Aerodynamic Characteristics of Two Waverider-Derived Hypersonic Cruise Configurations

NASA Technical Paper 3559

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

July 1996

NASA Technical Paper 3559

Aerodynamic Characteristics of Two Waverider-Derived Hypersonic Cruise Configurations

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 Administration Langley 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 National Technical Information Service (NTIS) 800 Elkridge Landing Road 5285 Port Royal Road Linthicum Heights, MD 21090-2934 Springfield, VA 22161-2171 (301) 621-0390 (703) 487-4650 Abstract

An evaluation was made of the effects of integrating the required aircraft components with hypersonic high-lift configurations known as waveriders to create hypersonic cruise vehicles. Previous studies suggest that waveriders offer advantages in aerodynamic performance and propulsionairframe integration (PAl) characteristics over conventional non-waverider hypersonic shapes. A wind-tunnel model was developed that integrates vehicle components, 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 fluid 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 aerodynamic characteristics of the fully integrated configurations, including control surface effectiveness. The aerodynamic performance of the fully integrated configurations 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 integrating all vehicle components. The final objective was to evaluate the controllability of each of the fully A waverider is any shape designed such that the bow integrated vehicles and the effectiveness of the control- shock generated by the shape is perfectly attached along surface design. These objectives were accomplished the outer leading edge at the design flight condition. The using results from wind-tunnel testing and a limited num- waverider design method leads to several potential ber of computational fluid dynamics (CFD) solutions. advantages over conventional non-waverider hypersonic The CFD predictions were obtained for the pure wave- concepts. The attached leading-edge shock wave con- rider shapes only and provide comparisons with experi- fines the high-pressure region to the lower surface and mental data and design-code predictions. Two wind- results in high lift-drag ratios. Several design predictions tunnel models were designed that integrate canopies, suggest that waveriders may offer an aerodynamic per- engine packages, and control surfaces with two Mach 4.0 formance advantage in terms of higher lift-drag ratios pure waverider shapes. The models were tested in the over non-waverider hypersonic concepts (refs. 1 and 2). Langley Unitary Plan Wind Tunnel (UPWT) at NASA In addition, the flow field below the waverider bottom Langley Research Center. surface is uniform and, in the case of waveriders derived from axisymmetric flow fields, there is little or no cross- This report describes the waverider aerodynamic flow in this region, making these shapes attractive candi- design code used and discusses the method used in the dates for engine integration. These advantages have led development of the wind-tunnel models. The details of to interest in using waverider shapes for the forebody the experimental study are then presented as well as the geometries of hypersonic airbreathing engine-integrated airframes. Waveriders have been considered for various computational method used to obtain the CFD predictions. The results are analyzed in three sections. First, the types of missions including hypersonic cruise vehicles, results of the pure waverider shapes without integrated single-stage-to-orbit vehicles, airbreathing hypersonic vehicle components are presented. These results include missiles, and various space-based applications (ref. 3). flow-field characteristics from CFD solutions and experi- The purpose of the current study is to examine the mental flow-visualization data as well as aerodynamic aerodynamic characteristics of two waverider-derived characteristics from the experiment and CFD predictions. hypersonic cruise vehicles. No experimental data cur- Second, the experimental results of adding aircraft com- rently exist that address the integration of realistic vehi- ponents to the pure waverider shapes are presented. The cle components with waverider shapes. Therefore, the effects of the canopy, engine components, and control objectives of this study were threefold. The first was to surface additions on aerodynamic performance and create an experimental and computational database for stability are examined. Finally, the aerodynamic charac- waverider-derived configurations. The second was to teristics of the fully integrated waverider-derived config- examine the effects of individual vehicle components on urations are examined and compared with those of the pure waverider performance and to determine the differ- pure waverider shapes. Control-surface effectiveness is ences in aerodynamic characteristics that result from also addressed in this section. 2. Symbols p density, lbm/ft 3

B.L. buttline of model (distance from centerline in Subscripts: spanwise direction), in. c conditions at first cell center next to solid drag coefficient boundaries G rolling-moment coefficient c.g. center-of-gravity location OC t flee-stream conditions Ct_ rolling-moment derivative, _

lift coefficient G 3. Configuration Design and Model cM pitching-moment coefficient Development cn yawing-moment coefficient

_C n 3.1. Waverider Design Method C_ yawing-moment derivative, _l 3 A specific waverider shape is uniquely defined by moment reference length, in. free-stream conditions, the type of generating flow-field body, and a leading-edge definition (ref. 1). The shapes C L of the upper and lower surfaces of the configuration fol- L/D lift-drag ratio, low from these parameters. The free-stream conditions, including Mach number and Reynolds number or alti- Mach number tude, are selected based on mission criteria. The design model station (distance from nose in stream- method used in this study involves a specific design wise direction), in. point. The generating flow-field body is used to define the shock shape upon which the leading edge of the P pressure, lbf/fi 2 waverider is constructed. Although any arbitrary body in roll rate, deg/sec supersonic or hypersonic flow can be used as a generating flow-field body, this study focuses specifically on the Re Reynolds number class of conical-flow-derived waveriders, in which the planform area, ft 2 generating flow-field body is a right circular cone in supersonic or hypersonic flow. At the outset of this U velocity component, ft/sec research effort, this option was the best available for the V total volume, ft3; velocity, ft/sec application of interest. Other possible generating flow V2/3 fields include osculating cone flow fields (ref. 4), hybrid cone-wedge generated flow fields (ref. 5), and inclined volumetric efficiency, Sref circular and elliptic conical flow fields (ref. 6). The W.L. waterline of model (distance from zero refer- length of the generating cone, length of the waverider, ence in vertical direction), in. and semiapex angle of the cone are specified by the designer. The selection of these parameters can signifi- moment reference center location XC.g. cantly affect the shape of the waverider generated as well X,F,Z Cartesian coordinates, in. as the aerodynamic performance of the configuration. y+ inner law variable Figure 1 illustrates the design of a conical-flow-derived waverider. The planform shape, or leading edge, is angle of attack, deg defined on the shock wave produced by the cone. The sideslip angle, deg lower surface of the configuration is defined by tracing streamlines from the leading edge to the base of the cone. angle of aileron deflection (trailing edge down The result is that the lower compression surface is a positive), deg stream surface behind the conical shock wave. The con- angle of elevon deflection (trailing edge down figurations studied here have an upper surface that is positive), deg designed as a constant free-stream pressure surface. distance from solid boundary to first cell However, other techniques may be used, such as shaping center, in. the upper surface as an expansion or compression surface. The conical flow field, defined behind the viscosity coefficient, lbf-sec/ft 2 shock wave, exists only below the lower surface of computational coordinates the waverider.

2 The resultingconfigurationoffers two possible The design code used in this study is the (University advantagesover non-waveriderhypersonicconfigura- of) Maryland Axisymmetric Waverider Program tions.Thefirst is a potentialaerodynamicperformance (MAXWARP) (refs. 1, 2, and 9). The MAXWARP code advantage(refs.1, 2, and7). Theoretically,theshock is an inviscid design method that includes an estimate for waveisperfectlyattachedalongtheouterleadingedgeat skin friction in the design process. Various volumetric thedesignMachnumber.Theresultis thatthehigh- constraints may also be imposed by the user in order to pressureregionbehindtheshockwaveisconfinedtothe produce waveriders with desirable structural characteris- lowersurface,andnoflow spillagefromthelowersur- tics and component packaging. These constraints include faceto theuppersurfaceoccurs.Themaximumlift-drag aspect ratio, slenderness ratio, and total volume. For ratiosthismethodproducespromisetoexceedthoseof the case of conical-flow-derived waveriders, the Taylor- existinghypersonicconfigurations.Figure2,takenfrom Maccoll equation, which describes the flow field behind reference2, showsthetraditional"L/D barrier"in the a conical shock wave (ref. 10), is integrated using a supersonic/hypersonicregimeforconventionalvehicles. fourth-order Runge-Kutta method to compute the invis- Thiscorrelationisempirical,basedonactualflightvehi- cid conical flow field behind the shock wave. The cone cleexperienceatsubsonicandlowsupersonicspeedsand semiapex angle and length of the flow-field generating extrapolatedto hypersonicMachnumbers(ref.7).The body are specified by the user along with free-stream symbolsin figure2representpredictionsforavarietyof conditions. The code starts with an initial leading-edge conical-flow-derivedwaveridershapesgeneratedusing definition on the conical shock wave and creates a thecurrentmethod,whichisdescribedin detailin refer- waverider shape from this initial leading edge. The pres- ence2.Thewaveridershapesrepresentedhereareonly sure distributions on the surface of the configuration are theforwardportionsof possiblehypersonicconfigura- integrated to calculate lift and drag coefficients. An esti- tionsandthereforearenotrealisticvehicles.Thepredic- mate for skin friction is also included so that force coeffi- tionsshownassumethattheconfigurationhaszerobase cient predictions include both inviscid and viscous dragin orderto removetheeffectof thebluntbase, effects. This estimate is based on the reference tempera- whichwill beeliminatedin a fully integratedvehicle, ture method, which is described in reference 11. The andshowonlytheperformanceof theforwardportionof effect is to generate shapes for which wetted surface area sucha vehicle.In otherwords,thepredictionsassume is minimized to reduce skin friction drag. The code uses thatfree-streamstaticpressureactsatthebase,makinga a simplex optimization routine (ref. 1) to optimize directcomparisonof thelift-dragratiosfor waveriders waveriders for a given figure of merit: maximum lift- andthoseof existingsupersonic/hypersonicconfigura- drag ratio or minimum drag. More recent versions of the tionsdifficult.Furthermore,thewaveridersrepresented code allow the user to construct various other objective heredo nothavelevelsof volumetricefficiencycompa- functions. At each iteration in the optimization process, rabletothoseofthevehiclesusedin theL/D barrier cal- an updated leading-edge definition is used to generate a culation and may not have been obtained at similar new waverider shape that progresses toward the desired flight-scaled Reynolds numbers. Although the lift-drag figure of merit. This process continues over a number of ratios of a fully integrated waverider configuration with iterations until the optimum shape is found without viola- the blunt base closed would likely be lower than those tion of any of the user-specified volumetric constraints. for the pure waverider shape, these predictions suggest that waveriders may offer an aerodynamic performance 3.2. Waverider Shape Description advantage over non-waverider vehicle concepts. Another advantage of axisymmetric waverider flow fields is that The pure waverider shapes used in this study, which the lower surface flow field is uniform, and there is pure define the forward portions of the waverider-derived conical flow in this region for a perfectly attached shock vehicles, were designed using the MAXWARP design wave. Therefore, a known uniform flow field can be code. Free-stream conditions and optimization parame- delivered to scramjet engine modules on the lower sur- ters were chosen based on the applicability of this study face, providing a benefit in propulsion/airframe integra- to a hypersonic cruise vehicle, with available ground- tion (PAl) (ref. 8). The osculating cone and cone-wedge based test facility limitations taken into account. The concepts mentioned previously may provide an even design free-stream Mach number was 4.0 and the design greater benefit over conical-flow-derived waveriders Reynolds number was 2.0 x 106 per foot. Although the (refs. 4 and 5). The aerodynamic performance and PAI specific cruise Mach number for this type of vehicle benefits suggested in previous research efforts have gen- would be higher, Mach 4.0 was selected as the design erated interest in using waveriders for various hypersonic point based on the limitations of the UPWT and the vehicle designs. range of data desired. The Mach number range of this facilityis1.47to4.63.A designpointofMach4.0would of free-streamconditions.Therefore,an attemptwas permitthevalidationof thewaveriderconceptat the madeto increasethevolumetricefficiencyasmuchas designMachnumberandalsoallowfor thedetermina- possiblewhileacceptinga minimumpenaltyin maxi- tion of aerodynamiccharacteristicsat off-designMach mumlift-dragratio.Finally,aconfigurationwithaflator numbers.Theuseof endothermicfuelson thisvehicle slightlyconvexbottomsurfacein thecrosssectionwas classis expectedto drivetheselectionof cruiseMach desiredfor easein propulsionsystemsintegration.In numberto approximately5.0to 5.5.No significantdif- additiontothesethreeprimarydesignguidelines,acon- ferencesin theflow physicsareexpectedbetweenthe figurationfreeof substantialcurvatureovermostof the ultimatedesignMachnumberandtheMachnumber crosssectionwasalsodesiredtoprovidefortheinclusion rangeinvestigatedin thisstudy.TheReynoldsnumber of aninternalsparin anactualaircraft.Furthermore,the chosenis basedon nominalfacilityoperatingconditions targetvalueof span-to-lengthratiowas0.8.Information in theUPWTandisnotrepresentativeof aflightcruise from previousstudiesshowsthat largerspan(higher altitude.Theconfigurationwasoptimizedformaximum aspectratios)waveridersprovidehigherlift-dragratios lift-dragratioatthedesignpointbecausethisquantityis butaremoredifficult to integrateasa full waverider- moreappropriatethanminimumdragasa hypersonic basedvehicle(ref.12). cruiseperformanceparameter. A three-viewdrawingandan obliqueviewof the A fullyturbulentboundarylayerandawalltempera- straight-wingpuretheoreticalwaveridershapegenerated tureof 585°Rwerespecifiedin thedesign.Thiswall bythedesigncodeareshownin figure3.Table1sum- temperaturewasselectedbasedonpreviousexperimental marizesthecharacteristicsof thisshape.Thespan-to- datafrommodelstestedin theUPWT.It isnotlikelythat lengthratiois 0.83.Thelowersurfaceof thestraight- fully laminarconditionscouldbemaintainedin experi- wingconfigurationhasa slightconvexcurvaturethat mentaltestingattheconditionsof interest,andtransition facilitatesintegrationof the propulsionsystem.The is difficultto predict.Fullyturbulentconditionscanbe lengthselectedfor the waveriderconfigurationwas achievedandmaintainedbytheapplicationofboundary- 24.0in. basedon thesizeof thetest sectionin the layertransitiongrittothemodelsurface. UPWT.Thelengthof thegeneratingconewasselected to fix thelocationof thewaveriderleadingedgeon the Twodifferentpurewaveridershapesweredeveloped conicalshockwaveto achievethedesigncriterianoted for thisstudy.Thefirst is referredto asthe"straight- previously(48.0in. for thisapplication).A selectionof wing" shapeandwasdesignedusingtheMAXWARP differentlocationson theconicalshockwavewould optimizationroutine.The second,referredto as the resultin waveriderswithmuchdifferentgeometricchar- "cranked-wing"shape,wascreatedbyadjustingthelead- acteristicsandmayresultin thegenerationof unrealistic ing edgeof the straight-wingwaveriderto createa shapesthatcouldnotbe integratedinto vehicles.The curvedwingtipshapethathadincreasedaspectratiobut volumetricefficiency,Vef f, of this configuration is 0.11 still maintainedshockattachmentalongtheouterleading with a predicted maximum lift-drag ratio of 6.9. edgeat the designfree-streamcondition.The term "cranked"inthiscontextreferstoawingshapein which A three-view drawing and an oblique view of the thesweepanglenotonly changesbutalsoexhibitsa cranked-wing pure theoretical waverider shape generated largeoutboarddihedralanglein theplaneof thebase. by the design code are shown in figure 4. The cranked The cranked-wingshapewas designedto provide leading edge still lies on the same conical shock wave improvementsin subsonicaerodynamicperformance produced by the generating cone used to design the (becauseof increasedaspectratio) and in lateral- straight-wing waverider. The characteristics of the directionalstability(becauseof dihedraleffect)while cranked-wing waverider shape are summarized in maintaininghigh performancein the supersonic/ table 2. The span-to-length ratio is 0.96, which represents hypersonicregime. an approximately 16 percent increase in aspect ratio. This increase in aspect ratio should improve the subsonic Threeprimarydesigncriteriawereusedtoselectthe aerodynamic performance over the straight-wing wave- bestwaveridershapedesignsfor thisapplication.First, rider while maintaining the structural characteristics themaximumlift-dragratiowaschosentobeashighas of the straight-wing waverider near the centerline of possiblewhilenotviolatingotherdesignguidelines.This the configuration. The volumetric efficiency of this criteriondrivestheselectionof theconesemiapexangle configuration is 0.108 with a calculated maximum lift- for the generatingflow field. A valueof 8.1° was drag ratio of 6.7. This configuration represents only a selectedforthisapplication.Second,thevolumetriceffi- slight decrease of both parameters from the straight-wing ciency(V2/3/Sref)waschosento beashighaspossible. waverider. The slight convex curvature of the bottom An inverserelationshipexistsbetweenthevolumetric surface is maintained toward the centerline of the efficiencyandthemaximumlift-dragratioforagivenset model. The dihedral angle of the aft cranked section is approximately28°whenmeasuredfromthecenterlineof surface to integrate some type of engine system and is thissection. not intended to be a realistic propulsion simulation. The Thevaluesfor maximumlift-dragratiogivenarefor inlet capture area, expansion ramp turning angle, and nozzle exit area were designed for full-scale Mach 4.0 thepurewaveridershapesonly. The waveriders were conditions. The compression surface shown in figure 8 is subsequently altered to close the blunt base and add con- required for additional precompression of the flow enter- trol surfaces. The predictions assume that free-stream ing the inlet. The non-flow-through configuration static pressure are acting at the base of the unaltered pure attempts to model the external cowl drag present on a waverider shape, so that only forebody drag values are realistic flow-through nacelle, but does not have the included in the performance predictions. As will be associated internal drag. Two different nozzle/expansion shown later, the incorporation of aftbody closure is a sig- ramps were designed for the model. The first was used nificant issue in hypersonic vehicle development. with the pure waverider configurations with the nacelle attached and the second was used with configurations 3.3. Wind-Tunnel Model Designs that had control surfaces attached. These nozzles are Two slight modifications to the design-code shapes referred to as the "short" and "long" nozzles, respec- were implemented in the wind-tunnel model design in tively (figs. 10(a) and 10(b)). Identical nozzles with static order to accommodate model support hardware and addi- pressure taps were also fabricated in order to obtain sur- tional vehicle components. A smooth ogive-cylindrical face pressure measurements on the nozzle. The non- fairing was blended on to the upper surface of the pure instrumented ramps were used for force and moment waverider shapes to accommodate the sting and balance runs. necessary to measure the aerodynamic loads on the model during testing. This volume was added to the Control surfaces were provided to examine their upper surface rather than the lower surface because pre- effects on waverider aerodynamic performance as well as vious research indicates that modifications to the lower the effectiveness of the control concept. The control sur- surface have an affect on the PAI characteristics of the faces were sized based on control-volume trends from waverider (ref. 13). Figures 5 and 6 show tunnel installa- supersonic fighter aircraft to extend and close the blunt tion photographs of the straight-wing and cranked-wing base of the configurations. Elevon deflections of 0% pos- pure waverider models with the upper surface fairing. itive 20 ° (trailing edge down), and negative 20 ° (trailing The lower surface of the theoretical waverider shape was edge up) were incorporated. A set of outboard ailerons modified slightly by creating an inboard expansion sur- having the same three deflection angles was designed for face with an angle of approximately 10 °, beginning the straight wing. Because of the curved surface of the approximately 22 in. aft of the nose of the configuration cranked wing and the small thickness of the outboard and measuring approximately 3.5 in. in the spanwise leading edge, the set of ailerons for the cranked-wing direction. The lower surfaces follow the waverider theo- configuration consisted of an inboard aileron, which retical stream surface up to this point. This modification remained fixed at 0 °, and a set of outboard ailerons, was made in order to facilitate the integration of engine which were deflected at 0 ° and +_20°. A vertical tail sur- components and to reduce the closure angle necessary for face was also designed in order to augment directional control surfaces. Figure 7 shows a photograph of the stability. Figures 8 and 9 show photographs of the model lower surface of the cranked-wing waverider with the components with the various control surfaces. Figure 11 expansion on the aft end of this surface. shows three-view drawings of the elevons, straight-wing ailerons, cranked-wing inboard ailerons, and cranked- A realistic canopy was designed for the waverider- wing outboard ailerons. This figure indicates the perti- based configuration. The canopy was provided with fac- nent dimensions and shows the hinge-line locations for eted surfaces to resemble the canopy for a hypersonic each control surface. vehicle. The aft portion of the canopy was designed to blend with the cylindrical fairing on the upper surface The model design allowed for testing of the straight- discussed previously. Figures 5 and 6 show the pure wing and cranked-wing pure waverider models, which waverider models with the ogive-cylindrical fairing are defined as configurations with no engine components attached (i.e., canopy-off configuration). Figures 8 and 9 or control surfaces. A configuration build up of the show the model with the faceted canopy attached. waverider models with different vehicle components The engine package for this configuration included a could also be tested up to and including the fully inte- compression ramp, a non-flow-through engine module grated waverider-derived configurations, which are with side walls, and a nozzle/expansion ramp. The defined as configurations with engine components, con- engine-package-on configuration provided an indication trol surfaces, and the canopy. Table 3 shows the pertinent of the effect of modifying the theoretical waverider lower model geometry for each configuration tested. Figure 12

