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NASA Technical Memorandum 81872

NASA-TM-81872 19800024834

J CONCEPTUALSTUDYOFAN ADVANCEDSUPERSONICTECHNOLOGYTRANSPORT(AST-107)FORTRANSPACIFICRANGEUSINGLOW-BYPASS-RATIOTURBOFAN

ENGINES _z_,'-_ _ ,-=,._ _-_'_'_,'_'_._

ShelbyJ. Morris, Jr., Willard E. Foss,Jr.and Milton J. Neubauer,Jr.

September1980 ,_,_ q,_ q[)_y

OCT1 7 1980

'I-AD;GLEYIRESEAR.2HCENTEr,LIBRARY, NASA

NASA :"="National Aeronautics andSpace Administration

Langley Research CenterHampton, Virginia 23665

CONCEPTUALSTUDYOF AN ADVANCEDSUPERSONICTECHNOLOGYTRANSPORT(AST-I@7)FOR TRANSPACIFICRANGEUSINGLOW-BYPASS-RATIOTURBOFANENGINES

By

Shelby J. Morris,Jr.; WillardE. Foss, Jr.;and Milton J. Neubauer,Jr.

- NASA LangleyResearchCenterHampton,Virginia23665

SUMMARY

A new advancedsupersonictechnologyconfigurationconceptdesignatedthe AST-107, using a low bypass-ratio-turbofanengine,is describedand analyzed. Theaircraft had provisionsfor 273 passengersarrangedfive abreast. The cruise Machnumberwas 2.62. The missionrange for the AST-107was 8.48 Mm (4576 n.mi.)and an averagelift-dragratio of 9.15 during cruisewas achieved. The avail-able lateralcontrnlwas not sufficientfor the required15.4m/s (30 kt) cross-wind landingcondition,and a crosswindlandinggear or a significantreductionin dihedraleffectwould be necessaryto meet this requirement. The lowestcomputednoise levels,includinga mechanicalsuppressornoise reductionof 3EPNdB at the flyoverand sidelinemonitoringstations,were 110.3 EPNdB (side-line noise),113.1EPNdB (centerlinenoise) and 110.5 EPNdB (approachnoise).

INTRODUCTION

Technologyadvancessince 1972 have promptedseveraladvancedsupersonictech-nology vehicleintegrationstudies(for example,refs. 1-3). Subsequenttothese studies,a number of significantadvanceshave occurred. These advancesincludethe developmentof an expandedaerodynamig data base coveringlow-speedtrim, longitudinalstability,and control (ref. 4) and low-speedlateral-directionalstabilityand control(ref. 5); the developmentof techniquestosize the wing area and thrust-weightratio of supersoniccruiseaircraftforoptimumperformance(ref. 6); and the developmentof techniquesto computeairportnoise for these configurations(refs.7 and 8).

The objectiveof the presentstudy is to apply this new technologyto the con-ceptualdesignof an advancedsupersonictransportfor transpacificrange (com-parableto a San Francisco-Tokyoflight),and, then, to subjectthe concepttothe latest analyticaltechniquesfor performance,noise,and economicevalua-tions. In addition,this detailedsystemsintegrationstudy has been used tosupporta recentnoise-sensitivitystudy (ref. 9) by providingthe systemweightand performanceused therein.

The configurationdescribedin this report is very similarto that reportedinreference1 exceptfor differentengines. The engineused was a low-bypass-ratio turbofanincorporatingboth a mechanicalsuppressorin the outer exhaustjet and an invertedjet-velocityprofilefor coannularnoise reduction. The

I basic criteriaused in this study are: the use of five abreast seatingof 273passengersin an all-tourist-classseatingarrangementwith a seat pitch of0.864 m (34 in.); a range of 8.33 Mm (4500 n.mi.)at a cruiseMach number of2.62 (8C above standard-dayatmosphere);the abilityto land and takeoffonexistingrunwayswith a tire footprintno greaterthan that of a DC-8-50;sta-bility sufficientto trim for minimum trim-dragthroughoutthe flightenvelopewith no significantpitch-upin the takeoffor landingmodes; and satisfactoryshort-periodcharacteristicsat approach.

SYMBOLS

Computationsin the course of this study were performedin U.S. CustomaryUnits.Resultswere convertedto the InternationalSystem of Units (SI) by using con-version factorsgiven in reference10.

A aspect ratio

b wingspan,m (ft)m

c mean aerodynamicchord,m (ft)m

Cref referencemean aerodynamicchord,m (ft)

c.g. center of gravity

CD drag coefficient,dragqS

CDp° parasite drag coefficient associated with camber, protuberances,interference, and separated flow

CD wave-dragcoefficientw

CL lift coefficient,liftqS

C_ rollingmoment, rollingmomentqSb

C_B rollingmoment coefficientdue to sideslip,per degree

Cm pitchingmoment coefficient,pitchin9 momentqSc

Cn yawing momentcoefficient, yawin9 momentqSb

CnB yawing moment coefficient due to sideslip, per degree

CyB side force coefficient due to sideslip, per degree

DOC direct operating cost, cents/passenger km (cents/passenger statutemile)

EPNL effective perceived noise level, EPNdB

g acceleration due to gravity, m/s2 (ft/s 2)w

h altitude, m (ft)

HSAS hardened stability augmentation system

Ix,Iy,l z moments of inertia about X, Y, and Z body axes, respectively,

kg-m2 (slug-ft 2)

IXZ product of inertia, kg-m2 (slug-ft 2)

KI directional-control flexibility factor

Lm lift per unit angle of attack per unit momentum(qs/aircraft

momentum)_CL , per second

L/D lift-drag ratio

M Mach number

n load factor

P pressure, Pa (Ibf/ft 2)

Psp period of longitudinal short-period oscillation, s

Pd period of Dutch roll oscillation

PR pilot rating

Pph period of longitudinal phugoid oscillation

q dynamic pressure, Pa (Ibf/ft 2)

S wing area, m2 (ft 2)

Sref wing reference area, m2 (ft 2)

San static normal acceleration gust sensitivity, g/(m/s), (g/(ft/s)

SCAS stability and control augmentation system

t2 time to doubleamplitude,s

tl/2 time to damp to one-halfamplitude,s

t@=30o time requiredto roll 300, s

TOC total operatingcost, cents/seatkm (cents/seatstatutemile)

V airspeed,m/s (ft/s)

V1 aircraftvelocityat engine failurefor balancedfield length,m/s(ft!s)

q

V2 aircraft velocity at obstacle, m/s (ft/s)

VR aircraft velocity at rotation, m/s (ft/sec)

V tail-volume coefficient

x longitudinal coordinate, m (ft)

y lateral coordinate, m (ft)

z vertical coordinate, m (ft)

angle of attack (deg or rad, as noted)

_wrp angle of attack of wing reference plane, deg

B angle of sideslip, deg

APmax maximumsonic boomoverpressure, Pa (Ibf/ft 2)

_a aileron deflection, positive for right-roll command, deg

_i deflection of ith flap (see fig. 6 for definition of i) deg

6r rudder deflection, deg

_t horizontal-tail deflection, positive when leading-edge is deflectedup, deg

aa aileron deflection rate, positive for right-roll command, deg/s

ailerondeflectionrate, positivefor right-rollcommand,deg/s

6i deflectionrate of ith flap (see fig. 6 for definitionof i), deg/s

_r rudderdeflectionrate, deg/s

6t horizontaltail deflectionrate, positivewhen leading-edgeisdeflectedup, deg/s

Cd Dutch roll dampingratio

Cph longitudinalphugoidmode dampingratio

_sp longitudinalshort-periodmode dampingratio

_@ dampingratio of numeratorquadraticof a transferfunction

0 pitch angle, deg

e pitch rate, deg/s

pitch acceleration,rad/s2

@ roll angle,deg

@ roll rate, deg/s

rollingangularvelocitiesat the first and second peaks of a roll@I, @2 rate-oscillation,deg/s

roll acceleration,rad/s2

yaw angle, deg

yaw rate, deg/s

,°

_b yaw acceleration, rad/s 2

TR time constant of roll mode, s

md undampednatural frequency of Dutch roll mode, rad/s

5

msp longitudinalshort-periodundampednatural frequency,rad/s

_ph undampednaturalfrequencyof phugoidmode, rad/s

_@ undampednaturalfrequencyappearingin numeratorquadraticof@/_a transferfunction,rad/s

Subscripts:

app approach

c.g. center of gravity

elastic elasticairplane

f friction

LG landinggear

max maximum

min minimum

r roughness

rigid rigid airplane

req required

ss steady state

trim trimmedcondition

RESULTSAND DISCUSSION

CONFIGURATION

The presentAST-107 (fig. 1) is geometricallysimilarto the conceptof reference1. Geometriccharacteristicsare definedin Table I. The fuselagelength is92.96 m (305 ft) with provisionsfor 273 passengersarrangedfive abreastat0.86 m (34 in.) pitch. Passengerbaggageand cargo volume is providedunder thefloor forwardof the wing structuralbox. Figure2 shows the inboardprofileof the configuration. Fuel tanks are locatedin the wing, under the fuselagefloor in the center-wingcarry-throughstructure,and in the rear of the fuselage.The fuselagetank providesfor aircraftbalancecontroland fuel reserve. Thetank arrangementand capacitiesare shown in figure3.

The arrow-wingplanformis retainedfrom the configurationof reference1;however,the wing area was modifiedto obtain a wing loadingof 4.07 kPA(85 psf) at a design gross weight of 3.15 MN (709,000Ibm). The wing thickness-ratio variesfrom approximately3 percentat the root to 2.5 percentat the tip

6

(fig. 4). Contours of wing thickness are shown in figure 5. Wing control sur-faces (fig. 6) are geometrically similar to those of reference I except foradjustments for the difference in wing area.

The aircraft is powered by four General Electric GE21/JIO-B5 low bypass-ratioturbofan engines. The nacelle shape is based on a Boeing configuration sizedto the appropriate engine thrust level. The inlet used is a mixed compressionaxisymmetric design designated as the NASA"P" Inlet. The nacelles are mountedunder the wing at the trailing edge with the nozzles and thrust reversers suffi-ciently rearward to provide for clearance.

The main landing gear is a two-strut arrangement with 12 wheels per strut andretracts forward into the wing. Wheels are 0.80 m x 0.28 m size with the appro-priate ply rating to satisfy the equivalent single-wheel load and the 0.61 m(24 in.) flexible pavement thickness flotation criteria. The strut locationprovides for a 13o flare angle from the static ground line and a 14.5 o flareangle with the gear fully extended. These flare angles provide a maximumof0.30 m (12 in.) clearance for the engine nacelle, wing tip, and horizontal tailunder normal takeoff-and-landing conditions.

The nose landing gear is a single-strut, two-wheel arrangement and retractsrearward into the fuselage. Wheel size is 0.69 m x 0.19 m at the required plyrating to satisfy loading and floatation requirements.

Flight crew provisions and the visor nose are the same as those in reference i.With the visor nose in the downward position, sufficient clearance is availablefor a ground handling tractor.

The normal area distribution of the AST-I07 is presented in figure 7. Theutilization of aircraft volume for subsystem and accon_nodations is also shown inthis figure.

AERODYNAMICCHARACTERISTICS

Low Speed

The low speed aerodynamic characteristics are based on the wind tunnel data ofreference 4. Since the data of reference 4 are unaffected by Reynolds numberabRve that achieved at tunnel dynamic pressures of approximately 527 Pa (II Ibf/ftL), the data for these higher Reynolds number were used directly. The fric-tion drag from the wind tunnel tests was adjusted to full scale using the Somersand Short T' method (ref. ii) assuming flat-plate fully turbulent boundary-layerflow. Form drag corrections, which depend on thickness-chord ratio for surfacesand the fineness ratio for bodies, were applied. An additional three percentof friction drag was added for roughness, and 5 percent of friction drag wasadded for gaps and irregularities. Trim drag was obtained from reference 4 withappropriate corrections for tail-size and wing-reference-area differences. Flapincrements to lift and drag were obtained from reference 4. The landing-geardrag coefficient increments (ref. 12) are shown in figure 8. Ground effect wasestimated using the procedure of reference 2. The resulting lift curves anddrag polars are shown in figures 9 and I0. The corresponding lift-drag ratiosare shown in figure 11.

