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CONTENTS

1. Abstract

2. List of symbols

3. Introduction

4. Review of Design project-I

5. V-n diagram for the design study

Introduction

a. V-n diagram flight envelope

b. Unintentional maneuvers

c. Gust load

d. Maneuvering loads

e. Conclusion

6. Gust and Maneuverability In Envelopes

a. Maneuver diagram

b. Gust diagram

c. Addition computations

d. Conclusion

7. V-n diagram graph calculation

8. Aircraft Structural Design

a. Introduction

b. Structural concept

c. Design Criteria

d. Wing structure

e. Basic function of wing structural members

f. Wing box configurationg. Rib construction

h. Air loads on wing

i. Conclusion

9. Design of the components of wing and fuselage

a. Structural analysis of fuselage

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b. Stress analysis

c. Fuselage structure

10.Computation of load estimation of wing and fuselage

11.Material selection

a. Composites

b. Metals

c. Mechanical Properties

d. Conclusion

12. Balancing and maneuvering load on tail plane

a .Maneuverability load on primary control surface

b. Trim tab effect

c. Gust load condition

d. Unsymmetrical load

e. Aileron

f. Conclusion

12.Materials

a. Miscellaneous number

b. Structural optimization and design

c. Metallic materials

d. Non metallic materials

e. Conclusion

14 .Three view diagrams

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15. Conclusion

16. Bibliography

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1. ABSTRACTAircraft design project is about the estimation or new ideas and values

implementing into given empirical formulas. The ADP II is continuation of ADP1 which was

done and the weight calculations and efficiency of four seat light aircraft was estimated.

As a continuation of previous one the ADP II is meant for more advanced and

estimation of loads on four seat light aircraft. Advancement of technology is leading to a new era

of aircrafts and this design project is an added value to it.

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LIST OF SYMBOLS

A Area

Compressive area of the fuselage

Tension area of the fuselage

b Span

B.M Bending Moment

C Chord

C.G Centre of Gravity

Coefficient of lift

Maximum lift coefficient

FOS Factor of Safety

H Height of the fuselage

I Moment of Inertia

Reaction of the front spar

Maximum lift

Reaction of the rear spar

Reaction of the tail plane

M Bending Moment

N Reaction factor

n Load factor

Maximum load factor

q Shear flow

R Resultant

S Area of the wing

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T Torque

Gust Velocity

Cruise Velocity

Dive Velocity

Positive Stall Velocity

Negative Stall Velocity

Weight of the aircraft

Weight of the aircraft

Weight of the fuselage

Weight of the wing

Angle of attack

Change in the angle of attack

Change in the lift coefficient

L Change in lift

Twist

Density

Ultimate Stress

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INTRODUCTION

Aircraft Design Project-II is a continuation of Aircraft Design Project-I. In our

Aircraft Design Project-I, we have performed a preliminary and conceptual analysis. We have

carried out a weight estimation, engine selection, weapon loading and aerodynamic parameter

selection and analysis. Apart from these, we have also determined performance parameters such

lift, drag, range, endurance, thrust and power requirements.

The purpose of ADP-II is to enhance the knowledge in continuation of the

design project given in ADPI. Also, 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 V-n diagram, structural analysis of fuselage and wings and the appropriate

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

the landing gear position, retraction and other accompanying systems and mechanisms have also

been done.

Thus, by imposing all the performance parameters in our ADP-I, structural analysis

of our goal is done in this project. Albert Einstein once said Do not worry about your problems

with mathematics; I assure you mine are far greater. He said this to imply on the significance of

mathematics to reduce complicated things into simpler ones. Hence, a lot of attention is given to

calculations in this report.

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General notes on "FAR-25" Airworthiness Requirements

Part A - General

Part B - FLIGHT - General - Performance - Controllability and Manoeuvrability

Stability - Stall - Ground Handling Characteristics -Misc.

Part C - STRUCTURE - GeneralFlight Manoeuvre and Gust Conditions -

Ground Loads - Emergency Landing ConditionsFatigue Evaluation.

Part D - DESIGN & CONSTRUCTION - General - Materials - Production of

Structure - Control Systems - Personnel and Cargo Accommodation-

Environment Control.

Part E - POWER PLANT - General - Induction System - Power Plant Controls andAccessories - Fuel System - Exhaust System - Power Plant Fire Protection. Part

Performance Requirements (in FAR - 25)

(a) STALL - A minimum operational speed must be established to assure

controllability and manoeuvrability in all reasonable operating conditions. This

must provide for a minimum level of safety to account for gust upsets, inadvertent

operations, evasive manoeuvring and production tolerances.

(b) TAKE-OFF - The Take-Off speed is dependent on runway conditions. (Dry,

slippery, etc.) A 35 foot obstacle (height) must be cleared at the end of the runway

for the "ENGINE-OUT" case.

(c) LANDING - Landing distance is the horizontal distance from a point where the

main wheels of the airplane are 50 feet above the runway surface to the point

where the aircraft has come to a complete stop. The landing must be preceded by a

steady approach down to the threshold height (50 feet) at a gradient of descent not

greater than 5.2% (approximately works to 3). Velocity of descent must be

specified. The rate of sink at TOUCH-DOWN should not exceed 3 feet per

second. (For HF-24 aircraft, this is taken as 7 feet per sec). Braking should not be

initiated till all the wheels are firmly on the ground. The landing distance must be

corrected for head-wind (not more than 50%) and tail-wind (not more than 150%).

(d) TRIM - Trim requirements must encompass the prescribed minimum operational

speeds and airplane configuration selected. Undue control forces, overshoots, or

objectionable airplane

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Structure Requirements (in FAR - 25)

(a) LOADS - Loads are critically affected by MANOEUVRES, stability of the

airplane, and also a wide range of structural temperature (particularly in SST).

Loads must be determined under all specified conditions.

(b) FACTORS OF SAFETY - The LIMIT LOAD or the UNFACTORED LOAD is

the load which is the largest load likely to occur in the operational conditions. It is

usually not a static load, but generally represents some DYNAMIC

MANOEUVRING CONDITION. (V-n diagram). The PROOF LOAD is obtained

by multiplying the Limit Load by a number, sometimes as low as 1.0, and

sometimes as high as 1.33. The structure must withstand the proof load without

DETRIMENTAL DISTORTION. By multiplying the limit load by the ultimate

factor of safety, usually from 1.5 to 2.0, the ULTIMATE LOAD is obtained.

Ultimate load is the load which the structure must withstand without collapse.

RESERVE FACTOR = A / B (British practice) - for optimum performance.

Where, A" is the load which the structure is established to be capable ofwithstanding. B" is the load which the structure is required to withstand.

An additional Factor of Safety of 1.25 should be applied in case of thermal problems to obtain

combined load and thermal ultimate strains (in SST). This is in addition to the usual factor of

safety mentioned earlier.

(c) PROOF OF STRUCTURE - TESTING - TESTING must be the primarymethod for proof of structure (Static and fatigue) due to complexities and

uncertainties. The ULTIMATE TEST PROGRAM is a necessity, but the length

of time needed to execute it will be very large. All the airplanes and 50% of the

major components (like Wing, Fuselage, Fin, Stabilizer, etc.) experience major

test failures under flight loads less than the design ultimate.

Static tests to ultimate strength levels for both NORMAL and FAIL-SAFE

conditions must be the primary method of proof of original and residual static

strength. Proof of adequate residual strength for repeated loads should be

demonstrated by tests and analyses, as well as by supporting tests.

