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NASA Technical Paper 3559

Aerodynamic Characteristics of TwoWaverider-Derived Hypersonic CruiseConfigurations

Charles E. Cockrell, Jr., Lawrence D. Huebner, and Dennis B. Finley

July 1996

https://ntrs.nasa.gov/search.jsp?R=19960045290 2020-07-02T23:02:56+00:00Z

NASA Technical Paper 3559

Aerodynamic Characteristics of TwoWaverider-Derived Hypersonic CruiseConfigurations

Charles E. Cockrell, Jr. and Lawrence D. Huebner

Langley Research Center • Hampton, Virginia

Dennis B. Finley

Lockheed-Fort Worth Company • Fort Worth, Texas

National Aeronautics and Space AdministrationLangley Research Center • Hampton, Virginia 23681-0001

July 1996

Available electronically at the following URL address: http:lltechreports.larc.nasa.govlltrslitrs.html

Printed copies available from the following:

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(703) 487-4650

Abstract

An evaluation was made of the effects of integrating the required aircraft compo-

nents with hypersonic high-lift configurations known as waveriders to create hyper-

sonic cruise vehicles. Previous studies suggest that waveriders offer advantages inaerodynamic performance and propulsion�airframe integration (PAl) characteristics

over conventional non-waverider hypersonic shapes. A wind-tunnel model was devel-

oped that integrates vehicle components, including canopies, engine components, and

control surfaces, with two pure waverider shapes, both conical-flow-derived wave-

riders for a design Mach number of 4.0. Experimental data and limited computationalfluid dynamics (CFD) solutions were obtained over a Mach number range of 1.6

to 4.63. The experimental data show the component build-up effects and the aero-

dynamic characteristics of the fully integrated configurations, including control sur-

face effectiveness. The aerodynamic performance of the fully integrated configura-

tions is not comparable to that of the pure waverider shapes, but is comparable to

previously tested hypersonic vehicle models. Both configurations exhibit good lateral-

directional stability characteristics.

1. Introduction

A waverider is any shape designed such that the bow

shock generated by the shape is perfectly attached alongthe outer leading edge at the design flight condition. The

waverider design method leads to several potential

advantages over conventional non-waverider hypersonic

concepts. The attached leading-edge shock wave con-

fines the high-pressure region to the lower surface and

results in high lift-drag ratios. Several design predictions

suggest that waveriders may offer an aerodynamic per-formance advantage in terms of higher lift-drag ratios

over non-waverider hypersonic concepts (refs. 1 and 2).In addition, the flow field below the waverider bottom

surface is uniform and, in the case of waveriders derived

from axisymmetric flow fields, there is little or no cross-

flow in this region, making these shapes attractive candi-dates for engine integration. These advantages have led

to interest in using waverider shapes for the forebody

geometries of hypersonic airbreathing engine-integratedairframes. Waveriders have been considered for various

types of missions including hypersonic cruise vehicles,

single-stage-to-orbit vehicles, airbreathing hypersonic

missiles, and various space-based applications (ref. 3).

The purpose of the current study is to examine the

aerodynamic characteristics of two waverider-derived

hypersonic cruise vehicles. No experimental data cur-

rently exist that address the integration of realistic vehi-

cle components with waverider shapes. Therefore, the

objectives of this study were threefold. The first was tocreate an experimental and computational database for

waverider-derived configurations. The second was to

examine the effects of individual vehicle components on

pure waverider performance and to determine the differ-

ences in aerodynamic characteristics that result from

integrating all vehicle components. The final objective

was to evaluate the controllability of each of the fullyintegrated vehicles and the effectiveness of the control-

surface design. These objectives were accomplished

using results from wind-tunnel testing and a limited num-

ber of computational fluid dynamics (CFD) solutions.

The CFD predictions were obtained for the pure wave-rider shapes only and provide comparisons with experi-

mental data and design-code predictions. Two wind-

tunnel models were designed that integrate canopies,

engine packages, and control surfaces with two Mach 4.0

pure waverider shapes. The models were tested in the

Langley Unitary Plan Wind Tunnel (UPWT) at NASA

Langley Research Center.

This report describes the waverider aerodynamic

design code used and discusses the method used in thedevelopment of the wind-tunnel models. The details of

the experimental study are then presented as well as the

computational method used to obtain the CFD predic-

tions. The results are analyzed in three sections. First, theresults of the pure waverider shapes without integrated

vehicle components are presented. These results include

flow-field characteristics from CFD solutions and experi-

mental flow-visualization data as well as aerodynamiccharacteristics from the experiment and CFD predictions.

Second, the experimental results of adding aircraft com-

ponents to the pure waverider shapes are presented. The

effects of the canopy, engine components, and control

surface additions on aerodynamic performance and

stability are examined. Finally, the aerodynamic charac-

teristics of the fully integrated waverider-derived config-

urations are examined and compared with those of the

pure waverider shapes. Control-surface effectiveness isalso addressed in this section.

2. Symbols

B.L.

G

Ct_

G

cM

cn

C_

L/D

P

Re

U

V

W.L.

XC.g.

X,F,Z

y+

buttline of model (distance from centerline in

spanwise direction), in.

drag coefficient

rolling-moment coefficient

OC t

rolling-moment derivative, _

lift coefficient

pitching-moment coefficient

yawing-moment coefficient

_C n

yawing-moment derivative, _l 3

moment reference length, in.

C L

lift-drag ratio,

Mach number

model station (distance from nose in stream-

wise direction), in.

pressure, lbf/fi 2

roll rate, deg/sec

Reynolds number

planform area, ft 2

velocity component, ft/sec

total volume, ft3; velocity, ft/sec

V2/3

volumetric efficiency, Sref

waterline of model (distance from zero refer-

ence in vertical direction), in.

moment reference center location

Cartesian coordinates, in.

inner law variable

angle of attack, deg

sideslip angle, deg

angle of aileron deflection (trailing edge down

positive), deg

angle of elevon deflection (trailing edge down

positive), deg

distance from solid boundary to first cellcenter, in.

viscosity coefficient, lbf-sec/ft 2

computational coordinates

2

p density, lbm/ft 3

Subscripts:

c conditions at first cell center next to solid

boundaries

c.g. center-of-gravity location

flee-stream conditions

3. Configuration Design and Model

Development

3.1. Waverider Design Method

A specific waverider shape is uniquely defined by

free-stream conditions, the type of generating flow-field

body, and a leading-edge definition (ref. 1). The shapes

of the upper and lower surfaces of the configuration fol-low from these parameters. The free-stream conditions,

including Mach number and Reynolds number or alti-

tude, are selected based on mission criteria. The design

method used in this study involves a specific design

point. The generating flow-field body is used to definethe shock shape upon which the leading edge of the

waverider is constructed. Although any arbitrary body in

supersonic or hypersonic flow can be used as a generat-ing flow-field body, this study focuses specifically on theclass of conical-flow-derived waveriders, in which the

generating flow-field body is a right circular cone in

supersonic or hypersonic flow. At the outset of thisresearch effort, this option was the best available for the

application of interest. Other possible generating flow

fields include osculating cone flow fields (ref. 4), hybrid

cone-wedge generated flow fields (ref. 5), and inclined

circular and elliptic conical flow fields (ref. 6). The

length of the generating cone, length of the waverider,

and semiapex angle of the cone are specified by the

designer. The selection of these parameters can signifi-

cantly affect the shape of the waverider generated as well

as the aerodynamic performance of the configuration.

Figure 1 illustrates the design of a conical-flow-derived

waverider. The planform shape, or leading edge, isdefined on the shock wave produced by the cone. The

lower surface of the configuration is defined by tracingstreamlines from the leading edge to the base of the cone.

The result is that the lower compression surface is astream surface behind the conical shock wave. The con-

figurations studied here have an upper surface that is

designed as a constant free-stream pressure surface.

However, other techniques may be used, such as shaping

the upper surface as an expansion or compression

surface. The conical flow field, defined behind the

shock wave, exists only below the lower surface ofthe waverider.

The resultingconfigurationoffers two possibleadvantagesovernon-waveriderhypersonicconfigura-tions.Thefirst is a potentialaerodynamicperformanceadvantage(refs.1, 2, and7). Theoretically,theshockwaveisperfectlyattachedalongtheouterleadingedgeatthedesignMachnumber.Theresultis thatthehigh-pressureregionbehindtheshockwaveisconfinedtothelowersurface,andnoflowspillagefromthelowersur-faceto theuppersurfaceoccurs.Themaximumlift-dragratiosthismethodproducespromisetoexceedthoseofexistinghypersonicconfigurations.Figure2,takenfromreference2, showsthetraditional"L/D barrier"in thesupersonic/hypersonicregimeforconventionalvehicles.Thiscorrelationisempirical,basedonactualflightvehi-cleexperienceatsubsonicandlowsupersonicspeedsandextrapolatedto hypersonicMachnumbers(ref.7).Thesymbolsin figure2representpredictionsforavarietyofconical-flow-derivedwaveridershapesgeneratedusingthecurrentmethod,whichisdescribedindetailin refer-ence2.Thewaveridershapesrepresentedhereareonlytheforwardportionsof possiblehypersonicconfigura-tionsandthereforearenotrealisticvehicles.Thepredic-tionsshownassumethattheconfigurationhaszerobasedragin orderto removetheeffectof thebluntbase,whichwill beeliminatedin a fully integratedvehicle,andshowonlytheperformanceof theforwardportionofsucha vehicle.In otherwords,thepredictionsassumethatfree-streamstaticpressureactsatthebase,makingadirectcomparisonof thelift-dragratiosfor waveridersandthoseof existingsupersonic/hypersonicconfigura-tionsdifficult.Furthermore,thewaveridersrepresentedheredonothavelevelsof volumetricefficiencycompa-rabletothoseofthevehiclesusedin theL/D barrier cal-

culation and may not have been obtained at similar

flight-scaled Reynolds numbers. Although the lift-drag

ratios of a fully integrated waverider configuration withthe blunt base closed would likely be lower than those

for the pure waverider shape, these predictions suggest

that waveriders may offer an aerodynamic performance

advantage over non-waverider vehicle concepts. Another

advantage of axisymmetric waverider flow fields is that

the lower surface flow field is uniform, and there is pure

conical flow in this region for a perfectly attached shockwave. Therefore, a known uniform flow field can be

delivered to scramjet engine modules on the lower sur-

face, providing a benefit in propulsion/airframe integra-

tion (PAl) (ref. 8). The osculating cone and cone-wedge

concepts mentioned previously may provide an evengreater benefit over conical-flow-derived waveriders

(refs. 4 and 5). The aerodynamic performance and PAI

benefits suggested in previous research efforts have gen-

erated interest in using waveriders for various hypersonic

vehicle designs.

The design code used in this study is the (Universityof) Maryland Axisymmetric Waverider Program

(MAXWARP) (refs. 1, 2, and 9). The MAXWARP code

is an inviscid design method that includes an estimate for

skin friction in the design process. Various volumetric

constraints may also be imposed by the user in order to

produce waveriders with desirable structural characteris-

tics and component packaging. These constraints include

aspect ratio, slenderness ratio, and total volume. For

the case of conical-flow-derived waveriders, the Taylor-

Maccoll equation, which describes the flow field behind

a conical shock wave (ref. 10), is integrated using a

fourth-order Runge-Kutta method to compute the invis-cid conical flow field behind the shock wave. The cone

semiapex angle and length of the flow-field generating

body are specified by the user along with free-stream

conditions. The code starts with an initial leading-edgedefinition on the conical shock wave and creates a

waverider shape from this initial leading edge. The pres-

sure distributions on the surface of the configuration areintegrated to calculate lift and drag coefficients. An esti-mate for skin friction is also included so that force coeffi-

cient predictions include both inviscid and viscous

effects. This estimate is based on the reference tempera-ture method, which is described in reference 11. The

effect is to generate shapes for which wetted surface area

is minimized to reduce skin friction drag. The code uses

a simplex optimization routine (ref. 1) to optimize

waveriders for a given figure of merit: maximum lift-

drag ratio or minimum drag. More recent versions of the

code allow the user to construct various other objective

functions. At each iteration in the optimization process,

an updated leading-edge definition is used to generate a

new waverider shape that progresses toward the desired

figure of merit. This process continues over a number of

iterations until the optimum shape is found without viola-tion of any of the user-specified volumetric constraints.

3.2. Waverider Shape Description

The pure waverider shapes used in this study, which

define the forward portions of the waverider-derived

vehicles, were designed using the MAXWARP design

code. Free-stream conditions and optimization parame-

ters were chosen based on the applicability of this studyto a hypersonic cruise vehicle, with available ground-

based test facility limitations taken into account. The

design free-stream Mach number was 4.0 and the designReynolds number was 2.0 x 106 per foot. Although the

specific cruise Mach number for this type of vehicle

would be higher, Mach 4.0 was selected as the designpoint based on the limitations of the UPWT and the

range of data desired. The Mach number range of this

facilityis1.47to4.63.A designpointofMach4.0wouldpermitthevalidationof thewaveriderconceptat thedesignMachnumberandalsoallowfor thedetermina-tionof aerodynamiccharacteristicsat off-designMachnumbers.Theuseof endothermicfuelson thisvehicleclassis expectedto drivetheselectionof cruiseMachnumberto approximately5.0to 5.5.Nosignificantdif-ferencesin theflow physicsareexpectedbetweentheultimatedesignMachnumberandtheMachnumberrangeinvestigatedin thisstudy.TheReynoldsnumberchosenis basedonnominalfacilityoperatingconditionsin theUPWTandisnotrepresentativeof aflightcruisealtitude.Theconfigurationwasoptimizedformaximumlift-dragratioatthedesignpointbecausethisquantityismoreappropriatethanminimumdragasa hypersoniccruiseperformanceparameter.

A fullyturbulentboundarylayerandawalltempera-tureof 585°Rwerespecifiedin thedesign.Thiswalltemperaturewasselectedbasedonpreviousexperimentaldatafrommodelstestedin theUPWT.It isnotlikelythatfully laminarconditionscouldbemaintainedin experi-mentaltestingattheconditionsof interest,andtransitionis difficultto predict.Fullyturbulentconditionscanbeachievedandmaintainedbytheapplicationofboundary-layertransitiongrittothemodelsurface.

Twodifferentpurewaveridershapesweredevelopedfor thisstudy.Thefirst is referredto asthe"straight-wing" shapeandwasdesignedusingtheMAXWARPoptimizationroutine.The second,referredto as the"cranked-wing"shape,wascreatedbyadjustingthelead-ing edgeof the straight-wingwaveriderto createacurvedwingtipshapethathadincreasedaspectratiobutstill maintainedshockattachmentalongtheouterleadingedgeat the designfree-streamcondition.The term"cranked"inthiscontextreferstoawingshapein whichthesweepanglenotonly changesbutalsoexhibitsalargeoutboarddihedralanglein theplaneof thebase.The cranked-wingshapewas designedto provideimprovementsin subsonicaerodynamicperformance(becauseof increasedaspectratio) and in lateral-directionalstability(becauseof dihedraleffect)whilemaintaininghigh performancein the supersonic/hypersonicregime.

Threeprimarydesigncriteriawereusedtoselectthebestwaveridershapedesignsfor thisapplication.First,themaximumlift-dragratiowaschosentobeashighaspossiblewhilenotviolatingotherdesignguidelines.Thiscriteriondrivestheselectionof theconesemiapexanglefor the generatingflow field. A valueof 8.1° wasselectedforthisapplication.Second,thevolumetriceffi-ciency(V2/3/Sref)waschosento beashighaspossible.An inverserelationshipexistsbetweenthevolumetricefficiencyandthemaximumlift-dragratioforagivenset

of free-streamconditions.Therefore,an attemptwasmadeto increasethevolumetricefficiencyasmuchaspossiblewhileacceptinga minimumpenaltyin maxi-mumlift-dragratio.Finally,aconfigurationwithaflatorslightlyconvexbottomsurfacein thecrosssectionwasdesiredfor easein propulsionsystemsintegration.Inadditiontothesethreeprimarydesignguidelines,acon-figurationfreeof substantialcurvatureovermostof thecrosssectionwasalsodesiredtoprovidefortheinclusionof aninternalsparin anactualaircraft.Furthermore,thetargetvalueof span-to-lengthratiowas0.8.Informationfrom previousstudiesshowsthat largerspan(higheraspectratios)waveridersprovidehigherlift-dragratiosbutaremoredifficult to integrateasa full waverider-basedvehicle(ref.12).

A three-viewdrawingandanobliqueviewof thestraight-wingpuretheoreticalwaveridershapegeneratedbythedesigncodeareshownin figure3.Table1sum-marizesthecharacteristicsof thisshape.Thespan-to-lengthratiois 0.83.Thelowersurfaceof thestraight-wingconfigurationhasa slightconvexcurvaturethatfacilitatesintegrationof the propulsionsystem.Thelengthselectedfor the waveriderconfigurationwas24.0in. basedon thesizeof thetest sectionin theUPWT.Thelengthof thegeneratingconewasselectedto fix thelocationof thewaveriderleadingedgeon theconicalshockwaveto achievethedesigncriterianotedpreviously(48.0in. for thisapplication).A selectionofdifferentlocationson theconicalshockwavewouldresultinwaveriderswithmuchdifferentgeometricchar-acteristicsandmayresultin thegenerationof unrealisticshapesthatcouldnotbe integratedinto vehicles.Thevolumetricefficiency,Vef f, of this configuration is 0.11

with a predicted maximum lift-drag ratio of 6.9.

