Flying Wing Cargo Airplane

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N A S A Technical Paper 1158

Preliminary Study ofa Large Span-Distributed-Load ,Flying-Wing Cargo Airplane Concept

C A SLloydS. Jernell " ' - , - , , ', MAY1978

N

c.-

f V J A S A


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SUMMARYA preliminary study has been conducted of an aircraft capable of trans-orting containerized cargo over intercontinental distances. The specifica-tions forpayloadweight, density, and dimensions in essence configure thewingand establish unusually low values of wing loading and aspect ratio. Thestructural weight comprises only about 18 percent of the design maximum grosseight. Although the geometric aspect ratiois4.53,it isestimated thatthewinglet effectof thewing-tip-mounted vertical tails increases theeffectiveaspect ratiotoapproximately7.9.Sufficient controlpower tohandlethe large rolling moment of inertiadictates arelatively high minimum approach velocity of 315km/hr(170knots).Theairplanehas acceptable spiral, Dutch roll, androll-damping modes. A

hardened stability augmentation system is required.The most significant noise source isthatof theairframe. However, forboth take-off and approach, the levels are below the FAR-36 limitof 108 dB.The design mission fuel efficiency is approximately 50 percent greaterthan thatof themost advanced, currently operational, large freighterair-craft. The direct operating cost is significantly lower than that of currentfreighters, the advantage increasing as fuel price increases.

INTRODUCTIONLarge,long-range, subsonic cargo aircraft of the future probably willuselarge cargo containers and have payload capabilities much greater than thoseofpresent-day aircraft. Adesignconceptwhich holds promise forsuch anair-plane accommodates payload distribution along thewingspan to counterbalancetheaerodynamic loads,with a resultant decrease in the in-flight wing bendingoments and shear forces. Decreased loading of the wing structure, coupled withthevery thick wing housing the cargo, isexpected toresultinrelatively lowoverall structural weight in comparison withthatof conventional aircraft.There are many potential problem areas associated with this type aircraft,

including aerodynamicefficiency, control (particularly in roll due to the highoment of inertia aboutthataxis), and airport handling because of its largesize. Inorder toevaluate someofthese problems, the preliminary studyofa large distributed-load cargo airplane was performed. Portions of this workwereconducted by theVought Corporation -Hampton Technical Center (contractNAS1-13500)under the technical direction of the Vehicle Integration Branch,Aeronautical SystemsDivision, Langley Research Center. Vought personnelincluded were C. B. Quartero, leader and mission analyst; G. F. Washburn,structures; P. Baldasare, mass properties; L. A. Bodin and R. R. Combs,Jr.,aerodynamics;G. E.Martin, stabilityandcontrol;W. A. Lovell, propulsion;andJ. W. Russell, noise. The results of this study are summarized herein.


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NASA-funded studiesof alternate configurationsaredocumented inreferences 1 to 3. A review of the current airfreight system and its futureprospectsis presented in reference 4.

SYMBOLSValuesarepresented inbothSI andU.S. Customary Units. The measure-ments were madein U.S. Customary Units.

A aspectratiob wingspan

DragCD drag coefficient, qSLiftCT lift coefficient,qS

Cm- -__ pitching-moment coefficient about0.25c\ J* OCYawingmomentCn yawing-moment coefficient, qSb

C,v, directional stability parameter, dCn/df3,per degc localchordc mean aerodynamic chordca speedof soundat ambient conditionscr rudder localchordcvt vertical taillocalchordD diameter; alsodragFb blade passing frequencyg gravitational constanth altitude,alsoheight of vertical tailnweb heightof wing-spar webL/D lift-drag ratioM Mach number


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1 swept-wing-axesnetlimit bending momentAlimit torsion about wing-box elastic axisYea

f o number of fan blades dynamic pressures wing-box-skin shear flowv wing vertical beam shear floweb beam-web shear flow radius wing reference areav vertical-tail area2 . swept-wing-axesnetlimit shearz thrustt a ambient absolute temperaturet jet Jet total absolute temperatureTt,2\ ratioof total absolute temperatureat low-pressure turbine\Tt,I/engine discharge station to thatat fan inlet station/Tt,2\ I ratioof total absolute temperatureat fandischarge station\Tt,l/fan to that at fan inlet station thickness;alsotime2 timetodouble amplitude/c wing-section thickness ratio velocity;alsovertical shear gross weight distance from nacelle lip measured parallel to center line cross-sectionalneutral axis

t angleofattack, referencedtoairfoil center line,deg


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3 angle of sideslip, deg6e elevon deflection, positive fortrailingedge down,deg6f flap deflection,deg6r rudder deflection, positivefor trailing edge left,deg damping ration . wing station, measured from fuselage center line along center lineof wing box9 nose-downpitching acceleration at minimum demonstrated velocity andmaximumgross weight,radians/sec2

< t > roll angle, d e gc o n natural frequencySubscripts:elastic nonrigid structureIe leading edgemax maximummin minimumrigid rigid structuretrim trimmed conditionNotation:BPR ratio of inlet air mass flow of fan to that of core enginee.g. center of gravitydB decibels, referencedto 2 x10~5N/m2EAS equivalent airspeedEPNL effective perceived noise levelHSAS hardened stability augmentation systemHz HertzMLW maximum landing weight


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minimum demonstrated velocity overall airframe sound pressure level

operating weight, emptym revolutions perminute

reserve-fuel landing weight sound pressure level take-off gross weight thrust specific fuel consumption

zero-fuel weight

BASIC DESIGN CRITERIAThe study required the preliminary design of a span-distributed loadair-nces. The basic design criteria are asfollows:Configuration - flying wing, with wing-tip vertical tails and a relativelysmall fuselage for flight deck and crew accommodationWingplanform- 30sweep, no taperAirfoil- t/c =0.20,one ofseveral Langley-developed airfoilsormodifications thereofCargo-compartment dimensions -sufficientto handle2.44m x2.44m(8ft x 8 ft)cargo containersofassorted lengths

*Payload weight- 2 668 933 N(600000Ibf)Payload density-1571 N/m (10Ibf/ft-*),including containerPropulsion - current-production turbofan engines, scaled if necessaryRange-5926 km (3200n.mi.)CruiseMachnumber - at least0.7Runway length-3658m (12 000 ft)maximumCargo-compartment pressurization-noneCargo loading location- at wing tips


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moment. Two airfoils were selected as candidates for application to thepresent design study. Oneutilizedamoderate amountofpositive camber overapproximately theforward70percentof thechord,but was reflexed overtheremaining rearwardsection to provide anessentiallyzero pitching moment aboutthequarter chord. The other airfoil was a modification of an early super-critical airfoilinwhich thecamberwasremovedand thethickness ratioincreased to 0.20.Early in the design study, unpublished wind-tunnel data became availableona 30sweep, distributed-load cargo aircraft model incorporatingthe afore-mentioned reflexed airfoil. These data indicated that at cruise Mach numberandangle of attack, boundary-layer separation existed over roughly the rear-ward30 percent of the upper surface (the region of the reflexedsurface).Sufficient data werenotavailableto ascertain whether theseparationwas aMach number effector due simplyto the lowtestReynoldsnumber. Furthermore,theoretical data fromtheanalysisprogramof reference5(which computes theflow field aboutanairfoil atsupercritical Mach numbers) predicts thatforanassumed lift coefficient of0.40, thedrag-rise Mach number for thereflexed

airfoil is0.03 less than thatfor thesymmetrical airfoil. Inaddition, pre-liminarylayouts of the wing structure for both airfoils showed thatthe sym-metrical airfoil was slightly more efficient in terms of wing volume utili-zation for the cargo compartment. Hence, the symmetrical airfoil was selectedfor the design study.Dihedral.- A wing dihedral angle of 3 wasemployed to alleviate the needfor the relatively long main landing gear required to provide for ground clear-ance of the wing tip and deflected eleven during landing and take-off.

FuselageThefuselagewasoriginally configured solelyforflight deck, crewaccom-modation,andnose gear. However, it was later found necessary to installafuel tankin theunused volumeso as toprovideagreater rangeofcenter-of-gravity management.

VerticalTailsThe wing-tip-mounted vertical tails, designed according to the suggestedguidelines of reference 6, have a quarter-chordsweepof 30, ataper ratio of

0.30, and an aspect ratio of 2.31. The airfoil used is an 8-percent-thickmodification of theGA(W)-1airfoil (17-percent thickness) described in refer-ence7. Thenonplanar lifting surface methodofreference 8 wasusedtoopti-mize cant and toe-inanglesof the fins for the best combination of aerodynamicefficiency and structural weight.

Engines andNacellesThe configurationhas sixturbofan engines (scaled fromthe JT9D-7engine)to provide the required thrust. The engines, mounted onpylonsabove the wing,


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were originally positioned sothat roughly80percentof thenacellewasaheadof the wing leading edge. Later it was necessary to move the nacelles rearwardtolessen thelarge adverse effectof thenacelles andpylons ondirectionalstability aswellas toavoid possible adverse interference drag frompylonslocated within the supercritical flow region of the upper surface of the wing.In thefinalposition,the nacelleinletlip islocated at approximately the35-percent local-chord station.

Controlsand High-Lift SystemThe elevens haveachord equalto 20percentof thewing chord,andextendfrom the 60-percent semispan station to the vertical tails. Maximum elevendeflection is40. The spoilers, requiredto augment roll control becauseofthe high inertia in roll, have a chord equal to 15 percent of the wing chordandare located inboard on the wing as shown in figure 1. The nonsplit ruddershaveachord equalto 20percentof thevertical-taillocalchordand amaximumdeflection of+40. Thehigh-lift system consistsofsimple trailing-edge flaps

having a chord equal to 15 percent of the wing chord and extending from the wingcenter lineto the60-percent semispan station. Maximum flap deflection is20.

FuelTanksThe fuel tanksare located asshown infigure2. Thewing tanksare posi-tioned ahead of the front wing box beam and behind therearbeam. The forwardtanks extend outward to the 50-percent semispan station, whereas the rearwardtanks extend to theinboard main gear wheelwells. The fuselage tankwaspro-vided to widen the range of control over the center of gravity.