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Some Special Aspects of Requirements from Point of View of Structures (for all

aircraft)

(a)FACTORS OF SAFETY - The designer first estimates the APPLIED LOADS, i.e. the actualmaximum loads imposed on the components. Then the sign is done such that, for these loads,

the stresses in any part the structure should not exceed the yield point stresses and it shouldnot have a permanent set.

Then, Design Loads = 1.5 x Applied Loads

(Or Ultimate Loads) (Or Limit Loads)

The Structure should carry these design loads without failure or collapse, although it

may deform considerably.

Note:

(1) For Aluminium alloys, permanent set should not exceed 0.2%

(2) For current high duty aluminium alloys, this ratio can be taken as1.35.

FITTING FACTORS IN DESIGN

For FITTINGS, the Fitting Margin of Safety can be taken as about 15% for

Military Aircraft and about 20% for Commercial Aircraft (for STATIC case).

The above is necessary due to the following unforeseen factors - Nature of

load distribution - Correct fitting and bearing of bolts or rivets on the parts jointed -

Abrupt changes in cross sections of fittings - Stress Concentration Factors around holes

- Defects in manufacture.

Special fitting factors are necessary for repeated loads, vibration and impact

loads.

Important Factors in Structural Design

(i) Ductility of material - RESILIENCE of the component (strain energy density at

FRACTURE).

(ii) Effective Stress Concentration Factors (Fatigue)

(iii) Load paths, Fatigue and Fail-Safe

(iv) Load spectrum and Random loading

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(v) Reliability of stress analysis in Systems approach

(vi) Scatter in local stress values for the given load and Environment.

(vii) Lubrication of contact surfaces, Chemical engineering.

(viii) Grade of Workmanship, Production processes, Human efforts.

(ix) The Factors of safety

Proof Factor Ultimate Factor

General structures ........ 1.125 ------ > 1.5

1 1/3

Special structures ......... 1.5 ------ > 2.0

1 1/3

Extreme cases of structures (E.g. Canopies) VERY LARGE FACTORS

(d) SAFE-LIFE versus FAIL-SAFE DESIGN (Dr. F H Hooke)

Pronouncement of LIFE for a SPECIFIED FAILURE RATE still remains a very

complicated problem.

SAFE-LIFE........ Undiagnosed disease

FAIL-SAFE...... Diagnosable and Remediable disease

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3. REVIEW OF DESIGN PROJECT-I

1. WEIGHT SPECIFICATIONS

Total takeoff gross weight =1069.66 kg

Empty weight = 664.66 kg

Fuel weight = 137.66kg

2. GEOMETRIC SPECIFICATIONS

Wing span = 10.93m

Wing area = 16.2 m2

Aspect ratio = 7.35

Wing loading = 66.82 N/m2

Maximum length = 8.11 m

Maximum height = 2.45m

3. PERFORMANCE SPECIFICATIONS

Range = 1065 km

Maximum cruise speed = 217 kmph

Stalling speed = kmph

Service ceiling = 6200 m

Rate of climb = 300 m/min (At sea level)

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V-N DIAGRAM

FOR STUDY

DESIGN

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V-N DIAGRAM FOR STUDY DESIGN

Introduction

A v-n diagram shows the flight load factors that are used for the structural design as a function of

the airspeed . these represent the maximum expected loads that the aircraft will experience .

these load factors refered to us the limit of the load factor. Load factor standards for aircraft are

covered by the FAR-25 for transport and FAR-23 for normal ,utility ,aerobatic ,and commuter

aircraft. The military specification is covered in MIL-A-8661A .A summary of load limits is

given

AIRCRAFT TYPE LOAD FACTOR

General aviation normal -1.25 TO 3.1General utility -1.8 TO 4.4

General aerobatic -3 TO 6

Home built -2 TO 5

Commercial aircraft -1.5 TO 3.5

Fighter -4.5 TO 7.75

An example of the vn diagram for maximum maneuver loads factor is shown .the curve from

n=0 to the point A represent the maximum normal component load produced by the high angle of

attack flight the equation of the curve is

The maximum value of the point a is determined

by the FAR standard for the particular type of

the aircraft .the limit corresponds to the

horizontal line A-D

n = qCl /w/s

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Point D

occurs at the highest flight velocity which is dive velocity .Recall the v dive=vcruise .the point vs in

the figure cruise velocity .at cruise n=1 which is shown . the in intersection of that line with the

OA curve corresponding to the stall velocity which is minimum speed at the aircraft can

maintain level flight .the line HF represents the largest negative load factor for aparticular type of

the aircraft .These are generaly less than the maximum positive load factor

V-n DIAGRAMFLIGHT ENVELOPE

The V-n diagram is a graphical representation of an aircrafts flight envelope. This plotgives us a clear indication of the structural and aerodynamic limitations of an aircraft. The V-n

diagram is basically a plot of velocity (equivalent air speed) to the load factor.

The greatest air loads on an aircraft usually come from the generation of lift during high-

g maneuvers. Even the fuselage is always structurally sized by the lift of the wing rather than theair pressures produced directly on the fuselage. Aircraft load factor expresses the maneuvering of

an aircraft as a multiple of the standard acceleration due to gravity. At lower speeds, the highest

load factor that an aircraft may experience is limited by the maximum lift available. At higher

speeds, the maximum load factor is limited to some arbitrary value based upon the expected use

of the aircraft.

Using the V-n diagram two important load factor values can be plotted, which are

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1) Limit load factor- Value of load factor corresponding to which there is Permanent structural

deformation

2) Ultimate Load factor Value of load factor corresponding to which there is outrightstructural failure.

UNINTENTIONAL MANEUVERS

The loads experienced when the aircraft encounters a strong gust can exceed the

maneuverings loads in some cases. Civil aircraft flying near thunderstorms or encountering high-

altitude clear air turbulence or while performing low altitude attack runs, dog fighting etc.experience load factors due to gusts. When an aircraft experiences gust, the effect is an increase

or decrease in the angle of attack.

Thus in order to establish the safe flight envelope of our aircraft, we have plotted as per FAR 25

norms,

1) V-n maneuvering diagram

2) V-n gust diagram

GUST LOAD

Gust loads are the unsteady aero dynamics load that are produced by the atmospheric turbulence

. they represent a load factor that are added to the added to the aerodynamic load

The effect of turbulence gust produced for a short time of change in the effective angle of attack

This changes may positive or negative there may produce increase and decrease in the wing lift.

MANEUVERING LOADS

The greatest air loads on an aircraft usually comes from the generation of lift during

high-g maneuvers. Even the fuselage is almost always structurally sized by the lift of the wings

rather than by the pressures produced directly on the fuselage. Aircraft load factor (n) expresses

the maneuvering of an aircraft as a standard acceleration due to gravity

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

the maximum lift available. At higher speeds the maximum load factor is limited to some

arbitrary value based upon the expected use of the aircraft. The maximum lift load factor equals

1.0 at levels flight stall speed. The aircraft can be stalled at a higher speed by trying to exceed the

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available load factor, such as in steep turn. The point labeled high A.O.A (Angle of Attack) is

the slowest speed at which the maximum load can be reached without stalling.

This part of the flight envelope is important because the load on the wing is

approximately perpendicular to the flight direction, not the body-axis vertical direction. At high

angle of attack, the load direction may be actually forward to the aircraft body-axis vertical

direction, causing a forward load component of the wing structure.

The aircraft maximum speed, or dive speed at right of the V-n diagram represents the

maximum dynamic pressure and maximum load factor is clearly important for structural sizing.

