(c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

AIAA Atmospheric Flight MechanicsConference and Extent

6-9 August 2001 Montreal, Canada A01-37229

AIAA 2001-4008BASELINE DESIGN CONFIGURATION FOR MILITARY JET TRAINER

THROUGH CORRELATIONS BASED ON DESIRABLE FLIGHT CHARACTERISTICS

K Ibrahim*

ABSTRACT

Configuration and performance data of jet trainers andlight attack airplanes has been analyzed. Correlations,based on desirable flight characteristics, have beenevolved for choosing wing airfoil section, thickness tochord ratio, quarter chord sweep, aspect ratio, taperratio, dihedral, aileron size and trailing flap area;fuselage length; horizontal tail area, arm, aspect ratioand elevator area; and vertical tail area, arm and rudderarea for baseline design configuration of military jettrainers and light attack airplanes. These correlationsbring out that a) benefits of supercritical sections interms of improvement in maximum Mach number taperoff at higher Mach numbers, b) aspect ratios of theairplanes with lower sweepback tend to diverge fromthe pitch-up boundary, c) wing planforms of most of theairplanes are less tapered compared to sweep/taperboundary for minimum induced drag and d) fuselagefineness ratios are higher than those optimum for thestreamlined bodies of revolution.

NOMENCLATURE

AAhtb

Wing aspect ratio, b2/SHorizontal tail aspect ratio, bht2/ShtWing span, m

bht Horizontal tailplane span, mC Wing mean aerodynamic chord, mCl, C2 & C3 ConstantsQp Airplane lateral static stability(Qp)r Lateral static stability due to dihedral(Cip)A Lateral static stability due to sweepback

Lateral static stability due to wing locationLateral static stability due to vertical tail

Cma Airplane longitudinal static stability(Cma)fus Longitudinal static stability due to fuselage(Cmo)ht Longitudinal static stability due to H.TailCnp Airplane directional static stability(Cnp)fus Directional static stability due to fuselage

* Former Deputy General Manager (Design),Aircraft Research and Design Centre,Hindustan Aeronautics Ltd., Bangalore, India.

Copyright © 2001 by Ibrahim. Published by theAmerican Institute of Aeronautics and Astronautics, Incwith Permission.

M

hfIfSSanSeieSfShtSrSvtTOt/cVsWWfAc/4F

Horizontal tailplane arm, mVertical tailplane arm, mMaximum Mach numberDirectional static stability due to vertical tailMax depth of fuselage, mFuselage length, mReference wing area, sqmAileron area, sqmElevator area, sqmTrailing edge flap area, sqmHorizontal tailplane area, sqmRudder area, sqmVertical tailplane area, sqmSea level static thrust at ISA conditionsThickness to chord ratio, percentStall speedAirplane weightMax width of fuselage, mQuarter chord sweep, degDihedral angle, degTaper ratio

INTRODUCTION

Past experience constitutes the primary source forselection of baseline design configuration, initial inputsto the iterative process of aircraft design. It is a cut anddried method and therefore time consuming. If, on theother hand, initial inputs are chosen based on thedesirable flight characteristics, conceptual design cycletime can be reduced significantly. There is, therefore, aneed to evolve easy to use design tools for selection ofbaseline design configuration.

Selection of wing airfoil type, thickness to chord ratio,quarter chord sweep, aspect ratio and taper ratio foradvanced jet trainers and light attack airplanes is dealtby Ibrahim1. Whereas a comprehensive exposition ofaircraft conceptual design with simplified but completeset of first order analytical methods with specialemphasis on aircraft configuration layout is provided byRaymer2. Statistical correlations for selecting baselinedesign configuration for advanced jet trainers and lightattack airplanes are evolved by Ibrahim3. Roskam4

presents a rapid method for preliminary sizing in termsof gross take-off weight, empty weight, mission fuelweight, take-off thrust/power, wing area and aspect

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ratio for homebuilt airplanes to supersonic cruisevehicle to a given mission specification. Roskam5

provides configuration data for 12 types of airplanesand also suggests a range for each parameter.