5 shows a three-view drawing of the fully integrated ify that yawing and rolling moment values are linear over configurations. this range (ref. 15). Based on these results, stability derivatives were calculated from data obtained at the two 4. Experimental Method fixed sideslip angles.

The facility used in this study was the UPWT at The data obtained from the wind-tunnel tests include NASA Langley Research Center. The UPWT is a closed- six-component force and moment data, static pressure circuit, continuous-flow pressure tunnel with two 4- by readings on the blunt base of the model, static pressure 4- by 7-ft test sections, which were both used in this data on the nozzle surfaces, and flow-visualization data. study. The Mach number range of the facility is 1.47 The balance used in this case was the NASA-LaRC- to 4.63, with a range in the low Mach number test section designated UT-50-B balance, which is a six-component of 1.47 to 2.86 and a range in the high Mach number test strain gauge balance. Unless otherwise noted, the section from 2.30 to 4.63. Continuous variation of Mach moment reference center for all configurations was number is achieved by using asymmetric sliding block located 16.623 in. aft of the nose. A total of 11 5-psi nozzles to vary the nozzle throat-to-test-section area pressure transducers were used to measure the static ratio. The Reynolds number range of this facility is pressure along the blunt base of the configurations and in 0.5 x 106 to 8.0 x 106 per foot. However, the nominal the cavity surrounding the sting. Integrated areas were Reynolds number for most tests is 2.0 x 106 per foot. assigned to each tap or averaged group of taps and used A detailed description of the UPWT can be found in to calculate the base axial force. All of the force data pre- reference 14. sented is corrected to assume free-stream static pressure acting at the base. This procedure is carried out so that The configurations tested ranged from the straight the data may be presented showing only the upper and and cranked pure waverider models to the fully inte- lower surface lift and drag values and eliminating the grated waverider-derived vehicles. The test configura- effect of the blunt base. This procedure is necessary tions were chosen to show pure waverider performance; to isolate the effects on waverider aerodynamic perfor- because the base will be eliminated in any realistic waverider-derived configuration. The method of assum- mance of the canopy, engine package, and control surfaces; and to show the aerodynamic performance and ing free-stream pressure at the base is consistent with the design-code method and with previous studies showing stability characteristics of the fully integrated configurations. Only the cranked-wing configurations were tested predictions for waverider aerodynamic performance in the low Mach number test section. The data were cor- (refs. 2, 9, and 16). Details on the procedure used are rected for flow angularity in the test sections. Calibration included in reference 15. For configurations with both data for the UPWT shows that the flow in both test sec- engines and control surfaces, only two base and two tions has an upflow angle generally within 0.5 ° of the chamber pressures were measured. A 32-port, 5-psi tunnel centerline (ref. 14). In each run, either six- external electronically scanned pressure (ESP) module component force and moment data, nozzle pressure data, was used to measure the static pressure on the nozzle surface for four runs. Figure 10 shows the locations of or vapor-screen photographs were obtained. Schlieren pressure taps on the nozzle surfaces for the short and photographs were taken during the force and moment runs. long nozzles. Recall that the short nozzle is used with configurations having no control surfaces and the long The test conditions were chosen to investigate the nozzle is used with configurations with control surfaces. aerodynamic performance and stability of each configu- A total of 12 pressure taps were located on the short noz- ration at both the design Mach number and at off-design zle and 24 pressure taps were located on the long nozzle. Mach numbers. Data were obtained at Mach numbers of The data are used to correct the nozzle surface pressures 2.3, 4.0, and 4.63 for all configurations studied and, addi- to assume free-stream static pressure acting on these sur- tionally, at Mach numbers of 3.5 and 4.2 for some con- faces for some configurations. figurations. Data for the cranked-wing configurations were also obtained at Mach numbers of 1.6, 1.8, and 2.0. Schlieren and laser-vapor-screen photographs were The free-stream Reynolds number for most runs was taken in order to examine flow-field features includ- 2.0 x 106 per foot. Some runs were made at Reynolds ing the shock attachment characteristics for various numbers of 1.5 x 106 per foot and 3.0 x 106 per foot in configurations. For the vapor-screen runs, the laser was order to investigate the effects at off-design Reynolds positioned outside of the test section window and the numbers. The angle-of-attack range studied was -6 ° to light sheet was projected across the model surface in the 10° at fixed sideslip angles of 0 ° and 3°. Data were spanwise direction, illuminating one cross section at obtained over a sideslip angle range of-5 ° to 5 ° for the a time. The camera was mounted inside of the test sec- first configuration run in each test section in order to ver- tion above and behind the model. This setup gives a cross-sectionalviewof thewaveriderflow fieldin the k-computational direction runs from the surface of the vapor-screenphotographs. configuration to the outer boundary. The grids for each Theaccuracyof theUT-50-Bbalance,basedon a of the two pure waverider shapes contained 91 points in May1993calibration,is0.5percentof full scaleforeach the _ direction, 111 points in the 11 direction, and componentto within95-percentconfidence.Thefull- 91 points in the _ direction. Blunt leading edges were scaleloadlimits were600lbf normal,40 lbf axial, modeled for each configuration in order to provide a bet- 1500in-lbfpitchingmoment,400in-lbfrollingmoment, ter comparison with experimental data. Grid points were 800in-lbfyawingmoment,and300ibfsideforce.Asan also clustered near the surface of each configuration in example,usingthemethodof root-mean-squaressum- order to adequately resolve the boundary-layer flow. The amount of grid spacing needed is judged by examining mationtocombineindependenterrorsources,theselim- the grid spacing parameter, y÷, which is given by itscorrespondto arangeof uncertaintyin lift coefficient of 0.0053at_ =0° to 0.0054atot= 10° andanuncer- taintyrangein dragcoefficientof 0.00036at o_=0° y+= / pcucA_ to0.001ato_= 10° fortheMoo = 4.0 and Reoo = 2.0 × 106 6/ 0c (1) per foot condition. The repeatability of measurements was observed to be better than these uncertainties. There- where Pc, Uc, and _tc are the density, velocity, and viscosity at the first cell center next to the solid surface and A t fore, differences less than the indicated ranges for com- is the distance from the first cell center to the body sur- parisons with data from different configurations in the face. Previous research has shown that y÷ values on the same test, could be considered significant. However, order of I provide accurate solutions (ref. 21). comparisons between independent measurements are only good to within the quoted uncertainty ranges. Tran- The CFD solutions were obtained using the General sition grit (no. 60 size sand grit in the low Mach number Aerodynamic Simulation Program (GASP), version 2.2 tests section and no. 30 size grit in the high Mach number (refs. 22 and 23). GASP is a finite volume code capable test section) was applied in a 0.1-in-wide strip to the of solving the full Reynolds-averaged Navier-Stokes model upper and lower surfaces along the outboard lead- (RANS) equations as well as subsets of these equations, ing edge at a location approximately 0.4 in. from the including the parabolized Navier-Stokes (PNS), thin- leading edge in the streamwise direction. These proce- layer Navier-Stokes (TLNS), and Euler equations. Time dures were established for models tested in the UPWT integration in GASP is based on the integration of primi- based on unpublished transition experiments conducted tive variables, and convergence to a steady-state solution in the UPWT and the methods of references 17 to 19. is obtained by iterating in pseudotime until the L2 norm of the residual vector has been reduced by a sufficient 5. Computational Method amount. GASP also contains several flux-split algo- rithms and limiters to accelerate convergence to steady Computational grids were developed for each of the state. Mesh sequencing is available as a means to accel- pure waverider configurations by first developing a erate convergence. numerical surface description and then creating 3-D volume grids. Numerical surface descriptions of the In this study, each configuration was modeled as a straight-wing and cranked-wing wind-tunnel models two-zone problem, as illustrated in figure 13. The first were obtained from computer-aided design (CAD) zone includes the blunt nose of the configuration. The descriptions of the model parts. Three-dimensional vol- flow in this region is a combination of subsonic and ume grids were created for each configuration using the supersonic flow because a small area of subsonic flow GRIDGEN software package, which uses algebraic exists behind the detached bow shock. Therefore, the transfinite interpolation methods with elliptic interior TLNS equations are solved over the first zone using a point refinement (ref. 20). Only the pure waverider global iteration procedure. The second zone encom- shapes with no integrated vehicle components were mod- passes the remainder of the configuration, extending eled for the CFD analysis. from the zonal boundary to the base of the configuration. The flow in this region is computed by solving the PNS The computational grids for each of the two pure equations. These equations are valid for regions of waverider shapes model only half of the configuration predominately supersonic flow with no streamwise because each is symmetric about the centerline. The grid separation. A no-slip boundary condition is applied to all orientation is shown in figure 13. The k-computational solid boundaries with a fixed wall temperature of 585°R, direction runs from the nose of the configuration to the which is identical to that specified in the MAXWARP base in the streamwise direction. The q-computational optimization routine when designing the waverider direction begins at the upper centerline and wraps around shapes. Free-stream conditions are applied at the outer the leading edge, ending at the lower centerline. The boundary, second-order extrapolation from interior cells is applied at the last streamwise plane, and symmetry sion surface begins, so the flow entering the inlet would boundary conditions are applied at the center plane. The be highly compressed. Similar data are shown in fig- Baldwin-Lomax algebraic turbulence model was used in ure 15 for the cranked-wing pure waverider model. The these solutions to model turbulent boundary layers, and shock can be seen in the right-hand side of the photo- convergence to a steady state was obtained by reducing graph to be very near the outer leading edge of the the L2 norm of the residual vector by 5 orders of model. The lower surface is again highlighted by the magnitude. laser light. The full cross-sectional view is not shown because of the poor quality of the photographs. The In order to make appropriate comparisons, the condi- experimental data confirm the qualitative shock location tions at which solutions were obtained were chosen at the outer leading edge, which is predicted by the CFD based on conditions at which experimental data were solution for this case as well. Figure 16 further illustrates available. Solutions were obtained at Mach 4.0 at angles that the shock is slightly detached at the outer leading of attack of -6 °, 0°, 2°, 4 °, and 8° for the straight-wing edge for both models. This figure shows a close-up view model. Solutions were obtained at Mach 4.0 at angles of of the outer leading edge at the base of the cranked-wing attack of-6 °, 0°, and 8 ° for the cranked-wing model. and straight-wing waverider shapes from CFD solutions Solutions were also obtained at off-design Mach num- at Mach 4.0 and 0 ° angle of attack. Both of the views bers of 2.3 and 4.63 at 0 ° angle of attack for each in figure 16 are to the same length scale, and non- configuration. dimensional static pressure contours are shown in each view. 6. Flow-Field and Aerodynamic Characteristics of Pure Waverider Models The flow-field characteristics of each pure waverider shape at off-design Mach numbers can also be illustrated 6.1. Flow-Field Characteristics by examining experimental flow-visualization data and CFD solutions. Figure 17 shows a comparison of a The flow-field characteristics of the pure waverider vapor-screen photograph and a CFD solution for the models at the design Mach number can be illustrated by cranked-wing shape at Mach 2.3 and 0° angle of attack. examining computational solutions of each configuration The free-stream Reynolds number is 2.0 × 106 per foot. and laser vapor-screen photographs from wind-tunnel The data shown in this figure and orientation of the tests. Figure 14 shows a laser vapor-screen photograph of camera in the test section are the same as in figures 14 the flow at the base of the pure straight-wing waverider and 15. At Mach numbers below the design Mach num- model and nondimensional static pressure contours at the ber of 4.0, the shock-wave angle is larger and the detach- base of the same configuration from a CFD solution at ment distance should be much larger than at the design Mach 4.0, 0° angle of attack, and free-stream Reynolds Mach number. This outcome is predicted by the CFD number of 2.0 x 106 per foot. The model lower surface is solution and confirmed by the experimental data. Fig- highlighted in the photograph by the laser light sheet on ure 18 shows similar views of the same configuration at the surface. The bow shock is indicated by the contrast Mach 4.63. The photograph in this figure was taken with between light and dark regions below the light sheet. On the laser light sheet approximately 5 in. upstream of the the left-hand side of the photograph, the shock is base because the quality of the photograph taken with the observed to be very near the edge of the lower surface. light sheet at the base was poor. At Mach numbers Thus, the vapor-screen photograph confirms the qualita- greater than the design Mach number, the shock moves tive shock location predicted by the CFD solution. A closer to the leading edge than at the design condition, as small detachment distance exists even at the design point illustrated in both the vapor-screen photograph and pre- caused by blunt leading edge and boundary-layer dicted by the CFD solution. A large high-pressure region displacement effects. These effects are not accounted for still exists in the bottom-surface flow field of this config- in the design code. The CFD predictions also indicate uration at Mach 4.63. The qualitative shock locations can that the high-pressure region remains mostly confined be further illustrated by examining planform schlieren below the model lower surface. A large low-pressure photographs of the cranked-wing model. Figure 19 region (P/P** of 0.95 or less) exists near the centerline of shows schlieren photographs of the cranked-wing pure the model below the bottom surface because of the bot- waverider model in a planform view at Mach 2.3 (top), tom surface expansion present on the model. However, Mach 4.0 (middle), and Mach 4.63 (bottom). The right the remainder of the bottom surface flow field is a side of the figure shows a close-up view near the leading smooth, conical flow field, so the presence of this slight edge at each Mach number. The schlieren images in this expansion surface does not degrade the favorable PAl figure have been enhanced by computer imaging tech- characteristics offered by the waverider. Engine modules niques in order to show the shock structure more clearly. would be placed upstream of the point where the expan- At Mach 2.3, the schlieren photograph shows that the shockis detachedfromtheleadingedge.Theoutermost occurs at an angle of attack greater than 0 ° for the wave- shockin the top view representsthe bowshock.At rider configurations studied in these references. Mach 4.0, the shock is much closer to the outer leading edge, but a small detachment distance still exists. At A direct comparison of the experimental aero- Mach 4.63, the photograph does not show the presence of dynamic performance of the two pure waverider models a shock wave near the leading edge, possibly because the is shown in figure 22. The experimental data show that shock is attached at this condition. 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 6.2. Aerodynamic Performance slightly higher lift coefficients. Aside from these obser- vations, there are no significant differences between the The aerodynamic performance characteristics of the two configurations. two pure waverider models are examined here using experimental force and moment data and computational The off-design performances of the straight-wing predictions. Off-design Mach number and Reynolds pure and cranked-wing pure waverider models are shown number effects are evaluated using experimental data. in figures 23 and 24, respectively. Each of these figures The longitudinal and lateral-directional stability charac- shows the experimental lift, drag, and lift-drag ratio at all teristics are also examined for each configuration using Mach numbers studied as well as maximum lift-drag experimental data. Unless otherwise indicated, all of the ratio versus Mach number. The data indicate that there is experimental and computational data presented have no significant performance degradation at off-design been corrected to a condition of free-stream pressure act- Mach numbers. Both configurations show higher maxi- ing at the blunt bases of the configurations, as previously mum lift-drag ratios than the design point value at Mach discussed. numbers less than 4.0, using the assumption of free- stream pressure acting at the base. Similar results have The aerodynamic performance of the straight-wing been found in previous waverider studies (refs. 13, and cranked-wing pure waverider shapes at the design 16, and 24) and are also typical for non-waverider Mach number is shown in figures 20 and 21. These fig- supersonic/hypersonic configurations. The cranked- ures show experimental data, CFD predictions, and wing waverider shape provides better aerodynamic per- design-code predictions for the lift, drag, and lift-drag formance at Mach numbers of 4.0 and below. At higher ratios of each configuration at Mach 4.0 and a Reynolds Mach numbers, there are no significant differences number of 2.0 x l06 per foot. The computational values between the performance of the two configurations. were obtained by integrating surface pressure and skin The effects of Reynolds number on aerodynamic friction predictions from CFD solutions. Because the performance of the straight-wing and cranked-wing con- data are corrected to eliminate the base drag, these data figurations are shown in figures 25 and 26, respectively. should be interpreted as the performance of the forward No significant effects of Reynolds number variation were portion (or forebody) of a possible hypersonic configura- observed for either configuration in the range studied, tion and not that of a realistic hypersonic vehicle. In gen- except for a slight increase in maximum lift-drag ratio at eral, agreement is good between the experimental data the 3.0 x 106-per-foot condition for both configurations. and computational predictions. Both the computational This result is most likely because the skin friction coeffi- predictions and experimental data show lower lift and cient decreases as Reynolds number increases, resulting higher drag values than the predicted design-code values, in decreased drag and thus increased lift-drag ratios at and these differences can be attributed to several causes. higher Reynolds numbers. The decrease in drag observed The flow-visualization data and CFD flow-field solutions experimentally is approximately equal to the decrease in showed that a slight detachment distance exists at the viscous drag predicted by the reference temperature outer leading edge even at the design condition, which method (ref. 11). Computational solutions at Mach 4.0 results in a lift loss and a drag decrease. However, the and a Reynolds number of 2.0 x 106 per foot show that design code assumes an infinitely sharp leading edge the viscous drag contribution is approximately 34 percent with a perfectly attached shock wave. An additional lift of total drag. By comparison, the MAXWARP design loss results from the expansion ramp on the bottom sur- code predicts a viscous contribution of approximately face of the waverider, and an increase in drag results 38 percent to the total drag. from the additional volume added to the upper surface of the model. The experimental data also show that the The pitching-moment characteristics of the straight- maximum lift-drag ratio occurs near 2 ° angle of attack wing and cranked-wing pure waverider configurations for each configuration. This finding is also consistent are shown in figure 27. This figure shows the pitching- with previous studies, such as those in references 13 moment coefficient versus angle of attack at each Mach and 16, which show that the maximum lift-drag ratio number studied. Both configurations are longitudinally