A three-view drawing and an oblique view of the

cranked-wing pure theoretical waverider shape generatedby the design code are shown in figure 4. The cranked

leading edge still lies on the same conical shock wave

produced by the generating cone used to design thestraight-wing waverider. The characteristics of the

cranked-wing waverider shape are summarized in

table 2. The span-to-length ratio is 0.96, which represents

an approximately 16 percent increase in aspect ratio.

This increase in aspect ratio should improve the subsonic

aerodynamic performance over the straight-wing wave-

rider while maintaining the structural characteristics

of the straight-wing waverider near the centerline of

the configuration. The volumetric efficiency of this

configuration is 0.108 with a calculated maximum lift-

drag ratio of 6.7. This configuration represents only a

slight decrease of both parameters from the straight-wingwaverider. The slight convex curvature of the bottomsurface is maintained toward the centerline of the

model. The dihedral angle of the aft cranked section is

approximately28°whenmeasuredfromthecenterlineofthissection.

Thevaluesfor maximumlift-dragratiogivenareforthepurewaveridershapesonly. The waveriders were

subsequently altered to close the blunt base and add con-trol surfaces. The predictions assume that free-stream

static pressure are acting at the base of the unaltered purewaverider shape, so that only forebody drag values are

included in the performance predictions. As will be

shown later, the incorporation of aftbody closure is a sig-nificant issue in hypersonic vehicle development.

3.3. Wind-Tunnel Model Designs

Two slight modifications to the design-code shapes

were implemented in the wind-tunnel model design in

order to accommodate model support hardware and addi-

tional vehicle components. A smooth ogive-cylindrical

fairing was blended on to the upper surface of the pure

waverider shapes to accommodate the sting and balance

necessary to measure the aerodynamic loads on the

model during testing. This volume was added to the

upper surface rather than the lower surface because pre-vious research indicates that modifications to the lower

surface have an affect on the PAI characteristics of the

waverider (ref. 13). Figures 5 and 6 show tunnel installa-

tion photographs of the straight-wing and cranked-wing

pure waverider models with the upper surface fairing.

The lower surface of the theoretical waverider shape was

modified slightly by creating an inboard expansion sur-

face with an angle of approximately 10 °, beginningapproximately 22 in. aft of the nose of the configuration

and measuring approximately 3.5 in. in the spanwisedirection. The lower surfaces follow the waverider theo-

retical stream surface up to this point. This modification

was made in order to facilitate the integration of engine

components and to reduce the closure angle necessary for

control surfaces. Figure 7 shows a photograph of thelower surface of the cranked-wing waverider with the

expansion on the aft end of this surface.

A realistic canopy was designed for the waverider-

based configuration. The canopy was provided with fac-

eted surfaces to resemble the canopy for a hypersonic

vehicle. The aft portion of the canopy was designed to

blend with the cylindrical fairing on the upper surface

discussed previously. Figures 5 and 6 show the pure

waverider models with the ogive-cylindrical fairing

attached (i.e., canopy-off configuration). Figures 8 and 9

show the model with the faceted canopy attached.

The engine package for this configuration included a

compression ramp, a non-flow-through engine module

with side walls, and a nozzle/expansion ramp. The

engine-package-on configuration provided an indication

of the effect of modifying the theoretical waverider lower

surface to integrate some type of engine system and isnot intended to be a realistic propulsion simulation. The

inlet capture area, expansion ramp turning angle, and

nozzle exit area were designed for full-scale Mach 4.0

conditions. The compression surface shown in figure 8 is

required for additional precompression of the flow enter-

ing the inlet. The non-flow-through configuration

attempts to model the external cowl drag present on arealistic flow-through nacelle, but does not have the

associated internal drag. Two different nozzle/expansionramps were designed for the model. The first was used

with the pure waverider configurations with the nacelle

attached and the second was used with configurationsthat had control surfaces attached. These nozzles are

referred to as the "short" and "long" nozzles, respec-

tively (figs. 10(a) and 10(b)). Identical nozzles with static

pressure taps were also fabricated in order to obtain sur-

face pressure measurements on the nozzle. The non-

instrumented ramps were used for force and momentruns.

Control surfaces were provided to examine their

effects on waverider aerodynamic performance as well asthe effectiveness of the control concept. The control sur-faces were sized based on control-volume trends from

supersonic fighter aircraft to extend and close the blunt

base of the configurations. Elevon deflections of 0% pos-

itive 20 ° (trailing edge down), and negative 20 ° (trailing

edge up) were incorporated. A set of outboard ailerons

having the same three deflection angles was designed forthe straight wing. Because of the curved surface of the

cranked wing and the small thickness of the outboard

leading edge, the set of ailerons for the cranked-wing

configuration consisted of an inboard aileron, whichremained fixed at 0 °, and a set of outboard ailerons,which were deflected at 0 ° and +_20°. A vertical tail sur-

face was also designed in order to augment directional

stability. Figures 8 and 9 show photographs of the model

components with the various control surfaces. Figure 11

shows three-view drawings of the elevons, straight-wing

ailerons, cranked-wing inboard ailerons, and cranked-

wing outboard ailerons. This figure indicates the perti-

nent dimensions and shows the hinge-line locations foreach control surface.

The model design allowed for testing of the straight-wing and cranked-wing pure waverider models, which

are defined as configurations with no engine components

or control surfaces. A configuration build up of the

waverider models with different vehicle components

could also be tested up to and including the fully inte-

grated waverider-derived configurations, which are

defined as configurations with engine components, con-

trol surfaces, and the canopy. Table 3 shows the pertinent

model geometry for each configuration tested. Figure 12

5

shows a three-view drawing of the fully integrated

configurations.

4. Experimental Method

The facility used in this study was the UPWT at

NASA Langley Research Center. The UPWT is a closed-

circuit, continuous-flow pressure tunnel with two 4- by

4- by 7-ft test sections, which were both used in this

study. The Mach number range of the facility is 1.47

to 4.63, with a range in the low Mach number test section

of 1.47 to 2.86 and a range in the high Mach number testsection from 2.30 to 4.63. Continuous variation of Mach

number is achieved by using asymmetric sliding block

nozzles to vary the nozzle throat-to-test-section area

ratio. The Reynolds number range of this facility is0.5 x 106 to 8.0 x 106 per foot. However, the nominal

Reynolds number for most tests is 2.0 x 106 per foot.

A detailed description of the UPWT can be found inreference 14.

The configurations tested ranged from the straight

and cranked pure waverider models to the fully inte-

grated waverider-derived vehicles. The test configura-

tions were chosen to show pure waverider performance;to isolate the effects on waverider aerodynamic perfor-

mance of the canopy, engine package, and control sur-faces; and to show the aerodynamic performance and

stability characteristics of the fully integrated configura-tions. Only the cranked-wing configurations were testedin the low Mach number test section. The data were cor-

rected for flow angularity in the test sections. Calibrationdata for the UPWT shows that the flow in both test sec-

tions has an upflow angle generally within 0.5 ° of the

tunnel centerline (ref. 14). In each run, either six-

component force and moment data, nozzle pressure data,

or vapor-screen photographs were obtained. Schlieren

photographs were taken during the force and momentruns.

The test conditions were chosen to investigate the

aerodynamic performance and stability of each configu-

ration at both the design Mach number and at off-designMach numbers. Data were obtained at Mach numbers of

2.3, 4.0, and 4.63 for all configurations studied and, addi-

tionally, at Mach numbers of 3.5 and 4.2 for some con-

figurations. Data for the cranked-wing configurationswere also obtained at Mach numbers of 1.6, 1.8, and 2.0.

The free-stream Reynolds number for most runs was

2.0 x 106 per foot. Some runs were made at Reynolds

numbers of 1.5 x 106 per foot and 3.0 x 106 per foot in

order to investigate the effects at off-design Reynolds

numbers. The angle-of-attack range studied was -6 ° to

10° at fixed sideslip angles of 0 ° and 3°. Data were

obtained over a sideslip angle range of-5 ° to 5 ° for the

first configuration run in each test section in order to ver-

ify that yawing and rolling moment values are linear over

this range (ref. 15). Based on these results, stabilityderivatives were calculated from data obtained at the two

fixed sideslip angles.

The data obtained from the wind-tunnel tests include

six-component force and moment data, static pressure

readings on the blunt base of the model, static pressuredata on the nozzle surfaces, and flow-visualization data.The balance used in this case was the NASA-LaRC-

designated UT-50-B balance, which is a six-componentstrain gauge balance. Unless otherwise noted, the

moment reference center for all configurations was

located 16.623 in. aft of the nose. A total of 11 5-psipressure transducers were used to measure the static

pressure along the blunt base of the configurations and in

the cavity surrounding the sting. Integrated areas were

assigned to each tap or averaged group of taps and used

to calculate the base axial force. All of the force data pre-

sented is corrected to assume free-stream static pressureacting at the base. This procedure is carried out so that

the data may be presented showing only the upper and

lower surface lift and drag values and eliminating the

effect of the blunt base. This procedure is necessary

because the base will be eliminated in any realistic

waverider-derived configuration. The method of assum-

ing free-stream pressure at the base is consistent with the

design-code method and with previous studies showing

predictions for waverider aerodynamic performance

(refs. 2, 9, and 16). Details on the procedure used are

included in reference 15. For configurations with both

engines and control surfaces, only two base and two

chamber pressures were measured. A 32-port, 5-psi

external electronically scanned pressure (ESP) module

was used to measure the static pressure on the nozzlesurface for four runs. Figure 10 shows the locations of

pressure taps on the nozzle surfaces for the short and

long nozzles. Recall that the short nozzle is used with

configurations having no control surfaces and the long

nozzle is used with configurations with control surfaces.

A total of 12 pressure taps were located on the short noz-

zle and 24 pressure taps were located on the long nozzle.

The data are used to correct the nozzle surface pressuresto assume free-stream static pressure acting on these sur-

faces for some configurations.

Schlieren and laser-vapor-screen photographs weretaken in order to examine flow-field features includ-

ing the shock attachment characteristics for various

configurations. For the vapor-screen runs, the laser was

positioned outside of the test section window and the

light sheet was projected across the model surface in the

spanwise direction, illuminating one cross section ata time. The camera was mounted inside of the test sec-

tion above and behind the model. This setup gives a

cross-sectionalviewof thewaveriderflow fieldin thevapor-screenphotographs.

Theaccuracyof theUT-50-Bbalance,basedonaMay1993calibration,is0.5percentof full scaleforeachcomponentto within95-percentconfidence.Thefull-scaleloadlimits were600lbf normal,40 lbf axial,1500in-lbfpitchingmoment,400in-lbfrollingmoment,800in-lbfyawingmoment,and300ibfsideforce.Asanexample,usingthemethodof root-mean-squaressum-mationtocombineindependenterrorsources,theselim-itscorrespondto arangeof uncertaintyin lift coefficientof 0.0053at_ =0° to 0.0054atot= 10° andanuncer-taintyrangein dragcoefficientof 0.00036at o_=0°to0.001ato_= 10° fortheMoo = 4.0 and Reoo = 2.0 × 106

per foot condition. The repeatability of measurementswas observed to be better than these uncertainties. There-

fore, differences less than the indicated ranges for com-

parisons with data from different configurations in thesame test, could be considered significant. However,

comparisons between independent measurements are

only good to within the quoted uncertainty ranges. Tran-

sition grit (no. 60 size sand grit in the low Mach number

tests section and no. 30 size grit in the high Mach number

test section) was applied in a 0.1-in-wide strip to themodel upper and lower surfaces along the outboard lead-

ing edge at a location approximately 0.4 in. from the

leading edge in the streamwise direction. These proce-dures were established for models tested in the UPWT

based on unpublished transition experiments conductedin the UPWT and the methods of references 17 to 19.

5. Computational Method

Computational grids were developed for each of the

pure waverider configurations by first developing a

numerical surface description and then creating 3-Dvolume grids. Numerical surface descriptions of the

straight-wing and cranked-wing wind-tunnel models

were obtained from computer-aided design (CAD)

descriptions of the model parts. Three-dimensional vol-

ume grids were created for each configuration using theGRIDGEN software package, which uses algebraic

transfinite interpolation methods with elliptic interior

point refinement (ref. 20). Only the pure waveridershapes with no integrated vehicle components were mod-

eled for the CFD analysis.

The computational grids for each of the two pure

waverider shapes model only half of the configuration

because each is symmetric about the centerline. The grid

orientation is shown in figure 13. The k-computational

direction runs from the nose of the configuration to the

base in the streamwise direction. The q-computational

direction begins at the upper centerline and wraps around

the leading edge, ending at the lower centerline. The

k-computational direction runs from the surface of the

configuration to the outer boundary. The grids for each

of the two pure waverider shapes contained 91 points in

the _ direction, 111 points in the 11 direction, and

91 points in the _ direction. Blunt leading edges were

modeled for each configuration in order to provide a bet-

ter comparison with experimental data. Grid points were

also clustered near the surface of each configuration in

order to adequately resolve the boundary-layer flow. The

amount of grid spacing needed is judged by examiningthe grid spacing parameter, y÷, which is given by

y+= / pcucA_6/ 0c (1)

where Pc, Uc, and _tc are the density, velocity, and viscos-ity at the first cell center next to the solid surface and A t

is the distance from the first cell center to the body sur-face. Previous research has shown that y÷ values on the

order of I provide accurate solutions (ref. 21).

The CFD solutions were obtained using the General

Aerodynamic Simulation Program (GASP), version 2.2

(refs. 22 and 23). GASP is a finite volume code capable

of solving the full Reynolds-averaged Navier-Stokes

(RANS) equations as well as subsets of these equations,

including the parabolized Navier-Stokes (PNS), thin-

layer Navier-Stokes (TLNS), and Euler equations. Time

integration in GASP is based on the integration of primi-

tive variables, and convergence to a steady-state solution

is obtained by iterating in pseudotime until the L2 norm

of the residual vector has been reduced by a sufficient

amount. GASP also contains several flux-split algo-

rithms and limiters to accelerate convergence to steady

state. Mesh sequencing is available as a means to accel-

erate convergence.

In this study, each configuration was modeled as a

two-zone problem, as illustrated in figure 13. The first

zone includes the blunt nose of the configuration. The

flow in this region is a combination of subsonic and

supersonic flow because a small area of subsonic flowexists behind the detached bow shock. Therefore, the

TLNS equations are solved over the first zone using a

global iteration procedure. The second zone encom-

passes the remainder of the configuration, extendingfrom the zonal boundary to the base of the configuration.

The flow in this region is computed by solving the PNS

equations. These equations are valid for regions ofpredominately supersonic flow with no streamwise

separation. A no-slip boundary condition is applied to all

solid boundaries with a fixed wall temperature of 585°R,

which is identical to that specified in the MAXWARP

optimization routine when designing the waverider

shapes. Free-stream conditions are applied at the outer

boundary, second-order extrapolation from interior cells

is applied at the last streamwise plane, and symmetry

boundary conditions are applied at the center plane. The

Baldwin-Lomax algebraic turbulence model was used in

these solutions to model turbulent boundary layers, and

convergence to a steady state was obtained by reducing

the L2 norm of the residual vector by 5 orders ofmagnitude.

In order to make appropriate comparisons, the condi-tions at which solutions were obtained were chosen

based on conditions at which experimental data were

available. Solutions were obtained at Mach 4.0 at angles

of attack of -6 °, 0 °, 2 °, 4 °, and 8 ° for the straight-wing

model. Solutions were obtained at Mach 4.0 at angles of

attack of-6 °, 0 °, and 8 ° for the cranked-wing model.

Solutions were also obtained at off-design Mach num-

bers of 2.3 and 4.63 at 0 ° angle of attack for each

configuration.

6. Flow-Field and Aerodynamic

Characteristics of Pure Waverider Models

6.1. Flow-Field Characteristics

The flow-field characteristics of the pure waverider

models at the design Mach number can be illustrated byexamining computational solutions of each configuration

and laser vapor-screen photographs from wind-tunnel

tests. Figure 14 shows a laser vapor-screen photograph ofthe flow at the base of the pure straight-wing waverider

model and nondimensional static pressure contours at thebase of the same configuration from a CFD solution at

Mach 4.0, 0 ° angle of attack, and free-stream Reynolds

number of 2.0 x 106 per foot. The model lower surface is

highlighted in the photograph by the laser light sheet on

the surface. The bow shock is indicated by the contrast

between light and dark regions below the light sheet. On

the left-hand side of the photograph, the shock is

observed to be very near the edge of the lower surface.