In this paper configuration and performance data of 20jet trainers and light attack airplanes (Table I)1'3 hasbeen analyzed. Correlations based on desirable flightcharacteristics, have been evolved for wing airfoilsection, thickness to chord ratio, quarter chord sweep,aspect ratio, taper ratio, dihedral, aileron area andtrailing flap area; fuselage length; horizontal tailplanearea, arm length, aspect ratio and elevator size; andvertical tailplane area, arm length and rudder size.Multi-variable least square technique is employed.Where dispersion of data did not yield meaningfulcorrelations support has been drawn from militarycombat aircraft data. These correlations bring out thata) benefits of supercritical sections in terms ofimprovement in maximum Mach number taper off athigher Mach numbers, b) aspect ratios of the airplaneswith lower sweepback tend to diverge from the pitch-upboundary, c) wing planforms of most of the airplanesare less tapered compared to sweep/taper boundary forminimum induced drag and d) fuselage fineness ratiosare higher than those optimum for the streamlinedbodies of revolution. The correlations can be effectivelyutilized for selection of baseline design configurationfor military jet trainers and light attack airplanes.

BASELINE DESIGN CONFIGURATION

Selection of baseline design configuration includeschoosing appropriate combination of wing loading(W/S) and thrust to weight ratio (To/W) based on pointperformance requirements in terms of take-off distance,maximum rate of climb at sea level, maximum levelspeed at sea level as well as altitude and landingdistance. Training sortie duration, range and radius ofaction in a light attack role form the criteria forcomputing clean airplane weight and fuel fractionemploying statistical trends. Subsequently variousparameters describing the wing, fuselage andempennage configuration are selected. Correlations tochoose these parameters are dealt in the followingsections.

WING AIRFOIL SECTION, THICKNESS TOCHORD RATIO AND QUARTER CHORD

SWEEP

For military jet trainers and light attack airplanesmaximum Mach number is a measure of dragdivergence Mach number. Parameters that influence the

drag divergence Mach number are, type of airfoil(Conventional NACA/Supercritical) and its thickness tochord ratio, wing sweepback, aspect ratio and operatinglift coefficient.

NACA 6-series sections, when smooth, supportextensive laminar boundary layer thereby givingrelatively low drag coefficients around the design liftcoefficients. These airfoils, therefore, are preferred forhigh performance aircraft. Drag divergence Machnumber of NACA 6-series sections is higher than thatof its predecessors. Amongst the NACA 6-series,NACA 64-series offer minimum drag, relatively higherdrag divergence Mach number and favorablycomparable maximum lift coefficient.

A survey of airfoil sections employed on variousairplanes reveals that about half of them employmodern airfoils. Supercritical airfoils developed byNASA and aircraft companies have improved lift todrag ratio, higher drag divergence Mach number, highermaximum lift coefficient and consistently gentle stall.These airfoils, therefore, are optimally suited formilitary jet trainers and light attack airplanes.Incorporating these airfoils results in a configurationthat will have a thicker wing (less structural weight andmore fuel volume) lesser wing sweep (lower weight andbetter handling and low speed performance) for aspecified drag divergence Mach number. As a resultcruise speed, sustained manoeuvrability, range andflying qualities are improved. On the debit side,because of the increased aft loading, the pitchingmoments for a given lift coefficient are substantiallymore negative than those for the NACA 6-seriesairfoils. Since adequate data, documentedsystematically like that of classical NACA airfoils toenable selection of supercritical airfoils for applicationsis not yet available, wind tunnel tests ought to becarried out beforehand to ensure desired aerodynamiccharacteristics.