At this condition, the aircraft is at fairly low angle of attack because of the high

dynamic pressure, so the load is approximately vertical in the body axis. For a subsonic aircraft,

maximum speed is typically 50% higher than the level

The loads experienced when the aircraft encounters a strong gust can exceed themaneuver loads in some cases. For transport aircraft flying near thunderstorms encountering

high-altitude clear air turbulence, it is not unheard of to experience load factors due to gustsranging from a negative 1.5 to a positive 3.5g or more. When an aircraft experiences a gust, the

effect is an increase (or decrease) in A.O.A

Gust tends to follow a cosine like intensity increase as the aircraft flies through,

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allowing it more time to react to the gust. This reduces the acceleration experienced by the

aircraft by as much as 40%.To account for this effect a statistical gust alleviate on factor(K)

has been devised and applied to measuring gust load

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CONCLUTION:

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Thus the design of v-n diagram was studied, and also study about v-n diagram flight

envelope, gustload, maneuver load etc..

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GUST AND

MANEUVERABILITY

ENVELOPE

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Gust and Manuver ability in envelopes

Maneuver DiagramThis diagram illustrates the variation in load factor with airspeed for maneuvers. At

low speeds the maximum load factor is constrained by aircraft maximum CL. At higher speeds

the maneuver load factor may be restricted as specified by FAR Part 25.

The maximum maneuver load factor is usually +2.5 . If the airplane weighs less than

50,000 lbs., however, the load factor must be given by: n= 2.1 + 24,000 / (W+10,000)

n need not be greater than 3.8. This is the required maneuver load factor at all speeds up to Vc,unless the maximum achievable load factor is limited by stall.

The negative value of n is -1.0 at speeds up to Vc decreasing linearly to 0 at VD

.Maximum elevator deflection at VA and pitch rates from VA to VD must also be considered.

Gust Diagram

Loads associated with vertical gusts must also be evaluated over the range of

speeds.

The FAR's describe the calculation of these loads in some detail. Here is a

summary of the method for constructing the V-n diagram. Because some of the speeds (e.g. VB)

are determined by the gust loads, the process may be iterative. Be careful to consider the

alternative specifications for speeds such as VB.

The gust load may be computed from the expression given in FAR Part 25. This

formula is the result of considering a vertical gust of specified speed and computing the resulting

change in lift. The associated incremental load factor is then multiplied by a load alleviation

factor that accounts primarily for the aircraft dynamics in a gust.

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with: a = (dCL

Ue = equivalent gust velocity (in ft/sec)

Ve = equivalent airspeed (in knots)

Kg = gust alleviation factor

Additional Notes on Computations

1) The lift curve slope may be computed from the DATCOM expression:

- -M2)

and is an empirical correction factor that accounts for section lift curve slopes different from 2.

In practice is approximately 0.97. This expression provides a reasonably good low-speed lift

curve slope even for low aspect ratio wings. The effect is an important one as can be seen from

the data for a DC-9 shown below. The maximum lift curve slope is about 50% greater than its

value at low Mach numbers.

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2) Recall CLmax may vary with Mach number as discussed in the high-lift section.

CONCLUTION:

Thus the study of gust and maneuverability envelope has been studied, by using gust

diagram and maneuver diagram.

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V-N DIAGRAM

GRAPH CALCULATION

V-n DIAGRAM GRAPH CALCULATION

Determination of +1g Stall speed (VS )

VS= {2(GW/s)/ CN max}

GW-flight design gross weight in lbs

S-Wing area in ft2

-Air density in slugs/ft3

CNmax= {(CLmax)2+ (CD at CLmax)

2}

1/2

CNmax = maximum normal force coefficient

CNmax= {(CLmax)2+ (CD at CLmax)

2}

1/2

= {(1.8)2+ (0.6)

2}

1/2

=1.897

Vs = {2(80,000/1347.964)/ (1.225*1.897)}1/2

= 7.146 knots

Determination of Design Cruising speed (VC )

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Vc must be sufficiently greater than VB to provide for inadvertent speed increases likely

to occur as a result of severe atmospheric turbulence

VC >> VB+45 knots

VC>kc (GW/S)1/2

Vc=design cruising speed

Kc=36, for acrobatic aircraft w/s=20 psi varies up to 100 psi

Vc= (36) (80000-/1347.964)1/2

=277.33 knots

Determination of VD:

Vd >1.25 Vc

Vd=design diving speed

=1.25(277.33)

= 346.66 knots

Determination of VA:

VA>VS (lim)

VA=design maneuvering speed

lim =load factor limit

lim pos>(2.1+ (24000/(GW+10000))

VA=7.146 (2.3667)1/2

=11 knots

VA need not exceed VC

Determination of design speed for maximum gust intensity (VB)

VB need not be greater than VC

VB may not be less than the speed determined by the intersection of CN max lines and the

gust line VB

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Determination of negative stall speed VSneg

Vs neg= {2(gw/s)/ ( CN max neg)2}

1/2

Where, CN max neg=1.1 CL max neg

In preliminary design

CL max neg = negative coefficient of lift

CN max neg =1.1 CL max neg

=1.1*1.897 = 2.0867

Vs neg = {2(80000/1347.964)/ (1.225*2.0867)2}

1/2

=4.262 knots

Determination of design limit load factor lim

The positive limit maneuvering load factor is determined from the following:

lim pos>(2.1+ (24000/(GW+10000))

lim pos = (2.1+(24000/(80000+10000))

= 2.3667

Exceptions:

lim pos> 2.5 at all times

lim pos need not be greater tan 3.8 at Wto

The negative design limit load factor is determined from:

lim neg >-1 up to VC

Construction of gust load factor lineslim= (1+kg Ude VCL/498(gw/s))

kg=gust alleviation factor

kg =0.88Mg/ (5.3+Mg)

Mg=2(gw/s)/CgCL

=2(80000/1347.964)/1.225*1.518*1.82)

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=35.072

Kg =0.88*35.072/ (5.3+35.072)

=0.764

limneg= (1+0.7781*1.4*33.28*277.33/(80000/1347.964))

=1.21

Conclusion

Thus calculation of V-n diagram was done and using this the load factor limits wereknown of designed passenger aircraft. Also the velocity limits was known for design purpose.

AIRGRAFT

STRUCTURAL DESIGN

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Aircraft Structural Design

Introduction

Although the major focus of structural design in the early development of aircraft was on

strength, now structural designers also deal with fail-safety, fatigue, corrosion, maintenance and

inspectability, and producability.

Structural Concepts

Modern aircraft structures are designed using a semi-monocoque concept- a basic load-

carrying shell reinforced by frames and longerons in the bodies, and a skin-stringer construction

supported by spars and ribs in the surfaces.

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Proper stress levels, a very complex problem in highly redundant structures, are

calculated using versatile computer matrix methods to solve for detailed internal loads. Modern

finite element models of aircraft components include tens-of-thousands of degrees-of-freedom

and are used to determine the required skin thicknesses to avoid excessive stress levels,

deflections, strains, or buckling. The goals of detailed design are to reduce or eliminate stress

concentrations, residual stresses, fretting corrosion, hidden undetectable cracks, or single failure

causing component failure. Open sections, such as Z or J sections, are used to permit inspection

of stringers and avoid moisture accumulation.

