85

7

114600.98

8

141801.31

9

172262.34

10

206168.6

11

243704.57

12

285054.76

12.84

322871.54

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Tab 14:Span wise VS Torque at Normal force TORQUE DUE TO CHORDWISE FORCE: Torque per unit length T2 = FC×0 T2 = 0 TORQUE DUE TO MEAN AERODYNAMIC CHORD: Torque due to Ma.c = YZ[\ ×0.5×ρ×V2×C×C T3 = -3347.9 ×C2 T3 = -77.671 x3- 21768.04 x- 2253.806 X2 Span wise

MEAN AERODYNAMIC CHORD

0

0

1

-24099.51

2

-53172.67

3

-87685.49

4

-128104.0

5

-174894.22

6

-228522.19

7

-289453.92

8

-358155.45

9

-435092.80

10

-520732.01

11

-615539.06

12

-719980.03

12.84

-815496.45

P a g e | 38 [email protected]

Tab 15: Spanwise VS Mean Aerodynamic Chord NET TORQUE: Then the different torque components are brought together in a same graph to make a comparison The net torque will be sum of all the above torques (i.e.) torques due to normal force, chordwise force, powerplant and aerodynamic moment.

400000 200000 0 1 -200000

2

3

4

5

6

7

8

9 10 11 12 13 14 Series1 Series2

-400000

Series3

-600000 -800000 -1000000

Graph.7 Span vs Net Torque Force

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LOAD ESTIMATION OF WINGS WING STRUCTURAL LAYOUT Specific Roles of Wing (Mainwing) Structure: The specified structural roles of the wing (or main plane) are: Ø 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. Ø To react against: Landing loads at attachment points Loads from pylons/stores Wing drag and thrust loads Ø To provide: Fuel tank age space Torsional rigidity to satisfy stiffness and aero-elastic requirements. Ø 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 v Form the main span wise beam v Transmit bending and torsional loads v Produce a closed-cell structure to provide resistance to torsion, shear and tension loads. In particular: v Webs – resist shear and torsional loads and help to stabilize the skin. v Flanges - resist the compressive loads caused by wing bending. SKIN v To form impermeable aerodynamics surface v Transmit aerodynamic forces to ribs & stringers v Resist shear torsion loads (with spar webs). v React axial bending loads (with stringers). STRINGERS v Increase skin panel buckling strength by dividing into smaller length sections. v React axial bending loads RIBS v Maintain the aerodynamic shape v Act along with the skin to resist the distributed aerodynamic pressure loads v Distribute concentrated loads into the structure & redistribute stress around any discontinuities v Increase the column buckling strength of the stringers through end restraint v Increase the skin panel buckling strength. SPAR DEFINITION: The maximum bending moment from previous section was found to be as 2897784.51 Nm. Therefore we define 3 Spars with front spar at 15% of chord, middle spar at 45% of chord and rear spar at 70% of chord. The position of the three spars from the leading edge of the root chord is given below as follows:

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Front spar - 15% of chord = 2.442 m Middle spar - 45% of chord = 7.326 m Rear spar - 70% of chord = 11.396 m Bending moment M = Max BM * FOS * n = 2897784.51 × 1.5 × 3.8 = 16517371.71Nm The Structural load bearing members in the wing are the Spars and Stringers. The bending moment carried by the Spars is 70% and that of Stringers is 30% of the total Bending Moment. Bending Moment taken by Spars is = 0.7 x 16517371.71 = 11562160.19 Nm The cross section of the spar chosen here is an I-section For each spar we are determining the following parameters: A. B. C. D.