Thus, the vapor-screen photograph confirms the qualita-tive shock location predicted by the CFD solution. A

small detachment distance exists even at the design point

caused by blunt leading edge and boundary-layerdisplacement effects. These effects are not accounted for

in the design code. The CFD predictions also indicate

that the high-pressure region remains mostly confined

below the model lower surface. A large low-pressure

region (P/P** of 0.95 or less) exists near the centerline ofthe model below the bottom surface because of the bot-

tom surface expansion present on the model. However,the remainder of the bottom surface flow field is a

smooth, conical flow field, so the presence of this slight

expansion surface does not degrade the favorable PAl

characteristics offered by the waverider. Engine modules

would be placed upstream of the point where the expan-

sion surface begins, so the flow entering the inlet would

be highly compressed. Similar data are shown in fig-ure 15 for the cranked-wing pure waverider model. The

shock can be seen in the right-hand side of the photo-

graph to be very near the outer leading edge of the

model. The lower surface is again highlighted by the

laser light. The full cross-sectional view is not shown

because of the poor quality of the photographs. The

experimental data confirm the qualitative shock location

at the outer leading edge, which is predicted by the CFD

solution for this case as well. Figure 16 further illustrates

that the shock is slightly detached at the outer leading

edge for both models. This figure shows a close-up view

of the outer leading edge at the base of the cranked-wing

and straight-wing waverider shapes from CFD solutions

at Mach 4.0 and 0 ° angle of attack. Both of the views

in figure 16 are to the same length scale, and non-

dimensional static pressure contours are shown in eachview.

The flow-field characteristics of each pure waverider

shape at off-design Mach numbers can also be illustratedby examining experimental flow-visualization data and

CFD solutions. Figure 17 shows a comparison of a

vapor-screen photograph and a CFD solution for the

cranked-wing shape at Mach 2.3 and 0 ° angle of attack.The free-stream Reynolds number is 2.0 × 106 per foot.

The data shown in this figure and orientation of thecamera in the test section are the same as in figures 14

and 15. At Mach numbers below the design Mach num-ber of 4.0, the shock-wave angle is larger and the detach-

ment distance should be much larger than at the design

Mach number. This outcome is predicted by the CFD

solution and confirmed by the experimental data. Fig-

ure 18 shows similar views of the same configuration at

Mach 4.63. The photograph in this figure was taken with

the laser light sheet approximately 5 in. upstream of the

base because the quality of the photograph taken with thelight sheet at the base was poor. At Mach numbers

greater than the design Mach number, the shock moves

closer to the leading edge than at the design condition, as

illustrated in both the vapor-screen photograph and pre-

dicted by the CFD solution. A large high-pressure region

still exists in the bottom-surface flow field of this config-

uration at Mach 4.63. The qualitative shock locations can

be further illustrated by examining planform schlieren

photographs of the cranked-wing model. Figure 19

shows schlieren photographs of the cranked-wing pure

waverider model in a planform view at Mach 2.3 (top),

Mach 4.0 (middle), and Mach 4.63 (bottom). The right

side of the figure shows a close-up view near the leading

edge at each Mach number. The schlieren images in this

figure have been enhanced by computer imaging tech-

niques in order to show the shock structure more clearly.

At Mach 2.3, the schlieren photograph shows that the

shockis detachedfromtheleadingedge.Theoutermostshockin the top view representsthe bowshock.At

Mach 4.0, the shock is much closer to the outer leadingedge, but a small detachment distance still exists. At

Mach 4.63, the photograph does not show the presence of

a shock wave near the leading edge, possibly because theshock is attached at this condition.

6.2. Aerodynamic Performance

The aerodynamic performance characteristics of the

two pure waverider models are examined here usingexperimental force and moment data and computational

predictions. Off-design Mach number and Reynolds

number effects are evaluated using experimental data.

The longitudinal and lateral-directional stability charac-

teristics are also examined for each configuration using

experimental data. Unless otherwise indicated, all of the

experimental and computational data presented have

been corrected to a condition of free-stream pressure act-

ing at the blunt bases of the configurations, as previouslydiscussed.

The aerodynamic performance of the straight-wing

and cranked-wing pure waverider shapes at the design

Mach number is shown in figures 20 and 21. These fig-ures show experimental data, CFD predictions, and

design-code predictions for the lift, drag, and lift-drag

ratios of each configuration at Mach 4.0 and a Reynolds

number of 2.0 x l06 per foot. The computational values

were obtained by integrating surface pressure and skinfriction predictions from CFD solutions. Because the

data are corrected to eliminate the base drag, these data

should be interpreted as the performance of the forward

portion (or forebody) of a possible hypersonic configura-tion and not that of a realistic hypersonic vehicle. In gen-

eral, agreement is good between the experimental data

and computational predictions. Both the computationalpredictions and experimental data show lower lift and

higher drag values than the predicted design-code values,and these differences can be attributed to several causes.The flow-visualization data and CFD flow-field solutions

showed that a slight detachment distance exists at the

outer leading edge even at the design condition, which

results in a lift loss and a drag decrease. However, the

design code assumes an infinitely sharp leading edgewith a perfectly attached shock wave. An additional lift

loss results from the expansion ramp on the bottom sur-face of the waverider, and an increase in drag results

from the additional volume added to the upper surface of

the model. The experimental data also show that the

maximum lift-drag ratio occurs near 2 ° angle of attack

for each configuration. This finding is also consistent

with previous studies, such as those in references 13

and 16, which show that the maximum lift-drag ratio

occurs at an angle of attack greater than 0 ° for the wave-

rider configurations studied in these references.

A direct comparison of the experimental aero-

dynamic performance of the two pure waverider models

is shown in figure 22. The experimental data show that

the cranked-wing shape has a slightly higher maximum

lift-drag ratio than the straight-wing shape. At positive

angles of attack, the straight-wing shape produced

slightly higher lift coefficients. Aside from these obser-

vations, there are no significant differences between the

two configurations.

The off-design performances of the straight-wing

pure and cranked-wing pure waverider models are shownin figures 23 and 24, respectively. Each of these figures

shows the experimental lift, drag, and lift-drag ratio at all

Mach numbers studied as well as maximum lift-dragratio versus Mach number. The data indicate that there is

no significant performance degradation at off-design

Mach numbers. Both configurations show higher maxi-mum lift-drag ratios than the design point value at Mach

numbers less than 4.0, using the assumption of free-

stream pressure acting at the base. Similar results have

been found in previous waverider studies (refs. 13,

16, and 24) and are also typical for non-waverider

supersonic/hypersonic configurations. The cranked-

wing waverider shape provides better aerodynamic per-

formance at Mach numbers of 4.0 and below. At higherMach numbers, there are no significant differences

between the performance of the two configurations.

The effects of Reynolds number on aerodynamic

performance of the straight-wing and cranked-wing con-figurations are shown in figures 25 and 26, respectively.

No significant effects of Reynolds number variation were

observed for either configuration in the range studied,

except for a slight increase in maximum lift-drag ratio atthe 3.0 x 106-per-foot condition for both configurations.

This result is most likely because the skin friction coeffi-

cient decreases as Reynolds number increases, resultingin decreased drag and thus increased lift-drag ratios at

higher Reynolds numbers. The decrease in drag observed

experimentally is approximately equal to the decrease in

viscous drag predicted by the reference temperature

method (ref. 11). Computational solutions at Mach 4.0and a Reynolds number of 2.0 x 106 per foot show that

the viscous drag contribution is approximately 34 percent

of total drag. By comparison, the MAXWARP design

code predicts a viscous contribution of approximately

38 percent to the total drag.

The pitching-moment characteristics of the straight-

wing and cranked-wing pure waverider configurations

are shown in figure 27. This figure shows the pitching-

moment coefficient versus angle of attack at each Machnumber studied. Both configurations are longitudinally

9

unstableat all Mach numbers studied. The moment refer-

ence center location here is an arbitrarily selected

location at the approximate location of the center of grav-

ity of the fully integrated model. This moment referencecenter location (16.623 in. aft of the nose) is used for all

configurations studied unless otherwise stated. The

cranked-wing pitching moment curve is more nonlinear

than that for the straight-wing shape, indicating that theshock may be detached at higher angles of attack for the

cranked-wing configuration. The yawing moment char-

acteristics are shown in figure 28. This figure shows the

yawing moment derivative versus angle of attack at each

Mach number studied for both configurations. The

straight-wing configuration is directionally unstable at all

Mach numbers studied at angles of attack of 8 ° and

below. The cranked-wing configuration is directionallystable at all Mach numbers studied above an angle of

attack of 4 °. Both configurations experience a destabiliz-

ing effect as Mach number increases. The cranked-wingconfiguration was expected to provide improved direc-

tional stability from the increased dihedral along the out-

board leading edge. The rolling moment characteristics

are shown in figure 29 for each configuration. The

cranked-wing waverider shows better lateral stabilitycharacteristics than the straight-wing model. The

cranked-wing configuration exhibits positive effective

dihedral above 0 ° angle of attack at all Mach numbers.

The straight-wing model is unstable at angles of attackbelow 6.0 ° at Mach numbers of 4.0 and 4.63 and is unsta-

ble at angles of attack below 4° at a Mach number of 2.3.

7. Component Build-Up Effects

tion when the faceted canopy is used. Similarly, a

5.1-percent reduction in maximum lift-drag ratio occurs

for the cranked-wing configuration. The data indicate

that a penalty was incurred for the canopy, and therefore

attention should be paid to the canopy design in a hyper-sonic waverider-based vehicle.

7.2. Effect of Engine Package

The engine component effects are evaluated by com-paring experimental data from engine-on and engine-off

configurations. Figures 32 and 33 show the effects of

adding the engine package (ramp, inlet, and nozzle com-

ponents) to the straight-wing and cranked-wing configu-rations, respectively. The data shown here are for

configurations with the canopy and no control surfaces.

The data are corrected to assume free-stream static pres-

sure acting at the base. No correction is applied to these

data for the nozzle surface pressures. Each figure showslift and drag coefficients as well as lift-drag ratios at

Mach 4.0 and the maximum lift-drag ratio at comparative

Mach numbers for engine-on and engine-off configura-

tions. The addition of engine components results in a

slight increase in lift and a significant increase in drag atMach 4.0. These effects are caused by the inlet compres-

sion surface and the increase in projected frontal area and

produce a decrease in lift-drag ratio at positive values of

lift and a reduction in maximum lift-drag ratio over the

Mach number range studied. The straight-wing engine-on configuration shows a 19.7-percent reduction in the

maximum lift-drag ratio at Mach 4.0 over the engine-off configuration. The cranked-wing model shows a

17.7-percent reduction at the same condition.

7.1. Effect of Canopy

The effects of adding the canopy on the aerodynamic

performance of the pure straight-wing and cranked-wing

waverider models are illustrated in figures 30 and 31,

respectively. These data were obtained for configurations

that have no control surfaces or engine componentsattached, and the data are corrected to assume free-

stream static pressure acting at the base. Each figure

shows the lift and drag coefficients as well as lift-drag

ratios at Mach 4.0 and the maximum lift-drag ratio at

each comparative Mach number studied for the canopy-

off and faceted-canopy configurations. The canopy-off

configurations have the ogive-cylindrical fairing on the

upper surface, as discussed previously. Both the straight-wing and cranked-wing configurations show little differ-

ence in lift when the canopy is added. The canopy-on

configurations show slightly higher drag than those with

no canopy and an accompanying decrease in lift-dragratios at positive values of lift over the Mach number

range studied. The maximum lift-drag ratio at Mach 4.0

is reduced by 3.6 percent for the straight-wing configura-

7.3. Effect of Control Surface Addition

The effects of adding undetected control surfaces

are illustrated by comparing data for configurations withno control surfaces to those with undetected ailerons and

elevons attached. Each configuration includes the canopy

and engine components. Data for both the straight-wingand cranked-wing configurations are shown. The coeffi-

cient data are reduced by the planform areas of each cor-

responding configuration so the effects of increased

planform are accounted for in the normalization of these

data. The plots showing drag and lift-drag data includethree separate data sets. The first is the data for the

controls-off configuration corrected to assume free-

stream pressure at the base. Therefore, only forebodydrag values are included in these data and base drag isnot included. The second data set is the controls-on data

and the third set is the controls-off data with base dragincluded (i.e., uncorrected data from wind-tunnel mea-

surements), so that these data include the effect of the

blunt base. A comparison between the second and third

data sets shows the aerodynamic effect of adding control

10

surfacestotheconfiguration,andacomparisonbetweenthefirst twodatasetsshowstherelativeperformancebetweentheclosedconfigurationsandthatof thefore-bodysurfaceonlywithouttheeffectof thebluntbase.

Theeffectof addingundeflectedcontrolsurfacestostraight-wingwaveriderconfigurationwith thecanopyandenginecomponentsattachedis summarizedin fig-ure34.Theadditionof controlsurfacescausesa slightdecreasein lift coefficientatMach4.0.Thisdecreaseispartiallycausedby the largeexpansionanglethat ispresenton theelevonlowersurfacesanda 16-percentincreasein referenceareafor thecontrols-onconfigura-tion.A comparisonof thecontrols-offdatawithbasedragandthecontrols-ondatashowsadecreaseindragatagivenlift-coefficientvalue.Thereisaslightincreaseinlift-dragratiosat low positiveanglesof attackandanincreasein maximumlift-dragratioswhen0° controlsurfaceswereaddedto theconfiguration.However,acomparisonof thecontrols-ondatawiththecontrols-offdatawithnobasedragshowsthattheclosedconfigura-tionhassignificantlyhigherdragvaluesandlowermaxi-mumlift-dragratiosthantheforebody-onlyvalues.Thisresultindicatesthattheinclusionof aftbodyclosurepre-sentsa significantchallengein theintegrationof purewaveridershapesintohypersonicvehiclesandthatthisaspectof theconfigurationdeservesspecialconsider-ationin thedesignprocess.It is likely thatthelift-dragratiosof aclosedconfigurationcannotapproachthoseofpurewaveridershapesbecausetheeffectof basedragisoftennotincludedin lift-dragvaluesfortheseconfigura-tions.Theeffectsof controlsurfaceadditionaresimilarfor thecranked-wingconfigurationasindicatedin fig-ure35. Forreference,the baseareais approximately8.3percentof theplanformareafor thestraight-wingmodelwith no control surfacesand approximately9.1percentforthecranked-wingmodel.

Thecontrolsurfacedesignfortheconfigurationusedin thisstudywasasomewhatarbitrarydesignbasedonlyon trendsfromvarioussupersonicfighteraircraft.Amoreoptimumdesigncouldminimizetheperformancedegradationcausedby theclosureof thebluntbase.Aperformanceimprovementcouldbeobtainedbyinclud-ingtheaftbodyclosurein thedesign/optimizationpro-cess.Previousstudieshaveexaminedthepossibilityofusingblunttrailingedgesoncontrolsurfacesasameansof enhancingthe aerodynamicperformance(refs.25to27).Thebluntbasereducesthestrengthof thebaserecompressionshockandproperdesignof thetrailingedgecanresultin an increasein basepressureandadecreasein drag.A controlsurfacedesignthattakesadvantageof theseeffectswouldenhancethe aero-dynamic performanceof the configuration.Thisenhancementcouldbe accomplishedby reducingthethicknessof thebasebymaintainingthelowersurfaceas

a waverider stream surface all the way to the base while

designing the upper surface as an expansion surface.

Longer control surfaces would also reduce the closure

angle and enhance the pitch control power of theconfiguration.

8. Characteristics of Fully-Integrated

Waverider-Derived Hypersonic Cruise

Configurations

8.1. Aerodynamic Performance

Aerodynamic characteristics of each of the fullyintegrated waverider-derived configurations are exam-

ined over the Mach number range using experimental

data, and the performance of these configurations are

compared to that of the pure waverider shapes. The

fully integrated configurations are defined here to have

the canopy, the engine components, the undeflected aile-rons, the undeflected elevons, and the vertical tail

attached. The aerodynamic characteristics of the straight-

wing and cranked-wing configurations are presented first

followed by comparisons to the corresponding pure

waverider configuration.

The aerodynamic performance of the straight-wing

and cranked-wing waverider-derived hypersonic cruise

configurations are shown in figures 36 and 37, respec-

tively. The data presented here have the nozzle surface

pressures corrected to assume free-stream pressure acting

on the nozzle surface. The data are presented using this

method to show the aerodynamic characteristics without

any propulsive effect on the nozzle surface. The force

data were corrected by assigning integration areas to

each pressure tap measurement and computing the cor-

rected coefficients. The locations of pressure taps areshown in figure 10. The straight-wing configuration has

a maximum lift-drag ratio of 4.69 at Mach 4.0 and the

cranked-wing configuration has a value of 4.56 when the

nozzle surface pressures are corrected to free-stream

pressure. The aerodynamic performance of each configu-

ration does not vary significantly at off-design Mach

numbers. The maximum lift-drag ratio for each configu-

ration also occurs near 2 ° angle of attack at Mach 4.0.

The angle of attack for maximum lift-drag ratio increasesas Mach number decreases. At Mach numbers of 2.0 and

below, the maximum lift-drag ratios for the cranked-

wing configuration do not follow the general trend of

increasing maximum lift-drag ratio with decreasing

Mach number. This situation results from lift curve slopevalues that show similar inconsistencies at Mach num-

bers less than 2.3.

