evaluation of class 1 fighter

7/25/2019 Evaluation of Class 1 Fighter

1/121

EVALUATION OF CLASS 1 FIGHTER

A PROJECT REPORT

Submitted by

DHANASEKARAN.K

ELAKKIYA.G

in partial fulllment for the the award of the degree

of

BACHELOR OF ENGINEERING

in

AERONAUTICAL ENGINEERING

PARISUTHAM INSITUTE OF TECHNOLOGY AND SCIENCE

ANNA UNIVERSITY:: CHENNAI 600 025

AIRCRAFT DESIGNPROJECT-2


2/121

NOVEMBER 2013

ANNA UNIVERSITY : CHENNAI 600 025

BONAFIDE CERTIFICATE

Certied that this project reportEVALUATION

CLASS 1 FIGHTER

is the bonade work ofELAKKIYA.G! who carried out

the project work under my supervision.

SIGNATURE SIGNATURE

DR.S.BHARATHIRAJA MR.RAJIV

HEAD OF THE DEPARTMENT SUPERVISOR

DEPARTMENT OF AERONAUTICAL LECTURER

ENGINEERING DEPARMENT OF AERONAUTICAL



3/121

PARISUTHAM INSTIUTE OF ENGINEERING

TECHNOLOGY AND SCIENCE PARSUTHAM INSTITUTE OF

TECHNOLOGY AND SCIENCE

ACKNO"LEDGEMENT

We are very much grateful to our beloved chairman , Mr. S.P.Antonisamy, for

Given us an opportunity to study in this wonderful institute and for the various excellent

facilities to learn, develop and excel ourselves in various fields.

We wish to record our deep sense of gratitude to our Dean-Academics

Mrs. J. Nirmalafor her untiring work and dedication towards the growth of the students

community in all aspects.

We feel very grateful to express my thanks t sincere to our ead of the Department,

Dr. S. Bharathiraja, for his continuous encouragement, valuable guidance and support which

always motivated us to be in a groove of the learning process.

!ur heartfelt no bounds to express our sincere thanks to our pro"ect guide Mr. R. Rajiv. Who

was always been with us during the entire course of our pro"ect work. !ur special mention aboutAIRCRAFT DESIGNPROJECT-2


4/121

his involvement in successful completion of our pro"ect.

#ast but not least$ we bow our heads to honor our beloved parents who are our first teachers in

this world and our supreme guide for all our activities and a source of finishing this pro"ect

successfully.

INDE#

Serial.No. Topic Page.no

% &n D'AG(A) *!( + D'G+/D0

1 G/+ AD )A/&(A2'#'+0 !

4 5('+'5A# #!AD'G 3(*!()A5 AD

*'A# &n G(A3 5A#5/#A+'!

6 +(/5+/(A# D'G 7+!(0 A33(!A5

8 #!AD +')A+'!) !* W'G

9 #!AD +')A+'!) !* */#AG

: 2A#A5'G AD )A/&('G #!AD! +A'# 3#A



5/121

; D+A'#D +(/5+/(A# #A0!/+

< D'G !* !) 5!)3!+

%= 3(3A(A+'! !* A D+A'#D D'G(3!(+ W'+ 5AD D(AW'G

ABBREVAT!N

A.(. - Aspect (atio

b - Wing pan

5swell - 5hord of the Airfoil

5root - 5hord at (oot

5tip - 5hord at +ip

5

- )ean Aerodynamic 5hord

5D - Drag 5o-efficient

5D,= - >ero #ift Drag 5o-efficient

5" - pecific fuel consumption



6/121

5# - #ift 5o-efficient

D - Drag

- ndurance

e - !swald efficiency

g - Acceleration due to gravity

G - *actor due to ground effect

?A, ?+ - ymbols

h - eight from ground

h!2 - !bstacle height

k% - 3roportionality constant

kuc - *actor depends on flap deflection

@A , @+ - ymbols

# - #ift

loiterD

#

- #ift-to-drag ratio at loiter

cruise

D#

- #ift-to-drag ratio at cruise

) - )ach number of aircraft



7/121

mff - )ission segment fuel fraction

- +ime between initiation of rotation and actual

( - (ange

(e - (eynolds umber

(5 - (ate of climb

- Wing Area

a - Approach distance

ab - Distance reBuire to clear an obstacle after becoming

airborne

f - *lare distance

g - Ground (oll

ref. - (eference surface area

wet - Wetted surface area

+ - +hrust

3 - 3ower

3cruise - +hrust at cruise

3take-off - +hrust at take-off



8/121

loiterW

P

- +hrust-to-weight ratio at loiter

cruiseW

P

- +hrust-to-weight ratio at cruise

takeoffW

P

- +hrust-to-weight ratio at take-off

&cruise - &elocity at cruise

&stall - &elocity at stall

! - #ift off peed

&+D - +ouch down speed

Wcrew - 5rew weight

Wempty - mpty weight of aircraft

Wfuel - Weight of fuel

Wpayload - 3ayload of aircraft

W= - !verall weight of aircraft

-

W

- Wing loading

- Density of air



9/121

- Dynamic viscosity

r - 5o-efficient of rolling friction

- +apered ratio

!2 - Angle between flight path and take-off

- +urning angle

- Gliding angle

(5 - (ate of climb



10/121

AIM OF THE PROJECT

AIM OF THE PROJECT



11/121

+here 1 classes of fighter aircraft. +hey are class-% and class-1 fighter aircrafts. +he class one

fighter is officially an air superiority fighter. )ost of it can function either as multi role fighters and

ground attack. Air superiority fighter mainly does the function to gain air space control over the enemy

territory so that the bombers can bomb their targets safely, and give support for the ground units. +hey

literally make the enemy air space home ground for the invaders aircrafts. 5lass1 fighters mainly

concentrate on electronic warfare and ground attack along with surveillance.

