AEDC .. TDR .. 64-137

SUMMARY REPORT O.N AERODYNAMIC CHARACTERISTICS OF STANDARD

MODELS HB-1 AND HB-2

1. Don Gray

von Karman Gas Dynamics Facility ARO, Inc.

TECHNICAL DOCUMENTARY REPORT NO. AEDCwTDR-64-137

July 1964

Program Element 65402034

(Prepared under Contract No. AF 40(600}.1000 by ARO, Inc., contract operator of AEDC, Arnold Air Force Station, Tenn.)

lD EN INEERI PME TEE

AIR fORCE SYSTEMS COMMAND

UNITE STATES AI fORCE R

UB 40[600}1000

NOl'lCBS Qualified requesters may obtain copies of this report from DDC, Cameron Station, Alexandria, Va. Orders will be expedited if placed through the librarian or other staff member designated to request and receive documents from DDC.

When Government drawings, specifications or other data are used for any purpose other than in connection with a definitely related Government procurement operation, the United States Government thereby incurs no responsibility nor any obligation whatsoever; and the fact that the Government may have formulated, furnished, or in any way supplied the said drawings, specifications, or other data, is not to he regarded by implication or otherwise as in any manner licensing the holder or any other person or corporation, or conveying any rights or permission to manufacture, use, or sell any patented invention that may in any way be related thereto.

I

AF - AEOC Arnold AFS Ten"

SUMMARY REPORT ON AERODYNAMIC

CHARACTERISTICS OF STANDARD

MODELS HB-1 AND HB-2

By

J. Don Gray I I

von Karman Gas Dynamics Facility

ARO, Inc.

a subsidiary of Sverdrup and Parcel, Inc.

July 1964

ARO Project No. VT2116

AEDC-TDR-64-137

AEDC-TDR-64-137

ABSTRACT

All available test data on the standard hypersonic ballistic correla­tion models. HB-1 and HB-2. have been examined to establish reference characteristics for use in the correlation of experimental results among the various aerodynamic facilities. The data presented are comprised primarily of static stability and zero -lift forebody axial-force coefficient results at Mach numbers from about 1. 5 to 20 for a Reynolds number range from 0.01 to 2.7 x 106 based on centerbody diameter. Selected pressure distribution. dynamic stability. and heat-transfer data are in­cluded. Comparisons of data from various test facilities show excellent correlation in spite of the appreciable Reynolds number dependence of test results for model HB-2 with natural boundary-layer transition.