Centroid Moment of Inertia Bending Moment Bending Stress

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FRONT SPAR: ª Height of the spar = 38 cm ª Breadth of the spar = 16 cm ª Thickness of the spar = 4.5 cm

Fig 11: Cross Section of Rear Spar To find out the centroid, the following calculations are made: Element

Area(A) (cm2 )

x (cm)

y (cm)

Ax (cm3 )

Ay (cm3 )

Ax2 (cm4 )

Ay2 (cm4 )

Icx (cm4 )

Icy (cm4 )

1

72

8

2.25

576

162

4608

364.5

121.5

1536

2

130.5

8

19

1044

2479.5

8352

47110.5

9145.8

220.22

3

72

8

35.75

576

2574

4608

92020.5

121.5

1536

Total

274.5

2196

5215.5

17568

139495.5

9388.87

3292.22

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Front Spar Calculations: Centroid = X =

]^_ ]^

= 8 cm;

Y=

]^` ]^

= 19 cm

I xx = Σ Icx + ΣAy2 – ΣAY2 I xx = (9388.87) + (139495.5) – (274.5)(19)2 I xx = 49789.88 cm4

I yy = Σ Icy+ ΣAx2 – ΣAX2 I yy = (3292.22) + (17568) – (274.5) (8)2 I yy = 3292.22 cm4 The FRONT SPAR carries 35 % of the BM carried by the Spars. Thus, Front spar BM = 0.35 x 1156216019 N-cm

= 404675606.7 N cm Front Spar Bending Stress: a Bending Stress, σ z = c by

POINTS

COORDINATES (y) (cm)

BENDING STRESS (N/cm2)

A

19

154425.68

B

14.5

117851.18

C

14.5

117851.18

D

-14.5

-117851.18

E

-14.5

-117851.18

F

-19

-154425.68

The bending stress at various points whose co-ordinates are determined with centroid as the origin are calculated from above formula and tabulated.

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MIDDLE SPAR: ª Height of the spar = 41.6 cm ª Breadth of the spar = 18 cm ª Thickness of the spar = 5 cm

Fig 12: Cross Section of Middle Spar

P a g e | 45 [email protected]

To find out the centroid, the following calculations are made: Element

Area(A) (cm2 )

x (cm)

y (cm)

Ax (cm3 )

Ay (cm3 )

Ax2 (cm4 )

Ay2 (cm4 )

Icx (cm4 )

Icy (cm4 )

1

90

9

2.5

810

225

7290

562.5

187.5

2430

2

158

9

20.8

1422

3286.4

12798

68357.12

13147.7

329.17

3

90

9

39.1

810

3519

7290

137592.9

187.5

2430

Total

338

3042

7030.4

27378

206512.5

13522.7 5189.17

Middle Spar Calculations Centroid = X =

]^_ ]^

= 9 cm;

Y=

]^` ]^

= 20.8 cm

I xx = Σ Icx + ΣAy2 – ΣAY2 I xx = (13522.7) + (206512.5) – (338) (20.8)2 I xx = 60467.7 cm4

I yy = Σ Icy+ ΣAx2 – ΣAX2 I yy = (5189.17) + (27378) – (338) (9)2 I yy = 5189.17 cm4 The bending moment carried by the middle spar is 40% of the total bending moment carried by the spars. Middle Spar BM = 462486407.6 N-cm Bending Stress, σ z =

d e

y

P a g e | 46 [email protected]

POINTS

COORDINATES (y) (cm)

BENDING STRESS (N/cm2)

A

20.8

159088.52

B

15.8

120846.09

C

15.8

120846.09

D

-15.8

-120846.09

E

-15.8

-120846.09

F

-20.8

-159088.52

The bending stress at various points whose co co-ordinates ordinates are determined with centroid as the origin are calculated from above formula and tabulated. REAR SPAR ª Height of the spar = 17.72 cm ª Breadth of the spar = 7.6 cm ª Thickness of the spar = 2.5 cm

P a g e | 47 [email protected]

To find out the centroid, the following calculations are made: Element

Area(A) (cm2 )

x (cm)

y (cm)

Ax (cm3 )

Ay (cm3 )

Ax2 (cm4 )

Ay2 (cm4 )

Icx (cm4 )

Icy (cm4 )

1

19

3.8

1.25

72.2

23.75

274.36

29.6875

9.896

91.45

2

31.8

3.8

8.86

120.84

281.748

459.19

2496.287

428.76

16.56

3

19

3.8

16.47

72.2

312.93

274.36

5153.957

9.896

91.45

Total

69.8

265.24

618.428

1007.9

7679.932

448.552

199.46

Rear Spar Calculations Centroid = X =

]^_ ]^

= 3.8 cm;

