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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:
NASA Center for AeroSpace Information
800 Elkridge Landing Road
Linthicum Heights, MD 21090-2934
(301) 621-0390
National Technical Information Service (NTIS)
5285 Port Royal Road
Springfield, VA 22161-2171
(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
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1. Bowcutt, K. G.; Anderson, J. D.; and Capriotti, D.: Viscous
Optimized Hypersonic Waveriders. AIAA-87-0272, Jan. 1987.
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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
(b) Base view of CFD solution.
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
(b) Base view of CFD solution.
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
(b) Base view of CFD solution.
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
.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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
[] No controls (base drag)
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
[] No controls (base drag)
-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
Lift coefficient, C L
.--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
[] No controls (base drag)
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
Lift coefficient, C L
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
0 Pure waverider (no base drag)
A Pure waverider (base drag)
[] Fully integrated
, , I
--,1
A
DOz5
[]0
_0
°, J , I I I
0 .1 .2
Lift coefficient, C L
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
Lift coefficient, C L
@,..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)
A Pure waverider (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)
A Pure waverider (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
0 Pure waverider (no base drag)
A Pure waverider (base drag)
[] Fully integrated I:_°
_0_0
_0_0
_0@_ _0O__n 0
_00000_
-.1 0 .!
Lift coefficient, C L
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
Lift coefficient, C L
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.
Figure 43. Concluded.
,-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.
Figure 44. Concluded.
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.
Figure 45. Concluded.
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.
Figure 46. Concluded.
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
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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
18. SECURITY CLASSIFICATION
OF THIS PAGE
Unclassified
19. SECURITY CLASSIFICATION
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