Aspect ratios of most of the trainers and light attackairplanes lie in the range of 4 to 6.68. Averageincrement in the drag divergence Mach number due tothree-dimensional effect is about 0.0256 and is assumedto remain same for all the airplanes analyzed. Liftcoefficients corresponding to maximum Mach numberflight condition are less than 0.2. At these liftcoefficient there is no significant effect on the dragdivergence Mach numbers7. Thus drag divergenceMach number of jet trainers and light attack airplanes isprimarily influenced by the type of airfoil, its thicknessratio and wing sweepback. Correlation of wing sectionand planform parameters with their corresponding

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maximum Mach numbers for 27 airplanes is presentedin Figure 1.

Suttend

Der

*T-2A

ar

*Gnat T\*Myster

jee^Hunt

Supercritical AirfoilsMmax = 1.10635 - 0.027151 (t/c)% Cos2Ac/4

e IV

\^AIphdj<

^<

^

>7^^

"%N

skra ^AA

Conventional AirfoilsMmax = 1.090675 - 0.035556 (t/c)% Cos2Ac/4

kT-3^C/>^/

x^^3 105\

Xm339A

arnP x\

Ciran l]j\

Jet Prove

.Aviojet

T-46A

er Saeta

Fouga 90

(t/c)% Cos2Ac/4

Fig. 1 Choice of quarter chord sweep and airfoilthickness to chord ratio

Data segregates broadly into two groups, one belongingto conventional NACA airfoils and the other tosupercritical sections. Equations of the linearlyregressed lines are given below.

Mmax = 1.0907 - 0.0356 (t/c) Cos2Ac/4

for conventional airfoils and

Mmax = 1.1064 - 0.0272 (t/c) Cos2Ac/4

for supercritical airfoils.

...(1)

...(la)

Besides, the correlation brings out another significantaspect of wing design. Converging of the two regressedlines towards higher Mach number suggests that thebenefits of supercritical sections in terms ofimprovement in maximum Mach number taper off athigher Mach numbers.

WING ASPECT RATIO

Aspect ratio influences lifting ability of the wing at agiven angle of attack, the induced drag and angle ofattack at stall. For military jet trainers and light attackairplanes transonic speed capability forms additionalcriterion for selecting aspect ratio.

In a sweptback wing local lift coefficients in theoutboard region are higher than those in the inboardregion. Since the onset of flow separation is stronglydependent on the local lift coefficient, a highlysweptback wing is, therefore, prone to early tip stall.The higher the sweepback and aspect ratio combination,

the farther aft the tip is from the centre of gravity andgreater the forward movement of centre of pressurewhen lift is lost at the tips. Airplane, therefore,undergoes a rapid and uncontrolled nose-up movement,which constitutes the phenomenon of pitch-up. Thussweepback and aspect ratio combination is the primaryfactor in establishing pitch-up tendencies.

Pitch-up boundary in terms of aspect ratio and quarterchord sweep evolved by NASA8 is shown in Figure 2.For pitch-up free flight characteristics it is necessary tochoose a right combination of sweepback and aspectratio, such that sufficient margin is incorporated fromthe pitch-up boundary. Sweepback angles and aspectratios of 27 airplanes are superposed on the NASAboundary. Aspect ratios of most of the airplanes aresufficiently below the NASA boundary. However,aspect ratios of the airplanes with lower sweepbacktend to diverge from the pitch-up boundary. Most ofthese airplanes have been designed for transonic speedcapability required for penetration. High aspect ratiospermitted by the pitch-up boundary impose largetransonic drag penalty and inhibit transonic penetrationspeeds. On the other hand, lower aspect ratios reducethe severity of transonic wave drag and hence thedivergence between the design aspect ratios and pitch-up boundary. Equation of the regressed line is as under.