PUBLICATION REVIEW

This report has been reviewed and publication is approved.

~~~ Darreld K. Calkins Major, USAF AF Representative, VKF DeS/Test

iii

/\ rf}~' (I .-\',-.. jUfn'l { ~ .. -'ll ~ .J~an A: Jack .' v Colonel, USAF DCS/Test

AEDC-TDR-64-137

COMTEMTS

ABSTRACT .•. NOMENCLATURE

1. 0 INTRODUCTION. 2.0 TEST EQUIPMENT . . • • . 3. 0 RESULTS AND DISCUSSION

3. 1 Force Measurements 3. 2 Pressure and Heat-Transfer Measurements

4.0 CONCLUSIONS . . • . . REFERENCES . • . • . • • . . • • . . . • • .

TABLE

1. Summary of Experimental Results 0

ILLUSTRATIONS

Figure

1. Standard Model and Sting Configurations a. Models HB-1 and HB-2 .... O. Sting Geometry . . • 0 • 0 •

2. Normal-Force Coefficient versus Angle of Attack for Model HB-1

a. Mach Numbers 1. 5. 2. 3. 4. and 5 b. Mach Numbers 6. 7. 8. and 10 . 0

3. Pitching-Moment Coefficient versus Angle of Attack for Model HB-1

Page

iii ix 1 2

3 6 7' 8

9

11 12

13 14

a. Mach Numbers 1. 5. 2. 3. 4. and 5. . . . • 15-b. Mach Numbers 6, 7, 8, and 10 . . . . . . •. 16

4. Normal-Force Coefficient versus Angle of Attack for Model HB-2

a. Mach Numbers 1. 5. 2. 3. 4. and 5. . . • • 17 b. Mach Numbers 6, 7. 8. 10 and 16.5 .•• 0 18

5. Pitching-Moment Coefficient versus Angle of Att~k for Model HB-2

a. Mach Numbers 1. 5, 2, 3, 4, and 5. • . b. Mach Numbers 6. 7. 8, 10. and 16.5 ••.

v

19 20

AEDC-TDR-64-137

Figure

6, Initial Slope of Normal-Force Curve versus Reynolds Number for Model H~-l

a. Mach Numbers 1. 5. 2. 3. 4. and 5. . . • • 21 b. Mach Numbers 6. 7. 8. and 10. . . • • 22

7. Center of Pressure versus Reynolds Number for Model HB-1

a. Mach Numbers 1. 5. 2. 3. 4, and 5. • • 23 b. Mach Numbers 6. 7, 8. and 10. • • • • • ., 24

8. Zero-Lift Forebody Axial-Force Coefficient versus Reynolds Number for Model HB-1

a. Mach Numbers 1. 5. 2. 3. 4. and 5. • . . b. Mach Numbers 6, 7. 8. 10. and 20 •...

9. Zero-Lift Base Axial-Force Coefficient versus Reynolds Number for Model HB-1

25 26

a. Mach Numbers 1. 5, 2. 3. 4, and 5. . . • • •. 27 b. Mach Numbers 6, 7. 8. and 10. . • • • . • 28

10. Initial Slope of Normal-Force Curve versus Reynolds Number for Model HB-2

a. Mach Numbers 1. 5. 2. 3. and 4 . . • • . b. Mach Numbers 5. 6. 7. 8. and 10 • . . •

11. Center of Pressure versus Reynolds Number for Model HB-2

a. Mach Numbers 1. 5. 2. 3. 4. and 5. . . . b. Mach Numbers 6. 7. 8. and 10. . . • .

12. Zero-Lift Forebody Axial-Force Coefficient versus Reynolds Number for Model HB-2

a. Mach Numbers 1. 5. 2. 3. 4. and 5. • • . b. Mach Numbers 6. 7. 8. 10. 16. 5. a~d 20 .

13. Zero-Lift Forebody Axial-Force Coefficient versus Viscous Interaction Parameter for Models HB-1 and HB- 2 . . . . . . . . . . . . . . . . . . .

14. Zero-Lift Base Axial-Force Coefficient versus Reynolds -Number for Model HB-2

29 30

31 32

33 34

35

a. Mach Numbers 1. 5. 2. 3. 4. and 5. . . . • 36 b. Mach Numbers 6. 7. 8. and 10. . . . • • • •. 37

15. Comparison of Mach Number Effect on Models HB-1 and HB-2

a. Initial Slope of Normal-Force Curve . • • b. Center of Pressure. . . . . . . • . • . c. Zero-Lift Forebody and Base Axial-Force

Coefficients • .• ...••••••.

vi

38 39

40

AEDC·TDR·64·137

Figure

16. Damping-in-Pitch Derivatives versus Reynolds Number for Model HB-2 at Mach Numbers 2, 3, and 4. . . . .. 41

17. Longitudinal Pressure Distributions at Zero Lift for Model HB-2 at Mach Numbers 1. 5, 2, 3, 4, 5, 8, and 18. . . . . . . . . . . . . .

18. Longitudinal Pressure Distributions at Zero Lift for

42

Model HB-1 at Mach Numbers 8 and 20. . . . • • . 43

19. Comparison of Longitudinal Pressure Distributions at Zero Lift for Model HB-2 from VKF, DAL, and ONERA . . . . . . . . . . . . . . . . . . . . . 44

20. Longitudinal Heat-Transfer Distribution at Zero Lift for Models HB-1 and HB- 2 at Mach 18. . . • . • . • 45

21. Longitudinal Pressure Distributions for an Angle of Attack of 10 deg at Three Circumferential Positions for Model HB-2

a. Mach Number 1. 5. b. Mach Number 3. . . c. Mach Number 5. . d. Mach Number 8. . .

22. Local Normal-Force and Axial-Force Coefficient Distributions on Model HB - 2 at an Angle of Attack of 10 deg for Mach Numbers 1. 5, 3, 5, and 8

46 46 46 46

a. Local Nonnal- Force Coefficient Distribution. 47 b. Local Axial- Force Coefficient Distribution 47

vii

CNa

D

d

L

f

p

Pm

NOMENCLATURE

Reference aref' 1Td2 14, in. 2

Base area. in.~

Chapman-Rubesin constant. iJwTm/iJmTw

AE DC-TDR-64-137

Zero-lift forebody axial-force coefficient. (CAt-CAb)

Zero-lift base axial-force coefficient. Cpb(Ab I A)

total axial force Total zero-lift axial-force coefficient.

Cleo A

Pitching-moment coefficient (reference at 1. 95d from nose). pitching moment

CleoAd

Damping-in-pitch derivative. a Cm la(qd/2Va). l/rad

Damping-in-pitch derivative. a Cmla (ad/2Vm ). 1/rad

., normal force Normal-force coefflclent. Cleo A

Slope of normal-force curve. (a CN I aa) a = 0

Pressure coefficient. (p - Pm) I Cleo

1T Local axial-force coefficient. (2/A) J Cpr tan o-dc/>. i/in.

o

1T Local normal-force coefficient. (21 A) J Cpr cos ~ #. 1 lin.

o

Sting diameter at model base. in.

Reference centerbody diameter (see Fig. 1). in.

Length of sting of constant diameter. D. in.

Overall model length (4. 9d). in.

Free-stream Mach number

Surface pressure, psia

Free-stream pressure. psia

Base pressure, psia

ix

AEDC-TDR-64-137

r

S

x

xcp/.R.

'Y

e

(J

w

Model stagnation pressure. psia

Pitching velocity. rad/ sec 2

Free-stream dynamic pressure. ('Y /2) Pm Mm. psia

Heat-transfer rate relative to model stagnation point heat­transfer rate

Reynolds number bas ed on centerbody diameter and free­stream conditions

Blunted nose radius (0. 3d). in.

Model radius. in.

Wetted length. in.

Free-stream temperature. OR

Wall temperature, OR

Free-stream velocity. ft/sec

Viscous parameter. Moo..J coo/~ Red

Axial distance. in.

Center of pressure. [1. 95 - (dCm/dCN)a= o]d/.R.

Angle of attack. deg

Time rate-of-change of u. rad/sec

Sting windshield half-angle. deg

Ratio of specific heats

Angular displacement. deg:

Viscosity of air evaluated at free-stream temperature. lb-sec /ft 2

Viscosity of air evaluated at wall temperature, lb-sec /ft 2

Local surface inclination, deg

Circumferential position. deg

Angular frequency. rad/ sec

x

AEDC-TDR-64-137

1.0 INTRODUCTION

The necessity for a standard for the evaluation of data obtained in aerodynamic test facilities is well recognized. Such a standard may be established through the use of identical calibration models in the greatest number of test facilities.

