ADP

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A PROJECT REPORTSubmitted by

PRAVEEN.K 96409101047RAM GANESH.T 96409101054VIGNESH.K 96409101073VISWANATHAN.A 96409101075VIVEK.R 96409101076

In partial fulfillment of the requirement for the award of the degree of

BACHELOR OF ENGINEERINGIn

AERONAUTICAL ENGINEERING

PSN COLLEGE OF ENGINEERING AND TECHNOLOGYANNA UNIVERSITY : CHENNAI 600 025

AIRCRAFT DESIGN PROJECT – 2BUSINESS AIRCRAFT

NOVEMBER 2012

BONAFIDE CERTIFICATE

Certified that the project report “DESIGN PROJECT ON BUSINESS JET

FIGHTER AIRCRAFT”, is the bonafide work of PRAVEEN.K, RAM

GANESH.T,VISWANATHAN.A,VIGNESH.K,VIVEK.R, who carried out the

project work under my supervision.

SIGNATURE SIGNATURE

HEAD OF THE DEPARTMENT SUPERVISOR

Asst. prof. Mr.Suresh Kumar.M.E., Asst. prof. Ms.D.Shruthi.M.E.,

Dept. of Aeronautical Engineering Dept. of Aeronautical engineering

PSN COLLEGE OF ENGG&TECH PSN COLLEGE OF ENGG&TECH

Tirunelveli. Tirunelveli.

Submitted for the B.E project work Viva – Voce held at PSN College of

Engineering and Technology, Tirunelveli, on ……………..

INTERNAL EXAMINER EXTERNAL EXAMINER

ACKNOWLEDGEMENT

We express our deep and sincere thanks to chairman of our college

Dr.P. Suyambu, for giving us the inspiration and for making all the facilities

during the identification of this project.

our college for having provided the necessary infrastructure for the successful

completion of the project.

We are much greatful to Asst.Prof. Mr.Suresh Kumar, Head of the Aeronautical

Department for this encouraging decision, valuable comments and many

innovative ideas in this project. Without his help, it would have been impossible

for us to complete the report.

We acknowledge in no less terms the qualified and excellent assistance rendered

by Asst. Prof. Ms. Shruthi, Department of Aeronautical Engineering. We owe a

debt of gratitude for his valuable suggestion, kind inspiration and encouragement.

We most sincerely thank our staff members, for their constant inspiration and

suggestions.

ABSTRACT

The purpose of the project is to make the detailed design of the

BUSINESS AIRCRAFT which is preliminarily designed in Aircraft Design

Project-1.This aircraft will possess a low wing attachment, cylindrical fuselage

with cabin, tricycle landing gear arrangement and a T-tail arrangement. In

this project a detailed study of load and structural analysis of wing is made

including the Flight Envelope curves for Maneuver and Gust Loads. Layout

of fuselage along with cabin design is also designed. Design of landing gear

is made along with the amount of load that can be taken by the landing

gear. Critical loading performance analysis of the aircraft is studied and then

Shear Force & Bending Moment for the aircraft is calculated respectively and

diagrams are drawn. Finally a three view diagram of our aircraft is drafted using

Auto-Cad software with exact dimensions.

Thisproject is centered towards a design of efficient and safe business

jetaircraft.The objectiveof this project is to provide a new better design by

manipulating the previous designs of various aircrafts.

INTRODUCTIONDesign is a process of use age of creativity with the knowledge of science where we try to get the best things available and to overcome the pitfalls the previous design has. It is an iterative process to idealism towards which everyone is marching still.Design of any is system is successfully application of fundamental physics .Thus the airplane design incorporates the fundamentals of aerodynamics,structures, performance and stability & control and basic physics .These are based on certain degree of judgment and experience . Every designer has the same technical details but each design retains its own individuality and the mode of the designer.Here the preliminary design has been done of an executive transport aircraft. The basic requirements are the safe, comfortable and economic transport mode with reasonable timeperiod of flight. Here comfort and safety are given primary importance.Here the most possible considerations have been taken .The flight parameters limitations are studied. The modern day calls for the need of latest aircraft for the use of passenger transport which aims mainly at improving the aerodynamic characteristics as well as the passenger comfort. This design project also looks at the above aspects in a lot more closer way. Also the design project has been classified into different stages for easier approach and achieving better performance. The different stages in our design will be as follows.1 .Introduction to V-n diagram.2 .V-n diagrams for our design study.3. Gust and maneuverability envelops.4 .Critical loading performance and final V-n graph calculation.5. Structural design study.6. Load estimate of fuselage.7. Load estimation of wings.8. Balancing and maneuvering loads and tail plane, aileron and rudder loads.9. Details structures layouts.10. Design of some components of wings and fuselage.11. Preparation of a detailed design report with drawings.

TABLE OFCONTENTS

CHAPTER TITLE PAGE NO.