7

High Speed

Method of analysis- The procedureused to establishthe high-speeddrag valuesparallelsthe techniquesused in referencesI, 2, and 3. The common data baseused for each of those analyseswas also used for the presentconfiguration.The drag buildupconsistsof computingthe wave drag, skin frictionand rough-ness drag, and parasitedrag; each at zero lift, and then adding the induceddrag and the trim drag (for the appropriatelift coefficientand pitchingmoment) to this value.

Wave dra9 - Zero-liftwave drag coefficientsfor the AST-107were computedusing the supersonicarea-ruletechniqueof reference13. The Mach 2.62 equi-valent area distributionsdevelopedby the area rule for both the fuselageandcompleteconfigurationare shown in figure 12. Wave drag as a functionof Machnumber for the overallconfigurationis presentedin figure 13.

Skin frictionand roughnessdraB - Skin frictiondrag has been computedusingthe T' method describedin reference11. The frictiondrag for a given Machnumber-altitudecombinationwas computedby representingthe variousconfigura-tion componentsas appropriatewetted areas and referencelengths. Smooth flat-plate, adiabaticwall, fully turbulentboundarylayer conditionswere assumed.Configurationcomponentssuch as the wing or tail, which may exhibitsignificantvariationsin referencelengths,were furthersubdividedinto strips for a moreaccurate determinationof the frictiondrag. In additionto the skin frictiondrag, a separation-dragcomponentwas estimatedfor subsonicMach numbersusingan empiricalmethod (ref. 14). A subsonicform-dragfactor relatingthe skinfrictionon an airfoilor fusiformbody to the flat plate skin frictionwasdeterminedfor each configurationcomponentand used to obtain the subsonicform drag.

The configurationroughness-dragincrementwas assumedto be 6 percentof thefrictiondrag for the Mach 2.62 cruisecondition. For other Mach numbers,theratios of roughnessdrag to skin frictionpreviouslydevelopedin reference2were used.

A summaryof the configurationwetted areas is presentedis Table II. Skinfrictionand roughness-dragcoefficientsare shown in figure 14 as a functionofMach number. The altitudecorrespondingto each Mach number is defined by thenominalclimb scheduleof the aircraft.

Parasitedrag _ The parasitedrag coefficientwas derivedfrom the analysisofreference2 and is shown in figure 15. Parasitedrag includescamber drag,interferencedrag, excrescences,locallyseparatedflow, and other miscella-neous zero-liftdrag contributions.

Trim drag - The incrementaldrag-due-to-liftvalues for the AST-107 horizontaltail were obtainedby correctingthe values of referencei for differencesinwing and tail areas. These prior resultswere obtainedfrom a detailedanalysisof the horizontal-tailincidencerequiredfor maximumconfigurationperformance.

Total drag - The variousdrag componentsdiscussedabove were combinedtoobtain the overalldrag characteristicsof the AST-107. Figure 16 presentsthe

8

variationinminimumdragcoefficientacrosstheMachnumberrange. Figure17presentstypicalpolarsat Machnumbersof 0.6, 1.2,and 2.62. Figure18 pre-sents(L/D)maxas a functionof Machnumber.

PROPULSION

Engine

A GeneralElectricadvanced,augmented,low bypass-ratioturbofanengine,GE21/J10-B5,was selectedfor this study. This enginewas designedfor cruiseat a Mach number of 2.55 and an altitudeof 18.3 km (60,000ft) on a standardday. For the presentstudy,the enginewas sized to meet a sea-level-staticinstalledthrust-weightratio of 0.268 at temperatures8C above standard. Theresultingsea-level-staticmaximumairflowwas 298.4 kg (657.8Ibm). TheGeneralElectricGE21/J10-B5has a sea-level-staticdry thrustof 211 kN(47,504Ibf). The engine has a designoverallpressureratio of 16.3 and abypass ratio of 0.30. The technologyrequiredfor this engine is projectedtobe availablein 1985. The engine has a low-temperature(1,533K) augmentor.The exhaustsystemconsistsof an annular,translating,plug nozzlewith athrust reverserand a mechanicalsound suppressorinstalledin the outer streamof the nozzle. The mechanicalsound suppressorfurnishesan estimatedincre-mental noise suppressionof 3 EPNdB at maximumpower and is assumedto weigh2.22 kN (500 Ibf) per engine.

The engine weightwas determinedby scalingthe baseline-engineweight as the1.2 power of thrust. The length and diameterof the engine was assumedtoscale as the squareroot the thrust. The externalconfigurationand envelopeof the engine are shown in figure 19. The weight of the engine is 55.55 kN(11,590Ibf) includingthe engine,nozzle,thrust reverser,and mechanicalnoisesuppressor.

Nacelleand Inlet

The nacellehousingthe scaledGE21/J10-B5engine is shown in figure20. Thefixed downwardcant of the nozzle (8°) is providedso that the thrust line ofthe enginewill pass throughthe airplanerearmostcenter-of-gravityduringtakeoffand landingconditions. Locationof the nacelleson the aircraftisshown schematicallyin figure21.

The inlet is the NASA-Ames"P" inlet (ref. 15). It is a axisymmetric,mixedcompressiondesignwith a translatingcenterbodysized for supersoniccruiseconditions. Allowancewas made in the inlet design to provide2 percentof theinlet-systemairflowfor nacellecoolingand ventilation. Inlet performanceispresentedin figure22.

Performance

The installed engine performance is based on the 1962 U.S. standard atmosphereusing the inlet total-pressure recovery of reference 15 and a fuel lower heatinavalue of 42.8 MJ/kg (18,400 BTU/Ibm). An engine bleed of 0.906 kq/s (2 Ibm/s)per engine below 6.1 km i20,000 ft) and 0.459 kg/sec (I Ibm/s) per engine above6.1 km (20,000 ft) was used to account for required cooling air. A power ex-traction of 149 kW (200 HP) per engine was also included.

The installed engine performance presented in this report includes the effectof inlet pressure recovery, compressor bleed air, power extraction, the nozzlevelocity coefficient (with the mechanical noise suppressor), as well as boattaildrag from the engine nozzle-nacelle connection point rearward (stations D throughG of figure 20). At all engine operating conditions, engine performance hasalso been corrected for the effects of inlet spillage, bleed, and bypass drag.The nacelle skin friction, interference, and wave drags are accounted for in theaircraft drag polars. The resulting engine performance data for maximumclimb,maximumcruise, and part-power throttle setting is shown in figure 23 through25.

Mechanical Noise Suppressor

The mechanical noise suppressor used on the AST-I07 is a 40 shallow-chute outer-stream design which is assumed to weigh an additional 2.22 kN (500 Ibf) perengine and to supply 3 EPNdBof noise suppression at the sideline and fly-overnoise stations. The nozzle velocity coefficient was estimated to be 0.92 whenthe suppressor was deployed because of the flow losses associated with the sup-pressor.

MASSCHARACTERISTICS

Method of Analysis

Configuration selection and sizing of the AST-I07 was accomplished through theuse of the techniques described in reference 6 and in accordance with themission requirements specified in section entitled "Mission Analysis." Themass-properties prediction technique of reference 6 is based on data from pre-vious similar configurations and shows good correlation with data generated byairframe manufacturers.

The sizing and configuration selection was performed within the computer program "of reference 6 by producing a matrix of conceptual aircraft with an array ofdesign gross weights ranging from 3.05 to 3.67 MN (685,000 to 825,000 Ibf); withsea-level installed thrust-weight ratio varying from 0.20 to 0.44; and wingloading varying from 3.4 to 4.8 kPa (70 to I00 Ibf/ft2). These conceptual air-craft were then subjected to mission performance evaluations. The results ofthe evaluation were used to generate plots referred to as "thumbprints" (fig.26). The aircraft configuration selected for more detailed analysis has adesign gro_s weight of 3.15 MN (709,000 Ibf), a wing loading of 4.07 kPa(85 Ibf/ftL), and a sea-level-static thrust-weight ratio of 0.268.

I0

The aircraft primary structure was assumed to be all titanium. Special designfeatures included: stressed-skin titanium skin and core sandwich for all wingand other aerodynamic surfaces; skin, stringer, and frame fuselage constructionof titanium; and landing gear of high-strength steel-alloy construction.

Weight and Balance

After the design gross weight, wing loading, and aircraft thrust-weight ratiowere determined, an estimate of the weight and location of the different air-craft components was made using data from previous designs and the standardscaling techniques described in reference 6. The result of this analysis ispresented in Table III.

The aircraft components were located, their individual weights were estimated,and an analysis of the balance characteristics of the aircraft was made. Thedesired center-of-gravity limits for this aircraft, as determined in a latersection of this paper (Stability and Control), were a forward center-of-g_avitylimit of. 42.55 percent of T_^f_ and. an rearward, limit, of 60.01 percent of c_f.For minlmum trim drag, it was deslrable to malnta_n the center of gravity _about 50 percent Cref..

Combinations of different fuel usage and fuel-transfer sequencing were investi-gated to determine attainable center-of-gravity boundaries. These boundaries,together with the desired center-of-gravity trace during a typical mission, arepresented in figure 27. It was determined that, if the wing apex was located atfuselage station 15.24 m (50 ft), all points along the center-of-gravity pathwould lie within the required boundaries, provided that proper fuel managementwas utilized. It should be noted that the attainable rearward center-of-gravityexceeds the stability limit defined by 60.1 percent-c_f- For center-of-gravitylocations behind this point, the aircraft would not stable with the current tailsize.

Inertial Characteristics

Inertial characteristics were computed for maximumgross weight and for normallanding weight. The inertias of the individual components were computed abouttheir respective centroids and then translated to the aircraft's center-of-gravity location for both conditions. The inertia information is summarized inTable IV.

STABILITY ANDCONTROL

Requirements

Longitudinal - The requirements for longitudinal stability and control duringtakeoff were that the forward center-of-gravity limits should be set at theposition of neutral stability, and the center-of-gravity range during takeoffshould be at least 66 cm (26 in.). It was also required that the main landinqgear struts be located at least 6 percent of Cref behind the rear center-of-

Ii

limit. Furthermore,it must be possibleto maintain controlwithin the abovecenter-of-gravitylimitsat takeoff,both in and out of ground effect,with nosignificantpitch-up.

The requirementfor longitudinalstabilityand controlduring landingwas thatthe rear center-of-gravitylimit be set to providea nose-downpitchingaccel-erationof 0.08 rad/sL at minimumdemonstratedflight speed and normal landingweight. It was also requiredthat acceptabledynamicshort-periodcharacteris-tics be achievedat approachspeedswith an operationshardenedstabilityaug-mentationsystem,(HSAS),and that no significantpitch-upbe presentunderthese conditions.

Lateral-directional- For all cases, it was requiredthat a negativeroll resultfrom the occurrenceof positivesideslip(positivedihedraleffect). Duringtakeoff,controlpower sufficientto maintaindirectionalcontrolin a 15 m/s(30-knot)crosswindat 900 was required. Directionalcontrolshould also besufficientto counteractthe yaw caused by the loss of an outboardengine attakeoff. A 300 roll responsein 2.5 secondsafter initiationof a rapid full-controlinput at, or above, normal approachspeedswas also required. Controlpower must be adequateto trim the aircraftat _ = 0° in a 15 m/s (30-knots),900 crosswindusing not more than 75 percentof full lateralcontrol. Further-more, the aircraftshould possessinherentDutch-rollstabilitywith acceptablelevelsof undampednaturalfrequency(mH > 0.4 rad/sec) At supersoniccruise,the yawing moment due to sideslipmust 6e-greaterthan,"or equal to, zero for a2.5 g manuever.

Data Base

Low-speed, high-lift, longitudinal stability and control data were obtained fromreference 4. Transonic and supersonic longitudinal data were obtained fromreferences 16 and 17. Low-speed, high-lift, lateral-directional stability andcontrol data were taken from reference 5 and from recent unpublished wind-tunneltest results. Supersonic lateral-directional data were taken from reference 6.All data were corrected for tail-volume differences from the configuration ofreference 2.

Wing flexibilityassociatedwith lateralcontroldeflectionswas estimatedusingthe resultsof reference18. Fuselagetransversebendingdue to verticaltaildeflectionwas taken from reference19.