9 unstableat all Mach numbers studied. The moment refer- tion when the faceted canopy is used. Similarly, a ence center location here is an arbitrarily selected 5.1-percent reduction in maximum lift-drag ratio occurs location at the approximate location of the center of grav- for the cranked-wing configuration. The data indicate ity of the fully integrated model. This moment reference that a penalty was incurred for the canopy, and therefore center location (16.623 in. aft of the nose) is used for all attention should be paid to the canopy design in a hyper- configurations studied unless otherwise stated. The sonic waverider-based vehicle. cranked-wing pitching moment curve is more nonlinear than that for the straight-wing shape, indicating that the 7.2. Effect of Engine Package shock may be detached at higher angles of attack for the cranked-wing configuration. The yawing moment char- The engine component effects are evaluated by com- acteristics are shown in figure 28. This figure shows the paring experimental data from engine-on and engine-off yawing moment derivative versus angle of attack at each configurations. Figures 32 and 33 show the effects of Mach number studied for both configurations. The adding the engine package (ramp, inlet, and nozzle com- straight-wing configuration is directionally unstable at all ponents) to the straight-wing and cranked-wing configurations, respectively. The data shown here are for Mach numbers studied at angles of attack of 8° and configurations with the canopy and no control surfaces. below. The cranked-wing configuration is directionally stable at all Mach numbers studied above an angle of The data are corrected to assume free-stream static pres- attack of 4 °. Both configurations experience a destabiliz- sure acting at the base. No correction is applied to these ing effect as Mach number increases. The cranked-wing data for the nozzle surface pressures. Each figure shows configuration was expected to provide improved direc- lift and drag coefficients as well as lift-drag ratios at tional stability from the increased dihedral along the out- Mach 4.0 and the maximum lift-drag ratio at comparative board leading edge. The rolling moment characteristics Mach numbers for engine-on and engine-off configura- are shown in figure 29 for each configuration. The tions. The addition of engine components results in a cranked-wing waverider shows better lateral stability slight increase in lift and a significant increase in drag at characteristics than the straight-wing model. The Mach 4.0. These effects are caused by the inlet compres- cranked-wing configuration exhibits positive effective sion surface and the increase in projected frontal area and dihedral above 0 ° angle of attack at all Mach numbers. produce a decrease in lift-drag ratio at positive values of The straight-wing model is unstable at angles of attack lift and a reduction in maximum lift-drag ratio over the below 6.0 ° at Mach numbers of 4.0 and 4.63 and is unsta- Mach number range studied. The straight-wing engine- ble at angles of attack below 4° at a Mach number of 2.3. 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 7. Component Build-Up Effects 17.7-percent reduction at the same condition.

7.1. Effect of Canopy 7.3. Effect of Control Surface Addition

The effects of adding the canopy on the aerodynamic The effects of adding undetected control surfaces performance of the pure straight-wing and cranked-wing are illustrated by comparing data for configurations with waverider models are illustrated in figures 30 and 31, no control surfaces to those with undetected ailerons and respectively. These data were obtained for configurations elevons attached. Each configuration includes the canopy that have no control surfaces or engine components and engine components. Data for both the straight-wing attached, and the data are corrected to assume free- and cranked-wing configurations are shown. The coeffi- stream static pressure acting at the base. Each figure cient data are reduced by the planform areas of each cor- shows the lift and drag coefficients as well as lift-drag responding configuration so the effects of increased ratios at Mach 4.0 and the maximum lift-drag ratio at planform are accounted for in the normalization of these each comparative Mach number studied for the canopy- data. The plots showing drag and lift-drag data include off and faceted-canopy configurations. The canopy-off three separate data sets. The first is the data for the configurations have the ogive-cylindrical fairing on the controls-off configuration corrected to assume free- upper surface, as discussed previously. Both the straight- stream pressure at the base. Therefore, only forebody wing and cranked-wing configurations show little differ- drag values are included in these data and base drag is ence in lift when the canopy is added. The canopy-on not included. The second data set is the controls-on data configurations show slightly higher drag than those with and the third set is the controls-off data with base drag no canopy and an accompanying decrease in lift-drag included (i.e., uncorrected data from wind-tunnel mea- ratios at positive values of lift over the Mach number surements), so that these data include the effect of the range studied. The maximum lift-drag ratio at Mach 4.0 blunt base. A comparison between the second and third is reduced by 3.6 percent for the straight-wing configura- data sets shows the aerodynamic effect of adding control

10 surfacestotheconfiguration,andacomparisonbetween a waverider stream surface all the way to the base while thefirst twodatasetsshowstherelativeperformance designing the upper surface as an expansion surface. betweentheclosedconfigurationsandthatof thefore- Longer control surfaces would also reduce the closure bodysurfaceonlywithouttheeffectof thebluntbase. angle and enhance the pitch control power of the Theeffectof addingundeflectedcontrolsurfacesto configuration. straight-wingwaveriderconfigurationwith thecanopy andenginecomponentsattachedis summarizedin fig- 8. Characteristics of Fully-Integrated ure34.Theadditionof controlsurfacescausesa slight Waverider-Derived Hypersonic Cruise decreasein lift coefficientatMach4.0.Thisdecreaseis Configurations partiallycausedby the largeexpansionanglethat is presenton theelevonlowersurfacesanda 16-percent 8.1. Aerodynamic Performance increasein referenceareafor thecontrols-onconfigura- tion.A comparisonof thecontrols-offdatawith base Aerodynamic characteristics of each of the fully dragandthecontrols-ondatashowsadecreasein dragat integrated waverider-derived configurations are exam- agivenlift-coefficientvalue.Thereisaslightincreasein ined over the Mach number range using experimental lift-dragratiosat low positiveanglesof attackandan data, and the performance of these configurations are increasein maximumlift-dragratioswhen0° control compared to that of the pure waverider shapes. The surfaceswereaddedto theconfiguration.However,a fully integrated configurations are defined here to have comparisonof thecontrols-ondatawiththecontrols-off the canopy, the engine components, the undeflected aile- datawith nobasedragshowsthattheclosedconfigura- rons, the undeflected elevons, and the vertical tail tionhassignificantlyhigherdragvaluesandlowermaxi- attached. The aerodynamic characteristics of the straight- mumlift-dragratiosthantheforebody-onlyvalues.This wing and cranked-wing configurations are presented first resultindicatesthattheinclusionof aftbodyclosurepre- followed by comparisons to the corresponding pure sentsa significantchallengein theintegrationof pure waverider configuration. waveridershapesintohypersonicvehiclesandthatthis The aerodynamic performance of the straight-wing aspectof theconfigurationdeservesspecialconsider- and cranked-wing waverider-derived hypersonic cruise ationin thedesignprocess.It is likelythatthelift-drag configurations are shown in figures 36 and 37, respec- ratiosof aclosedconfigurationcannotapproachthoseof tively. The data presented here have the nozzle surface purewaveridershapesbecausetheeffectof basedragis pressures corrected to assume free-stream pressure acting oftennotincludedin lift-dragvaluesfortheseconfigura- on the nozzle surface. The data are presented using this tions.Theeffectsof controlsurfaceadditionaresimilar method to show the aerodynamic characteristics without for thecranked-wingconfigurationasindicatedin fig- any propulsive effect on the nozzle surface. The force ure35. Forreference,the baseareais approximately data were corrected by assigning integration areas to 8.3percentof theplanformareafor thestraight-wing each pressure tap measurement and computing the cor- modelwith no control surfacesand approximately rected coefficients. The locations of pressure taps are 9.1percentforthecranked-wingmodel. shown in figure 10. The straight-wing configuration has Thecontrolsurfacedesignfortheconfigurationused a maximum lift-drag ratio of 4.69 at Mach 4.0 and the in thisstudywasasomewhatarbitrarydesignbasedonly cranked-wing configuration has a value of 4.56 when the on trendsfromvarioussupersonicfighteraircraft.A nozzle surface pressures are corrected to free-stream moreoptimumdesigncouldminimizetheperformance pressure. The aerodynamic performance of each configu- degradationcausedby theclosureof thebluntbase.A ration does not vary significantly at off-design Mach performanceimprovementcouldbeobtainedbyinclud- numbers. The maximum lift-drag ratio for each configu- ingtheaftbodyclosurein thedesign/optimizationpro- ration also occurs near 2 ° angle of attack at Mach 4.0. The angle of attack for maximum lift-drag ratio increases cess.Previousstudieshaveexaminedthepossibilityof as Mach number decreases. At Mach numbers of 2.0 and usingblunttrailingedgesoncontrolsurfacesasameans of enhancingthe aerodynamicperformance(refs.25 below, the maximum lift-drag ratios for the cranked- to27).Thebluntbasereducesthestrengthof thebase wing configuration do not follow the general trend of recompressionshockandproperdesignof thetrailing increasing maximum lift-drag ratio with decreasing Mach number. This situation results from lift curve slope edgecanresultin an increasein basepressureanda values that show similar inconsistencies at Mach num- decreasein drag.A controlsurfacedesignthattakes bers less than 2.3. advantageof theseeffectswould enhancethe aerodynamic performanceof the configuration.This A direct comparison of the straight-wing and enhancementcouldbe accomplishedby reducingthe cranked-wing fully integrated vehicles is shown in fig- thicknessof thebasebymaintainingthelowersurfaceas ure 38. The straight-wing configuration produces slightly