A direct comparison of the straight-wing and

cranked-wing fully integrated vehicles is shown in fig-

ure 38. The straight-wing configuration produces slightly

11

highervaluesof maximumlift-dragratio than thecranked-wingconfigurationatMachnumbersof 2.3andhigher.Thestraight-wingmodelalsoshowshigherliftcoefficientvaluesatMach4.0.Thestraight-wingmodelshowsa maximumlift-dragratio thatis 3.0 percenthigherthanthatofthecranked-wingconfigurationatthedesignMachnumberof4.0.

Comparisonsof theaerodynamicsof thestraight-wingpurewaveridermodelandthefullyintegratedcon-figurationareshownin figure39.Thisfigureshowsliftanddragcoefficientsaswellaslift-dragratiosatMach4.0andthemaximumlift-dragratiosateachMachnum-berstudied.Asin figures34and35,thesedatasetsarepresentedforcomparisonin thedragandlift-dragplots.Thefirst datasetrepresentsthepurewaveridershapewithnobasedragincluded.Thesecondrepresentsthewaveridershapewithbasedrag,andthethirdrepresentsthefully integratedconfiguration.Thenozzlesurfacepressuresarecorrectedtoassumefree-streampressureonthenozzlesurfacefor thefully integratedvehicles.Acomparisonof thepurewaveriderdatawith basedragandthefully integrateddatashowsthattheaerodynamicperformanceof thepurewaveridershapeis degradedwhenallof thevariousvehiclecomponentsareadded.Areductionin lift coefficientfor thefully integratedcon-figurationis observedat Mach4.0above0° angleofattack,whichincreasesasangleof attackincreases.Anincreasein dragis observedwhenall componentsareintegratedwiththepurewaveridermodel.Theseeffectsresultinadecreasein lift-dragratiosatMach4.0andinmaximumlift-dragratiosatcomparativeMachnumbersof 4.0andabove.At Mach2.3,thereisaslightincreaseinmaximumlift-dragratiowhenall vehiclecomponentsareadded.Thisincreaseismostlikelycausedbythenoz-zlesurfacepressurecorrectionto free-streampressure.Thefree-streamstaticpressureincreasesasMachnum-berdecreases.However,theaerodynamicperformanceof thefully integratedvehicleis significantlydegradedfromthatofthepurewaveridershapeonlywithnobasedragincluded,becauseof thedragproducedbythecon-trol surfaceaddition.The maximumlift-dragratioatMach4.0for thefully integratedvehicleis 4.69,com-paredto6.68forthepurewaveridershape.

A comparisonof thefully integratedcranked-wingconfigurationand the pure cranked-wingwaveridermodelyieldsconclusionssimilartothoseofthecompari-sonof thestraight-wingconfigurations.Figure40showsthe aerodynamicperformanceof the cranked-wingwaveriderforebodyandthecranked-wingfully inte-gratedconfiguration.Theadditionof vehiclecompo-nentscausesa slightdegradationin the aerodynamicperformance,but the lift-drag ratios observed for the

fully integrated model are significantly lower than those

for the pure waverider shape only with no base drag. The

12

maximum lift-drag ratio at Mach 4.0 for the fully inte-

grated configuration is 4.56, compared to a value of 6.72

for the fully integrated vehicle.

From these results, it can be concluded that the max-

imum lift-drag ratios of a fully integrated waverider-

derived configuration with aftbody closure likely cannot

approach those of pure waverider shapes. Theoretical

predictions for waverider configurations do not include

the effects of aftbody closure. However, it will be shown

that the fully integrated waverider-derived configurations

studied here are comparable in aerodynamic performance

to previously tested hypersonic models with performance

improvements possible through enhanced control surfaceand propulsion system designs.

In order to characterize the lift-drag values of the

configurations studied here, a comparison is made

between data for the present cranked-wing fully inte-

grated waverider-derived configuration and experimentaldata from six hypersonic vehicle wind-tunnel models

previously tested in NASA Langley facilities (refs. 28 to

33) in figure 41. Although direct comparisons of these

data are not possible here because of different conditions,

geometries, levels of volumetric efficiencies, and force

accounting methodologies, a range of values can be

obtained to compare with the data from the current study.

As shown in figure 41, the waverider falls within the

same general range of lift-drag values as the non-

waverider hypersonic configurations. The lift-drag ratios

of the waverider configurations studied could be

improved significantly through a better design of the pro-

pulsion system and control surface closure. Therefore,

the waverider configurations studied here offer at least

comparable aerodynamic performance and perhaps a

modest advantage over conventional non-waveriderhypersonic vehicles.

8.2. Longitudinal Control Effectiveness and Trim

Both fully integrated configurations are longitudi-

nally unstable at each Mach number studied. The

pitching-moment coefficient data as a function of angle

of attack at each Mach number studied are shown in fig-ure 42. Data for the straight-wing and cranked-wing fully

integrated waverider-derived hypersonic cruise configu-rations are shown. The moment reference center is

located at 62.5 percent of the centerline chord. At higherangles of attack, the cranked-wing configuration shows a

destabilizing increase in the pitching moment curve. This

increase indicates that the shock may have detached from

the leading edge of the outer cranked portion of the wing

at higher angles of attack. The longitudinal instability of

these configurations may be addressed in one of two

ways. First, it may be possible to shift the center-of-gravity location for a fully integrated flight vehicle to a

locationthatwouldprovideatleastneutralstabilityoverthe Machnumberrange.Recommendationsfor suchlocationsarepresentedlaterin thissection.Second,theadditionof afully functioningpropulsionsystemwouldenhancethelongitudinalinstabilitybyincreasingtheaft-bodylowersurfacepressures.

Thepitchcontroleffectivenessof theelevonsanddevon/aileroncombinationfor thestraight-wingconfig-urationisshownin figure43.Dataareshownforthreetrimsettings.Thefirst isonewithboththeelevonsandaileronsat 0°, thesecondwith a positive20° elevondeflection(_E) and a 0° aileron deflection (_A), and thethird with both elevons and ailerons deflected at 20 °. The

effectiveness of the elevon decreases as Mach number

increases, as evidenced by the smaller increments in lift

and pitching-moment coefficients produced by eachdeflection. The ailerons were more effective than the

elevons in pitch control because of the shadowing of theelevon behind the thick wing shape and the location of

the elevon in an expansion flow field. The CFD flowfield solutions showed that the bottom surface flow field

expands to pressure below free-stream pressure in the

region where the elevons are placed. Also, the closure

angle for the elevon was severe because of the thick base

of the waverider. Each aileron has only 70 percent of the

planform area of the elevon but at higher angles of attack

generates substantially more pitching moment. These

characteristics may be unacceptable and indicate that the

pitch control concept should be redesigned.

The pitch control effectiveness of the elevons for the

cranked-wing configuration is shown in figure 44. Each

figure shows data for 0°, 20 °, and -20 ° elevon deflec-tions with 0 ° ailerons. No runs were made with both aile-

rons and elevons deflected at the same angle because of

the shape of the trailing edge for the cranked-wing con-

figuration. The elevon pitch control power for this con-figuration also decreases as Mach number increases.

However, in contrast to the straight-wing pitch control

data, the cranked-wing pitching moment curves are non-

linear. This factor makes the elevon pitch control power

even more critical for this configuration than for the

straight-wing vehicle. These data indicate that the nose-

down pitch control power of this configuration is not suf-

ficient. Either symmetric ailerons must be used to pro-

vide additional pitch control power or the elevon areashould be increased.

Because of the combination of unstable pitching

moment characteristics and low pitch control power

observed in the experimental data, the configurations

should be balanced such that they are at least neutrally

stable to ensure adequate pitch control power throughout

the angle-of-attack range. For a realistic full-scale flightvehicle, it should be possible to control the center-of-

gravity location through packaging. Also, it may be pos-

sible to control the shift in static margin from subsonic to

supersonic speeds using fuel transfer. Neutral stability

can be achieved by placing the center of gravity at a loca-

tion equal to 58 percent of the centerline chord for the

fully integrated straight-wing configuration and 59 per-

cent of the centerline chord for the cranked-wing config-

uration. Data for lift and pitching-moment coefficients

referenced to these center-of-gravity locations are shown

in figure 45 for the straight-wing vehicle and in figure 46

for the cranked-wing vehicle. In figure 45, the data for

the trailing-edge-up elevon deflections were extrapolated

from the cranked-wing data and applied to the straight-

wing configuration. Also note that all of the data pre-

sented here are for unpowered conditions. The addition

of a functioning propulsion system will enhance the lon-

gitudinal stability of the vehicle even further. These data

are presented only to indicate the effects of an alternative

choice of center-of-gravity locations. Subsequent data

are presented at the original moment reference centerlocation of 62.5 percent of the centerline chord.

8.3. Lateral-Directional Stability and ControlEffectiveness

The lateral-directional stability of the straight-wing

and cranked-wing hypersonic cruise vehicles are shown

in figures 47 and 48, respectively. Each figure shows

yawing and rolling moment derivatives at each Mach

number studied. Both configurations are directionally

stable at all Mach numbers investigated, with the

cranked-wing model providing higher stability levels

than the straight-wing model. The cranked-wing fullyintegrated configuration is laterally stable across the

angle-of-attack range at all Mach numbers studied. The

straight-wing fully integrated configuration is laterallyunstable at angles of attack below 6 ° (at Mach 4.0). This

roll instability may be caused in part by the high place-ment of the balance in the model. No transfer distance in

the vertical direction was applied to the moment refer-

ence center in the presentation of data.

Figures 49 and 50 show the effect of the vertical tail

on yawing moment derivative and rolling moment deriv-

ative values for each configuration. The effect of the ver-

tical tail is to significantly enhance the directional

stability of both the straight-wing and cranked-wing con-

figurations, indicated by the large positive shift in yaw-ing moment derivatives when the vertical tail is added toeach model. No rudder control effectiveness runs were

made in this study, so it is not clear whether sufficient

yaw control power exists to augment stability. The addi-

tion of the vertical tail does not cause any significant

change in the lateral stability characteristics of either

configuration.

The effectiveness of a 20 ° aileron deflection on the

straight-wing configuration is shown in figure 51. A 20 °

13

ailerondeflectionindicatedhereimpliesoneaileronwitha20° trailing-edge-down deflection and the other with a

20 ° trailing-edge-up deflection. The elevons remained

fixed at 0 ° for these runs. Figure 51 shows rolling

moment and yawing moment increments between the

deflected and nondeflected runs. Additionally, the

DATCOM computer code was used to estimate the

steady state roll rates for this configuration (ref. 34).

Table 4 shows the steady roll rate capabilities as pre-

dicted by this method. The roll rate is shown as deg/sec

of roll, normalized by flight velocity. For most vehicles

of this type, excess roll-control power is available at

lower angles of attack. The requirements for pitch androll control surfaces for the waverider-derived vehicles

may be driven by low-speed flying qualities. These qual-

ifies include roll-rate capabilities at subsonic speeds andcrosswind landing requirements.

Figure 52 shows the effectiveness of the ailerons for

the cranked-wing fully integrated configuration. How-

ever, a significant difference exists between these results

and those for the straight-wing configuration. The

cranked-wing ailerons produce considerably moreadverse yaw at 0 ° angle of attack than the straight-wing

configuration, as evidenced by the large negative values

of AC t. The adverse yaw produced by the cranked-wing

ailerons will further drive the control power requirementsof the rudder.

Figure 53 shows the aileron effectiveness on lateral-

directional stability with the ailerons deflected at 20 ° for

the cranked-wing fully integrated configuration. Rolling

moment and yawing moment increments for a positive20 ° elevon deflection and a +_20° aileron deflection are

shown. A comparison of these data shows that a 20 °

elevon deflection has no effect on roll control power,

indicating that interaction between controls is minimal atthe Mach numbers studied here.

9. Concluding Remarks

The aerodynamic performance and stability and con-trol characteristics of two Mach 4.0 waverider-derived

hypersonic cruise configurations were examined. Experi-mental force, moment, and flow-visualization data were

obtained for the two Mach 4.0 waverider configurations

in both test sections of the Langley Unitary Plan Wind

Tunnel (UPWT). The wind-tunnel models were designed

to allow testing of various configurations ranging from

pure waveriders to fully integrated vehicles. The two

pure waverider shapes were referred to as the straight-

wing pure and the cranked-wing pure waveriders. Exper-

imental data as well as limited computational solutions

were used to examine the flow field and aerodynamic

characteristics of the two pure waverider shapes, the

component build-up effects, and the aerodynamic and

14

controllability characteristics of the fully integratedhypersonic cruise vehicles.

The flow-field characteristics and aerodynamic per-

formance of the two pure waverider shapes were

examined using experimental and computational data.

Computational fluid dynamics (CFD) predictions and

laser vapor-screen photographs of the straight-wing and

cranked-wing pure waverider models confirmed theshock attachment/detachment characteristics of each

configuration. The shock was slightly detached from the

outer leading edge at the design Mach number of 4.0

and 0 ° angle of attack. This detachment distance exists

because of boundary-layer displacement effects as well

as blunt leading-edge effects. The design code assumesan infinitely sharp leading edge and does not account for

the physical presence of a boundary layer. Comparisons

between experimental force data and CFD predictions

were generally good. The maximum lift-drag ratios

observed experimentally were lower than the design-

code predictions, as expected. These lower lift-dragratios were caused by a loss of lift and an increase in drag

caused by the shock not being perfectly attached as well

as to loss of lift from the lower-surface expansion and an

increase in drag from the additional volume added to the

upper surface to accommodate model support hardware.

The maximum lift-drag ratio for each configuration

occurs at an angle of attack above 0 °. Both the CFD pre-

dictions and experimental data showed that there were no

significant performance degradations at off-design Mach

numbers. The cranked-wing pure waverider model

exhibited slightly better aerodynamic performance at the

comparative Mach numbers studied than the straight-wing model.

Component build-up effects of waverider-derived

vehicles were examined by comparing experimental

force and moment data. The primary effect of individu-

ally adding the canopy and the engine package was to

increase the drag of the configuration, thereby resulting

in a degradation in aerodynamic performance. The aero-

dynamic effect of adding control surfaces was to increase

the maximum lift-drag ratios slightly at each Mach

number studied. However, the aerodynamic performance

of the controls-on configurations was significantly

degraded from that of the pure waverider shape only by

the addition of aftbody closure and the associated drag

production. The values presented for the pure waverider

model show the performance of the waverider surface

only and do not include base drag. These results indicate

that additional consideration should be applied to

the design of control surfaces and aftbody closure in

waverider-based hypersonic cruise configurations. A

control surface configuration with a less severe closure

angle or controls with blunt trailing edges may result in

improved performance. Inclusion of the aftbody closure

in the optimization process for the waverider shape may

also improve the performance significantly.

The characteristics of the fully integrated waverider-

derived hypersonic cruise vehicles were also examined

by comparisons of experimental force and moment data.

The aerodynamic performance of each fully integrated

waverider model (straight-wing and cranked-wing con-

figuration) was significantly degraded from that of the

pure waverider shapes, because of the inclusion of aft-

body closure in the fully integrated configuration. The

straight-wing fully integrated configuration provided

slightly better aerodynamic performance than the

cranked-wing fully integrated model. The maximum lift-

drag ratios at Mach 4.0 were 4.69 for the straight-wing

model and 4.56 for the cranked-wing model. The wave-

rider concept also provides a uniform compressed flow

field to the inlet, which offers potential advantages for

airbreathing propulsion systems integration. The use of

different generating flow fields, such as osculating-cone

and cone-wedge flow fields, may further improve these

characteristics. Furthermore, the results of this study

have identified areas where design improvements could

enhance performance, such as control surfaces, aftbody

closure, and propulsion system design.

Both fully integrated vehicles are longitudinally

unstable across the Mach number range studied for

unpowered conditions with the selected reference

moment center. Additionally, locations were recom-

mended for placement of the center of gravity in each

configuration in order to ensure at least neutral stability

across the Mach number range. The pitch-control effec-

tiveness of the elevons was judged to be unacceptable for

both configurations, and the data indicate that the pitch

control concept should be redesigned. The ailerons were

significantly more effective than the elevons for pitch

control. The cranked-wing vehicle shows significantly

better lateral-directional stability than the straight-wing

vehicle. The straight-wing configuration was unstable at

angles of attack below 6 ° at Mach 4.0. The vertical tail

has a significant stabilizing effect on directional stability,

but very little effect on lateral stability. The ailerons are

also highly effective for the cranked-wing vehicle, but

produce a significant amount of adverse yaw.

NASA Langley Research Center

Hampton, VA 23681-0001

May 6, 1996

References

1. Bowcutt, K. G.; Anderson, J. D.; and Capriotti, D.: Viscous

Optimized Hypersonic Waveriders. AIAA-87-0272, Jan. 1987.

2.

3.

4.

7,

8.

9.

10.

11.

12.

13.

14.

15.

16.

Corda, Stephen; and Anderson, John D., Jr.: Viscous Opti-

mized Hypersonic Waveriders Designed From Axisymmetric

Flow Fields. AIAA-88-0369, Jan. 1988.