ACKNOWLEDGEMENT iii

ABSTRACT iv

TABLE OF CONTENTS

LIST OF SYMBOLS xi

1. STUDY OF V-n DIAGRAM

1.1 INTRODUCTION

1.2 LIMIT LOAD FACTOR

1.3 ULTIMATE LOAD FACTOR

1.4 VELOCITY LOAD FACTOR DIAGRAM

2. GUST AND MANEUVER ENVELOPES

2.1 DETRMINATION OF +1g STALL SPEED

2.2 DESIGN CRUISE SPEED

2.3 DESIGN DIVING SPEED

2.4 DESIGN MANEUVERING SPEED

2.5 FOR GUST DIAGRAM

2.6 FOR GUST LINE MARK(VC)

2.7 FOR GUST LINE MARK(VD)

2.8 FOR GUST LINE MARK(VB)

3. AIRCRAFT LOADS

3.1 AIR LOADS

3.2 LANDING

3.3 INERTIA LOADS

3.4 POWER PLANT

3.5 TAKE OFF

3.6 TAXI

3.7 MANEUVERING LOADS

3.8 GUST LOADS

3.9 AIR LOADS ON WING

4. LOAD ESTIMATION OF WINGS

4.1 CHORD LENGTH

4.2 AIR LOADS

4.3 TO FIND THE VALUE OF K

4.4 TO FIND LIFT LOAD INTENSITY

4.5 TO FIND THE STRUCTURAL LOAD

4.6 TO FIND THE RESULTANT LOAD

4.7 SHEAR FORCE DIAGRAM

4.8 BENDING MOMENT DIAGRAM

5. WING STRUTURAL LAYOUT

5.1 SPECIFIC ROLES OF WING STRUCTURE

5.2 BASI FUNCTIONS OF WING STRUCTURAL

MEMBERS

5.2.1 SPARS

5.2.2 SKIN

5.2.3 STRINGERS

5.2.4 RIBS

5.3 WING BOX CONFIGURATION

5.4 TYPES OF SPARS

5.5 RIBS

5.6 RIBS CONSTRUCTION AND

CONFIGURATION

5.7 RIB ALLIGNMENT POSSIBILITIES

6. STRUCTURAL ANALYSIS OF WING

6.1 SPAR CALCULATIONS

6.2 SPAR LOCATIONS

6.3 BENDING MOMENT CALCULATIONS

6.4 BENDING MOMENT DISTRIBUION

6.5 TO DIMENSIONS FOR EACH SPAR

6.5.1 FRONT SPAR

6.5.2 MID SPAR

6.5.3 REAR SPAR

6.6 CALCULATION OF BENDING STRESS

6.7 SHEAR FLOW ANALYSIS ON WINGS

7. BALANCING AND MANEUVERING LOADS ON

TAIL PLANE, AILERON AND RUDDER

7.1 HIGH LIFT SYSTEMS

7.2 FLAPS

7.3 SLOTS AND SLATS

7.4 DEFLECTED SLIPSTREAM AND JET WASH

7.5 LEADING EDGE DEVICES

7.6 WING TIPS

8. FUSELAGE LAYOUT ANALYSIS

8.1 CALCULATION OF WEIGHT OF FUSELAGE

8.2 SPECIFICATION OF OUR FIGHTER

AIRCRAFT

8.3SEAT DIMENSIONS

9. DESIGN OF SOME OF COMPONENTS IN

AIRCRAFT( LANDING GEAR)

9.1 DESIGN OF LANDING GEAR SELECTION

9.2 TYPES OF LANDING GEAR

9.3 ARRANGEMENTS OF LANDING GEAR

9.4 TYRE SIZING

9.5 PERFORMANCE PARAMETER

9.6 GEAR RETRACTION GEOMETRY

9.7 LANDING GEAR WEIGHT

9.8 LANDING GEAR DIMENSIONS

9.9 TO LOCATE THE LANDING GEAR

9.10 TO FIND THE CENTER OF GRAVITY

9.11 TO FIND THE LOAD ACTING ON THE

LANDING GEAR

9.12 STICK DIAGRAM

10. CRITICAL LOADING PERFORMANCE

10.1 FITTING AND CONNECTIONS

10.2 BOLTS AND RIVETS

10.3 SECONDARY IN FITTING DESIGN

10.4 GENERAL RULES

10.5 JOGGLED MEMBERS

10.6 PROTRUDING HEAD

10.7 FLUSH HEAD

10.8 SHEAR CLIPS

10.9 AIRCRAFT NUTS

10.10 FAILURE BY INTER RIVET BUCKLING

10.11 RIVET WRINKLING

10.12 AIRWORTHINESS REQUIREMENTS

10.13 TO DETERMINE THE WEIGHT OF THE WING

10.14 DETERMINATION OF SHEAR FORCE AND

BENDING MOMENT DIAGRAM

11. THREE VIEW DIAGRAM OF AN AIRCRAFT

LIST OF SYMBOLS

n Load Factor

IAS Indicated Air Speed

ρ Density

V∞ Free Stream Velocity

S Surface Area of the Wing

CL Co-efficient of Lift

W Weight of the Aircraft

CLmax Maximum Co-efficient of Lift

CNmax Maximum Yawing Co-efficient

M Mach Number

VS +1g Stall Speed

VA Design Maneuvering Speed

VC Design Cruise Speed

VD Design Diving Speed

WFDWG Flight Design Gross Weight

n Load Factor

CLmax Maximum Lift Co-efficient

Kg Gust Elevation

h Altitude of an Aircraft

AOA Angle Of Attack

g Acceleration due to gravity

b Span of the wing

WS,W Structural Weight of the wing

y0 Root Chord

Cx Chord at each Point

k Lift Load Intensity

L Lift

W Weight

FOS Factor Of Safety

M Bending Moment

A Area

I Moment Of Inertia

x ,y Centroid

Q Shear Flow

ds Sectional Length

QS-O Corrected Shear Flow

LF Length of the Fuselage

DF Diameter of the Fuselage

LFC Length of the Fuselage Cone𝜃FC Angle of the Fuselage Cone

^ 25% Mean Aerodynamic Chord Swept angle

SF Wetted Surface Area of the Fuselage

WF Weight of the Fuselage

WW Weight of the Landing Gear

WTO Take-Off Weight

C.G Center of Gravity

P Load

lm,ln Length of Landing Gear distance

R Reaction Force

S Shear Force

CHAPTER-1

STUDY OF V-n DIAGRAM

1.1 INTRODUCTION:

Airplanes may be subjected to a variety of loading conditions in flight.The

structural design of the aircraft involves the estimation of the various loads on the

aircraft structure and designing the airframe to carry all these loads,providing

enough safety factors, considering the fact that the aircraft under design is a fighter

aircraft.It is obviously impossible to investigate every loading condition that the

aircraft may encounter, it becomes necessary to select a few conditions such that

each one of these conditions will be critical for some structural member of the

airplane.

Using the V-n diagram to important load factor values can be plotted which

are

1.2 LIMIT LOAD FACTOR:

Value of load factor corresponding to which there is permanent structural

deformation.

