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Aerospace Science and Technology, 1998, no. 8, 505-514

Improvement of forebody/in.let intentionfor hypersonic vehicleN. C. Bissinger a*, N. A. Blagoveshchensky b, A. A. Gubanov b,

V. N. Gusev b, V. P. Starukhin b, N. V. Voevodenko b, S. M. Zadonsky ba Daimler Benz Aerospace AG, MT633,81663 Miinchen, Germanyb Central Aerohydrodynamic Institute, TsAGI , 140160 Zhukovsky, Moscow region, Russia

(Received 25 November 1995, revised and accepted 21 September 1998)

B&singer N. C., Blagoveshchensky N. A., Gubanov A. A., Gusev V. N., Starukhin V. P.,Voevodenko N. V., Zadonsky S. M., Aerospace Science and Technology, 1998, no. 8, 505-514.AbSWWt Results of a numerical (CID) study of the influence of the forebody shape on local flow parameters ata bottom-mounted inlet entrance are presented. The free-stream Mach number is assumed to be 3.5-7.0.Some recommendations on forebody shape optimization are provided. Main characteristicsof the air inletare evaluated. 0 Elsevier, Par is

hypersonic inlet charac teristics / forebody-inlet integration / hypersonic mu&disturbance theory /Godunovs methodZusammenfmg Verbesserung der Vork&rper/EinIauf-Integration fir Hyperschsdl-Flaggerlite . Es werden dieErgebnisse von numerischen Rechnungen (CPD) tiber den Einflu6 der Vork&perform auf dieStronnmgsparameter an einem Unterrumpfeinlauf vorgestellt. Die untersuchten Flugmachzahlen liegenin dem Bereich von 3.5 bis 7.0. Aus den Ergebnissen werden Empfehhmgen zur Vork&peroptimierungabgeleitet. Die wichtigs ten charakteristischen Gr66en des Einlaufs werden ermittelt. 0 Elsevier, Paris

EiWufckaraktiri$tiken im Hyperschall / Vorkthper-Elnlauf Integration / Theorie kleiner Stirmgenim I-IyperschaII / Gdunovs Metbode

1. ItiroddonAerodynamic qualities of a flying vehicle designedfor high supersonic speeds depend to a great extenton the degree of integration between the airframe andthe propulsion system. One integration measure is theuse of the forebody lower surface as a precompressionramp for the inlet. In this respect it is important toshape the forebody so as to ensure the most eff icientintegration of the airframe and the propulsion system.Various aspects of optimization are considered, for* Correspond ence and reprintsAemspace Science and Techndogy, 1270-9638, 98/08/O Elsevier, Paris

example, in [l, 3, 4, 61. The resul ts provided hereinindicate a potential for improving essentially the inletentrance flowfield by properly adjusting the forebodyshape.The study involves two versions of an unmannedhypersonic research vehicle, HYTEX and HYTEX-M,that were included in a list of 10 alternatives for useas a ramjet flying testbed [S].Major geometry parameters of these vehic les (e.g.wing size, body length, volume etc.) were fixed


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Section A-AR I Seclipn B-B

Figure 1. HYTEX configuration.or could be varied only within strong limitations.Therefore for this study the following geometricparameters of the airplane forebody were assumedto be design variables:

- the longitudinal curvature of the lower surface,- the fuselage cross-sectional shapes,- the angle between the forebody lower surface and

the airplane longitudinal axis,- the presence of wing strakes.The influence of viscosity on inlet entranceflowfields is assumed to be similar for theconfigurations considered, and therefore an analysisis carried out within the frame of an invisc id idealgas model.The major flow parameters that govern (or indicate)the influence of the airplane forebody on inletcharacteristics, include:- precompression intensity (measured by changesin Mach number),- losses in total pressure,- flow uniformity at the inlet entrance.

2. Descri@ion of HYTEX & HYTEX-McoA general view of HYTEX is presented in figure 1.This aircraft has the length L=15.4 m, the wing spane= 3.7 m, and the height H= 2.9 m. The ramjet inletentrance plane is at a distance za =9 m from theforebody nose tip; the width of the inlet entrancesection is 0.8 m, and the height is 0.5 m.