Eggers, A. J., Jr.; Ashley, Holt; Springer, George S.; Bowles,

Jeffrey V.; and Ardema, Mark D.: Hypersonic Waverider Con-

figurations From the 1950's to the 1990's. Proceedings of the

1st International Hypersonic Waverider Symposium, John D.

Anderson, Jr., Mark J. Lewis, Stephen Corda, and Isaiah M.

Blankson, eds., Univ. of Maryland, 1990.

Sobieczky, H.; Dougherty, E C.; and Jones, K.: Hypersonic

Waverider Design From Given Shock Waves. Proceedings of

the 1st International Hypersonic Waverider Symposium,

John D. Anderson, Jr., Mark J. Lewis, Stephen Corda, and

Isaiah M. Blankson, eds., Univ. of Maryland, 1990.

Takashima, Naruhisa; and Lewis, Mark J.: Waverider Configu-

rations Based on Non-Axisymmetric Flow Fields for Engine-

Airframe Integration. AIAA-94-0380, Jan. 1994.

Rasmussen, Maurice L.: Waverider Configurations Derived

From Inclined Circular and Elliptic Cones. J. Spacecr &

Rockets, vol. 17, no. 6, Nov.-Dec. 1980, pp. 537-545.

Kuchemann, D.: The Aerodynamic Design of Aircraft. Perga-

mon Press, Inc., 1978.

O'Neill, Mary K. L.; and Lewis, Mark J.: Optimized Scramjet

Integration on a Waverider. J. Aircr., vol. 29, no. 6, Nov.-Dec.

1992, pp. 1114-1121.

Corda, Stephen; and Seifert, E. Scott: User Information for

Maryland Axisymmetric Wave Rider Program (MAXIWARP).

Univ. of Maryland, Jan. 1989.

Anderson, John D., Jr.: Modern Compressible Flow--With

Historical Perspective. McGraw-Hill, Inc., 1982.

Anderson, John D., Jr.: Hypersonic and High Temperature Gas

Dynamics. McGraw-Hill, Inc., 1989.

Beuerlein, David L.: Optimization of Waveriders to Maximize

Mission Performance. Proceedings of the 1st International

Hypersonic Waverider Symposium, John D. Anderson, Jr.,

Mark J. Lewis, Stephen Corda, and Isaiah M. Blankson, eds.,

Univ. of Maryland, 1990.

Cockrell, Charles E., Jr.: Interpretation of Waverider Perfor-

mance Data Using Computational Fluid Dynamics. AIAA-

93-2921, July 1993.

Jackson, C. M., Jr.; Corlett, W. A.; and Monta, W. J.: Descrip-

tion and Calibration of the Langley Unitary Plan Wind Tunnel.

NASA TP-1905, 1981.

Cockrell, Charles Edward, Jr.: Vehicle Integration Effects on

Hypersonic Waveriders. NASA TM-109739, 1994.

Bauer, Steven X. S.; Covell, Peter F.; Forrest, Dana K.; and

McGrath, Brian E.: Analysis of Two Viscous Optimized

Waveriders. Proceedings of the 1st International Hypersonic

Waverider Symposium, John D. Anderson, Jr., Mark J. Lewis,

Stephen Corda, and Isaiah M. Blankson, eds., Univ. of

Maryland, 1990.

15

17. Braslow, A. L.; and Knox, Eugene C.: Simplified Method for

Determination of Critical Height of Distributed Roughness

Particles for Boundary-Layer Transition at Mach Numbers

From 0 to 5. NACA TN-4363, Sept. 1958.

18. Braslow, A. L.; Hams, R. V., Jr.; and Hicks, R. M.: Use of

Grit-Type Boundary-Layer-Transition Trips on Wind-Tunnel

Models. NASA TN D-3579, 1966.

19. Stallings, Robert L., Jr.; and Lamb, Milton: Effects of Rough-

ness Size on the Position of Boundary-Layer Transition and on

the Aerodynamic Characteristics of a 55 Degree Swept Delta

Wing at Supersonic Speeds. NASA TP-i 027, 1977.

20. Steinbrenner, John P.; and Chawner, John R.: The GRIDGEN

Version 8 Multiple Block Grid Generation Software. MDA

Engineering Report 92-01, 1992.

21. Richardson, Pamela F.; McClinton, Charles R.; Bittner, Robert

D.; Diiley, A. Douglas; and Edwards, Kelvin W.: Hypersonic

CFD Applications for the National Aero-Space Plane. SAE

892310, Sept. 1989.

22. McGrory, W. D.; Huebner, L. D.; Slack, D. C.; and Waiters,

R.W.: Development and Application of GASP 2.0. AIAA-

92-5067, Dec. 1992.

23. McGrory, William D.; Slack, David C.; Applebaum,

Michael P.; and Waiters, Robert W.: GASP Version 2.2--The

General Aerodynamic Simulation Program. AeroSoft, Inc.,1993.

24. Takashima, Naruhisa: Navier-Stokes Computations of a Vis-

cous Optimized Waverider. NASA CR- 189658, 1992.

25. Bushnell, D. M.: Supersonic Aircraft Drag Reduction. AIAA-

90-1596, June 1990.

26. Chapman, Dean R.: Reduction of Profile Drag at Supersonic

Velocities by the Use of Airfoil Sections Having a Blunt

Trailing Edge. NACA TN-3503, 1955. (Supersedes NACA

RM A9H11.)

27.

28.

29.

30.

31.

32.

33.

34.

Dutton, J.; Herrin, J.; Molezzi, M.; Mathur, T.; and Smith, K.:

Some Problems in Transonic Aerodynamics. AIAA-95-0476,Jan. 1995.

Penland, Jim A.; Hallissy, James B.; and Dillon, James L.:

Aerodynamic Characteristics of a Hypersonic Research Air-

plane Concept Having a 70 ° Swept Double-Delta Wing at

Mach Numbers From 0.80 to 1.20, With Summary of Data

From 0.20 to 6.0. NASA TP-1552, 1979.

Dillon, James L.; and Pittman, Jimmy L.: Aerodynamic Char-

acteristics at Mach 6 of a Wing-Body Concept for a Hyper-

sonic Research Airplane. NASA TP- ! 249, 1978.

Penland, Jim A.; Edwards, Clyde L. W.; Witcofski, Robert D.;

and Marcum, Don C., Jr.: Comparative Aerodynamic Study of

Two Hypersonic Cruise Aircrafi Configurations Derived From

Trade-OffStudies. NASA TM X-1436, 1967.

Small, William J.; Kirk.ham, Frank S.; and Fetterman,

David E.: Aerodynamic Characteristics of a Hypersonic

Transport Configuration at Mach 6.86. NASA TN-D-5885,

1970.

Watts, Joe D.; Olinger, Frank V.; Weidner, John P.; Johnson,

Stuart K.; Sanders, Bobby W.; and Keyes, J. Wayne: Mach 5

Cruise Aircraft Research. Proceedings of the Langley Sympo-

sium on Aerodynamics, Volume 2, Sharon H. Stack, Compiler,

NASA CP-2398, 1986, pp. 285-304.

Ellison, James C.: Investigation of the Aerodynamic Charac-

teristics of a Hypersonic Transport Model at Mach Numbers

to 6. NASA TN D-6191, 1971.

Williams, John E.; and Vukelich, Steven R.: The USAF

Stability and Control Digital Datcom. Volume 1: User's Man-

ual. AFFDL-TR-79-3032-VOL-I, Apr. 1979. (Supersedes

AFFDL-TR-73-23, AFFDL-TR-74-68, and AFFDL-TR-76-

45-VOL-1.)

16

Table 1. Characteristics of Straight-Wing Waverider Designed by MAXWARP

Waverider length, in .......................................... 24.0

Span/length ................................................. 0.83

Base height/length .......................................... 0.092

Volumetric efficiency (Veff) ................................... 0.112

Planform area, Sre f, ft 2 ........................................ 1.89

Predicted maximum LID ....................................... 6.9

Base area, ft 2 .............................................. 0.136

Table 2. Characteristics of Cranked-Wing Waverider Designed by MAXWARP

Waverider length, in .......................................... 24.0

Span/length ................................................. 0.96

Base height/length .......................................... 0.092

Volumetric efficiency (Veff) ................................... 0.108

Planform area, Sre f, ft 2 ........................................ 2.05

Predicted maximum LID ....................................... 6.7

Base area, ft 2 .............................................. 0.153

Table 3. Reference Quantities for Various Configurations

Length, in. Base Xc.g.,

Configuration Sre f, ft 2 Span, in. _ area, ft 2 percent of _"

Straight-wing pure model 1.894 19.80 24.0 0.1580 69.3

Straight-wing pure model with engine 1.894 19.80 24.0 0.1481 69.3components

Straight-wing fully integrated model 2.202 19.80 26.60 0.0194 62.5 a

Cranked-wing pure model 2.052 23.016 24.0 0.1860 69.3

Cranked-wing pure model with engine 2.052 23.016 24.0 0.1745 69.3components

Cranked-wing fully integrated model 2.346 23.016 26.60 0.0194 62.5 b

aFor some data: 58.0.bFor some data: 59.0.

Table 4. Steady-Roll-Rate Capabilities Calculated From

DATCOM for Straight-Wing Fully

Integrated Configuration

Mach number Pss per unit velocity,

2.3 0.119

4.0 0.095

4.63 0.095

deg/secft/sec

17

ZConicalshockwave

II X

Waverider leading edge

Bottom surface (stream surface)

Figure 1. Design of conical-flow-derived waverider.

18

16

14

12

10E

8

6

4

2

L/Dma x = 4(M + 3)/M "L/D Barrier"Conical-flow-derived waveriders

.... I , , , , I , , , , I .... I , , , i I , , , , I

5 10 15 20 25 30Mach number

Figure 2. Comparison of L/Dma x values of conventional vehicles and waveriders.

18

........_z_,,-i:!ii;ii?iii

_' ,,i _i,_ i__iii :iiiii,iii!iiiii!iiiii!ii_iiiiiiiiiii_iiii_iiii_iiii_iiiiiiiiiiiiiiii_......

; i_ii_i!J_ii_ii!ii_ii!ii_i!_!_i]iiii!iiii!ii!iiii!iiii_i_iiiiiiiiiiii!i!ii_i_i_ii_i!;!!_!_i_i'

, i i i il,ii ii iiiiiiii!iiii!iiiiiiiiiiiiiiiii !iiiiii iii:iiii i!!!!!! ....../

iiiiiiiiiiiiiiililiiii!iii_ili_i!iiiii!ii_iiiii!_ ¸_

_ _ _ _i,_iiiii,i_iiiii!_iiiiiiii!ili!!!ii;_,_"

,_: _ iiili ii_,i_iiiiiiiiiii]'i!_ii__y'_

ii i iiiiii!_[_i_ ¸y

Plan form Oblique

Profile Rear

Figure 3. Straight-wing pure waverider shape designed by MAXWARP.

j_

.... i

, : 9:

. ¢ :: 3,_ -

_ _;_._ i_ -:_

.::.. !: ;_:_::i:,i_:::::_ '_

Planforrn Oblique

Profile

.............................._i_!iiill_ii!iii!!!_ii!_iii_ii_iiiii_i_i_!iiii!i_iiii!iiiiii_i_!iiill!i!iiiii_i!_ii_iii_:,.....................Rear

Figure 4. Cranked-wing pure waverider shape designed by MAXWARP.

19

Figure5. Straight-wingpurewaveridermodelinUPWT.

2O

E_

.E

0

E

c_

o_

°_

21

_-- Expansionsurface

Figure7. Lowersurfaceofcranked-wingpurewaveridermodel.

22

0

c_

\

0E

p_'r_

c_

LT_

23

0

E0

0

0

E

N._

t:m

o .o

24

1.971"

Top view

I

I

I

I

I

I

I

I

I

l

I

B.L. 0.0

Tap No. Model Sta. B.L. W.L.1 22.030 -1.353 -1.785

2 22.030 -1.015 -1.788

3 22.030 -0.676 -1.792

4 22.030 -0.338 -1.796

5 22.791 -1.353 -1.244

6 22.791 -1.015 -1.252

7 22.791 -0.676 -1.254

8 22.791 -0.338 -1.255

9 23.552 -1.353 -0.964

10 23.552 -1.015 -0.974

11 23.552 -0.676 -0.974

12 23.552 -0.338 -0.975

Short ramp

(No-controls configurations)

M.S. M.S.

21.779" 23.750"

--_ 1.386"

t

Side view

3.5"

W.L. 0.0

Taps 9-12--...._ o o o ool

Taps 5-8_ o ooo1

°

Rear view

Pressure taplocations on nozzle

surface

(a) Short expansion ramp used with no-controls configurations.

Figure 10. Three-view drawings of expansion ramps.

25

B.L. 0.0

W.L. 0.0

Top view

Tap No. Model Sta. B.L. W.L.13 24.313 -1.353 -0.724

14 24.313 -1.015 -0.73115 24.313 -0.676 -0.734

16 24.313 -0.338 -0.745

17 25.075 -1.353 -0.516

18 25.075 -1.015 -0.523

19 25.075 -0.676 -0.532

20 25.075 -0.338 -0.73921 25.836 -1.353 -0.332

22 25.836 -1.015 -0.339

23 25.836 -0.676 -0.451

24 25.836 -0.338 -0.739

Locations for taps 1 to 12 are given on

previous page.

Long ramp

(Fully integrated configurations)

M.S. M.S.

21.779 26.597

4_ 4.818" _

.... - 2.01

Sting sl

Side view

_-- 3.5"

Taps 21-22 _ I

Taps 17-19_ _o----'_'_-- --_ W.L. 0.0Taps 13-16 oOoO_ [

Taps 9-12 _ o °oo °oi _ Taps 20,

_°°°°i I 23,24Taps 5-8 I_ [ located on

Taps 1-4 _ ' B.L. 0.0 sting sleeve

x____ Pressure tap locationson nozzle surface

Rear view

(b) Long expansion ramp used with fully integrated configurations.

Figure 10. Concluded.

26

Hinge line

.22'z---_

lnboard side

_'-_ __._ \ /---Trailing edge\ " _t

(a) Elevons.

U _

----3.760"-------,- 10 o

0.588-| "_

_ "_Trailing edge

(b) Straight-wing ailerons.

27__ edge railing edge

3

(c) Cranked-wing inboard ailerons. (d) Cranked-wing outboard ailerons.

Figure 11. Dimensions of elevons and ailerons.

27

Cranked//

/ wing

11.5"

Cranked wing

Canopy --

Moment ref. center

62.5 percent ofmodel length

9.9 _l

Straight wing

1Cylindrical upppersurface fairing

Vertical tail

i2.81,,

/"

/"

Waverider /stream surface _jr

Compression ,/surface

L 0.87" radius

Elevon surface

","_ Sidewalls -- Aileron surface

""--- Lower surface of engine module

Figure 12. Three-view drawing of fully integrated waverider model.

\ Nozzlesurface

' Straightwing

28

PNS marching zone

7 TLNS globaliteration zone

Figure 13. Coordinates and computational scheme for waverider CFD solutions.

29

Shock

(a) Vapor-screenphotographofbase.

Lowersurfaceofmodel

Baseofconfiguration

\

P/Poo

1.71

1.62

1.54

1.46

1.37

1.29

_ 1.20

1.12

1.03

0.95

(b) Base view of CFD solution.

Figure 14. Comparison of base-view vapor-screen photograph and CFD nondimensional static pressure contours ofstraight-wing pure waverider model at M = 4.0 and cx = 0 °.

30

Shock

(a) Vapor-screenphotographofbase.

Lowersurfaceofmodel

Baseofconfiguration

i_iii_iiiiiiiiii_!iii!_i!ii_

i:i:i;!iii!ii:iiii!ilii i

P/P_

1.61

1.54

1.47

1.39

1.32

1.24

1.17

1.10

1.02

0.95


Figure 15. Comparison of base-view vapor-screen photograph and CFD nondimensional static pressure contours ofpure cranked-wing waverider model at M = 4.0 and _x= 0%

31

(a)

/Cranked-wing pure waverider model.

P/P,,_

2.18

2.05

1.92

1.79

1.66

1.53

1.39

1.26

1.13

1.00

\

P/Po,,

2.18

2.05

1.92

1.79

1.66

1.53

1.39

1.26

1.13

1.00

(b) Straight-wing pure waverider model.

Figure 16. Comparison of CFD nondimensional static pressure contours near leading edge at base of cranked-wing and

straight-wing pure waverider models at M = 4.0 and ot = 0°.

32

Sho(

Upper surfaceof model

(a) Vapor-screen photograph at base.

Base ofconfiguration

\,\\

iii!iiiiiiii_iiii_iii_i_ii

P/Poo

1.28

1.24

1.21

1.17

1.13

1.10

1.06

1.02

0.99

0.95


Figure 17. Comparison of base-view vapor-screen photograph and CFD nondimensional static pressure contours ofcranked-wing pure waverider model at M = 2.3 and o_ = 0 °.

33

Shock

Uppersurfaceof model

(a) Vapor-screen photograph 5 in. upstream of base.