Y=

]^` ]^

= 8.86 cm

I xx = Σ Icx + ΣAy2 – ΣAY2 I xx = (448.552) + (7679.932) – (69.8) (8.86)2 I xx = 2649.184 cm4

I yy = Σ Icy+ ΣAx2 – ΣAX2 I yy = (199.46) + (1007.9) – (69.8) (3.8)2 I yy = 199.46 cm4 Rear Spar carries 25 % of the spar Bending Moments. Bending Moment = 289054004.8 N-cm d

Bending Stress, σ z = e y

P a g e | 48 [email protected]

The bending stresses at various points are obtained as: Rear Spar Bending Stress POINTS

COORDINATES (y) (cm)

BENDING STRESS (N/cm2)

A

8.86

966719.74

B

6.36

693943.29

C

6.36

693943.29

D

-6.36

-693943.29

E

-6.36

-693943.29

F

-8.86

-966719.74

The bending stress at various points whose co-ordinates are determined with Centroid as the origin are calculated from above formula and tabulated.

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4 STRUCTURAL ANALYSIS AND DESIGN OF FUSELAGE FUSELAGE STRUCTURAL LAYOUT: The fuselage is the main structure, or body, of the aircraft. It provides space for personnel, cargo, controls, and most of the accessories. The power plant, wings, stabilizers, and landing gear are attached to it. There are two general types of fuselage construction construction—welded welded steel truss and monocoque designs. The welded steel truss was used in smaller Navy aircraft, and it is still being used in some helicopters. onocoque design relies largely on the strength of the skin, or covering, to carry The monocoque various loads. The monocoque design may be divided into three classes - monocoque, semimonocoque and reinforced shell.

Fig 13:Semimonocoque And Monocoque

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v The true monocoque construction uses formers, frame assemblies, and bulkheads to give shape to the fuselage. However, the skin carries the primary stresses. Since no bracing members are present, the skin must be strong enough to keep the fuselage rigid. The biggest problem in monocoque construction is maintaining enough strength while keeping the weight within limits. v Semimonocoque design overcomes the strength-to-weight problem of monocoque construction. In addition to having formers, frame assemblies, and bulkheads, the semimonocoque construction has the skin reinforced by longitudinal members. v The reinforced shell has the skin reinforced by a complete framework of structural members. Different portions of the same fuselage may belong to any one of the three classes. Most are considered to be of semimonocoque-type construction. The semimonocoque fuselage is constructed primarily of aluminum alloy, although steel and titanium are found in high-temperature areas. Primary bending loads are taken by the longerons, which usually extend across several points of support. The longerons are supplemented by other longitudinal members known as stringers. Stringers are more numerous and lightweight than longerons. The vertical structural members are referred to as bulkheads, frames, and formers. The heavier vertical members are located at intervals to allow for concentrated loads. These members are also found at points where fittings are used to attach other units, such as the wings and stabilizers. The stringers are smaller and lighter than longerons and serve as fill-ins. They have some rigidity but are chiefly used for giving shape and for attachment of skin. The strong, heavy longerons hold the bulkheads and formers. The bulkheads and formers hold the stringers. All of these join together to form a rigid fuselage framework. Stringers and longerons prevent tension and compression stresses from bending the fuselage. The skin is attached to the longerons, bulkheads, and other structural members and carries part of the load. The fuselage skin thickness varies with the load carried and the stresses sustained at particular location. There are a number of advantages in using the semimonocoque fuselage. v The bulkhead, frames, stringers, and longerons aid in the design and construction of a streamlined fuselage. They add to the strength and rigidity of the structure. v The main advantage of the semimonocoque construction is that it depends on many structural members for strength and rigidity. Because of its stressed skin construction, a semimonocoque fuselage can withstand damage and still be strong enough to hold together.