The forward portion of the forebody is a bodyof revolution with a maximum cross-section radiusR=0.55 m and a length of Ax, = 4.4 m. The secondpart of the forebody has a length Axs =3.2 m andis a transition piece smoothly couphng the forwardportion (with circular cross-sections) and the thirdportion, whose cross section is- a semi-circle (with a radius R = 0.55 m) at the top,- a rectangle (with a width b=2R=l.l m and aheight h =0.6 m) at the bottom,

the sides and the bottom being matched by circularcylinders with a radius ~0.05 m (see the fuselagecross-section at a location x= Axr + Ax:2 =7.6 m infisure I). The lower surface of this portion is set atan angle of 0.9 to the airplane longitudinal axis.The third portion of the HYTRX forebody (fromx = 7.6 m to 5 =x0 = 9 m, seefigure 1) has a constantcross section.A general view of the HYTEX-M aircraft ispresented in figure 2. It possesses the length

L = 14.64 m, the wing span e=4 m, and the heightH= 3.16 m. The inlet entrance plane is at the. locationx0 =7.84 m as measured from the fuselage nose tip;the inlet dimensions are the same as on the HYTEXconfiguration (that is, 0.8 by 0.5 m).The HYTRX-M forebody has a Rat lower surfaceinclined by an angle @a with respect to the aircraftlongitudinal axis (two versions were considered, onewith 80 =O, and the other, with 8. =3). Forebodycross sections are com#osed of an upper semi-circ leand a lower rectangle of the same width. Thetwo lower comers of the rectangle are rounded, seefigure 2.

Aerospace Science and Technology


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Improvement of forebody/inlet integration for hypersonic vehicle/Verbesserung der Wrk&-per/Einlauf-Integration fir Hyperschall-Fluggeriite 507

t= -ISection A-A

Figure 2. HYTEX-M conf iguration.The HYTEX-M configuration with Ba= 0 was usedto evaluate the effects of wing strakes on the flowconditions near the inlet. These strakes were of atriangular planform with a flat lower surface parallelto the airplane longitudinal axis; the span of the strakesin the inlet entrance plane was 2.1 m, the leading edgesweep angle was 83. The HYTEX-M wing itself iscompletely downstream of the inlet entrance plane andhas not been simulated in the flowfield calculations.

3. Features of the flow around the forebodiesSome recommendations on the optimization of thehypersonic airplane forebody shape can be formulatedon the basis of a preliminary qualitative analysis ofsupersonic flow features.The first thing that can be noticed about the flowaround the HYTEX forebody is a notable loss of totalpressure in the bow shock wave. The semi-vertexangle w. of the nose cone is 15. This means thatthe bow shock intensity (as a function of the angle

between a body surface and the free-stream velocityvector) will be high, even if angles of attack (cy) arelow. Therefore the total pressure losses due to thisshock wave will be large. With increasing incidence(i.e. with increasing values of the angle betweenthe lower surface and free stream, (YL = wu + a)these losses become larger and the total pressure ofthe flow entering the inlet wil l be decreased evenfurther. Because the maximum angle between thelower surface and free stream, OL = 60 + a, is smallerfor HYTEX-M its shock losses can be expected to bemuch smaller.Secondly, the convex lower surface of HYTBXcauses the flow to expand, thus increasing the local

1998. no. 8

Mach number at the inlet entrance; the precompressioneffects become partly lost. The HYTEX-M lowersurface without longitudinal curvature eliminates thisdrawback.Thirdly, compared with HYTEX the flat lowersurface of the HYTEX-M forebody is expected toprovide a more uniform flowfield.

4. Computational methodThe influence of the forebody shape on flowparameters in the vic inity of the inlet was evaluatedusing the numerical method of [lo]. The latter isbased on the numerical integration of the equations ofthe hypersonic small-disturbance theory [5].This theory reduces the 3-D Euler equations tothe 2-D unsteady Euler equations (it is assumedthat the longitudinal component of the flow velocityis u = U, COSQ, U, is the velocity of the freestream, and the longitudinal coordinate 2 is replacedby time t : z = U,t cos (Y), which are integrated by

Godunovs numerical method.Assuming that:

andMm > 1, r N S = d/L < 1,

Moor - 1 or M,7> 1V& c is the free-stream Mach number, 7 is themaximum inclination of the body surface to thefree-stream direction, d is the maximum transversedimension of the body, L is the length of the body),it is possible to simplify the equations of the idealgas motion (3-D Euler equations). As follows from


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the hypersonic small-disturbance theory, in accordancewith an order-of-magnitude estimate, dimensionlessvariables (with subscript 0) are introduced:the coordinates:2 = La-J, y = dye, z = dz,,

the flow velocity components:21= U&l + S%(J), w = u,sva, w = U&W&

the density:P = POOPO,

the static pressure:P = Pmu~~2Po.