Base ofconfiguration

--\\

P/Poo

1.71

1.62

1.54

i .46

1.37

1.29

1.20

1.12

1.03

0.95


Figure 18. Comparison of base-view vapor-screen photograph and CFD nondimensional static pressure contours ofcranked-wing pure waverider model at M = 4.63 and o_= 0 °.

34

Shock

wave

Figure 19. Comparison of planform schlieren photographs of cranked-wing pure waverider model at M = 2.3, 4.0,and 4.63.

35

.30

.25

.20

.15

.10

.058

0

-.05

-.10

-.15

C)

C) Experimental

dictions

C] MAXWARP prediction

_.208...,,,,,,,,,,,,,,,,...,,.,,_.,,.,.,,,,,,- -6 -4 -2 0 2 4 6 8 10 12

.07

.06

.05O.re-

._ .04U

§ .03

g .02

.01

6

.2

4

2

0

-2

-4

-6

-8-.10

Q) Experimental (3

CFD predictions(3

[] MAXWARP prediction

0 I I I I [ I I I I [ .... [ .... [ .... [ .... I .... [ .....

-. 10 -.05 0 .05 .10 .15 .20 .25 .30

Lift coefficient, C L

8

[]6

ental

_ _ CFD predictions

/_) [] MAXWARP prediction

,¢i i i i I i i i i [ .... I , i . . I .... I .... [ , i i . I .... I

-.05 0 .05 .10 .15 .20 .25 .30


Figure 20. Comparison of experimental data, CFD predictions, and design-code predictions for aerodynamic perfor-mance of straight-wing pure waverider model at M = 4.0 and Reynolds number of 2.0 x 106 per foot.

36

.30

.25

.20

.15

.10=

¢.J

.05

0

-.05

-.10

-.15

-.20-8

.07

.06

.05

,..r

_= .04

§ .03

_ .02

.01

013

,.-1

1

_ CFD predictions

[] MAXWARP prediction

-6 -4 -2 0 2 4 6 8 10 12

-2

-4

-6

-8-.10

C) Experimentaldl&

I, CFD predictions O

[] MAXWARP predictionO

.... I , , J , I , , , t I , , , , I .... I L t _ i I i i t _ t j j L i I0-.10 -.05 0 .05 .10 .15 .20 .25 .30


8

6

4

2

0

rimental

_ CFD predictions

/ [] MAXWARP prediction

.... I .... I .... I .... I .... I .... t .... I .... I

-.05 0 .05 . I 0 .15 .20 .25 .30


Figure 21. Comparison of experimental data, CFD predictions, and design-code predictions for aerodynamic perfor-

mance of cranked-wing pure waverider model at M = 4.0 and Reynolds number of 2.0 x 106 per foot.

37

.30

.25

.20,..a.15•" .10e-,

.05o

-.05

-.10

-.15

-.20

.07

.O6

.05r,.)..r¢..,

.N .04

8 .03

.02

#,..1

[] Straight wing []

OCran_OZ)(__

-6 --4 -2 0 2 4 6 8 10 12

.01

0-.10

4

2

0

-2

-4

-6

-8-.10

[] Straight wing

O Cranked wing

[]

©[]

Q[]

i L , , i .... i .... i .... i .... i .... i .... t .... I

-.05 0 .05 .10 .15 .20 .25 .30


O_U_)_) _)_)%

[] _t_

0

[]

0 [] Straight wing

[] 0 Cranked wing

O[]

.... I .... 1 .... i .... I .... I .... I .... 1 .... I

-.05 0 .05 .10 .15 .20 .25 .30


Figure 22. Comparison of aerodynamic performance of straight-wing and cranked-wing pure configurations at M = 4.06

and Reynolds number of 2.0 x 10 per foot.

38

¢o.3

ca

"3

ca

.4

.3

.2

.1

0

-.1

-.2

DD

D

> %

j D Mach 2.3

<3 Mach 3.5

Mach 4.0

* Mach 4.2

D © Mach 4.63

i , ,_l , , I, ,,I , ,L[t _IILL_IIL ' I ' ' ,ll , ,I , , , I-4 -2 0 2 4 6 8 10 12


10

8

6

4,.d6 2

•_ o

,=-2

"7 -4

45

-8

-10

10.0

_> t> l> D 9.5

__ 9.0t>E>D 8.5

80<> D Mach 2.3 _ 7.5

< Mach 3.5 "_7.0

<_ _ Mach 4.0

<_ + Mach 4.2©Mach 4.63

DD

, , , I n i , I h J , I _ , , l , _ , I , * ,

-. 1 0 . 1 .2 .3 .4


O

©

6.5 © 0

6.0 ©

5.5

, J , i i i L i I , i _ i i

5"02.0'' '2'.5''' 3.0 '3t.5 ' 4.0 '415' '5'.0

Mach number, Moo

Figure 23. Aerodynamic performance of straight-wing pure waverider configuration across Mach number rangestudied.

39

,.d

e-

Q

,.d

.4

.3

.2

.1

0

--.1

--.2

[]

[z []_>

[]

[]b

[] b

b

[] Mach 1.6

_ Mach 2.3

<_ Mach 3.5

_ Mach 4.0

+ Mach 4.2

Mach 4.63

8 10 1245 -4 -2 0 2 4 6

..Jr_

8et0t_

.07

.06

.05

.04

.03

13 Mach 1.6

b Mach 2.3

<1 Mach 3.5

Mach 4.0

* Mach 4.2

© Mach 4.63 b

.02

.01

[]

[]

0.2 -.1 6 .! .L .3 .;Lift coefficient, C L

6

14

12

10

8

6

4

2

0

-2

-4

-6

-8

-10

-12

-14-2

[] []

1_ Ib _> b U [][] []

bb

J[] []

_- O Mach 4.63

-. 1 0 . l .2 .3 .4

[] Mach 1.6

D Mach 2.3

<1 Mach 3.5

Mach 4.0

+ Mach 4.2


E

t_

12.0

11.5

11.0

10.5

10.0

9.5

9.0

8.5

8.0

7.5

7.0

6.5

6.0

5.5

O

OO0

0

5°1_5'''2'.6'''2'.5'' '3'.6"'315' :_Y i5 .... 5'.0Mach number, Moo

Figure 24. Aerodynamic performance of cranked-wing pure waverider configuration across Mach number range

studied.

4O

,.d

"_ .10e-

.05

8 o

,d -.05

.30

.25 (_(_

.20

.15 (_

__1. 65XlO per foot

(_(_ [] 2.0 x 106 per foot0 3.0 × 106 per foot-.10

-.15

f_-'2_-8 _ -4 -2 0 2 4 6 8 I0 12

ot

.07

/k 1.5 × 106 per foot t_

.06 [] 2.0 x 106 per foot 0r_

0 3.0 × 106 per foot 0

0

0

d

02.01

.... I .... i .... I .... I .... I .... I , . , _ i L L i i I0-.10 -.05 0 .05 .10 .15 .20 .25 .30


10

8

64

2 0

0

0 /_ 1.5 x 106 per foot-2

eq [] 2.0 × 106 per foot

-4 0 0 3.0 x 106 per foot

-8

--lO , , , I .... I .... I .... I . , . t I .... I .... I .... I

-.10 -.05 0 .20 .25 .30

.05

e.,

._ .04

8 .o3

6

-?

,.d

.05 .10 .15


Figure 25. Effects of Reynolds number on aerodynamic performance of straight-wing pure waverider configuration atM= 4.0.

41

.30

.25

.20

,2 .15

.10=.o

.05

0.2 -.05

-.10

-.15

-.20

.07

.06

.05

,,d

._ .04

_ .03

.02

.01

6

,.-1

©00

0

00

O O /X 1.5 x 106 per foot

{_ [] 2.0 x 106 per foot

O _ O 0 3.0 x 106 per foot

ik*llhllLiLl,,,I,,,I,,,l*,,I,,,I,*,l,,,I

-6 -4 -2 0 2 4 6 8 10 12

0-.10

10

8

6

4

2

0

-2

-4

-6

-8

-10-.!0

/_ 1.5 x 106 per foot

[] 2.0 x 106 per foot

O 3.0 x 106 per foot

_ OO#OOOO

, , , , i .... i .... I .... i .... i .... i .... i .... I

-.05 0 .05 .10 .15 .20 .25 .30


O t_¢_f_t _

{_ /X 1.5 x 106 per foot

[] 2.0 x 106 per foot

0 3.0 x 10 6 per foot

.... I .... I .... I .... I .... I .... I .... i .... i

-.05 0 .05 .10 .15 .20 .25 .30


Figure 26. Effects of Reynolds number on aerodynamic performance of cranked-wing pure waverider configuration at

M= 4.0.

42

L)

O

_0e-

.030

.025

.020

.015

.010

.005

0

-.005

-.010-8

Mach 2.3

<3 Mach 3.5

0 Mach 4.0

* Mach 4.2

© Mach 4.63

[>

_>

D,O

b b <s _oO

AJ

-6 -4 -2 0 2 4 6 8 10 12

(a) Straight wing.

0

¢.J

,fiEO

.030

.025

.020

.015

.010

.005

0

-.005

-.010

[] Mach 1.6[]

t> Mach 2.3 []

Mach 3.5[] t>

0 Mach 4.0 [] b<_

÷ Mach 4.2 [] I> .q (_

0 Mach 4.63 [] t>

°; ©

[] P

b

E>

q

_,,I,,,l,,_l,_,l,,,I,J,l_l,,,l,,,]tll I

-6 -4 -2 0 2 4 6 8 10 12

(b) Cranked wing.

Figure 27. Pitching moment characteristics of pure waverider configurations.

43

.0004

.0003

.0002

.0001.r-u

0

E

,_ -.0001

'_ -.0002

-.0003

--.000_

-6

Stable b

O> _ _ _ _a_.3_ Mach 4.0

b B0 Mach 4.63

-4 -2 0 2 4 6 8 10 12

(a) Straight wing.

.0008

.0006

r,.),; .0004

_" .0002

oEd_

.- -.0002

--.0004

[]

b[]

[] 8lStab,o [] _ e

b [] Mach 1.6

Mach 2.3

Mach 4.0

0 Mach 4.63

L , , I , , _ I _ , , I . , , I , , , I . _ i I e A i I e _ , I , _ , I

-4 -2 0 2 4 6 8 10 12

a

(b) Cranked wing.

Figure 28. Yawing moment characteristics of pure waverider configurations.

44

.0016

.0012e,_

d .0008

_ .0004

E 0

._ -.00O4

-.0008

-.0012

D" Mach 2.3

_ Mach 4.0

_(_ O Mach 4.63

g> % o

0 0P 0 0

I Stable O

D

_'I'J'I''LI'''I'''IllLiIIII_I_I

-2 0 2 4 6 8 10 12

(a) Straight wing.

.0008

.0004

e_

_- oo"

._ -.0004

-r"-.0008

e-,

-.0012o

_-.0016

-6 -.0020e¢

-.0024

_ _ _ 8 8 i Stable

SE>

[]

[][] Mach 1.6

Mach 2.3 []

Mach 4.0 []

0 Mach 4.63

P

-2 0 2 4 6 8 10 12

t_

(b) Cranked wing.

Figure 29. Rolling moment characteristics of pure waverider configurations.

45

.30

.25

.20

,..1 .15

,.z .10e-

.05_ o._ -.05

-.10

-.15

-'20-8 -6 -4 -2 0 2 4 6 8 10 12

.07 Moo = 4.0O

.06 []

0[]

0[]

[]o

_ (2) Canopy off.01 [] Canopy on

o_o........................................I -.05 0 .05 .10 .15 .20 .25 .30

r_ .05r_

.__ .04ej

8 .o3

.02


6

,.d

8

6

4

2

0

-2

-4

-6

-8-.10

Moo = 4.0

0 8.5

[] _ 8.0

o S 75E

7.0

06.5

[]

_S_IS)DO O Canopy off 6.0[] Canopy on 5.5

.... I .... I .... i .... I .... I .... i .... i .... I

-.05 0 .05 .10 .15 .20 .25 .30


10.0

9.5

9.0

0 Canopy off

[] Canopy on

8

0D 0 o

[] []@

5.020.............................• 2.5 3.0 3.5 4.0 4.5 5.0

Mach number, Moo

Figure 30. Effect of canopy on aerodynamic performance of straight-wing pure waverider configuration.

46

.30

.25

.20

..a .15¢0

.10e-

.05

8 0

..a -.05

-.10

-.15

-.20

Mo_ = 4.0

O

O0 °

0 ©

O0 °0

0

O0 O00 0 Canopy offO

[] Canopy on

' '6"S "-'2...................- 0 '2'' 4 6 8 'l_O'' 1_2

.07Moo = 4.0

.06

{3

0

©

©

©0 0 Canopy off

/DO ° [] Canopy onmO 1

°_ ............. , .... , .... , ...............20.l -.05 0 .05 .10 15 .20 .25 .30


.05¢.),.te-,

._ .04

_ .03

_ .02

.16

i

,-.1

2

0

-2

-4

O

ID

M = 4.0

O0 0

©

© Canopy off

[] Canopy on

--8 .... I .... I .... i .... I .... I .... I I I L I I J _ L , I

-.10 -.05 0 .05 .10 .15 .20 .25 .30


E

E

14

13

12

11

12

10

9

8

7

6

O Canopy off

[] Canopy on

D

23

5L5....................... ,...........2.0 2.5 3.0 3.5 4 0 4.5 5.0

Mach number, M

Figure 31. Effect of canopy on aerodynamic performance of cranked-wing pure waverider configuration.

47

.30

.25

.20

,..z .15

"-: .10ID

.05o

,.-: -.05

-.10

-.15

-.20-8

Moo = 4.0

u- [] Engine on

-6 -4 -2 0 2 4 6 8 10 12

.07

.06

.05r,.).,..r

.04

8 .03

.02

.01

0-.10

Moo = 4.0

%%

IODdz3ODO%% _%E3Ofi_D[_IZ_ O Engineoff[] Engine on

-.05 0 .05 .10 .15 .20 .25 .30


6

,.d

8

6

4

2

0

-2

-4

-6

-8-.10

Moo = 4.0

0000 Or-,,Drl oo [] ,--,

[] _c_D o

(D°

[]

4P

%[]

[] O O Engine off

0 0 [] Engine on

- i _ , i i .... I .... I .... t _ i _ i I i , i i I i i i i i .... I

-.05 0 .05 .10 .15 .20 .25 .30


E

10

O Engine off

[] Engine on

O

OOo

[] 0

[]Iz] []

[]

4 .... ] , , , t t L , , i i .... J .... i .... J

2.0 2.5 3.0 3.5 4.0 4.5 5.0

Mach number, M

Figure 32. Effect of adding engine package on aerodynamics of straight-wing pure waverider model with canopy.

48

.30

.25

.20

,.d .15

.10

"U.05

o-.05

-.10

-.15

-.20

Moo = 4.0

O_3_3

O _ 0 Engine off0

[] Engine on

+tI'''I'''JL'+I'+,IL,,I,++I,,,L,,,]+ 0 ,,I-6 -4 -2 0 2 4 6 8 12

.07Moo = 4.0

.05

e-,.04

_28 .03

_ .02

D

.06 ©[]

DO

m©

DO

m©

m©DO O Engine off

5_ [] DO [] Engine

.01 (_ (_ (_ (_(]_£0 0 on

0 , , , , I , , J , I , L , , I , , , , I .... I _ + L k l i i i i I i L i _ _

-.!0 -.05 0 .05 .10 .15 .20 .25 .30


d

8 M°° = 4.0

6 00_00

4 % [] [] [] D(_ (_ (_(_ (_

2

0 []

©-2

[]

-4 []O

D_ 0 Engine off[]

-60 w [] Engine on

-8 + , l .... i .... i + + , , [ i , , , i .... i , , i t [ i i , i i

-.10 -.05 0 .05 .10 .15 .20 .25 .30


10

_ 8

B_ 7

m6

Q

[]

O

0 Engine off

[] Engine on

[] OOo

0

[] [][]

4 , t, , I .... I .... _ i + i i I L i + + t,, + t I,,,, I

1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0

Mach number, M

Figure 33. Effect of adding engine package on aerodynamics of cranked-wing pure waverider model with canopy.