Through an appreciation of the many fundamental factors involved and an awareness of the characteristics of the instrumentation used, a realistic confidence in the validity of test data may be then established. The Advisory Group for Aeronautical Research and Development (AGARD) calibration models are well- known examples of the utility of this approach (Ref. 1).

The desire for a ballistic correlation model was first voiced at the joint meeting of AGARD and STA (Supersonic Tunnel Association) held in France during September 1959, and the initiative for selecting a suitable

, I

configuration was assumed by Lukasiewicz of the von Karman Gas Dy-namics Facility (VKF), Arnold Engineering Development Center (AEDC), Air Force Systems Command (AFSC). This resulted in adoption by the STA, in March 1960, of two hypervelocity ballistic model configurations, designated HB-1 and HB- 2 (Fig. 1a), which consisted of a blunt cone­cylinder with and without a flare. The latter shape was included to pro­vide a model that would be less sensitive to viscous effects and that would also aid in the evaluation of these effects. A model support sting geom­etry was also specified as to the maximum diameter and minimum length necessary to provide turbulent base pressure data essentially free of interference (Fig. Ib). These configurations were also submitted for the consideration of the AGARD Fluid Dynamics Panel; at its July 1963 meet­ing this panel expressed interest in the addition of these models to the AGARD series of standard models, pending review of the already avail­able data on shapes HB-1 and HB-2.

Since the publication of Ref. 2, some additional tests with Models HB-1 and HB- 2 have been conducted at the VKF as well as by other organizations. It appears, therefore, that a diverse enough sampling of test measurments has become available to justify a comparison of data. To date, Gther organizations have tested only the HB- 2 configuration with the predominant emphasis being on force measurements.

In particular, the new data consist of force measurements in the 12-in. hypersonic tunnel (Gas Dynamic Wind Tunnel, Hypersonic (E» and

Manuscript received June 1964.

1

AEDC-TDR-64-137

50-in. hypervelocity tunnel (Gas Dynamic Wind Tunnel, Hypersonic (H» and selected pressure measurements in the 40-in. supersonic tun-nel (Gas Dynamic Wind Tunnel, Supersonic (A», the 50-in. hypersonic Tunnel B, and Tunnel H of the VKF. Data from other organizations con­sist notably of force results from ONERA in France (Ref. 3), as well as the Sandia Corp. (unpublished) and RARDE (Ref. 4). Some pressure distribution results were obtained by the Douglas Aerophysics Labora­tory (DAL) (Ref. 5) as well as ONERA. In addition, some VKF ~ynamic stability measurements at supersonic speeds and heat-transfer data: at Mach 18 are also included. Pertinent details of the test conditions may be found in Table 1.

2.0 TEST EQUIPMENT

Both configurations are axisymmetric, blunted cone-cylinders distinguishable by the addition of a flare to the HB-l model to yield the HB-2 model. The geometry of these shapes is detailed in Fig. la, and it may be noted that the junctures of the nose and flare with the cylinder are made by smooth radius fairings. As a matter of interest, the tangency locations for the various sections of the HB- 2 configuration are given below:

Tangency Locations

Section xli-Hemisphere - Cone I 0.0353 I

Cone-Shoulder O. 1064

Shoulder- Cylinder 0.167

Cylinder- Radius 0.583

Radius - Flare O. 725

As shown in Table 1, many model sizes have been tested, and it is believed that all conform to the geometry shown in Fig. la. Obviously there are small variations to be expected among the models; 'however, it is considered that they are the usual fabrication tolerances and thus should be a negligible source of discrepancies because of the blunt nose.

All test models were supported by a sting, the geometry of which was a variable among the various tests, as detailed also in Table 1. The sting proposed by the STA (Fig. lb) was specified to have a constant diam­eterno more than O. 3d for a length of at least 3. Od with a windshield of

2

AEDC-TDR-64-137

20-deg half-angle in order to ensure negligible interference on the base pressure for turbulent flow.

Only test data for which the model was "smooth" are presented; thus, transition location was freely determined by the test facility and model conditions. Only the French (Ref. 3) have investigated the effect of roughness to fix transition; these data are not included herein.

All of these test results were obtained in the wind tunnels briefly described in Table 1. More information on the tunnel characteristics may be found in the references cited.

Inasmuch as diverse facilities are involved in the comparisons to be made, any quotations regarding measurement precision are con­sidered of little value, especially since the basic data required are generally unavailable. It is indeed unfortunate that no widely accepted and common method exists for evaluating data precision as applied to wind-tunnel testing. In certain cases where deviations are clearly re­lated to precision, they will be so noted in the discussion. Otherwise, it is considered that precision is only one of many indeterminate values being judged by the comparisons.

3.0 RESULTS AND DISCUSSION

3.1 FORCE MEASUREMENTS

The normal-force and pitching-moment coefficient characteristics of model HB-1 are shown in Figs. 2 and 3. At most Mach numbers, data from two facilities and two model sizes are available for comparison, whereas at Mach 5 and 8 results from several sources are compared. The Reynolds numbers were selected as close as possible to being the same at each nominal Mach number; thus, the differences shown should be attributable, for the most part, to precision of measurement factors. With the exception of Cm at Mach 2 (Fig. 3a), the curves have been faired with due consideration given all sources to establish a correlation of the data. Although a significant discrepancy in Cm at Moo = 2 exists for the 4.0-in. model, a thorough examination of all possible causes failed to re­veal a satisfactory reason. As will be documented later in Figs. 6 and 7, these curves may be considered applicable for a fairly broad Reynolds number range.

The static stability characteristics of HB- 2 are documented in Figs. 4 and 5. More data are available for comparison because of the

3

AEDC-TDR-64-137