High-LiftDevicesand Controls

The high-liftdevices includetwo-segmentleading-edgeflaps at the wing apexand outboardKruegerflaps near the wing tips. The trailing-edgeflaps consistof an inner flap betweenthe fuselageand the inboardengine,a flap betweenthe inboardand outboardengine,and anotherflap betweenthe outboardengineand the aileronat the outermosttrailingedge. The high-liftdevicesandnomenclatureare given in figure6. The presenthigh-liftconfigurationisdefined by the two leadingedge inner flaps (labeled9 and 10) deflected300;

12

the two outer leading-edge Krueger flaps (labeled 13 and 14) deflected 45o; thetwo outer trailing edge flaps (labeled 5 and 6); and the ailerons (labeled 7 and8) deflected 5o. Longitudinal control is an all-movable horizontal tail with ageared elevator, and directional control is an all-movable vertical tail. TableV presents the mass and dimensional characteristics of the aircraft, and themaximumcontrol-surface deflections and the deflection rates assumed for thepresent analysis.

Control Power

Longitudinal - Longitudinal control power was established by the use of datafrom reference 3. Based on the assumed gearing between tail incidence andelevator deflection for the configuration of reference 3 (fig. 28), longitudinalcontrol capability was determined for the horizontal-tail-volume coefficientV of reference 4.

For takeoff and approach, the maximumlift-drag ratio was achieved with 20o oftrailing-edge flap deflection (fi_. 11), The rearmost center-of-gravity limitwas chosen to be 60.] percent of Cref, enabling the aircraft to achieve an angleof attack of 8.6 o at the 158 knot approach speed. The tail area was chosen tobe 83 m2 (893 ft 2) in order to achieve a nose-down pitch acceleration of 0.08rad/s 2 (ref. 2__0)at minimum approach speed. With-the landing gear located at66.1 percent Cref (6 percent behind 60.1 percent _ _), the minimum nose-wheellift-off speed is approximately 77 m/s (150 knots)_eTThis speed can be comparedwith the 103 m/s (200 knot) nose-wheel lift-off speed required for takeoff.

Lateral - Lateral control power was determined using the wind-tunnel data ofreference 5. Lateral-control flexibility factors for each surface (fig. 29)were established from the results of reference 18 for a stiffness-sized (flutter-free) wing design.

The steady-state sideslip, bank angle, rudder deflection, and lateral controlrequired for approaches with sideslip at an airspeed of 81 m/s (158 knots)are shown in figure 30. It can be seen that 75 percent of the availablelateral control is required for a crosswind component of approximately 12 m/s(22.5 knots) rather than the 15 m/s (30 knots) required by the criteria. Itwould require a reduction of about 20 percent in dihedral effect to achieve15 m/s (30 knot) crosswind capability.

Acceptable crosswind landing capability could be achieved by the use of a cross-wind landing gear or by a substantial reduction in dihedral effect. Geometricanhedral (ref. 21) would reduce positive dihedral effect. The models of refer-ences 4 and 5 were tested in the high-lift configuration with the wing shapeof the flexible wing at supersonic cruise. Reference 18 indicates that aero-elastic effects would result in comparatively greater anhedral at low speeds.The increased anhedral would alleviate the low-speed crosswind problem but mightnecessitate longer main landing-gear struts.

Directional - The piloted simulator study of reference 20 was conducted on aconfiguration similar to the present AST-I07. The vertical tail size was

13

chosen as 32.9 m2 (354 ft2) to providethe increaseddirectionalcontrolincrosswindsrecommendedby reference20. The directional-controlflexibilityfactor presentedin figure 31 is based on the data of reference19.

The directionaltrim requiredduring takeoffground-rollin a 900 crosswindisshown in figure 32. For a 12 m/s (22.5 knot) crosswind,directionalcontrolcannot be maintainedat ground speeds less than approximately44 m/s (85 knots).For lower speedsor highercrosswinds,nose-wheelsteeringor differentialthrustwould be required.

Static Stability

Longitudinal- The low-speedlongitudinalstatic-stabilityanalysiswas basedon wind-tunneltests (ref.4) of the configurationof reference2. Figures33 to 36 presentthe calculatedtrim and stabilityresultswithoutground effectfor trailing-edgeflap deflectionsof O, 10, 20, and 30 degrees. The horizon-tal tail was assumedto producemaximum lift-dragratio (tail upload)forclimb, accelerationto cruise,deceleration,and descentfrom cruise. Fromfigure 35, it can be seen that the trimmablecenter-of-gravityrange withtrailing-edgeflap deflectionsof 200 for flaps 1, 2, 3, and 4 (of fig. 6) isfrom 42.55 to 60.10 percentof Cref-

Supersoniclongitudinalstatic stabilitywas estimatedfrom the aerodynamiccenter data of reference3. The resultingflexible-airplaneaerodynamic-centerlocation is shown in figure 37.

Lateral-directional- The low-speedlateral-directionalstatic-stabilityanaly-ses were based on the wind-tunnelresultsof reference5. Supersonic-cruiselateral-directionalstatic-stabilitycharacteristics(estimatedfrom ref. 3)for the flexibleAST-107are presentedin figure38. Figure 38 shows thatthere is barely sufficientdirectionalstabilityto meet the criterionof

CnB _ 0 in a 2.5 g pull-upmaneuver.

DynamicStabilityand HandlingQualities

Extensivesimulatorstudies (ref. 12) have been conductedon the configurationof reference3 which was equippedwith a HardenedStabilityAugmentationSystem(HSAS)and a Stabilityand ControlAugmentationSystem (SCAS). Detailsof thesesystemsand the associatedautothrottleare given in reference12. Similarstudies,includingpilot ratings,have not been conductedon the presentcon-figuration;however,the stability,control,and handlingqualitiesmay beinferredby a comparsionwith the resultsof reference12.

The dynamicstabilitycharacteristicsof the AST-107are presentedin TableVl and the controlresponsecharacteristicsare given in Table VII. Thelocationof the AST-107with respectto the short-periodfrequencyrequirementsof reference22 is comparedwith the resultsof reference12 in figure 39(a).

A similarcomparisonusing the criteriaof reference23 is shown in figure 39(b).The low-speedpitch time-historyis comparedwith the requirementof reference

14

24 in figure 40. In all cases, the aircraft is indicated to be at least"acceptable" based on these criteria.

Roll control is compared with the criteria of reference 22 in figure 41(a)and with the criterion of reference 25 in figure 41(b). Provision of greatercontrol power would markedly improve the lateral handling qualities of theAST-I07.

AIRPORTNOISE

The jet-only noise characteristics of the AST-I07 were calculated for selectedtakeoff and approach profiles (using the methods of refs. 7 and 8) at the threemeasuring stations prescribed in reference 26. The relative location of thesestations on the airport runway is shown in figure 42. The flight trajectoriesused for the noise calculations were required to meet the balanced field-lengthand climb-gradient requirements of the federal safety regulations.

Computed jet-only noise for the trajectories described as case numbers I through3 of Table VIII and figure 43 are summarized in Table IX. Case 1 did not useany power cutback and the normal takeoff-power setting was used throughout thetakeoff and climbout. Case 2 used the normal takeoff-power setting in takeoffand climbout until the aircraft reached 5.94 km (19,500 ft) from brake release,at which point the throttle was cut back to 62.6 percent of normal takeoff power.In case 3, the takeoff power was set such that the takeoff Federal-Aviation-Regulation field-length was 3.81 km (12,500 ft). This throttle setting (approxi-mately 93 percent of normal) was used throughout the takeoff and until the air-craft reached an altitude of 213 m (700 ft) (which occurred at 6.4 km (20,995 ft)from brake release). At that point, the power setting was further reduced toabout 65.2 percent for the remainder of the climbout.

The takeoff of case 1 (Table IX) resulted in a maximumsideline noise level of113.8 EPNdBand a centerline noise level of 119.4 EPNdB. The takeoff of case 2resulted in a maximumsideline noise level of 113.3 EPNdBand a centerline noiseof 116.1EPNdB. The takeoff of case 3 resulted in a maximumsideline noise of113.4 EPNdBand a centerline noise of 120.7 EPNdB. All of these levels shouldbe reduced by an estimated 3 EPNdBto account for the mechanical jet-noise noisesuppressor. Of the trajectories considered, case 2 produced the least noise,which, including the increment fer the mechanical suppressor, resulted in amaximumsideline noise of 110.3 EPNdBand a centerline noise of 113.1EPNdB.The mechanical-suppressor noise increment is only an estimate and should beconsidered with caution. Approach noise, using a standard 3o glide slope atan approach speed of 158 knots was 110.5 EPNdB.

Jet-noise-only contours for case I and case 2 are shown in figures 44 and 45.The results of the jet-noise-only calculations for the three cases are summarizedin figure 46.

15

SONICBOOM

Equivalent area distributions due to volume and lift which are required forthe sonic boom analysis were computed by the use of method described inreference 13. These methods were modified prior to their use in references Iand 3, and in the present study, to include the effect of angle of attack. Theresulting equivalent areas and the procedures of references 27 and 28 wereemployed to define near-field pressure signatures below the aircraft for threedifferent fuselage lengths. These near-field pressure signatures were extra-polated to ground level using the methods described in reference 29 whichinclude the effects of variations in atmospheric properties, of aircraft accel-eration, and the variation in aircraft flight path angle. A reflection factorof 1.9 was used through the analysis.

Sonic boomsignatures were computed for a series of Mach numbers from transonicclimb through supersonic cruise. The maximumoverpressure (Ap_v) was then com-

puted as a function of Mach number (fig. 47) fo_ the aircraft _jectory. Themaximumsonic-boom overpressure (APmax) is abou. 130 Pa (2.7 Ibf/ft2). Basedon figure 47 and information from previous analyses, a focused boom (ref° 30)could occur on the ground as a result of the acceleration during the transonicclimb phase of the flight (about M = I°I0). This focused boom effect can bepositioned well offshore (about 160 km (88 n.mi.) from brake release) of anyland mass for many of the airports to be used by a supersonic transport.

MISSIONANALYSIS

Mission Requirements

The design objective was to optimize a supersonic transport aircraft concept forthe minimum weight required to achieve the mission goals. Mission analysis wasrequired to establish the weight of the aircraft required to meet these goals.

The objective of the study was an aircraft concept which could cruise at a Machnumber of 2.62 (at standard day plus 8C temperature) for 8.33 Mm(4,500 n.mi.)carrying a payload of 273 passengers and their baggage (a total payload weightof 253.8 kN (57,057 Ibf). The balanced field length for takeoff was not toexceed 3.81 km (12,500 ft) (at standard day plus IOC temperature). The approachspeed was not to exceed 81.3 m (158 knots).

The fuel requirement included that required to complete the base mission plusreserve fuel sufficient to overcome the effect of headwinds (5 percent of tripfuel); the fuel required for 30 minutes in a holding pattern at 3.05 km (I0,000ft); the fuel required for one missed approach and the necessary "go-around";and the fuel required for cruise to an alternate airport 463 km (250 n.mi.)away (using the best cruise Mach number and altitude). The fuel reservesdescribed above are based on the recommendations of reference 31.

16

Method of Analysis

The methods of reference 6 were used to establish the configuration of theaircraft and to determine its size. An array of candidate aircraft was studiedby varying the thrust-weight ratio and wing loading. This resulted in a matrixof aircraft of different gross weights. The takeoff field length, approach speed,fuel volume, and climb acceleration margin were included as constraints on thedesigns. Data generated for this matrix of related aircraft were plotted toestablish the "thumbprint" sizing plot shown on figure 26.

Since the study objective was to determine the minimum weight of the aircraftrequired to fly a given payload at fixed range with the above constraints, thepresentation of figure 26 is in the form of constant takeoff-gross-weight con-tours as a function of thrust-weight ratio and wing loading. From this thumb-print plot, a candidate aircraft having the minimum gross weight within therequired constraints was selected. This aircraft had an estimated takeoff grossweight of 3.15 MN (709,000 Ibf) with an installed thrust-to-weight ratio of 0.268and a design wing loading of 4.07 kPa (85 Ibf/ftL).

This configuration was then subjected to a more detailed analysis to determinepertinent performance characteristics and to evaluate its compliance with missiongoals. The conceptual aircraft was "flown" within the computer program (ref. 6)in accordance with the selected mission profile. For each segment of the profile,the program determined enroute details such as thrust and fuel required, altitude,speed, and the end point time of each segments. The profile used in this study(fig. 48) was composed of the following segments: a lO-minute warmup and taxi-out fuel allowance; a takeoff of one minute at full takeoff thrust a climb andacceleration using the climb schedule shown in figure 49, where the flaps aredeflected 30o below 213 m (700 ft.) altitude and are retracted above 213 m(700 ft.) altitude; a cruise at either optimum Brequet range factor or at climbceiling; a descent in accordance with figure 49,

Design Mission Performance

Results of the mission performance evaluation are summarized in Table X.Cruise lift-drag ratio, an indicator of aerodynamic efficiency, averaged 9.15.Additional evaluations were performed for reduced payload conditions to producethe payload-range characteristics shown in figure 50.