49

.30

.25

.20

.15L)'_ .10

"U.05

o,2 -.05

-.10

-.15

Moo = 4.0.07

O

OAO/x .06

0 ¢_OA _ .05L)

¢-_ .04

8 .03

No controls (no base drag) _ .02zx_ /_ 0 ° controls

[] No controls (base drag) .01

-6 -4 -2 0 2 4 6 8 10 12

Moo = 4.0

O No controls (no base drag) AS]L_ 0 ° controls

[] No controls (base drag) _3O

_yO

_0

_0

_y_ _t_O _0

_O--d 0 0 _

o_o........ .... ..........................1 -.05 . 5 .10 .15 .20 .25 .30


6

,.d

M = 4.08

6

4

2

0

-2

-4

-6

-8-.10

O 0

(Z 0

/k

Z_O0 0 No controls (no base drag)/x 0 o controls

[] No controls (base drag)

-.05 0 .05 .10 .15 .20 .25 .30

Li_ coefficient, C L

E

9.0

8.5

8.0

7.5

7.0

6.5

6.0

5.5

5.0

4.5

4.02.0

0 No controls (no base drag)/k 0 ° controls


O

O

O

2.5 3.0 3.5 4.0 4.5 5.0

Mach number, Moo

Figure 34. Effect of undeflected control surface addition on aerodynamics of straight-wing pure waverider configura-

tion with canopy and engine package attached,

5O

.3O

.25

.20

,.d .15

'_ .10

"U.05

_ 0,_ -.05

-.10

-.15

-.20

Mo_ = 4.0

O @

Q

_@@_q u._='_.05.04

@ @ Q _ .03

O No controls (no base drag)A 0ocontrols _ .02


-6 -4 -2 0 2 4 6 8 10 12

.07 Moo = 4.0 []

.06 O No controls (no base drag) _ O

/_ 0 ° controls _O[] No controls (base drag)

6o

65o45o

E_ _Dso_°

°-.1o'-.;5; .............................05 .1o _5 So .25 .3oLift coefficient, C L

,.dd

i.

..d

8 Mo_ = 4.0

6

,-., 0000,_

4 o__K:s_t_ @

o _9

-2 []_

[]/x 0 No controls (no base drag)-4 _ O z_ 0 °controls

-6 O O [] No controls (base drag)

-8 _ , . i .... ] .... i , . , I i i L , , i .... i .... [ i i i i [

-.10 -.05 0 .05 .10 .15 .20 .25 .30


.--i

8.5

8.0

7.5

7.0

6.5

6.0

5.5

5.0

4.5

4.0

O

0 No controls (no base drag)/_ 0 ° controls


O

O

O

/x _ A

3.5 t , , , I , * , , I .... I , i i i I i i j L I i i L _ I _ i _ i I

1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0

Mach number, Moo

Figure 35. Effect of undeflected control surface addition on aerodynamics of cranked-wing pure waverider configura-

tion with canopy and engine package attached.

51

.30

.25

.20

,-.I .15

"_" .10CA

.058

0

,_ -.05

-.10

-.15

-.20-I

D Moo = 2.3

M =4.0oo

O Moo = 4.63 i> D _8

D

-6 -4 -2 0 2 4 6 8 10 12

.09

.08

.07

.06

-_ .05

_ .04

.03

.02

D Moo = 2.3

Moo = 4.0

O Moo-- 4.63

.01 --v_

0 + + _ I i + + I i L + I h _ I I

-.2 -.1 0 .1 .2


I

.3

8 D Moo = 2.3

Moo = 4.06

O Moo = 4.63

_Q

-6

-8 i i h i , +

-.2 -.1

4

26

'_ 0ex(_

-2

,.d-4

O>

@

0 .1


I + i i I

.2 .3

8.0

7.5

7.0

6.5

6.0

E= 5.5

s.0

4.5

00 0

4.0

3.5

3"02 d ' ' '2'.5.... 3t.O .................... 3.5 4.0 4.5 5.0

Mach number, Moo

Figure 36. Aerodynamics of fully integrated straight-wing configuration.

52

,..J

E

,d

.4

.3

.2

.1

0

--.2

--.3

[] M = 1.6

/x Moo = 1.8 .[3

M:20 V

l> M =23 &_ I_ _,_

O M =40 A_i7 _z9oo • ^n t_ _))U_)

.... -6 .......-4 -2 '''l'''''''jO 2 4 ...........6 8 1'0''1'2

Ga

¢-

'ra

¢J

t_

.10

.09

.08

.07

.06

.05

.04

.03

.02

.01

[] M = 1.6

/x Moo = 1.8

V Moo = 2.0 []

D Moo = 2.3 '_

M =-4.0ooO Moo = 4.63 _

L.),

0 , , , I . . . I i i _ [ _ , i I i _ i I , , _ I , , , I

-.3 -.2 -.1 0 .1 .2 .3 .4


6.,.._

,--i

[] Moo = 1.6

Moo = 1.8

V Mo =2.0

l> Moo = 2.3

Moo = 4.0

O Moo = 4.63

2

0

-2

-.3 -.2 -. 1 0 .1 .2 .3 .4


E

8.0

7.5

7.0

6.5

6.0

5.5

5.0

4.5

4.0

3.5

3.01.5

0 0) 0 0 0

2.0 2.5 3.0 3.5 4.0 4.5 5.0

Mach number, Moo

Figure 37. Aerodynamics of fully integrated cranked-wing configuration.

53

.30M = 4.0

.25

.20,

,_ .15

.10e_

.05

-_ -.05

-.10

-.15

Oo °

0 o0

0 8

0 o

[] Straight wing

0 Cranked wing

,,I,,,I,,,I,,,I,,,I, ,,I,,,I,,,I,,,I,,,I

-'20-8 _5 -4 -2 0 2 4 6 8 10 12

tX

.07

.06

r_ .05

e-

.N .04

8 .o3

M=4.0

.02

.01

0-.10

[] Straight wing (_

0 Cranked wing (5]

illll_j,,I .... I .... I .... 1 .... I .... I .... I

-.05 0 .05 .10 .15 .20 .25 .30


6

e_D

,J

M,, ° = 4.0

6

4

2

0

-2

-4

-6

-8-.10

(90

(9

0[] Straight wingOF

(ff_3 0 Cranked wing

.... I .... i .... I .... t .... i .... i .... i .... i

-.05 0 .05 .10 .15 .20 .25 .30


E

.e_

12

11

10

9

8

7

[] Straight wing

0 Cranked wing

)o ° 0 0 0

1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0

Mach number, M

Figure 38. Comparison of aerodynamics of straight-wing and cranked-wing fully integrated configurations.

54

..7e-

,.-1

.4M = 4.0

.3

.2

.1

0

--.1

-.2

0 Pure waverider (no base drag)

/_ Pure waverider (base drag) _20

-'3-8'''-6J'''-4''''-_2'''O_'''2''''4L'''6''''8J'''I_(I ''1_2

¢,a

o.0t_

.09

.08

.07

.06

.05

.04

.03

.02

.01

0_.2

M = 4.0


A Pure waverider (base drag)

[] Fully integrated

, , I

--,1

A

DOz5

[]0

_0

°, J , I I I

0 .1 .2


I

.3

d-.--r

t_

,.d

8

6

4

2

0

-2

-4

-6

-8--.2

, J

M = 4.0

O000Oo

[]

_0

eq_ 0 Pure waverider (no base drag)

0 A Pure waverider (base drag)

0 0 [] Fully integrated

J I , , _ I , _ , I _ , , I J , , I

-.1 0 .1 .2 .3


@,..1

E

9.0

8.5

8.0

7.5

7.0

6.5

6.0

5.5

5.0

4.5

4.0

3.5

3.0

2.0

O0 Pure waverider (no base drag)


[] Fully integrated

0O0

0

/x /x /k A[] []

t,,,l .... I .... l*il,lliiilliltl

2.5 3.0 3.5 4.0 4.5 5.0

Mach num_r, M

Figure 39. Comparison of aerodynamic performance of straight-wing fully integrated and straight-wing pure waveriderconfigurations.

55

,..1

,.r

8

,.d

.4M o = 4.0

.3

.2

.1

0

--,l

--.2

@@@@

@@ @@

I_¢D QO®@

(2) Pure waverider (no base drag)


[] Fully integrated

-'3_8' I ' I ' ' ' I ' ' ' 12' L* I * ' k L _ ' ' L ' ' ' I ' ' ' I ' ' ' I _ ' _I'--6 --4 -- 0 2 4 6 8 lO 12O_

.09

.08

.07

.06

o .05"5

8 .04

.03

.02

.01

0--.2

Moo : 4.0



[] Fully integrated I:_°

_0_0

_0_0

_0@_ _0O__n 0

_00000_

-.1 0 .!


1 I

.2 .3

6

,.d

8

6

4

2

0

-2

-4

-8--.2

Moo = 4.0

000000 -

g_(3r_

g_(D

rq

_t A_(3 (3 Pure waverider (no base drag)A Pure waverider (base drag)

0 [] Fully integrated(3(3

L I I _ , J I I I

-.1 0 .1 .2 .3


E

.e.

12

11

10

9

8

7

6

5

4

31.5

(3 Pure waverider (no base drag)

Zk Pure waverider (base drag)

[] Fully integrated

0 O00

[] [] /k /xA ix[_ [] [] []

i i i I b _ i , I .... I .... I .... I .... I .... i

2.0 2.5 3.0 3.5 4.0 4.5 5.0

Mach number, M

Figure 40. Comparison of aerodynamic performance of cranked-wing fully integrated and cranked-wing pure waverider

configurations.

56

E= 4e.

6

Present fully integratedcranked-wing waverider

• Reference 28

• Reference 29• Reference 30

• Reference 31

• Reference 32• Reference 33

2 3 4 5 6 7 8

Mach number, M._

Figure 41. Assessment of aerodynamic performance of waverider-derived vehicle.

57

.020

.015¢0

t-

.010

.005EO

,_ 0

_= -.005

M_ = 2.3

M =4.0

0 M =4.63b

b

...... _ h ; _........... _-'010-8 -6 - 4 6 '1;' ' '1'2

(a) Straight wing.

.025

.020

.015

,_ .010

E

_, .oo5.=.¢..,

h2 0

-.005

[] Moo = 1.6

A Moo = 1.8

v M_ = 2.0

D M = 2.3

O M,_ = 4.0

[]

[]

[] z_vE] &v

[][] Av

0 M =4.63 0 DO A _V 0

DD_DDO _b_b©QQ

bbv_ _bbbb_

................................... _d"-6 --4 -2 0 2 4 6 8 12

(b) Cranked wing.

Figure 42. Longitudinal stability of each fully integrated waverider-derived configuration with undeflected elevons andailerons.

58

,.d

t-

8

.3

.2

-.1

Mo_ = 2.3

_ 5 A = 0 °, 8 E = 20 °O 5 A = 20 °, fE = 200

-'2-8'"] '''1'' ' [2'_'1'''1 _ ''] '''_ '''1"' _' ''1--6 -4 - 0 2 4 6 8 10 12

.025

.020

r,.)

,_ .015¢-

"0.010

8

,_ .005

o 0,e

.=- -.005

._,-.010

Mo_ = 2.3

[] _A = 0% fiE = 0°

_A = 0°' _E = 200[]

@ fA = 200' 8E = 200 [] [][]

•D D _

••D ©0

00© 00000

ooOO 000

[][]

[][]

[]

o

GO ©

0 oo0

0

0

4).15 '''l'''l'''l'''l'L'l_'l'''llllLIL_lLlll

-8 -6 -4 -2 0 2 4 6 8 10 12

_d

e-

8

M = 4.0

20

-2 t@888 , =oo-4 _ OSA = 0°, fE = 20°

0 fA = 20% _E = 20°-6

-8 -6 -4 -2 0 2 4 6 8 10 12

.O25M• = 4.0

.020

,.s .015e-

.OlO

N .005

o 0

_ -.005

-.010

-.015-8

[] fA = 0% fE = 0°

_A = 0°' _E = 200[]

08 A=20 ° ,8 E=20 ° [] []

[]DaD•O© 0

aDDS• _ 000[] 0000 O0 O 0

_O O 0000O00@QO0 00°0

-6 --4 -2 0 2 4 6 8 10 12

(a) M_ = 2.3 and 4.0.

Figure 43. Pitch control effectiveness of elevons and elevon/aileron combination for straight-wing fully integratedconfiguration.

59

,d

e_

.1

.3M,, ° = 4.63

.2

.1

-.1

.O25

•-_ .015

.010

i .005_ _&__oo ___oo ,E o

o _A: 0%_: 20° .=.-.oo5e.

©Sa : 2°°, _E = 20° _-.010

281 ' ' [ ' ' ' _ ' ' ' ' ' ' ' _ ''12 . ' ' ] ` ' [ _ 1 ' ' J ' ' j I ' t t I _0.15-" -6 -4 -2 0 4 6 8 10 12 -8ot

M = 4.63

[] _A = 0°' _E = 0°

O 8 A = 0 °, _E = 200[]

© _A = 20°, q_E = 20° [] []

[]E3

[] O0 0

o•[]D_©

oog_,O0 0

ooooooooO°°°°

-6 -4 -2 0 2 4 6 8 10 12

(b) M.. = 4.63.


,-d

...$

O

.4

.3

.2

.1

0

--.1

--.2

Moo= 1.6

0oSo

8o _A--°°,_ --2°°o o 8A=0o,fie=_200

-" -6 -4 -2 0 4 6 8 10 12

.040

.035

.030

.025

.o20

8 .015

.010E

.005

,.= 0

._ -.005t_

-.010

Moo = 1.6

0 o°0

00000 []O000OO []

DDDD_DODG DD_D 0 o

IS 8A = 0°, fiE = 0°

+ 8 A = 0 °, 8 E = 20 °

O 8A = 0°' 8E = -20°

-.015 '''''''''''''''J'''''''''''''' ''',_,,,_-8 -6 -4 -2 0 2 4 6 8 10 12

,..Ak)

e-

,.d

Moo = 1.8

.4

o-.1 _ D (SA = 0°, 8E = 0°

-.2 _ +8A = 0°' fie = 20°

08 A = 0 °, 8 E =-20 °

-8 -6 -4 -2 0 2 4 6 8 10 12Ot

.040

.035

.030

'_ .025

.020

.015

o .010

O

.005

.= 0

-.005

-.010

-.015

-8

M = 1.8

0 o0 o

O00000oO0 Oo _D13

E3 8A = 0°, 8E = 0°

8 A = 0% fiE = 20°

0 8 A = 0% fie : -20 °

t,.I,,,I ,,.llllll,.i,_ ,i,,_l.,,i,,.lllll

-6 -4 -2 0 2 4 6 8 10 12

(a) M.o = 1.6 and 1.8.

Figure 44. Pitch control effectiveness of elevons for cranked-wing fully integrated configuration.

61

.4Moo = 2.0

.3

.2

.1

0

--.1

--.2O _5A 0 °, _)E = 20°

O 6 A = 0% _E = -200

-'3--8'''_ ......-4 -'2 ...............0 2 4 _,,,_ ........10 12

.040

.035

rj_ .030

.025

.020

.015

.010

_ .005

.= 0r.,

._ -.oo5

Moo = 2.0

OOQOQOOOQO 00000

DE)[_ 0 00DC]OD OC] FIE) C]

m _A = 0°, 8E = 0°

<_ 8A = 0°' _E = 20°

-.010 O8 A = 0% qY)E = -20 °

_.015_8,,,,,,,L,,,L2,..,*.,,,,,J,,,,,,, L,.._...,-6 -4 - 0 2 4 6 8 10 12ot

,-1

,,dI:I

0

.1

.4

.3

.2

.1

0

--,l

--.2

M =2.3

896:f 8 g gg 88[] a A = 0 o, i_E = 0 o

_A = 0°' $)E = 200

O _A = 0% _ = -20 °

-'38' -6 -4 -2 0 2 4 6 8 10 12

.040

.035

.030

.025.o_9

.020

.015

.OlO

_ .005

._ 0

_ -.005

-.010

-.015-8

M_o = 2.3

oooooooooOO°°!°°°13131313 00 °

00[3013 rq ° ° u.... 0<_0

OO OOOOOOOv[]aA=O°,_=O °

O _iA = 0 °, i_E = 20 °

O BA = 0% _5E = -200

-6 -4 -2 0 2 4 6 8 l0

(b) Mo. = 2.0 and 2.3.

Figure 44. Continued.

12

62

,..jL_

.=.,1

.4Moo = 4.0

.3

.2

.1

0

--.1

--,2

8_

_0000_

_80 _

8_ 0g []_A__oo__o o

G 8 A = 0% 8 E = 20 °

O8 A = 0 °, 8 E =-20 °

, I , , , i , , , , , , I , , , I , , , J .... i

-'38' ' _ _- ''2'- ()' 2 4 6 8' '1'() '1'2

o[

.040

.035

.030¢9

.025o

.020

8 .o15

.010

,_ .005

," 0

._-.oo5e_

-.010

-.015

Moo

[] 8 A = 0 o, 8 E = 0 o

8 g = 0 °, 8 E = 20 °

08 A = 0 °, 8 E =-20 °

= 4.0

oooooooo ©O o_VID t_

,,,r,,,I,,,t2,,,J,,,I,,,I,,,I,,,I,,,I,,,i-6 -4 - 0 2 4 6 8 10 12

..d

e_

8

..d

Moo = 4.63.4

.3

.2

0 _ 00_

-.1 [] 8A = 0o, $_E = 0o

8 A = 0 °, 8 E = 20 °--,2

0 8 A = 0 °, fiE = -20°

3_8,,, t,,.,,.,L2,,, I,,, I,,,J,,,t,,,,,..,,,,1,2-' -6 -4 - 0 2 4 6 8 10ot

.040

.035

.030

.025

.020

,015

.010

,_ .005

,_ 0e-

-.005e_

-.010

Moo = 4.63

_ 8A = 0°, 8E = 0°

8 A = 0 °, 8 E = 20 °

08 A =0 °, 6 E =-20 °

oooooooo0 O BSB8888_l_o_°°_

_.015_,,,,,,,_,,.,,,,_,,,,,,,_,,.,,.,_,,,,,,,,-8 -6 -4 -2 0 2 4 6 8 10 12

O[

(c) M.. = 4.0 and 4.63.