11 highervaluesof maximumlift-dragratio than the maximum lift-drag ratio at Mach 4.0 for the fully inte- cranked-wingconfigurationatMachnumbersof 2.3and grated configuration is 4.56, compared to a value of 6.72 higher.Thestraight-wingmodelalsoshowshigherlift for the fully integrated vehicle. coefficientvaluesatMach4.0.Thestraight-wingmodel showsa maximumlift-dragratio thatis 3.0 percent From these results, it can be concluded that the max- higherthanthatofthecranked-wingconfigurationatthe imum lift-drag ratios of a fully integrated waverider- designMachnumberof4.0. derived configuration with aftbody closure likely cannot approach those of pure waverider shapes. Theoretical Comparisonsof theaerodynamicsof thestraight- predictions for waverider configurations do not include wingpurewaveridermodelandthefullyintegratedcon- the effects of aftbody closure. However, it will be shown figurationareshownin figure39.Thisfigureshowslift that the fully integrated waverider-derived configurations anddragcoefficientsaswellaslift-dragratiosatMach studied here are comparable in aerodynamic performance 4.0andthemaximumlift-dragratiosateachMachnum- to previously tested hypersonic models with performance berstudied.Asin figures34and35,thesedatasetsare improvements possible through enhanced control surface presentedforcomparisonin thedragandlift-dragplots. and propulsion system designs. Thefirst datasetrepresentsthepurewaveridershape with nobasedragincluded.Thesecondrepresentsthe In order to characterize the lift-drag values of the waveridershapewithbasedrag,andthethirdrepresents configurations studied here, a comparison is made thefully integratedconfiguration.Thenozzlesurface between data for the present cranked-wing fully inte- pressuresarecorrectedtoassumefree-streampressureon grated waverider-derived configuration and experimental thenozzlesurfacefor thefully integratedvehicles.A data from six hypersonic vehicle wind-tunnel models comparisonof thepurewaveriderdatawith basedrag previously tested in NASA Langley facilities (refs. 28 to andthefully integrateddatashowsthattheaerodynamic 33) in figure 41. Although direct comparisons of these performanceof thepurewaveridershapeis degraded data are not possible here because of different conditions, whenallof thevariousvehiclecomponentsareadded.A geometries, levels of volumetric efficiencies, and force reductionin lift coefficientfor thefully integratedcon- accounting methodologies, a range of values can be figurationis observedat Mach4.0above0° angleof obtained to compare with the data from the current study. attack,whichincreasesasangleof attackincreases.An As shown in figure 41, the waverider falls within the increasein dragis observedwhenall componentsare same general range of lift-drag values as the non- integratedwiththepurewaveridermodel.Theseeffects waverider hypersonic configurations. The lift-drag ratios resultin adecreasein lift-dragratiosatMach4.0andin of the waverider configurations studied could be maximumlift-dragratiosatcomparativeMachnumbers improved significantly through a better design of the pro- of 4.0andabove.At Mach2.3,thereisaslightincrease pulsion system and control surface closure. Therefore, in maximumlift-dragratiowhenall vehiclecomponents the waverider configurations studied here offer at least areadded.Thisincreaseismostlikelycausedbythenoz- comparable aerodynamic performance and perhaps a zle surfacepressurecorrectionto free-streampressure. modest advantage over conventional non-waverider Thefree-streamstaticpressureincreasesasMachnum- hypersonic vehicles. berdecreases.However,theaerodynamicperformance of thefully integratedvehicleis significantlydegraded 8.2. Longitudinal Control Effectiveness and Trim fromthatofthepurewaveridershapeonlywithnobase dragincluded,becauseof thedragproducedbythecon- Both fully integrated configurations are longitudi- trol surfaceaddition.The maximumlift-dragratio at nally unstable at each Mach number studied. The Mach4.0for thefully integratedvehicleis 4.69,com- pitching-moment coefficient data as a function of angle paredto6.68forthepurewaveridershape. of attack at each Mach number studied are shown in figure 42. Data for the straight-wing and cranked-wing fully A comparisonof thefully integratedcranked-wing integrated waverider-derived hypersonic cruise configu- configurationand the pure cranked-wingwaverider rations are shown. The moment reference center is modelyieldsconclusionssimilartothoseofthecompari- located at 62.5 percent of the centerline chord. At higher sonof thestraight-wingconfigurations.Figure40shows angles of attack, the cranked-wing configuration shows a the aerodynamicperformanceof the cranked-wing destabilizing increase in the pitching moment curve. This waveriderforebodyandthecranked-wingfully inte- increase indicates that the shock may have detached from gratedconfiguration.Theadditionof vehiclecompo- the leading edge of the outer cranked portion of the wing nentscausesa slightdegradationin the aerodynamic at higher angles of attack. The longitudinal instability of performance,but the lift-drag ratios observed for the these configurations may be addressed in one of two fully integrated model are significantly lower than those ways. First, it may be possible to shift the center-of- for the pure waverider shape only with no base drag. The gravity location for a fully integrated flight vehicle to a

12 locationthatwouldprovideatleastneutralstabilityover sible to control the shift in static margin from subsonic to the Machnumberrange.Recommendationsfor such supersonic speeds using fuel transfer. Neutral stability locationsarepresentedlaterin thissection.Second,the can be achieved by placing the center of gravity at a loca- additionof afully functioningpropulsionsystemwould tion equal to 58 percent of the centerline chord for the enhancethelongitudinalinstabilitybyincreasingtheaft- fully integrated straight-wing configuration and 59 per- bodylowersurfacepressures. cent of the centerline chord for the cranked-wing config- Thepitchcontroleffectivenessof theelevonsand uration. Data for lift and pitching-moment coefficients devon/aileroncombinationfor thestraight-wingconfig- referenced to these center-of-gravity locations are shown urationisshownin figure43.Dataareshownforthree in figure 45 for the straight-wing vehicle and in figure 46 trimsettings.Thefirst isonewithboththeelevonsand for the cranked-wing vehicle. In figure 45, the data for aileronsat 0°, thesecondwith a positive20° elevon the trailing-edge-up elevon deflections were extrapolated from the cranked-wing data and applied to the straight- deflection(_E) and a 0° aileron deflection (_A), and the third with both elevons and ailerons deflected at 20 °. The wing configuration. Also note that all of the data pre- effectiveness of the elevon decreases as Mach number sented here are for unpowered conditions. The addition of a functioning propulsion system will enhance the lon- increases, as evidenced by the smaller increments in lift gitudinal stability of the vehicle even further. These data and pitching-moment coefficients produced by each deflection. The ailerons were more effective than the are presented only to indicate the effects of an alternative choice of center-of-gravity locations. Subsequent data elevons in pitch control because of the shadowing of the are presented at the original moment reference center elevon behind the thick wing shape and the location of location of 62.5 percent of the centerline chord. the elevon in an expansion flow field. The CFD flow field solutions showed that the bottom surface flow field 8.3. Lateral-Directional Stability and Control expands to pressure below free-stream pressure in the Effectiveness region where the elevons are placed. Also, the closure angle for the elevon was severe because of the thick base The lateral-directional stability of the straight-wing of the waverider. Each aileron has only 70 percent of the and cranked-wing hypersonic cruise vehicles are shown planform area of the elevon but at higher angles of attack in figures 47 and 48, respectively. Each figure shows generates substantially more pitching moment. These yawing and rolling moment derivatives at each Mach characteristics may be unacceptable and indicate that the number studied. Both configurations are directionally pitch control concept should be redesigned. stable at all Mach numbers investigated, with the cranked-wing model providing higher stability levels The pitch control effectiveness of the elevons for the than the straight-wing model. The cranked-wing fully cranked-wing configuration is shown in figure 44. Each integrated configuration is laterally stable across the figure shows data for 0°, 20 °, and -20 ° elevon deflections with 0 ° ailerons. No runs were made with both aile- angle-of-attack range at all Mach numbers studied. The straight-wing fully integrated configuration is laterally rons and elevons deflected at the same angle because of unstable at angles of attack below 6 ° (at Mach 4.0). This the shape of the trailing edge for the cranked-wing con- roll instability may be caused in part by the high place- figuration. The elevon pitch control power for this con- ment of the balance in the model. No transfer distance in figuration also decreases as Mach number increases. the vertical direction was applied to the moment refer- However, in contrast to the straight-wing pitch control ence center in the presentation of data. data, the cranked-wing pitching moment curves are nonlinear. This factor makes the elevon pitch control power Figures 49 and 50 show the effect of the vertical tail even more critical for this configuration than for the on yawing moment derivative and rolling moment deriv- straight-wing vehicle. These data indicate that the nose- ative values for each configuration. The effect of the ver- down pitch control power of this configuration is not suf- tical tail is to significantly enhance the directional ficient. Either symmetric ailerons must be used to pro- stability of both the straight-wing and cranked-wing con- vide additional pitch control power or the elevon area figurations, indicated by the large positive shift in yaw- should be increased. ing moment derivatives when the vertical tail is added to each model. No rudder control effectiveness runs were Because of the combination of unstable pitching made in this study, so it is not clear whether sufficient moment characteristics and low pitch control power yaw control power exists to augment stability. The addi- observed in the experimental data, the configurations tion of the vertical tail does not cause any significant should be balanced such that they are at least neutrally change in the lateral stability characteristics of either stable to ensure adequate pitch control power throughout configuration. the angle-of-attack range. For a realistic full-scale flight vehicle, it should be possible to control the center-of- The effectiveness of a 20 ° aileron deflection on the gravity location through packaging. Also, it may be pos- straight-wing configuration is shown in figure 51. A 20 °

13 ailerondeflectionindicatedhereimpliesoneaileronwith controllability characteristics of the fully integrated a20° trailing-edge-down deflection and the other with a hypersonic cruise vehicles. 20 ° trailing-edge-up deflection. The elevons remained The flow-field characteristics and aerodynamic per- fixed at 0 ° for these runs. Figure 51 shows rolling formance of the two pure waverider shapes were moment and yawing moment increments between the examined using experimental and computational data. deflected and nondeflected runs. Additionally, the Computational fluid dynamics (CFD) predictions and DATCOM computer code was used to estimate the laser vapor-screen photographs of the straight-wing and steady state roll rates for this configuration (ref. 34). cranked-wing pure waverider models confirmed the Table 4 shows the steady roll rate capabilities as pre- shock attachment/detachment characteristics of each dicted by this method. The roll rate is shown as deg/sec configuration. The shock was slightly detached from the of roll, normalized by flight velocity. For most vehicles outer leading edge at the design Mach number of 4.0 of this type, excess roll-control power is available at and 0° angle of attack. This detachment distance exists lower angles of attack. The requirements for pitch and roll control surfaces for the waverider-derived vehicles because of boundary-layer displacement effects as well as blunt leading-edge effects. The design code assumes may be driven by low-speed flying qualities. These qual- an infinitely sharp leading edge and does not account for ifies include roll-rate capabilities at subsonic speeds and crosswind landing requirements. the physical presence of a boundary layer. Comparisons between experimental force data and CFD predictions Figure 52 shows the effectiveness of the ailerons for were generally good. The maximum lift-drag ratios the cranked-wing fully integrated configuration. How- observed experimentally were lower than the design- ever, a significant difference exists between these results code predictions, as expected. These lower lift-drag and those for the straight-wing configuration. The ratios were caused by a loss of lift and an increase in drag cranked-wing ailerons produce considerably more caused by the shock not being perfectly attached as well adverse yaw at 0 ° angle of attack than the straight-wing as to loss of lift from the lower-surface expansion and an configuration, as evidenced by the large negative values increase in drag from the additional volume added to the of AC t. The adverse yaw produced by the cranked-wing upper surface to accommodate model support hardware. ailerons will further drive the control power requirements The maximum lift-drag ratio for each configuration of the rudder. occurs at an angle of attack above 0°. Both the CFD predictions and experimental data showed that there were no Figure 53 shows the aileron effectiveness on lateral- significant performance degradations at off-design Mach directional stability with the ailerons deflected at 20 ° for numbers. The cranked-wing pure waverider model the cranked-wing fully integrated configuration. Rolling exhibited slightly better aerodynamic performance at the moment and yawing moment increments for a positive comparative Mach numbers studied than the straight- 20 ° elevon deflection and a +_20° aileron deflection are wing model. shown. A comparison of these data shows that a 20 ° elevon deflection has no effect on roll control power, Component build-up effects of waverider-derived indicating that interaction between controls is minimal at vehicles were examined by comparing experimental the Mach numbers studied here. force and moment data. The primary effect of individu- ally adding the canopy and the engine package was to 9. Concluding Remarks increase the drag of the configuration, thereby resulting in a degradation in aerodynamic performance. The aero- The aerodynamic performance and stability and con- dynamic effect of adding control surfaces was to increase trol characteristics of two Mach 4.0 waverider-derived the maximum lift-drag ratios slightly at each Mach hypersonic cruise configurations were examined. Experi- number studied. However, the aerodynamic performance mental force, moment, and flow-visualization data were of the controls-on configurations was significantly obtained for the two Mach 4.0 waverider configurations degraded from that of the pure waverider shape only by in both test sections of the Langley Unitary Plan Wind the addition of aftbody closure and the associated drag Tunnel (UPWT). The wind-tunnel models were designed production. The values presented for the pure waverider to allow testing of various configurations ranging from model show the performance of the waverider surface pure waveriders to fully integrated vehicles. The two only and do not include base drag. These results indicate pure waverider shapes were referred to as the straight- that additional consideration should be applied to wing pure and the cranked-wing pure waveriders. Exper- the design of control surfaces and aftbody closure in imental data as well as limited computational solutions waverider-based hypersonic cruise configurations. A were used to examine the flow field and aerodynamic control surface configuration with a less severe closure characteristics of the two pure waverider shapes, the angle or controls with blunt trailing edges may result in component build-up effects, and the aerodynamic and improved performance. Inclusion of the aftbody closure