Fail-safe design is achieved through material selection, proper stress levels, and

multiple load path structural arrangements which maintain high strength in the presence of a

crack or damage. Examples of the latter are:

a)Use of tear-stoppers

b)Spanwise wing and stabilizer skin splices

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Analyses introduce cyclic loads from ground-air-ground cycle and from power

spectral density descriptions of continuous turbulence. Component fatigue test results are fed

into the program and the cumulative fatigue damage is calculated. Stress levels are adjusted to

achieve required structural fatigue design life.

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Design Life Criteria -- Philosophy

Fatigue failure life of a structural member is usually defined as the time to initiate a crack

which would tend to reduce the ultimate strength of the member.

Fatigue design life implies the average life to be expected under average aircraft utilization

and loads environment. To this design life, application of a fatigue life scatter factor accounts for

the typical variations from the average utilization, loading environments, and basic fatigue

strength allowables. This leads to a safe-life period during which the probability of a structural

crack occurring is very low. With fail-safe, inspectable design, the actual structural life is much

greater.

The overall fatigue life of the aircraft is the time at which the repair of the structure is no

longer economically feasible.

Scatter factors of 2 to 4 have been used to account for statistical variation in component

fatigue tests and unknowns in loads. Load unknowns involve both methods of calculation and

type of service actually experienced.

Primary structure for present transport aircraft is designed, based on average expected

operational conditions and average fatigue test results, for 120,000 hrs. For the best current

methods of design, a scatter factor of 2 is typically used, so that the expected crack-free

structural life is 60,000 hrs, and the probability of attaining a crack-free structural life of 60,000

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hrs is 94 percent as shown in the following figure and table.

s.f. = N / NpProbability of

Survival (%)

Np (Flight Hours)

(N = 120,000 hrs)

Np (Years)

(3,000 flight hrs / year)

2.0 94.0 60,000 20

2.5 97.5 48,000 16

3.0 98.8 40,000 13.3

3.5 99.3 34,300 11.4

4.0 99.54 30,000 10.0

With fail-safe design concepts, the usable structural life would be much greater, but in

practice, each manufacturer has different goals regarding aircraft structural life.

FUNCTION OF THE STRUCTURE:

The primary functions of an aircrafts structure can be basically broken down into thefollowing:

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To transmit and resist applied loads.

To provide and maintain aerodynamic shape.

To protect its crew, passenger, payload, systems, etc.

For the vast majority of aircraft, this leads to use of a semi-monocoque design (i.e. a thin,stressed outer shell with additional stiffening members) for the wing, fuselage & empennage.

These notes will discuss the structural layout possibilities for each of these main areas, i.e. wing,

fuselage & empennage.

WING STRUCTURE:

The specific structural roles of the wing are:

To transmit:

Wing lift to the root via the main spanwise beam.

Inertial loads from the powerplants, undercarriage, etc. to the main beam.

Aerodynamic load generated on the aerofoil, control surfaces & flaps tothe main beam.

To react against:

Landing loads at attachment points.

Loads from pylons/stores.

Wing drag and thrust loads.

To provide:

Fuel tank space.

Torsional rigidity to satisfy stiffness and aeroelastic requirements.

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

Span wise members (known as spars or booms).

Chord wise members (ribs).

A covering skin.

Stringers.

BASIC FUNCTIONS OF WING STRUCTURAL MEMBERS

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

independently as:

Spars:

Form the main span wise beam Transmit bending and torsional loads

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

In particular:

Websresist shear and torsional loads and help to stabilize the skin. Flanges - resist the compressive loads caused by wing bending.

Skin:

To form impermeable aerodynamics surface Transmit aerodynamic forces to ribs & stringers

Resist shear torsion loads (with spar webs).

React axial bending loads (with stringers).

Stringers:

Increase skin panel buckling strength by dividing into smaller length sections.

React axial bending loads

Ribs:

Maintain the aerodynamic shape

Act along with the skin to resist the distributed aerodynamic pressure loads

Distribute concentrated loads into the structure & redistribute stress around any

discontinuities

Increase the column buckling strength of the stringers through end restraint

Increase the skin panel buckling strength.

Wing Box Configurations

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

Mass boom concept

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

construction

Multi-Spar

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Single spar D-nose wing layout

Mass Boom Layout

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

booms or flanges. A two-boom configuration is usually adopted but a single spar D-nose

configuration is sometimes used on very lightly loaded structures. The outer skins only reactagainst the shear loads. They form a closed-cell structure between the spars. These skins need to

be stabilized against buckling due to the applied shear loads; this is done using ribs and a small

number of span wise stiffeners.

Box Beam or Distributed Flange Layout:

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 zsection) to support the highly stressed skin panel area. The resultant use of a large number ofend-load carrying members improves the overall structural damage tolerance.

Design Difficulties Include:

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

stringers or the load path must be broken. Some examples of common design solutions

are shown in figure

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

costs & stress concentration areas.

The concept described above is commonly known as built-up construction method. An

alternative is to use a so-called integral construction method. This was initially developed for

metal wings, to overcome the inherent drawbacks of separately assembled skin-stringer built-up

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

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

construction method over the traditional built-up method include:

Simpler construction & assembly

Reduced sealing/jointing problems

Reduced overall assembly time/costs

Improved possibility to use optimized panel tapering

Disadvantages include:

Reduced damage tolerance so that planks are used

Difficult to use on large aircraft panels.

Types of spars:

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In the case of a two or three spar box beam layout, the front spar should be located as far

forward as possible to maximize the wing box size, though this is subject to there being:

Adequate wing depth for reacting vertical shear loads.

Adequate nose space for LE devices, de-icing equipment, etc.

This generally results in the front spar being located at 12 to 18% of the chord length. For a

single spar D-nose layout, the spar will usually be located at the maximum thickness position of

the aerofoil section. For the standard box beam layout, the rear spar will be located as far as aft

as possible, once again to maximize the wing box size but positioning will be limited by various

space requirements for flaps control surfaces spoilers etc

This usually results in a location somewhere between about 55 and 70% of the chord length.

If any intermediate spars are used they would tend to be spaced uniformly unless there are

specific pick-up point requirements

Ribs:

For a typical two spar layout, the ribs are usually formed in three parts from sheet metal by

the use of presses and dies. Flanges are incorporated around the edges so that they can be riveted

to the skin and the spar webs Cut-outs are necessary around the edges to allow for the stringers to

pass through Lightening holes are usually cut into the rib bodies to reduce the rib weight and also

allow for passage of control runs fuel electrics etc.

Rib construction and configuration:

The ribs should be ideally spaced to ensure adequate overall buckling support to spar

flanges .In reality however their positioning is also influenced by

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

plants, stores, undercarriage attachments etc

Positions of fuel tank ends, requiring closing ribs

A structural need to avoid local shear or compression buckling.

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Rib Alignment Possibilities:

There are several different possibilities regarding the alignment of the ribs on swept-

wing aircraft

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

axis while the remainder is aligned perpendicularly to the rear spar

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

root region

(c) Gives good torsional stiffness characteristics but results in heavy ribs and

complex connections

AIR LOADS ON WING

With the V-n diagram complete, the actual loads and load distribution on the wing can be

determined. Before the actual structural members can be sized and analyze, the loads they willsustain must be determined. Aircraft loads estimation, a separate discipline of aerospace

engineering, combines aerodynamics, structures and weights.

Initially we have to calculate the lift produced by the wings. Once the lift on the wings is

known, the span wise and chord wise load distributions can be determined.

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According to classical wing theory, the span wise lift or load distribution is proportional

to the circulation at each station. A vortex lifting line calculation will yield the span wise liftdistribution. For an elliptical plan form wing, the lift and load distributions is of elliptical shape.