extensive tests reported by the ONERA for the HB- 2 model., As a matter of interest and additional verification, values derived from integration of longitudinal- circumferential pressure distributions (Ref. 9) are included for a nominal 10-deg angle of attack. In general, it appears that excel­lent agreement exists between the force data. The most significant dif­ferences are apparent for the pitching-moment coefficient in the hyper­sonic range, Moo \ 2:; \ 6. However at Mach 8, for example, the variations in Cm for a\;": 12deg (see Fig. 5b) amount to less than a ±3-percent variation in center of pressure.

Inasmuch as the forebody axial force is not greatly influenced by angle-of-attack variation, a comparison of angle-of-attack data is con­sidered unnecessary. However, typical Mach number effects on CA versus a may be found in Ref. 1. It is shown that CA decreases slightly with a for Moo<. 3, whereas it increases with angle of attack at Mach num­bers above 3.· .

The influence of Re~nolds number on the parameters eNa' xcp/.Q, CA, and CAb are shown in Figs. 6 through 9 for model HB-1 and in Figs. 10 through 13 for model HB-2. With regard to the influence on test data for model HB-1, it may be said, in general, that the Reynolds number effect is quite small. At Mach 20, however, the forebody drag (Fig. 8b) is especially sensitive to Reynolds number variations because of the large friction contribution associated with the very low local Reyn­olds numbers. Fewer data are shown for base drag in Fig. 9a, since those data are excluded for which the sting did not approach the specifi­cation. The base drag coefficient appears to be rather insensitive to Reynolds number, when considering the trends for IVlach 8 and 10.

Conversely, model HB-2 aerodynamic characteristics are shown in Figs. 10, 11, and 12 to be significantly dependent upon the Reynolds number. At supersonic speeds (Moo \:5' 5), these effects are caused by the occurrence of laminar boundary-layer separation induced by the inter­action of the boundary layer with the adverse pressure gradient at the flare. Pitching perturbations produce circumferential variations in the extent of separation, thus causing abnormal differential pressure loadings, particularly in the region near reattachment. Since the relative extent of this separation is an inverse function of the Reynolds number (Ref. 10), the variations are to be expected. At hypersonic speeds, flow separations tend to disappear, but when they do exist the longitudinal pressure gra­dients are significantly less than those found at intermediate supersonic speeds. Thus, the rapid changes in pressure between the windward and leeward sides of the flare do not occur. As a consequence, Reynolds number has a small effect on the initial normal-force curve slope above Mach 5 (see Fig. lOb). The initial center-of-pressure location is affected

4

AEDC·TDR·64·137

in a similar manner, of course, as is evident in Fig. 11. From these data in Figs. 10 and 11, it may be observed that the agreement among the various data sources is excellent; hence, the fairings are judged to be suitable reference correlation curves.

The forebody axial-force coefficient of model HB-2 at zero lift is given in Fig. 12 as a function of Reynolds number, Red, for the Mach number range, 1. 5 to 20. Wave drag coefficients (pressure integra-tions) at maximum Reynolds number, primarily from Ref. 9, are in­cluded and show, as expected, that the friction drag is nearly insignifi­cant at these Reynolds numbers up to Mach 8. An interesting difference in the influence of Reynolds number is shown by the data at Mach numbers greater than and less than Mach 5. At low supersonic speeds, for example Mach 3, decreasing Reynolds number is in the direction for increasing extent of flow separation. Flow separation as well as thickened boundary layers decrease the initial pressure rise on conical frustums; thus, the forebody axial-force coefficient would be expected to decrease with a reduction in Reynolds number because the drag contri­bution of the flare is correspondingly reduced. However, at hypersonic speeds the flare pressure distribution is basically unaffected by com­parable Reynolds number variations (see Fig. 17). Hence, the wave drag remains nearly fixed, and the low local Reynolds number, which is caused by the increased bow-wave total pressure losses, now produces a rela­tively large skin friction contribution which is the reason for an increasing drag coefficient as Reynolds number is reduced. That this effect becomes quite appreciable is especially clear at Moo ~. 16, which is shown in Fig. 12b. These data as well as the data for HB-1 from Fig. 8b are re­plotted in Fig. 13 in terms of the viscous interaction parameter, v 00' It should be noted that for infinite Reynolds number (when voo equals zero) these data extrapolate to the value indicated for the wave drag only.

The base axial-force coefficient variation (for a = 0 deg) with Reyn­olds number is indicated in Fig. 14 for model HB-2. Since the data at each Mach number are fairly limited, a true correlation is not possible. However, the data at Mach 5 and 8 indicate that excellent correlation is to be anticipated.

The effect of Mach number on the parameters discussed is shown in Fig. 15 for both model HB-1 and model HB-2 at the condition of maximum Reynolds. number as indicated by the double abscissa. For reference, the magnitude of these parameters as predicted by the modified Newtonian theory at Mach 10 is indicated. Only the drag ,is c1oselyestimated. The declining base drag contribution at high Mach numbers is particularly evident in Fig. 15c; however, the trend at Mach 10 suggests that the assumption of zero base drag for Mach 20 may not entirely be justified.

5

AEDC-TDR-64-137

It should be noted that these base drag data were obtained by measuring the cavity pressure inside the model.

Dynamic stability data for model HB-2 are shown in Fig. 16 in terms of the damping-in-pitch derivatives versus Reynolds number at Mach 2, 3, and 4. These data were obtained by Ward of the VKF using a small-amplitude, forced-oscillation, balance system which incorpo­ra:te~ a cross flexure for pitch oscillations only. The data were obtained at the natural frequency of the model-balance system and were corrected for model damping in still air. Although no other results are available, these data show the usual decrease of damping-in-pitch with Mach number increase and a pronounced dependence-on Reynolds number at Mach 4.

3.2 PRESSURE AND HEAT.TRANSFER MEASUREMENTS