Off-Design Operation

The aircraft was also investigated while performing subsonic off-design missions.The trade between the length of the subsonic (M = 0.9) and length of the super-sonic (M = 2.62) flight segments for passenger load factors of 60 percent andi00 percent is shown in figure 51. Figure 51 indicates that at 60 percent loadfactor the all-subsonic range is about 7.60 Mm(4200 n.mi.) and the all-supersonic range is about 9.50 Mm(5200 n.mi.).

17

Sensitivity Analysis

The sensitivity of the design range to changes in total drag coefficient, enginespecific fuel consumption, and structural weight was also investigated. Theresults of this sensitivity analysis are presented in figure 25.

ECONOMICEVALUATION

An airline, as a potential purchaser and user of transport aircraft, mustconsider the profit-making potential of any aircraft which it might introduceinto its fleet. Parameters important in determining the productivity, and thus,in effect, the profit-making capability of any aircraft, include the direct andtotal operating costs of the aircraft. The fuel burned per mission also isimportant because of the current difficulty in predicting the future cost of fuel.Block fuel and seat miles per gallon are used to evaluate the fuel efficiency ofthe aircraft.

The direct operating cost was computed as described in the Air TransportAssociation model of reference 32. The indirect operating cost and the totaloperating cost (sum of direct operating cost and indirect operating cost) wascomputed as described in reference 33. Monetary values were in terms of con-stant 1976 dollars. The ticket price was computed with a return-on-investmentof 15 percent. Analyses were performed for both 60 and i00 percent passengerload factors at subsonic (M = 0.90) and supersonic (M = 2.62) cruise for variousranges. At 1976 dollars, the AST-I07 is economically viable producing 15 per-cent return-on-investment with 60 percent load factor and an average ticketprice of approximately 300 dollars per passenger at a range comparable to aflight from San Francisco to Tokyo or 200 dollars per passenger for a New Yorkto London flight.

The direct operating cost, total operating cost and ticket price are presentedin figures 53 to 55. From the comparison of ticket prices, it is apparent thatoperation of the AST-I07 in the subsonic cruise mode would impose a significantburden on the fare-paying passenger. The fuel cost used in the foregoing cal-culations was 10.25 cents per liter (38.7 cents per gallon). The block fueland passenger statute miles per gallon are shown in figures 56 and 57. The AST-107 obtains about 14.5 passenger kilometers per liter (34.3 passenger statute-miles per gallon) at reduced ranges (fig. 57). :The effect of varying the fuelcost and the hours of utilization per year on the total operating cost of thev6hicle is shown in figures 58 and 59. The fuel cost and hours of utilizationsensitivities shown in figures 58 and 59 are for cruise at a Mach number of 2.62and a range of between 4500 and 4600 nautical miles.

CONCLUSIONS

A new advanced supersonic technology configuration designated the AST-107 usinga low bypass-ratio turbofan has been described and analyzed. The results ofthis study indicate that:

18

1. Mission range for the AST-107 was 8.48 Mm(4576 n.mi.). The averagelift-drag ratio during cruise was 9.15.

2. The available lateral control is not sufficient for the 15.4 m/sec(30 kt) crosswind requirement. A crosswind landing gear or a significantreduction in dihedral effect would be required to meetthis criteria.

3. Centerline and maximumsideline noise (jet only) were computed to be110.3 EPNdB(sideline noise) and 113.1EPNdB (centerline noise) including a3 EPNdBbenefit for a mechanical noise suppressor. The approach noise wasestimated to be 110.5 EPNdB.

4. Off-design operation of the AST-I07 in the subsonic cruise mode wouldnot be economically attractive because of the poor subsonic specific fuel con-sumption of this particular engine.

5. The AST-I07 obtains as much as 14,5 seat-kilometers per liter (34,3seat-miles per gallon) at reduced ranges.

6. At 1976 dollars, the AST-I07 is economically viable producing 15percent return on investments with 60 percent load factor and an average ticketprice of approximately 300 dollars per passenger at a range comparable to aflight from San Francisco to Tokoyo or 200 dollars per passenger for a New Yorkto London flight.

19

REFERENCES

1. Baber, Hal T., Jr.: Characteristics of the Advanced Supersonic TechnologyAST-I05-1 Configured for Transpacific Range With Pratt and WhitneyAircraft Variable Stream Control Engines. NASATM_78818, March 1979.

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

3. Baber, Hal T., Jr.; and Swanson, E. E.: Advanced Supersonic TechnologyConcept AST-IO0 Characteristics Developed in a Baseline - Update Study.NASATM X-72815, 1976.

4. Smith, Paul M.: Low-Speed Aerodynamic Characteristics From Wind-TunnelTests of a Large-Scale Advanced Arrow-Wing Supersonic-Cruise TransportConcept. NASACR-14580, 1980.

5. Coe, Paul L., Jr.; Smith, Paul M.; and Parlett, Lysle P.: Low-Speed WindTunnel Investigation of an Advanced Supersonic Cruise Arrow-WingConfiguration. NASATM-74043, 1977.

6. Fetterman, David E., Jr.: Preliminary Sizing and Performance Evaluationof Supersonic Cruise Aircraft. NASATM X-73936, 1976.

7. Foss, Willard E., Jr.: A Computer Program for Detail Analysis of theTakeoff and Approach Performance Capabilities of Transport CategoryAircraft. NASATM-8012, June 1979.

8. Raney, John P.: Noise Prediction Technology for CTOLAircraft.NASATM-78700, 1978.

9. Staff of the Langley Research Center: Preliminary Noise Tradeoff Studyof a Mach 2.7 Cruise Aircraft. NASATM-78732, April 1979.

10. Mechtly, E.A.: The INternational System of Units - Physical Constants andConversion Factors (Second Revision) NASASP-7012, 1973.

11. Sommer, Simon C.; and Short, Barbara, J.: Free Flight Measurements ofTurbulent Boundary Layer Skin Friction in the Presence of SevereAerodynamic Heating at Mach Numbers From 2.8 to 7.0. NACATN 3391,1955.

12. Grantham, William D.; Nguyen, Luat T.; Deal, Perry L.; Neubauer, Milton, J.,Jr.; Smith, Paul M.; and Gregory, Frderick D.: Ground-Based and In-FlightSimulator Studies of Low-Speed Handling Characteristics of Two SupersonicCruise Transport Concepts. NASATP-1240, July 1978.

13. Harris, Roy V., Jr.: An Analysis Correlation of Aircraft Wave Drag. NASATM X-947, 1964.

2O

14. Hoerner, S. F.: Fluid-Dynamic Drag. Hoerner Fluid Dynamics (Bricktown,N.J.), c. 1965.

15. Koncsek, J.L.; and Syberg, J.: Transonic and Supersonic Test of a Mach 2.65Mixed Compression Axisymmetric Intake. NASACR-1977, March 1972.

16. Decker, John P.; and jacobs, Peter F.: Stability and PerformanceCharacteristics of a Fixed Arrow-Wing Supersonic Transport Configuration(SCAT 15F-9898) at Mach Numbers From 0.60 to 1.20. NASATM-78726, June1978.

17. Morris, Odell A.; Fuller, Dennis E.; and Watson, Carolyn B.: TheAerodynamic Characteristics of a Fixed Arrow-Wing Supersonic CruiseAircraft Configuration at Mach Numbers of 2.30, 2.70, and 2.95. NASATM-78706, June 1978.

18. Wrenn, G. A.; McCullers, L. A.; and Newsom,R., Jr.: Structural andAeroelastic Studies of a Supersonic Arrow Wing Configuration. NASACR-145325, 1977.

19. Douglas Aircraft Company: Studies of the Impact of Advanced TechnologiesApplied to Supersonic Transport Aircraft. MDCJ-4394, 1973.

20. Grantham, William D.; Nguyen, Luat T.; Neubauer, M. J., Jr.; and Smith,Paul M.: Simulator Study of the Low-Speed Handling Qualities of aSupersonic Cruise Arrow-Wing Transport Configuration During Approach andLanding. NASACP-O01, Part 1, 1976, pp. 215-248.

21. Lockwood, Vernard E.: Effect of Leading-edge Contour and Vertical-TailConfiguration on the Low-Speed Stability Characteristics of a SupersonicTransportModel Having a HighlySwept Arrow-Wing. NASA TM-78683,1978.

22. Anon: FlyingQualitiesof PilotedAirplanes. MilitarySpecificationNIL-F-8785B(ASG),1969.

23. Shomber,H. A.; and Gertsen,W.M.; LongitudinalHandlingQualitiesCriteria: An Evaluation. Journalof Aircraft,Vol. 4. No. 4, July-August 1967.

24. Suddreth,RobertW.; Bohn, Jeff G.; Caniff,MartinA.; and BennettGregoryR.: Developmentof LongitudinalHandlingQualitiesand CriteriaFor Large AdvancedSupersonicAircraft. NASA CR-137635,1975.

25. Anon: AerospaceRecommendedPractice- DesignObjectivesfor FlyingQualitiesof Civil TransportAircraft. ARAP 842, SAR, 1964.

26. DOT/FAANoise Standards: AircraftType and AirworthinessCertification.FAR Part 36, June 1074.

27. Carlson,Harry W.: Correlationof Sonic Boom Theory With Wind Tunnel andFlightMeasurements. NASA TRR-213,1964.

21

28. Middleton, Wilbur D.; and Carlson, Harry W.: A Numerical Method forCalculating Near-Field Sonic BoomPressure Signatures. NASATN D-3982,1965.

29. Thomas, Charles L.: Extrapolation of Sonic BoomPressure Signatures bythe Wave Form Parameter Method. NASATN D-6832, 1972.

30. Haglund, George T.; and Kane, Edward J.: Effect of SST OperationalManeuvers on Sonic Boom. Journal of Aircraft. Vol. 9, No. 8, August1972, pp. 563-568.

31. Anon.: An Airline's View of Reserve Fuel Requirements for the SupersonicTransport. Leckheed-California LR 26133, 1973.

32. Anon.: Standard Method of Estimating Comparative Direct Operating Costs ofTurbine Powered Transport Airplanes. Air Transport Assoc. of America,1967.

33. Anon.: Indirect Operating Expense Coefficients, Years 1963 Through 1979.Lockheed-California CompanyAA/2917, June 1980.

22

TABLE 1.- AST 107 GEOMETRIC CHARACTERISTICS

Geometry Wing Horizontal Vertical Wing fintail tai 1

m2 853.96 82.96 32.84 18.12Geometric area

(ft2) (9192 ) (893) (354) (195)

Mean Aerodynamic m 31.19 8.97 8.90 7.19chord t c (ft) (102.32 ) (29.43) (29.20) (23.58)

Wing reference area rtf 774.16

sref (ft2) (8333)

Reference mean aero- m 26.82dynamic chord t cref (ft) (87.99)

Exposed area rTf 57.60 32.89 18.12(ft2) (620 ) (354) (195)

Span t b m 38.39 10.54 4.16 2.99(ft) (125.97) (34.57) (13.66) (9.83)

Aspect ratio based on 1.721 1.338 .527 .495above geometric area

Aspect ratio based on 1.904wing reference area

Leading edge sweep Deg 74 70.84 60 55 68.2 73.42

Root chord m 51.02 12.96 12.77 10.64(ft) (167.38) (42.53 ) (41.90) (34.91)

Tip chord m 4.92 2.79 3.03 1.45(ft) (16.13) (9.14) (9.93) (4.75)

Root thickness/chord % 3.000 2.996 2.996

Tip thickness/chord % 3.000 2.996 2.996

Taper ratio .215 .237 .136

Dihedral Deg -15-

Vol. coeff. (gross)t V * .113 .041 .012 (each)

Vol. coeff. (ref. L ij ** .146 .046 .014 (each)

*Based on gross wing characteristics**Based on reference wing characteristics 23

24·

TABLE 11.- WETTED AREA SUMMARY

Component Wetted aream2 (ft2)

Wing 1400.71 (15077.09)

Fuselage 733.63 (7896.77)

Nacell es (4) 204.53 (2201.52)

Wing fins (2) 72.78 (783.42)

Vertical tail 66.11 (711.65 )