1.3 ULTIMATE LOAD FACTOR:

Value of load factor corresponding to which there is out right structural

damage/failure.

1.4 VELOCITY - LOAD FACTOR (V-n) DIAGRAM:

The various external loads on the airplane are usually represented on a graphof the

limited load factor and plotted against the indicated air speed (IAS).

FIGURE-1

This diagram is known better as the V-n diagram. The indicated air speed

used since all air loads are proportional to q. The value of q is the same density for

air and actual air speed at altitude, as it is for the standard sea level density and

IAS.

Then V-n diagram is therefore the same for all altitudes of indicated air

speeds are used. However in this design case, corrections involving compressibility

have to be taken in to consideration while calculating the time true air speeds from

indicated air speeds. Therefore calculations involving high speeds have been

performed with respect to sea level conditions only

n = ½ ρV∞2SCL/ W

The load factor is basically the ratio of wing lift produced to the weight of

the aircraft and Hence, represents the amount of acceleration produced along Z

axis of the plane for fighter aircraft. The ultimate positive load factor ranges from

3.1 to 4.4 and negative load factor between -1.25 and -2.3. The positive and

negative load factors are arbitrarily chosen as 3.1 and -1.25 respectively.

For level flight an unit load factor the value of V corresponding to

CL max would be the stability speed of the airplane. In accelerated flight the

maximum lift coefficient can be achieved at higher speeds. The wing is usually

analyzed for a coefficient of 1.25 CL maxand various values of n are obtained by

varying the velocity, until the ultimate positive load factor is reached. It can be

made out from this boundary that it is impossible to maneuver at speeds and load

factors corresponding to points above or to the left of line because this would

represent positive high angle of attack (+HAA). This load factor is usually arrived

at by considering both aerodynamics and structural design capabilities.

In a similar manner, the maneuver boundary can be carried to the negative

load factor region which is indicative of inverted flight. The negative maneuver

boundary is seldom made use of in transport aircraft. However the gust load in the

negative region are indispensable and can be more severe than the maneuver load

factor itself.

Thus, In order to establish the safe flight envelope of the aircraft we have

plotted as per FAR-25 norms

RESULT:

Thus the V-n diagram has been studied successfully.

CHAPTER 2

CALCULATION OF V-n DIAGRAM FOR THE DESIGNED AIRCRAFT

2.1DESCRIPTIONThe V-n diagram is very much useful for determining the structural and load

limitations of an aircraft. It is used to find at what velocity the aircraft can travel without any structural damage. It also tells the maximum loading of aircraft which it can withstand without any damage. It also tells us about the maximum possible lift without any failure or damage.

2.2FORMULAS

For positive load factor [n (+)],Velocity (V) = √2n (w/s)/ρCL

For negative load factor [n (-)],Velocity (V) = √2n (w/s)/ρCL*0.75

2.3CALCULATION

Wing loading = W/S N/m2

Wing loading (W/s) = 258.9 N/m2

ρ = 1.225kg/m2

CL = 0.434V∞ = √2 (3.8)*258.9/1.225*0.434 = 60.83 m/sV∞ = √2 (3)*258.9/1.225*0.434 =54.25 m/sV∞ = √2 (2)*258.9/1.225*0.434 = 44.13 m/sV∞ = √2 (1)*258.9/1.225*0.434 = 31.20 m/sV∞ = √2 (0)*258.9/1.225*0.434 = 0 m/sV∞ = -√2 (1)*258.9/1.225*0.434 = -4.335 m/sV∞ = -√2 (1.5)*258.9/1.225*0.434 =- 3.75 m/s

2.4TABULATION:

N V∞

m/s

+3.8

+3

+2

+1

0

-1

-1.5

60.83

54.25

44.13

31.20

0

4.335

3.75

RESULT

Thus the Vn diagram is drawn for our designed aircraft.

CHAPTER 3

GUST AND MANEUVERABILITY ENVELOPE

DESCRRIPTION:In aerodynamics, the flight envelope, service envelope, or performance

envelope of an aircraft refers to the capabilities of a design in terms of airspeed and load factor or altitude. The term is somewhat loosely applied, and can also refer to other measurements such as maneuverability. When a plane is pushed, for instance by diving it at high speeds, it is said to be flown "outside the envelope", something considered rather dangerous.

FORMULAS:

Positive stall velocityVS (+) =√2 (W/S)/ρCL max

Negative stall velocityVS (-) =√2 (W/S)/ρCL max*0.75

Critical VelocityVCR = Vs (+) +100

Dive VelocityVD=1.5Vcr

Gust velocity with constantUde (stall) =20xVs

Ude (critical) =15xVcr

Ude (D) =7.5xVD

Velocity of gustU=Kde

Where,K- Gust effectivenessϻ-mass ratio

∆n=ρVCLαU/2(W/S)Where,∆n-change in load factorU-velocity of gust

TABULATION:

Sl.No. V (m/s) Ude U(m/s)=K Ude

∆n 1+∆n 1-∆n

1 Vs=31.20 624 517.116 11.51 12.51 -10.51

2 Vcr=131.20 1968 1630.88 152.71 153.71 -151.71

3 VD=196.8 1476 1223.16 171.80 172.80 -170.80

RESULT

Thus the gust and maneuvering envelope for Vn diagram is drawn.

CHAPTER 4

CRITICAL LOADING PERFORMANCE

4.1 FITTING AND CONNECTIONS:

Shock absorber analysis

Fuselage bending stress analysis

Various failures

Airworthiness requirements

4.2 BOLTS AND RIVETS:

The connection involves primary and secondary connection such as

fittings, bolts, rivets, welding etc. No tough that main or primary fitting involves

more weight and costs than any other aerospace structure.

4.3 SECONDARY IN FITTING DESIGN:

In a wing structure fitting involving in main load carrying structure is

move costly to design as well as to manufacture for economy at fabrication the

structural designed should have a fabrication the structural designed should have a

good knowledge of shop processes and operations .