+oday, complex sets of reBuirements and ob"ectives include specification and research

studies are set for the pro"ectC

Airplane performance, afety, (eliability )aintainability, ubsystems properties 3erformance and!thers.



12/121

ABSTRACT

ABSTRACT



13/121

+he purpose of this pro"ect is to evaluate the class% fighter -

of specifications such as,

)aximum range C 1=== @ilometres

ndurance C %= ours

3ay load C 9=== @ilogram

5ruise altitude C %6%:8 )eters

)aximum )ach number C 4

umber of crew C 1=

And make sure that all factors are off the reBuired factors for the safe flight, under all conditions and

allowable datas are off of the standard of the perfect flying and give the maximum performance.



14/121

INTRODUCTION



15/121

NTR!D"#T!N

Airplane design is both an art and a science.

When we look at aircraft, it is easy to observe that they have a number of commonfeaturesC wings, a tail with vertical and horiEontal wing sections, engines to propel them through the air,and a fuselage to carry passengers or cargo. 'f, however, you take a more critical look beyond the grossfeatures, we also can see subtle, and sometimes not so subtle, differences. What are the reasons forthese differencesF What was on the mindH of the designers that caused them to configure the aircraftin this way for the perfect performance according to the standards.

+he design process is indeed an intellectual activity, but a rather special one that is

tempered by good intuition developed via experience, by attention paid to successful airplane designsthat have been used in the past, and by generally proprietaryH design procedures and databaseshandbooks, etc.,H that are a part of every airplane manufacturer.

De$ining a ne% &esign

+he design of an aircraft draws on a number of basic areas of aerospace engineering. Asshown in the illustration, these include aerodynamics, propulsion, light-weight structures andcontrol.

ach of these areas involves parameters that govern the siEe, shape, weight and

performance of an aircraft. Although we generally try to seek optimum in all these aspects, with anaircraft, this is practically impossible to achieve. +he reason is that in many cases, optimiEing onecharacteristic degrades another basis on their performance.

'n most cases, the design ob"ectives are not as focused. )ore often, the nature of an aircraftdesign is compromise. +hat is, the goal is to balance the different aspects of the total performance whiletrying to optimiEe a few or oneH based on well-defined mission reBuirements.

+here are many performance aspects that can be specified by the mission reBuirements.+hese includeC

+he aircraft purpose or mission profile$ +he typesH and amount of payload$ +he cruise and maximum speeds$ +he normal cruise altitude$ +he range or radius with normal payload$



16/121

+he endurance$ +he take-off distance at the maximum weight$ +he landing distance with 8= percent of the maximum fuel weight$ +he purchase cost$ And other reBuirements considered important$

And there are many performance measuring factors which determines the mission reBuirements.+hese includeC

& 7 n Diagram Gust (oll stimation )anoeuvering load estimation #oad stimation of Wings #oad stimation of *uselage tructural Design Analysis hear and 2ending moment analysis for Wings hear and 2ending moment analysis for *uselage

+he starting point of any new aircraft is to clearly identify its purpose. Withthis, it is often possible to place a design into a general category. uch categories includecombat aircraft, passenger or cargo transports, and general aviation aircraft. +hese may also befurther refined into subcategories based on particular design ob"ectives such as range short orlongH, take-off or landing distances, maximum speed, etc. +he process of categoriEing is usefulin identifying any existing aircraft that might be used in making comparisons to a proposeddesign.

+hus, by analysis the design aspects of the proposed deign factors the aircraft.

DES'N PR!JE#T(

'nternal discussions Discussion with prospective customers Discussion with certification Authorities Deciding upon a 2(!AD !/+#' to start the A5+/A# D'G, which

will consist of 5onstruction of )ock-up tructural and *unctional testing *inal estimation and 3erformance calculation



17/121

3(!3!D D'G *A5+!(

5#)axI 4

#anding distance, #andI %==m

Approach velocity, &aI 6=.;1 ms

tall velocity,&sI 4%.6= ms

WH#andI %;%%.:=

#andI %:.


18/121

Wtake-off I 81414.8; kg

wetI 4


19/121

*or horiEontal tail,

Aspect ratio, A( I 4

+aper ratio, +( I =.4

*or vertical tail,

Aspect ratio, A( I %.8

+aper ratio, +( I =.8

EN'NE SE)E#T!N

As our aircraft is the fighter aircraft flying at a speed of )ach 4, we selected the turbo fan

engines

+wo engines are located in the wing

+he dry thrust is assumed to be of :9.6 @ %:,%;8 lbfH

+he thrust with after burner is assumed to be of %=


20/121

AR#RA*T DES'N !BJE#TVES AND #!NSTRANTS

ss+e Military

Dominant design criteria )ission accomplishment and

survivability

3erformance AdeBuate range and response

!verall mission accomplishment

Airfield environment

hort-to-moderate runways

All types of runway surfaces

!ften partan A+5, etc.