A - 5.852 - 0.005 (Ac/4)1.7 (2)

Aspect Ratio = 5.85219 - 0.004957 (Ac/4)1'

10 15 20 25 30 35 40Quarter Chord Sweep (Ac/4), deg

Fig. 2 Quarter chord sweep and aspect ratioboundary for transonic capability

WING TAPER RATIO

At moderate to high lift coefficients corresponding tosubsonic cruise and manoeuvre conditions, induceddrag forms a significant (> 50%) part of airplane totaldrag. The induced drag is because of spanwise

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variations of wing loading. It is well known that ellipticload distribution along the span offers minimuminduced drag. Spanwise load distribution is primarilydependent on the wing planform. The larger the aspectratio the lower will be the induced drag. Aspect ratioscannot be larger than those permitted by the NASApitch-up boundary. Increasing the wing sweep causesthe loading to move outboard thereby distorting it fromthe ideal elliptic. Likewise taper causes the local liftcoefficient at the tip to increase. However, because ofreduction in the local chord, spanwise loading will tendto move towards the elliptic. Hence by choosing anappropriate combination of wing sweepback and taperit is possible to achieve near elliptic loading.

A relation between wing quarter chord sweep and taperratio required for approximately elliptical loading,derived by NASA9 is presented in Figure 3 forplanforms without twist, along with design data of 26military jet trainers and light attack airplanes. It isobserved that wing planforms of most of the airplanesare less tapered compared to NASA sweep/taperboundary. It can be deduced that in the case of jettrainers and light attack airplanes a practical lower limitto taper ratio is imposed by structural height required atthe tip to provide adequate room for ailerons and theircontrol elements. Besides the points in Figure 3 lie in aregion where wing loading moves outboard. In majorityof theses designs certain amount of wing twist(washout) has been incorporated. This suggests thatmodern trend is to go for less tapered wing planformsand alleviate the outboard loading through a suitabletwist. Less tapered wings also offer more room forinstallation of external stores and associated controlelements, which are essential for performing lightattack roles.

Aviojet*Parr

MB.326

L2$\^

CL.419 I-

\xSkyf°*\

MB. 339*im^lskra _

t6A^\.

^Kiron

\x^

Taper ratio = 0.55881 Exp(-0.01731 Ac/4>

~~"\. S.21

II

"\

*

~^-\

H

\^

^sr\*G-4

~^~~-

AlvIX

ha jet

""

^

--~4AR.93>

---^

*T 2

~~^~~fn^ary" 921

Gnat T

Ajeet T

HF-24

Quarter Chord Sweep (\/4), deg

Fig. 3 Quarter chord sweep and taper ratio boundaryfor minimum induced drag

Equation of the exponential regression fit nearlyparallel to the NASA sweep/taper boundary is givenbelow.

Taper Ratio (X) = 0.5588 exp (-0.0173 Ac/4) ... ..(3)

WING DIHEDRAL

Airplane lateral static stability, rolling moment due tosideslip, can be expressed as sum total of thecontributions from various sources as below.

(vt ... ... ... (4)

It is observed that all the military jet trainers currentlyeither in development or in series production havestraight or mildly sweptback and tapered low/high wingmonoplanes. Sweep effect, therefore, is relatively small.Airworthiness requirements in regard to crew vision,operational requirements in respect of accommodating aspecified percentile of pilot population and assurance ofstructural integrity by means of carry-through wing donot accord much freedom to the airplane designer toadopt diverse cross-section shapes for trainer fuselages.Besides, closer examination reveals that the relativewidths (w/b) of the trainer fuselages are about thesame. Therefore for jet trainers and light attackairplanes order of magnitude of the effect of low/highwing on the lateral static stability is assumed to remainsame. Order of magnitude of the vertical tailplanecontribution to lateral static stability is small. In view ofthe above, expression for aircraft lateral static stabilitycan be rewritten as under.

Qp - (Cip)r + (Qp)A + Constant (5)

Assuming that for a class of airplanes order ofmagnitude of lateral static stability remains same,substituting for the contributions of dihedral andsweepback in terms of taper ratio, aspect ratio andquarter chord sweep and simplifying yields thefollowing expression for the dihedral.