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Loads and its distribution: To find out the loads and their distribution, consider the different cases. The main components of the fuselage loading diagram are: v Weight of the fuselage v Engine weight v Weight of the horizontal and vertical stabilizers v Tail lift v Weight of crew, payload and landing gear v Systems, equipment, accessories Symmetric flight condition, steady and level flight: (Downward forces negative) Values for the different component weights are obtained from aerodynamic design calculations. S.No.

Equipment and Component

Length from Ref. point

Weight (N)

Moment

Shear Force

Bending Moment

1

Nose

0.58

3933.81

2281.609

3933.81

2281.809

2

Pilot(2)

2.02

1962.00

3963.24

5895.81

6244.84

3

Cockpit

2.62

4944.24

12953.908

10840.05

19198.757

4

Wing

6.62

54426.46

360303.16

65266.51

379501.91

5

Passenger(3)

6.63

3237.30

21463.29

68503.81

400965.2

6

Passenger(3)

7.92

3237.30

25639.41

71741.11

426604.61

7

Passenger(2)

9.85

2060.10

20291.98

73801.21

446896.59

8

Crew(2)

10.32

2060.10

21260.23

75861.31

468156.82

9

Weight of Fuselage Sheet

10.50

2000.68

21007.14

77861.99

489163.96

10

Power Plant

14.23

3937.83

56036.74

81799.92

545200.72

11

Horizontal Tail

17.10

13232.807

226280.9

95032.72

771481.607

12

Vertical tail

17.60

8971.34

157895.58

104004.06

929377.18

SHEAR FORCE AND BENDING MOMENT DIAGRAM

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To determine the shear force and bending moment diagram for the wing we assume that the wing is a cantilever beam with the root end fixed while the tail end is free. For a cantilever beam the shear force is a given by, Shear Force = Rx fQ g Bending Moment = Tabulation for the values of shear force and bending moment at various positions along the span is as follows.

Graph 9. Length from Ref. point Vs Shear Force 120000

Shear force

100000 80000 60000 40000 20000 0 0.58 2.02 2.62 6.62 6.63 7.92 9.85 10.32 10.5 14.23 17.1 17.6 Length from Ref. point

Bending moment

Graph 10. Lengtth from Ref. point Vs Bending Moment 1000000 900000 800000 700000 600000 500000 400000 300000 200000 100000 0 0.58 2.02 2.62 6.62 6.63 7.92 9.85 10.32 10.5 14.23 17.1 17.6 Leght from Ref. point

FUSELAGE STRUCTURAL ANALYSIS

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Structural analysis of fuselage like that of wing is of prime importance while designing an aircraft. As the fuselage is the one which houses the pilot, the power plant and also part of the payload its structural integrity is a matter of concern. While analyzing the fuselage structure the section must be idealized. Idealization involves the conversion of a stringer and its accompanying skin thickness into a concentrated mass known as a boom. The shear flow analysis of the fuselage simulating flight conditions is shown below. X (m) 1.005

Y (m) 0

0.985

0.26

0.88

0.48

0.72

0.72

0.48

0.88

0.26

0.985

0

1.005

-0.26

0.985

-0.48

0.88

-0.72

0.72

-0.88

0.48

-0.985

0.26

-1.005

0

-0.985

-0.26

-0.88

-0.48

-0.72

-0.72

-0.48

-0.88

-0.26

-0.985

0

-1.005

0.26

-0.985

0.48

-0.88

0.72

-0.72

0.88

-0.48

0.985

-0.26

1.005

0

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Stringer location in Fuselage 1.5

1

0.5

0 -1.5

-1

-0.5

0

0.5

-0.5

-1

-1.5

The stringer used is of Z type. The following are its dimensions Cross sectional area of each stringer is 100mm2 Fig 14: Cross section of Z Z-section

1

1.5

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The above stringer section is uniformly used throughout the fuselage as shown above in order to provide the fuselage the required load carrying capacity. The diagram showed adjacent is of the idealized fuselage structure. The idealization process is carried out in the following way. STRESS ANALYSIS: IDEALIZATION: The boom 1 is given by

where, B1 èArea of Boom 1 tD èThickness of skin panel b è Circumferential distance between 2 stringers

By Symmetry, B1 = B9, B2 = B8, B10 = B16, B3 = B7 , B11 = B15, B4 = B6 = Bl2 = B14 ,B5 = B13 .h

B1=100+ (0.65×1.37×

/

) [2+

,..