For a non elliptical wing, a good semi-empirical method for span wise load estimation is

known as Schrenks Approximation. This method assumes that the load distribution on anuntwisted wing has a shape that is the average of the actual plan form shape and an elliptical

shape of same and area. The total area under the lift load curve must sum to the required total

lift.

Schrenks approximation for various Wing Planforms

CONCLUTION:

Thus the structural design studytheory approach has been studied,and also study

about wing structure , rib construction,wing box configuration.

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DESIGN OF THE COMPONENTS OF WING AND FUSELAGE

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DESIGN OF THE COMPONENTS OF WING AND FUSELAGE

STRUCTURAL ANALYSIS OF FUSELAGE

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

analysing 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.

Fig. 13 Stringer Position on Fuselage

The stringer used is of Z type. The following are its dimensions

Cross sectional area of each stringer is 200 mm^2

5m

1.5m

1.386m

1.06m

0.574m

0.589m

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Fig. 14 Cross section of Z-section

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 = Area of Boom 1

tD = Thickness of skin panel

b = Circumferential distance between 2 stringers

B1 = 200+ (5X (589/6)) [2+ (1386/1500)] + (5X (589/6)) [2+ (1386/1500)]

4mm

40mm

100mm

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= 3070.3933 mm2

Similarly for boom 2 we get

B2 = 200+ (5X (589/6)) [2+ (1060/1386)] + + (5X (589/6)) [2+ (1300/1386)]

= 3070.3933 mm2

Thus solving we B1:B16 = 3070.3933 mm2. But B5 = B13 = 0

We know that Ixx = Ay2

Ixx = [(2 X 15002) + (4 X 1386

2) + (4 X 1060

2) + (4 X 574

2)] X 3070.3933

= 5.5255 X 1010

mm4

We know that the maximum bending moment is B.M = 77650 N-m

Hence the bending moment acting on the fuselage M = B.M X n X FOS

= 77650 X 3 X 1.5

= 349425 N-m

The value of stress acting is given by the expression:

= 0.07589 * y

The following is the shear flow diagram over a fuselage.

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Fig. 15 Shear Flow diagram of Fuselage

FUSELAGE STRUCTURE:

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

the payload, crew, equipment, systems and (possibly) the powerplant. Furthermore, it must react

against the in-flight manoeuvre, pressurisation and gust loads; also the landing gear and possibilyany powerplant loads. Finally, it must be able to transmit control and trimming loads from the

stability and control surfaces throughout the rest of the structure.

Fuselage Layout Concepts

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

mass boom and longeron layout

semi-monocoque layout

Mass Boom & Longeron layout

This is fundamentally very similar to the mass-boom wing-box concept discussed in

previous section. It is used when the overall structural loading is relatively low or when there are

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

(longeron), reacting against any direct stresses caused by applied vertical and lateral bending

loads. Frames or solid section bulkheads are used at positions where there are distinct direction

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changes and possibly elsewhere along the lengths of the longeron members. The outer shell helps

to support the longerons against the applied compression loads and also helps to support the

longerons against the applied compression loads and also helps in the shear carrying. Floors are

needed where there are substantial cut-outs and the skin is stabilized against buckling by the use

of frames and bulkheads.

Semi Monocoque Layout

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

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

skin thickness determined by pressurization, shear loading & fatigue considerations.

Longitudinal stringers provide skin stabilisation and also contribute to the

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

often used around edges of cut-outs. Less integral machining is possible than on an equivalent

wing structure.

Frames are used to stabilize the resultant skin-stringer elements and also to

transmit shear loads in the structure. They may also help to react against any pressurization loads

present. They are usually manufactured as pressings with reinforced edges. Their spacing (pitch)

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is usually determined by damage tolerance considerations,I.e.crack-stoppingrequirements.

CONCLUTION:

Thus the design of components of wing and fuselage has been studied,by using stress

and bending moment diagram.

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COMPUTATION OF LOAD ESTIMATION ON WING AND FUSELAGE

LOAD ESTIMATION ON WINGS AND FUSELAGE

Wing loading or otherwise W/S plays a very important role,which determines the velocity

at which the aircraft can fly at different conditions

Wing loading at take off

Take off parameter: TOP=(W/S)*(1/CLmax)*(W/T)to*(1/)

For different values of (W/S)to we get

W/S(Kg/m2) TOP(Kg/m2)

0 0

125 610.125

130 634.5

135 658.935

140 683.34

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Take Off Distance

Sto=20.9*(TOP)+87*((TOP)*(T/W))

For Different Values Of TOP we get Sto values

W/S(Kg/m2) STO(m)

0 0

610.125 4740.67

634.5 4930.30

658.93 5119.93

683.34 5309.5518

TOP values at different T/W AND Clmax

TOP=(W/S)*(1/Clmax)*(W/T)to*(1/ )

Clmax TOP(m)at T/W=.2 TOP(m)at T/W=.24 TOP(m)at T/W=.321.2 7632.523 7670.6 7770.6

1.8 6866.32 6842.85 6262.89

2 6143.75 6177.8 6238.42

WE plot a graph for various value of Clmax Vs Sto at different T/W values.

Sto=20.9*(TOP)+87*(TOP)*(T/W)

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WING LOADING ON LANDING

LANDING PARAMETER

LP=(W/S)*(1/Clmax)

For different values of Clmax

Clmax LP1(kg/m2)

1.2 1320.9

1.8 1188.81

2.0 1080.73

For calculating landing system, the formula used is

Landing distance=SL=118(LP)+400ft

Clmax SL1(m)

1.2 2087.55

1.8 1890.98

2.0 1730.15

CONCLUSION:

Thus the load estimation of wing and fuselage was done.

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MATERIAL SELECTION

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MATERIAL SELECTION

The materials on our aircraft are a careful blend of strength and minimum cost. With a

256002 lb aircraft capable of manoeuvres above 2 gs, the structural requirement dictate a

material with a very high strength to weight ratio such as composites. The general considerations

guiding material selection on our aircraft are listed in order of importance are:

1) Strength to weight ratio

2) Cost

3) Manufacturing

4) Availability

Two types of materials are used, metal alloys and composites. The metals will be used

in low load areas where large amounts of material are required such as frames, bulkheads, ribs

and the fuselage skin. This is due to the higher availability of metals compared to composites, as

well as easing manufacturing costs. Composites are used in the high load areas in the wings,

control surfaces, and longerons.

COMPOSITES

Aluminium 7075 was also considered for low stress areas. High stress areas require

materials such as fibre based composites with either a thermoplastic or thermo set matrix. The

wing of the aircraft needs very high strength components, so varieties of composites are

considered. The materials considered for the spar are listed below. The skin along with panels is

made out of Hexcel composite materials. The HexPly 8551-7 Epoxy Matrix is used as the

matrix material in the structural components. The material properties can be seen below. The

wing spars use HexTow IM9 Carbon Fibres as the fibre as it has a higher specific strength. The

matrix material for the HexTow IM9 Carbon Fibres parts is HexPly 8551-7 Epoxy Matrix, a low

temperature matrix material that has a mould temperature of 93C. The spars also contain

honeycomb core to prevent buckling of the load carrying fibre by increasing the stiffness and

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increasing the strength. The core increases the strength up to 7 times the strength of the panel and

up to 17 times the stiffness, if the core is 4 times the plates thickness. The honeycomb used is

aluminium based material because it has the highest strength to weight ratio. The honeycomb

core used is HexWeb Cr III Micro-Cell and the Adhesive used is Redux 328H.