= Cl [1+C2/A] + C3 Tan Ac/4 ... (6)

Regression of aspect ratio, taper ratio and quarter chordsweepback for low wing aircraft provides the followingexpression for dihedral.

T - 47.39 [1- 4.89/A] - 2.294 Tan Ac/4... (6a)

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AILERONS

The ailerons provide roll control to the airplane.Aileron power must be adequate a) to provide desiredroll rate to the airplane, b) to overcome the rollingmoments arising during cross wind take-off andlanding, and c) to effect required bank angle for acoordinated turn under specific conditions of flight.Requirement at a) generally forms the design criterion.

While providing a steady roll rate, ailerons mustovercome wing damping due to roll. Accordingly,correlation of aileron area with wing area yields thefollowing equation.

= 0.069 S (7)

TRAILING EDGE FLAP AREA

Trailing edge flaps provide additional lift requiredduring take-off and landing. They do not increase theangle of stall. In fact they tend to reduce the stall angleby increasing the adverse pressure gradient over the topof the airfoil, which promotes flow separation. Trailingedge flaps commonly employed on military jet trainersare plain flap, split flap and slotted flap.

The operational requirement of stall speed with flapsforms the design driver for sizing of the flap area andselecting the deflection angle. However, a survey of theflap areas for the military jet trainers and light attackairplanes suggests that the flap area is related to grosswing area. Regarding deflection angle, with a view toderive full benefit, designer prefers to employ angles(35 to 50 deg.) corresponding to the largest incrementin maximum lift coefficient. Correlation of flap areawith wing area (Figure 4) results in a broadband.

Sf, sqm = -1.9365 + 0.2711 S

VJet Squall

Pampa

/

,s

T-4-6AX

Alphajet,

/y/K-8

/Hawk

x

G-4/i

/

+L-39

+IAR.99

MB.33Sx /•

*AMX

ryci

+T-4

Aviojet */

'

/

/

AT- 3

x^ IAR.93

S f , sqm = -1.3991 4- 0.1816 S

\ 16 18 20 22 24Wing area (S), sqm

Fig. 4 Trailing edge flap size

The following regression lines define the upper andlower bounds.

Sf =-1.94 + 0.278 (upperbound) ...

Sf =-1.40 + 0.188 (lowerbound) ...

.-.(8)

Depending upon the requirement of stall speed in thelanding configuration suitable training edge flap arealying within the specified bounds can be chosen.

FUSELAGE LENGTH

Based on practicable considerations, for sizing of thefuselage length two criteria a) practicable fineness ratioand b) recovery from spin are suggested.

The optimum fineness ratio for a fuselage with aspecified frontal area and volume are around 3 and 5respectively. However, in view of the interference fromthe adjoining components, requirement of adequatespace for installation of equipment and appropriate armlengths for the tailplanes, fuselages in practice arefound to be longer than the optimum.

Cockpit configuration in terms of side-by-side ortandem seats along with percentile pilot population tobe accommodated uniquely determines the maximumwidth of the fuselage. Maximum depth, however, isinfluenced by the vision requirements at the forwardand aft crew stations. In regard to length, forebodyhouses equipment, center fuselage accommodates thecrew stations and rear fuselage generally houses theengine. Thus fuselage length also varies dependingupon the type of equipment, crew station and engineemployed. Correlation of fuselage length withmaximum depth for tandem seat configurations yieldsthe following regression line.

Lf=6.6h f (9)

The correlation suggests that fineness ratio of fuselagesfor jet trainers and light attack airplanes is higher thanthat of the streamlined bodies of revolution.