Horizontal tail 119.73 (1288.72)

Total 2597.49 (27959.15)

TABLEIII.- GROUPWEIGHTSUMMARY

kN I bf

Wing 336.850 75727Horizontal tail 29.394 6608

. Vertical tail 17.962 4038Vertical fin 9.461 2127Canard 0.0 0Fuselage 226.935 51017Landing gear 122.268 27487Nacelle 47.200 10611

Structural total (790.070) (177615)

Engines 206.220 46360Thrust reversers 0.0 0Miscellaneous systems 7.918 1780Fuel system - tanks and plumbing 32.681 7347

- insulation 0.0 0Propulsion total (246.819) (55487)

Surface controls 41.311 9287Auxiliary power 0.0 0Instruments 7.811 1756Hydraulics 25.898 5822Electrical 22.321 5018Avionics 12.273 2759Furnishings and equipment 88.048 19794Air conditioning 38.553 8667Anti-icing 0.934 210

Systems and equipment total (237.149) (53313)

Weight empty 1274.038 286415

Crew and baggage - flight, 3 3.003 675- cabin, 9 6.606 1485

Unusable fuel 8.972 2017Engine oil 2.180 490Passenger service 36.760 8264Cargo containers, 6 11.921 2680

Operating weight 1343.480 302026

Passengers, 273 200.370 45045Passenger baggage 53.432 12012Cargo 0.0 0

Zero fuel weight 1597.282 359083

Mission fuel 1556.508 349917

Takeoff gross weight 3153.790 709000

25

TABLE IV.- INERTIA SUMMARY

Condition

Weight, kN, lbf

Horizontal e.g., m, in

Percent of Cref

Roll inertia, Ix, Gg_m2, sl u9-ft2

Pitch inertia, Iy , Gg-m2, sl ug-ft2

Yaw inertia, Iz, Gg_m2 2, slug-ft

Product of inertia, Ixz , Gg_m2 2, 51 ug-ft

Principle axis angle of inclination, rad, deg

26

Takeoff Normalgross wei ght landing weight

3153.790 1827.153709000 410500

52.923 52.6452083.58 2072.65

60.00 58.97

11.089x 106 6.081

8.179 4.485 x 106,

77 .904 73.80857.459 x 106 54.438 x 106

86.565 77.72163.847 x 106 57.423 x 106

-2.562 6 -2.259 6-1.890 x 10 -1. 666 x 10

-0.0338 -0.0316-1.94 -1.81

TABLE V.- MASS AND DIMENSIONAL CHARACTERISTICS OF THE AST-I07

Reference wing area, m2 (ft2) ••••••••••••••••••••••Wing span, m (ft) .Wing leading-edge sweep, deg (see fig. 1) •••••••••••Reference mean aerodynamic chord, m (ft) ••••••••••••Center-of-gravity location, percent Cref ••••••••••••Static margin, percent

774.16 (8333)38.39 (125.964)

74.00/70.84/60.0026.82 (87.985)

60.10-3.7

Takeoff weight, mN (lbf)Ix, kg-m2 (slug-ft2)Iy, kg-m2 (slug-ft2)Iz, kg-m2 (slug-ft2)Ixz , kg-m2 (slug-ft2)

Landing weight, mN (lbf)Ix, kg-m2 (slug-ft2 )Iy, kg-m2 (slug-ft2)Iz, kg-m2 (SlU9-ft2~Ixz , kg-m2 (slug-ft )

· .·.................... ..·.................... ..·.................... ..·.................... ..·.................... ..· ~ ..·.................... ..·.................... ..·.................... ..

3.154 (709,000)11,089,230 (8,179,000)

77,903,940 (57,459,000)86,564,910 (63,847,000)-2,562,500 (-1,890,000)

1.827 (410,500)6,080,840 (4,485,000)

73,808,020 (54,438,000)77,720,910 (57,324,000)-2,258,790 (-1,666,000)

Maximum control surface deflections in degrees:Horizontal tail (6t ) .Flap (61 and 62) •..•.•••.••..••..••.•.•...•••......•••...Aileron (07 and 08) •••••••••••••••••••••• ;•••••••••••••••••••Outboard flaperon (05 and 06) ••••••••••••••••••••••••••••••Inboard fl aperon (03 and 04) ••••••••••••••••••••••••••••••Rudder (or)

Maximum control surface deflection rates in degtees per secondHorizontal tail .Fl ap • • • • • • • • • • • • • • • • • • • • • • • • • •• • • • • • •• • ••••••••••••••••••Aileron .Outboard flaperon .•••.•••.•••••.•••••••••••••.•••••.••••••••Inboard flaperon •••••••••••.••••.••••••.•••••••••••••••••••Rudder ••••••••••••••••••••••••••••.•••.••••••.•••••••••••••

±20o to 30

±25±30±10±25

±50±10±70±40±40t50

27

TABLE VI.- DYNAMICSTABILITY CHARACTERISTICSOF AST-107

(158 knots approach speed)

AugmentationSatisfactory Acceptable

criterion criterionParameters None HSAS SCAS ModifiedSCAS

Short-period mode

I_sp, rad/s .167 .704 1.394 1.394 See fig. 39 See fig. 39Psp, s 44.88 12.63 24.63 24.636sp .542 .708 .983 .983 0.35 to 1.30 0.25 to 2.00L_/_sp .... 2.91 .688 .348 .348 See fig. 39 See fig. 39n/s, g un1_s/raa 4.02 4.02 4.02 4.02 See fig. 39 See fig. 39

Long-perlod (aperiodic) mode

t 2 , s 5.32 oo oo oo >5

Long-period (periodic) mode

_ph, rad/s .089 .087 .087

h, s 84.98 89.74 89.74ph .555 .597 .597 _0.04 _0

Rol I mode

t R 1.536 .282 .321 .117 _1.4 _3.0

Spiral mode

tl/2, s 32.46 50.26 oo oo

Dutch roll mode

_d, rad/s .937 .541 1.117 .740 20.4 20.4_d .109 .359 .220 .229 _0.08 _0.02_dmd, rad/s .102 .194 .243 .169 _0.15 20.05Pd, s 6.75 12.43 5.77 8.72 - -@/P 2.81 2.68 .78 .78 - -

Roll-control parameters

_@/(_d .638 1.106 1.017 1.006 0.80 to 1.15 0.65 to 1.356_/6 d 1.90 .636 1.202 1.004

28

TABLE VII.- CONTROL RESPONSE CHARACTERISTICS OF AST-107

(158 knots approach speed)

AugmentationParameters HSAS a SCASa Sat i sfactory Acceptable

None Modified criteri on criterion

SCASa

Longitudi nal

¢max, rad/s2 b_.070 b_.049 b_.070 Same as SCASa b_0.08 b_O.O.

<P/<P ss - - - - - - See fig. 40 See fig. 40

Lateral

<Pmax , rad/s2 .248 .209 .194 .20 See fig. 41 See fig. 41

<P max , deg/ s • 16.57 10.51 23.25 21.10 See fig. 41 See fig. 41

t<p = 30°, s 2.90 3.80 2.82 2.86 :S2.5 S3.2

aAutothrottle onbMinimum demonstrated speed of 129 knots

wo

TABLE VIII.- SUMMARY OF TAKEOFF PARAMETERS

Case number 1 2 3 4.Takeoff performance

Rotational velocity, m/s (kts) 100.3 (195) 100.3 (195) 97.7 (190) 100.3 (195)Liftoff velocity, m/s (kts) 107.3 (208.5) 107.3 (208.5) 104.3 (202.7) 107.3 (208.5)Velocity at obstacle, m/s (kts) 110.8 (215.4) 110.8 (215.4) 108.3 (210.5) 110.8 (215.4)Engine out (VI), m/s (kts) 91.2 (177.2) 91.2 (177.2) 88.4 (171.8) 91.2 (177.2)Engine out (V2), m/s (kts) 107.0 (208.0) 107.0 (208.0) 106.2 (206.4) 107.0 (208.0)Engine out balanced

field length, m (ft) 3518 (11544) 3518 (11544) 3739 (12270) 3518 (11544)

FAR field length, m (ft) 3568 (11706 ) 3568 (11706) 3810 (12500) 3568 (11706)

Cutback performance

Altitude at cutback, m (ft) 250 (822 213 (700) 235 (773)Distance to cutback, m (ft) No 5944 (19500) 6399 (20995) 5944 (19500)Velocity at cutback, m/s (kts) Cutback 125.5 (244) 122.4 (238) 121.7 (236.5)CL at cutback .424 .436 .445LID at cutback 8.5 8.2 8.1Percent thrust, after cutback .626 .641 .652

6.49 km (3.5 n.mi.) flyover poi nt 307 (1008) 273 (897 ) 217 (712 ) 259 (851)Altitude, m (ft)

Above data assumes 20 degree flap deflection.

TABLE IX.- EFFECTIVEPERCEIVEDNOISE LEVELS

Jet EPNdBCase 1 Takeoff, normal throttle setting, accelerating climb,

no power cutback:maximumsideline noise 113.8centerline noise 11.9.4

Case 2 Takeoff, normal throttle setting, accelerating climb,power cutback 5944 m (19,400 ft):

maximumsideline noise 113.3centerline noise 116.1

Case 3 Takeoff, reduced throttle setting*, accelerating climbpower cutback limited by altitude 213 m (700 ft),cutback at 6399 m (20995 ft)


Case 4 Takeoff, normal throttle setting, cutback at obstacleto 93 percent power setting, accelerating climb


Approach, standard 3 degree glide slope, constantspeed 813 m/sec (158 kt):

centerline noise 110.5

Values listed are jet only data and the suppression increment is not yet removedfrom these EPNdBvalues. Text assumes that an increment will be removed.

*A throttle setting of 93 percent was used for the initial portion of Case 3.

31

kN (1 bf) 3153.790 (709000)kN (1 bf) 1343.480 (302026)kN (1 bf) 200.370 (45045)kN (1 bf) 53.432 (12012 )

kN (lbf) 253.802 (57057)m2 (ft2) 774.161 (8333)m2 (ft2) 853.965 (9192 )

TABLE X.- MISSION PERFORMANCE

Mission: Supersonic cruise at Mach 2.62

Aircraft characteristics:

Design gross weight

Operating weight emptyPayload - Passengers, 273

- Passenger, baggageTotal payload weightWing area - reference

- grossGE21/J10-B5 enginos (4); sea level

static (std. +8 C day) installed thrustper engine, N (lbf)

Initial installed thrust to weight ratioInitial wing loading - reference kPa, (lbf/ft2)

- gross kPa, (lbf/ft2)

211. 304

.2684.0743.693

(47503)

.268(85.08)(77.13)

14.946 (3360 ) 0 0 10.0

21.280 (4784) 6 (3) 1.4

379.656 (85350) 970 (524) 36.4

893.928 (200963 ) 6932 (3743 ) .146.7

18.042 (4056) 572 (309) 28.6

0 0 0 0 0

1312.906 (295153 ) 8474 (4576) 213.11327.852 (298513 ) 8480 (4579) 223.1

Mission segment Operating Weightsor condition kN 1bf

Ramp gross weight 3153.790 (709000)Warm-up &taxi-out

Takeoff gross weight 3138.843 (705640)Takeoff run

Begin ascent 3117.563 (700856 )Climb &accelerate

Begin cruise 2737.907 (615506)Cruise segment

End cruise 1843.979 (414543)Descent &decelerate

End descent 1825.937 (410487)Landing &taxi-in

End mission 1825.937 (410487)Trip fuel, range &timeBlock fuel, range &time

FuelkN lbf

Rangekm n.mi.

Timemi n.

32

TABLE XI.- BASELINE DATA FOR DIRECT OPERATING COST ANALYSIS

(Monetary values in constant 1976 dollars)

ww

Gross takeoff weight, kN (lbf)Range, km (n.mi.)Cruise speed, Mach numberNumber of enginesThrust per engine, kN (lbf)Seats (passengers)Load factor, %Fuel cost, cents/liter (cents/gal)Insurance rate, %of purchase price (average value)Year dollarsDepreciation period, yearsResidual value, %Utilization rate, hrs/yr block (flight)CrewPurchase price:

Aircraft, complete millions of dollarsAirframe, millions of dollarsEngines, millions of dollarsSpares, millions of dollars

Crew pay relative to subsonic, %Down payment (5%/6 month intervals)Return on investmentInterest rateSalvage valueTax rateEconomic lifetime

3153.88480.82.62

4211.3273

10010.250.51976

1610

40003

88.5272.0216.59.2711720%15%10%10~~

48%16 yr

(709,000)(4579)

(47503)

(38.7)(1. 0)

(3820)

f/ i 2.25 °

.26_.re f / .