4.4 GENERAL RULES:

4.4.1USAGE OF BOLTS

Bolts thread should not be placed in or shear the length of the bolts

should be such that not more than one thread external fitting surface which can be

done by washer.

Bolts less than 3/8’’ inch dia should not be used in major fittings.

For steel bolts 3/10 inch should be small size to be used.

Bolts connecting parts having relative motion or stress reversal should have

closed tolerance to decrease shock loads.

Protruding head level of rivet.

The flush type rivet.

For many years round head rivet was, all sections but wind tunnel

experiment. It was that these produce more drag so rivet head are changed to flush

type which produce lesser drag.

4.5 JOGGLED MEMBERS:

A joggle is an offset formed in a member it usually involves one or more

flanges of a member of open cross section type joggles are quite common in

typical airplane structured they are used most often when it is desired to fasten to

gather two intersecting members without using an extra part at joint there will be

slight loss in stiffness of joggled member.

4.6 PROTRUDING HEAD:

FIGURE-17

4.7 FLUSH HEAD:

4.8 SHEAR CLIPS:

There are hundreds of these in a typical military airplane they are used

for joining together both primary and secondary structural components such as

equipments mounting brackets etc. The function of shear clip is to transfer shear

load from one part to another .It is not intended to transfer axial load or bending or

twist.

Fillers are used to fill up a void when they become a part of the structural

path that they held particular attention quite common in complicated metal

structure.

4.9 AIRCRAFT NUTS:

Three standard steel nuts shown in figure nut material is more ductile than bolt

material, thus when the nut is tightened the threads will deflect to seat on the bold

thread. The nut is probably the most common aircraft nut.

FIGURE-19

The shear nut is one half as thick the cast head nut and has threads only

enough to develop one half bolds tensile stress.

4.10 FAILURE BY INTERRIVET BUCKLING:

The effective sheet area is considered to act monolithically with

stiffness. However if the rivets are spot welded that fasten the rivet to the stiffener

are spaced to far apart sheet will buckle before the crippling stress of the stringers

is placed. In order to prevent this sort of buckling rivet spacing has to be selected

on the upper surface of the wing. Rivet spacing is closer than on the lower surface

because the compressive loads act on the top of the wing.

4.11RIVET WRINKLING:

The rivet spacing is relatively large the sheet will start buckling belt rivets this

buckling belt rivets this buckling does not deform in flange to which the sheet is

attached. The rivet spacing is such that prevent the inter-rivet buckling then the

failure of the sheets occurs by wrinkling. It is also known as forced wrinkling it is

also known as forced wrinkling of the riveted pads.

4.12 AIRWORTHINESS REQUIREMENTS:

Airworthiness of an aircraft is concerned with the standards of safety incorporated

in all aspects of its construction. These range from structural strength to the

provision of certain guards in the event of crash landing and include design

requirements related to aerodynamics, performance and electrical hydraulic

systems. The selection of minimum standards of safety is largely the concern of air

worthiness authorities who prepare handbooks of official requirements. In UK the

relevant publications are AVP970 for military aircraft and British civil air

worthiness requirements of civil aircraft. The handbooks include operation

requirements, minimum safety requirements recommended practices and design

data.

Clearly airworthiness implies a certain level of safety like saying that

the ship is sea worthy and it takes little fore thought to release that there must be

some yardstick against which air worthiness can be assessed. We might start with a

general, all embracing design requirements. An airplane shall be designed and built

to fly safely. Unfortunately we cannot then dust our hands and get on with the job,

believing that in one swipe we have got rid of government and other official

interference and struck below for freedom.

To be awarded a certificate of air worthiness an aircraft must be

demonstrated to be air worthy. Air worthiness can be defined as the contribution

made by the aircraft itself to be safety of the flight when the pilot has been

removed from the man machine loop. It is concerned with those aspects of design,

construction maintenance and the provision of all related limitations and essential

information which together determine fitness for flight, thus aircraft of

Airworthiness is awarded to an aircraft and its equipment, although under certain

circumstances the award may also be conditional upon the aircraft being operated

under the control of certain named persons or perhaps just one individual.

RESULT

Thus the critical loading performance has been studied successfully.

CHAPTER 5

STRUCTURAL DESIGN STUDY

5.1 WING DESIGN The front of the airfoil is defined by a leading-edge radius that is tangent to the upper and lower surfaces. an airfoil designed to operate in supersonic will have a sharp or nearly- sharp leading edge to prevent a drag producing bow shock. the chord of the airfoil is the straight line from the leading edge to the trailing edge. it is very difficult to build a perfectly sharp trailing edge, so most airfoils have a blunt trailing edge with some small finite thickness. “camber” refers to the curvature characteristic of most airfoils. the “mean camber line” is the line equidistant from the upper and lower surface. total airfoil camber is defined as the maximum distance of the mean camber line from the chord line, expressed as a percent of the chord. the thickness distribution of the airfoil is the distance form the upper surface to the lower surface, measured perpendicular to the mean camber line, and is function of the distance from the leading edge. the “airfoil thickness ratio” refers to the maximum thickness of the airfoil divided by its chord. 5.2 FUSELAGE: once the takeoff gross weight has been estimated, the fuselage, the wing. and tail can be sized. many methods exist to initially estimate the required fuselage size. for certain types of aircraft, the fuselage size is determined strictly by “real world constraints”. for example, a large passenger aircraft devotes most of its length to the passenger compartment. once the number of passengers is known and the number of seats across is selected, the fuselage length and diameter are essentially determined. 5.3 WING: actual wing size can now be determined simply as the take off weight divided by takeoff wing loading. remember that this the reference area of the 35

THEORETICAL, TRAPEZOIDAL WING, AND INCLUDES THE AREA EXTENDING INTO THE AIRCRAFT CENTER LINE. 5.4 TAIL VOLUME CO-EFFICIENT:

For the initial layout, the historical approach is used for the estimation of the tail size. The effectiveness of a tail in generating a moment about the centre of gravity is proportional to the force produced by the tail and to the tail moment arm. The primary purpose of the tail is to counter the moments produced by the wing