] + (0.65×1.37×

.h /

)[2+

,..

]

=815582.12 Similarly for boom 2 ,

B2 = 815582.12 mm2 Similarly B3 = 815582.12 mm2, B4 =815582.12 mm2. We note that stringers 5 and 13 lie on the neutral axis of the section and are therefore unstressed; the calculation of boom areas B5 and B13 does not then arise. Thus, we have B1:B16 = 815582.12 mm2

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We know that, Ixx = By2

c__? = 24.67 m4; c__g = 13.77 m4; c__R = 6.12 m4; c__i = 1.11 m4 Maximum bending moment = 2897784.51 Nm Hence the Bending moment acting on the fuselage M = Max.B.M × n× FOS =2897784.51 × 3.8×1.5 =16517371.71 Nm Ixx = 24.67 m4 The value of stress acting is given by the expression: =

./.++.+.$` ,/+

STRINGER/BOOM

Y (m)

STRESS x 10 6 (Nm-2)

1

5.5

3.68

2, 16

4.11

2.75

3, 15

2.74

1.83

4, 14

1.37

0.9

5, 13

0

0

6, 12

-1.37

-0.9

7, 11

-2.74

-1.83

8, 10

-4.11

-2.75

9

-5.5

-3.68

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5 Design of Miscellaneous Members Wing fuselage intersection The 8 seater business jet aircraft has low wing configuration, thus the entire wing structure continues in the way of airplane body. Four pin design concept: This concept is adopted as it is the most simple and straight forward method used in Horizon 4000 transport, during 1950s. The lift and moment loads can be carried between the wing and fuselage by simple shear on the four pins. The drag and thrust is taken by breather web. This design allows the wing spar and fuselage bulkheads to deflect independently of each other such that no spar moment is directly transferred to the bulkheads.

The wing-body juncture produces aerodynamic interference which in turn promotes flow separation with its attendant higher drag and unsteady buffeting. This adverse pressure gradient and consequent flow separation can be minimized using contoured surface called fillet. Engine mount An engine mount is a frame that supports the engine and holds it to the fuselage or nacelle. Usually it is made of built up sheet metal, welded steel tubing. The turbofan engine, “HONEY WELL TFE731-20” is wing mounted. A typical turbofan engine installation for a low wing aircraft configuration similar to that of this aircraft is shown below,

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The pylon has three spars (longerons) – Upper, middle and lower- and three major bulkheads, and is attached to the wing at four primary points. These are two mid mid--spar fittings, an upper link and a diagonal brace (drag strut). The attachment pins are secured with “fuse” bolts which are hollow carbon steel devices that have been heat treated to shear fail at a definite load. In the landing break way condition (wheels (wheels-up landing); anding); the sequence is designed to fail the upper and lower links so that the pylon rotates around the mid mid-spar spar and upward. The wing pylon design provides considerable load path redundancy such that an upper link can fail, partially or completely, and there ere is an alternate path path- lower diagonal brace. The below figure shows the engine mounts.

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Empennage Design Horizontal Stabilizer: The horizontal tail of the aircraft is conventional and consists of a fixed tail box. The horizontal stabilizer is usually a two spar structure consisting of a Centre structural box section and two outer sections. The stabilizer assembly is interchangeable (symmetrical airfoil section) as a unit at the fuselage attach points and the outer sections are interchangeable at the attachment to the center box. The two basic horizontal stabilizer box constructions for modern transports are 1. Box constructions with spars, closer light rib spacing (usually less than 10 inches) and surface (may be tapered skins) without stringer reinforcement. The feature of this design is the low manufacturing cost and high torsional stiffness require by the flutter analysis.

2. Box construction with spar stronger ribs and surface skins with stringer reinforcements (skin-stringer or integrally stiffened panels) is a lighter weight structure.