#imited space available

ystem complexity and mechanical

design

#ow maintenance- availability

issue

Acceptable system cost

(eliability and survivability

Damage toleranceGovernment regulations and community

acceptance

)ilitary standards

#ow noise desirable

--Good neighbor in peace

--Dectability in war



21/121

V$% DIAGRAM FOR THE

DESIGN STUDY



22/121

V,n DA'RAM *!R T-E DES'N ST"D

V,n DA'RAM

*A#T!R !* SA*ET / *)'-T ENVE)!PE

+he control of weight in aircraft design is of extreme importance. 'ncreases in weight reBuire

stronger structures to support them, which in turn lead to further increases in weight and so on.

xcesses of structural weight mean lesser amounts of payload, thereby affecting the economic viability

of the aircraft. +he aircraft designer is therefore constantly seeking to pare is aircrafts weight to the

minimum compatible with safety. owever, to ensure general minimum standards of strength and

safety, airworthiness regulations Av.p.


23/121

apply the positive and negative limit loads, corresponding to n%and n4, without stalling the aircraft so

that A5 and * represent &cthe cut-off lines 5D% and D1 relive the design cases to be covered since it

is not expected that the limit loads will be applied at maximum speed. &alues of n %,n1 and n4 are

specified by the airworthiness authorities for particular aircraft$ typical load factors laid down in 25A(are shown.

A particular flight envelope is applicable to one altitude only since 5#max is generally reduced

with an increase of altitude, and the speed of sound decreases with altitude thereby reducing the critical

)ach number and hence the design diving speed &D. *light envelopes are therefore drawn for a range

of altitudes from sea level to the operational ceiling of the aircraft.



24/121

MANE"VER Diagram+his diagram illustrates the variation in load factor with airspeed for maneuvers. At low speeds themaximum load factor is constrained by aircraft maximum 5#. At higher speeds the maneuver loadfactor may be restricted as specified by *A( 3art 18.



25/121

+he load factor, n, is defined as the ratio of the lift to weight, nI#W .'n level flight, the lift producedby the wings eBuals the weight, so that nI%. owever, during )anoeuvres such as climb to altitude,acceleration to high speed, or sustained or instantaneous turns associated with combat, significantlylarger load factors can occur. ince these set the limit on the internal structure, it is important that the

maximum load factor be determined.

n 0 2234214s5546 2#D74855

Where again, B is the dynamic pressure,

5D= is the base drag coefficient for the wing.

And k I1

Ae with e 9=.;

NSTANTANE!"S T"RN RATE

With 'nstantaneous +urn (ate, the load factor was given in below eBuation. +his is reproduced in n 062:2 instV4g5;

ere the turn rate inst, is the 'nstantaneous +urn (ate, which has units of radians per second.

S"STANED T"RN RATE

(ecall that in a sustained turn, the speed and altitude are maintained so that the thrust eBuals the drag,and the load factor is constant.

An expression for the maximum sustained load factor as a function of the wing loading, which isneeded to achieve a specified sustained turn rate, was given in eBuation this reproduced in

n 0 :3?Ae4214S5:2T415ma@/ 3#D7214S5>>=4;

'n terms of the maximum sustained turn rate, the load factor is

n 0 6 2:2 s+stV4g5;

Where sustis the maximum sustained turn rate with units of radians per second.

#)MB

'n the analysis of wing loading effect on climb, it was assumed that nI%. owever, an expression canbe derived, which relates the climb rate to the load factor

2y definition,



26/121

n 0 )41 0 2#)3S5 4 1

+he climb gradient is given as

' 0 sin 0 2T,D541ubstituting for DW in the above eBuation and solving for n, we obtain

D41 0 n2#D74#5)< 2#)4n?Ae5

n 0 :22T415,'5>:22T415,'5;,2 #D74?Ae5>7.C4:; #D74#)>

With the condition that

T41 '


27/121

We can again illustrate this by using the conditions of the conceptual 2?. With 5 #5ruise I =.1,under a dive condition, the load factor would be n I %.81.

+he maximum maneuver load factor is usually L1.8 . 'f the airplane weighs less than 8=,=== lbs.,

however, the load factor must be given byC nI 1.% L 16,=== WL%=,===Hn need not be greater than 4.;. +his is the reBuired maneuver load factor at all speeds up to &c, unless

the maximum achievable load factor is limited by stall.

+he negative value of n is -%.= at speeds up to &cdecreasing linearly to = at &D.

)aximum elevator deflection at &A and pitch rates from &Ato &Dmust also be considered.

V 0 2;nma@4FG#) ma@5 H 214S5

Where,

nma@01

2 FGVG; H :#) Ma@4214S5>

Where,

MJI 1.1:;8N%=-%

5# )axI 4

W I 4=== m1

'f &J I %%== ms

nmax I O N 1.1:;8 N %=-% N %%==H1N 4H N1

3000

nmax I %4:.;6

'f &J I %=8= ms



28/121

nmax I O N 1.1:;8 N %=-% N %=8=H1N 4H N1

3000

nmax I %18.9=

'f &J I %=== ms

nmax I O N 1.1:;8 N %=-% N %===H1N 4H N1

3000

nmax I %%4.