63

,.d

O

...1

.4

.3

.2

.1

0

--.1

--.2

M = 2.3

888888 8888

o 6A = 0°, _E = 20°

© 6A = 0o, _ = _20°

,,,i,,,i,,,l,,,i,,,i.,.i,,,),,,i,,,i,,,i-" -6 -4 -2 0 2 4 6 8 10 12

.040

.035

r_ .030

.025O

E .020O

8 .015

.010E

_ .005

.= 0

._ -.005

-.010

-.015

M. =2.3

[] _A = 0°' _ = 0°

_A = 0°, _E = 20°

O _A = 0°' _E = -20°

OOQOQO00QO0000000

CDCDDDDODODDDDO00

-6 -4 -2 0 2 4 6 8 10 12¢t

.1

.r

"0

0

8

,.d

.4

.3

.2

.1

0

--.2

Moo = 4.0

8

OOO00o

O8 A = 0 o, _ = 20 °

O 8A = 0 °, _5E = -20 °

-'38' --6 -4 -2 0 2 4 6 8 lO 12

o_

.040

.035

.030

= .025

.020

.015

o .010E

E, .005

.= o

._ -.oo5

-.OlO

-.015-8

M._ = 4.0

[] 8A = 0°, 8E = 0°

QSA = 0°, 8E = 20°

O _A = 0°' _E = -20°

00000000000000888DFIr_ mDC)[] O DDI_ _ _ r_

-6 -4 -2 0 2 4 6 8 10 12O_

(a) Mo. = 2.3 and 4.0.

Figure 45. Pitch control effectiveness of elevons for straight-wing fully integrated configuration with moment referencecenter at 58 percent of body length.

64

8.=.,.-I

.4 Mo= = 4.63

.3

.2

.1

0

--.l

--.2

O 8 A = 0 o, _ = 0 o

+ _A = °°, _ = 2°°

o 8A = 0%_ = -20 °

-8 -6 -4 -2 0 2 4 6 8 10 12

.040

.035

.030r,.)

.025

.020.o15

L_

.010E

.005

.n 0

:_-.oo5

-.010

M o = 4.63

[] 8A = 0°' _E = 0°

6A = 0°' _E = 20°

O fiA = 0°' fiE = -20°

°°°°oooooo98888998<SS>8<Saoooooooooooo

-.015 "'''_'''_'''''''_'" _'''_'_' J'''_,, ,'.,, _-8 -6 -4 -2 0 2 4 6 8 10 12

(b) M_ = 4.63.


65

.g

.4

.3

.2

.1

0

--.1

--.2

--.3

Moo = 1.6

°

<_o []_A_-oo,__-oo ._o ° <>8A_-oo,__-2ook-O 0 8A = 0°' 8E = -20°

-6 -4 - 0 2 4 6 8 10 2

M, ° = 1.64_5

D 8A = 0°, 8E = 0°.040

O _A = 0°' 8E = 20°

.035 O 8 A = 0 °, _E = -20°

.030 O

©©

.025 O00000000000oO

.020[]

[]

.015 n n D FI E313r7 _D DFII313 fin.010

©

.005 _©_©_©_©©_<_0

--.005 , ,_1, , ,I,_,lt J ,I,, ,iJ J,IJl_ll,,E_,,Ih, ,1

-8 -6 -4 -2 0 2 4 6 8 10 12

,1

c-

o

8

,.d

.41

.3

.2

.1

0

--.1

--.2

--.3

-8

M = 1.8.040

R .030

o= .025"0

'= .020

0,,.010

.oo5_ _ o_:oo,_:2oo

o_:oo,_:-2oo -005,,',,,',,,_,,,_,,''_''''''''''''''''''_ -.010

-6 -4 -2 0 2 4 6 8 10 12 -8

Moo = 1.8

D 8A =0°, 8E =0°

8 A = 0 o, 8 E = 20 °

O8 A = 0% 8 E =-20 °

OOOo000000OO00000

[] D[] [] n D rIC_C)_C) D(_ [] C]N R

-6 -4 -2 0 2 4 6 8 10 12

(a) M** = 1.6 and 1.8.

Figure 46. Pitch control effectiveness of elevons for cranked-wing fully integrated configuration with moment referencecenter at 59 percent of body length.

66

,d

_9

8

Moo = 2.0.4

.2

o

'-.2 _JA =0 ' _E 20o

(2) _iA = 0% gE = -20°

-'--83 , , i , , , [ , , , i , , , _, , , i , , , i , , , i , , , i , , , i , , ,1_6 --4 -2 2 4 6 8 10 12

.040

.035

.030

.025"5

.020

8- .015

E .0100

E.005

.E

_ o-.005

Moo = 2.0

O 8A = 0°, _E = 0°

_A = 0°' f)E = 20°

© 8A = 0°' _E = -20°

QOOOOOOooooooOQO 0

(3 CI D [3 CI C][31_ C]I_I V113 D V1CI rl FI

00000'_00©©0.0©,00©0

-.Ol0 ,,i,,,_,,,.,,,t,,._.,,,,,,,,,... ,_,,,i-8 -6 -4 -2 0 2 4 6 8 10 12

,..]

,d

,.d

.4

.3

.2

.1

0

--.|

--.2

--.3

Moo = 2.3.040

.035

8888188

.030

._ .025_9

8 _ 0208 88_ 8 _, .015 OlO

oo,6A = 0°' 815 = 20° ._ 0

tLO _A = 0% _E = -20° -.005

-6 -4 -2 0 2 4 6 8 10 12 -8

M =2.3

D _A = 0°, 6E = 0°

_5A = 0% fiE = 20°

O6 A = 0 °, 6 E =-20 °

O000Oo00000000000

-6 -4 -2 0 2 4 6 8 10 12

(b) M._ = 2.0 and 2.3.

Figure 46. Continued.

67

,.d

o

8

,.d

.4

.3

.2

.1

0

--.1

--.2

--.3

Moo = 4.0 .040 M = 4.0

.035

.030

88_8 "oo .025

8088_ __ .020

80080 _ o15888 _ olo

[]_A=0%_E=0o ._ 005o

+ 8A = 0 °, _ = 20 °

O 8A = 0% _E = -20° -.005

-6 -4 -2 0 2 4 6 8 10 12IX

[] _A = 0°' _)E = 0°

•_8 A = 0% 8 E : 20 °

0 8 A : 0 °, _E = -20°

°°°OOooooooooooo 8_[]••OrZDnDDDDDDO<P<}_'_O<D'OQCpO_O<D,O©O

-.010 ,, , ', ,, ',,, _,,, J,,, '... ',,, J,,, ..,, ,,.,-8 -6 -4 -2 0 2 4 6 8 10 12

IX

,..d

e-:

"0

8

,...1

Moo = 4.63.4

.3

.2 O0

.1 00000

o 0 0_0

-.1 _ 0_0013 8A = 0°, qSE= 0°

O 8A = 0 °, 8 E = 20 °--.2

O8 A = 0 °, 8 E =-20 °

-.3 tttl'''l'''l ' ''['/'I'''I'L'I'J'I '''l'ljJ

-8 -6 -4 -2 0 2 4 6 8 10 12IX

.04O

.035

.030

.025"0

.0208- .015e-

_: .010Q

E& .005

0

ta_

-.005

-.010-8

Moo = 4.63

D 8A = 0°, 8E =0°

8 A = 0 o, 8E = 20 °

©8 A = 0 o, 8 E =-20 °

°°Ooooooooooo9999[]D(3D(3U

-6 -4 -2 0 2 4 6 8 10 12IX

(c) M**= 4.0 and 4.63.


68

.0008

.0007

_= .0006

> .0005

'r_ .0004cD

._ .0003_D

E .0002@

,E.0001

.,=0

>,-.000l

-.0002

.0020

NM =2.3.0016

M = 4.0 I> _-

OM = 4.63 D _ .0012

[ >'_ .0008

b b

l> D D 4 .0004

0<? E

° 8 _ _ _ t _,-.oo_O O 0 0 Stable

_2 -.0008O

-.0012

[]M =2.3

QM = 4.0

OM = 4.63

b

et>

I Stable

e 8 8P

B>

>i>

I>

-4 -2 0 2 4 6 8 10 12 -6 -4 -2 0 2 4 6 8 10 12

Figure 47. Lateral-directional stability of straight-wing fully integrated configuration.

.O022

.0020

'_-.OOl8r..)d .0016

"_ .0014

_ .oo12.0010

E .0008oE .0006

.. .0004

.0002

0

-.0002-6

[] M =1.6 [] M =1.6

A Moo= 1.8 S M = 1.8

M o=2.0 V M =2.0

D M =2.3 D M =2.3

<_ M =4.0 @ M =4.0

O M =4.63 O Moo=4.63

[]

[]A

[][] A v

[] s /" v D[] [] /x V

A S V [>v v V l>

t> > l> t>

8 8 o_8 6 8Stable I

,,,I,.,l_,,I,,,I,,,I,.,ll_lljlllll_l

-4 -2 0 2 4 6 8 10 12

.0024

.0020

ca..0016

.0012d> .0008

..-_

;_ .0004.r"ID"_ 0

g-.ooo4-.0008

&-.O012e-.

;-=-.0016

-.0020

-.0024

-.0028

[],,_' _> g g.D _ v b

v D[] s D

s V DV

[] s[] A

[]

-4 -2 0 2 4 6 8 10 12

Figure 48. Lateral-directional stability of cranked-wing fully integrated configuration.

69

.0010

.0008

,; .0006

t_;_ .0004

.r-

.0002_A

Eo 0E

•=--.0002

--,0004

-.0006-6

O Vertical tail off

[] Vertical tail on

E2[] [] [] []

[][]

Stable [O

0 0 0 0 0 0

.0016

.0012

d .0008

;_ .0004

_ 0

_ -.0004i

_ -.0008ID

-.0012

,,,_,,, I,,, J,,,,,., i,,, ,,,, J,,, i,,, I -.0016-4 -2 0 2 4 6 8 10 12

O Vertical tail off

[] Vertical tail on

g

i Stable

-4 -2 0 2 4 6 8 10 12

Figure 49. Effects of vertical tail on lateral-directional stability of straight-wing fully integrated configuration at

M_ = 4.0.

.0010 .0016

,_, .0008e_

_; .0006

e_.0004

o

"fi .00O2

o 0E,

•_ -.0002

--.0004

0 Vertical tail off

[] Vertical tail on

[]

©

[]

[][] O

[] D [] ©© t

O O O Stable /

.0012

u-..0008

.0004-r"Q.I

•_ 0

_ -.0004

;= -.0008

-.0012

-.000% .........-4 - 2'"0 .... 2' ..............4 6 8 .........10 ;2 -.0016 _

0 Vertical tailoff

[] Vertical tail on

t0 rl _ _Stable

@@

--4 -2 0 2 4 6 8 10 12

Figure 50. Effects of vertical tail on lateral-directional stability of cranked-wing fully integrated configuration atM_, = 4.0.

7O

.016

I> M = 2.3

.014 (> Moo = 4.0

O Moo = 4.63

t> b E>E>t> _ b _>D b

D2> D I>D D2>

.002

L8 4 4 -2 0 2 4 6 8 10 12

.012w-<]

- .010

fio .008E

" .006-6"¢ .004

gEO

E

.010

.008

.006

.004

.002

0

_> Moo = 2.3

© Moo = 4.0

O Moo = 4.63

-.002.=.

--.004

-.006

-.008

f_-'01_-8 -6 -4 -2 0 2 4 6 8 10 12

Figure 51. Aileron effectiveness on lateral-directional stability of straight-wing fully integrated configuration;

fiA = +20° and _E = 0°-

,J

EO

E

=O

,v

.016

.014

.012

.010

.008

.006

.004

.002

Moo= 1.6

Moo= 1.8

Moo= 2.0

Moo = 2.3

Moo = 4.0

Moo = 4.63 .010

.008[]

[] [] [3 [] [] .006n

A /X _ .004A Eli

VV V V_ _, A /X [3 _ .002

0

88 reac; . v -.002D _ -.004

_8888_@00001>00 O0 * -.006

-.008

Moo= 1.6

Moo = 1.8

Moo = 2.0

Moo = 2.3

Moo = 4.0

Moo = 4.63

eooo@@@

u(3mNN[]NN

08''''' _' ' '' ' I '''''''' _' _'''' ' '''' 't*''ll 2 -.010 '''''''''''''''+'''+'''''''_'''''''_'''_--6 -4 -2 0 2 4 6 8 10 -8 -6 -4 -2 0 2 4 6 8 10 12

Figure 52. Aileron effectiveness on lateral-directional stability of cranked-wing fully integrated configuration;

_A = +20° and _5E = 0 °.

71

.016

.014

.012&

.010e_

_ .008E

" .006

0

_" .004

.002

Moo= 1.6

Moo = 1.8

Moo = 2.0

Moo = 2.3

Moo = 4.0

Moo = 4.63

[3[3DO DO0

AAAAA []AAA []

V VV VVVszzA []

08, , , i _, , ] , , , _, _, i , , , i , , , i , , , i _, , _, , , t , , ,i,...6 --4 -2 0 2 4 6 lO 12

.010

.008

.006

J .004

.-: .002

Eo 0E

._ -.002

-.006

-.008

Mo= 1.6

Moo = 1.8

Moo = 2.0

Moo = 2.3

Moo = 4.0

Moo = 4.63

o@@@@g@f>v - @@@@

[3 VI VI FI £)_ 13 i- I,,,i,, ,i ,,,I,,,It,,/,ILI,,,I,,,II,,I,, ,I

-'010-8 -6 -4 -2 0 2 4 6 8 10 12

Figure 53. Combined roll/pitch effectiveness on lateral-directional stability of cranked-wing fully integrated configura-tion; _A = +20° and 8E = 20°-

72

!

REPORT DOCUMENTATION PAGE I ro,_ A_o,o,_,_I ouBNo.oro4-ola8

Public re_i;.g burden for this collectio_ of klformetion is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources,gathering and maintaining the data needed, and completing and revlewing the collection of information Send comments regarding this burden esttmete or any other aspect of thiscollection of Informetion, including suggestions for reducing this burden, to Washington Headquarters Services, Direclorste for Information Operations end Reports, 1215 JeffersonDavis Highway, Suite 1204, Arlington, VA 2220_-4302. and to the Office of Managernenl and Budget, Pape_¥ork Reduction Project (0704-0188), Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

July 1996 Technical Paper4. TITLE AND SUt:IIi/LE S. FUNDING NUMBERS

Aerodynamic Characteristics of Two Waverider-Derived Hypersonic CruiseConfigurations WU 466-02-01-01

S. AUTHOR(S)

Charles E. Cockrell, Jr., Lawrence D. Huebner, and Dennis B. Finley

7. PP-HFORM;t;GORGANIZATIONNAME(S)ANDADDRESS(ES)

NASA Langley Research CenterHampton, VA 23681-0001

9. SPONSORING/MONITORINGAGENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space AdministrationWashington, _ 20546-0001

8. PERFORMING ORGANIZATION

REPORT NUMBER

L-17479

10. SPONSORING/MONITORINGAGENCY REPORT NUMBER

NASA TP-3559

11. SUPPLEMENTARY NOII::_

Cockrell and Huebner: Langley Research Center, Hampton, VA; Finley: Lockheed-Fort Worth Company,Fort Worth, TX.

12a. Di_iHIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited

Subject Category 02Availability: NASA CASI (301) 621-0390

12b. DISTRIBUTION CODE

13. A_,_IPIACT (Maximum 200 words)

An evaluation was made on the effects of integrating the required aircraft components with hypersonic high-liftconfigurations known as waveriders to create hypersonic cruise vehicles. Previous studies suggest that waveridersoffer advantages in aerodynamic performance and propulsion/airframe integration (PAl) characteristics over con-ventional non-waverider hypersonic shapes. A wind-tunnel model was developed that integrates vehicle compo-nents, including canopies, engine components, and control surfaces, with two pure waverider shapes, both conical-flow-derived waveriders for a design Mach number of 4.0. Experimental data and limited computational fluiddynamics (CFD) solutions were obtained over a Mach number range of 1.6 to 4.63. The experimental data show thecomponent build-up effects and the aerodynamic characteristics of the fully integrated configurations, includingcontrol surface effectiveness. The aerodynamic performance of the fully integrated configurations is not compara-ble to that of the pure waverider shapes, but is comparable to previously tested hypersonic models. Both configura-tions exhibit good lateral-directional stability characteristics.

14. SUBJECTTERMSHypersonic cruise; Waveriders; Airbreathing vehicles

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

NSN 7540-01-280-5500


OF THIS PAGE

Unclassified


OF ABSTRACT

Unclassified

15. NUMBER OF PAGES

7316. PRICE CODE

A04

20. LIMITATION

OF ABSTRACT

Standard Form 298 (Rev. 2-89)Prescnbed by ANSI Std. Z39-18298-102

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