•43.53 [142.83) : 36.58 (120.01)

i _ 41.53 (136.24}= 38.39(125.97) _22.88 {75.06)------I /

25_:- / .25_ /

Airplane/Wing Reference Line-_ " \ I \I

5° _-_----- - -- ......:::...... •.......... "..... _I _.______.._'_ / ._'o TJ---LL_ __:_ .........................._'_- ' - ', _ I1.83" 14.73

(6.00)/'_76oors2.so__I -__-_--__ .. _ .-----I'_ I {,.833_

92.96{305.00}

Figure 1.- General arrangement.

Galtey -40 PassengersGalley- ers _ GatleYl..---130 Passengers

I =J I_:i__::_.........o,_ Lavatory

I -] _ •....

FuelTan., {i.i J

.508(20) Min Fuel Tanks " -

1.981(78)_YP _ _Main Landing Gear

A;A

- Avionics BayFlight xl.829(32x72) Entry Door

Station _ 1.067 x 1.829 (42x72) Airplane/Wing |.,(_.'.'._i /- Entry Door I

_o._ .......... _ _ _ , /_,:9 _.. _1 ......... __ Reference___Line. 1.06?x1.829(t,2x72) .... ._ "Entry Door j

Landinc Baggage and CargoGear LA Fuet Fuel TankFuel Tank (188.5)

Environmental Control Hydrautics BayDistribution Bay

Static Ground Main Landing Gear 'Electrical BayEnvironmental Control Bay

Figure2.- Inboardprofile.

Main Landing Gear

14

12

-] - _ ,_ 1-,_--J-+- -.11

13

Tank Fuel Weight Per TankNumber kN Ibf

1 - 2 84.525 19002

• 3 - 4 97.251 21863

5 - 6 129.377 29085

7 - 8 75.900 17063

9 -10 98.110 22056

11 - 12 158.054 35532

13 -14 92.719 20844

15 -16 38.811 8725 '

17 200.535 45082

18 134.123 30152

19 119.212 26800

Total 2003.364 450 374

Figure 3. - Fuel tank locations and capacities.

36

b/2, feet

0 . 5 10 15 20 25 30 35 40 45 50 55 60 65i i l i l i i | I I J I I ,

,.o f ,eoO,o_O_e,II BreakI JTrailing Edge I Trailing Edge Leading Edge

I Break

Ireak Break Tip

Wing 3.0 _ _"- i _ _ I

Thickness- _ChordRatio,

% 2.0I I

' I1.0I

I!

0 2 4 6 8 10 12 14 16 18 20

b/2, meters

Figure 4.- Spanwise thickness distribution.

co

Figure 5.- Contours of wing thickness, dimensions in meters (ft.).

®

Flap Area, m2(ft 2)Number Each

1 -2 11.687 (125.8)

3-/, 8.070 (86.9)

5 -6 z,.673 (50.3)

7- 8 7.634 (82.2)

9 - 10 15.379 (165.5)

11-12 16.332 (175.8)

13-lZ, 8.421 (90.6)

Figure 6.- Wing control surface geometry and flap definition.

39

4::=O

Aircraft Length, ft0 20 40- 60 80 100 120 110 160 180 200 220 2/.0 260 280 300 320i i I i i I I i ! ! i i i i i I i

30

30025

_u[sion 125020 Galley and Lavator and Lavatory

Section nvironmenta! Control -1200 Section

Area, System / Afrt2a'm2 Wing Fuel Passengers

15 Control

Electrical llS010 drau[ics 1001

Avionics _ossengers Passengers J

5 ]50Baggage and Cargo _ody Fuel

00 10 20 30 /*0 50 60 ?0 80 90 100

Nose Gear Aircraft Length. m

Figure 7.- Volume utilization of configuration.

J e J ;

.010,

.009_

CDLG

.008

"0070 .2 .4 .6 .8 1.0 1.2

cL

Figure8.- Landinggear incrementaldrag coefficientas a functionof CL-in or out of ground effect.

41

1.2

1.o _7 _

CL B1= _2=33=3/., Deg

.6 _--10

X-2o

o

.2-10 -5 0 5 10 15 20 25

e, deg

Figure 9.- Trimmedlift curves (outof ground effect),69 = 610 = 611 = 612 = 300,

613 = 614 = 450, 65,66,67,68= 50, center of gravityof 60.1 percentof Cref.

42

100% Leading Edge Suction

L_" =__==_' LeadingEdge/,_ Suction

cL / 30/J

0 .05 .10 .15 .20 .25 .30 .35 .40 .45 °

cD

Figure 10.- Trimmeddrag polar (outof ground effect), landinggear retracted,center of_gravityat60.1 percentmean aerodynamicchord, 69 = 610 = 611 = 612 = 30U, 613 = 614 = 45U,

65 = 66 = 67 = 68 = 50.

IO I I(_1= 32= 33=_/., deg

e _ /---o

o(a) Landing Gear Extended

12 //_f-\ I I

.__ j $1=_2=_3=_4, dego

n_ 10 10

p_" // 20

30

2- 8 _ •

°_

1--

///0 .2 .Z. .6 .8 1.0 1.2

cL(b) Landing Gear Retracted

Figure 11.- High-lift configuration aerodynamic performance (out of

ground effect), center of gravity at 60.1 percent, °f7c6 68f"69 = 610 = 611 = 612 = 300, 613 = GI4 = 450, 65 = 66 = re = 50.

44

Airplane Equivalent Length, ft

0 40 80 120 160 200 240 280 320 360 400I I I I 1 I I I I I I

24

1,,o--Total Configuration -1200E

d

>, 11600rn

_ 12 S Fuselage"5 r _ -

'" 8

_ -_80 uJ

0 10 20 30 40 50 60 70 80 90 100 110 120

Airplane Equivalent Length, m

_n Figure 12.- Equivalent area distribution, M = 2.62.

.0040

.0030

Cow_---.._.0020 _ _

.0010

01.0 1./-, 1.8 2.2 2.6 3.0

M

Figure 13.- Wave drag coefficient as a function of Mach number.

46

.0070

0061

"-_. Start of Cruise--_

oo CDf

.0040L,.

Q

c" .0030o

co

c .00200

Start of Cruise

-u_c

.oolo /- CDr \

0•4 .8 1.2 1.6 2.0 2.4 2.8

M

Figure 14.- Skin frictionand roughnessdrag coefficientas a function ofof Mach numberfor the design missiontrajectoryof figure49.

47

.0032

.0028

•CDp°

.0020

J

.0016

!

.0012 __y

.0009.4 .8 1.2 1.6 2.0 2.4 2.8

M

Figure 15.- Parasite drag coefficient as a function of Machnumber.

48

.0120

"0110W

.0100

CDmin

.0090

oo_o_ _..0070

.0060.z. .8 1.2 1.6 2.0 2.4 2.8

M

Figure 16.- Minimumdrag coefficient as a function of Machnumber.

49

tj-io

.08 ' " _-Typical Operating Points

0 .004 .008 .012 .016 .020 .024 .028 .032 .036 .040

CD

Figure 17.- Typical drag polars at several Mach numbers for the clean configuration.

15

13

12 1(LID)max

,o _\

9 Start of Cruise

8.4 .8 1.2 1.6 2.0 2.4 2.8

M

Figure 18.- Maximumlift-to-drag ratio as a function of Mach number.

51

O-I

6.358m{20.858ft) Supersonic Cruise _l6.219 m(20.408 ft) Takeoff

Reverse---- 3.241m(10.633 ft) _ _ 2.395 m (7.858ft) =-

'2.027m (6.650 ft) : II_. Rear Mount "_ --Supersonic Cruise

= 1.974 +-.095 m -- .254 m._] 1.534 m(5.033 ft} __](6.L,77+-.313ft) -- (.833 ft_ _ Takeoff I

I _ I' --"'- "'--- .... "'_ _F1.465m I 1.692 m

(/..808 ft) I (5.550 ft)Diameter

Diameter _ _ _- , I - Ic.g.

1.808 m (5.933 ft)Diameter 2.982 m 19.783ft)

Takeoff3.117 m (10.225 ft)

Supersonic Cruise and Reverse

Figure 19.- The General Electric GE211J10-B5 low-bypass ratio engine.

R

C.g.

A C D_EF G

---X

X R AreaStation m2m ft rn ft ft 2

A 0 0 0 0 0 0B 1.9/.8 6.391 .308 1.012 .299 3.217

C 1.9/.8 6.391 .790 2.591 1.959 21.090D 8.936 29.319 .911 2.988 2.606 28.0/.9E 9.933 32.588 .885 2.905 2./.63 26.511

F 11.091 36.389 .8/.6 2.775 2.2/.8 2/..192

(3 12.181 39.96/. 0 0 0 0c. g. (Engine) 7.666 25.152

Figure 20.- GE211J10-B5 engine nacelle.

Aircraft Centerline

WingStation

_- Fuselage Station--1_ Inboard j , , _ "75°

Outboard :

F1.5t-)),- I Aircraft Reference Line _I

Vertical Station

1_I- Fuselage Station -S_L___ _.__/. __80____ -'I IJ

Inboard Nacelle in ft rn

Fuselacle Station 2328.220 194.018 59.137Wing Station 248.736 20.728 6.318Vertical Station -211.583 -1'7.632 -5.37/.

Outboard Nacelle

Fuselage Station 2372.590 197.716 60.264

Winc_Station 407.508 33.959 10.351Vertical Star.ion -204.462 -17.038 - 5.193

Figure 21.- Engine-nacelle location detail.

54

.98

.°/ra: .g6

ou

_ .94o IL,.-i

,. .g2 --_o.

a _J- .go

€-

.88

(a) "P" Inlet Total Pressure Recovery 700

3O0

u __ 600 u__

-_ 250 Ev

5oo irr', u.

200 ==o o

400 z

•=" _ "-

= 150 "

" 300 €Ill UJ

1000 .5 1.0 1.5 2.0 2.5 3.0

M

(b) Air Flow Schedule

Figure22.- NASA/Ames"P" inletperformance.

55

Pressure Altitude, ft

0 10 20 30 40 50 60 70 80x103I' i I I I I I I I

250

"_F-17.54 50x103

200 i_0

z

v. 150 M=1.6m 2.2_- 30 -

_- \ .95 1.2 2.62o_ _

LU 100 9_y \_ .c

tm

20 L_z

•

5010

00 5 10 15 20 25 -

Pressure Altitude. km

(a) Maximum Climb Power Setting

Figure 23 - GE21/J10-B5 installed net thrust at standard day plus eight degreescentigrade ambient temperature.

56

Pressure Altitude, ft

0 10 20 30 40 50 60 70 80x103

25O

iix3200 '.54

\.:,

150-z

M=1.6 -J30 "Q

" \ "/!= \ 2.2._ .8P" \.95 12 2._" 100"_ \ 2.62: -12o-_I.U

Z .9 '_

50 _ Jl000 5 10 15 20 25

Pressure Altitude, km

(b) Maximum Cruise Power Setting

Figure 23 - Concluded.

57

Pressure Altitude, ft0 10 20 30 40 50 60 70 80x103I I I I I I l J I

35

\.3

0 70x103

30, .7

\ -. 60

25 '

20 _ )_ 2.62 ._

1.2 \E

d _c.__ "__ C

c 15 uJLU

30Q_

3o _o__ u_

-_ 10 u.u. 20

5 10

00 5 10 15 20 25

Pressure Altitude, km

(a) Maximum Climb Power Setting

Figure 24 - GE21/J10-B5 installed fuel flow rate at a standard day plus eightdegrees centigrade ambient temperature.

58

Pressure Altitude. ft

o 10 20 30 40 50 60 70 80x103'-,---T"j---Tj----,..----,,----j,------rj---""'1;'-"'---',

35....-------..------r------,------r-------,

30~-----4------+--------l-----+-------1

60

10

50

....r:. ..." .r:.Cl "~ E

.aCII 40c: gjCl c:c: ClW c:... WCII ...a.. CII

a.~ 300 ~

u. 0

u.CII::l CIIU. ::l

10 u.