14 in the optimization process for the waverider shape may 2. Corda, Stephen; and Anderson, John D., Jr.: Viscous Opti- also improve the performance significantly. mized Hypersonic Waveriders Designed From Axisymmetric Flow Fields. AIAA-88-0369, Jan. 1988. The characteristics of the fully integrated waverider- 3. Eggers, A. J., Jr.; Ashley, Holt; Springer, George S.; Bowles, derived hypersonic cruise vehicles were also examined Jeffrey V.; and Ardema, Mark D.: Hypersonic Waverider Con- by comparisons of experimental force and moment data. figurations From the 1950's to the 1990's. Proceedings of the The aerodynamic performance of each fully integrated 1st International Hypersonic Waverider Symposium, John D. waverider model (straight-wing and cranked-wing con- Anderson, Jr., Mark J. Lewis, Stephen Corda, and Isaiah M. figuration) was significantly degraded from that of the Blankson, eds., Univ. of Maryland, 1990. pure waverider shapes, because of the inclusion of aft- 4. Sobieczky, H.; Dougherty, E C.; and Jones, K.: Hypersonic body closure in the fully integrated configuration. The Waverider Design From Given Shock Waves. Proceedings of straight-wing fully integrated configuration provided the 1st International Hypersonic Waverider Symposium, slightly better aerodynamic performance than the John D. Anderson, Jr., Mark J. Lewis, Stephen Corda, and cranked-wing fully integrated model. The maximum lift- Isaiah M. Blankson, eds., Univ. of Maryland, 1990. drag ratios at Mach 4.0 were 4.69 for the straight-wing Takashima, Naruhisa; and Lewis, Mark J.: Waverider Configu- model and 4.56 for the cranked-wing model. The wave- rations Based on Non-Axisymmetric Flow Fields for Engine- rider concept also provides a uniform compressed flow Airframe Integration. AIAA-94-0380, Jan. 1994. field to the inlet, which offers potential advantages for airbreathing propulsion systems integration. The use of Rasmussen, Maurice L.: Waverider Configurations Derived different generating flow fields, such as osculating-cone From Inclined Circular and Elliptic Cones. J. Spacecr & and cone-wedge flow fields, may further improve these Rockets, vol. 17, no. 6, Nov.-Dec. 1980, pp. 537-545. characteristics. Furthermore, the results of this study 7, Kuchemann, D.: The Aerodynamic Design of Aircraft. Perga- have identified areas where design improvements could mon Press, Inc., 1978. enhance performance, such as control surfaces, aftbody closure, and propulsion system design. 8. O'Neill, Mary K. L.; and Lewis, Mark J.: Optimized Scramjet Integration on a Waverider. J. Aircr., vol. 29, no. 6, Nov.-Dec. Both fully integrated vehicles are longitudinally 1992, pp. 1114-1121. unstable across the Mach number range studied for 9. Corda, Stephen; and Seifert, E. Scott: User Information for unpowered conditions with the selected reference Maryland Axisymmetric Wave Rider Program (MAXIWARP). moment center. Additionally, locations were recom- Univ. of Maryland, Jan. 1989. mended for placement of the center of gravity in each 10. Anderson, John D., Jr.: Modern Compressible Flow--With configuration in order to ensure at least neutral stability Historical Perspective. McGraw-Hill, Inc., 1982. across the Mach number range. The pitch-control effectiveness of the elevons was judged to be unacceptable for 11. Anderson, John D., Jr.: Hypersonic and High Temperature Gas both configurations, and the data indicate that the pitch Dynamics. McGraw-Hill, Inc., 1989. control concept should be redesigned. The ailerons were 12. Beuerlein, David L.: Optimization of Waveriders to Maximize significantly more effective than the elevons for pitch Mission Performance. Proceedings of the 1st International control. The cranked-wing vehicle shows significantly Hypersonic Waverider Symposium, John D. Anderson, Jr., better lateral-directional stability than the straight-wing Mark J. Lewis, Stephen Corda, and Isaiah M. Blankson, eds., vehicle. The straight-wing configuration was unstable at Univ. of Maryland, 1990. angles of attack below 6 ° at Mach 4.0. The vertical tail 13. Cockrell, Charles E., Jr.: Interpretation of Waverider Perfor- has a significant stabilizing effect on directional stability, mance Data Using Computational Fluid Dynamics. AIAA- but very little effect on lateral stability. The ailerons are 93-2921, July 1993. also highly effective for the cranked-wing vehicle, but produce a significant amount of adverse yaw. 14. 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.

NASA Langley Research Center 15. Cockrell, Charles Edward, Jr.: Vehicle Integration Effects on Hampton, VA 23681-0001 Hypersonic Waveriders. NASA TM-109739, 1994. May 6, 1996 16. Bauer, Steven X. S.; Covell, Peter F.; Forrest, Dana K.; and McGrath, Brian E.: Analysis of Two Viscous Optimized References Waveriders. Proceedings of the 1st International Hypersonic Waverider Symposium, John D. Anderson, Jr., Mark J. Lewis, 1. Bowcutt, K. G.; Anderson, J. D.; and Capriotti, D.: Viscous Stephen Corda, and Isaiah M. Blankson, eds., Univ. of Optimized Hypersonic Waveriders. AIAA-87-0272, Jan. 1987. Maryland, 1990.

15 17. Braslow, A. L.; and Knox, Eugene C.: Simplified Method for 27. Dutton, J.; Herrin, J.; Molezzi, M.; Mathur, T.; and Smith, K.: Determination of Critical Height of Distributed Roughness Some Problems in Transonic Aerodynamics. AIAA-95-0476, Particles for Boundary-Layer Transition at Mach Numbers Jan. 1995. From 0 to 5. NACA TN-4363, Sept. 1958. 28. Penland, Jim A.; Hallissy, James B.; and Dillon, James L.: 18. Braslow, A. L.; Hams, R. V., Jr.; and Hicks, R. M.: Use of Aerodynamic Characteristics of a Hypersonic Research Air- Grit-Type Boundary-Layer-Transition Trips on Wind-Tunnel plane Concept Having a 70 ° Swept Double-Delta Wing at Models. NASA TN D-3579, 1966. Mach Numbers From 0.80 to 1.20, With Summary of Data From 0.20 to 6.0. NASA TP-1552, 1979. 19. Stallings, Robert L., Jr.; and Lamb, Milton: Effects of Rough- ness Size on the Position of Boundary-Layer Transition and on 29. Dillon, James L.; and Pittman, Jimmy L.: Aerodynamic Char- the Aerodynamic Characteristics of a 55 Degree Swept Delta acteristics at Mach 6 of a Wing-Body Concept for a Hyper- Wing at Supersonic Speeds. NASA TP-i 027, 1977. sonic Research Airplane. NASA TP- ! 249, 1978. 20. Steinbrenner, John P.; and Chawner, John R.: The GRIDGEN Version 8 Multiple Block Grid Generation Software. MDA 30. Penland, Jim A.; Edwards, Clyde L. W.; Witcofski, Robert D.; Engineering Report 92-01, 1992. and Marcum, Don C., Jr.: Comparative Aerodynamic Study of Two Hypersonic Cruise Aircrafi Configurations Derived From 21. Richardson, Pamela F.; McClinton, Charles R.; Bittner, Robert Trade-OffStudies. NASA TM X-1436, 1967. D.; Diiley, A. Douglas; and Edwards, Kelvin W.: Hypersonic CFD Applications for the National Aero-Space Plane. SAE 31. Small, William J.; Kirk.ham, Frank S.; and Fetterman, 892310, Sept. 1989. David E.: Aerodynamic Characteristics of a Hypersonic 22. McGrory, W. D.; Huebner, L. D.; Slack, D. C.; and Waiters, Transport Configuration at Mach 6.86. NASA TN-D-5885, 1970. R.W.: Development and Application of GASP 2.0. AIAA- 92-5067, Dec. 1992. 32. Watts, Joe D.; Olinger, Frank V.; Weidner, John P.; Johnson, 23. McGrory, William D.; Slack, David C.; Applebaum, Stuart K.; Sanders, Bobby W.; and Keyes, J. Wayne: Mach 5 Michael P.; and Waiters, Robert W.: GASP Version 2.2--The Cruise Aircraft Research. Proceedings of the Langley Sympo- General Aerodynamic Simulation Program. AeroSoft, Inc., sium on Aerodynamics, Volume 2, Sharon H. Stack, Compiler, 1993. NASA CP-2398, 1986, pp. 285-304.

24. Takashima, Naruhisa: Navier-Stokes Computations of a Vis- 33. Ellison, James C.: Investigation of the Aerodynamic Charac- cous Optimized Waverider. NASA CR- 189658, 1992. teristics of a Hypersonic Transport Model at Mach Numbers to 6. NASA TN D-6191, 1971. 25. Bushnell, D. M.: Supersonic Aircraft Drag Reduction. AIAA- 90-1596, June 1990. 34. Williams, John E.; and Vukelich, Steven R.: The USAF 26. Chapman, Dean R.: Reduction of Profile Drag at Supersonic Stability and Control Digital Datcom. Volume 1: User's Man- Velocities by the Use of Airfoil Sections Having a Blunt ual. AFFDL-TR-79-3032-VOL-I, Apr. 1979. (Supersedes Trailing Edge. NACA TN-3503, 1955. (Supersedes NACA AFFDL-TR-73-23, AFFDL-TR-74-68, and AFFDL-TR-76- RM A9H11.) 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 69.3 components 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 69.3 components 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

deg/sec Mach number Pss per unit velocity, ft/sec 2.3 0.119 4.0 0.095 4.63 0.095

17 Z Conicalshockwave

I I X

Waverider leading edge

Bottom surface (stream surface)

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

14 L/Dma x = 4(M + 3)/M "L/D Barrier" Conical-flow-derived waveriders 12

10 E 8

.... I , , , , I , , , , I .... I , , , i I , , , , I 5 10 15 20 25 30 Mach 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.

.... i

, : 9:

. ¢ :: 3,_ - _ _;_._ i_ -:_

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

Planforrn Oblique

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

Profile Rear

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

19 Figure5. Straight-wingpurewaveridermodelin UPWT.

2O E_

0 E

°_

21 _-- Expansionsurface

Figure7. Lowersurfaceofcranked-wingpurewaveridermodel.

22 0 E

p_ 'r_

LT_ 0 c_ \

23 0 E 0

0 E

N ._

t:m

o .o

24 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 Top view 6 22.791 -1.015 -1.252 7 22.791 -0.676 -1.254 8 22.791 -0.338 -1.255 I I I 9 23.552 -1.353 -0.964 I I 10 23.552 -1.015 -0.974 B.L. 0.0 I 11 23.552 -0.676 -0.974 I I 12 23.552 -0.338 -0.975 I l I

Short ramp (No-controls configurations)

M.S. M.S. 3.5" 21.779" 23.750" 1.971" W.L. 0.0

Taps 9-12--...._ o o o ool --_ 1.386" Taps 5-8_ o ooo1 t ° Pressure tap locations on nozzle surface Side view Rear view

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

Figure 10. Three-view drawings of expansion ramps.

25 Tap No. Model Sta. B.L. W.L. 13 24.313 -1.353 -0.724 14 24.313 -1.015 -0.731 15 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 -0.739 Top view 21 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

B.L. 0.0 Locations for taps 1 to 12 are given on previous page.

Long ramp (Fully integrated configurations)

_-- 3.5" M.S. M.S. 21.779 26.597 Taps 21-22 _ I 4_ 4.818" _ W.L. 0.0 _o----'_'_-- --_ W.L. 0.0 Taps 17-19_ Taps 13-16 oOoO_ [ .... - 2.01 Taps 9-12 _ o °oo °oi _ Taps 20,

Taps 5-8 _°°°°i I_ [ I located23,24 on Sting sl Taps 1-4 _ ' B.L. 0.0 sting sleeve x____ Pressure tap locations on nozzle surface Side view Rear view

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

Figure 10. Concluded.

26 Hinge line

.22'z---_

lnboard side ----3.760"------,- 10 o _'-_ __._ \ /---Trailing edge \ " _t 0.588-| "_ U _

_ "_Trailing edge

(a) Elevons. (b) Straight-wing ailerons.

27__ edge railing edge 3

Figure 11. Dimensions of elevons and ailerons.

27 Cranked // / wing

11.5" Cranked wing

Canopy --

\ Nozzle surface

Moment ref. center 9.9 _l 62.5 percent of model length Straight wing

1 ' Straight wing

Cylindrical uppper Vertical tail surface fairing i2.81,,

L 0.87" radius /" /" Elevon surface Waverider / Compression ,/ stream surface _jr surface ","_ Sidewalls -- Aileron surface ""--- Lower surface of engine module

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

28 PNS marching zone

7 TLNS global iteration zone

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

29 Shock Lowersurface ofmodel

(a) Vapor-screenphotographofbase.

Baseof P/Poo configuration 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 of straight-wing pure waverider model at M = 4.0 and cx = 0 °.

30 Shock

Lowersurface ofmodel (a) Vapor-screenphotographofbase.

P/P_ Baseof configuration 1.61

1.54

1.47

1.39

1.32

1.24

i_iii_iiiiiiiiii_!iii!_i!ii_ 1.17

i:i:i;!iii!ii:iiii!ili1.10 i i 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 of pure cranked-wing waverider model at M = 4.0 and _x= 0%

31 P/P,,_

2.18

2.05

1.92

1.79

1.66

1.53

1.39

1.26

1.13 / 1.00

(a) Cranked-wing pure waverider model.

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 surface of model

(a) Vapor-screen photograph at base.

P/Poo

Base of 1.28 configuration

\,\ 1.24 \

1.21

1.17

1.13

1.10 iii!iiiiiiii_iiii_iii_i_ii 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 of cranked-wing pure waverider model at M = 2.3 and o_ = 0 °.

33 Shock

Uppersurface of model

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

P/Poo

Base of 1.71 configuration 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 of cranked-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 C) .20

.15

.10

.05 8 0

-.05 C) Experimental -.10 dictions

-.15 C] MAXWARP prediction

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

.07 Q) Experimental (3 .06 CFD predictions (3 [] MAXWARP prediction .05 O

.r e- ._ .04 U

§ .03

g .02

.01

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

2 6 0

-2 ental .2 -4 _ _ CFD predictions /_) [] MAXWARP prediction -6 ,¢

-8 i i i i I i i i i [ .... I , i . . I .... I .... [ , i i . I .... I -.10 -.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 performance 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 1 -.05 _ CFD predictions -.10 [] MAXWARP prediction -.15

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

.07 C) Experimental

.06 dl& I, CFD predictions O [] MAXWARP prediction .05 O

,..r _= .04

§ .03

_ .02

.01

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

Lift coefficient, C L 8

0 013

-2 rimental ,.-1 -4 _ CFD predictions

-6 / [] MAXWARP prediction

-8 .... I .... I .... I .... I .... I .... t .... I .... I -.10 -.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 performance of cranked-wing pure waverider model at M = 4.0 and Reynolds number of 2.0 x 106 per foot.