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Vertical Stabilizer: The structural design of the vertical stabilizer is essentially the same as for the horizontal stabilizer is essentially the same as for horizontal stabilizers. The vertical stabilizer box is a two or multi spar structure (general aviation airplanes usually use single spar design) with cover panels (with or without ribs). The root of the box is terminated at the aft fuselage conjuncture with fittings or splices.

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WING FUEL TANKS: In addition to providing the required strength and stiffness, the structural box almost always has to provide fuel space. Integral tanks, as opposed to separate internally supported types, are preferred since their use enables the maximum advantage to be taken of the available volume. Integrally machined or moulded constructions, which use a small amount of large components, are obviously an advantage since sealing is reduced to a minimum. The major problem occurs at tank end ribs, particularly in the corners of the spar web and skins, and at lower surface access panels. The corner difficulty is overcome by using special “suitcase” corner fittings. Access panels should be large enough for a person to get through so that the inside can be inspected and resealed if necessary. On shallow section wings, the access has to be in the lower surface so that the operator can work in an acceptable way even if the depth is insufficient to climb in completely. Apart from the sealing problems, lower surface access panels are in what is primarily a tension skin and so introduce stress concentrations in an area where crack propagation is a major consideration. The access panels are arranged in a span-wise line so the edge reinforcing can be continuous and minimum stress concentration due to the cut-outs. Access panels are often designed to carry only shear and pressure loads, the wing bending being reacted by the edge reinforcing members. A deep wing can avoid these problems by using upper surface access panels but this is not a preferred aerodynamic solution. 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: v 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. v The control hinge loads and the resulting shear forces and bending moments should be equalized as far as is possible. v Structural failure of a single hinge should be tolerated unless each hinge is of fail-safe design and can tolerate cracking one load path.

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These points suggest the use of a relatively large number of discrete hinges but there are difficulties associated with this solution there are the obvious loads likely to be induced in the control by the distortion under load of the main surface to which it is attached may be significant. These problems do not arise if only two hinge points are used as any span-wise distortion or misalignment can be accommodated by designing one of the hinges so that it can rotate about a vertical axis. When more than two hinges are used the „floating hinge concept cannot fully overcome the problems. However, it is possible to design the control surface so that it is flexible in bending and indeed the more hinges there are the easier this is to accomplish. One hinge must always be capable of reacting side loads in the plane of the control surface. The hinges are supported near to the aft extremities of the main surface ribs. 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. Thus the structural layout may consist of an integral root rib or pivot or stub spar arrangement to which is attached a number of shear webs fanning out towards the extremities of the surface, possibly in conjunction with full depth honeycomb. High skin shear loading is inevitable due to the need to bring the loads to the two concentrated points. Shear loads due to torsion may be limited by locating the operating point on the root rib some distance away from the pivot. 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 local geometry allows minimizing 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 majority of flaps and slats are split into span wise segments of no greater lengths than can be supported at two or three locations. As with control surfaces, the locations of the support points are established so as to minimize local deformations since the various slots are critical in determining the aerodynamic performance. Sometimes the actuation may be located at

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a different pan wise position from the support points. This is often a matter of convenience, layout clearances, and the like. 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. Figure shows a cross section of a typical slotted flap of metal construction but the same layout applies if composite materials are used. In many cases the slipstream or afflux from power plants impinges upon a flap and this is likely to require special consideration in the design. Additional stiffness is not necessarily the answer because acoustic fatigue characteristics are often worse at higher panel frequencies. However the extensive local support offered by sandwich construction, either in panel or full depth configuration, is usually beneficial. This leads naturally to the application of reinforced plastic materials. Trailing edge flaps tends to be prone to damage by debris thrown up by the landing gear and it may be desirable to use Kevlar or glass rather than carbon fibers for the lower surface, but material compatibility needs to be considered. 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, the resulting constraint stresses being allowed for. It is usual for the wing structure of large aircraft to include a production joint at the side of the fuselage and this is virtual essential for swept wings. 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. Similar remarks also apply to the attachment of the horizontal stabilizer when the incidence setting is fixed. If the surface is also used for trimming or control, some special consideration is necessary in the location of the pivot and actuation fittings. These usually require a relatively heavily loaded rib or a pair of ribs, and where possible at least one of the attachment points should be close to the rib or spar intersection. It is desirable to arrange for the lateral distance between the pivots to be as great as possible to minimize pivot loads resulting from asymmetric span-wise loading. When the controls are manually operated, it is simplest if the elevator-hinge line and pivot coincide. Fins are usually built integrally with the rear fuselage. This is mainly due to the different form of loading associated with the geometric asymmetry.