Selected longitudinal pressure disL .. JJution results for zero-degree angle of attack are presented in Fig. 17 for Mach numbers from 1. 5 to 18 at high and low Reynolds numbers. Although more detailed results may be found in Ref. 9, these data show the variation of the Reynolds number in­fluence with Mach number for model HB-2. Since the flare angle is small, only laminar flow separations are possible (Ref. 10). The first indication of separation is to be noted at Mach 2 by the slight pressure in­crease near the beginning of the flare (xl i. '" 0.6). This is the plateau pressure rise caused by the stable reverse flow region ahead of the flare. This pressure rise clearly increases at Mach 3 and 4 for the low Reynolds number condition. Since there is no flow separation at the maximum Reynolds number, the boundary layer obviously was turbulent. At Mach 5, flow separation was present even at the maximum Reynolds number, since boundary-layer transition was delayed sufficiently by the Mach number in­crease. However, data at Mach 8 and 18 demonstrate quite clearly the elimination of laminar flow separation through the influence exerted by the appreciable Mach number increase. Thus, as noted already in Figs. 10 and 11, significant Reynolds number effects caused by flow separation are to be expected only in the supersonic speed range.

Data obtained at only Mach 8 and 20 are currently available for model HB-1, and these are given in Fig. 18. They are also included in Fig. 17 to illustrate better the flare pressure rise. A comparison of available pressure data is made in Fig. 19. The differences arenegli­gible at Mach 2 and 3, whereas they become increasingly pronounced with Mach number. This situation exists simply because of the resolution capabilities of the transducer system employed. At Mach 8, however, a higher resolution transducer was used for stations at xl i. = O. 16, 0.30, and O. 45, and the agreement is markedly improved. The agreement,

6

AE DC-TDR-64-137

however, is much better between the ONERA Mach 16. 5 data and VKF data at Mach 18 because the transducers used were chosen specifically for one pressure range.

The only reliable heat-transfer data presently available are those at zero lift, for about Mach 18, shown in Fig. 20. Results for both. models are presented in terms of the heat transfer relative to the stag-, nation level versus the nondimensional surface distance. The heat-transfer rates, which were measured with calorimeter-type gages, for both models are in excellent agreement.

Longitudinal pressure distributions at three circumferential posi­tions are given in Fig. 21, at an angle of attack of approximately 10 deg, for model HB- 2. The corresponding distributions of local normal-force coefficient and local axial-force coefficient are shown in Fig. 22. It may be observed that at Mach 3 or greater the load distributions are qualitatively the same.

4.0 CONCLUSIONS

All of the available experimental results on the standard ballistic models HB-1 and HB- 2 have been examined in some detail to establish correlation curves which document the aerodynamic performance of these configurations. A few general observations, to highlight the most noteworthy characteristics of these models, follow:

1. Although the aerodynamic parameters for model HB-l are only slightly affected by Reynolds number, the greate~t effect is ob­served on drag at Mach 20. Mach number variations change the stability parameters to an appreciable extent for Moo S' 5, whereas the zero-lift forebody axial-force coefficient is rela­tively invariant for Moo S 10.

2. Because of laminar boundary-layer separation, all aerodynamic (stability and drag) parameters for model HB-2 in the supersonic range (2 is Moo'S' 5) are sensitive to Reynolds number variations; however, the data are satisfactorily correlated by the Reynolds number, Red, at all Mach numbers regardless of model size. At high Mach numbers (Moo> 8) the zero-lift axial-force coef­ficient becomes sensitive to Reynolds number variation because of viscous effects.

3. In general, excellent correlation was found among the test data from various facilities in the United States and France, and from from one facility in England.

7

AEDC-TDR·64-137

REFERENCES

1. Hills, R. "A Review of Measurements on AGARD Calibration Models. " AGARDograph 64, November 1961.

2. Gray, J. Don and Lindsay, E. Earl. "Force Tests of Standard Hypervelocity Ballistic Models HB-1 and HB- 2 at Mach 1. 5 to 10." AEDC-TDR-63-137, August 1963.

3. Private Communication from P. Carriere, ONERA, to J. Lukasiewicz, AEDC-VKF, January 1964.

4. Cable, A. J. and Bowman, J. E. "Hypersonic Wind Tunnel Measure­ments of the Normal Forces and Centres of Pressure of Three C9ne-Cylinder-Flare Models." RARDE Memorandum (B) 1962.

5. Deem, R. E. "Pressure and Heat-Transfer Distribution Tests on a Standard Ballistic-Type Model HB-2 at Mach 2 to 5, Including the Effects of Shock Impingement." Douglas Aircraft Co, Inc., Report S-R6, June 1961.

6. Test Facilities Handbook (5th Edition). "von K~rm~n Gas Dynamics Facility, Vol. 4." Arnold Engineering Development Center, July 1963.

7. Maydew, R. C. "Sandia Laboratory Aerodynamic Test Facilities. " SC-4937(M}, September 1963.

8. Rebuffet, P. "Les Souffleries Hypersoniques de Recherches de l' O. N. E. R. A." Jahrbuch der W. G. L. " 1960.

9. Jones, J. H. "Pressure Tests on the Standard Hypervelocity Ballistic Model HB-2 at Mach 1. 5 to 5." AEDC-TDR-64-, to be published.

10. Gray, J. Don. "A Correlation of Axisymmetric Laminar Flow Separation Characteristics." Paper presented in AIAA 1st Annual Meeting in Washington, D. C., June 29 - July 2, 1964.

8

CO

acility Facility Ref.

AEDC, VKF. 40 in. x 40 in. Supersonic Tunnel (A)

AEDC, VKF. 40 in. x 40 in. Supersonic Tunnel (A)

AEDC. VKF, 40 in. x 40 in. Supersonic Tunnel (A)

AEDC. VKF. 12 in. x 12 in. Supersonic Tunnel (D)

AEDC, VKF. 50 in. Mach 8 Tunnel (B)

AEDC. VKF. 50 in. Mach 8 Tunnel (B)

AEDC. VKF, 50 in. Mach 10 Tunnel (C)

AEDC 1 VKF. 12 in. x 12 in. Hypersonic Tunnel (E)

AEDC, VKF. 50 in. Hypervelocitv Tunnel (H)

Sandia Corporation, 18-in.-diam, Hypersonic

RARDE. No. 3 Hypersonic Gun Tunnel

ONERA 85 Chalais, O. 3 m x O. 3 m

LBRA C4Vernon.0.4mxO.4m

ONERA SR 2 Chalais, O. 3-m Diam

ONERA SR'3 Chalais, 0.12-m Diam

ONERA ARC 1 Fontenay, O. 5-m Diam

! AEDC, VKF, 40 in. x 40 in. Supersonic Tunnel (A) 6

AEDC, VKF. 50 in. Mach a Tunnel (B) 6

AEDC, VKF. 50 in. Mach a Tunnel (B) 6

AEDC. VKF. 50 in. Hypervelocity Tunnel (H) 6

Douglas Aerophysics Laboratory Supersonic 4-ft Tunnel 5

ONERA Arc 1 Fontenav. O. 5-m Diam 3