Military jet trainers are required to be capable ofspinning and recovery from the spin upon application ofrudder, a classical recovery control input. Complianceof spin recovery requirement necessitates ensuring ofmore than 1/3 of unshielded (by the horizontaltailplane) rudder area at attitudes relevant to spin aswell as equal distribution of mass along the wing andfuselage (zero loading category wherein lyy). The

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former is ensured through appropriate relative locationof horizontal and vertical tailplanes. Whereasaccomplishing distribution of a given set of massestypical to military jet trainers and light attack airplanesalong the wing and fuselage to achieve Ixx « lyy suggeststhat a correlation ought to exist between the wingspanand fuselage length. Correlation (Figure 5) brings forththe following results.

lf= 1.18b (10)

for military jet trainers and

lf= 1.18+1.25 b ... ...

for light attack airplanes.

(lOa)

,etS

Wing span (b), m

Fig. 5 Fuselage length

While choosing fuselage length care ought to beexercised to ensure that it is not very much at variancewith the value specified by the spin recovery criterion.

HORIZONTAL TAIL SIZE AND ARM

Major contributions to airplane longitudinal staticstability come from fuselage and horizontal tailplane.While providing the required level of longitudinal staticstability to the airplane, the horizontal tail overcomesthe destabilizing contribution of the fuselage.Therefore, airplane longitudinal static stability can bebroadly expressed as

c ~ (Cma)ht + (Cma)fhs

Assuming that the order of magnitude of longitudinalstability for a class of airplanes remains same andsubstituting for contributions of horizontal tailplane andfuselage in equation (11) yields the followingexpression.

[(Shtlht/SC)] = Constant+[(wf2lf/SC)] ... ... (12)

Correlation of horizontal tailplane area with fuselagelength and width yields the following regression.

Sht = 0.862 [(SC/lf) +0.536 wf2] ... ... ... (13)

Horizontal tailplane arm upon correlation with fuselagelength turns out as below.

lht = 0.382 lf ... ... ... ... ... ... ... ... (14)

HORIZONTAL TAIL ASPECT RATIO

Horizontal tailplane with a relatively large geometricaspect ratio enhances longitudinal static stability byvirtue of its high lift curve slope. Besides, large aspectratio reduces the area exposed to the fuselage wakethereby improving horizontal tailplane effectiveness.Both these effects cause a reduction in horizontaltailplane area. Increase in structural weight due toincrease in aspect ratio may balance out with thatcaused by reduction in area. However, in order to retaincontrol of the airplane throughout the flight envelope,horizontal tailplane ought to stall at angles of attackgreater than those at which the wing stalls. This isaccomplished through choosing a tailplane of aspectratio smaller than that of the wing. Accordingly, aspectratios of the horizontal tailplanes are correlated withthose of wing in Figure 6.

.2 9

= 0.6414 + 0.6991 A

Jet Squa us

i^MB.339

3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0Wing aspect ratio

Fig. 6 Horizontal tailplane aspect ratio

The correlation suggests a broad band with upper andlower bounds. Equations for the regressed lines are asunder.

Ah t=- 0.641 + 0.699 (upperbound) ... ... ...(15)

Aht = - 1.218 + 0.420 (lower bound) ... ... ... (15a)

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ELEVATOR SIZE

In a classical design, elevator is sized to provideadequate control at the most forward C.G. location athigh angles of attack. Yet another design practiceAFWAL-TR-82-308110 suggests that during take-offand landing flight phases elevator control power fromtrimmed condition should be adequate to cause a pitchacceleration of -0.18 rad/s (nose down), and 0.13 rad/s(nose up). These methods, however, demand detailedknowledge of aerodynamic, mass and inertiacharacteristics of the configuration.

The higher the level of stability, the larger will be thecontrol power required to execute a manoeuvre.Accordingly correlation (Figure 7) of elevator area withhorizontal tailplane area results in a broad band withupper and lower bounds. Following are the expressionsfor the regressed lines.