Discarding terms of the order of r2 as comparedwith unity, we can obtain relations identical to 2-Dunsteady flow equations, if the variable zo is replacedby to ($0 = to):dP0 dPov0 + apowTg+- -dY0

=o >0 azo 1 ape~+vo~+wo~+--dl,0 aYo 0awe &w, p. hoTg + 210z& +wo awe 1 aPo-+---~o,0 azo p. azo

Here, K = 1.4.The approximate conditions on the surface ofthe body (g(xa, yc,xo) = 0 is the equation whichdetermines this surface) will have the form:69 ag agat +voa~o-+wodz=o. 0The transformation of the conditions at the shockwave (the equation F(za, ye, ~0) = 0 represents theshape of the shock-wave which is being determinedduring the numerical calculations) leads to thefollowing relations:

[PO%] =o, [PoTJog +PoE] = 0,

=o,

Here, the expressions in square brackets denote thedifferences in the corresponding magnitudes ahead andbehind the shock, and as the flow velocity ahead ofthe shock wave we have taken the free-stream velocitycomponent in the plane z =const.This is the problem of the 2-D unsteady motion ofa gas caused by the expansion of a 2-D piston [2].The theory described above has been extended tothe case of high angles of attack (up to 90) byV.V. Sychev [9]. Sychevs theory leads to the same2-D unsteady equations as Hayess theory [5]. Thenumerical method used here is based on both theories[5] and [9].The hypersonic small-disturbance theory equations(HSDTE) are integrated by Godunovs first-orderapproximation method. The bow shock wave is fittedby the procedure given in [7] and is assumed to bea boundary of the calculation region. The disturbedregion is confined between the body surface and the

bow shock wave. Accordingly, the calculation grid isattached to the body surface and the bow shock wave.If the calculation region is geometrically complex thisregion may be divided into simple subregions. Ineach subregion the calculation grid is built such thatgrid points coincide on the boundaries of neighboringsubregions. The shape of the cross section determinesthe number of subregions and their arrangement.The advantages of the numerical method based onhypersonic small-disturbance theory are the following :- significant saving of computer time in comparisonwith the numerical integration of the completeEuler equations, as a result of the successfulamalgamation of hypersonic smal l disturbancetheory with Godunovs method;- the numerical method based on HSDTE is morestable in operation and more robust than 3-DEuler space-marching methods because HSDTEare always hyperbolic. (Their longitudinalvelocity component stays constant.) In contrast,3-D Euler space-marching methods often loosetheir stab ility due to the change of the equationtype when the longitudinal velocity componentbecomes subsonic;- the HSDTE method even allows the calculationsof flows with small subsonic flow regions.Because of the very small bluntness of the HYTEX

and HYTEX-M forebodies (radi i 0.02 m and 0.025 mrespectively) this method could be applied withsufficient accuracy.The numerical method and program based onHSDTE have been tested on a great variety ofconfigurations. The results of those calculations havebeen compared with experimental data and with resul tsof various numerical methods including 3-D Eulermethods with Godunov and McCormack flow salvers.The results of such investigations are presented in [lo].The basic theory is asymptotic, but in practice it hasbeen shown that the applicability range of this methodis: 2 5 1M, 5 10, (a]


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Improvement of forebodykdet integration for hypersonicvehicle/Verbesserung der VorkiirperfEinlauf-Integration j?ir Hyperschall-Fluggeriite 509

The present numerical method calculates the flow wing strakes on the mass flow ratio is insignificantparameters p, p, w and w. After the completion (the related increments in f do not exceed 5 %), andof the marching procedure the longitudinal velocity the removal of the strakes may he compensated bycomponent u can be obtained using Bernoul lis increasing the forebody lower surface inclination angleequation: 0s by about 1.u2 + ?J2+ w2 v; K pm

2 +2-P-=K-1 p 2f-- K-1 pooThis procedure has been used to predict the massflow ratio of the inlet because it improves theaccuracy of the conservation of mass by taking intoaccount the most important higher-order terms ofthe basic relations. Comparisons with the results ofcomputations based on 3-D Euler equations for similarconfigurations with flow precompression in front ofinlets showed that differences in mass flow ratios donot exceed 1% in the Mach number range from 2 to

7 and for angles of attack of 0 < Q 5 10.