20

o 5 10 15Pressure Altitude. km

20 25o

(bl Maximum Cruise Power Setting

Figure 24 - Concluded.59

o Net Engine Thrust, Ibf

0 10 20 30 40x103I I = I I

25

20 ¢ 1 50x103

\o_ Altitude= 13716 m(z_000 ft) _a; 13716(45 000 ft)

._.o ,.ooo,,,\" /f1.1.1

_. 10 16 746 m (55000 ft) _l t/_

2LL

-_" _7620m (25000ft)u_ 5

f . 110 _JO

0 50 100 150 200

Net Engine Thrust, kN

Figure 25.- Installed fuel flow for maximumand part power cruise at a standardday plus eight degree centigrade ambient temperature.

Wing Loading, psf

100 95 90 85 80 75 70J I I I I i I

Gross Weight.40 MN (Ibf)

3.670 (825000)

3.559 (800 000)3.447 (775000)

.36 _ 3.336 (750000)o 3.225 (725000)

o 3.203 (720000)n_

._ .32 3.181 (715000)t-

.__m \ 3.158 (710000)

/s \!

_ .28

F-

.2_

Design Pointr Code=100

Absolute Fuel Limit Power Code=70

Takeoff Field Length.20 3810m (12500ft) Standard +10 C Day

t I I I I I I I I4.8 4.6 4.4 4.2 4.0 3.8 3.6 3.4 3.2

Wing Loading, kPa

Figure 26 - Aircraft sizing "thumbprint" for the AST-107 carrying 273 passengers overa range of 8334 km (4500 n.mi.) at M = 2.62 hot day cruise conditions.

750x103-_'_kk Forward and Aft

eoff and Landing Limits3200 . Takeoff

/, Gross Weight 700/

/3000 1 1

f

/ 650 _-

2800 1Begin

Cruise <600

2600

Desired Cruise Aft

2/.00 c.g, Path For _ Stability 550z MinimumDrag Limit

€..2200 500 _-_

2000 450

End /Cruise _ ........ ._>1800 Normal - 400

Landin(Weigt

1600Zero 350Fuel

Weight

1400 Operatingt 300

• EmptyWeight

1200 l I = , _ = T f l = I I I I38 40 42 44 46 48 50 52 54 56 58 60 62 64 66

Center-of-grevity, Percent F.ref

Figure 27. - Center-of-gravity travel diagram.

62

25

2O

150

u

a

ul

-oilJ

"_ 5

O' 5 10 15 20

±Horizontal Tail Deflection, deg

Figure 28. - Assumedhorizontal tail/elevator deflection relationship.

63

Equ=valentAirspeed, knots

120 1/.0 160 180 200 220 240 260= I I I | I I I o

1.0

elastic

_,_ _8 5 on( 86.5

_8_ on( 8e

.260 70 80 90 100 110 120 130 1/.0

' Equivalent Airspeed, m/sec

Figure 29. - Estimated lateral control flexibility factors.

64

Steady 90° Crosswind Velocity, knots0 5 10 15 20 25 30u I I I I I =

Percent Lateral Control Required----_ /_r2C r / 100

_,/,,//_ 8015 (Lateral Control o_

8r,/_,_,deg Limit) • _;-60 _o

10

5 201-

._1

1 o0 3 6 9 12 15

Steady 90 Crosswind Velocity, m/sec

Figure 30. - Crosswind trim capability at an approach speed of 81.3 m/sec (158 knots).

65

Equivalent Airspeed, knots -120 140 160 180 200 220 240 260I = = i I I I J

01.oo io,.i=

v

I-O

u .98Cl

.D

" .96

-u.

o

_- .94 _t-O

a

c .92O

•- _

u '_8 =25°._,= rma>.90

60 70 80 90 100 110 120 130 141

Equivalent Airspeed, m/sec

Figure 31. - Estimated directional control factors due to fuselage bending.

66

Equivolent Airspeed, knots40 60 80 100 120 140 160 180

I I I I I I I I

_ 3oL

€-

•£ 20"6

- _

"O

020 30 40 50 60 70 80 90 100

Equivalent Airspeed, m/sec

Figure 32. - Estimated directional control factors due to fuselage bending.

67

co

1.2 Center of Gravity Geared St, deg

in Percent F.ref / 24• 6o1__k / o lo

-10-_1.0 • -20 -20

.8 -16

CL otrim -12

.6 _ awr p, deg8

.4

.2 t -4 --40 ' I I I

/ , i-8-.2

\ToilOff

-10 0 10 20 30 .15 .10 .05 0 -.05 -.10

awr p, deg Cm about.6Ol_.re f

Figure33. - Directionaltrim required in 90-degreecrosswindat maximumgrossweight for an angle of attackof the wing referenceplane of -50 and a

crosswindvelocityof 11.58 m/sec (22.5 knot).

Geared St' deg1.2 Center of Gravity

in Percent _ref / 0 10 _24

6o.1--_ / -2__1.0

.8

CL -12trim

.6 awrp, deg-8

.4 Z.

2 I0 _/ = ' I

-.2 \Tail Off

-10 0 10 20 30 .15 ,10 .05 0 -.05 -.10

awrp, deg Cm about .601F.ref

Figure 34. - High-lift trim and stability (out of ground effect) with power off,

65 : 66 = 67 = 68 = 30o , 61 = 62 : 63 : 64 : 20o .

' O

Center of Gravity Geared8t, dee 2010 v..241.2 in PercentF,ref 0 246 '__'_'_""

-20 [.nstantaneous

Minimum , 160 =-.08 rad/sec

Demonstration7 _".8 Speed / _'_ 12

CLtrim c.g. Limit_ =wrp, dee

.6 Approach ._

.4

•0 ' I , J , __ , _--_\ _,_ \\ \ ,\ .._r_ -8

/ \-.2 - Tail Off

-10 0 10 20 30 .15 .10 .05 0 -.05 -.10

awrp, dee Cm about .601_,ref

Figure35. - High-lifttrim and stability(out of ground effect),with power off,69 = 610 = 611 = 612 = 300, 613 = 614 = 450, 65 = 66 = 67 = 68 = 50,

61 = 62 = 63 = 64 = 200.

Geared 8 t, deg 20

1.2 Center of Gravity 0 10 -2/,

in Percent Cref / -10 -'

1.0

.8. =wrp' deg

CLtrim8

.6

-Z,

.z,

.2 / "_ ,-80 1 i I I ,

-.2

-10 0 10 20 30 . .15 .10 .05 0 -.05 -.10

awrp, de9 Cm about .601F.ref

Figure 36. - High-lift trim and stability at thirty degrees (out of ground effect)

with power off, 69 = =_1 611 = _12 = 3_v, 613 = 614 = 450 ,61 °2=:63= 4 : 30.

.65IU

C

oU

o jJ .60

E .55

E

< .500 .5 1.0 1.5 2.0 2.5 3.0

M

Figure 37. - Flexible airplane aerodynamic center location variation with Mach numberwith all flaps undeflected except 613 = 614 = 20o at Mach numbers less than 1.2.

72

.10

f_

,05 __ jj

-.05

c_-.10 Cl/_

cy_rad -1

-.15

-.20--1 g Cruise --2.5g MQneuver

-.25 _ __

-.30 I0 .05 .10 ,15 .20 .25 .30

CL

Figure 38. - Flexible static lateral-directional stability at M = 2.62.

73

Stability and Control Augmentation System

a Hardened Stability Augmentation System

50 Open symbols denote configuration of reference 3

Solid symbolsdenote AST-107

10 Unacceptable

j Unacceptable

.3- I I I I ! I lllJ I I i I = a Ill1 2 3 10 20 30 100

n/a, g Units/rad

(a) Longitudinalshort-period frequency requirements of reference 12

1.2 _ Stability and Control Augmentation Systemo Hardened Stability Augmentation System

Unacceptable1,0 Open symbols denote configuration of reference :3

Solid symbols denote AST-107

_ Subsonic jet transports

La/°JsP .8 /_table

Per rod .6 (

.z. _ ( Satisfact°ry &

I I I I I I I I I I I I I I I I

0 .2 ./. .6 .8 1.0 1.2 1.Z. 1.6

Short-period Damping Ratio

(b) Shomber-Gertsen longitudinal handling qualities criteria of reference 23

Figure 39.- Comparison of longitudinal short-period characteristics with handlingqualities criteria.

74

3.5

3.0

2.5 Stability and Control Augmentation System

AST-1072.0 ..... of reference 12

1.5

, ,, ',,,0 1 2 3 4 5 6 7 8 9 10

Time, sec

Figure40.- Comparisonof low-speedpitch rate responseusing the stabilityand control augmentationsystemcriteriafrom reference24.

75

10O Unaugmentedr'l Hardened Stability Augmentation SystemA Stability and Control Augmentation System

Modified Stability and Control AugmentationSystem

Open symbols denote configuration of reference 3Roll Acceleration, 1 Solid symbols denote AST-107

rad/sec 2

.0_1 10

Roll Mode Time Constant, sec !

(a) Roll acceleration response boundaries for large aircraft.Boundaries from reference 2z.

O Unaugmented70 O Hardened Stability Augmentation System

A Stability and Control Augmentation SystemModified Stability and Control Augmentation System6O

u Open symbols denote configuration of reference 3Solid symbols denote AST-107

50"10

a; /+0 Acceptable Areao +n,-

DC-8-_ 30On_

E 20 • L3& +_-_ 707-E-_ Fr'T'T4_ \ s\s_ \_ \ \s ,_\ s s s _ _-_, - v O

o 10 • Q_: Unacceptable Area

I I i I I I I I I = I I I I I I I I I I

0 .2 .4 .6 .8 1.0 1.2 1.Z, 1.6 1.8 2.0

Roll Mode Time Constant, sec

(b} Roll-rate capability criterion for transport aircraft.Boundaries from reference 25

Figure 41.- Comparison of roll response with criteria boundaries.

76

Approach Takeoff

Threshold Point / Thrust Cutback_t_/_ S

Altitude= 15.2/+m (50 ft) /

Liftoff -------z J

_Brake Release Runway Centerline/For Takeoff / 7 / Measurement

/" _

_ej n 6_y 9__

(0.35 n.mi)

_,_ 6/,86 m r(3.5 n.mi.)

Measure t

Measurement Point 3 Point 2

_NOTE; Sideline noise is measured where noise level after liftoff is greatest.

• ",,.I

"_ Figure 42.- Noise measurement locations prescribed in reference 26 for approachand takeoff.

L=t"I- 120E-_ 100(E o'_ 80 to t=_ b o"_ 60"-

"_ 4o-U

o_ ®_Normal Throttle (Case 1)E] .... Reduced Throttle (Case 2)

Distance From Brake Release. ft

0 10 20 30 40x103I I I I I

1200 4x10 3

Climb at V2+ 18 m/sec 135

1000

3

Climb at V2 + 14 m/sec (278OOE

<_

400 "_

1

2O0

I I I 00 2 4 6 8 10 12

Distance From Brake Release, km

Figure 43.- Comparisonof normalthrottleand reducedthrottletakeoffprocedures.

78

Downrange Distance From Brake Release, ft

0 5 10 15 20 25 30 35 40 45 50 55x103I I I I I i I I I I I I

-_ 6.486 km _ 8x103(3.5 n. mi)

2.0 /--EPNL= 108 dB

1.5 _ _ _ sq. km (16.03 sq. mi.)/.

1.0

.5 Runway PNL=115 dB 0649 km--_- Area 8 11sq (0 3 a;,-° 5n ml ) u

f 1 °a 0 0

= .5- _5.E mm C

m 1.0- .'g_L/)

4

1.5 - __

2.0-

I I I I I I I I I I I I I I I , I I -80 1 2 3 4 5 6 7 8 9 10 11 12 13 1/. 15 16 17

Downrange Distance From Brake Release, km

Figure 44.- Noise contours for Case 1 type takeoff (jet noise only).

Coo

Downrange Distance From Brake Release, ft0 5 10 15 20 25 30 35 40 45 50 55 60x10"J I i l I i I 1 i I I I I

6.486 km _ 8x103{3.5 n. mi.)

2.0-

1.5 /--EPNL = 108 dB

Area=32.37 sq. km (12.5 sq. mi.) 4

1.0 _ ,-

a; - /-EPNL = 0.649 km_-u /-Runway _/Area =5.57 sq. km 12.15 sq. mi.) (0.35 n. mi) '-oo

0 I ) - 0aD m

_ .c_._c .5•_o u3u)

1.04

1.5

2.0

I I I I I I i I I I I i I I I I I I I ' I -8

0 1 2 3 z. 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19

Downrange Distance From Brake Release, km

Figure 45.- Noise contours for Case 2 type takeoff (jet noise only).