= -1.865 + 0.962 Sht (upper bound) (16)

Seie = -1.178 + 0.607 Sht (lower bound) ... ... (16a)

2.4 2.8 3.2 3.6 4.0 4.4 4.8 5.2Horizontal tailplane area (Sht ), sqm

Fig. 7 Elevator size

While the upper bound corresponds to configurationswith elevator deflection only, the lower boundcorresponds to those with elevator deflection andvariable incidence tailplane.

VERTICAL TAIL AREA AND ARM

Contribution to airplane directional static stability fromwing and its relative location are negligibly small andmajor contributions come from fuselage and vertical

tailplane. While providing the required level ofdirectional static stability to airplane, vertical tailplaneovercomes the destabilizing contribution of fuselage.Therefore, airplane directional static stability can bebroadly expressed as

Assuming that the order of magnitude of directionalstatic stability for a class of airplanes remains same,substituting for the contributions of fuselage andvertical tail plane yields the following expression.

(SvJvt/Sb) = Constant + (hf2lf/Sb) (17a)

Correlation of vertical tailplane area with fuselagelength and maximum depth brings out the followingexpression.

vt = 0.099[(Sb/lf+4.53hf2)] (18)

Vertical tailplane arm upon correlation with fuselagelength provides the following regression line.

1 = 0.341^ (19)

Comparison of horizontal tail and vertical tail armssuggests that the horizontal and vertical tails arestaggered with the latter being ahead of the former,primarily for ensuring more than 1/3 of unshielded (bythe horizontal tailplane) rudder area at attitudes relevantto spin.

RUDDER AREA

The rudder provides directional control to the airplane.Classical design norm is to size the rudder powerfulenough a) to overcome adverse yaw to enable theairplane to execute a coordinated turn at low speeds,b) to trim the airplane to maintain alignment with therunway during crosswind landing and c) to opposerotation to facilitate recovery from spin. Another designpractice suggested in AFWAL-TR-82-308110 prefers tosize the rudder to provide a yaw acceleration of not lessthan 0.2 rad/s2 at low speeds (1.2VS). These practicesdemand detailed knowledge of aerodynamic, mass andinertia characteristics of the configuration.

Like in the case of elevator, rudder size is also found tobe primarily influenced by the vertical tailplane area.Areas of vertical tailplane and rudder have beencorrelated in Figure 8.

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Sr, sqm = 0.0402 -h 0.327 S^

Jet S

L-2S

/dlriat-

qualusX/Kiran

-hIAF

Hawk

-T

//

II

T99xXlpJ

^ter-1

/+

y^r(

L-

IAR-9C

/

a Jet^B-339

\T-3

t

^ AMX

Sr, sqm = -0.17 + 0.289 S^

1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0

Vertical tail plane area (Svt ), sqm

Fig. 8 Rudder size

Following are the expressions for the regressed lines.

Sr = 0.04 + 0.327 S^ (upper bound) (20)

Sr = -0.17 + 0.289 S^ (lower bound) (20a)

CONCLUSIONS

Configuration and performance data of 20 jet trainersand light attack airplanes has been analyzed to evolvesimple and easy to use correlations based on desirableflight characteristics. These correlations bring out thata) benefits of supercritical sections in terms ofimprovement in maximum Mach number taper off athigher Mach numbers, b) aspect ratios of the airplaneswith lower sweepback tend to diverge from the pitch-upboundary, c) wing planforms of most of the airplanesare less tapered compared to sweep/taper boundary forminimum induced drag and d) fuselage fineness ratiosare higher than those optimum for the streamlinedbodies of revolution. The correlations can be effectivelyutilized for selection of baseline design configuration.

ACKNOWLEDGMENT

The author expresses his deep sense of gratitude toProf M R Ananthasayanam from the Indian Institute ofScience Bangalore, India, for his continued inspiration,constructive criticism and helpful suggestions duringthe preparation of this paper.

REFERENCES

1. K Ibrahim, "Wing Design for Advanced JetTrainers and Light Attack airplanes - A StatisticalApproach", 3rd National Conference onAerodynamics, Indian Institute of Technology,Bombay, India 3 November 1987.