Landing GearThe landing gear(see fig.1) iscomposed of a twenty-wheel, four-strutmain gear and a two-wheel nose gear. The inboard pair of main gear, utilizingsix-wheel bogies, are located rearward of the wing-box rearbeam at approxi-mately the 33-percent semispan station. Theoutboard gear have four-wheelbogies and are positioned forward of the wing box front beam at approximatelythe 77-percent semispan station. Tofacilitate landing load distribution,theoleo-pneumatic suspensionsof thepairon each sideare interconnected. Alanding gearofthis configuration might require steeringof atleastonepair

ofmain gear; however, suchananalysis wasbeyond thescope ofthis study.STRUCTURES

Since all structural components other thanthewingbox are of conventionaldesign, thestructuralanalysesof these items were confined to component layoutanddetermination ofmass properties using statistical data. Becauseof theunique geometry and loading requirementsof thewing,adetailed studywasper-formed wherein thewing-box structural concept wasdeveloped and thedimensionsof its structural components were analytically determined.8


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The final wing-box design, shown in figure 3, incorporates conventionalstiffened, stressed-sheet structure constructed primarilyof2024-T3aluminum,7075-T6aluminum being employed wherethehigher allowable stresscan beusedtoadvantage. Two vertical beams, reinforced by vertical stiffeners, are con-nected by beam-type upperandlower ribcaps which, inturn,are supportedbytensiontubes locatedbetweenthe cargo bays. The ribcapssupport thestringer-stiffened wing skins. Thelower ribcaps also support the spanwisebeamsof thecargo subfloor. Figure4,which shows across sectionof thewingnormal to the leading edge, provides additional details of the wing boxat a typical rib station.

Maximum design loads criteria established earlyin thestudyare(a) 2.5-g balanced flight maneuver at maximum gross weight and cruise Machnumber andaltitudeand(b) 2.0-g taxiatmaximum gross weight

The final structural analysis is based on the following additionalconditions:(1)Maximum design gross weightof 6 052 250 N (1 360 600 Ib).(2)For 2.5-g flight maneuver,thecenterofgravityislocated at

0.295, M =0.75, and altitude, 8595 m (28 200 ft)(3) For 2.0-g taxi,thecenterofgravity positioned at0.35cThe procedures employed in the design of the wing box are based on themethods of reference 8. Although the analyses are of comparatively limited

scope, the results are considered to be adequate for preliminary design pur-poses. The values of wing shear, bending moment, and torsion, calculated forthemaximum-design-load conditions, are shown infigures5, 6, and 7, respec-tively. The airloads for 2.5-g flight maneuver were calculated by using acomputer program based on the method of reference 9.Because of the simple wing-box geometry and the desire to minimizecompo-nentgage changes, structural analyses were conducted only at the eight struc-tural semispan stations shown in figure 8. As will be noted, two stationsrepresent the ribs supporting the inboard and outboard main gear. A shear flowdiagram similar to that of figure 9 was generated at each station to determine

beam-web and skin thicknesses.The vertical shear is distributed equally between the two vertical beams.Thebeam webs arepermittedtobuckleand aredesigned tocarrytheverticalshear and the wing-skin shear flow due to torsion with the webs in the diagonaltension-field condition. The variation of web thickness along the structuralsemispan isshown infigure10. The web stiffeners, spaced at38.10 cm (15in.)intervals, are of the geometry shown in figure 11. The spanwise variation ofstiffener cross-sectional areaispresented infigure12.


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The beamcaps,stringers,andskinsaredesigned to carryallbendingloads. Inaddition, the skins, whichare allowed tobuckle,alsosupport thechordwise shear loads due totorsion. Hence, the sizingofthese componentsand determination ofstringer spacing required several iterations. The loadson the beam caps and stringers (including effective skin) were calculated atthe eight semispan stations usingadistanceof307.34cm(121.00 in.) betweenbeam-capcentreids and anaverage distanceof365.76cm(144.00in.) betweenstringer centroids. Sectional geometriesofthese components are shown infigure11. The spanwise variationofbeam-cap and stringercross-sectionalareas arepresented in figures13 and 14,respectively. For agiven skin thick-ness,the allowable buckling chordwise shear stress is proportional to stringerspacing; therefore, theclosestringer spacing (20.42cm (8.04 in.)) allowsarelatively high buckling stress. The variation of skin thickness along thesemispan isshowninfigure15.

The wing box structure includes 130 frame-type ribs. In addition, fourbeam-type ribs of heavier forged aluminum are located at the main-gear attach-mentpoints. All ribsarespaced at76.20-cm (30.00-in.)intervals. Theupperand lower I-beam rib caps are designed for the load resulting from the 2.5-gflight maneuver. Theanalysis andsizing were performed only at wing station2001.32cm(787.92 in.) (measured along the wingbox center line), which is thelocationof the rib supporting the inboard main gear. The rib cap loads at thispoint were assumed to be typical of those throughout the wing box.

The cargo subfloor structure consists of the lower rib cap, which alsoserves as the main chordwise subfloor beam, and four spanwise beams locatedbelow eachof thethree cargo bays. The spanwise beams, consisting ofupperand lower caps andstiffened webs, havea25.40 cm (10.00in.) depth deter-mined by designlayout. No structural analyses were performed on the subfloorcomponents.Although the study airplane exhibits a low ratio of structural weight togrossweight in comparison with conventional cargo aircraft, weight reductionislimited since neither weight nor the external loads are uniformly distributedalong thespan. Component weights of such items as propulsion units, fuel andtanks, and landing gear cause considerable spanwise variation of weight, andrealistically, even the assumed uniform distribution of payload weight is anideal case which would rarely be encountered. With regard to external loads,theairloads are not uniform because of the aforementioned impracticability ofutilizing wing twist. The results of the studies indicated that the extremedepth of the spars is not as advantageous as might be expected since the fail-

ure modes occur in buckling with very low maximum allowable stress. Preliminaryestimates, wherein extrapolations of empirical data representing all-aluminumstructures were utilized, indicated a wing structural weight of approximately287 N/m2 (6lbf/ft2);however, detail design studies predictedan all-aluminumweightofapproximately 421 N/m2 (8.8lbf/ft2). Further studies, whereinitwas assumed that90percentof thewing secondary structure, control surfaces,and flaps could be constructed of epoxy composite material, indicated thattheoverall wing weight could bereducedto 402N/m2 (8.4lbf/ft2).

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MASSPROPERTIESThe mass properties analysis consisted of the determination of aircraftments of inertia, and center-of-gravity ranges. Mass properties ofewing box were obtained analytically by using data generatedduring the design. Thoseof thewing secondarystructure, controlsurfaces,dflaps were estimated by using statistical data, with an adjustment for a componentweight reduction throughthe use ofepoxy compositeial for 90 percent of the structure. The fuselage properties are those of typical subsonic transport forebody, adjusted forstructural modificationduethe increased loadsof the nose gear and fuselage fuel tank. Data for thetails, landing gear, nacelles, and fuel system were obtained statisti-ly with the use of a computer program developed by Vought-Hampton. Themass JT9D-7engines were calculated with the use ofdata and scale factors provided by the manufacturer. Mass properties

in operation, with adjustments applied where appropriate.The weight breakdown by component and by group is listed in table I. Thehas anoperating empty weight of 1 719 682 N(386600Ibf)and adesignight of 6 052 250 N (1 360 600Ibf). A bar graph of the weight break-n is provided in figure 16. The structural weight comprises only aboutpercent of the maximum gross weight and exemplifies the magnitude of struc-efficiency achievable through the utilization of the span-distributedconcept. Unpublished in-house studies by severalairframemanufacturerswould notpressurization; therefore, no studies were conducted to determine theciated with pressurizing the cargo compartment.The moments of inertia about the stability axes and the product of inertiacon-ons. Of course, the roll inertial moment is relatively muchgreater thanose of conventional cargo aircraft which carry the payload in the fuselage.The center-of-gravity gross-weight envelope is presented in figure 17 forassumed uniform design-payload distributionandalsofor theferrymission. limitrepresents the restriction imposed by thecontrolpower foraircraft rotation during take-off. Therearwardrepresent longitudinal dynamic stability restrictions. As will be noted,both thedesign-payload andferry missions, therearward dynamic limitsthe approach mode severely restrict utilization of the reserve fuel.of thestudy. The optimum cruise center-of-gravity position (zero elevon is0.29c. The fuel distributionforvarious pointson thecenter-ss weight(GW)envelope are presented in tableIII.

AERODYNAMICSBecause of the high span and inherent low wing loading associated withconfiguration, both span and chord were held to the minimum required for

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cargo containers, container clearance, andstructural thickness. Basedon theresults of the final structural analysis, values chosen for the span and stream-wise chord are88.39m (290.00ft) and19.51 m (64.00ft), respectively. Theresultant aspect ratiois4.53.Incomparison tocurrent cargo aircraft,theconfigurationhasnumerousunconventional features which affect the aerodynamic characteristics, includingthe low values of wing loading and aspect ratio, ahigh section-thicknessratio, wing-tip-mounted vertical tails, no horizontal tail, and upper-surface-mounted engines.Incomparison to conventional cargo aircraft, the study airplane exhibitsrelatively lowinduced dragdue to the lowspan loading. However, theconfigu-ration develops relativelyhighprofile dragdue to thehighthickness ratio.Also,the high thickness-ratio wing posed a design challenge because of thelarge adverse pressure gradients over the rearward surfaces at cruise condi-tions, which resulted in an increased tendency for flow separation. The reso-lution of the separation problem was complicated by the requirement for very

low pitching moment, which negated full implementation of supercritical airfoiltechnology. In thelatter partof thestudy,aneffortwas madetoemployasmall amount of camber to improve the aerodynamic efficiency; however, thisapproach was abandoned becauseof center-of-gravity limit problems. Frictiondragwas calculated by standard methods, using flat-plate turbulent frictioncoefficients adjusted for the effects of supervelocity, interference, pro-tuberances, gaps,and boundary-layer separation near lifting-surface trailingedges. Nacelledragwas also adjusted for boattail effects andlossofleading-edge suction.The induced drag was calculated by using the method of reference 9. Inthis method the configuration was represented as planar surfaces conforming tothe camber planes of the wing and vertical tails. Although the geometric aspectratio is only 4.53, the effect of the wing-tip-mounted vertical tails is toincrease the effective aspect ratiotoapproximately 7.9.Atailless design incurs large trim dragpenaltiesif thetrimming momentsareobtained by means of theelevens. This effect is even more pronounced inthe present configuration since a moderate upward deflection of the elevensignificantly decreases the induced efficiency increment of the vertical fin.Thus,trim is obtained by fuel management wherein fuel is pumped between tanksso as to maintain zero eleven deflection; in cruise, therefore, trimdrag iszero. At take-off and landing, dynamic stability limits the allowable travel

of the center of gravity. Appropriatetrim-drag penalties were assessed againstthe aircraftin thelandingand take-off configurations.The increase indragdue tolocalizedsupersonic flow was determined fromtwo-dimensionalairfoil calculations by using the computer program of refer-ence5. Adjustments were madeforthree-dimensional effects using simple sweeptheory. Drag-rise increments of the fuselage, and engine nacelles and pylonswere neglected since sufficient experimental data werenot available. However,itis believed that the contributions of the components are relatively small.