Other areas that use composites are the longerons and control surfaces. The forward

fuselage longerons are made of Spectra fibres and the rear longerons are made of Hexcel. The

Hexcel is required for the rear longerons as they undergo the heat produced by the engine. The

control surfaces are made of a Hexcel based skin, and have a honeycomb core. To assemble the

composite sections, thermoplastic flanges will be attached to the bulkheads. The thermoplastic

flange and adjoining piece can then be heated locally to bond the flange and adjoining piece

together.

Matrix Properties:

Fibre Properties:

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Honeycomb Core Properties:

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Adhesive Properties:

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METALS

Metals such as aluminium are a staple in aircraft structures. Aluminium is a material

with strength to weight ratio better than steel and is much cheaper than composites. Stainless

steel is a very high strength material and has very good high temperature properties, so it will be

used around the engine to shield the composite longerons from the extreme engine heat.

Aluminium lithium is lighter than typical 7075 aluminium and will therefore be used in the high

load bulkheads where the weight savings over aluminium 7075 would be the greatest.

Aluminium lithium is also used for the skin of the fuselage and tails for the same reason.

Aluminium 7075 is acceptable for the frames.

Mechanical properties

The mechanical properties of 7075 depend greatly on the temper of the material.

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7075-0

Un-heat-treated 7075 (7075-0 temper) has maximum tensile strength no more than 40,000psi (276 MPa), and maximum yield strength no more than 21,000 psi (145 MPa). The material

has elongation (stretch before ultimate failure) of 9-10%.

7075-T6

T6 temper 7075 has an ultimate tensile strength of 74,000 - 78,000 psi (510 - 538 MPa)

and yield strength of at least 63,000 - 69,000 psi (434-476 MPa). It has elongation of 5-8%.

7075-T651

T651 temper 7075 has an ultimate tensile strength of at least 67,000 - 78,000 psi (462 -538 MPa) and yield strength of 54,000 - 67,000 psi (372-462 MPa). It has elongation of 3-9%.

The 51 suffix has no bearing on the mechanical properties but denotes that the mate rial isstress relieved by control stretching.

Table: 14 Aluminium alloy comparisons

DESIGN OF LANDING GEAR

We have designed the landing gear characteristics by following a step by step method.

1) Landing gear SystemWe have chosen a Retractable system landing gear which will be retracted in to the

fuselage after the take off.

2) Landing Gear Configuration

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The landing gear configuration we have adapted is the Conventional type or Tri-cycle

type with a nose wheel in front. From an ease of ground manoeuvring viewpoint as well as

ground looping the nose wheel configuration is preferred.

3) Preliminary landing gear strut dispositionThere are two geometric criteria which are required to be considered on deciding

the disposition of landing gear struts are:

A)Tip-over criteria

B) Ground clearance criteria

A) Tip-over Criteria :a) Longitudinal Tip-over Criterion :For tricycle gears the main landing gear must be behind the aft CG location. The 15

deg angle as shown in the Fig. represents the usual relation between main gear and the aft CG.

Fig. 16 Longitudinal tip over criterion

b) Lateral Tip-over Criterion :The lateral tip-over is dictated by the angle in the Fig.

15 degrees

C.G.

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Fig. 17 Lateral Tip-over Criterion

a) Longitudinal Ground Clearance Criterion :

Fig. 18 Longitudinal Ground Clearance Criterion

Lateral Ground Clearance Criterion :

>15 degrees

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Fig.19 Lateral Ground Clearance Criterion

Number of Wheels :

Nose landing gear - 1

Main landing gear - 2

Maximum Static load per Strut :

Conclusion

Before constructing an aircraft there are many things that should be taken

account. One of the important thing is the material selection that we have seen above .

This field basically involves about structures and most the engineers today prefe

the composite materials

>5 degree

Pn

C.G.

Pm

WTO

In Im

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BALANCING AND MANEUVERING ON TAIL PLAIN

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BALANCING AND MANUVERING LOAD ON TAIL PLANE

MANEUVERABILITY LOADS ON PRIMARY CONTROL SURFACES

The control surfaces must be designed for the limit loads resulting from the flightconditions and the ground gust conditions, considering the requirements for --

(a) Loads parallel to hinge line,

(b) Pilot effort effects,

(c) Trim tab effects,

(d) Unsymmetrical loads, and

(e) Auxiliary aerodynamic surfaces

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PILOT EFFORT EFFECTS

(a) General. The maximum and minimum pilot forces, specified in paragraph (c) ofthis section, are assumed to act at the appropriate control grips or pads (in a manner

simulating flight conditions) and to be reacted at the attachment of the control system

to the control surface horn.

(b)Pilot effort effects. In the control surface flight loading condition, the air loads on

movable surfaces and the corresponding deflections need not exceed those that wouldresult in flight from the application of any pilot force within the ranges specified in

paragraph (c) of this section. Two-thirds of the maximum values specified for the

aileron and elevator may be used if control surface hinge moments are based onreliable data. In applying this criterion, the effects of servo mechanisms, tabs, and

automatic pilot systems, must be considered.

(c)Limit pilot forces and torques. The limit pilot forces and torques are as follows:

---------------------------------------------------------------------Maximum forces or Minimum forces or

Control torques torques

---------------------------------------------------------------------Aileron:Stick. 100 lbs. 40 lbs.

Wheel \1\ 80 D in.-lbs \2\.. 40 D in.-lbs.

Elevator:Stick .. 250 lbs. 100 lbs.

Wheel (symmetrical) 300 lbs. 100 lbs.

Wheel (unsymmetrical) \3\... ............ 100 lbs.Rudder. 300 lbs 130 lbs.---------------------------------------------------------------------

\1\ The critical parts of the aileron control system must be designed

for a single tangential force with a limit value equal to 1.25 timesthe couple force determined from these criteria.

\2\ D=wheel diameter (inches).

\3\ The unsymmetrical forces must be applied at one of the normalhandgrip points on the periphery of the control wheel.

TRIM TAB EFFECTS

The effects of trim tabs on the control surface design conditions must be

accounted for only where the surface loads are limited by maximum pilot effort. Inthese cases, the tabs are considered to be deflected in the direction that would assist

the pilot, and the deflections are --

(a) For elevator trim tabs, those required to trim the airplane at any point within the

positive portion of the pertinent flight envelope, except as limited by the stops; and

(b) For aileron and rudder trim tabs, those required to trim the airplane in the critical

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unsymmetrical power and loading conditions, with appropriate allowance for rigging

tolerances.

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Control surfaces and supporting hinge brackets must be designed for inertia loads

acting parallel to the hinge line.

In the absence of more rational data, the inertia loads may be assumed to be equal

to KW, where --

(1) K=24 for vertical surfaces;

(2) K=12 for horizontal surfaces; and

(3) W=weight of the movable surfaces.

The effects of trim tabs on the control surface design conditions must be

accounted for only where the surface loads are limited by maximum pilot effort. In these

cases, the tabs are considered to be deflected in the direction that would assist the pilot,and the deflections are --

(a) For elevator trim tabs, those required to trim the airplane at any point within the

positive portion of the pertinent flight envelope in 25.333(b), except as limited by the

stops; and

(b) For aileron and rudder trim tabs, those required to trim the airplane in the critical

unsymmetrical power and loading conditions, with appropriate allowance for rigging

tolerances

GUST LOAD CONDITION

(a) The control system must be designed as follows for control surface loads due to ground

gusts and taxiing downwind:

(1) The control system between the stops nearest the surfaces and the cockpit controls must bedesigned for loads corresponding to the limit hinge moments H of paragraph (a)(2) of this

section. These loads need not exceed --

(i) The loads corresponding to the maximum pilot loads for each pilot alone; or

(ii) 0.75 times these maximum loads for each pilot when the pilot forces are applied in the samedirection.