2. Daniel P Raymer, "Aircraft Design - AConceptual Approach", AIAA Education Series,1989.

3. K Ibrahim, "Selection of Baseline DesignConfiguration for Advanced Jet Trainers andLight Attack Airplanes", - Unpublished.

4. Jan Roskam, "Airplane Design ^ Part I:Preliminary Sizing of Airplanes",DARcorporation, 1997.

5. Jan Roskam, "Airplane Design - Part II:Preliminary Configuration Design and Integrationof the Propulsion System", DARcorporation,1997.

6. R C Feagin et al, "Delta Method, an EmpiricalDrag Build-up Technique", NASA CR 15971,December 1978.

7. Gerald Corning, "Supersonic and SubsonicAirplane Design", College Park, Maryland, 1953.

8. Joseph A Shorthall and Bernard Maggin, " Effectof Sweepback and Aspect Ratio on LongitudinalStability Characteristics of Wings at LowSpeeds", NACA TN 1093.

9. John D Young and Charles W Harper,"Theoretical Symmetric Span Loading atSubsonic Speeds for Wings Having ArbitraryPlanform", NACA TR 921, 1951.

10. Roger H. Hoh et al., "Proposed MIL Standard andHandbook - Flying Qualities of Air Vehicles",AFWAL -TR-82-3081 Volume II, November1982.

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Table 1. Design Data of Military Jet Trainer and Light Attack Airplanes

No.

123

45

6789

1011

12

13

1415

16

17

181920

Country

ArgentinaBelgiumCzechItaly/BrazilFrance/GermChina/PakYugl/RomaniaItalyItalyJapanPolandRomaniaSpainTaiwanU.KU.S.AU.S.AYugoslaviaIndiaIndia

AirplaneName

PampaJet SqualusAeroL-39AMXAlpha JetK-8IAR93MB.339AS.211T-41-22IAR-99AviojetAT-3HawkSkyfoxT-46AG-4Kiran IIAjeet T

EmptyWeightWe, kg

28211300345565503345

2557615031251850370037823364350038553750

36652540325029662579

TakeoffWeightWt,kg

370020004525

96005000350088264400275055006082

4630

457052165150

73653357

476042123629

ThrustatSL

To,kN

15.57

5.9216.8749.1026.48

16.0135.5817.8011.1332.7421.5817.7916.4631.1425.3532.9211.8417.8018.5316.54

Mmax

0.728-

0.6500.865

0.860

-0.9300.7700.598

0.900-

-0.7150.8500.880

-0.6940.8000.6000.907

Wing

AreaS,m2

15.6313.5818.8021.0017.5017.0226.0019.3012.6021.0019.92

18.7120.0021.9316.6922.2215.4019.5019.0012.69

Spanb, m

9.699.049.468.879.119.639.62

10.868.439.949.60

10.1610.6010.469.39

11.8411.779.88

10.706.73

AspectRatioAR6.006.005.203.75

4.805.453.606.105.104.704.60

5.525.605.005.30

6.689.005.006.033.57

SweepAc/4

deg2.950.001.75

27.5028.00

-

35.008.48

15.5027.5018.00

6.581.887.33

21.50

-

3.50

22.007.47

40.00

TaperRatio

X0.590.530.520.500.40

-

0.330.560.460.36

-0.510.600.500.340.380.510.370.400.45

FuselageLength

If, m

10.909.36

12.1312.5511.7510.5013.4110.979.31

11.9613.22

11.0112.5012.5010.50

14.029.00

11.8610.6010.05

MaxWidthw f,m1.90

---

1.801.78

-2.053

-

---

2.061.651.90

--

1.651.371.48

Maxdepthhf,m1.87

---

1.73--

1.69---

-

1.821.951.75

-

-

1.79-

1.52

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