David Taylor Model Basin, 13.5-in. Hypersonic Tunnel -

*Types of Tests: F - Static Force (except as noted) P - Pressure Distribution H - Heat-Transfer Distribution

TABLE 1

SUMMARY OF EXPERIMENTAL RESULTS

Data Model Sting Nominal Ref. Corlfig. d, in. Did Lid I~. de.

Red x 10- 6 Ct, deg Mach Number

F

HB-I &-2 4.00 S' O. 3 ~·3, 0 20 1.5, 2, 3, 4, & 5 0.35 to 2.53 -1 to 14 x Unpub. HB-2 4.00 2, 3, 4 0.40 to 2.00 x

HB-1 &-2 1. 25 0.6 & O. 8 6.4 1.5, 2, 3. 4, & 5 0.11 to 0.78 -1 to 14

HB-1 &-2 1. 25 0.6 & O. 8 6.4 1.5, 2, 3, 4, & 5 0.04 to 1. 68 -1 to 14 I x

HB-I &-2 7.50 0.3 3.0 20 0.55 to 2.18 -2 to 14

HB-I &-2 4.00 0.3 3.0 20 0.29 to 1. 12 -2 to 14 I x

HB-l &-2 7.50 0.3 3.0 20 10 D. 19 to 1. 36 -2 to 14Jx Unpub. HB-l &-2 2.50 0.3 3.0 20 5, 6. 7, & 8 D. 60 to 2.65 -2 to 12

Unpub. HB-l &-2 2.00 -0.5 -6.0 18 - 21 0.01 to D. 09

Unpub. HB-l &-3 2.00 0.38 2.7 5.3&7.5 O. 10 to 0.20 -2 to 20

HB-:2 1. 00 SO.5 <4.0 8.4 & 10.3 0.14 & 0.30 +3 to 21

HB-:2 1. 18 0.5 2.0 14 1.6, 2.2, 3.1 O. 16 to 0.23 -5 to a HB-:2 1. 18 0.5 1.5 2.1, 3.2, 4.3 0.41 & 0.82 -4 to 8 I x HB-2 1. 18 0.5 1.9 6, 7 0.82 -5 to 10 HB-2 1. 18 0.5 4.2 7.2 10 0.20 -5 to 10 I x HB-2 2.35 0.4 5.3 16.5 0.16 -1 to 10

9 HB-2 4.00 <0.3 >'3.0 . , 2, 3, " & , O. "3 to 2.81 -3 to 15

Unpub. HB-1 7.50 0.3 3.0 8 0.62t02.13 1-15to 15

Unpub. HB-l &-2 4.00 0.3 3.0 8 0.18 to 1. 07 1-15 to 15

Unpub. HB-1 & -2 4.00 nO.5 n6.0 17.6 to 20 0.03 to O. 12

5 HB-2 3.75 0.53 3.0 2, 3. 4, & 5 3.8 to 7.1 I o & 20

3 HB-2 2.35 - - 16.5 0.16 1 o & 5 Unpub. HB-l & -2 2.75 0.3 - 7 0.10 to 0.63

Tests* p H

Remarks

Dynamic Stability Data

Axial-Force Measurements Only

No Axial-Force Measurements

With Boundary-Layer Trips Also

Preliminary Data

]IT--IVL = 2 & 4, a = 0 only

Preliminary Data

Unresolved Discrepancies > m o () , -I o ;;0 , 0-~ . -' W .....,

.300d..J

AEDC-TDR-64-137

1---------4.900d---------I

1-------4.238 d,------------I 1-----4. 083d---------I

LO.700d I \.-~-1. 950d-----i Moment Reference

HB-l

O.362d 40 900d-----:--------I-1

,. i~~ g~~ ~~'-.,..----------1--'"":1'00----1. 702 d' -1 To. 200-::....:::d~---4IQ- _~ ___ _

I .?~o i t - ri I - t - ---~1.600d

O.300d-l~ i 1\ 1 ~. III LO.700d I ~ f '. 1. 950d-"'i 4.00d

Moment Reference 4.500d! 100

1.+2.050d

---t

\101788\ HB-2

(I. Models HB·l and HB-2

fig. 1 Standard Model and Sting Configurations

11

A E DC- T DR-64-137

I i"'"114~--L= 3.0d-----I1IIoo-I

-.,....-----.....

D = 0.3 d

b. Sting G!!!ometry

Fig. 1 Concluded

12

1.2

0.8

0.4

o

z u -~ 0 s::: Q)

u ~ -Q) 0 u Q) U L.. 0 u... I

0 ro E L.. 0

Z

o

o

-0.4 -4

l.

II )r

~rI' l/ V p

~ V )~ ~

J I{ H if r"

/ V / ~ ~

..J::t" ~ ,/ C1' ;/

/ V / ~

d p.c

~ ~ if

/ Y

V" ./

~ V <Y V{

......

LJ r ,..A V

~p-~ "AjV.

o 4 8 12 Anqle of Attack, a, deg

16

Sym M '"

0 1.49

• 1.50 /),. 1.45

0 2.00·

• 2.00 /),. 2.00

o 3.00

• 3.01 /),. 2.99

o 4.01

• 4.03 /),. 3.97

0 5.05

• 5.08 /),. 5.02 V 4.98 Q 5.30

a. Mach Numbers 1.5. 2, 3, 4, and .5

Red x 10

0.51 0.43 0.32

0.86 0.54 0.82

1. 23 0.78 1.00

0.55 0.62 0.58

0.44 0.34 0.35 0.88 0.20

-6

AEDC.TDR·64-137

d, in. Facility

4.00 VKF (A) 1.25 VKF {A) 1.25 VKF (D)

4.00 VKF (A) 1. 25 VKF (A) 1.25 VKF (D)

4.00 VKF (A) 1. 25 VKF (A) 1. 25 VKF (D)

4.00 VKF (A) 1. 25 VKF (A) 1.25 VKF (D)

4.00 VKF (A) 1. 25 VKF (A) 1. 25 VKF (D) 2.50 y-J{F (E) 2.00 Sandia

Fig. 2 Normal.Force Coefficient versus Angle of Attack for Model HB-l

13

AEDC-TDR-64-137

0.8

0.4

o ~'rN'

z u ...... c Q)

0 u ~

t,..N" -Q) 0

U Q) u "-0 u.. I

ro E 0 "-0 z ~~

o <~

-0.4 -4

II ./ V

...; Vfl

~ ~jV

... V V f';f"

-""

~ F7' d

~ ~ ~ ~

Lw ~ ~

~ IfY

~ V yAf" P"'"

..t:Y'

o 4 8 12 Angle of Attack, a, deg

.l1li bJf

A

16

Sym M~ Red x 10-6 d. in. Facility

V 6.04 0.98

V 7.13 1.00

C 8.07 1.05 • 8.09 1.12 V 8.02 1.06 a 7.50 0.17

o 10.18 1.36

2.50 VKF (E)

2.50 VKF (E)

7.50 VKF (B) 4.00 VKF (B) 2.50 VKF (E) 2.00 Sandia

7.50 VKF (C)

b. Mach Numbers 6, 7, 8, and 10

Fig. 2 Concluded

14

o.

o.

E u

-c: Q)

R v

4

0

0

E 0 o :E

I 0> c:

..r:::.. C,,)

:=: c..

o

o

-0.4 -4

II In 0

~ ~ '"lI~

.~

~ rr

~ ?' 0""

.......-4 ~ !O ~ ~

.d ~p o'\~

.0'" ()

~ pf" 0......

V' ~ ~

~ ......- ,r .--Z 1-.J"-..,.,j

I'"\. .Il:lO ~ r' ., .... r<7 v' ....,

~J

~-

o 4 8 12 16 Angle of Attack, a. deg

Sym M 00

0 1. 49

• 1. 50 t:. 1.45

o 2.00

• 2.00 e. 2.00

o 3.00

• 3.01 t:. 2.99

o 4.01

• 4.03 b. 3.97

0 5.05

• 5.08 t:. 5.02 V 4.98

" 5.30

a. Mach Numbers 1.5, 2, 3, 4,· and 5

Red x 10

0.51 0.43 0.32

·0.86 0.54 0.82

1. 23 0.78 1. 00

0.55 0.62 0.58

0.44 0.34 0.35 0.88 0.20

-6

AEDC-TDR-64-137

d, in.

4.00 1. 25 1. 25

4.00 1. 25 1. 25

4.00 1. 25 1. 25

4.00 1. 25 1. 25

4.00 1. 25 1. 25 2.50 2.00

Facility

VKF (A) VKF (A) VKF (D)

VKF (A) VKF (A) VKF (D)

VKF (A) VKF (A) VKF (D)

VKF (A) VKF (A) VKF (0)

VKF (A} VKF (A) VKF (D) VKF (E) Sandia

11017921

Fig. 3 Pitching-Moment Coefficient versus Angle of Attack for Model HB.1

15

AEDC.TDR·64·137

0.4 II ~

'\ .,.AT' o

E u ..... r

0 c Q)

'u ~ ,....-v"

E Q) 0

U ..... C Q)

E 0 ~ 0 ~ i j""1T c::

..c:: u

:!:: 0...

-"""" .,.-<:T o

-0.4 -4 o

'rV'"' iV v

;-v-rr

J'1IIIIoi ~ -'I -v 'V

~ ~ lIT" -v

4 8 12

1'41

v

16

Sylll

V

1I ., 6.04

V 7.13

o 8.07 II1II 8.09 V 8.02 " 7.50

<> 10.18

Angle of Attack, a, deg

b. Mach Numbers 6, 7, 8, and 10

Fig. 3 Concluded

16

-6 Red x 10

0.98

1.00

1.05 1.12 1.06 0.17

1.36

d, in. Facility

2.50 VKF {E)

2.50 VKF (I»

7.50 VKF (B) 4.00 VKF (B) 2.50 VKF (E) 2.00 Sandia

7.50 VKF (C)

1.2

0.8

0.4

o

z u ...... c .~ 0 :.;::: -Q,) o

U Q,) u '­o u... I

ro E '­o Z

o

o

\

-0.4 -4

AEDC-TDR-64-137

• It

~ F In

-'" ~ JJ ./ v ~

d ~ J~

~ P I f L

\

Red' x 10-6 Syrn \.t 00

d,. in. Facility ---

0 1. 50 0.99 ~.OO VKF (A) • 1. 50 0.45 1. 25 VKF (A) A 1. 45 0.69 1.~5 VKF (D) + Pressure Integration (See Fig. 22)

d' / 'fl ~ / ~ /

0 1.99 0.82 4.00 VKF (A) • 2.00 0.54 1. 25 VKF (A) A 2.00 0.52 1. 25 VKF (D) \) 2.12 0.41 1.18 LRBA (C4)

~~ /F' ~ / ~.

o 3.00 1.21 4.00 VKF (A) @ 3. 01 O. 79 1. 25 VKF (A) A 2.99 1. 01 1. 25 VKF (D) + Pressure Integration (See Fig. 22)

If' / ~ \) 3.20 0.82 1.18 LRBA (C4)

/' ~ j{ ~ ~ V

~~ V 0 4.01 0.60 4.00 VKF (A) ., 4.03 0.62 1.25 VKF (A) A 3.97 0.61 1. 25 VKF (D) \) 4.26 0.82 1.18 LRBA (C4)

r5 )(' .) V V

;;I ~ /f> 0 5.05 0.44 4.00 VKF (A) .. 5.08 0.34 1.25 VKF (A) A 5.02 0.35 1.25 VKF (D) V 4.98 0.85 2.50 VKF (E) £I 5.30 0.10 2.00 Sandia + Pressure Integration (See Fig. 22)

L /f> '(f" :/

rtf

0'" itf

~ ~ zV

o 4 8 12 16 Angle of Attack, a, deg

a. Mach Numbers 1.5, 2, 3, 4, and 5

Fig • .4 Normal.Force Coefficient versus Angle of Attack for Model HB.2