(2) The control system stops nearest the surfaces, the control system locks, and the parts of thesystems (if any) between these stops and locks and the control surface horns, must be designed

http://www.flightsimaviation.com/data/FARS/part_25-333.htmlhttp://www.flightsimaviation.com/data/FARS/part_25-333.htmlhttp://www.flightsimaviation.com/data/FARS/part_25-333.htmlhttp://www.flightsimaviation.com/data/FARS/part_25-333.html

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for limit hinge moments H, in foot pounds, obtained from the formula, H=.0034KV cS, where

--

V=65 (wind speed in knots)

K=limit hinge moment factor for ground gusts derived in paragraph (b) of this section.

c=mean chord of the control surface aft of the hinge line (ft);

S=area of the control surface aft of the hinge line (sq ft);

(b) The limit hinge moment factor K for ground gusts must be derived as follows:

---------------------------------------------------------

Surface K Position of controls

---------------------------------------------------------(a) Aileron. 0.75 Control column locked

Or lashed in mid-position.

(b) ......do...... \1\ 1 Ailerons at full throw.0.50

(c) Elevator...\1\ 1 Elevator full down.

0.75(d) ......do... \1\ 1 Elevator full up.

0.75

(e) Rudder...... 0.75 Rudder in neutral.

(f) ......do..... 0.75 Rudder at full throw.

\1\ A positive value of K indicates a moment tending to depress thesurface, while a negative value of K indicates a moment tending to

raise the surface.

UNSYMMETRICAL LOADS

(a) In designing the airplane for lateral gust, yaw maneuver and roll maneuver conditions,account must be taken of unsymmetrical loads on the empennage arising from effects such as

slipstream and aerodynamic interference with the wing, vertical fin and other aerodynamic

surfaces.

(b) The horizontal tail must be assumed to be subjected to unsymmetrical loading conditionsdetermined as follows:

(1) 100 percent of the maximum loading from the symmetrical maneuver conditions and the

vertical gust conditions acting separately on the surface on one side of the plane of symmetry;

and

(2) 80 percent of these loadings acting on the other side.

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(c) For empennage arrangements where the horizontal tail surfaces have dihedral angles greater

than plus or minus 10 degrees, or are supported by the vertical tail surfaces, the surfaces andthe supporting structure must be designed for gust velocities specified in) acting in any

orientation at right angles to the flight path.

(d) Unsymmetrical loading on the empennage arising from buffet conditions of) must be takeninto account.

AILERONThe ailerons must be designed for the loads to which they are subjected

(a) In the neutral position during symmetrical flight conditions; and(b) By the following deflections (expect as limited by pilot effort), during

unsymmetrical flight conditions:

(i)Sudden maximum displacement of the aileron control at VA. Suitable allowance may

be made control system deflections.(ii)Sufficient deflection at VC, where VC is more than VA, to produce a rate of roll not less

than obtained in paragraph (a)(2)(i) of this section.

(iii)Sufficient deflection at VD to produce a rate of roll not less than one-third of that

obtained in paragraph (a)(2)(i) of this section.

Conclusion

Every control surfaces are important and unavoidable part of an aircraft. All the

moments such as rolling, yawing and pitching occurs only due to this control surfaces.

By knowing about this it easy to understand the working and constructing features ofcontrol surfaces.

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MATERIALS

MATERIALS

An aircraft must be constructed of materials that are both light and strong. Early

aircraft were made of wood. Lightweight metal alloys with a strength greater than wood were

developed and used on later aircraft. Materials currently used in aircraft construction are

classified as either metallic materials or nonmetallic materials.

Choice of materials emphasizes not only strength/weight ratio but also:

Fracture toughness

Crack propagation rate

Notch sensitivity

Stress corrosion resistance

Exfoliation corrosion resistance

Acoustic fatigue testing is important in affected portions of structure.

Doublers are used to reduce stress concentrations around splices, cut-outs, doors, windows,

access panels, etc., and to serve as tear-stoppers at frames and longerons.

Generally DC-10 uses 2024-T3 aluminum for tension structure such as lower wing skins,

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pressure critical fuselage skins and minimum gage applications. This material has excellent

fatigue strength, fracture toughness and notch sensitivity. 7075-T6 aluminum has the highest

strength with acceptable toughness. It is used for strength critical structures such as fuselage

floor beams, stabilizers and spar caps in control surfaces. It is also used for upper wing skins.

For those parts in which residual stresses could possibly be present, 7075-T73 material is

used. 7075-T73 material has superior stress corrosion resistance and exfoliation corrosion

resistance, and good fracture toughness. Typical applications are fittings that can have

detrimental preloads induced during assembly or that are subjected to sustained operational

loads. Thick-section forgings are 7075-T73, due to the possible residual stresses induced during

heat treatment. The integral ends of 7075-T6 stringers and spar caps are overaged to T73 locally.

This unique use of the T73 temper virtually eliminates possibility of stress corrosion cracking in

critical joint areas.

Miscellaneous Numbers

Although the yield stress of 7075 or 2024 Aluminum is higher, a typical value for design

stress at limit load is 54,000 psi. The density of aluminum is .101 lb / in3

Minimum usable material thickness is about 0.06 inches for high speed transport wings.

This is set by lightning strike requirements. (Minimum skin gauge on other portions of the

aircraft, such as the fuselage, is about 0.05 inches to permit countersinking for flush rivets.

On the Cessna Citation, a small high speed airplane, 0.04 inches is the minimum gauge on

the inner portion of the wing, but 0.05 inches is preferred. Ribs may be as thin as 0.025 inches.Spar webs are about 0.06 inches at the tip.

For low speed aircraft where flush rivets are not a requirement and loads are low, minimum

skin gauge is as low as 0.016 inches where little handling is likely, such as on outer wings and

tail cones. Around fuel tanks (inboard wings) 0.03 inches is minimum. On light aircraft, the spar

or spars carry almost all of the bending and shear loads. Wing skins are generally stiffened. Skins

contribute to compression load only near the spars (which serve as stiffeners in a limited area).

Lower skins do contribute to tension capability but the main function of the skin in these cases is

to carry torsion loads and define the section shape.

In transport wings, skin thicknesses usually are large enough, when designed for bending,

to handle torsion loads.

Fuel density is 6.7 lb/gallon.

Structural Optimization and Design

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Structures are often analyzed using complex finite element analysis methods. These

tools have evolved over the past decades to be the basis of most structural design tasks. A

candidate structure is analyzed subject to the predicted loads and the finite element program

predicts deflections, stresses, strains, and even buckling of the many elements.

The designed can then resize components to reduce weight or prevent failure. In

recent years, structural optimization has been combined with finite element analysis to

determine component gauges that may minimize weight subject to a number of constraints.

Such tools are becoming very useful and there are many examples of substantial weight

reduction using these methods. Surprisingly, however, it appears that modern methods do not

do a better job of predicting failure of the resulting designs, as shown by the figure below,

constructed from recent Air Force data.

METALLIC MATERIALS

The most common metals used in aircraft construction are aluminum, magnesium,

titanium, steel, and their alloys.