29/121

nmax I O N 1.1:;8 N %=-% N ;8=H1N 4H N1

3000

nmax I ;1.4%

'f &J I ;== ms

nmax I O N 1.1:;8 N %=-% N ;==H1N 4H N1

3000

nmax I :1.


30/121

nmax I O N 1.1:;8 N %=-% N 9==H1N 4H N1

3000

nmax I 6%.=%

'f &J I 88= ms

nmax I O N 1.1:;8 N %=-% N 88=H1N 4H N1

3000

nmax I 46.69

'f &J I 8== ms

nmax I O N 1.1:;8 N %=-% N 8==H1N 4H N1

3000

nmax I 1;.6;

'f &J I 68= ms

nmax I O N 1.1:;8 N %=-% N 68=H1N 4H N1

3000

nmax I 14.=9

'f &J I 6== ms

nmax I O N 1.1:;8 N %=-% N 6==H1N 4H N1

3000

nmax I %;.11



31/121

nlimit

5nmax I %.%5# max

&s I P1N%N81414.8;H

P1.1:;8N%=-%N


32/121

&s LveI P1N 1.88=N=81414.8;

P %.118N


33/121

&D supersonicI Q%L1=RH N cruising speedS

&D supersonic I Q%L=.1HN;;8.1=S

&D supersonicI %=.91.16 ms

Altitude I %6=== m

Altitude I 68


34/121

*or &c 1==== ft to 8==== ftH

&De I 99.9:- =.===;44NhH

&De I 99.9:-=.===;44 N 68


35/121

*or guest velocity, &b

)aximum value, &bI%=.46 ms

)inimum value, &bI8.4= ms



36/121

*or &b I %=.46 ms

nlimit 0 = :2IgH VDe Vc)54 2KH 2%4s55>

&I%=.46 ms

&De I;6.9: L =.===


37/121

=.


38/121

+ is the engine thrust, assumed here to act parallel to the direction of flight in order to simplifycalculation.

AIRCRAFT LOADS IN LEVEL FLIGHT

+he loads are in static eBuilibrium since the aircraft is in a steady state level flight condition. +hus forvertical eBuilibrium

) < P / 1 0 7

*or horiEontal eBuilibrium

T / D 0 7

And taking moments about the aircrafts centre of gravity in the plane of symmetry

) a / D / T c , Mo/ Pl 0 7

*or a given aircraft weight, speed and altitude, the above eBuations may be solved for theunknown lift, drag and tail loads. owever, other parameters in these eBuations, such as )= dependupon the wing incidence V which in turn is a function of the reBuired wing lift so that, in practice, amethod successive approximation is found to be the most convenient means of solution.

As a first approximation we assume that tail load 3 is small compared with the wing lift # sothat, from the above eBuation # 9 W. *rom aerodynamic theory with the usual notation

) 01(V2 S C L)

2



39/121

ence

1(V2 S C L)2

9 1

+he above eBuation gives the approximate lift coefficient 5# and thus from 5#7 V curvesestablished by wind tunnel testsH the wing incidence V. +he drag load D follows knowing & and VH andhence we obtain the reBuired engine thrust + from above eBuation also )o, a, b, c l may be calculatedagain since & and V are knownH and the eBuation can be solved for 3.As a second approximation thisvalue of 3 is substituted in above eBuation to obtain a more accurate value for # and the procedure isrepeated. /sually three approximations are sufficient to produce reasonably accurate values.

'n most cases 3, D and + are small compared with the lift and aircraft weight. +herefore, fromabove eBuation # 9 W and substitution in the above eBuation gives, neglecting D and +

3 X1 22a4l5,Mo4l55

We see from above return eBuation that if a is large then 3 will most likely be positive. 'n otherwords the tail load acts upwards when the centre of gravity of the aircraft is far aft. When a is small ornegative, that is, a forward centre of gravity, then 3 will probably be negative and act downwards.

Lc P+ll o+t manoe+vres(

'n a rapid pull-out from a dive a downward load is applied to the tail plane, causing the aircraftto pitch nose upwards. +he downward load is achieved by a backward movement of the control

column, thereby applying negative incidence to the elevators, or horiEontal tail if the latter is all-moving. 'f the manoeuvre is carried out rapidly the forward speed of the aircraft remains practicallyconstant so that increases in lift and drag result from the increase in wing incidence only. ince the liftis now greater than that reBuired to balance the aircraft weight, the aircraft experiences an upwardacceleration normal to its flight path. +his normal acceleration combined with the aircrafts speed in thedive results in the curved flight path shown in above figure. As the drag load builds up with an increaseof incidence the forward speed of the aircraft falls since the thrust is assumed to remain constant duringthe manoeuvre. *or steady level flight n I %, giving %g flight, although in fact the acceleration is Eero.What is implied in this method of description is that the inertia force on the aircraft carrying out an ngmanoeuvre is nW. We may therefore replace the dynamic conditions of the accelerated motion by an

eBuivalent set of static conditions in which the applied loads are in eBuilibrium with the inertia forces.+hus, in above figure, n is the manoeuvre load factor while f is a similar factor giving the horiEontalinertia force. ote that the actual normal acceleration in this particular case is n-%H g.