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Thelift characteristics, including flap andelevendeflections, were byusing the computer programofreference 10. This method calculates aerodynamic characteristics ofwing-body-tail combinations insubsonic and flow. Thewing andfuselage of theconfiguration are as alarge numberofpanels, eachofwhich contains aerodynamicities. Because of the limitations of the program, the enginenacelles pylons were not included in the input geometry; however, the effects ofcomponentson lift are believed to be minor because of enginelocation.thod of reference 11 was employed to account for the effects of engine

Lift-drag polars,withandwithout ground effect(h/b 0.1 and h/b 1), shown in figures 18 and 19 for the landing and take-off modes, respectively.deflection of 20 is used for both take-off and landing. The differencepolars for the twoflight modesis due tothrust effects on trim require- Figure 20presentsthe variation oflift withangleof attack for the

Cruise lift-drag polars are shown in figure 21. The corresponding lift- ratios areshown in figure 22. The curve for M =0.75,which has a maxi-of19.00, compares favorably withthecombination of lift- ratio(L/D=18.65)and specific fuel consumption computed for thecruiseAs will be discussed in the section "Mission Analysis," these optimumespond to the maximum range as determined from the Breguet range

STABILITYANDCONTROLThe static and dynamic analysesof the aircraft stability and control are ondata generatedby themethodofreference10, thedataofreference 12,epreviously discussed aerodynamic and mass-properties data, and the methodsreference13.

CriteriaThecriteria employed indeterminingthestability andcontrol require-14,withtheexception of thelongitudinalwhich are based on unpublished data. The longitudinalare asfollows:(a)For all weights and center-of-gravity positions, the time to double be greater than2seconds(SASrequirement because aircraftsstatically unstable over most of operational center-of-gravity range).(b)The forward center-of-gravity position during take-off shall beability toprovidetherequired control power foraircraft and tomaintain take-off lift coefficient.

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(c) Therearward center-of-gravitypositionduringapproach shallbedetermined by the ability to provide anose-downpitching acceleration of0.08 rad/sec^atminimum demonstrated velocityandmaximum gross weight.The criteriafordeterminingthe lateral-directional stabilityandcontrolrequirements are asfollows:

(a) The aircraft shall have positive effective dihedral.(b) The aircraft shall be directionally stable for all flight modes.(c)Thereshallbe adequate on-the-ground directionalcontroltoprovide trimin a56-km/hr (30-knot),90cross wind.(d)Themin imu m cross-wind control velocity shallbesufficientlylowtoallownose-wheelsteering.(e) There shall be adequate directional control to counteract an

outboard engine failureatmaximum-thrust engine-failure velocity.(f) At approach velocity, the lateral control shallbesufficienttoprovidearoll-responsecapabilityof 30within2.5secondsafter initiationof arapid, full lateral control input.


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data alsoshow the significant effect ofcenter-of-gravitypositionon elevendeflection requiredfor trim during climb-out and approach.The longitudinal-control capabilities for the climb-out and approach modesare replotted infigures27 and 28,respectively/ along withthe staticallydetermined limitsfor center-of-gravity traveland thecorresponding trimmedliftcoefficients. Theclimb-out forward static center-of-gravity limitof0.28cwas determined by the control power required to rotate the aircraft atavelocity of 263km/hr (142 knots) at maximum take-off gross weight. Therearward static center-of-gravity limitforboth climb-out andapproach isO.SOcand isbasedon theability to provideanose-downpitching accelerationof 0.08 rad/sec2atminimum demonstrated velocity andmaximum gross weight. Theapproach forward static center-of-gravity limit of0.235is not determined bymaximumcontrol power, but on the ability to attain a lift coefficient 1.5 timesthe approach lift coefficient. Since the aircraft is statically unstable overmost of the center-of-gravity range, a hardened stability augmentation system(HSAS) isrequiredfor stability. (HSASis abackup stability augmentationsystem havinga reliability comparable tothatof theprimary structure.)Controls-fixed dynamic analyses of the aircraft were conducted for theclimb-out and approach modes. The estimated time required to double amplitudeasa function of center-of-gravity position is shown in figure 29. According tounpublished data,a2-secondminimum timetodouble amplitude is thelimitforwhich a currentHSASwould be able to provide adequate stability. The result-ing rearward center-of-gravity limitfor the initial climb-out atmaximum grossweight is0.3095. For the approach mode, the rearward limits are0.3045and0.3185, respectively, for themaximumand reserve-fuel gross weights. Theselimits, which impose greater restrictions on rearward center-of-gravity travelthanthe aforementioned static limit, prevent the use of eleven settings opti-mized for maximum lift during take-off and landing. Therefore, efficient opera-tion of the aircraft in these flight modes would require the development of avery rapid reaction (fast response) control system. However, sucha system,which might include small secondary surfaces on the elevens, was not analyzed inthe present study.The rearward dynamic center-of-gravity limit for the clean configurationduring the climband acceleration mode isshown infigure30 (fortheminimumtimeof 2 seconds to double amplitude). The rate of change of the rearwardlimitwith aircraft velocity is sufficiently low toallowthe use of fuel trans-fer formaintaining thecenter ofgravity withintherequired limits.

Lateral-Directional Stability andControlThe methods of reference 13 were employed in determining the lateral-directional characteristics. Although the engine nacelles and pylons generatealarge partof theside force, thesecomponentshavea relatively small effecton yawing moment since their longitudinal position is near the aircraft centerof gravity. The vertical tails are considerably larger than those requiredtomeet the criterion that Cn~ > 0. However, tail design was not based on

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directional stability minimum requirements. Instead, the tails were designedprimarily to increase aerodynamic efficiency by following the winglet designguidelines. (See ref. 6.)Figure31 exhibits the effects of control deflection and local rudder-tailchord ratio on the directional control capability of the aircraft. Also shownisthe minimumcontrolpower necessary to meet the requirement of maintaining astraight flight path during take-off withanoutboard engine inoperative. Basedon these data,arudder-tail chord ratioof 0.2 and amaximum deflectionof40were selected.The lateral response of the aircraft was estimated by solving the single-degree-of-freedom equation of motion in roll for a maximum step control input.Theresults arepresented infigure32 for thethree levelsofflying quali-ties requirements specified inreference14. Thedata indicate thattheair-planehassatisfactoryroll responseat anapproach velocitygreater than315 km/hr (170 knots); this is a considerablyhigher velocity than the 278 km/hr(150 knots) believed to be desirable. The level-2 requirement can be met at a

velocity of 248 km/hr (134 knots). These speeds, rather thanmaximum lift,controlthe aircra ft approach speed. A lower approach velocity could beattained by aroll-responserequirement reduced from those of reference 14.Extensive developmentof more powerful lateral control systems would be neces-sary to reduce the landing speed further; however, such development is beyondthescopeof this study.The ability of the aircraft to meet steady-sideslip trimrequirementsduring take-off andlandingwasestimatedby solving the two-degree-of-freedomequations for roll and yaw steady-state trim. The results indicate that 36 per-cent of the maximumrudder deflection and 46 percent of the maximum elevendeflection arerequiredtomaintainawings-level approach with a 10sideslipangle. This iswellwithin the specified allowable limit of 75-percent maximumcontrol deflection.Figure33showstheminimum eleven andrudder deflections requiredforlateralanddirectional control during thetake-off ground runwitha 56-km/hr(30-knot), 90crosswind. It will be noted that adequate control is availableabout both axes at amini mum velocityof 145km/hr (78knots). Control atlowervelocities can be accomplished by nose-wheel steering.An examination of the roots of the characteristic equation of motion indi-catesthat the airplane has acceptable spiral and Dutch roll modes and an accept-

able roll-damping mode. Table IVpresentsacomparison of theinherent lateral-directional characteristicsof theairplane withtheassumed requirements.