37 .30 .25 [] Straight wing [] .20 ,..a.15

•e-," .10

.05

o Cran_OZ)(__O -.05

-.10

-.15

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

.07 [] Straight wing [] .O6 O Cranked wing

.05 © r,.) [] ..r ¢.., Q .N .04 []

8 .03

.02

.01

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

Lift coefficient, C L

% O_U_)_) _)_) 4 [] _t_

2 0

[] 0 0 [] Straight wing # -2 [] 0 Cranked wing ,..1 -4 O [] -6

-8 .... I .... 1 .... i .... I .... I .... I .... 1 .... I -.10 -.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 = 4.0 6 and Reynolds number of 2.0 x 10 per foot.

38 .4 D D .3 D > %

¢o .2 .3

ca "3 .1

0 <3 Mach 3.5 j D Mach 4.02.3

-.1 * Mach 4.2 D © Mach 4.63

-.2 i , ,_l , , -4I, ,,I -2 , ,L[t 0 _IILL_IIL2 4 ' 6I ' ' ,ll 8 , ,I 10, , , 12I

Lift coefficient, C L

10 10.0

8 _> t> l> D 9.5 6

4 __ t>E>D 9.08.5 ,.d O 6 2 80

•_ o <> D Mach 2.3 _ 7.5 ,= < Mach 3.5 "_ 7.0 -2 © "7 -4 <_ _ Mach 4.0 6.5 © 0

45 6.0 © <_ ©Mach+ Mach 4.634.2 -8 DD 5.5

-10 , , , I n i , I h J , I _ , , l , _ , I , * , , J , i i i L i I , i _ i i -. 1 0 . 1 .2 .3 .4 5"02.0'' '2'.5''' 3.0 '3t.5 ' 4.0 '415' '5'.0

Lift coefficient, C L Mach number, Moo

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

39 .4 .07 [] 13 Mach 1.6 b [] b b Mach 2.3 .3 .06 <1 Mach 3.5 [] b ,.d .05 Mach 4.0 .2

..J * Mach 4.2 e- [z []_> r_ .04 © Mach 4.63 b .1 [] Mach 1.6 .03 [] Q _ Mach 2.3 8 0 et0 [] <_ Mach 3.5 t_ ,.d .02 _ Mach 4.0 [] --.1 + Mach 4.2 .01

Mach 4.63 --.2 45 -4 -2 0 2 4 6 8 10 12 0.2 -.1 6 .! .L .3 .;

Lift coefficient, C L

14 12.0 11.5 12 [] [] 10 11.0 8 10.5 1_ Ib _> b U [] 6 [] [] 10.0 bb 4 9.5 6 2 9.0 O 0 8.5 [] Mach 1.6 E -2 8.0 D Mach 2.3 t_ -4 7.5 -6 <1 Mach 3.5 7.0 O J O0 -8 Mach 4.0 6.5 0 6.0 -10 [] [] + Mach 4.2 -12 _- O Mach 4.63 5.5 -14 -2 -. 1 0 . l .2 .3 .4 5°1_5'''2'.6'''2'.5'''3'.6"'315' :_Y i5.... 5'.0

Lift coefficient, C L Mach number, Moo

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

4O .30

.25 (_(_ .20

.15 (_ ,.d

"_e- .10

.05 __1. 6 8 o 5XlO per foot

,d -.05 (_(_ 0[] 3.02.0 x× 106 perper 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 0 r_ 0 3.0 × 106 per foot 0 .05

e., 0 ._ .04 0 8 .o3 d

02.01

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

Lift coefficient, C L 10

6 4

6 2 0 0 0 /_ 1.5 x 106 per foot -? -2 eq [] 2.0 × 106 per foot ,.d -4 0 0 3.0 x 106 per foot

-8

--lO , , , I .... I .... I .... I . , . t I .... I .... I .... I -.10 -.05 0 .05 .10 .15 .20 .25 .30

Lift coefficient, C L

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

41 .30

.25 ©0 .20 0 0 ,2 .15

= .10 .o 0 .05 0 0 O O /X 1.5 x 106 per foot {_ [] 2.0 x 106 per foot .2 -.05 O _ O 0 3.0 x 106 per foot -.10

-.15

-.20 ik*llhllLiLl,,,I,,,I,,,l*,,I,,,I,*,l,,,I -6 -4 -2 0 2 4 6 8 10 12

.07 /_ 1.5 x 106 per foot .06 [] 2.0 x 106 per foot

.05 O 3.0 x 106 per foot

,,d ._ .04

_ .03

.02

.01 _ O O#OOOO 0 , , , , i .... i .... I .... i .... i .... i .... i .... I -.10 -.05 0 .05 .10 .15 .20 .25 .30

Lift coefficient, C L 10

4 O t_¢_f_t _

6 2 0 {_ /X 1.5 x 106 per foot -2 [] 2.0 x 106 per foot ,.-1 -4 0 3.0 x 10 6 per foot -6

-8

-10 .... I .... I .... I .... I .... I .... I .... i .... i -.!0 -.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 .030 [> Mach 2.3 _> .025 <3 Mach 3.5 D ,O L) .020 0 Mach 4.0 * Mach 4.2 b b

~~.010~~

~~O .005~~

~~_0 e- 0 AJ~~

~~-.005~~

~~-.010 -8 -6 -4 -2 0 2 4 6 8 10 12~~

~~(a) Straight wing.~~

~~.030 [] Mach 1.6 [] b .025 t> Mach 2.3 [] E> Mach 3.5 .020 [] t> q 0 Mach 4.0 [] b 0 .015 <_ ÷ Mach 4.2 [] I> .q (_~~

~~¢.J ,fi .010 0 Mach 4.63 [] t> © E O .005 [] °;P~~

~~-.005~~

~~-.010 _,,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 Stable b 0 E ,_ -.0001~~

~~'_ -.0002 O> _ _ _ _a_.3 _ Mach 4.0 -.0003 b B 0 Mach 4.63~~

~~--.000_ -6 -4 -2 0 2 4 6 8 10 12~~

~~(a) Straight wing.~~

~~.0008~~

~~[] .0006 b r,.) [] ,; .0004 [] 8 _" .0002 lStab,o [] _ e o E d_ b [] Mach 1.6 .- -.0002 Mach 2.3~~

~~Mach 4.0 --.0004 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~~

~~(b) Cranked wing.~~

~~Figure 28. Yawing moment characteristics of pure waverider configurations.~~

~~44 .0016~~

~~.0012 D" Mach 2.3 e,_ _ Mach 4.0 d .0008 _(_ O Mach 4.63~~

~~_ .0004 g> % o~~

~~E 0 0 0 P 0 0 I Stable O ._ -.00O4 D~~

~~-.0008~~

~~-.0012 _'I'J'I''LI'''I'''IllLiIIII_I_I -2 0 2 4 6 8 10 12~~

~~(a) Straight wing.~~

~~.0008~~

~~.0004~~

~~e_ _- o o" _ _ _ 8 8 i Stable ._ -.0004 -r" -.0008 S e-, E> -.0012 [] o P _-.0016 [] [] Mach 1.6~~

~~-6 -.0020 Mach 2.3 [] e¢ Mach 4.0 [] -.0024 0 Mach 4.63~~

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

~~(b) Cranked wing.~~

~~Figure 29. Rolling moment characteristics of pure waverider configurations.~~

~~45 .30 .07 Moo = 4.0 .25 O .06 [] .20 0 [] ,..1 .15 r_ .05 r_ 0 ,.z .10 [] e- .__ .04 ej .05 _ o 8 .o3 []o~~

~~._ -.05 .02 -.10 .01 _ (2)[] Canopy offon -.15~~

~~-'20-8 -6 -4 -2 0 2 4 6 8 10 12 o_o.I ...... -.05 0 .05 .10 .15 .20 .25 .30~~

~~Lift coefficient, C L~~

~~8 Moo = 4.0 10.0~~

~~9.5 6 0 Canopy off 9.0 [] Canopy on 4 0 8.5~~

~~2 [] _ 8.0 8 6 0 o S 75 E 7.0 -2 0 0 6.5 D 0 o ,.d [] [] -4 [] @~~

~~-6 _S_IS)DO O[] Canopy offon 5.56.0~~

~~-8 .... I .... I .... i .... I .... I .... i .... i .... I -.10 -.05 0 .05 .10 .15 .20 .25 .30 5.020• ...... 2.5 3.0 3.5 4.0 4.5 5.0~~

~~Lift coefficient, C L Mach number, Moo~~

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

~~46 Mo_ = 4.0 .30 .07 Moo = 4.0 .25~~

~~.20 O0 ° .06~~

~~..a .15 0 © .05 {3 ¢0 ¢.) .10 ,.t e- O0 ° e-, 0 0 ._ .04 0 .05 © 8 0 _ .03 O0 O0 © ..a -.05 © O 0 0 Canopy off _ .02 O 0 0 Canopy off -.10 [] Canopy on~~

~~mO 1 /DO ° [] Canopy on -.15~~

~~-.20 °_ ...... , .... , .... , ...... ' -'6"S "-'2...... 0 '2'' 4 6 8 'l_O'' 1_2 20.l -.05 0 .05 .10 15 .20 .25 .30~~

~~Lift coefficient, C L~~

~~M = 4.0 14~~

~~13 O Canopy off~~

~~O0 0 12 [] Canopy on 11 D 23 .1 2 © 6 12 E 0 10 E O 9 i -2 ,-.1 8 -4 ID © Canopy off 7 [] Canopy on 6~~

~~--8 .... I .... I .... i .... I .... I .... I I I L I I J _ L , I -.10 -.05 0 .05 .10 .15 .20 .25 .30 5L5...... 2.0 2.5 3.0 3.5 4 ,.0 ...... 4.5 5.0~~

~~Lift coefficient, C L Mach number, M~~

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

~~47 Moo = 4.0 .30 .07 Moo = 4.0 .25 .06 .20~~

~~,..z .15 .05 r,.) "-: .10 .,..r % ID .04 .05 % o 8 .03~~

~~,.-: -.05 .02 -.10 u- [] Engine on [] Engine on .01 IODdz3ODO%% _%E3Ofi_D[_IZ_ O Engineoff -.15~~

~~-.20 0 -8 -6 -4 -2 0 2 4 6 8 10 12 -.10 -.05 0 .05 .10 .15 .20 .25 .30~~

~~Lift coefficient, C L~~

~~8 Moo = 4.0 10~~

~~6 0000 Or-,, O Engine off Drl o o [] ,--, [] Engine on 4 [] _c_D o O 2 (D° 6 [] E 0 4P O Oo -2 [] 0 ,.d -4 % [] [] Iz] [] [] O O Engine off -6 [] 0 0 [] Engine on~~

~~- i _ , i i .... I .... I .... t _ i _ i I i , i i I i i i i i .... I -8 4 .... ] , , , t t L , , i i .... J .... i .... J -.10 -.05 0 .05 .10 .15 .20 .25 .30 2.0 2.5 3.0 3.5 4.0 4.5 5.0~~

~~Lift coefficient, C L Mach number, M~~

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

~~48 Moo = 4.0 .30 .07 Moo = 4.0 .25 D~~

~~.20 .06 © [] ,.d .15 .05 DO~~

~~.10 e-, .04 m© "U .05 O _2 DO 8 .03 o _3_3 m© -.05 O _ 0 Engine off _ .02 m© 0 DO O Engine off -.10 [] Engine on 5_ [] DO [] Engine -.15 .01 (_ (_ (_ (_(]_£0 0 on~~

~~-.20 0 , , , , I , , J , I , L , , I , , , , I .... I _ + L k l i i i i I i L i _ _ +tI'''I'''JL'+I'+,IL,,I,++I,,,L,,,]+-6 -4 -2 0 2 4 6 8 0 ,,I 12 -.!0 -.05 0 .05 .10 .15 .20 .25 .30~~

~~Lift coefficient, C L~~

~~8 M°° = 4.0 10 Q 0 Engine off 6 00_00 [] Engine on 4 % [] [] [] D(_ (_ (_(_ (_ O~~

~~2 _ 8 d [] B 0 [] _ 7 [] O © m -2 Oo [] 6 0 -4 []O D_ 0 Engine off [] [] [] [] -60 w [] Engine on~~

~~-8 + , l .... i .... i + + , , [ i , , , i .... i , , i t [ i i , i i 4 , t, , I .... I .... _ i + i i I L i + + t,, + t I,,,, I -.10 -.05 0 .05 .10 .15 .20 .25 .30 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0~~

~~Lift coefficient, C L Mach number, M~~

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

49 Moo = 4.0 .30 .07 Moo = 4.0 .25 O OAO/x .06 .20 O No controls (no base drag) AS] L_ 0 ° controls .15 0 ¢_OA _ .05 [] No controls (base drag) _3 L) L) O '_ .10 ¢- _ .04 "U _yO .05 o 8 .03 _0 _0 ,2 -.05 No controls (no base drag) _ .02 zx_ /_ 0 ° controls _y_ _t_O _0 -.10 [] No controls (base drag) .01 _O--d 0 0 _ -.15

~~-6 -4 -2 0 2 4 6 8 10 12 o_.1o...... -.05 ..... 5...... 10 .15 .20 .25 .30~~

~~Lift coefficient, C L~~

~~8 M = 4.0 9.0~~

~~8.5 6 O 0 8.0 4 0 No controls (no base drag) 7.5 /k 0 ° controls (Z 0 2 [] No controls (base drag) 6 7.0 E 0 /k 6.5 O 6.0 -2~~

~~,.d 5.5 -4 O 5.0 Z_O0 0/x 0Noo controlscontrols (no base drag) O -6 [] No controls (base drag) 4.5~~

~~-8 4.0 -.10 -.05 0 .05 .10 .15 .20 .25 .30 2.0 2.5 3.0 3.5 4.0 4.5 5.0~~

~~Li_ coefficient, C L Mach number, Moo~~

~~Figure 34. Effect of undeflected control surface addition on aerodynamics of straight-wing pure waverider configuration with canopy and engine package attached,~~

5O .3O Mo_ = 4.0 .07 Moo = 4.0 [] .25 .06 O No controls (no base drag) _ O .20 /_ 0 ° controls _O [] No controls (base drag) ,.d .15 6o '_ .10 _@@_q u._='_.05 .04 "U .05 65o _ 0 @ @ Q _ .03 45o Q O No controls (no base drag) ,_ -.05 A 0ocontrols _ .02 E_ _Dso_° O @ -.10 [] No controls (base drag)

~~-.15~~

~~-.20 -6 -4 -2 0 2 4 6 8 10 12 °-.1o'-.;5; ...... 05 .1o _5 So .25 .3o~~

~~Lift coefficient, C L~~

~~8 Mo_ = 4.0 8.5 0 No controls (no base drag) 6 8.0 /_ 0 ° controls ,-., 0000,_ 7.5 O [] No controls (base drag)~~