Case Number

T118 -3.3

@-T---

_- -.5 -.4

114 _ --_ @ t @ t @

nO _ (_3 0 ("300Z .D l 0 _-O _-0 --

,,, 110 .x u _ --...: ca I _ ._ E a

-J _ _ "O .D n

tn o 0 (0 _ Q.•_ 106 = i m _ ,,- oZ 0 I 0 0

I Qj Qj I-- ap-

• l _= _= _ L. =-g-IO O O C

E u E E -,o u')• ,-- _ _ _ Q;

x i x _ E E u0 I 0 m --- .-- "1_: I _: _ x x -o

98Flyover Sideline Approach

Figure 46.- Comparison of noise results (jet noise only) for the differenttakeoff procedures.

82

Passenger Load Factor

100 %

140 3

120 -0 111Q. 0.

Ql eli.... 100 ....:J 2

:J111 111111 Begin 111Ql Ql.... 80

....0. 0.....

End Cruise~I....

Ql Ql

> >0 0

60 EE 00 00 1· CDCD

0 40 0

c c00 UlUl

20

0 01.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 2.8

M

Figure 47.- Maximum over-pressure during climb and cruise.

893.928 kN (200963 Ibf)

Cruise at optimum altitude

17 983m (59000ft) or climb ceiling _ f20510 m (67 290ft}

End cruise altitude

Begin cruise oltitude_ \

379.656 kN (85 350 Ibf) / \Climb,Accelerate --_ / \ 18.042 kN (/*056 Ibf)

111110r-_'1/"9l'6kN (3 360Ibf)minrainTakeoffTaxi --__4V _/_/ _Descent, DecelerateI 21.280kN(/, 78/*llbf)_f I

_--/-,_-I-- j_ Trip range 8/,74km (/*576n.mi.)

rj__ Trip fuel 1312.906 kN (295 153 Ibf)-_, Block fuel 1 327.852 kN (298513 Ibf)

_, Block time 223 min

Block range 8/.80 km (/.579 n.mi.)

NOTE: Civil Aeronautics Board Range=Trip range minus traffic allowance as specified forsupersonic aircraft.

(a) Primary Mission

M=0.8/,at 10062m (33011ft)

Cruise at best altitude and veiocity_ 69.570 kN (15 6/.0 Ibf)30rain Hold at f4=0.5

65.660 kN (1/.761 Ibf) \ 30/,8 m (10000_"

5 o/= Trip Reserve

I/.63 km (250 n.mi.) ,._]

6810 kN (1531 Ibf) v I- Missed approach To alternate airport

86.633 kN (19/.76 Ibf)

(b) Reserve Allowance Mission

Figure 48.- Mission profile and fuel weights associated with each profilesegment to indicate the fuel burned during that segment.

83

2270)(103--20 ----- 15018 ,,"

././

16 ./

// 50E 14 /

.Y. / -/ 40 -ai 12 / ai't:l/ 't:l:J

:J- 10 I -- 30 -;( I ;(8 I

IClimb6 / 20

I - - - - Descent4

I/

10/2 /

/0

0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2..4 2.6 2.8

M

Figure 49.- Climb and descent schedule as a function of Mach number for astandard day plus 8 degree centigrade ambient temperature atmosphere.

Range, n.mi.

0 1 2 3 4 5 6 7x103I I I I I I i I

60x10326O

240220 50

2OO

180 i _ - 40

z 140 ,,-"_ -30 -_"oo120 Cruise Mach Number _o"

o o_, _- 2.62 oo 100 .... o.go _"oo.. 20 n.

8O

°o i,,l40 10

2O

I I I I I I I I I 00 1 2 3 4 5 6 7 8 9 10 11 12

Range, Mm

Figure 50.- Payload-range capability variation with cruise Mach number.

85

Subsonic Flight Segment. n.mi.

0 2 3 4 5x103I I I I I

10.0

Cruise M=2.625.25x103

9.5

5.0060% Payload

9.0

4.75

8.5 'EE

4.50 c:~

a) aiCl

Cl8.0

cc aa 4.25 0::

0::

a a-- 00 l-I- 7.5

4.00100 % Payload I

I

7.03.75Cruise M=0.90

3.506.5

90 2 3 4 5 6 7 8

Subsonic Flight Segment. Mm

Figure 51.- Total range as a function of subsonic flight segment and passengerload factor.

86

2.2 1.2)(103

2.0 1.1Drag Coefficient

1.8 1.0

1.6 Specific Fuel Consumption.9

1.4.8

1.2.7

1.0 '".6

\.... \.

~..5

.8 \.\. .4

.6 ... "EStructural wei9h~~

.3 .-~ .4

E

oj ... ~ .2 C

Cl .2~ oj

c Cl

0.1 c

c:: 0

c 0 0c::

QI

C

Cl -.2 % Change in Parameter -.1 QI

c Cl

0c

..c -.40

t.)-.2 ..c

t.)

-.6 -.3

-.8-.4

-1.0-.5

-1.2-.6

-1.4-.7

-.8-1.6

-.9-1.8

87

Figure 52.- Range sensitivity to changes in drag, structural weight, and specificfuel consumption.

Range, n.mJ. =

0 1 2 3 4 5 6i , J i , I I

_.0 I I I100 % Load Factor

6m__ 60% Load Factor

__E3.5 E I I "-Ep, ',m "'El -J_ M=0.90" "'-'-B-,--B- ---Q- -'O""-'=-

3.0D ul13. O

Q, E "_

L_2.5 \ /. o

_ o

m O_ "_:3"" -_':t _ M:0.90 o_--'--"2.0 " --43 M=2.62- .-_"_) 3 of.-O. Q)

o Q-

._ e o1.5 _ *6

•- _ P

o _"_""""(_----(9-----_)-- ---(_)-----"Q M=2.62 2 o

1.00 1 2 3 4 5 6 7 8 9 10 11

Range, Mm

Figure 53.- Effect of passenger load factor and cruise speed on direct operating cost.

88

Range. n. mi.0 1 2 3 & 5 6x103I I I _ I I I 11

6.5

__ m 100% Load Factor 106.0 \ 60% Load Factor\

\\

5.5 I_k "-"

E \ E

t-in 5.0 \ "_. 8 =

= "'1_--E]--.-E_ Vl=0.90mu \ n

"_ 7 c_

o0 _

o

_- 3.s _.. •M=0.90

oo ='-e]_'--13......E}_ L .ELM=2.62

3.0

2.5 "-e..,- "v"" "-e-- "-e M=2.62

2.00 1 2 3 /. 5 6 7 8 9 10 11 3

Range. Mm

Figure54.-Effectof passengerloadfactorandcruisespeedon totaloperatingcost.

89

Range,n.mi.

0 1 2 3 4 5 6x103I I I I I I I

500

100% Load Factor

.... 60% Load Factor

/1_ I M=0.90soo /

/

/J

# //El M=2"62300 I /I_

o / M=0.90 i rK'

_ M=2.62/-,, /

._u200 I

r/E

100o

0 2 4 6 8 10 12

Range, Mm

Figure 55.- Effect of passenger load factor and cruise speed on ticket cost.

90

Range, n. mi.

0 1 2 3 4 5 6x103I I J I I i l

350x103

1.50 M=O._O M=2.62

2501.00

z

/ 200 __

•-." _t .a

u. .75 0_:Du LL

o j 150 "uorn

.50

//__//_ _100% Load Factor 100.... 60% Load Factor

.255O

0: .0 2 z. 6 8 10 12

Range, Mm

J

Figure 56.- Effect of passenger load factor and cruise speed on block fuel.

91

Range, n.mi.

0 1 2 3 4 5 6x103I I I I l I

16 38

36

14 "(9 M=2"62 32

_ 30 =0

..1 Q _ --

\ 12 o28 otn

-,%E M=0.90 26O

_' 10 24," u')

o ...6]" ---E_---'i2b.. 22

/ 20M=2.6 Or: I:L

8 i_.. __.Gj_____.1...19....qk 18--100% Load Factor "19G]M=o.g0 16.... 60% Load Factor

6 I I ,_0 2 4 6 8 0

Range, Mm

Figure 57.- Effect of passenger load factor and cruise speed on passengerstatute-miles per gallon.

92

Fuel Cost, Dollars/Gallon

0 1 2 3 4I I I i I I I i |

13'

20

12 .'1M=2.62 --100% Load Factor , 19

/.... 60% Load Factor / 18

11 //

17E / .-

_. 10 / 16 EQ) /t-- / €:_15

9 / _co / 14 mrl

m / nc 8 / I_M=2.62 m 13

/ ' "E

= 12 o

j "o 7 "_o J 11 =0m_ / o

6 / "-

o / 9 _,o / o•_. 5 8o / "6

.:{ j "°4 /

1/ / 6

3 r{/_ 5E 4

2; 0 10 20 30 40 50 60 70 80 90 100 3

Fuel Cost. Cents/Liter

Figure 58.- Sensitivity of total operating cost to variation in fuel cost, 1976 dollars.

93

6

"E...CII01CCII1/11/10

5 a.

"1/1-CCII

U

,.;1/10

U

01.!:-0...CIIa.a

.E0I-

4

60005000400030002000

0 I T-- 100 % Load Factor

---- 60 % Load Factor·

[~

5,

" ,,'1

~ .........., ,

'''{~''G M=2.62

c~

~~

- -0 M=2.62

2.5

2.01000

3.

4.

...CII01CCII1/11/1oa.

"1/1c8 3.0,.;1/1o

U01.!:-o...

CIIa.oo-oI-

utilization. hrs/yr

Figure 59.- Sensitivity of total operating cost (TOC) to variation in aircraftutilization per year.

94

i

1. Report No. 2. GovernmentAccessionNo. 3. Recipient'sCatalogNo.NASATM-81872

4. Title and Subtitle 5. Repor_ Date

CONCEPTUALSTUDYOF AN ADVANCEDSUPERSONICTECHNOLOGY September 1980TRANSPORT(AST-I07) FORTRANSPACIFICRANGEUSINGLOW- 6. PerformingOrganizationCodeBYPASS-RATIOTURBOFANENGINES

7, Author(s) 8. Performing Organization Report No.

Shelby J. Morris, Jr.; Willard E. Foss, Jr. andMiltnn ,1 Np_halmr..It 1o. Work Unit No.

9. Performing Organization Name and Address 517-53-43-01" NASALangley Research Center

Hampton, Virginia 23665 11 Contract or Grant No.

13. Type of Report and Period Covered

12. Sponsoring Agency Name and Address Technical MemorandumNational Aeronautics and Space AdministrationWashington, DC 20546 14 SponsoringAgency Code

15, _ppfementary Notes

This paper is partially based on unpublished memorandaby R.A. DaCosta, G.J. Espil,W.A. Lovell, J. W. Russell, P.M. Smith, E.E. Swanson and K.B. Walkley of theHampton Technical Center of Kentron International under contract NASI-16000.

16, Abstract

A new advanced supersonic technology configuration concept designated the AST-I07,using a low bypass-ratio-turbofan engine, is described and analyzed. The aircrafthad provisions for 273 passengers arranged five abreast. The cruise Mach number was2.62. The mission range for the AST-107 was 8.48 Mm(4576 n.mi.) and an averagelift-drag ratio of 9.15 during cruise was achieved. The available lateral controlwas not sufficient for the required 15.4 m/s (30 kt) crosswind landing condition,and a crosswind landing gear or a significant reduction in,dihedral effect would benecessary to meet this requirement. The lowest computed noise levels, including amechanical suppressor noise reduction of 3 EPNdBat the flyover and sideline moni-toring stations, were 110.3 EPNdB(sideline noise, 113.1EPNdB (centerline noise)and 110.5 EPNdB(approach noise).

7. Key Words (Sugg_ted by Author(s)) 18. Distribution Statement '

Aerodynamics; Aircraft Design, Testing, Star Category Oland Performance; Aircraft Propulsion and Unclassified - UnlimitedPower; Aircraft Stability and Control;Advanced Supersonic Transport; AircraftNoise; Sonic Boom

19. S_urity _assif. (of thisreport) 20. S_urity Classif.(of this _ge) 21. No. of Pa_s 22.' Dice°

UNCLASSIFIED UNCLASSIFIED 94 A05

N-305 ForsalebytheNationalTechnicallnformationService,Springfield,Virginia22161

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