PROPULSIONThe engines selected for the study are scaled JT9D-7turbofans which havebeen sized toprovideaninstalled static thrustof 240 200 N (54 000Ibf) each

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atsea-levelstandard atmosphere conditions. The engineis oftwo-spool,axial-flow design withhighbypassand compression ratios. The production engine max-imumambient temperature limits for constant thrust, as recommended by themanu-facturer, are asfollows:(a)Take-off thrust- standardday + 12 C(b) Maximum climb thrust-standardday + 10 C(c)Maximum cruise thrust-standardday + 15 CAlthough permitted within these recommended limits, constant thrust opera-tion above standard day temperature results in a considerable increase in fuelconsumption. A detailed description of theproduction engine, along with basic(uninstalled) performance data, is presented in reference 15.Theunsealed, installed engine performance data were generatedby correct-ingthe basic performance data for inlet recovery, service airbleed, and auxil-

iary power extractionby using themethods ofreference 15.Typical installation lossesin thestudy airplane engine performanceareasfollows:

M

0.40.75

hm0

6 10010 670

ft0

2000035000

Thrust rating

Take-offMaximum climbMaximum cruise

Atmosphericconditions

Standardday + 10 CStandarddayStandard day

Percent changeT

-8.5-5.3-5.5

TSFC1.62.74.6

With the exception of take-off performance, the data are based on sea-levelstandard atmospheric conditions. However, the take-off data were computed forsea-level standardday + 10 Csince, as specified in Federal AviationRegulations, Part 36, these are the atmospheric conditions at which enginenoise shall beevaluated. These atmospheric conditons alsomeettheFederalAviation Regulations Part 25 requirements for determining aircraft take-offperformance.Thedata indicate thatat take-off velocities andaltitudes, theprimaryand fannozzlesare operating subcritically; thatis, thefully expandedexhaust flow areas are equal to the respectivenozzle throat areas.Theinlet recovery presented infigure34 isbased on thegeometryof aninlet employed in the study documented in reference 16. Although the inlet wasoriginally designed for a cruise Mach number of 0.98, it was selected for thepresent study since it exhibits a relatively high pressure recovery of0.994atthe cruise Mach numbers considered herein. It was assumed that the inlet massflow isequal tothat requiredby theengine throughouttheflight envelope;

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hence, performance was not penalized for inlet spillage drag. The engineser-vice airbleed schedule isshown infigure35. Power extraction for electricalandhydraulicsystemswas held constant at48.5kW(65.0 hp). It was assumedthatthenozzleefficiencies of the scaled and reference engines are of equalmagnitude;therefore, no additional performance penalties were assessed fornozzles.Thecharacteristicsof thescaled engine were obtainedbyusingthemethodsof reference 17. Flow rate, exhaust gas massflow,and fully expanded exhaustgasarea were adjusted by therelative thrust ratio (ratioof required installedthrust toproduction engine installed thrust). Theeffectsofrelative thrustratioon fanrotational velocity (rpm),andengine weightanddimensionsareshown in figure 36. The weight of the scaled engine is 54 206 N (12 186Ibf),including manufacturer-furnished standard equipment. This weight does notinclude the inlet, fan cowling, nozzles,or engine-driven airframeaccessories.The installed performance datafor theclimbandcruise modesarepresented infigures 37 to 41. Datafor thetake-off andpart-power cruise modesareshowninfigures42 to 49.Thenacelle incorporates afull-length fanduct,and coplanar primaryandfannozzles. The inlet length isequal to themaximum inlet diameter. Thenozzlelengths are equal to 1.5 times the primary nozzle diameter. The maximumnacellediameter is equal to the maximum inlet diameter plus40.6cm(16.0 in.)forengine-driven accessories andnacelle ventilation. Thenacelle externaldimensions arepresented intableV.

MISSIONANALYSIS

Thedesign-mission criteria specify thattheaircraft shallbecapableoftransporting a 2 668 933 N(600000Ibf)payload aminimum distanceof 5926km(3200n.mi.)at a cruise Mach number of at least 0.7 and shall require arun-way length no greater than 3658 m (12 000ft). (Seesection "Basic DesignCriteria.") Aspreviously mentioned, theengine selectedfor thestudy is ascaledJT9D-7turbofan. The purpose of the mission analysis is to optimize therequired thrustforminimum fuel consumption and toobtaintherequired fuelweights andgross weights,aswellas todeterminetheperformance. Allperfor-mance characteristics arebasedonstandard atmospheric conditions, take-offandlanding data being calculated forsea-level altitude.

Performance CriteriaThecriteria employed indeterminingthevarious performance parametersare:Take-off.-The take-off distance, based on Federal Aviation Regulations,Part25, isdefinedas thegreaterofeither1.15timestheall-engine take-offdistanceor thebalanced field length withoneengine inoperative. Fuelallow-ance includes 10 minutes at taxi power and 1 minute at take-off power.

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Acceleration and climb.- A constant equivalent airspeed shall be maintaineduntilthecruise Mach numberisreached.Cruise.- A cruise climb shall be performed at altitudes optimized for min-imumfuel consumption unless constrained by the service ceiling.Reserve fuel.- The total mission fuel shall include the reserves recom-mended by the Air Transport Association for international flights, consistingof allowancesfor:(a)Increased trip timeof 10percent(b)Missed approach, followed by acceleration to climb velocity(c) Flighttoalternate airport, 370-km (200-n.mi.) distance(d)Hold for 30minutesat analtitudeof 457 m (1500 ft)

Method ofAnalysisThe take-off and landing performance data were generated with the use ofunpublished computer programs developed by the Vought Corporation, HamptonTechnicalCenter. Mission performance was evaluated with the use of an unpub-lished mission analysis computer program developed at the Langley ResearchCenter.

Performance CharacteristicsPreliminaryestimates indicated that becauseof therelativelylowwingloading, engine size isdeterminedbycruise ceiling rather thanby take-offfield length. Inordertodeterminethedesign mission engine sizeandfuelweight, several iterationsof themission performance calculations were required.Thefinal results, presented infigure50, show theeffects ofinstalled thruston take-off field length and design mission range. These data are based on amission fuel weightof 1 663 635 N (374000Ibf). The selected scale representsanengine which generates a sea-level standard day installed take-off thrust of240 204 N (54 000 Ibf). The corresponding design mission range is 5954 km(3215 n.mi.). Thetake-off field lengthatmaximum gross weight with20flapdeflection is 2499 m (8200ft), which is considerably less than the specified

maximumallowablefield length. Theeffectofengine sizeonoperating emptyweight and gross weight areshown in figures 51 and 52, respectively.Take-off rotationisinitiatedat avelocityof 252 km/hr (136knots).Lift-off is accomplished at an angle of attack of 5.5 and a velocity of282 km/hr (152knots). The climb segment is performed at an equivalent air-speedof 519 km/hr (280 knots).Theeffectsofcruise Mach numberon thedesign mission lift-drag ratio andrange areshown infigure53. A reductioninMach numberto 0.68 resultsin an

19


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increase in mission range to 6543 km (3533 n. mi.), which is 10 percent greaterthan thatfor acruise Mach numberof 0.75.The effects of gross weight on approach velocity and landing distance fora flap settingof 20 arepresented in figure54. Theapproach employs a 3glideslope. The relatively high approach velocityisdeterminedby therollresponsecapabilityof theaircraft. The spoilers assist inbraking during the

ground roll. The landing distance for the design mission landing weight is3018 m (9900ft). For thedesign mission take-off gross weight,thedistanceis3200m (10 500 ft).The variationof payload capabilitywithrangeisshown infigure55. Forthe lower payloads,the aircraft is capable of very long range because of thelarge wing volume availablefor fuel stowage.A summary of the design mission performance characteristics is presentedin tableVI. Ofparticular interestis thedesign mission fuel efficiency,which isestimated to be1.19 Mg-km/N (3.16 ton-n.mi./lbf)offuel burned.

This value is approximately 50 percent greater than that of the most advanced,currently operational, large freighter aircraft.

NOISETheengineandairframe noise characteristicsof thestudy airplaneduringtake-offand approachwere estimatedat themeasurement points (ref. 18)shownin figure56. Point 2represents thelocation ofmaximum sideline noise alonga line parallel to and 649 m (0.35 n. mi.) from the runway center line. Themethods employed and the results are discussed in the following sections.

Methodof AnalysisEnginenoise.-The noise characteristics of the fan and jet were evaluatedseparately and then combined to determine the overall engine noise level. Thefan noise characteristics were determined according to the method of refer-ence 19, which predicts the variation of fan sound pressure level SPL withfrequency and directivity at a source noise radius of 46 m (150 ft). Frequencyand directivity angle are treated as functions of fan performance factors.This technique assumes that a fraction of the mechanical work is converted intooutput sound power; hence, both the total temperature rise and mass flow of the

fanwere used indeterminingthe fansourcenoise. Fannoise isalsoaffectedby thedesignandoperating Mach numbersof therotor tips,thenumberofstator vanes, the blade-passing frequency, and the rotor-stator spacing ratio.Theblade passing frequencyis theproductof thenumberof fanbladesand fanrevolutions persecond. The rotor-stator spacing ratio is the average axialdistance between therotor bladesandstator vanes dividedby theaveragerotor-blade axial length. TableVIIlists typical input parameters usedtopredict thefan-source-noise sound pressure level for each engine. It should benoted that the fan total temperature rise, mass flow, and rotational velocityare dependent on engine performance, which variesduring the take-off mode.

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The jetnoisecharacteristics of the engines were predicted by using thennular and single jet methods of reference 20, which predict the variationjet noise sound pressure level with frequency and directivity angle at ae radius of 46 m (150 ft). The magnitude of the jetnoiseis dependent onvelocityand the flow characteristics of each jet, including exit area velocity, massflow,total temperature ratio,anddensity. Typical inputfor predicting coannular jet noise are listed in tableVIII.Following thrust cutback, the mass flow of the fan jet is considerably than that of the primary jet (see tableVIII). Therefore, for this

jet method to the fan exitflow. Figure 57 shows the variation of therce noise sound pressure levels with frequency at a directivity angle of forboththe fan and jetfollowing thrust cutback. Asindicated,jet is predominantat the lower frequencies, whereasfan source noise at thehigher frequencies. However, at theobserver locations,thenoise levels arepredominant because ofatmospheric attenuationof thehigh fannoise.Before combining the source noise values of the fan and jet, correctionsapplied to each spectra to account for the wing shielding effects due tothe engines above the wing. Based on preliminary data correlations the Pratt and Whitney Aircraft Group, noise reductions of 3 dB were appliedwith the fan or jet noise source directivity. Forstudy airplane, these reductions were applied to the fan source noise at

lessthan 100 relative to the inlet axis. The resulting levelsto obtain a total engine noise spectra over the applicable ranges of

For agiven instantduring take-off theengine source noise iscomputed ateobserver location along thedirectivity angle determinedby theobserverion relative to the aircraft. The prediction method includes the effectsreflection. Thus, at each observer station on the ground at a particulare is a preceived noise level generated by the engines. The time

(EPNL)at each observer position.Airframenoise.-Reference 21 presents the results of a study in which air-noise data were correlated formultiengine commercial andmilitary air-with aspect ratios from approximately 7 to 10. As part of the study docu- inreference22,airframe noisewasalso evaluatedfor an arrow-planformanaspect ratioofapproximately 1.9. thestudy airplanehas anaspect ratioof4.53, it wasassumed that air-noise could beapproximated byaveragingthevalues predictedby the Figure58showsthe minim um altitude as a function of fly-airframenoise level of 108 dB at the FAR 36 center-linepoint as predicted by the reference methods. Alsoshown is the representing the estimated noise level of the study airplane. However,previously mentioned in the section "Aerodynamics," the winglet effect of

21


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the vertical tails increases theeffective (aerodynamic) aspect ratio toapproximately 7.9. Hence/ it is believed that the average-value curverepresenting the study airplane may be somewhat conservative.