~~4 o__K:s_t_ @ 7.0 ,.d O d 6.5~~

~~o _9 .--i 6.0 i. 5.5 -2 []_ O ..d 5.0 []/x 0 No controls (no base drag) O -4 _ O z_ 0 °controls 4.5 /x _ A -6 O O [] No controls (base drag) 4.0~~

-8 _ , . i .... ] .... i , . , I i i L , , i .... i .... [ i i i i [ 3.5 t , , , I , * , , I .... I , i i i I i i j L I i i L _ I _ i _ i I -.10 -.05 0 .05 .10 .15 .20 .25 .30 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0

~~Lift coefficient, C L Mach number, Moo~~

~~Figure 35. Effect of undeflected control surface addition on aerodynamics of cranked-wing pure waverider configuration with canopy and engine package attached.~~

~~51 .30 D Moo = 2.3 .09 D Moo = 2.3 .25 M =4.0 oo .08 Moo = 4.0 .20 O Moo-- 4.63 O Moo = 4.63 i> D _8 .07 ,-.I .15 .06 "_" .10 CA -_ .05 .05 8 _ .04 0 .03 ,_ -.05 .02 -.10 D -.15 .01 --v_~~

~~-.20 0 + + _ I i + + I i L + I h _ I I I -I -6 -4 -2 0 2 4 6 8 10 12 -.2 -.1 0 .1 .2 .3~~

~~Lift coefficient, C L~~

~~8 D Moo = 2.3 8.0 Moo = 4.0 7.5 6 O Moo = 4.63 7.0 4 6.5 2 6.0 6 O> E '_ 0 @ = 5.5 ex(_ s.0 -2 0 0 0 ,.d _Q 4.5 -4 4.0~~

~~-6 3.5~~

~~I + i i I -8 i i h i , + 3"02 d ' ' '2'.5.... 3t.O ...... -.2 -.1 0 .1 .2 .3 . 3.5 4.0 4.5 5.0~~

~~Lift coefficient, C L Mach number, Moo~~

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

52 .4 [] M = 1.6 .10 [] M = 1.6 /x Moo = 1.8 .[3 .09 /x Moo = 1.8 .3 V M:20 .08 V Moo = 2.0 [] l> M =23 &_ I_ _,_ ,..J .2 .07 D Moo = 2.3 '_ Ga O M =40 A_i7 _z9 M =-4.0 oo • ^n t_ _))U_) .06 oo .1 ¢- 'ra O Moo = 4.63 _ E .05 L.), 0 ¢J .04

~~t_ ,d .03~~

~~.02~~

~~--.2 .01~~

~~--.3 0 , , , I . . . I i i _ [ _ , i I i _ i I , , _ I , , , I .... -6 ...... -4 -2 '''l'''''''jO 2 4 ...... 6 8 1'0''1'2 -.3 -.2 -.1 0 .1 .2 .3 .4~~

~~Lift coefficient, C L~~

~~[] Moo = 1.6 8.0~~

~~Moo = 1.8 7.5~~

~~V Mo =2.0 7.0 l> Moo = 2.3 6.5 2 Moo = 4.0 6 6.0 .,.._ O Moo = 4.63 E 0 5.5~~

~~5.0 0 0 ,--i -2 4.5 ) 0 0 0 4.0~~

~~3.5~~

~~3.0 -.3 -.2 -. 1 0 .1 .2 .3 .4 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 Lift coefficient, C L Mach number, Moo~~

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

~~53 M = 4.0 M=4.0 .30 .07~~

~~.25 .06 [] Straight wing (_ .20, 0 8 0 Cranked wing (5] ,_ .15 r_ .05 0 o e_ .10 e- 0 o .N .04 .05 0 Oo ° 8 .o3~~

~~-_ -.05 .02 [] Straight wing -.10 0 Cranked wing .01 -.15~~

~~,,I,,,I,,,I,,,I,,,I, ,,I,,,I,,,I,,,I,,,I 0 illll_j,,I .... I .... I .... 1 .... I .... I .... I -'20-8 _5 -4 -2 0 2 4 6 8 10 12 -.10 -.05 0 .05 .10 .15 .20 .25 .30~~

~~tX Lift coefficient, C L~~

~~M,, ° = 4.0 12 [] Straight wing 11 6 0 Cranked wing 10 4 (9 9 2 6 0 E 8 0 e_D (9 .e_ 7 -2 0 [] Straight wing ,J OF -4 (ff_3 0 Cranked wing )o ° 0 0 0 -6~~

~~-8 .... I .... i .... I .... t .... i .... i .... i .... i -.10 -.05 0 .05 .10 .15 .20 .25 .30 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 Lift coefficient, C L Mach number, M~~

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

~~54 M = 4.0 .4 .09 M = 4.0~~

.3 .08 0 Pure waverider (no base drag) 0 Pure waverider (no base drag) .07 A Pure waverider (base drag) A .2 [] Fully integrated DO .06 z5 ..7 []0 e- .1 .05 ¢,a _0 0 /_ Pure waverider (base drag) _20 .04

~~o.0 t_ .03 ,.-1 --.1 ° .02~~

~~-.2 .01~~

~~, , I , J , I I I I -'3-8'''-6J'''-4''''-_2'''O_'''2''''4L'''6''''8J'''I_(I ''1_2 0_.2 --,1 0 .1 .2 .3~~

~~Lift coefficient, C L~~

~~M = 4.0 8 9.0~~

~~8.5 0 Pure waverider (no base drag) 6 O000Oo O 8.0 A Pure waverider (base drag) [] Fully integrated 4 7.5~~

~~7.0 2 @ 0 [] d ,..1 6.5 O0 -.--r E 0 6.0 0~~

~~t_ 5.5 -2 5.0 ,.d _0 /x /x /k A -4 [] [] eq_ 0 Pure waverider (no base drag) 4.5 0 A Pure waverider (base drag) 4.0 -6 0 0 [] Fully integrated 3.5~~

~~-8 , J J I , , _ I , _ , I _ , , I J , , I 3.0 t,,,l .... I .... l*il,lliiilliltl --.2 -.1 0 .1 .2 .3 2.0 2.5 3.0 3.5 4.0 4.5 5.0~~

~~Lift coefficient, C L Mach num_r, M~~

~~Figure 39. Comparison of aerodynamic performance of straight-wing fully integrated and straight-wing pure waverider configurations.~~

~~55 M o = 4.0 Moo : 4.0 .4 .09~~

.08 .3 .07 0 Pure waverider (no base drag) .2 A Pure waverider (base drag) ,..1 @@@@ .06 [] Fully integrated ,.r I:_° .1 @@ @@ o .05 "5 _0 _0 0 8 .04 8 I_¢D QO _0 .03 ,.d ®@ _0 --,l (2) Pure waverider (no base drag) _0 .02 A Pure waverider (base drag) @_ _0 --.2 [] Fully integrated .01 O__n 0 _00000_ 0 1 I -'3_8' I 'I'--6I ' ' '--4I ' ' '--12' L* 0I * ' k2L _ ' ' 4L' ' ' 6I ' ' ' I8 ' ' ' lOI _ ' _12 --.2 -.1 0 .! .2 .3

~~O_ Lift coefficient, C L~~

~~Moo = 4.0 8 12~~

11 6 000000 - (3 Pure waverider (no base drag) 10 4 Zk Pure waverider (base drag) g_ 9 [] Fully integrated 2 (3 6 r_ E 8 0 g_ .e. (D 7 0 O0 -2 rq 0 6 ,.d -4 _t A_(3 (3A Pure waverider (base(no basedrag)drag) 5 [] [] /k /xA ix [_ [] [] [] 0 [] Fully integrated 4 (3(3 -8 L I I _ , J I I I 3 i i i I b _ i , I .... I .... I .... I .... I .... i --.2 -.1 0 .1 .2 .3 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0

~~Lift coefficient, C L Mach number, M~~

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

~~56 Present fully integrated cranked-wing waverider • Reference 28 • Reference 29 • Reference 30 • Reference 31 • Reference 32 6 • Reference 33~~

~~E = 4 e.~~

~~2 3 4 5 6 7 8 Mach number, M._~~

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

~~57 .020 M_ = 2.3 M =4.0 .015 ¢0 0 M =4.63 b t- b .010~~

~~.005 E O ,_ 0~~

~~_= -.005~~

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

~~(a) Straight wing.~~

~~.025 [] Moo = 1.6~~

~~A Moo = 1.8 [] .020 v M_ = 2.0 [] D M = 2.3 [] z_v E] &v .015 [] O M,_ = 4.0 [] Av 0 M =4.63 0 DO A _V 0 ,_ .010 E DD_DDO _b_b©QQ _, .oo5 .=. bbv_ _ ¢.., bbbb_ h2 0~~

~~-.005 ...... -6 --4 -2 0 2 4 6 8 _d"12~~

~~(b) Cranked wing.~~

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

~~58 Mo_ = 2.3 Mo_ = 2.3 .3 .025~~

[] .020 [] [] .2 [] _A = 0% fiE = 0° r,.) [] ,_ .015 _A = 0°' _E = 200 [] ¢- 0 ,.d [] o "0 @ fA = 200' 8E = 200 [] [] .010 [] GO © t- 8 •D D _ 0 ,_ .005 8 0 oo0 o 0 ••D ©0 ,e .=- -.005 00© 00000 -.1 _ 5 A = 0 °, 8 E = 20 ° ._, O 5 A = 20 °, fE = 200 -.010 ooOO 000

~~4).15 '''l'''l'''l'''l'L'l_'l'''llllLIL_lLlll -'2-8'"] '1--6'''1'' -4 '- [2'_'1'''10 2 _ ''] 4 '''_ 6 '''1"' 8 10_' ' 12 -8 -6 -4 -2 0 2 4 6 8 10 12~~

~~M = 4.0 M• = 4.0 .O25~~

~~.020 [] fA = 0% fE = 0°~~

,.s .015 e- _A = 0°' _E = 200 _d [] 08 A=20 ° ,8 E=20 ° [] [] .OlO e- []DaD•O© 0 02 N .005 8 aDDS• _ 000 o 0 [] 0000 O0 O 0 -2 t@888 , =oo _O O 0000 -4 _ OSA = 0°, fE = 20° _ -.005 O00@QO0 00°0 0 fA = 20% _E = 20° -6 -.010

~~-.015 -8 -6 -4 -2 0 2 4 6 8 10 12 -8 -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 integrated configuration.~~

~~59 M,, ° = 4.63 M = 4.63 .3 .O25~~

~~e_ .1 []E3 .010 [] O0 0 o•[]D_© i .005 oog_,O0 0 _ _&__oo ___oo ,E o .1 ooooooooO°°°° o _A: 0%_: 20° .=.-.oo5e. -.1 ©Sa : 2°°, _E = 20° _-.010~~

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

~~(b) M.. = 4.63.~~

~~Figure 43. Concluded. Moo= 1.6 .4 .040 Moo = 1.6~~

.035 .3 .030 0 o° 0 .025 ,-d .2 00000 [] O000OO [] ...$ .o20 .1 8 .015 DDDD_DODG DD_D 0 o .010 0 O E .005 --.1 0o So ,.= 0 IS 8A = 0°, fiE = 0° 8o _A--°°,_ --2°° ._ -.005 --.2 t_ + 8 A = 0 °, 8 E = 20 ° o o 8A=0o,fie=_200 -.010 O 8A = 0°' 8E = -20° -.015 '''''''''''''''J'''''''''''''' ''',_,,,_ -" -6 -4 -2 0 4 6 8 10 12 -8 -6 -4 -2 0 2 4 6 8 10 12

~~Moo = 1.8 M = 1.8 .4 .040~~

~~.035~~

~~.030~~

~~'_ .025 0 o ,..A k) 0 o .020 e- O00000oO0 Oo _D .015 13 o .010~~

~~O o .005 ,.d -.1 _ D (SA = 0°, 8E = 0° .= 0 E3 8A = 0°, 8E = 0°~~

~~-.2 _ +8A = 0°' fie = 20° -.005 8 A = 0% fiE = 20°~~

~~08 A = 0 °, 8 E =-20 ° -.010 0 8 A = 0% fie : -20 °~~

~~-.015 t,.I,,,I ,,.llllll,.i,_ ,i,,_l.,,i,,.lllll -8 -6 -4 -2 0 2 4 6 8 10 12 -8 -6 -4 -2 0 2 4 6 8 10 12 Ot~~

~~(a) M.o = 1.6 and 1.8.~~

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

~~61 Moo = 2.0 Moo = 2.0 .4 .040~~

~~.035 .3 rj_ .030~~

~~.025 .2~~

~~.020 OOQOQOOOQO 00000 .1 .015~~

~~.010 DE)[_ 0 00DC]OD OC] FIE) C] 0~~

~~_ .005~~

~~--.1 .= 0 m _A = 0°, 8E = 0° r., O _5A 0 °, _)E = 20° ._ -.oo5 <_ 8A = 0°' _E = 20° --.2 O 6 A = 0% _E = -200 -.010 O8 A = 0% qY)E = -20 °~~

~~-'3--8'''_ ...... -4 -'2 ...... 0 2 4 _,,,_ ...... 10 12 _.015_8,,,,,,,L,,,L2,..,*.,,,,,J,,,,,,,-6 -4 - 0 2 4 6 8L,.._..., 10 12 ot~~

~~M =2.3 M_o = 2.3 .4 .040~~

~~.035 .3 .030~~

~~.025 ,-1 .2 .o _9 .020 ,,d I:I .1 .015 oooooooooOO°°!°°° .OlO 13131313 00 ° 0 0 896:f 8 g gg 88 _ .005 00[3013 rq ° ° u.... 0<_0 .1 --,l [] a A = 0 o, i_E = 0 o ._ 0 OO OOOOOOOv[]aA=O°,_=O °~~

~~_A = 0°' $)E = 200 _ -.005 O _iA = 0 °, i_E = 20 ° --.2 O _A = 0% _ = -20 ° -.010 O BA = 0% _5E = -200~~

~~-.015 -'38' -6 -4 -2 0 2 4 6 8 10 12 -8 -6 -4 -2 0 2 4 6 8 l0 12~~

~~(b) Mo. = 2.0 and 2.3.~~

~~Figure 44. Continued.~~

~~62 Moo = 4.0 = 4.0 .4 .040 Moo~~

~~.035 .3 [] 8 A = 0 o, 8 E = 0 o .030 ¢9 8 g = 0 °, 8 E = 20 ° .2 .025 ,..j 8_ o 08 A = 0 °, 8 E =-20 ° L_ .020 .1 _0000_ 8 .o15~~

~~G 8 A = 0% 8 E = 20 ° ._-.oo5 --,2 e_ O8 A = 0 °, 8 E =-20 ° -.010~~

~~, I , , , i , , , , , , I , , , I , , , J .... i -.015 -'38' ' _ _- ''2'- ()' 2 4 6 8' '1'() '1'2 ,,,r,,,I,,,t2,,,J,,,I,,,I,,,I,,,I,,,I,,,i-6 -4 - 0 2 4 6 8 10 12 o[~~