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Flutter: Flutter as the dynamic instability of an elastic body in an airstream. It is found most frequently in aircraft structures subjected to large aerodynamic loads such as wings, tail units and control surfaces. Flutter occurs at a critical or flutters speed Vf which in turn is defined as the lowest airspeed at which a given structure will oscillate with sustained simple harmonic motion. Flight at speeds below and above the flutter speed represents conditions of stable and unstable (that is divergent) structural oscillation, respectively. Generally, an elastic system having just one degree of freedom cannot be unstable unless some peculiar mechanical characteristic exists such as a negative spring force or a negative damping force. However, it is possible for systems with two or more degrees of freedom to be unstable without possessing unusual characteristics. The forces associated with each individual degree of freedom can interact, causing divergent oscillations for certain phase differences. The flutter of a wing in which the flexural and torsional modes are coupled is an important example of this type of instability. Some indication of the physical nature of wing bending–torsion-flutter may be had from an examination of aerodynamic and inertia forces during a combined bending and torsional oscillation in which the individual motions are 90 out of phase. In a pure bending or pure torsional oscillation the aerodynamic forces produced by the effective wing incidence oppose the motion; the geometric incidence in pure bending remains constant and therefore does not affect the aerodynamic damping force, while in pure torsion the geometric incidence produces aerodynamic forces which oppose the motion during one-half of the cycle but assist it during the other half so that the overall effect is nil. Thus, pure bending or pure torsional oscillations are quickly damped out. This is not the case in the combined oscillation when the maximum twist occurs at zero bending and vice versa; i.e. a 90 phase difference. The type of flutter described above, in which two distinctly different types of oscillating motion interact such that the resultant motion is divergent, is known as classical flutter. Other types of flutter, non-classical flutter, may involve only one type of motion. For example, stalling flutter of a wing occurs at a high incidence where, for particular positions of the span wise axis of twist, self-excited twisting oscillations occur which, above a critical speed, diverge. Aileron Buzz: Another non-classical form of flutter, aileron buzz, occurs at high subsonic speeds and is associated with the shock wave on the wing forward of the aileron. If the aileron oscillates downwards the flow over the upper surface of the wing accelerates, intensifying the shock and

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resulting in a reduction in pressure in the boundary layer behind the shock. The aileron, therefore, tends to be sucked back to its neutral position. When the aileron raises the shock intensity reduces and the pressure in the boundary layer increases, tending to push the aileron back to its neutral position. At low frequencies these pressure changes are approximately 180 out of phase with the aileron deflection and therefore become aerodynamic damping forces. At higher frequencies a component of pressure appears in phase with the aileron velocity which excites the oscillation. If this is greater than all other damping actions on the aileron a high frequency oscillation results in which only one type of motion, rotation of the aileron about its hinge, is present, i.e. aileron buzz. Aileron buzz may be prevented by employing control jacks of sufficient stiffness to ensure that the natural frequency of aileron rotation is high. Buffeting: Buffeting is produced most commonly in a tail plane by eddies caused by poor airflow In the wing wake striking the tail plane at a frequency equal to its natural frequency; a resonant oscillation having one degree of freedom could then occur. The problem may be alleviated by proper positioning of the tail plane and clean aerodynamic design.