40/121

*or vertical eBuilibrium of the aircraft, we have, referring to figure where the aircraft is shownat the lowest point of the pull-out

) < P < T sin / n1 0 7

*or horiEontal eBuilibrium

T cos < $1 / D 0 7

And for pitching moment eBuilibrium about the aircrafts centre of gravity

) a / D / T c , Mo/ Pl 0 7

+he above eBuation contains no terms representing the effect of pitching acceleration of theaircraft$ this is assumed to be negligible at this stage. +he engine thrust + is no longer directly related tothe drag D as the latter changes during the manoeuvre. Generally the thrust is regarded as remaining

constant and eBual to the appropriate to conditions before the manoeurve began.

L& Stea&y p+ll,o+t(

#et us suppose that the aircraft has "ust began its pull-out from a dive so that it is describing acurved flight path but is not yet at its lowest point. +he load acting on the aircraft at this stage of themanoeuvre are shown in above figure. Where ( is the radius of curvature of the flight path. 'n this casethe lift vector must eBuilibrate the normal to the flight pathH component of the aircraft weight andprovide the force producing the centripetal acceleration &1( of the aircraft towards the centre ofcurvature of the flight path. +hus

) 0 21V;4gR5 < 1 cosY

!r, since # I nW

n 0 2V;4gR5 < cosY

At the lowest point of the pull-out, Y I =, and

n 0 2V;4gR5 < =



41/121

AIRCRAFT LOADS AND ACCELERATION DURING A STEADY PULL-OUT

We see from either from the above eBuations the smaller the radius of the flight path, that is themore severe the pull-out to over stress the aircraft by sub"ecting it to the loads which lie outside theflight envelope and which may even exceed the proof or ultimate loads. 'n practice, the control surfacemovement may be limited by stops incorporated in the control circuit. +hese stops usually operate onlyabove a certain speed giving the aircraft adeBuate manoeuvrability at lower speeds. *or hydraulicallyoperated controls Zartificial feel is built into the system whereby the stick force increases progressivelyas the speed increases$ a necessary precaution in this type of system since the pilot is merely openingand closing valves in the control circuit and therefore receives no direct physical indication of control

surface forces.

Alternatively, at low speeds, a severe pull-out or pull-up may stall the aircraft. Again safetyprecautions are usually incorporated in the form of stall warning devices since, for modern high speedaircraft, a stall can be disastrous, particularly at low altitude.

Le #orrectly anIe& t+rn(

'n this manoeuvre the aircraft flies in a horiEontal turn with no side slip at constant speed. 'f the

radius of the turn is ( and the angle of bank [, then the forces acting on the aircraft are those shown inthe figure. +he horiEontal component of the lift vector in this case provides the force necessary toproduce the centripetal acceleration of the aircraft towards the center of the turn. +hus

)sin 0 1V ;4gR



42/121

And for vertical eBuilibrium

)cos 0 1

!r

) 0 1sec

*rom the above eBuation we see that the load factor n in the turn is given by

n 0 sec

Also, dividing the eBuations

tan 0 V ;4gR

xamination of the above eBuation reveals that the tighter the turn the greater the angle of bankreBuired to maintain horiEontal flight. *urthermore, we see from n I sec H Buation that an increasein bank angle results in an increased load factor. Aerodynamic theory shows that for a limiting value ofn the minimum time taken to turn through a given angle at a given value of engine thrust occurs whenthe lift coefficient 5#is a maximum$ that is, when the aircraft on the point of stalling.

'"ST )!AD ESTMAT!N

;a '+st loa&s(Gust loads are unsteady aerodynamic loads that are produced by atmospheric turbulence. +hey

represent a load factor that is added to the aerodynamic loads, which were presented in the previoussections. +he effect of a turbulent gust is to produce short-time change in the effective angle of attack.



43/121

+his change can be either positive or negative, thereby producing an increase or decrease in the winglift and a change in the load factor, nI T#W.

2elow figure shows a model for the effect of a gust on an aircraft in level flight. Aircraft has a

forward velocity, &. +he turbulence gust produces small velocity components, v and u. At that instant,the velocity component in the aircraft flight direction is & L v. 'n level flight, the mean velocitycomponent normal to the flight direction is / I =. +herefore, the total normal velocity is u.

MODEL FOR GUST LOAD EFFECT ON A AIRCRAFT IN LEVEL FLIGHT

'n most cases, u and v are much less than the flight speed, &.+herefore, & L v X &. 2ased on thisassumption, the effective angle of attack is

O 0 tan,=2+4V5

2ecause u is small compared to &

O 9tan,=2+4V5

+he incremental lift produced by the small change in the angle of attack is

O) 01( V2 S C L a )

2

ubstituting the above eBuation gives

O) 01(V2 S C L a u)

2



44/121

+he incremental load factor is then

On 0 S4;1 2 uV C La

+he peak load factor is then the sum of the mean load factor at cruise nI%H and the fluctuationload factor, namely,

npeaI0 n < On

+he gusts that result from at atmospheric turbulence occur in a fairly large band of freBuencies.+herefore their effect on an aircraft depends on factors that affect its freBuency response. 'n particular,the freBuency response is governed by an eBuivalent mass ratio,U, defined as

Q 0 2;14S5 42 uV C La 5

Where c is the mean chord of the mail wing and g is the gravitational constant. ote that U isdimensionless, so that in 2ritish units, gc I 41.1f 7 lbm lbfH 7 s1is reBuired in the numerator.