~~Moo = 4.63 Moo = 4.63 .4 .040~~

~~.035 .3 .030 _ 8A = 0°, 8E = 0°~~

~~8 A = 0 °, 8 E = 20 ° ..d .2 .025 08 A =0 °, 6 E =-20 ° .020 e_ ,015~~

~~0 _ 00_ .010 8 oooooooo0 O ,_ .005 ..d -.1 [] 8A = 0o, $_E = 0o ,_ 0 BSB8888_l_o_°°_ e- 8 A = 0 °, 8 E = 20 ° -.005 --,2 e_ 0 8 A = 0 °, fiE = -20° -.010~~

~~_.015_,,,,,,,_,,.,,,,_,,,,,,,_,,.,,.,_,,,,,,,, -' 3_8,,, -6t,,.,,.,L2,,,-4 - 0I,,, 2I,,,J,,,t,,,,,..,,,,1,24 6 8 10 -8 -6 -4 -2 0 2 4 6 8 10 12 ot O[~~

~~(c) M.. = 4.0 and 4.63.~~

~~Figure 44. Concluded.~~

~~63 M = 2.3 M. =2.3 .4 .040~~

~~.035 .3 [] _A = 0°' _ = 0° r_ .030 _A = 0°, _E = 20°~~

~~,.d .2 .025 O _A = 0°' _E = -20° O E .020 O O .1 888888 8888 8 .015 OOQOQO00QO0000000 .010 0 E _ .005 CDCDDDDODODDDDO00 ...1 --.1 .= 0 o 6A = 0°, _E = 20° ._ -.005 --.2 © 6A = 0o, _ = _20° -.010~~

~~-.015 -" ,,,i,,,i,,,l,,,i,,,i.,.i,,,),,,i,,,i,,,i-6 -4 -2 0 2 4 6 8 10 12 -6 -4 -2 0 2 4 6 8 10 12 ¢t~~

~~Moo = 4.0 M._ = 4.0 .4 .040~~

~~.035 [] 8A = 0°, 8E = 0° .3 .030 8 QSA = 0°, 8E = 20° = .025 O _A = 0°' _E = -20° .1 .2 .020 .r .1 .015 "0 OOO00o o .010 0 0 8 E E, .005 00000000000000888 ,.d DFIr_ mDC)[] O DDI_ _ _ r_ .= o~~

~~O8 A = 0 o, _ = 20 ° ._ -.oo5~~

~~--.2 O 8A = 0 °, _5E = -20 ° -.OlO~~

~~-.015 -'38' --6 -4 -2 0 2 4 6 8 lO 12 -8 -6 -4 -2 0 2 4 6 8 10 12 o_ O_~~

~~(a) Mo. = 2.3 and 4.0.~~

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

~~64 Mo= = 4.63 .4 .040 M o = 4.63~~

~~.035 .3 [] 8A = 0°' _E = 0° .030 r,.) 6A = 0°' _E = 20° .2 .025 O fiA = 0°' fiE = -20° .020 .1 L_ .o15 .010 0 8 E .=. .005 ,.-I --.l °°°°oooooo9888899 O 8 A = 0 o, _ = 0 o .n 0 88~~

~~--.2 + _A = °°, _ = 2°° :_-.oo5 o 8A = 0%_ = -20 ° -.010~~

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

~~(b) M_ = 4.63.~~

~~Figure 45. Concluded.~~

~~65 Moo = 1.6 M, ° = 1.6 .4 4_5 D 8A = 0°, 8E = 0° .040 .3 O _A = 0°' 8E = 20°~~

~~.015 n n D FI E313r7 _D DFII313 fin .010 --.1 <_o []_A_-oo,__-oo . © _o ° <>8A_-oo,__-2oo --.2 k- .005 _©_©_©_©©_<_ O 0 8A = 0°' 8E = -20° 0~~

~~--.3 --.005 , ,_1, , ,I,_,lt J ,I,, ,iJ J,IJl_ll,,E_,,Ih, ,1 -6 -4 - 0 2 4 6 8 10 2 -8 -6 -4 -2 0 2 4 6 8 10 12~~

~~M = 1.8 Moo = 1.8 .41 .040 D 8A =0°, 8E =0°~~

~~.3 8 A = 0 o, 8 E = 20 ° R .030 O8 A = 0% 8 E =-20 ° .2 o= .025 ,1 "0~~

~~c- '= .020 .1 OOOo000000OO00000~~

~~o 8 0 0,,.010 [] D[] [] n D rIC_C)_C) D(_ [] C]N R ,.d --.1 .oo5 _ _:oo,_:2oo _ o --.2 o_:oo,_:-2oo -005~~

~~--.3 ,,',,,',,,_,,,_,,''_''''''''''''''''''_ -.010 -8 -6 -4 -2 0 2 4 6 8 10 12 -8 -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 reference center at 59 percent of body length.~~

~~66 Moo = 2.0 Moo = 2.0 .4 .040~~

~~.035 O 8A = 0°, _E = 0°~~

E .010 8 0 E (3 CI D [3 CI C][31_ C]I_I V113 D V1CI rl FI o .005 .E _ o 00000'_00©©0.0©,00©0 -.2' _JA =0 ' _E 20o -.005 (2) _iA = 0% gE = -20° -.Ol0 ,,i,,,_,,,.,,,t,,._.,,,,,,,,,... ,_,,,i -'--83 , ,1_6i , , ,--4[ , , ,-2i , , , _, , , 2i , , , 4i , , , 6i , , , 8i , , , 10i , , , 12 -8 -6 -4 -2 0 2 4 6 8 10 12

~~Moo = 2.3 M =2.3 .4 .040~~

~~D _A = 0°, 6E = 0° .3 .035 _5A = 0% fiE = 20° .030 O6 A = 0 °, 6 E =-20 ° ,..] .2 ._ .025 _9~~

~~,d .1 88 _8 020 8888188 O000Oo00000000000 0 8_ 8 _,OlO.015 ,.d --.| oo, 6A = 0°' 815 = 20° ._ 0 --.2 tL O _A = 0% _E = -20° -.005~~

~~--.3 -6 -4 -2 0 2 4 6 8 10 12 -8 -6 -4 -2 0 2 4 6 8 10 12~~

~~(b) M._ = 2.0 and 2.3.~~

~~Figure 46. Continued.~~

~~67 Moo = 4.0 .4 .040 M = 4.0~~

~~.035 .3 .030 [] _A = 0°' _)E = 0° •_8 A = 0% 8 E : 20 ° ,.d .2 o .025 88_8 "o 0 8 A : 0 °, _E = -20°~~

~~.1 8088_ __ .020~~

~~o 8 0 80080 _ o15~~

~~,.d 888 _ olo --.1 °°°OOooooooooooo 8 []_A=0%_E=0o ._ 005 _[]••OrZDnDDDDDDO o + 8A = 0 °, _ = 20 °~~

~~--.3 -.010 ,, , ', ,, ',,, _,,, J,,, '... ',,, J,,, ..,, ,,., -6 -4 -2 0 2 4 6 8 10 12 -8 -6 -4 -2 0 2 4 6 8 10 12 IX IX~~

~~Moo = 4.63 Moo = 4.63 .4 .04O~~

.035 .3 D 8A = 0°, 8E =0° .030 8 A = 0 o, 8E = 20 ° ,..d .2 O0 .025 "0 ©8 A = 0 o, 8 E =-20 ° .020 e-: .1 00000 "0 8 - .015 o 0 0_0 e- _: .010 8 Q E ,...1 -.1 _ 0_00 & .005 13 8A = 0°, qSE= 0° °°Ooooooooooo9999 []D(3D(3U 0 O 8A = 0 °, 8 E = 20 ° --.2 ta_ O8 A = 0 °, 8 E =-20 ° -.005

~~-.3 tttl'''l'''l ' ''['/'I'''I'L'I'J'I '''l'ljJ -.010 -8 -6 -4 -2 0 2 4 6 8 10 12 -8 -6 -4 -2 0 2 4 6 8 10 12 IX IX~~

~~(c) M**= 4.0 and 4.63.~~

~~Figure 46. Concluded.~~

68 .0008 .0020 []M =2.3 NM =2.3 .0007 .0016 QM = 4.0 M = 4.0 I> _- _= .0006 OM = 4.63 OM = 4.63 D _ .0012 > .0005 [ > b '_ .0008 'r_ .0004 b b cD e l> D D 4 .0004 t> ._ .0003 _D 0 e 8 8 E .0002 .,= O O 0 0 Stable _2 -.0008 > 0 O >, i> -.000l -.0012 I>

~~-.0002 -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.~~

[] M =1.6 [] M =1.6 .O022 A Moo= 1.8 .0024 S M = 1.8 .0020 M o=2.0 .0020 V M =2.0 '_-.OOl8 D M =2.3 ca..0016 D M =2.3 r..) .0012 d .0016 <_ M =4.0 [] d @ M =4.0 > .0008 "_ .0014 O M =4.63 [] ..-_ O Moo=4.63 A ;_ .0004 _ .oo12 [] .r" [] A v "_ID 0 .0010 [] s /" v D g-.ooo4 E .0008 [] [] /x V o A S V [> -.0008 [],,_' _> g g. E .0006 v v V l> D _ v b &-.O012 v D .. .0004 t> > l> t> e-. [] s D ;-=-.0016 s V D 8 8 o_8 6 8 V .0002 -.0020 [] s Stable I [] A 0 -.0024 [] -.0002 ,,,I,.,l_,,I,,,I,,,I,.,ll_lljlllll_l -.0028 -6 -4 -2 0 2 4 6 8 10 12 -4 -2 0 2 4 6 8 10 12

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

~~69 .0010 .0016 O Vertical tail off O Vertical tail off .0008 .0012 [] Vertical tail on [] Vertical tail on~~

~~,; .0006 d .0008~~

~~t_ ;_ .0004 ;_ .0004 .r- g [] .0002 E2 [] _ 0 [] [] [] [] _A i Stable E Stable [ o 0 O E _ -.0004 0 0 0 0 0 0 i •=--.0002 _ -.0008 ID~~

~~--,0004 -.0012~~

~~-.0006 ,,,_,,, I,,, J,,,,,., i,,, ,,,, J,,, i,,, I -.0016 -6 -4 -2 0 2 4 6 8 10 12 -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 0 Vertical tail off 0 Vertical tailoff .0012 ,_, .0008 [] Vertical tail on [] Vertical tail on e_ u-..0008 _; .0006 [] e_ [] .0004 .0004 [] O [] -r" o [] D [] © Q.I "fi .00O2 © t •_ 0 t O O O Stable / o 0 © 0 rl _ _Stable E, _ -.0004 @ ;= -.0008 @ •_ -.0002

~~--.0004 -.0012~~

~~-.000% ...... -4 - 2'"0 .... 2' ...... 4 6 8 ...... 10 ;2 -.0016 _ --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 at M_, = 4.0.~~

7O .016 .010 I> M = 2.3 _> Moo = 2.3 .008 .014 (> Moo = 4.0 © Moo = 4.0 .006 O Moo = 4.63 O Moo = 4.63 .012 w- .004 <] - .010 t> b E>E>t> _ b _> .002 D b g fi E o .008 O 0 E D2> D I>D D2> E -.002 " .006 .=. -6 --.004 "¢ .004 -.006 .002 -.008

~~f_ L8 4 4 -2 0 2 4 6 8 10 12 -'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°-~~

~~Moo= 1.6 Moo= 1.6~~

Moo= 1.8 Moo = 1.8 Moo= 2.0 Moo = 2.0 Moo = 2.3 Moo = 2.3 Moo = 4.0 .016 Moo = 4.0 Moo = 4.63 .010 Moo = 4.63 .014 .008 [] [] [] [3 [] [] .006 .012 n A /X _ .004 A Eli ,J .010 VV V V_ _, A /X[3 _ .002 E O .008 0 E .006 = eooo@@@ O 88 reac; . vD _ -.004-.002 ,v .004 _8888_@00001>00 O0 * -.006

~~.002 -.008 u(3mNN[]NN~~

~~08''''' --6 _' -4' '' '-2I '''''''' 0 2 _' _''''4 6 ' ''''8 't*''ll10 2 -.010 -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 Moo= 1.6 Mo= 1.6~~

~~Moo = 1.8 Moo = 1.8 Moo = 2.0 Moo = 2.0 Moo = 2.3 Moo = 2.3~~

Moo = 4.0 Moo = 4.0 .016 .010 Moo = 4.63 Moo = 4.63 .008 .014 [3[3 .006 DO DO0 .012 AAAAA [] & AA J .004 A [] .010 .-: .002 e_ V VV VVVszzA [] E _ .008 o 0 E E " .006 ._ -.002 o@@@@g@ 0 f>v - @@@@ _" .004 -.006

~~.002 -.008 [3 VI VI FI £)_ 13 i- I~~

~~,,,i,, ,i ,,,I,,,It,,/,ILI,,,I,,,II,,I,, ,I 08, , ,i,...6i _, ,--4] , , ,-2_, _, 0i , , , 2i , , , 4i , , , i6 _, , _, , ,lOt , , ,12 -'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 configuration; _A = +20° and 8E = 20°-~~

! REPORT DOCUMENTATION PAGE I ro,_ A_o,o,_,_ I ouBNo.oro4-ola8 Public re_i;.g burden for this collectio_ of klformetion is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and revlewing the collection of information Send comments regarding this burden esttmete or any other aspect of this collection of Informetion, including suggestions for reducing this burden, to Washington Headquarters Services, Direclorste for Information Operations end Reports, 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 2220_-4302. and to the Office of Managernenl and Budget, Pape_¥ork Reduction Project (0704-0188), Washington, DC 20503. 1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED July 1996 Technical Paper

~~4. TITLE AND SUt:IIi/LE S. FUNDING NUMBERS Aerodynamic Characteristics of Two Waverider-Derived Hypersonic Cruise Configurations 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) 8. PERFORMING ORGANIZATION REPORT NUMBER NASA Langley Research Center L-17479 Hampton, VA 23681-0001~~

~~9. SPONSORING/MONITORINGAGENCYNAME(S)ANDADDRESS(ES) 10. SPONSORING/MONITORING AGENCY REPORT NUMBER~~

~~National Aeronautics and Space Administration NASA TP-3559 Washington, _ 20546-0001~~

~~11. SUPPLEMENTARY NOII::_ Cockrell and Huebner: Langley Research Center, Hampton, VA; Finley: Lockheed-Fort Worth Company, Fort Worth, TX.~~

~~12a. Di_iHIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE~~

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

13. A_,_IPIACT (Maximum 200 words) An evaluation was made on the effects of integrating the required aircraft components with hypersonic high-lift configurations known as waveriders to create hypersonic cruise vehicles. Previous studies suggest that waveriders offer advantages in aerodynamic performance and propulsion/airframe integration (PAl) characteristics over conventional non-waverider hypersonic shapes. A wind-tunnel model was developed that integrates vehicle components, 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 fluid 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 aerodynamic characteristics of the fully integrated configurations, including control surface effectiveness. The aerodynamic performance of the fully integrated configurations is not comparable to that of the pure waverider shapes, but is comparable to previously tested hypersonic models. Both configurations exhibit good lateral-directional stability characteristics.

14. SUBJECTTERMS 15. NUMBER OF PAGES Hypersonic cruise; Waveriders; Airbreathing vehicles 73 16. PRICE CODE A04 17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF REPORT OF THIS PAGE OF ABSTRACT OF ABSTRACT Unclassified Unclassified Unclassified

~~NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89) Prescnbed by ANSI Std. Z39-18 298-102~~