17

AEDC-TDR-64-137

1.2 II 0.8

../ V 0.4

~ V

.JJ ~ o

,....

z: U

._r 0 C

CP

U ;;::

--ll V ~

ip'

~ ~ -CP 0 r

U CP U "- ~ 0

L..I.. I

ro 0 E "-0

W ~ t~

z:

~ ~ ~

ft ~ o

p

V 1/ l,.P'

iff o

-0.4 -4 o 4

Jl

./ r Vr-'

r7

~ ~

Ltt' Ff

'" V ,c

~ 'j1l"l.J

I""

A

D V " ~

;;;, y r

.,./ lP

8 12

~

16

Sym Moo Red x 10-6 d, in. Facility

V 6.04 (I 6.00

V 7.13 (I 7.00

0.98 0.82

1.00 0.82

2.50 1.18

2.50 1.18

VKF (E) ONERA R2 Chalaie

VKF (E) ONERA R2 Chalais

[] 8.07 1.07 7.50 VKF (B) III 8.09 1.12 4.00 VKF (B) V 8.02 1.07 2.50 VKF (E) £I 7.50 0.17 2.00 Sandia l:> 8.40 0.30 1.00 RARDE + Pressure Integration (See Fig. 22)

<> 10.18 l:> 10.30 II 10.00

o 16.5

1.36 0.14 0.20

0.16

7.50 1.00 1.18

1.18

VKF (C) RARDE ONERA R3 Chalaie

ONERA ARC 1 Fontenay

Angle of Attack, a, deg

b. Mach Numbers 6, 7, 8, 10 and 16.5

Fig. 4 Concluded

18

0.4

o

o

E u -~ c

0 Q,)

CJ :: -Q) 0

U .-c Q)

E 0

:2: I 0 0>

C

.c CJ

:!:::: a..

o

-0.4

-0.8

-1. 2 -4

II 0.

~ ~ l;l iI\-I..

~~ U""l~

h

~ ~ h~ ......... ~l

I~~

Q" -1\ ~ ~'a 1

~ ~ F\ ~

~ '" ~ ~ ~ ~ ~ . ~o

~ ~ l~ ~ ~ ~ ~

~ 'Ml ~

M .~

~

~ r'\

o 4 8 12 16 Angle of Attack, a, deg

Sym M", -6 Red x 10

AEDC-TDR·64·137

d, in. Facility

o 1.50 0.99 4.00 VKF (A) • 1.50 0.45 1.25 VKF (A) 6 1.45 0.69 1.25 VKF (D) + Pressure Integration-(See Fig. 22)

0--1.99

• 2.00 6 2.00 o 2.12

0.B2 0.54 0.52 0.41

4.00 1.25 1. 25 LIB

VKF (A) VKF (A) VKF (D) LRBA (C4)

o 3.00 1.21 4.00 VKF (A) • 3.01 0.79 1.25 VKF (A) A 2.99 1.01 1.25 VKF (D) + Pressure Integration (See Fig. 22) o 3.20 0.B2 1.lB LRBA (C4)

0 4.01 0.60 4.00 VKF (A) • 4.03 0.62 1.25 VKF (A) A 3.97 0.61 1.25 VKF (D) 0 4.26 0.B2 LIB LRBA (C4)

0 5.05 0.44 4.00 VKF (A) • 5.0B 0.34 1.25 VKF (A) 6 5.02 0.35 1.25 VKF (D) V 4.98 0.85 2.50 VKF (E) £I 5.30 0.10 2.00 Sandia + Pre~sure Integration (See Fig. 22)

110~ 7961

a. Mach Numbers 1.5, 2, 3, 4, and 5

Fig. 5 Pitching-Moment Coefficient versus Angle of Attack for Model HB.2

19

AEDC.TDR·64~131

0.4

o

E 0

U

-~ c:: Q)

.~ -'Q; 0

U ..... 0 c:: Q)

E 0

::E I 0> c::

.s::: J;:a a:::

0

o

-0.4

.~

~ II ~ ~ ~ ~

r~

~ n t:1"

~ ~ ~ ~

,

tM. -0 ~ ~

0

'he.-. ~ ~ ~

~ ~

D. ~ ~v.." ~ ~ Cr-"",

~

'<;.J ---~ I"'r"l

~

~

~ X ~

'. ~L

h.

P,\

r\

8ym M~ Red x 10-6 d, in. Facility

V 6.04 0.98 2.50 VKF (E) o 6.00 0.82 1.18 ONERA R2 Chalais

V 7.13 1.00 2.50 VKF (E) o 7.00 0.82 1.18 ONERA R2 Chalais

D 8.07 1.07 7.50 VKF (B) III 8.09 1.12 4.00 VKF (B)

V 8.02 1.07 2.50 VKF (E) a 7.5 0.17 2.00 Sandia .. 8.4 0.30 1.00 HARDE

+ Pressure Integration (See Fig. 22)

o 10.18 .. 10.30 Q 10.00

o 16.5

1.36 0.14 0.20

0.16

7.50 VKF (C) 1.00 RARDE 1.18 ONERA R'3 Chalais

1.18 O~RA ARC 1 Fontenay

-4 o 4 8 12 . 16 Angle of Attack, 0, deg

b. Mach Numbers 6, " 8, 10, and 16.5

Fig. 5 . Concluded

20

I:'-' 1~

Cl Q)

"C W Q) c..

~ z u w· §-V'I w c: ::::s u w ~ .f

I (ij E !.. o

Z

HB-l Model

::::1 J rl-erl f 0 1

0.03 II I I ! , ,J.,,--~

::: I· ~ I -~ 4 I-Iii gil I ! J._. I ! 04 'I O.

:::: I ., 1 1 * ~ I W 1 0 1 01

0.04 ' ! ! ! ... ----........!..--..----!..-----&

:': I -J ~ ~ 1m J_ .1 :::f oJl ~ If I tH~ I~ I 0.03 ~ 6

O. 1 O. 2 O. 4 O. 6 1. 0 2. 0 4. 0 x 10

Reynolds Number, Red

a. Mach Numbers 1.5, 2, 3, 4, and 5

Sym

o • A

o • A

o • l!.

o • l!.

o • 'l!. 'V Q

Moo

1.50 1. 50 1. 45

2.00 2.00' 2.00

3.00 3.00 2.99

4.02 4.02 3.97

5.08 5.08 5.02 5.00 5.30

d, in.

4.00 l.25 l.25

4.00 1.25 1.25

4.00 1.25 1.25

4.00 1.25 1.25

4.00 1.25 1.25 2.50 2.00

Fig. 6 initial Slope of Normal.Force Curve versus Reynolds Number for Model HB.l

Facility

VKF (A) VKF (A) VKF (D)

VKF (A) VKF (A) VKF (D)

VKF (A) VKF (A) VKF (A)

VKF (A) VKF (A) VKF (D)

VKF (A) VKF (A) VKF (D) VKF (E) Sandia

[l?n97] » m o o · .... o ;;a · 0-&>. · w .....

t..:l t..:l

g' "0 ~ Q;I c..

-:2 u Q;I.

8-V')

Q;I

c: ::J

U Q;I

~ .f ..L n::I E ~ o :z

HB-l Model

::: I I I I I I ] 1 lObi O. 03 ' !!!! I --1 !

:::11 Illhl ~vv I I

::: I ~ I j IIi80 k u Ie I

::: I II II Rio II 0.03 6

o. I o. 2 -0. 4 O. 6 1. 0 2. 0 4. 0 X 10

Reynolds Number, Red

b. Mach Numbers 6, 7, 8, and 10

Fig. 6 Concluded

Sym

v

v

o II1II

V £J

<>

Mea

6.06

7.07-7.20

8.07 8.07

7.80-8.06 7.50

10.15

d, in.

2.50

2.50

7.50 4.00 2.50 2.00

7.50

Facility

VKF (E)

VKF (E)

VKF (B) VKF (B) VKF (E) Sandia

VKF (C)

F i981

» m o n · -I o :::tI · 0-

"'" · ..... to> .....,

~ -c.. XU

c,-

.!2 -n:II

~ Q) II-~ en en Q)

I.\.:) II-

CJ,j Il.. I -0 I II-Q) -c, Q)

u

HB-l Model 0.2

0.1

0 I 'I rtfftE I I 0.3

0.2

0.1 ~j \IQLFm_:n __ ~

0.4

0.3

0.2 1·1 kltHf:lo I

0.4

0.3

0.2 I ! i \-1 If 1ft I

::: ~ a t X I ~ II t I + E I 0.2 6

O. 1 O. 2 O. 4 O. 6 1. 0 2. 0 4. 0)( 10 Reynolds Number, Red

la. Mach Numbers 1.5, 2, 3, 4, and .5

Sym

e • IA

e • IA

o • A

o • lJ.

o • IA '1 a

Moo

1.50 1. 50 1.45

2.00 2.00 2.00

3.00 3.00 2.99

4.02 4.02 3.97

5.08 5.08 5.02 5.00 5.30

Fig. 7 Center of Pre·ssure versus Reynolds Number .for Model HB-1

d, in.

4.00 1.25 1.25

4.00 1.25 1.25

4.00 1.25 1.25

4.00 1.25 1.25

4.00 1.25 1.25 2.50 2.00

Facility

VKF (A) VKF (A) VKF (D)

VKF (A) VKF (A) VKF (D)

VKF (A) VKF (A) VKF (D)

VKF (A) VKF (A) VKF (D)

VKF (A) VKF (A) VKF (D) VKF (E) Sandia

[lrn9~

» m o () 'J

-I 0, ::c 0.. .$:I. . w ...,

> m

::: I I I III! t IH

: MOO~I s~ .:. d~::. '::'::: ! 1:::1 I III f1t= I I V 7.~-7.20 2.50 ~ (E)

Q) II... :::l

~ ~ :::! I I II J i ~ V I I ; 7 t~.O. ::~ * ~~l L a dUd iI!i 7. . 50 2 . 00 Sandul.

~ 0.2 ' Q)

u

::: I I I III I I b I 9 '0.15 7.50 m (C)

0.1 0.2 0.40.6 1.0 2.0 4.Ox106 11018001

Reynolds Number, Red

b. Mach Numbers 6, 7, 8, and 10

Fig. 7 Concluded

0.6 ~ l:r-I ~ I ~ I ffjl ~ fA HB-l Model s~ ':0 d:.::. F:':~ 0.5 IT I r IA 1.46 1.25 VKF (D)

0.4 .

<C 0.6 u

'E~ 0.5 I ps I ~I~r=r= - .... I A 2.00 1.2,5 VKF (D) (I,)

u

li 0.4 c u

: ::: I j Q f I iEft I 0 I : ~:~~ t:~g ~ r.l ~ i 0.4

i ::: ~"'--3---'--""'I-f"-· :....,ffi---.r-.. 1ft I :;: ~~ t: ~g ig ~~l 0.6

0.5 E±fl rEf r~ J 0.4

0.1 0.2 0.4 0.6 1.0 2.0 4.0x 106

Reynolds NumbE!f. Red iii. Mach Numbers 1.5, 2, 3, 4, and 5

0 08 4.00 I::. 5.01 1.25 V 5.00 2.50 a 5.30 2.00

Fig. 8 Zero-Lift Forebody A,dai.Force Coefficient versus Reynolds Humber for Model HB.l

VKF (A) VKF (D) VKF (E) Sand1!1.

11018011

> m o o -I o AJ . 0-01>0. . w .....

~ m

0.61 0.5

~ 0.61

..,J' 0.5 c: Q)

:~ -as 0

U Q)

~ af

I iij

~ ~ .8 Q) ~

af ..--::i

I 0 ~ Q)

N

0.6

1

0.5

0.4

0.61 0.5

0.1

0.8

0.7

0.6

0.5 0.01

I

I

a-

a 1

"I

I -

I 0.2

0.02

HB-l Model

I I I [L Ivv I

i I LIe Im ___ ~

I I, rr Vo

1

0

I