5.5ENGINE SELECTION t/w directly affects the performance of the aircraft. an aircraft with a higher t/w will accelerate more quickly, climb more rapidly, reach higher maximum speed, and sustain higher turn rates. on the other hand, the larger engines will consume more fuel throughout mission, which will drive up the aircraft’s take-off gross weight to perform the design mission. t/w is not a constant. the weight of the aircraft varies during flight as fuel is burned. also, the engine’s thrust varies with altitude and velocity (as does the horsepower and propeller efficiency). t/wto ratio for jet transport aircraft is 0.4 overall weight of aircraft wto = 8125 kg 5.6 THRUST CALCULATION OF T/W RATIO (T/WTO) = 0.4 where (t/w)to = thrust to weight ratio at take off t = 0.4 * wto

t = 0.4 * 8125 t = 3250 kg = 31.87 kn since the thrust produced by the aircraft engine decreases with altitude due to decreasing air density, thus:- 41

tcruise = 0.25 tto

thrust required = 31.87 k n for entire aircraft. hence the thrust per engine is = 31.87/2 = 15.93 kn from the literature survey the nearest value of the thrust of an aircraft is Learjet 35a the characteristics of the selected engine of the aircraft is, thrust per engine=16 KN

number of engine = 2

type of engine = Garrett tfe731-2-2b turbofan engine

total thrust = 32 KN

Therefore, the engines selected are two Garrett tfe731-2-2b turbofan engines which is nearer to the required thrust per engine. 5.7 LOCATION OF ENGINE: All civilian aircraft are required by law to possess multiple engines so that in case

of failure of any one engine the aircraft can continue to fly on the other engine. The

position to locate the engines for Business Jet Aircraft is on the rear fuselage to

reduce noise and due to low wing configuration. The mounting of the engines on

the rear fuselage leads to the fuselage requiring to be strengthened.

RESULT

Thus the structural design study of our aircraft has been studied.

CHAPTER 6

LOAD ESTIMATION OF FUSELAGE

6.1 CALCULATION OF WEIGHT OF FUSELAGE:

FOR OUR FIGHTER AIRCRAFT,

Length of the fuselage

LF=AWOC

Constant values are,

A=0.286

C=0.43

LF=0.286(8125)0.43

LF=13.77m

For our aircraft,

LF/DF=9

DF=2m

LFC/DF=6.885

LFC=7m

θ FC=11o

For business aircraft, There will be 2 inches breadth thickness frames.

The view of the pilot will be 15o from the pilot seat.

The Head Up display will displayed on 45otop of the pilot seat.

WFUSELAGE=0.328 KDOOR KLG[WOn2]0.5 LF0.5SF

0.302[1+KWS]0.5[L/D]0.01

For business aircraft,

KDOOR=1.08

KLG=1

KWS=0.96[Wingspan * 3.28] tan ^/2

^ =55.23o

n2=1.5*nlimpos

^ = 25% meanaerodynamic chord swept angle.

W0=Overall weight

SF=Fuselage wetted surface

Sub all this values in the WF&Kws

KWS= 5.21275

WFUSELAGE=883 kg

HUD DISPLAY(Angle= 450)

PILOT VIEW (Angle= 150)

FIGURE-11

6.2 SPECIFICATION OF OUR FIGHTER AIRCRAFT:

The seat dimensions and locations can be designed.

The

The aircraft pilot should have the front view at the 15˙

The aircraft have the HUD at the 45˙ above the top view.

The fuselage will have the sharp edge at the leading edge of the

fuselage to produce or move with high speeds.

The fuselage will have the low fuselage cone angle because of shirter

length compared to the transport aircraft.

The seat will be fixed with some fitted angle to eject seat at the

dangerous condition.

The fuselage will have buried engine.

The fuselage will have the canopy above the pilot seat.

6.3 SEAT DIMENSIONS:

FIGURE-12

RESULT:

Thus the Fuselage Layout analysis have been done successfully

CHAPTER-7

LOAD ESTIMATION OF WINGS

7.1 CHORD LENGTH:

πaB/4 =Planform area/2

a=Span of wing/2

a=14.995/2

a=7.45 m

Sub this value in the above equation

π*7.45*B/4=14.995/2

B=1.281 m

x2/a2+y2/b2=1

y=b2(1 – x2/a2)

Span=14.9 m is divided in to 20 equal parts.

For the taper wing for the business aircraft

Root chord=2.87 m

Tip chord=1.14 m

Varying the x value and find the value of chord length y where,

a=7.45 m

b=1.281 m

Structural weight of our one Wing ,WS,W= WTO/2=8125/2

WS,W=4062.5 kg.

Air load at Root=WS/2*y0/Area of Schrenk’s curve

Area of Schrenk’s curve= 20 Strip Area

Area of the Rectangle=bh

Area=11.175 m2

Air load at Root={(8125/2)*2.87}/11.175

Air load at Root=1043.34 kg

Air load at nth Point= Load at Root*yn/y0

Air load at nth Point=363.53*y

SPAN Y(m)0 1.281

0.375 1.2790.75 1.274

1.125 1.51.5 1.254

1.875 1.2392.25 1.221

2.625 1.1983 1.172

3.375 1.1423.75 1.107

4.125 1.0664.5 1.021

4.875 0.9685.25 0.909

5.625 0.8396 0.759

6.375 0.6636.75 0.542

7.125 0.33

TABLE-1

GRAPH-3

7.2 AIR

LOADS:

At Point 1

at x=0 m

Air

load=6938.7 kg

At Point 2

at x=0.375 m Air load=6927.75 kg

At Point 3 at x=0.75m Air load=6900.786 kg

At Point 4 at x= 1.125 m Air load=6857.42 kg

At Point 5 at x= 1.226 m Air load=6792.46 kg

At Point 6 at x= 1.875 m Air load=6711.109 kg

At Point 7 at x= 2.25 m Air load=6613.66 kg

At Point 8 at x= 2.625 m Air load=6489.09 kg

At Point 9 at x= 3 m Air load=6348.29 kg

At Point 10 at x= 3.375 m Air load=6191.09 kg

At Point 11 at x= 3.75m Air load=5996.07 kg

At Point 12 at x=4.125 m Air load=5774.04 kg

At Point 13 at x=4.5 m Air load=5537.04 kg

At Point 14 at x= 4.875 m Air load=5243.31 kg

At Point 15 at x=5.25 m Air load=4923.705kg

At Point 16 at x= 5.25 m Air load=4544.5 kg

At Point 17 at x=6 m Air load=4111.208 kg

At Point 18 at x= 6.375 m Air load=3591.19 kg

0 1 2 3 4 5 60

1

2

3

4

5

6

7

8

9

10

At Point 19 at x= 6.75m Air load=2935.74 kg

At Point 20 at x=7.125 m Air load=0 kg

7.3 TO FIND THE VALUE OF K:

WS,W=∫k CX2 dx

k - Constant

CX - Chord at each point

CX =a+bx

At x=0 CX =a= CR =2.87 m

At x=7.45 CX = CT

Substitute this value in equation

0=7.45+bx

WS,W=0.1*WTO/2

WS,W= 406.25kg for one side of the Wing.

406.25=k∫CX2 dx

Integrating this equation,

406.25=k CX2 b

The values of b and CX are known. CX vary with different points.

The value of b is 7.45 m

k=406.25/CX2b

k = 12.83

Yn LIFT LOAD INTENSITY STRUCTURAL LOAD RESULTANT LOAD NET1.281 465.68 99.36 75.32 174.591.279 464.95 93.25 81.11 174.351.274 463.14 87.33 86.34 173.671.266 460.23 81.6 90.98 172.581.254 455.87 76.07 94.88 170.951.239 450.41 70.73 98.11 168.91.221 443.51 65.58 100.87 163.321.198 435.51 60.6 102.69 159.771.172 426.061 55.88 103.89 155.681.142 415.15 51.32 104.36 150.911.107 402.43 46.95 103.96 145.32

1.066 387.52 42.77 102.55 139.371.021 371.67 38.79 100.58 131.960.968 351.9 35.014 96.946 123.920.909 330.45 31.42 92.5 114.3750.839 305 28.02 86.35 103.470.759 295.92 24.82 78.65 90.880.663 241.02 21.81 69.09 87.990.542 197.03 23.01 73.88

TABLE-2

7.4 TO FIND LIFT LOAD INTENSITY:

Lift on each element=Air load at grid point*Distance between grid point

7.5 SHEAR FORCE DIAGRAM:

0 2 4 6 8 10 120

100

200

300

400

500

600

700

800

900

GRAPH-7

SHEAR FORCE = WX2/2L

= (267.47 * 14.9^2)/(2*14.9)

= 1992.65 N

7.6 BENDING MOMENT DIAGRAM:

BENDING MOMENT = WX3/6L

= 9896.83 N/m

RESULT:

Thus the load estimation on wings have been done successfully.

CHAPTER-8

BALANCING AND MANEUVERING LOADS ON TAIL PLANE,

AILERON AND RUDDER

8.1 HIGH LIFT SYSTEMS – INTRODUCTION:

A wing designed for efficient high-speed flight is often quite different from

onedesigned solely for take-off and landing. Take-off and landing distances are

strongly influencedby aircraft stalling speed, with lower stall speeds requiring

lower acceleration or deceleration andcorrespondingly shorter field lengths. It is

always possible to reduce stall speed by increasingwing area, but it is not desirable

to cruise with hundreds of square feet of extra wing area (andthe associated weight

and drag), area that is only needed for a few minutes. Since the stallingspeed is

related to wing parameters. It is also possible to reduce stalling speed by

reducingweight, increasing air density, or increasing wing CLmax . The latter

parameter is the mostinteresting. One can design a wing airfoil that compromises

cruise efficiency to obtain a goodCLmax, but it is usually more efficient to include

movable leading and/or trailing edges so thatone may obtain good high speed

performance while achieving a high CLmax at take-off andlanding. The primary goal

of a high lift system is a high CLmax. However, it may also bedesirable to maintain

low drag at take-off, or high drag on approach. It is also necessary to dothis with a

system that has low weight and high reliability.This is generally achieved

byincorporating some form of trailing edge flap and perhaps a leading edge device

such as a slat.

8.2 FLAPS:

Wing flaps can be found on virtually every modern airplane. The effect of

adding flapsto the trailing edge of the wing is equivalent to increasing the camber

of the wing. Some flapdesigns also increase the chord length of the wing. This

increases the area of the wing so thatmore air is diverted, thus reducing the angle

of attack needed for lift. There are many types offlaps. In the 1930’s and 1940’s

the split flap, shown in Figure was introduced and was one of thefirst types of flap

to appear in production airplanes. Splitting the last 20 percent or so of the

wingforms this type of flap. The top surface of the wing does not move while the

bottom surfacelowers. The split flap is effective in improving the lift, but it creates

a great deal of form drag, asshown in the figure. The split flap was used on the DC-

3. It was also used on WWII-era divebombers. Because it helped increase lift at

low speeds and slowed the airplane during the dive.

Multi slotted flaps are seen on many modern passenger jets, while large

airplanes usesingle-slotted flaps. Until the 1990’s airplane performance was the

key design criterion. Airplanecompanies were proud of sophisticated triple-slotted

flap systems. During the 1990’s a shifttoward reducing cost as a key design

criterion has pushed airplane companies to maximize theperformance of single-

slotted flaps. One technique that is used is to place vortex generators onthe leading

edge of the single slotted flap. When the flap is retracted, the vortex generators on

theflap are hidden in the wing. Thus, the vortex generators do not penalize the

airplane in cruise butare available for takeoff and landing. The next times you fly a

commercial airplane ask for awindow seat behind the wing. During the approach

and landing phase of the flight, watch thewing unfold. It is truly remarkable how

the wing evolves into a high-lift wing from its normalcruise configuration.