One should take into account both a likely negativeimpact of the wing strakes on the airplane longitudinalstability and their low efficiency; with this in view,we shall dwell on HYTEX-M with no strakes and aforebody lower surface inclination angle B0= 3. Atthis value of 00 the forebody bow shock wi ll notimpinge on the inlet lip if a 5 10. This is evidencedby the shock wave shapes yigure 4) calculated for theinlet entrance plane at M, =7.

5. Results of flowfield calculationsIn order to identify the most promising configurationof HYTEX-M for more detailed studies, figure 3represents the calculated inlet mass flow ratio f asa function of the angle of attack at the Mach numberM, = 7; hereinafter, f = A,/Ai, A, is the cross-sectional area of the free stream tube captured bythe inlet, Ai is the inlet capture area as measuredin the plane normal to the forebody lower surface.From the diagram it can be seen that the influence of

f/2. 5

051-0 5

/a=0

Figure 3. Inlet mass f low rat io for dif ferent co ncepts of HYTEX-M.f versus angle-of-at tack, cx. A& =7.1998, no. 8

Figure 4. Shock -wave co nfigurations at the inlet entrance plane,x=7.84 m for HYTEX-M. 00 = 3, without strakes, A4, =7.Local forebody flowfields at the inlet entrance cross-sections (x = za) of both HYTEX and HYTEX-M atM, = 7, a = 6 are presented in figures 5 to 10.These flowfields are represented by distributions ofthe local Mach number, Me, the ratio of local totalpressure, PQ to the free-stream total pressure, Pt,,

and the local mass flow ratio:fe = Pe uePee v,Pe is the local flow density at the inlet entranceplane, pm is the free-stream density, ue is the z-component of the disturbed flow velocity; U, is thevelocity of the free-stream. Figures 11 to 13 depictthe variation of the flow parameters mentioned aboveat the inlet entrance location along two vertical lines:near the aircraft symmetry plane (y =0,02 m) and nearthe side wall of the inlet (y=O.O38 m).These results make it clear that over a considerablepart of the HYTEX-M inlet entrance the local Mach


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Figure 5. Local Mach num ber, M, distribution at the inlet entranceplane, z = 9 m for HYTEX forebody. A& = 7, (Y= 6.

Fiire 6. Ratio of local total to free-steam total pressure , l,e/Pt,distribut ion at the inlet entrance plane, cz 9 m for HYTEX forebody.M, =7, a=6.numbers are lower, and values PQ/P,, and fe arehigher, than those of HYTEX. Also, the HYTEX-Mflowfield is more uniform, Each of these differencesindicates the superiority of the HYTEX-M overHYTEX concerning the onset-flow for the inlet whichwil l be detailed in chapter 6 below.

The configurations under study feature an insignif-icant dependence of inlet entrance flow parameterson the distance to the symmetry plane: the curvesfor y=O.O2 m and for y =0.38 m corresponding to aparticular forebody configuration differ only slightly.Al l of the results demonstrated in this paper havebeen obtained on a grid fine enough to produceresults that are grid independent. This grid has beondetermined by numerous calculations with severalmore and more refined grid densities. The final gridfor one half of the symmetric forebody had 41 x 81points in each cross-sectional plane (41 points on eachof 81 grid lines connecting the body surface with the

N. C. Rissimgeret al.

Figure 7. Local ma ss flow ratio, fp distribution at the inlet entranceplane, z = 9 m for HYTEX forebody. M, = 7, a = 6.

Figure 8. Local Mach numb er, Me , distribution at the inlet entranceplane, z = 7.84 m for HYTEX-M forebody. Mm = 7, a = 6.bow shock-wave). Comparisons of the results fromcalculations with this fine grid with results obtainedusing a grid of 21 x 41 points showed differences inall flow parameters considered to be less than 1% forthe HYTEX configuration.