PredictedNoise LevelsEngine noise at take-off.- In order to minimize engine noise at the center-line measurement point (located 6486m (3.5 n.mi.) fromthebreak-releasepoint), engine thrustwasreduced 5944m (19 500 ft)from brake release. Thethrust cutback point was determined from the results of a previous studyreported in reference 23. To optimize the noise level of the study airplane atthemeasurement points, the take-off profile was varied to evaluate the effectof cutback altitudeon theengine effective perceived noise level (EPNL)atboth the sideline and center-line measurement points. Figure 59 indicates thatas altitude is increased engine sideline noise increases and center-line enginenoise decreases. Figure 59 also exhibits a decrease in the overall airframesound pressure level (OASPL)at the center-line measurement point as thrust

cutback altitude is increased. It will be noted that airframe OASPL isapproximately 5 dB greater than the center-line engineEPNL. Since OASPL is aninstantaneous sound pressure level rather than a time-weighted value, it shouldbereduced slightly to correlate with the EPNL; however, the amount of reductionislessthan4 dB. Consequently,themost significant noise sourceof thestudyairplane is that of the airframe.Itwas determined thatat a cutback distance of5944m (19 500ft), themaximum allowable altitudeis 515 m (1691 ft), because aircraft accelerationcapability isinadequatetoattain higher altitudes. Toreach this maximumaltitude, the required all-engine take-off field length is 2248 m (7375 ft).At cutback, thrustis reducedto provideaclimb gradientof 4percentin

accordance with regulations of reference 18. The take-off profile and twomeasurement points (ref.18) areshown infigure60. The lift-off velocityis287 km/hr (155 knots). At the point-1 measurement station, the velocity is321 km/hr (173 knots), the lift coefficient is 0.75, and the lift-drag ratiois 19.1.The variationsofEPNL along therunway center lineandalong theside-line (649 m (0.35 n. mi.) from the center line) are shown in figures 61 and 62,respectively. Figure61exhibitsanairframe OASPL of 94.8 dB at the 6486-m(3.5-n.mi.) center-line measurement point. Therefore, without engine cutback,engine EPNL would exceed theairframe noise. Contour plots forengine EPNL

values of 90 dB and 100 dB at engine cutback are presented in figure 63.Engine noise during approach.- During approach, the engines operate atidle thrustandhavea considerably lower noise level thantheairframe. Fig-ure 56 shows the 3 approach profile and the reference 18 measurement pointwhich islocated 1853m (1.0 n.mi.) fromthe 15.2-m (50.0-ft)thresholdpoint. The aircraft landing weight is 4 704 439 N (1 057 600 Ibf) and thelanding velocity is 300km/hr (162 knots). For a 3approach profile,thealtitudeat the1853-m(1.0-n.mi.) pointis 112 m (369 ft). These valueswere employed in the airframe noise calculations.

22


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Airframenoise.-The airframeOASPLof the study aircraft was computed byaveraging the values generated by the methods of references 21 and 22. Theinputvaluesand thecorresponding OASPLvalues from each methodat the tworunway center-line measurement points are presented in table IX. As shown, thevalues of airframeOASPLat the measurement pointsduring take-off and approachare94.78dB and104.77dB,respectively. However, these valuesmay be conser-vative since theyarebasedon thegeometric aspect ratioof4.53rather thanonthe effective aspect ratioofapproximately7.9.

DIRECT OPERATING COSTThevariationof thestudy airplane unitcostwiththenumberof airplanesproduced is shown in figure 64 based on1977dollars. These costs were calcu-lated by use of the method of reference 24.Figure 65 presents the variation of productivity with fleet size, a util-ization rate of 3500 hours per year, a load factor of0.65,and a range of5426 km (3200n.mi.)being assumed. According toreference25, the1976 air-freighttrafficfor theInternational Civil Aviation Organization airlines-excluding theSoviet UnionandChina - was approximately13.4billion revenueton miles. Although the cargo traffic handled by these airlines has increased atanaverage annual rateofabout15percent since1960,various surveys forecastafuturegrowth rate ranging from roughly 6 to 12 percent annually for the periodinto the1990's. If a 9-percent growth rate is assumed, traffic volume in1995would beapproximately 116 * 10^Mg-km. However, it isexpected that mostofthis cargo would continue to be carried by conventional freighters and within thebelly holds of passenger aircraft. Hence, it would appear that there might bea ma rke t for a fleet of rougly 100 span-distributed load aircraft.The effect of range on di rect operating cost(DOC)is shown in figure 66for several fleet sizes and load factors. These data were generated by usingthestandard Air Transport Association method of reference 26 with adjustmentsto reflect1977costs. For allcases shown,the effectsofrangeon DOC aresmall.Figure67showstheeffectsoffleet sizeandfuel priceon DOC forloadfactors of1.00and0.65. For the lower rangeoffleet size, which appearstomeet future market requirements,DOC isvery sensitiveto thenumberofaircraftproduced. Also,DOC ishighly dependentonfuel price, which isextremelyunpredictablefor theoperational period of thestudy airplane. A comparisonof the data with those of a current freighter of slightly less design rangeindicates that for the anticipated fleet size, the study airplane has lowerDOC,thedifference increasing as load factor decreases and fuel price increases.Since thestudy airplane utilizes scaled versionsof acurrently producedengine, it is expected that anticipated advancements in engine design wouldafford a further improvement in the DOC of the distributed load airplane.How-ever, itshould bekept inmine thatfor theoperational time period considered,advancements intechnology mayprovide an improvement in the DOC offuturecon-ventional freighter designs to the extent thattheirefficiency may be compar-able with thatof thedistributed-load airplane.

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Lastly, itshould also berealized that therequirementfor exceptionallywide runways may lead to a very small initial production rate, and thus greatlyincreases thefinancial burdenof theairframe manufacturer.

CONCLUSIONSA preliminary study has been conducted of a large span-distributed loadcargo aircraft capable of transportinga 2 668 933 N(600000Ibf)payloadofcontainerized cargo over intercontinental distances. The conclusions are asfollows:1. The specifications for payload weight, density, and dimensions inessence configure the wing, and establish unusually low values of wing loadingand aspect ratio.2. Thestructural weight comprises only about 18percentof thedesign

maximum gross weightandexemplifies themagnitudeofstructural efficiencyachievable through the utilization of the span-distributed loadingconcept.3.Although the geometric aspect ratiois4.53,it isestimated thatthewinglet effect of thewing-tip-mounted vertical tails increases theeffectiveaspect ratiotoapproximately7.9.4. A lift-drag ratioofnearly19 isattainedduringcruise.5.Trim drag incruiseisnegatedbycontrolling the centerofgravityby fuel management.6. Ahardened stability augmentation system (HSAS)isrequired. Controls-fixedlongitudinal dynamic analyses for the take-off and approach modes indicatethat utilization of a currentHSASimposes restrictions on the rearward center-of-gravity travel which preclude the use of optimum eleven settings. Therefore,efficient operation of the aircraft in these flight modes would require thedevelopment of afaster reacting control system.7. Sufficient control power to handle the large rolling moment of inertiadictates arelatively high minimum approach velocityof 315km/hr(170knots).8. Theairplane has acceptable spiral, Dutch roll,and roll-damping modes.9.Becauseof therelatively lowwing loading, engine size isdeterminedbycruise ceiling rather thanby take-off field length.

10. Themost significant noise source isthatof theairframe. However,for both take-offand approach the levels are below the limit(108dB) ofFederalAir Regulations, Part36.11. Thedesign mission fuel efficiency isapproximately 50percent greaterthan that of the most advanced, currently operational, large freighter aircraft.

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12. The direct operating cost is significantly lower than that of currentfreighterswiththeadvantage increasingasfuel price increases.

Langley Research CenterNational Aeronautics andSpace AdministrationHampton, VA23665February24, 1978

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REFERENCES1.Whitlow, David H.; andWhitener,P. C.: Technical and Economic Assessmentof Span-Distributed Loading Cargo AircraftConcepts. NASACR-144963, 1976.2.TechnicalandEconomic Assessment ofSpan-Loaded Cargo Aircraft Concepts.

NASACR-144962, 1976.3. Johnston/ WilliamM.;Muehlbauer, JohnC.;Eudaily,Roy R.;Farmer,Ben T.;Honrath, JohnF.; and Thompson, SterlingG.: Technical and EconomicAssessment of Span-Distributed Loading CargoAircraftConcepts. NASACR-145034, 1976.4. Whitehead, Allen H., Jr.: The Promise of Air Cargo - System Aspects andVehicle Design. NASATM X-71981, 1976.5.Bauer,F.;Garabedian,P.; Korn,D.; andJameson,A.: SupercriticalWing Sections II - A Handbook. Volume 108 of Lecture Notes in EconomicsandMathematical Systems, Springer-Verlag, 1975.6. Whitcomb, RichardT.: ADesign ApproachandSelected Wind-Tunnel Resultsat High Subsonic Speeds for Wing-Tip Mounted Winglets. NASA TND-8260,1976.7. McGhee, RobertJ.; andBeasley, WilliamD.: Low-Speed AerodynamicCharacteristics of a 17-Percent-Thick Airfoil Section Designed forGeneral Aviation Applications. NASATND-7428, 1973.8. Bruhn, E. F.: Analysis and Design of Airplane Structures. Tri-State Offset

Co. (Cincinnati, Ohio),c.1949. (Revised, Jan.1952.)9.Goldhammer, M. I.: A Lifting Surface Theory for the Analysis of NonplanarLiftingSystems. AIAA Paper No..76-16,Jan. 1976.