FIGURE-6

8.3 SLOTS AND SLATS:

Leading edge devices like flaps, are sometimes used to increase the

camber of thewing and increase the stall angle of attack. But the details are

somewhat different. Other times,The purpose of the leading edge devices is much

like that of the slot in a slotted flap. Thesedevices allow the high-energy air from

below the wing to flow to the upper surface of the wing.This energizes the

boundary layer. Thus, the wing stalls at a higher angle of attack and themaximum

lift is increased. The simplest leading edge device is the fixed slot shown in

Figure.This is a permanent slot near the leading edge of the wing.

FIGURE-7

8.4 DEFLECTED SLIPSTREAM AND JET WASH:

One way to increase lift at slow flight speeds is to divert the propeller’s

slipstream or thejet engine’s exhaust down. To achieve a substantial lift increase

with a slipstream, the plane musthave engines mounted on the wings with large

propellers that generate a slipstream over asubstantial portion of the wing. The

wing must also have a multislotted flap system to deflect theslipstream effectively.

This technique has not found significant commercial applications. Theexhaust of a

turbofan-powered airplane can be diverted down to produce additional lift at

lowspeeds. One way to produce the diversion is to have the flaps extend down into

the exhaust whenfully extended. One problem with this technique is that the flap

extension into the jet exhaustexposed it to very high temperatures, creating a

significant design challenge. Another way todivert the jet exhaust is to mount the

engines on the top of the wing with the engine exhaustcrossing the top of the wing

as in Figure. Flaps behind the engines use the Coanda effect todivert the exhaust

down when extended. This gives a substantial increase in lift for takeoff

andlanding.

Modern high lift systems are often quite complex with many elements and

multi-barlinkages. Here is a double-slotted flap system as used on a DC-8. For

some time Douglas resistedthe temptation to use tracks and resorted to such

elaborate 4-bar linkages. The idea was that thesewould be more reliable. In

practice, it seems both schemes are very reliable. Current practice hasbeen to

simplify the flap system and double (or even single) slotted systems are often

preferred.

Slats operate rather differently from flaps in that they have little effect on the

lift at a given angleof attack. Rather, they extend the range of angles over which

the flow remains attached. This isshown in fig

8.5 LEADING EDGE DEVICES:

Leading edge devices such as nose flaps, Kruger flaps, and slats reduce the

pressurepeak near the nose by changing the nose camber. Slots and slats permit a

new boundary layer tostart on the main wing portion, eliminating the detrimental

effect of the initial adverse gradient.

8.6 WING TIPS:

Wing tips shape has two effects upon subsonic aerodynamic performance.

The tipshape affects the aircraft wetted area, but only to a small extent. A far more

important effect isthe influence the tip shape has upon the lateral spacing of the tip

vortices. This is largelydetermined by the ease with which the higher-pressure air

on the bottom of the wing can ”escape” around the tip to the bottom of the wing. A

smoothly-rounded tip easily permits the air to flow around the tip. A tip with

asharp edge makes it more difficult, thus reducing the induced drag. Most of the

new low-dragwing tips use some form of sharp edge. In fact, even a simple cut-off

tip offers less drag than arounded-off tip, due to the sharp edges where the upper

and lower surfaces end.

The mostly widely used low-drag wing tip is the Hoerner wingtip. This is a

sharpedgedwing tip with a upper surface continuing the upper surface of the wing.

The lower surfaces “undercut” and canted approximately 30 deg to the horizontal.

The lower surface may also be“under cambered”

The “drooped” and “upswept” wing tips are similar to the Hoerner wingtip

except thatthe tip is curved upward or downward to increase the effective span

without increasing the actualspan.

RESULT:

Thus the balancing and maneuvering loads on tail plane, aileron and

rudder was studied.

CHAPTER-9

STRUCTURAL LAYOUT

9.1 SPECIFIC ROLES OF WING (MAIN PLANE) STRUCTURE:

The Specified Structural roles of the wing or main plane are,

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

beam.

Inertial loads from the power plants undercarriage etc to the

main beam.

Aerodynamic loads generated on the aerofoil, control surface

and the flaps to the main beam.

To React against,

Landing loads at the attachment points.

Loads from pylons/stores.

Wing drag and trust loads.

To provide

Fuel tankage space.

Torsional rigidity to satisfy stiffness and aero elastic

requirements.

To fulfill these specific roles of wing layout will conventionally

compromise.

Span wise members (known as spars or booms).

Chord wise members (ribs).

Covering skin.

Stringers.

9.2 BASIC FUNCTIONS OF WING STRUCTURAL MEMBERS:

The structural functions of each of these types of members may

be considered independently as

FIGURE-4

9.2.1 SPARS:

Form the main span wise beam.

Transmit bending and torsional loads.

Provide a closed cell structure.

Provide resistance to torsion, shear and tension loads.

In particular,

Webs-resist shear and torsional loads helps to stabilize the skin.

Flanges-resist the compressive loads caused by wing loading.

9.2.2 SKIN:

To form impermeable aerodynamic surface.

Transmit aerodynamic forces to ribs and stringers.

Resist shear torsion loads (with spar webs).

React axial bending loads (with stringers).

9.2.3 STRINGERS:

Increase skin panel buckling strength by diving in to the

smaller length section.

React axial bending loads.

9.2.4 RIBS:

Maintain the aerodynamic shape.

Act along with the skin to resist the distributed

aerodynamic pressure loads.

Distribute concentrated loads in to the structure and

redistribute stress around any discontinuities.

Increase the column buckling strength by the stringers

through the skin.

Increase the skin panel buckling strength by the stringers

through the skin.

9.3 WING BOX CONFIGURATION:

Several basic configurations are in use nowadays

Mass beam concept.

Box beam (distributed flange) concept built up or integral

construction.

Multi spar.

Single spar D-nose wing layout.