6. Evaluation of air inlet characteristicsCharacteristics of inlets are usually estimatedpreliminarily in terms of area-averaged flow conditionsat the inlet entrance plane. Results of the calculationof averaged values of the local Mach number M,, andthe ratio of the averaged total pressure Pt, to the free-stream total pressure Pt, at various Mach numbersAG!, over the angle-of-attack range from 0 to 10may be seen in Jigures 14 and 15. It is evident thatthe averaged Mach numbers for both I-IYTEX andHYTEX-M differ insigticantly; the correspoud&gdifferences of A& between HYTEX and HVIBX-M



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Improvement of forebody/inlet integration for hypersonic vehicle/Verbesserung der VorkiirpedEinlauf-Integration fir Hyperschall-Fluggeriite

Figure 9. Ratio of local total to free-stream total pressure, Pte Pt,distribution at the inlet entrance plane, 2 = 7.84 m for HYTE X-Mforebody. M, =7, cy=6O.

Figure 10. Local mass f low rat io, fe distribution at the inlet entranceplane, z = 7.84 m for HYTEX-M forebody. M, = 7, (Y = 6O.are less than 0.1-0.2. However, these configurationsshow considerable differences of inlet entrance totalpressure, especially at low incidence. For example,at M o. = 7 and a, =0 the HYTEX-M configurationprovides a Pt,, which is greater by 30% than thatof HYTJZX. The differences of Pt, become smallerwith increasing a.

The slightly decreased average Mach numbers incombination with the increased values for the inletentrance total pressure of HYTEX-M produce anoticeable increase in both the mass flow ratio andthe total pressure recovery factor.The computed values of the inlet mass flowratio f (obtained by integration over the capture areameasured in the plane normal to the lower surface ofthe forebody) are presented in jgure 16. One can seeadvantages of HYTEX-M under all flow condit ions

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511

Figure 11. Local M ach number at the inlet entrance planes, Meversus distance from the forebody bottom surface, AZ. M, =7,a=6'=.

PPtX

0. E

0. ;

0. 1

0.

0.

111 PI I

r-

-

-I/ aom y=O.O2m/ / -cc- 038m Hy[D(IiItk-

- y=O.O2m038m W&II

III iI I I

0 0.25 0. 5 A2Figure 12. Ratio of local total to free-steam total pressu re at inletentrance planes, Pte JPt, versus distance from the forebody bottomsurface, AZ. M, =7, a =6.considered. At Mm = 7 and 0~ Q < 10 the mass flowratio increments attained are 30-40 %.


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/ -*- y= 0.02m-CC- 038m HylM1- y=O.O2m038m *MI

T- I

0 0.25 0.5 uFigure 13. Local ma ss flow ratio at the inlet entrance planes, ffversus distance from the forebody bottom surface, AZ. M, = 7,cx=6.

0 5 10 aP 1Figure 14. Average Mach number at inlet entrance locat ions, M,,versus angle-of-at tack, (2.

Pdptl

0.9

0.8

0.7

0.6

5 10 al"JFigure 15. Ratio of average total pressure at inlet entrance locationsto free-steam total pressure, I&/P,- versus angle-of-at tack, Q.

The total pressure ecovery actor v = Pt,f/PtmV& f is the total pressureat the engine face, andPt- is the total pressure f the free stream) s alsoregarded s one of the most importantcharacteristicsof inlets. In order to estimate his factor, one canfix the kinetic energy ecovery actor 7 defined orthe flow within the inlet (from the tip of the inlet tothe engine ace), for instance,at the level q=O.92,andperform he appropriate omputation y using heaveraged arameters&faVandP,,/P,- ) of the flowat the inlet entrance:

Py=tav 1+[-A

P t- +I$(1 - 77)1Computedvaluesof v for HYTEX and HYTEX-M, as well as the corresponding aluesof v for anisolated nlet in a uniform free streamare epresentedin jlgure 17.These results show that the HYTEX-M fore-bodyiinlet integration increases he total pressurerecovery actor n comparisonwith the isolated nlet,throughout he flight envelope onsidered.ncrementsin v between WIFX andNYTElX-M exceed 0% atMm=7 and O


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f2. 5

2. 0

1. 5

1. 0

0. 5

I----Hmx- h YE% M

0 5 10 aP 1Figure 16. Inlet mass flow ratio, f versus angle-of-attack, (Y.