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7 CONCLUSION In conclusion, the series of short range aircrafts incorporated many unique design of future that was never seen on an operational aircraft. The design of these aircrafts points the way for the design of future of very high mach airplanes. The airplane has gone through many design modifications since its early conceptual designs expected, among these was a growth in weight. The document to provide information on the trends in various aircraft characteristics that may influence general long-term airport planning and design. These are strong indications that future trends could see the coexistence of very high capacity aircraft modules of similar capacities for the long range/very long range operations. Cargo payloads, which include mail, express and freight, are increasing in size and weight as larger aircraft service with the airlines, To ensure continued growth in payload and the profitability of cargo operations, improvements in methods, equipment and terminal facilities will be required in order to reduce cargo handling costs and aircraft ground time and to provide improved service for the shippers. We have enough hard work for this design project. A design never gets completed in a flutter sense but it is one step further towards ideal system. But during the design of this aircraft, we learnt a lot about aeronautics and its implications when applied to an aircraft design.

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BIBLIOGRAPHY

Ø “AIRPLANE AERODYNAMICS AND PERFORMANCE” by JAN ROSKAM Ø E.F. Bruhn, “Analysis and Design of Flight Vehicle Structures”, Tristate Offset Co., 1980. Ø Lloyd R. Jenkinson and James F.Marchman III., “Aircraft Design project”, ButterworthHeinemann., Burlington, 2003. Ø Anderson, John D., Jr:Aircraft Performance and Design, McGraw-Hill, Boston, 1999. Ø Megson, T.M.G; Aircraft Structures for Engineering Students, Edward Arnold, 1989 Ø Peery, D.J. and Azar, J.J., Aircraft Structures, 2nd Edition, McGraw-Hill, New York, 1993. Ø McCornic, B.W, “Aerodynamics, Aeronautics & Flight Mechanics”. John Wiley, 1995. Ø G. Corning, “Supersonic & Subsonic Airplane Design”, II Edition, Edwards Brothers Inc., Michigan, 1953. Ø Ira h. Abbott, Albert e. Von Doenhoff,and Louis S. Stivers, Jr,”Summary of Airfoil Data”, National advisory committee for aeronautics, 1947. Ø John T.Lowry., “Performance of Light Aircraft”, American institute of aeronautics and astronautics, Washington.D.C, 1935. Ø Dr.lng.S.F.Hoerner,”Fluid –Dynamic Drag”,Hoerner fluid dynamics., 1992. Ø J.B. Russell,” Performance and Stability of Aircraft”. ISBN 0-340-63170-8. Arnold 1996. Ø Mark D. Ardema, Mark C. Chambers, Anthony P. Patron, Andrew S. Hahn, Hirokazu Miura, and Mark D. Moore, “Analytical Fuselage and Wing Weight Estimation of Transport Aircraft”,1996.

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WEBSITE REFERENCE ª http://www.worldofkrauss.com/ ª http://faculty.dwc.edu/sadraey/V-n%20diagram.pdf

ª http://www.aerostudents.com/files/aircraftStressAnalysisAndStructura lDesign/reader.pdf ª http://www.cta-dlr2009.ita.br/Proceedings/PDF/60272.pdf ª http://www.emteq.com/aircraft-structural-analysis-modifications.php ª http://www.biznet.org.au/member.asp?id=1094&pid=184 ª http://www.cp.berkeley.edu/cds_ucb/UCB-05100.pdf ª http://www.aer.ita.br/~bmattos/download/fuselagem-design.pdf ª http://adg.stanford.edu/aa241/fuselayout/sstfuse.html ª http://www.mat.ethz.ch/news_events/archive/materialsday/matday01/ pdf/TempusMD.pdf ª http://www.free-online-private-pilot-ground-school.com/aircraftstructure.html ª http://www.zenithair.com/stolch701/7-design-fuselage.html ª http://www.nusil.com/products/engineering/aircraft/documents/Aircra ft%20Selection%20Guide.pdf ª http://www.scielo.oces.mctes.pt/pdf/ctm/v20n3-4/v20n3-4a11.pdf ª http://www.faa.gov/library/manuals/aircraft/amt_handbook/media/FA A-8083-30_Ch05.pdf ª http://www.ppart.de/aerodynamics/profiles/NACA5.html ª http://www.desktop.aero/appliedaero/airfoils1/airfoilgeometry.html ª http://www.mh-aerotools.de/airfoils/jf_users_manual.htm

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