10.Woodward, F. A.: AnImproved Method for theAerodynamic Analysis of Wing-Body-Tail Configurations in SubsonicandSupersonicFlow. NASA CR-2228,pts. I-II, 1973.PartI -Theory andApplication.PartII -Computer Program Description.11. Putnam, Lawrence E.: AnAnalytical Studyof theEffectsofJets LocatedMore Than One Jet Diameter Above a Wing at Subsonic Speeds. NASATND-7754, 1974.12.Abbott, Ira H.; and Von Doenhoff, AlbertE.: Theoryof Wing Sections,Dover Publ. Inc.,c.1959.13. USAF StabilityandControlDatcom. ContractsAF 33(616)-6460and

F33615-74-C-3021,McDonnell Douglas Corp., Oct. 1960. (RevisedJan.1975.)

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14.Flying QualitiesofPiloted Airplanes. Mil. Specif.MIL-F-8785B(ASG),Aug. 7,1969.

15.JT9DCommercial Turbofan Engine Installation Handbook. Marketing Support.Pratt&Whitney Aircraft, United Technologies, Mar. 1967. (RevisedAug.1975.)

16.General Electric Co.: Propulsion System Studiesfor anAdvanced High Sub-sonic, Long RangeJetCommercial Transport Aircraft. NASACR-121016,1972.

17.Anderson,B. A.: ScalingtheJT9D Engine. TDM-1990Revised, Pratt&Whitney Aircraft, United AircraftCorp.,Apr. 1968.

18. Noise Standards: Aircraft Type and Airworthiness Certification. FederalAviationRegulations,pt. 36,FAA, June 1974.

19.Heidmann,M. F.: Interim Predictionfor Fan andCompressor Source Noise.NASATMX-71763,1975.20. Stone, James R.: Interim Prediction Methodfor JetNoise. NASATMX-71618,1974.21.Hardin,Jay C. ;Fratello, DavidJ.; Hayden, RichardE.; Kadman, Yoran;and

Africk,Steven: PredictionofAirframe Noise. NASATND-7821,1975.22.Baber,Hal T.; andSwanson,E. E.: Advanced SupersonicTechnologyConcept

AST-100Characteristics Developedin aBaseline-Update Study. NASATMX-72815,1976.

23.Advanced Supersonic Technology Concept Study Reference Characteristics.NASACR-132374,1973.

24. Oman,B. H.: Vehicle Design Evaluation Program. NASACR-145070,1977.25. Aerospace Industries Association of America,Inc.: Aerospace FactsandFigures,1975/76. Aviation Week&Space Technol.,[1975].26. Standard Method of Estimating Comparative Direct Operating Costs of TurbinePowered Transport Airplanes. AirTransp. Assoc.ofAmerica,Dec. 1976.

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TABLEI.-GROUP WEIGHT SUMMARY

Structures:Vertical tails

NacellesTotal/ structures.................Propulsion:

Fuel system, tanks, and plumbing

Systemsandequipment:

Hydraulics....................

Antiicing

WeightN

690 24433 94918727289 02346 813

1078 756

317 674454709 421

53 290425 855

747304 2709 946

39 76739 086100408 719890934

188 3821 692 993

4 00318 2384 337111

1 719 6822 668 9334 388 6151 663 6356 052 250

Ibf

155 1737 6324 21064 97510 524

242514

71 41610 2222 11811 98095 736

16 800960

2 2368 9408 7872 2571 960200210

42 350380 600

9004 10097525

386 600600 000986 600374000

1 360 60028


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TABLE IV.- LATERAL-DIRECTIONAL DYNAMIC CHARACTERISTICS

Inherent characteristics:MLWRLW

Stability modeDutch roll

min

0.080

0.2850.301

(CUn)min'rad/sec0.150

0.1680.171

wn,min'rad/sec0.400

0.5900.570

Rollfcm xsec1.400

1.3671.302

Spiralt2,min'sec

20.0

a-40.6a-42.3aNegative sign denotes timetohalf amplitudeof oscillation(spirally stable).

31


34/106

TABLEV.-ENGINE NACELLE DIMENSIONS

0

X

m0.000.025.127.254.508.7621.0161.5242.0322.769

in.015102030

406080109

rm

1.2671.3111.3671.4151.4781.5141.5391.5671.5801.585

in.49.951.653.855.758.259.660.661.762.262.4

X

m4.5215.0805.5886.0966.6047.1127.6208.1288.5098.738

in.178200220240260280300320335344

rm

1.5851.5771.5571.5191.4661.4051.3281.2241.1131.016

in.62.462.161.359.857.755.352.348.243.840.0

32


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TABLE VI.- MISSION PERFORMANCE[Taxi-infuel taken out of reserves at destination.Civil Aeronautics Board rangeequalstrip rangeminusallowancesfor maneu ver, traffic, andairway distance]

(a) Airc raft characteristicsTake-off gross weight, N(Ibf) 6 052 250 (1 360600)Operating weight, empty, N(Ibf) 1 719 682 (386 600)Payload,gross, N(Ibf) 2 668 933 (600 000)Wing area,m 2(ft2) 1 724 (18560)Sea-level static thrust , per engine, standardday:

Uninstalled, N(Ibf) 262 445 (59000)Installed, N(Ibf) 240 204 (54000)Take-off thrust-weight ratio 0.238Take-off wing loading,N/m2(lbf/ft2) 3509.62 (73.3)

(b) Design missionFlightmode

Take-off. .Start climbStart cruiseEnd cruise.End descent

Grossweight,N(Ibf)6 052 250 (1 360600)6 027 340 (1 355000)5 870 229 (1 319680)4 727 036 (1 062680)4 704 573 (1 057630)

AFuel,N(Ibf) ARange, ATime,km (n.mi.) min

24910 (5600) 0 11157 111 (35320) 370 (200) 31

1 143 193(257 000) 5213(2815) 38722 464 (5050) 370 (200) 20

Taxi-in 4 696 566 (1 055830) 8 007 (1800) 0 5Block fueland time 1 355 684(304 770) 454Tri p range 5953 (3215)

(c)Reserve fuel breakdown10-percent trip time,N(Ibf) 114 319 (25700)Missed approach, N(Ibf) 17 793 (4000)370 km(200n.mi.)to alternate airport, N(Ibf) 116 099 (26100)30 minutes holdingat 457 m(1500 ft),N(Ibf) 67 746 (15230)Total rese rve fuel 315 957 (71030)

(d) Initial cruise conditionsCL 0.3323CD 0.01782L/D 18.65TSFC,kg/N-hr (Ibm/lbf-hr) 0.0637 (0:625)Altitude, m(ft) 10 119 (33200)

(e) Fuel efficiencyPayload-distance per quantity of fuel burned,Mg-km/N (ton-n. mi./ Ibf) 1.19(3.16)

33


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TABLE VII.- INPUT PARAMETERS FOR PREDICTING TAKE-OFF FAN NOISE

Diameter, D, m(ft) ...................... 2.643(8.671)Fan total temperature rise, AC(AF) .............. 86.06(154.9)Mass flow, kg/sec(slugs/sec) ................. 763.35(52.31)Numberof fanblades .......................... 108Numberofstator vanes ......................... 46Rotor tip Mach numberatdesign .................... 1.287Rotor stator spacing ratio ....................... 1.267Fanrotor speed, u, rpm ........................ 2972Calculated values:Bladepassing frequency,FH= - , Hz .............. 5350Rotortipoperating Mach number,M^R= - ............ 1.212760c

TABLE VIII.- INPUT PARAMETERS FOR PREDICTING TAKE-OFF JET NOISE

Primaryexit flow characteristics:Area,m2(ft2) 0.8393(9.04)Mass flow, kg/sec (slugs/sec) . . 157.75(10.809)Velocity, m/sec (ft/sec) 369.65(1212.75)Density,3kg/m3 (slugs/ft3) 0.5102(0.00099)Absolute total temperature ratio,Tt,jet/Tt,a 2.66

Fanexit-flow characteristics:Area,m2(ft2) 2.3950(25.78)Mass flow, kg/sec (slugs/sec) 763.48 (52.315)Velocity, m/sec (ft/sec) 286.91(941.32)Density,3kg/m3(slugs/ft3) 1.1132(0.00216)Absolute total temperature ratio,Tt,jet/Tt,a 1.15

Aircraftvelocity, m/sec (ft/sec) .... 88.40(290.03)aDensity is computed byusing mass flow, velocity,and jetexitarea.

34


37/106

TABLEIX.-AIRFRAMENOISEDURINGTAKE-OFFAND APPROACH(AIRPLANEDIRECTLYOVERMEASUREMENTPOINT)

Take-off Approach(Measurement point 1) (Measurement point 2)Aircraftweight, W, N

Velocity, V,m/secWing area, S, m^Reference 21 airframeOASPL,dB(eq. (1))Reference 22airframeOASPL,dB(eq. (2))Study airplane airframeOASPL,3dB

6 052 25088.394.531542.9089.30

1724.2898.3891.0994.78

4 704 43988.394.531112.3683. 39

1 724.28109.24100.27104.77

aStudyairplane airframeOASPLassumed to be average of equations (1)and(2).

OASPLairframe= 10 log V3-34(W/9.807)-6b-63

h.83A3.03+56.14 (1)

=10V3.17(W/9.807)0.88h.62s0.16A2.06 +41.29 (2)

35


38/106

/_.06IPADC3.500DEG)

19.507/MAC(64.0QO).

60 52250 N(I3&0600 Ibf)DIHEDRAL .052 R AD(3.000D E G )

3.588(11.772)56.396"7185.0251"

Figure 1.- Span-distributed-load cargo airplane configuration.Dimensionsareshownin meters with feet in parenthesesexceptas noted.

36


39/106

cocoH-P< aoom0)z


40/106

Q.Q )OOOcti_34->O-,UC OXO

H

O

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OC--f- -rO6

Osrc \ j

:.vj \::"jujI

11

i

- i"~

i

cv.'yjC:.T,c;--3cc::(L-:DJcv.5c-vcCV.D

C

C DbOOC D

HTJcoC Do-pcooc

H.pO0>COCOnoLO

.= rCD

HCL,

39


42/106

1 0 0 0 x4000

8003000

600

2000400

200

I b f

1 0 0 0

k N

- 2 0 0 - 1 0 0 0

-400-2000

I M 9._... . . . _ LyJ. ._i_i_i.i .... .. . . ,. LJ ,.- , _ . . _ .-. _ _ 1 _ . __ . -. ., . .... .-. .. . .J* - - - - - - - - - - --j- - - ".. _IE . . ._ ..