However, it should be noted that the final selectionof the shape of the forebody for the vehic le in additionrequires the consideration of lift, drag and momentsacting on the whole vehicle.It is interesting to compare the inlet flow parametersobtained by numerical calculation of the flow pastthe HYTEX-M configuration and those for the inletlocated under a 2-D flat plate with unswept leadingedge, which is inclined to the free-stream direction bythe same angle as the lower surface of the HYTEX-Mforebody (Q~ = B. + a). The results correspondingto 19~ 3) Mm = 7, (Y= 6 are presented in Table I.It can be seen that, though both precompressionsurfaces considered here are flat, differences are signi f-icant, and therefore numerical flowfield investigationsare necessary for correct quantitative prediction of theinlet performance.

7. ConclusionThe above studies illustrate the considerablepotential for improving aerodynamic configurationsof high-speed air-breathing aircraft by means offorebody/inlet integration. The following conclusionscan be derived.1. Convex lower surfaces of forebodies should beavoided ; otherwise, the total pressure losses in

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V

o? b

0. 5

0. 4

0. 3

0. 2

0. 1

x=10

o=o1

3 4 5 6 ' M,Figure 17. Total pressure recovery factor, v versus Mach numb er,n/r,.Table I. Comparison of two configurations.

Inlet under a flat plate 5.41 0.786 2.38 0.205

the nose shock are large and, additional ly, theflow in front of the inlet is accelerated.2. A flat-bottomed forebody with the length-to-

width ratio of about 6.5 ensures a uniform localflow at the inlet entrance when the flight Machnumber is 3.5-7.0 and a=O-10.3. For an airplane with a flat lower surface ofthe forebody the wing strakes are poor meansfor flow precompression intensification. Withstrakes removed, the inlet mass flow ratiodecrease may be compensated for by increasingthe forebody lower surface inclinat ion angle byone degree.4. The HYTEX-M forebody configuration formedwith due account of inlet entrance flow conditionrequirements enables to increase the mass flowratio and the total pressure recovery factor by


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514 N. C. l%singer et al.

30%ati&,=7andOcac6,ascomparedwithHYTEX that has an axisymmettic forward part.

AcknowledgementMost of the work the results of which are reported inthis paper has been supported by the Bundesministerium furBildung, Wissenschaft, Forschung und Technologie (BMBF)within the German Hypersonic Technology Programme.

References[l] Beach H.L., Blankson I.M., Prospects for FutureHypersonic Air-Breathing Vehicles, AIAA Paper91-5009 (1991).[2] Cox R.N., Crabtree L.F., Elements of HypersonicAerodynamics, The English Universities Press LTD,London, 1965.[3 Gubanov A.A., Pritulo M.F., Ruchyev V.M., The-oretical Investigation of Airframe/Inlet Interference

and Integration for Hypersonic Veh icles, Z. Flugw iss.Weltraumforsch (1994) 18.[4] Gusev V.N. Aerospace Aerothermodynamics, TsAGIJournal l(1) (1994).[5] Hayes W.D. On Hypersonic Similitude, Quart. Appl.Math. 5 (1) (1947).[6] Hirsche l E.H., Aerofhermodynamics and PropulsionIntegration in the SANGER Technology Programme,AIAA Paper 91-5041 (1991).[7] Krajko A.N., Makarov V.E., Tilla yeva N.I. OnNumerical Construction of Shock-WaveFronts, J. Num.Math. and Math. Phys. 20 (3) (1980).[8] Kraus M., Lazarev V., Sacher P., Shkadov L.Hypersonic Flight Test Vehicle for Ramjet Testing.-IAC 94 (International Aerospace Congress, August15-19, 1994), Moscow, Russia.[9] S ychev V.V. Three-dimensional Hypersonic Gas FlowPast Slender Bodies at High Angles of Attack, J. Appl.Math. Mech. 24 (2) (1960).[ 101 Voevodenko N.V., Computation of Super-

sonic/Hypersonic Flow near Complex Configurations,ICAS-94-2.2.3, 1994.


4 improvement of forebodyin.let intention

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