+he mass ratio, U is a parameter in a response coefficient, @, which is defined differently forsubsonic and supersonic aircraft, namely,

8 0 7.KK Q 4 2C.LL < Q5 2S+sonic5

8 0 Q=.7L4 2.C < Q=.7L5 2S+personic5

STATISTICAL GUST VELOCITY VALUES



45/121

VARIATION OF GUST VELOCITY, , WITH ALTITUDE FOR DIFFERENT FLIGHT CONDITIONS



46/121

+he normal component of the gust velocity, u, is the product of the statistical average of valuestaken from flight data, \, and the response coefficient, or

+ 0 8 \

+he above table gives values of \. +he variation with altitude is presented in the figure.

5onsidering eBuation, we observe that turbulent gusts have a greater effect on aircraft with alower wing loading. +herefore, a higher wing loading is better to produce a ]smoother^ flight, as wellas in lowering the incremental structural loads.

; V,n Diagram '+st Envelope(

+he effect of the additive gust loads can be seen in the &-n diagram shown in figure. +his isshown in blue to contrast it with the load factor envelope for manoeuvres alone.

V,n DA'RAM S-!1N' T-E ENVE)!PE !* )!AD *A#T!RS N#)"DN' '"STS !NA MAN!E"VERN' AR#RA*T

3oint ]2^ corresponds to the maximum lift at the highest angle of attack plus the load factor for a gustwith \ I 99 fs.

3oint ]5^ refers to the load factor at the design cruise velocity, &cplus that for a gust with \ I 8= fs.3oint ]D^ corresponds to the load factor at the dive velocity, &Dplus that for a gust with \ I 18 fs.

3oints ]^, ]*^ and ]G^ correspond to the additional of loads from negative gusts at the velocitiescorresponding to dive, &D$ cruise, &5$ and maximum lift, &V ,respectively.



47/121

3lots like figure which superpose the manoeuvre loads with the gust loads, are important for determinethe conditions that produce the highest load factors. +he largest values are the ones used in thestructural design.

;c Design )oa& *actor(+he ]limit load factor^ denoted in the above figures is the highest of all the manoeuvring load

factors plus the incremental load due to turbulent gusts.

nlimit 0 nma@< On

'n order to provide a margin of safety to the structural design, the limit load factor is multipliedby a ]safety factor^, *. +he standard safety factor used in the aircraft industry is %.8. +his value wasoriginally defined in %


48/121

'"ST AND

MANE"VERAB)T

ENVE)!PES



49/121



50/121

'"ST AND MANE"VERAB)T ENVE)!PES

'"ST DA'RAM

'+stmay beC

Gust, short blast of wind

#oads associated with vertical gusts must also be evaluated over the range of speeds.

+he *A(_s describe the calculation of these loads in some detail. ere is a summary of the method for

constructing the &-n diagram. 2ecause some of the speeds e.g. &2H are determined by the gust loads,

the process may be iterative. 2e careful to consider the alternative specifications for speeds such as &2.

+he gust load may be computed from the expression given in *A( 3art 18. +his formula is the result of

considering a vertical gust of specified speed and computing the resulting change in lift. +he associated

incremental load factor is then multiplied by a load alleviation factor that accounts primarily for the

aircraft dynamics in a gust.

withC a I d5#daH

/eI eBuivalent gust velocity in ftsecH

&eI eBuivalent airspeed in knotsH

@gI gust alleviation factor

ote that c is the mean geometricchord here.

http://en.wikipedia.org/wiki/Wind#Measurementhttp://en.wikipedia.org/wiki/Wind#Measurement


51/121

+he *AA specifies the magnitude of the gusts to be used as a function of altitude and speedC

Gust velocities at 1=,=== ft and belowC

99 ftsec at &2

8= ftsec at &5

18 ftsec at &D.

Gust velocities at 8=,=== ft and aboveC

4; ftsec at &2

18 ftsec at &5

%1.8 ftsec at &D.

+hese velocities are specified as eBuivalent airspeeds and are linearly interpolated between 1==== and

8==== ft.

o, to construct the &-n diagram at a particular aircraft weight and altitude, we start with the maximum

achievable load factor curve from the maneuver diagram. We then vary the airspeed and compute the

gust load factor associated with the &2gust intensity. +he intersection of these two lines defines the

velocity &2. Well, almost. As noted in the section on design airspeeds, if the product of the %-g stall

speed, &s%and the sBuare root of the gust load factor at &5ngH is less than &2as computed above, we

can set &2I &s%sBrtngH and use the maximum achievable load at this lower airspeed.

ext we compute the gust load factor at &5and &Dfrom the *AA formula, using the appropriate gust

velocities. A straight line is then drawn from the &2point to the points at &5and &D.