::::::::::::::::::::::::::::::::::::::::: :::-7-f ;=lz , ... . II .... . .-

t _ . , h/ .,. -;

:::; ::::::::;?:::::::::::::::::::::: :::::::::::. 1 p, j .jj:>::::::::::::::::::::::::::::::::::::::::|-z:::mtttttttttttttt ^' tttttttttH-4^Wffl:i:i;-;:;ii;;;i:;;;::::;::;;i:;;;;i:|-|-;;,::::: : : : \ \ \ \ \ \ \ \ ^ \ \ l \ \ \ \ \ \ : \ \ \ Wing St

13X1 :\\::::\\:\::::l\::\:\\\ \:\^: \\ \:\

; 2.5 g Balanced maneuver-::s . . . . .

. _ _ . _ _ _ _, _ _. -~. .S r ? * - '-- ' 4;~*2


43/106

40 x 10

30

M i c .f t - l b f m - M N

T O

-10'

Figure 6.- Net limit swept axis bending moments.

41


44/106

20 x 10 C

20

10 ^m - M NM Y , e aft-lbf

10

-10

... 1 1 ... .

: M : : i 3000

2000

1000

0

1664 k N(374000Ibf

2669 k N(600000Ibf)

4 2 6 k N(95736Ibf)

1079 k N(242514Ibf)

2 1 5 k N

F u e l

P a y o a d

P r o p u l s i o n

S t r u c t u r e

S y s t e m s a n d

G r o s s w e i g h t ,p e r c e n t

(48349Ibf) o p e r a t i o n a l i t e m sFigure16.-Gross weight breakdown.

27.49

44.10

7.04

17.82

3.55

51


54/106

1600

1400

1200

1000

800

Aft dynamic limttake-offconfiguration

600

400

200

0

Aft dynamiclimt approach configurationitffOptimum Aft dynamic

S H c r u i s e cente Rlimit take-off> o f g r a v i t y- t r i m ) 3 c o n f i g u r a t i o nF o r w a r d c e n t e ro f g r a v i t yl i m i t d u e t oc o n t r o l p o w e rl i m i t a t i o n sd u r i n gt a k e - o f f

P r i m a r y m i s s i o n :

F e r r ym i s s i o nw i t h o u tb a l l a s t

= Z F W f o r t h e f e r r y m i s s i o nc = 19.507m (64.000 f t )

24.130m(79.167ft)

o >5

20 30 40C e nt e r o f g r a v i t y , p e r c e n t c

Figure1?.-Loadability envelope with flight limitations.(See table III for correlation of point number.)

52


55/106

o

oo

CO

PO0

oe v j

vo co

OOC M

... -aD.-H

H t,rH 0)C0)

0-HCu"

53


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u4 24-

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CM 73-P CDC M T3CO OOcd(DC M bOSbO.=r CCM -HC T3- cc dII rH10^r oo-, (0OrHO -O>,4JbO-HflS >-. AT3 t.

H OH LCM 0)O CI C U0) O 3HICT>

OCM CMO

54


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< u

oJ- OO CMII

O


58/106

oo

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COH3-,OC O0)

56


59/106

LD

24

20

16

1 2

; . 7 2

.78

.80

0 .2 .4 .6 .8

Figure 22.- Lift-drag variation with liftcoefficient for several cruise Mach numbers.S =1724m2 (18 560 ft2).

57


60/106

o t . d e g

Figure 23.- Effectofeleven deflectiononliftandpitching moment. 6f = 0; cruise.

58


61/106

bOCrHOHQ.OC

O>ao

COH4->o0)

OOC\JC M CO -H

0)

. i:ocoH4->H

C M

o0)C MUIinCM

H VO-, 00T3

Q JC 01 2O J3S -P

HCb

59


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T3a

SZ bOO Cc d C Do\- *~^O.CM0.43C d i-tb O Oco H int- inNT3 2c ina^ = ro

c\jbOC H 4Jco31H

60


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d e g

C e n t e r o f gravity, percent cFigure26.- Elevon requiredtotrim.

61


64/106

-"".** * ..*> i IUCXI

OI

HJOOJJ- >C D

S OH OH CM-U II

; ; ; ; ; ; ; ; ; ; ; ; ; ; s > ; ; ; ; ; ; ; 1 0 ; ; ; ; ; ; ; ; ; ; ; ; ; ; ; : : ;: : : : : : : : : : : : : : ? * ; : : : : : ; o o i | i : : : : : : : : : : : : :: ; ; ; ; ; ; ; ; ; ; ; ; ; 3 > ; ; ; ; ; ; " ; ; ; ; ; ; ; ; ; ; ; ; ; ; ; ; ;: : : : : : : : : : : : : : J J : : : : : : : O : : : : : ; : : : : : : : : : : :: : : : : : : : : ; : : : : 3 o . : : : : : : : : I J : ; : : : ; : : i i ; ; ; : ; ; ; ;: : : : : : : : : : : : : f _ - : : : : : :::|::: : : : : : : : : : : : : : :: ; ; ; ; ; ; ; ; ; ; ; ; ; ; o ; ; ; ; ; ; ; ; ; f ; ; ; ; ; = ; ; ; ; ; ; ; i i ; ; ;I i.I-H-H-Hm m-HM-H 4rW; ; ; ; ; ; ; ; ; ; ; ; : ^ ; : : : : : : ; : i ; : : ; ; : ; : : : ; ; : ; : = ; :MWB S fiH t f f i f f: : : : : : : : : : : : : 3 < J :::O ; j J i ; : : : : : : : :::::::::: : TTTTT ' ' * * * - 1 - : ' : : : : : : :. - - + I - - - F i*3hR.RRBHHBMBaakBH*.m m m m m m m*- H m m ^ i m m m m **m m m mmmmm*.- m m m m m m m* * * m m m m mmmmmmm* mmm mmm **mi m m m m m m m m m m m mmm mmm E ; ; ; ; ; i ; l t ; ; ; , ; i ; ; ; t t t ; i ; ; : : ; ; ; ; ; ; ; ; ; ; ; ; ; i ; i , ; ; ;

k.. m m m m m m t ' m u m m m m m m m m m m m m m f f m m m m m m m m m m m m m m m m a m taBKHHMBIBVBIk^BHkk.'a*I>M * *k. k1.'**.*-sba*MBaiaiaaBHBKBsi m m m m m m m m m m... *..*-a

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62


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66/106

s ec

A p p r o a c h , r e s e r v e f u e l w e i g h t

T a k e - o f f , m a x i m u m gr o s s w e i g h tA p p r o a c h , m a x i m u m g r o s s w e i g h t

M i n i m u m t i m e t o d o u b l e

3 1 3 2 3 3C en t er o f g r a v i t y , p e r c e n t c

Figure 29.- Estimated timetodouble amplitude forair-craftin approachand take-off configurations.

64


67/106


68/106

o0)

0)

66


69/106

60

50

40

R o l l a n g l e ,* , d eg 3 0

20

10

::::::::::::::::::: Level 3(4.0

, _..

, .

. . . . . . . 1 , , i ,, , . . . . . . , . , ,..- , , , . . . , . . _ . ,

, . . . . , , ,

, . . . ,.-_, ... ,

sec)ffi Leve l 2 (3.2 sec),. t .

i

iL:::: : : : : : : : : ::::::::::::: Level

...,I* .4...........E : : : : : : : : : : : ; i i : \ : : : ':'': : i i : : : : : : : : : R e c

..-.,..- .........

60 70 80 90 1 0 0 m / s e c

120 1 4 0 1 6 0 1 8 0E q u i v a l e n t a i r s p e e d

2 0 0 k n o t sFigure32.-Estimated aircraft roll responseto alateralstepinput. Landing approach. Approach configuration;cn.elastic/cn,rigid assumed equalto0.90;classIII;category C,reference14.

67


70/106

ITU:...........o' jI::::::::::::::::::::::: :;::::::::::::

6 0 E E E E i E E E ; E i E E E E E E E E i E E E E i i ; E E i E ; E i, I S

4 0 E E E E E E E E E i E E E E E E E E E i E E E i E E E E E E E E E E i E i E

2 0 E E E E ; E E E : ; E E E E E E E E E ; E ; E ; E E E E E E E E E E : E ; E Eo l i i i i i i i i i i i i i i i i i i i i i i i i i i i i i i i i i n H i i i

20 40 6

* , ,

i i : h i

0 80 100 I

40 80 120 160 20E q u i v a l e n t a i r s p e e d

m / s e c200 knots

Figure 33.- Estimated aircraft directional trim require-mentsin a90, 15.4 m/sec (30knots)cross wind.

68


71/106

01>ooa ta*3to / )aia.Po< uitt o

.6 .7 -8 .9C o r r e c t e d a i r f l o w r a t i o

1.0

Figure 31*.- Air-inlet total pressure recovery.

Sevcearbe

bmse

O

M

-P

o


72/106

i -oMO> a


73/106

50 xl O3

1 5 -

101-1 0 1 2 k m

1 0 2 0P r e s s u r e a l t i t u d e

3 0

Figure 37.- Installed maximum-climbnetengine thrust.Standardday to(standardday + 10 C).

40x10Jft

71


74/106

18x10"

16

14

I'2oO)

10

8

3

8 10 12 km

0 10_L

20 30 40 x 10 ftP r e s s u r e a l t i t u d e

F igu re 3 8 . - Installed maximum-climb f u e l f l o w . S t a n d a r d d a y .

72


75/106

4 5 x l 0 3

40

35

30

2501+J0)

20

15

10L

40 12 km

10 20P r e s s u r e a l t i t u d e

30 40 x 103ft

Figure 39.- Installed maximum-cruisenetengine thrust.Standard day to (standard day + 15 C).73


76/106

18x103 8

16

j _.cE

14

1 2

O 10

8 -

61-

i.V*\****T***

10 12 km

0 10 20 30 40 x KT ftP r e s s u r e a l t i t u d e

F igu re M O . - Installed maximum-cruise f u e l f l o w . Sta n d a r d d a y .