FIGURE-5

9.3.1 MASS BEAM LAYOUT:

In this design, all of the span wise bending loads are reacted

against by substantial booms or flanges reacted against by unsingle spar D-nose

configuration is sometimes used on very lightly loaded structures. The outer skin

only react against the shear loads. They form a closed cell structure between the

spars. These skins need to be stabilized against buckling due to the applied shear

loads. This is done using ribs and a small number of span wise stringers.

9.3.2 BOX BEAM OR DISTRIBUTED FLANGE LAYOUT:

This method is more suitable for aircraft wings with medium to

high load intensities and differs from the mass boom concept in that the upper and

lower skins also contribute to the span wise bending resistance.

Another difference is that the concept incorporates span wise

stringers (usually Z section) to support highly stressed skin panel area. The

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

structural damage tolerance.

Design difficulties include,

Inter reaction between the ribs and stringers. So that the each

rib either has to pass below the stringers or the load path must be broken. Some

examples of common design solutions are shown in figure.

Many joints are present leading to high structural weight,

assembly times, complexity, costs and stress concentration areas.

The concept desired above is commonly known as built up

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

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

is very popular now a day. The concept is simple in the skin stringer panel are

manufactured singly from large billets of metal.

9.4ADVANTAGES OF THE INTEGRAL CONSTRUCTION METHOD

OVER TRADITIONAL BUILT UP METHOD INCLUDE

Simpler construction and assembly.

Reduce overall assembly time/costs.

Improved possibilities to use optimized panel tapering.

Disadvantages include

Reduce damage tolerance, so that planks are used.

Difficult to use on large aircraft panels.

9.5 TYPES OF SPARS:

In the case of two or three spar box beam layout, the front spar

should be located as far forward as possible to maximize the wing box size.

Adequate wing depth for reacting vertical shear loads.

Adequate nose space for LE device deicing equipment

etc…

This generally results in the front spar being located at 12% to

18% of the chord length. For a single spar D-nose layout the spar will usually

located at the maximum thickness position of the airfoil section. For the standard

beam layout the rear spar will be located as for as possible.

This is usually in a location between about 55% and 70% of the

chord length. If any intermediate spars are used they would tend to be spaced

uniformly. Unless there are specific pick up point requirements.

9.6 RIBS:

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

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

around the edges, so that they can be riveted to the skin and the spar webs cut outs

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

lightening holes are usually out in to the rib bodies to reduce the rib weight and

also allow the passage for control runs fuel election act.

9.7 RIB CONSTRUCTION AND CONFIGURATION:

The rib should be ideally spaced to ensure adequate overall

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

influenced by facilitating attachment points for control surfaces, flaps, slats, spoiler

hinges, power plants, stores, undercarriage attachment etc…

Positions of fuel tank ends, requiring closing ribs.

A structural need to avoid local shear or compressive

buckling.

9.8 RIB ALIGNMENT POSSIBILITIES:

There are several different possibilities regarding to

alignment of the ribs on swept wing aircraft.

It is a hybrid design n which one or more inner ribs are

aligned with bthe main axis while the remainder is

aligned perpendicularly to the rear spar.

It is usually the preferred option but presents several

structural problems in the root design.

Gives good torsional stiffened characteristics but results

in heavy ribs and complex connection.

RESULT:

Thus the study of using structural layout have been done successfully.

CHAPTER-10

DESIGN OF SOME OF COMPONENTS IN AIRCRAFT

10.1 TYRE SIZING:

Nearly 90% of the load is carried by the main landing gear.

Only of 10% of aircraft is carried by nose wheel. But it experience

dynamic loads.

Nose wheel size could be 60-100% of size of main wheel.

But in the bicycle and quarter cycle configuration the size is same.

10.2 PERFORMANCE PARAMETER:

Operating a tyre at the lower internal pressure will greatly improve the tyre

life.

Largest tyre cause drag, weights the space occupied etc.

10.3 GEAR RETRACTION GEOMETRY:

In The Wing

Wing-Podded

In The Fuselage

Fuselage Podded

Wing/Fuselage Junction

In The Nacelle

FIGURE-14

10.4 LANDING GEAR WEIGHT:

WTO=8125 kg.

WWm=0.8*WTO/ns

WWn=0.2* WTO

Sub this value in the WWm&WWnFormula,

WWm= 3250 kg.

WWn=1625 kg.

10.5 LANDING GEAR DIMENSIONS:

Diameter = AWWB

Width = AWWB

The constant values for MAIN LANDING GEAR,

A=1.63

B=0.315

The constant values for NOSE LANDING GEAR,

A=0.1043

B=0.480

Sub this values in the Diameter & Width Formula,

The constant values for MAIN LANDING GEAR,

Diameter = 20.81 cm

Width = 5.058 cm

The constant values for NOSE LANDING GEAR,

Diameter = 16.73cm

Width= 3.62cm

10.6 TO LOCATE THE LANDING GEAR:

There are two methods followed to locate the landing gear system,

TIP OVER CRITERIA

Longitudinal Tip Over Criteria

Lateral Tip Over Criteria

GROUND CLEARANCE CRITERIA

Longitudinal Ground Clearance Criteria

Lateral Ground Clearance Criteria

The Landing Gear arrangement is Tricycle Landing Gear.

10.7 TO FIND THE CENTER OF GRAVITY:

FIGURE-15

C.G= 0.45*17

C.G=7.65 m

Fwd C.G=0.2*17

Fwd C.G=3.4 m

Aft C.G=0.45*17

Aft C.G=11.9 m

Height of the aircraft= 3.45 m

10.8 TO FIND THE LOAD ACTING ON THE LANDING GEAR:

lm=3.45 tan15o

lm=0.92m

Pm=WTO*ln/ns(lm+ln)

Pm=WTO*lm/(lm+ln)

WTO=18500 kg

Sub this values in the Pm&PnFormula

Pm=4625 kg

Pn=9250 kg

10.9 STICK DIAGRAM:

FIGURE-16

The Nose landing gear retracted in Front direction.

The Main landing gear retracted in Rear direction.

RESULT

Thus the design of some component in aircrafts have been done successfully

THREE VIEW DIAGRAM