# N 5osY I W

Where,

At ground, # I O M/15#



52/121

# I O N%.118N%=44.91N4


53/121

( I %:8.:: m

+urn rate, WI g N Pn1max-% &D

5onsider, nmax I

SCHEMATIC REPRESENTATION OF TWO SPANWISE LIFT DISTRIBUTIONS FOR AN

ELLIPTIC AND TRAPEZOIDAL PLANFORM SHAPE, AND THE AVERAGE OF THE TWO

LIFT DISTRIBUTIONS USING SCHRENKS (19!" APPRO#IMATION

Where again, cr is the root chard length and is the taper ratio. *ollowing the elliptic wing, we can takethe span wise lift distribution to vary like the span wise chord variation. +herefore,

)T2y5 0 )r := , 22;y452=,55>

Where #ris the local lift value at the location of the wing rooty I =H.

ow the total lift must eBual the value found by integrating the lift distribution in the span wisedirection. +herefore, evaluating the integral, we obtain

) 0 2)r 2=< 554;

With this, we have an expression for #r which gives the necessary total lift for the trapeEoidal liftdistribution, namely,

)r0 2;)5 4 2 2=< 55

And therefore,



65/121

)T2y5 0 ;)4 2 2=< 55 := , 2;y45 2=,5>

As a check, for a planar wing 2 0 =5 )T 2y5 0 )4, which is the correct lift per span.

+o use chrenks method, it is necessary to graph the span wise lift distribution given in the eBuationfor the elliptic platform and the above eBuation for the trapeEoidal planform. 'n each case, # is thereBuired total lift. +he approximated span wise lift distribution is then the local average of the twodistributions, namely,

)2y5 0 U:)T2y5 < )E2y5>

An example of this corresponds to the dotted curve in figure.

't should be pointed out that chrenks method does not provide a suitable estimate of the span wise lift

distribution for highly swept wings. 'n that instance, a panel method approach or other computationalmethod is necessary.

A&&e& $lap )oa&s(

#eading-edge and trailing-edge flaps enhance the lift over the span wise extent where they areplaced. +he lift force is assumed to be uniform in the region of the flaps and to add to the local spanwise lift distribution that is derived for the unflapped wing.

+he determination of the added lift force produced by the flaps reBuires specifying a velocity. *or this,the velocity is taken to be twice the stall value, 1&s, with flaps down.

Span %ise Drag Distri+tion(

+he drag force on the wing varies along the span, with a particular concentration occurring nearthe wing tips. An approximation that is suitable for the conceptual design is to assume that

%. +he drag force is constant from the wing root to ;= percent of the wing span and eBual to


66/121

component direction. +he design of the internal structure of the wing is then primarily driven by theneed to counter the wing-thickness bending moments.

#oncentrate& an& Distri+te& 1ing 1eights(

!ther loads on the wing, besides the aerodynamic loads, are due to concentrated weights, suchas wing-mounted engines, weapons, fuel tanks, etc., and due to distributed loads such as the wingstructure.

ince the structure is being designed at this step, it is difficult to know precisely what the finalweight will be. +herefore, historic weight trends for aircraft are used to make estimates at this stage ofthe design. A refined weight analysis will be done later as the initial step in determining the staticstability coefficients for the aircraft.

+he above table gives historic weights for the ma"or components of a range of different aircraft.+hese include the main wing, horiEontal and vertical tails, fuselage, installed engine and landing gear.+he weights of these components are determined from the table as

1 2l$5 0 M+ltiplier *actor

Where the ]multiplier^ is a number that corresponds to a general type and the ]*actor^ is areference portion of the aircraft, such as the wing planform area, W, or the fuselage wetted area, fuse-wetted.

ngines and landing gear mounted on the wing can be treated as concentrated loads. +he wingstructure will be considered as a distributed load. 't is reasonable to consider that the weight of a spanwise section of the wing would scale with the wing chord length, so that with a linear tapered wing, thedistributed weight would decrease in proportion to the local chord from the root to the tip.



67/121

S-EAR AND BENDN' M!MENT ANA)SS *!R 1N'S

+he wing structure can be considered to be a cantilever beam, which is rigidlysupported at the wing root. +he critical loads that need to be determined are the shear forces andbending moments along the span of the wing. +hese taken into account the loads on the wing producedby the aerodynamic forces and component weights, which were discussed in the previous section. Ageneric load arrangement is listed in the table and illustrated in figure.

+o determine the shear force and bending moments along the span, it is useful todivide the wing into span wise segments of width y. A schematic of such element is shown in figure.

SCHEMATIC REPRESENTATION OF SHEAR LOADS AND BENDING MOMENTS ON A

SPANWISE ELEMENT OF THE WING

As an example, the element shows a distributed load, WyH. +he resultant load acting on the element isthen WyH y. 'n the limit as y goes to Eero, y approaches the differential length, dy, and the

resultant load is WyH dy.

+he element shear force, &, is related to the resultant load as

1 0 &V4&y



68/121

+he bending moment, ), acting on the element is related to the shear force by

V 0 &M4&y

'ntegrals can be approximated by sums, namely

V 0 i

n

Wi y

And

M 0 i

n

Vi y ,

Where is the number of elements over which the wing span is divided. !f course,the sums approximate the integrals better as the number of elements becomes large$ however, areasonably good estimate for the conceptual design can be obtained with approximately twentyelements over the half-span of the wing.

'n order to make these definite integrals, the integration summationH needs to bestarted where the shear and moment are known. With the wing, this location is at the wing tipyIb1H,

where & b1HI) b1HI=. ote that in this case, the resultant load on an element is W iIWyH. y ywhich is the Buantity inside the sum in eBuation. 'f the index, ', in the above eBuation indicates theelements along the wing span, with iI% signifying the one at the wing tip, then

V= 07

V; 0 1= < 1;

VL 0 1= < 1;


69/121

wing span is subdivided. owever, with a large enough number of elements the difference should besmall.