Alloys

An alloy is composed of two or more metals. The metal present in the alloy in the

largest amount is called the base metal. All other metals added to the base metal are called

alloying elements. Adding the alloying elements may result in a change in the properties of the

base metal. For example, pure aluminum is relatively soft and weak. However, adding small

amounts or copper, manganese, and magnesium will increase aluminum's strength many times.

Heat treatment can increase or decrease an alloy's strength and hardness. Alloys are important to

the aircraft industry. They provide materials with properties that pure metals do not possess.

ALUMINIUM

Aluminium is the most widely used material in aircraft structures. Modern commercial

transports such as the Boeing 747 use aluminium for about 80% of the structure. Aluminium is

readily formed and machined, has reasonable cost, is corrosion resistant, and has an excellent

strength-to-weight ratio. In its pure aluminium are used, the most common being aluminium

2024, an alloy consisting of 93.5% aluminium, 4.4% copper, 1.5% manganese, and 0.6%

magnesium. This alloy is also called duralumin.

Density/Specific Gravity (g.cm-3 at 20C) 2.70

Melting Point (C) 660

Specific heat at 100 C, cal.g-1K-1 (Jkg-1K-1) 0.2241 (938)

Latent heat of fusion, cal.g-1 (kJ.kg-1) 94.7 (397.0)

Electrical conductivity at 20C (% of international annealed copper standard) 64.94

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Thermal conductivity (cal.sec-1cm-1K-1) 0.5

Thermal emissivity at 100F (%) 3.0

Reflectivity for light, tungsten filament (%) 90.0

STEEL:

For a typical commercial transport, steel makes up to 17% of the structure. It is used

in those areas requiring very high strength, such as wing attachment fittings, L/G, engine fittings

and flap tracks. Steel is an alloy of iron and carbon; typical steel alloys have about 1% carbon.

Stainless steel is an alloy of steel & chromium that has good corrosion-resistant properties.

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TITANIUM

Titanium has a better strength-to-weight ratio than aluminium and retains its strength

at higher temperatures. However, it is hard to form and machine and is expensive, costing about5 to 10 times more than aluminium. But some supersonic aircraft have to use titanium because of

the high skin temperatures due to aerodynamic heating. The SR-71 aircraft is such a case. This

airplane cruises at mach 3 and above; hence, it was the 1st

airplane to make extensive use of

titanium. Today, titanium is still a major consideration in the decision about the design mach

number of a 2nd

generation supersonic transport.

HIGH TEMPERATURE NICKEL ALLOYS

We note that hypersonic airplanes required advanced, high- temperature

materials to withstand the high rates of aerodynamic heating at hypersonic speeds. Some nickel-based alloys are capable of withstanding the temperatures associated with moderate hypersonic

speeds. The X-15, is such a case. This aircraft was designed to fly as fast as mach 7; hence, its

structure made extensive use of inconel, a nickel based alloy.

AlloyYoung's

modulus [GPa]

Yield strength

[Mpa]

Ultimate

strength [Mpa]

Ultimate

strain [%]

Ti pure - Grade 1

Ti pure - Grade 2

Ti pure - Grade 3

Ti pure - Grade 4

102.7

102.7103.4

104.1

170

275380

485

240

345450

550

24

2018

15

Ti-6Al-4V (Annealed) 110 - 114 825-869 895-930 6-10

Ti-6Al-7Nb

Ti-5Al-2.5Fe

Ti-5Al-1.5BTi-15Zr-4Nb-4Ta-0.2Pd

(Annealed)

Ti-15Zr-4Nb-4Ta-0.2Pd

(Aged)

114

112110

9994

880-950

895820-930

693806

900-1050

1020925-1080

715919

8-15

1515-17

2818

Ti-13Nb-13Zr (Aged)

Ti-12Mo-6Zr-2Fe

(Annealed)

Ti-15Mo (Annealed)

Ti-15Mo-5Zr-3Al

(Solubilized)

Ti-15Mo-5Zr-3Al (Aged)Ti-15Mo-2.8Nb-0.2Si

(Annealed)

Ti-35.3Nb-5.1Ta-7.1Zr

Ti-29Nb-13Ta-4.6Zr

(Aged)

79-84

74-85

7880

808355

80

836-908

1000-1060

544838

1000-1060945-987

547

864

973-1037

1060-1100

874852

1060-1100979-999

597

911

10-16

18-22

2125

18-2216-18

19

13.2

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NONMETALLIC MATERIALS

In addition to metals, various types of plastic materials are found in aircraft

construction. Some of these plastics include transparent plastic, reinforced plastic, composite,

and carbon-fiber materials.

Transparent Plastic

Transparent plastic is used in canopies,windshields, and other transparent

enclosures. You need to handle transparent plastic surfaces carefully because they are relatively

soft and scratch easily. At

approximately 225F, transparent plastic becomes soft and pliable.

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Reinforced Plastic

Reinforced plastic is used in the construction of radomes, wingtips, stabilizer

tips, antenna covers, and flight controls. Reinforced plastic has a high strength-to-weight ratio

and is resistant to mildew and rot. Because it is easy to fabricate, it is equally suitable for other

parts of the aircraft. Reinforced plastic is a sandwich-type material. It is made up of two outer

facings and a center layer. The facings are made up of several layers of glass cloth, bonded

together with a liquid resin. The core material (center layer) consists of a honeycomb structure

made of glass cloth. Reinforced plastic Is fabricated into a variety of cell sizes.

Composite and Carbon Fiber Materials

High-performance aircraft require an extra high strength-to-weight ratio material.

Fabrication of composite materials satisfies this special requirement. Composite materials are

constructed by using several layers of bonding materials (graphite epoxy or boron epoxy). These

materials are mechanically fastened to conventional substructures. Another type of composite

construction consists of thin graphite epoxy skins bonded to an aluminum honeycomb core.

Carbon fiber is extremely strong, thin fiber made by heating synthetic fibers, such as rayon, until

charred, and then layering in cross sections.

Conclusion

By knowing about the materials used in aircraft we understood that how much

constructing materials affect the aerodyanamic loading of an aircraft.

Cost effective material as well as efficient materials are needed for the aircraft and still

the research is going on it.The most advanced one is the Composite Materials that is used in

aircrafts.

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THREE VIEW

DIAGRAMS

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THREE VIEW DIAGRAMS

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CONCLUSION

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CONCLUSION

Thus the designing of long range passenger was done.Design

calculations,Aerodynamic loading and Structural stability of our designed aircraft were done

efficiently. Its also made us to understand the errors and whole designing criteria of an aircraft.

An aircraft is faced by many loads and especially during maneuverability.So such

loads were considered and calculations on loading were done.Eventhough that were complicated

one it was very successful and give knowledge about the loading as well as stability

criterians.We wish that it will be a beginning of new era in aeronautical field.

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BIOBLIOGRAPHY

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BIBLIOGRAPHY

WEBSITES

www.google.com

www.flyzon.com

www.airliners.net

www.wikipedia.org

REFERENCES

Airplane Design And Performance

Introduction to Flight by John D. Anderson, 2nd

edition.

Aircraft Performance and Design by John D. Anderson, 2nd

edition.

Theory of wing sections by Ira.H.Abbott and Albert E. Von Doenhoff, Dover edition.

http://www.google.com/http://www.google.com/http://www.flyzon.com/http://www.flyzon.com/http://www.airliners.net/http://www.airliners.net/http://www.wikipedia.org/http://www.wikipedia.org/http://www.wikipedia.org/http://www.airliners.net/http://www.flyzon.com/http://www.google.com/