7 4


77/106

coCD

I ICD: O: O

II: in; i1

rffl " ff

_ > ^^^ffi * - i i i i i i i i i iCMO :

O ^ gII ^ fO [ [ [M l 1: II ^f 11I 1[ t 11I 1

' :

^ 'iillllllllll

; ^ I T I i i i l l l l l l l l l l l: Jlllllllii'" 1 -% * m & j :: -C : j g :

:: : 1 If) CM::: : 1r*. .;Hc^; ^E lill I

XCD

oCO< uo>

oOJ

0)Oo..pJ-n )o.Ort3 rtB T3HX "O3 (Tja-P a(0 (dgr-


78/106

3.0xl03

o

2

2 . 8 : : : : : : : : : : : : : :2.6 J4" J * je2 . 4 i i i : : : i i i i i : : J J j | : j j i j j j * f |STfff2 . 2 : i i i : : = i : : i i ; : = :

( a ) ; i ; ; ; ; ; ; ; ; ; ; ; ; ;? . n : ^ : : : : : : : : : ; ; ; : ; ; ; ; : : ; : : ; ; : : ; : ; ; ; : ; ; ; ;I ; h = 0 m (0 ft)

a . o x i o 3

S . O x l O 3

h =609.6m(2000ft)

2 . 8 ; ; ; ; ; ; ; ; ; ; ; : ; ; ; ; = ; ; = ; i i - i i i : ; :

2 . 2 ; ; ; ; ; ; ; j i i ; : : i i i : ; i i : : : : : ; ; : :2 . 0 ' llllilllllllllllllllllllllllilM60 100 140

h = 1219.2 m (4000 ft )180 220 260 kN

20 3 0 4 0 5 0Net e n g i n e t h r u s t 6 0 x 1 03 I b f

FigureU2.-Installed fan rotorspeedattake-off and part-powercruise thrust. Standardday + 10 C.76


79/106

4.5

a:a.co

< a

toa.

6.5

5.5

4.5

h = 609.6 m (2000 ft )

6 . 5

5.5

4.5

:M = 0 . 30 :

: : : : ; ; ; ; ; ; : i iM = 0 . 1 5 : : : ; ; ;;;:"=:: :: : : ; ; (r\::::::: : : : :

i m m\ \ i \ i i i \ \ \ i i i i i i i \ i i ; h =

: ; ; ; ; ; I I ; ;fflffl))JM-n.snlj:: : i i i i i iIIIIHIJIIIIIH: ffi

Illlllllllllllllllllllll1219.2 m (4C:::::::::::::::::::::::i i l i i i i i i i - j j i i i i M=015:

)00t)::::::::::::::::= 0 . 0 : : : :

6 0 1 00 1 4 0 1 80 2 2 0 2 60 k N

2 0 30 40N e t e n g i n e t h r u s t

50 6 0 x l 0 3 l b f

Figure43.-Installed engine bypass ratio at take-offand part-power cruise thrust. Standard day + 10 C.77


80/106

10x10*

8

u< u oM10= J03.c:x< uc


81/106

l .B rx lO 3

i.o.5

0o(U10

oootocr >iro.CX< U

1.5xl03

1.0

.5

h = 609.6 m (2000 ft) i i i

J 1.5xl03 g

1.0

.5

0

Q. .4

.2

0 i i i : ( c ) \ ' < i i i

; ; ; : : : ; i i i j M = 0.30 : ; ; : ; ; ; ; ; ; j i , i i j j i i i i l ; J i = 0.15; ;

i ; h = 1219.2 m (4000

.0

f t ) ;60 10 0 140 180 220 2 6 0 k N20 30 40

Net engine thrust50 60x103lbf

Figure 45.- Installed primary exhaust-gas velocity at take-off iand part-power cruise thrust. Standard d a y + 10 C.

79


82/106

1.2

1 . 1

1.0


83/106

O)


84/106

uO)

o030>

2

2 .0x l0 3

1.5

1.02.0xl03

1.5

< Ja >CT>

(0oI

ne

1.02. Ox 103

X< ucL_

'5lilllP'l'WIilmM. Illllll IllUmllllllUlllk

i n M i l l i H i ti 1 1 1 1 1 n m m 1 1 1 1 1 1 1 1 1 1 1 itiiii""* '" 1 1 n 1 1 IIIITIIHBi h =609.6m (2 U O U f t ) Bm i ini iiiiiii , , , , , , - t t t fH

1.5

1.010 0 140 180 22 0 2 6 0 k N

20 30 40N e t e n g i n e t h r u s t

5 0 60xl03lbf

Figure48.-Installed fan exhaust-gas flow at take-offandpart-power cruise thrust. Standard day + 10 C.82


85/106

4x102

o< D10

O

> O0>I(OXO)

rt)

(UCD

4x102 . -2

0 L

4x1

0L

OI/)C O

C OXa >

s _Q_

h = 609 .6 m (2000 ft)

h = 1219.2 m (4000 ft)220 260 kN

20 30 40N et engine thrust 50 6 0 x 1 0 Ib f

Figure49.-Installed primary exhaust-gas flow at take-offandpart-power cruise thrust. Standard day + 10 C.

83


86/106

1 2 X 10 3

c-' IT'CDCO)

O


87/106

402xTO3

398

394

a t 39001r*O)01c

386

382

378

374 L.

O 3 i i ; ; ; ; i i i ; ; i ; ; i ; i ; ; i i i ; ; i i i i i ; ; ; ; i i i

* 1720 ; ; ; ; ; ; ; ; ; ; ; ; ; ; ; ; ; ; ; ; ; ; ; ; : : ; : : ; ; ;

o 1700 i ^ = i i ^ ' ^

1 6 8 0 i i i i i i i i i i ; i i i ; i i ; i i i i i i i i ; i i i i i i i i i i i i i i i i

1660*^

:: : :: : : : : : : : : : : : : : ::

230 240 250 260 2 7 0 k N

50 52 54 56 58 60 x 103I n s t a l l e d s e a l e v e l , s t an da r d d a y t a k e - o f fn e t t h r u s t p e r e n g i n e

Figure 51.- Effect of engine sizeon operating weight empty.

85


88/106

1378 x 103

CD

in< />oeno< uat

1 3 7 4

1 3 7 0

1366

136 2

1 3 5 8

1 3 5 4

13501-

_6080

6120

6000220 2 7 0 k N

50 52 54 56 53 60 x 10J IbfI n s t a l l e d s e a l e v e l , s t a nd ar d d a y t a k e - o f fn e t t h r u s t p e r e n g i n e

Figure 52.- Effect of engine size on gross weight.

86


89/106

25

20COr-l_O

1536x10

34

.32


90/106

Il.xlo33.4&

C< UT3"atCJ)C

T a k e - o f f g r o s s w e i g h tD e s i g n m i s s i o nl a n d i n g w e i g h t

5300 5700 6100 6500k N

1 0 1 1 1 2L a n d i n g w e i g h t

1 3 1 4 x 1 0 Ib f

Figure54.-Landing performance. Noreverse thrust; spoilers;20flaps. Field length =1.66?x take-off roll distance.

88


91/106

0.6 x 106 2.8

2.4

2.0

1.6

ag 1.2ft

2000 4000 6000 8000 10000 12000 14000km

0 1000 2000 3000

R a n g e4000 5000 6000 7000

Figure 55.- Variationofpayload with range.

89


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OHCMO0)X(flO(0OfflOO.aa )

noH-Uc doo0)s1SM(0rt0)C QHO

VOina;t,g)H

C4

90


93/106

r

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S

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o v0 ^O 0)c < D ^2 30 3 1S o *>o m 30 . 0

j; on7? 1 3Ji 0) t-< u ^ 52TO-i-5O" j

O ?* iZ 5 ^r -H MH10 HO CO0

O < HO cO

cin?g

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gp aunssaud punos91


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1000

2000r 600r

1000

500

o4->00100

60L

4 0 0 ;

200:

-glOO,3280i5 6 0 i

40;

20;

10

1 - 1-

1=

-yi*

\

t

f

rf

f

-j

\

H

\

Hc

N

-/

i0

/4-

J

gr

/

*

H

h -rO)- 95o >to

90

8 5

' I -

" s

ijiiiiiiiiiiiipiiiiiiii

:::::-A1rframe O A S P L a tF A R 36 measurement point 1::::::ref. 185,(. . .

::::::: Engine E P N L atF A R 36 measurement point 1:::::::ref. 18

IIIIIIIBlB::::;] E n g i n e E P N L a tF A R 36 measurement point 2:::::::Tef.182 7 5 3 2 5 3 7 5 4 2 5 4 7 5 525 m

I J_9 0 0Figure 59.

1100 1300C u t b a c k a l t i t u d e

1 5 0 0Effect of cutback altitude on span-loadertake-off noise levels.

j1700 ft

93


96/106

> J H m- - $ j- - - \ - - - - - 1 . - -_

;;:;; % i;;;;;;;-;;' ;:;;;;;:;;:.Q '*> ;::::: 3::::::::::::::::: s , . - . :::::: cj:::::::::: :::::::::: ::::: :;:i j - U - + j 1 1 1 1 1 1 1 1 1 1 1 1 1 H I 1 1 I I 1 1 1 1 1 1 1 1 F l: ::: : to::::::::::::::::::::::::::;;;;; 2;;;;;;;;;;;;;;;;;-:;;;: *-144-:-c :::::::::: :::::::::::::::: J1- :;;:;;;;-;;;;;:;;-;;:S~:::;:::::::i::::::;;:::::::::::>S40 -* :::=::::::::::::::;::::::::::::a,Ml l iHll l l l l l l l l l l l nilmillIII *< u H-H-H-H4 5 " ;;;:;;;;;;;;:;;;;;;:;;;:;;:;;;mlllllllllllllllllllllllllllllllllH< oLU

a >cJ_+Jca u- >^8 B= 3 ^r- rfP " . oI I I ( 4 J 1 I I 1 i i d 1 llttHII 1 1 1 1 1 I I P I * ~ lJ_.. -.-_ ,.._, . ._ :LJ .|:J:::;;::::: J::::::i:|::::j

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97/106

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Figure 64.- Variation ofaircraft unit cost with numberofunits produced, 1977 dollars.

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10 0 20 0 300 40 0 500Fleet size

Figure65.-Variation ofproductivity with fleet size.Utilization, 3500hr/yr; load factor,0.65; range,5926 km (3200 n.mi.).

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2000 3000R a

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