+he bending moment on the wing is given by eBn. *or the moments along the wingspan, one should also start at the wing tip where the moment on the element is Eero. +hen the followingformat in the above eBn.

M= 07

M; 0 V=


70/121

SPANWISE DISTRIBUTIONS OF LIFT FORCE, L$ WEIGHT, W$ SHEAR LOADS, V AND

BENDING MOMENT, M



71/121

An example of the application of these eBuations is shown in fig. +he loadscorrespond to those listed in the table and illustrated in fig.

+he top plot in fig. illustrates chrenks approximation of the span wise lift

distribution for the finite span wing. +he solid curve corresponds to the lift distribution for thetrapeEoidal wing. +his is constant along the span because the taper ratio H in this example is %.

+he total span wise load distribution, W, shown in fig. for generic wing includes all ofthe weight and lift components. *or this, the wing was divided into 1= span wise elements. +he sharpnegative spike in the load distribution marks the location of the engine. +he more gradual dip in theloads near 0b1H I =.6 corresponds to the outboard edge of the flaps.

+he span wise distribution of the shear load, &, comes from eBn. +his shows that thelargest shear is at the wing root, with the second largest shear being at the location of the engine.

+he moment distribution, ), in fig. is based on eBn. 't reflects the wing cantilever structure, wherebythe largest moment is at the wing root. +he small peak in the moment distribution near 0b1H is dueto the engine.



72/121

1N' SPAN VS 1N' )!ADN'

1N' SPAN

+he wingspan or "ust spanH of anairplaneor abird, is the distance from one wingtip to the

other wingtip.

MP)#AT!NS *!R AR#RA*T DES'N

+he lift from wings is proportional to their area, so the heavier the animal or aircraft the bigger

that area must be. +he area is the product of the span times the width mean chordH of the wing, soeither a long, narrow wing or a shorter, broader wing will support the same mass. *or efficient steady

flight the ratio of span to chord, the aspect ratio, should be as high as possible the constraints are

usually structuralH because this lowers the lift-induced dragassociated with the inevitablewingtip

vortices. #ong-ranging birds, like albatrosses, and most commercial aircraft maximiEe aspect ratio.

Alternatively, animals and aircraft which depend on maneuverability fighters, predators, the predated

and those who live amongst trees and bushes, insect catchers, etc.H need to be able to roll fast to turn,

and the high moment of inertiaof long narrow wings produces lower roll rates. *or them, short-span,

broad wings are preferred.

+he highest aspect ratio man-made wings are aircraft propellers, in their most extreme form as

helicopter rotors.

1N' )!ADN'

'n aerodynamics,wing loading is the loaded weight of the aircraft divided by the area of the

wing. +he faster an aircraft flies, the more liftis produced by each unit area of wing, so a smaller wing can carrythe same weight in level flight, operating at a higher wing loading. 5orrespondingly, the landing and take-off

speeds will be higher. +he high wing loading also decreases maneuverability.

General ellipse eBuation is,

http://en.wikipedia.org/wiki/Fixed-wing_aircrafthttp://en.wikipedia.org/wiki/Fixed-wing_aircrafthttp://en.wikipedia.org/wiki/Fixed-wing_aircrafthttp://en.wikipedia.org/wiki/Birdhttp://en.wikipedia.org/wiki/Birdhttp://en.wikipedia.org/wiki/Chord_(aircraft)http://en.wikipedia.org/wiki/Aspect_ratio_(wing)http://en.wikipedia.org/wiki/Induced_draghttp://en.wikipedia.org/wiki/Wingtip_vorticeshttp://en.wikipedia.org/wiki/Wingtip_vorticeshttp://en.wikipedia.org/wiki/Wingtip_vorticeshttp://en.wikipedia.org/wiki/Moment_of_inertiahttp://en.wikipedia.org/wiki/Aerodynamicshttp://en.wikipedia.org/wiki/Aerodynamicshttp://en.wikipedia.org/wiki/Lift_(force)http://en.wikipedia.org/wiki/Fixed-wing_aircrafthttp://en.wikipedia.org/wiki/Birdhttp://en.wikipedia.org/wiki/Chord_(aircraft)http://en.wikipedia.org/wiki/Aspect_ratio_(wing)http://en.wikipedia.org/wiki/Induced_draghttp://en.wikipedia.org/wiki/Wingtip_vorticeshttp://en.wikipedia.org/wiki/Wingtip_vorticeshttp://en.wikipedia.org/wiki/Moment_of_inertiahttp://en.wikipedia.org/wiki/Aerodynamicshttp://en.wikipedia.org/wiki/Lift_(force)


73/121

x1 a1H L y1 b1 H I % ----------------------------------------%H

2ut we took Zb in x-axis Za in y-axis

x1

b1

H L y1

b1

H I %

abI 1

a I 1b

a I ;8.86 1 N %9.=1H

aI 1.64

*rom %H

y I Pa1 %- x1b1H

y IPa1b1H N b17 x1H

y I P1.641%9.=11H N %9.=117 x1H

y I P=.=1N 189.96 7 x1H

y I P8.%4-=.=1 x1

x y

% 1.19

1 1.16

4 1.11

6 1.%

evaluation of class 1 fighter

Documents