I I III 0

0.4 0.6 1.0 2.0 4.0 X 106

Reynolds Number, Red

0.04 0.06 0.1 0.2 O. 4x 106

Reynolds Number, Red

b. Milch Numbers 6, 1, So 10, and 20

Fig. 8 Concluded

Sym ill", d, in. Facility >-m

VlQ!' (E) 0

V 6.06 2.50 (') · -I 0 :;0 · a.. &>0-• -

V 7.07-7.20 2.50 VKF (E) w ......

EJ 8.07 7.50 VKF (B) , 8.07 4.00 VKF (B) v 7.80-8.06 2.50 VKF (E) S 7.50 2.00 Sandia

~ 10.15 7.50 VKF (C)

a 19.3-21.3 2.00 VKF (H)

11018021

::: FIR III ~ rlj s: '":0 d~.:n. F:':~

: :; ~ 1 1111+ 1 0 1 I 0 ~.oo 4.0 m ~) ~

! ::; I III E II 0 I ~ I 0 3.00 4.0 m ~) t\.? CI.)

~ i _::: I 111m 1 1 0 4.02 4.0 m ~)

_::: f 0 j . I IfEFf E I 6 ~ Hg t~ i~d~~ ~ 0.1 0.2 0.4. 0.6 1.0 2.0 4.0xlO 1101803\ g

Reynolds Number, Red ~ o ;;0

a.. Mach Numbers 1.S, 2, 3, 4, and 5 0.. 01:>. ,

Fig, 9 Zero-Lift Base AxiialmForce Coefficient versus Reynolds Number for ModellHB.l w 'I

> m 0 (')

HB-I Model d, in. Facility · Sym Moo -I

~ 0.: I I I I II If I_ I

-- 0 ;;0 · 0-

V 6.06 2.50 VKF (E) .l>-· ~ w ......

« -0-1 u . ..,.;

0.: I c

I I IIHf I Q)

:~ VKF (E) - V 7.07-7.20 2.50 Qj

0 u Q)

~ -0.1 0 u... I

'iii

0.: I ' I I I Itl- r r< In I 'x a 8.07 7.50 VKF (B) « iii 8.07 . 4.00 VKF (B) t'-' Q) C

V 7.80-8.06 2.50 VKF (E) VI ():) CQ C 7.50 2.00 Sandia OC -- -0.1 :::i

I e

_::: I

Q)

I I II j QI I N

~ 10.15 7.50 VKF (C) ¢

II0180~ 0.1 0.2 0.4 0.6 1.0 2.0 4.0 X 106

Reynolds Number, Red

b. Mach Numbers 6, 7, 8, and 10

Fig. 9 Concluded

g' "C s-CI) c..

0 :z u Q)-

c.. ..E VI CI) c: :::lI u

~ CI)

co ~ gf

I

'"iii E s-o :z

HB-2 Model 0.09

0.08

0.07 1 i f I-:£! ~ 0 I

0.10

0.09 ,

0.08

0.07

0.11

0.10

0.09

0.08

0.10

0.09

0.08

0.07 0.1

.... ~ ....

'" '" • . -.....

0.2

.... 0 -~ • a Q

~ ...

"lQ.e ~, 0 ,.,

il ..

1

I.

r---~ ..! "'"""Ii~ Ii..

t) I .....

0.4 0.6 1.0 2.0 4.0 x 106

Reynolds INumber, Red

Ci. Mach Numbers 1.5, 2, 3, 4

Sym

o • 4

o if • o

o • • o

o 1 ¢I

Moo

1. 50 1. 50 1.45

2.00 2.00 2.00 2.12

3.00 3.00 2.99 3.20

4.02 4.02 3.97 4.26

d, in.

4.00 1.25 1.25

4.00 1.25 1.25 1.18

4.00 1.25 1.25 1.18

4.00 1.25 1.25 1.18

Fig. 10 Initial Slope of Normal-Force Curve versus Reynolds' Num,ber for Model HB-2

Facility

VKF (A) VKF (A) VKF (I)

VKF (A) VKF (A) VKF (I) LRBA (C4)

VKF (A) VKF (A) VKF (0) LRBA. (C4)

VKF (A) VKF (A) VKF (I) LRBA (C4)

1101805]

> m o () · -I o :;0 · 0-01:>. · w ......

! I-CI,) Q..

-:2 u CI,)-

8-en CI,)

c: ::l

W U 0 CI,)

~ .£

I c; E I-0 :z

HB-2Model;

:::L:IJ4iHJ IQ¥v I 00 08

1 ::: I 1111 B~ I v~ I

::1 I III 1 IT x II

:::1 JIIiI ~ I I 0.05 6

O. 1 O. 2 O. 4 O. 6 1. 0 2. 0 4. 0 x 10

Reynolrts Number, Red

Sym IrI 00

d, in. Faci]~ty

0 5.08 4.00 VKF (A)

• 5.08 1.25 VKF (A) 6/ 5.01 1 .• 25 VKF (D) V 5.00 2.50 VKF (E) .a 5.30 2.00 Sandia

V 6.06 2.50 VKF (E) 1) 6.00 1.18 ONERA R2 Chalais

<J 7.07-7.20 2.50 VKF (E) tt 7.00 1.18 ONERA R2 Chalais

1:11

II1II V t3 CiI

~ Do

8.07 8.07

7.80-8.06 7.50 8.40

10.15 10.00

7.50 4.00 2.50 2.00 1.00

VKF (8) VKF (8) VKF (E) Sandia RARDE

7.50 VKF (C) 1.18 ONERA R'3 Chalais

11018061

b. Mach Numbers, 5, 6, 7, 8, and 10

Fig. 10 Concluded

> m 0 n . -I 0 ;;0

0-01:>. . -W ''-1

HB-2 Model Sym 114 d, in. Facility Q6 . ~

::: I .. I ~ 1 +t I fi 1 1 : Ug tH ~ !~l 0.61 11 ~~ I e 2.00 4.00 VKF (A)

""'I 0.5 If' _ Q • 2.00 1.25 VKF (A) -c. .' . A 0 ~ 2.00 1.25 VKF (I») XU 0 2.12 1.18 LRBA (C4) _ 0.4-

I:

t 0.7

1 W~~El t 1

....J • 0 3.00 4.00 VKF (A) ~ 0.6 . • 3.00 1.25 VKF (A)

~ 0 5 ..~ ~ ~ .... 0 ~ ~:~~ ~:~: I~BA(~~4) IC.. •

t.I:l ~ I-l _

o

! :::1 fl- Hill J I ! ~:H U~ ti~L :::~I A-1i'1.J-~ In I ~ Uf H! ~~! ~ .. 1 []Y .=vv , S5.30 2.00 Sandia > 0.5 - ....J-.....J-. __ ~____ I m

0.1 0.2 0.4 0.6 1.0 2.0 4.0 X 106 1101807 Q -I

Reynolds Number, Red ~ ,

(ll. Mach Numbers 1.5, 2, 3, 4, and 5 ~ .... Fig. 11 Centel' elf F'ressure versus Reynolds Number for Model HB.2 ~

W l\:I

HB-2 Model

::: I I I III t t ~v :

0.4 ~

~O.61 I I IIJ ~~ .~ 0.5

~ ! !! ! L.... .9 0.4 L.-Q) .... :=I VI VI Q) ....

Q.. I

'0 I ....

$ c: Q)

:::1 00

1 jlliltRut I

U ,:::1 ilill tt; , I O. 1 O. 2 O. 4 O. 6 1.. 0 2. 0 4. Oxlob

Reynolds Number, RE!(!

Sym

V 0

V o

EI III V a

~ A

b. Mach Numbers 6, 7, 8, and 10

Fig. 11 Concluded

Moo dl, in. Facility

6.06 2.50 VKF (E) 6.00 LIS ONERA R2 Chalais

7.07-7.20 2.50 VKF (E) 7.00 1.lS ONERA R2 Chalais

S.07 S.07

7.80-8.06 7.50

10.15 10.00

7.50 4.00 2.50 2.00

7.50 1.18

VKF (B) VKF (B) VKF (E) Sandia

VKF (C) ONERA R'S Cha1ais

11009491

» m 0 () . -I 0 ;u . 0-.c:.. , w '-l

HB-2 Model Sym Moo d, in. Facility

::: I 61 :± 1 tffi ~b~~d ~ U~ t:~~ ~g ~l 0.7

u< 0.8 3 I ::: I :±TIll it *1 I ~ n~ U: a~;L) u

! 00.76 F ~I ~ltpH 1 0 __ ~ 0 3.00 4.00 VKF (A) 'x . 6. 2.99 1.25 VKF (D) « 0 3.20 1.18 LRBA (C4) »

tJJ :8 0.5

tJJ ~ 0 7 Uncertalinty Caused by

~ . t t f fero Shift at Max Red ~ I =? 0 6 I t~~ I =l -- 0 4.02 4.00 VKF (A) o • • is- - I 6. 3.99 1.25 VKF (D) ~ . 0 4.26 1.18 LRBA (C4)

Q5--~-· -

o 7 Uncertainty Caused by_r--__ -,

. ~~~ I Zero Shift at Max Red =1 ~ ~:g~ i:~~ ~~ ~~~ 0.6 _. __ A. II LL ~. v 5.00 2.50 VKF (E) . . -n £I 5.30 2.00 Sandia > 0.5 '--__ --..L __ I...--I...--i._.&..:- 6 ~

O. 1 O. 2 O. 4 O. b 1. 0 2. 0 4. 0 x 10 0 -I

Reynolds Number, Red! ~

a. Mach Numbers 1.5, 2, 3, 4, and 5 t . -eN Fig. 12 Zero·lift Forebody A:l(ilClii.Force Coefficient versus Reynolds Number for Model HB.2 "-I

HB-2 Model

0.6 1 I I I I I f: I~ 0.5

Sym MIX)

v 6.06

a> u L-0

u.... I

.~

« 0.6 r-----r--y-I"""""""''''''-----i 0.5 1 I I Iii It'" I iii 0.7 -. I 8

0.6 ~~~~ 0.5

V 7.07-7.20

IJ 8.07 III 8.07 V 7.80-8.06 ~ 7.50

X « >. 0.7

(.Aj "0

~ .8 a> L-0

u.... --~ I

0 L-

~

~ CAb Assumed'Zero -:>

0.6

0.5 O~ 1 0.2 O. 4 0.6 1. 01 2.01 4. Oxl06

<> 10.15 D. 10.00

a> N

::: I mtHb:ss~medzero o 1.8.1-19.8 <I 16.5

O. 01 O. 02 O. 04 O. 06 O. 1 O. 2 O. 4

Reynolds Number, Red

b. M'ach Numbers 6, 7, 8, 10, 16.5, and 20

Fig. 12 Concluded

d, in.

2.50

2.50

7.50 4.00 2.50 2.00

Facility

VKF (E)

VKF (E)

VKF (B) VKF (B) VKF (E) Sandia

7.50 VKF (C) 1.18 ONERA R'3 Chalais

2.00 VKF (H) 2.36 ONERA ARCI Fontenay

~09511

> m o () . -I o AI

0-~ . w '-I

v;l r..n

1.0

<C U

......, c Q)

u --Q) a

U Q)

0.5 u l0-a w-

I

co x

<C ......, .... :.::i

I a lo-Q)

N

Note:

I 0

0

o

Pressure Integrations

,...-

C A Ass u med Zero b

Model

6. HB-l o HB-2 o HB-2

Model

HB-2

HB-l

Moo

18.1 - 19.8 18.1 - 19.8 16.5

! I

0.10 0.20 M -!C - ool''-'oo

Voo = ---

·tRed

d, in.

2.0 2.0 2.36

Facility

VKF (H) VKF (H) ONERA ARC I

[l?1810 I

Fig. 13 Zero-Lift Forebody Axial.Force Coefficient versus Viscous Interaction Parameter for Models HB-l and HB·2

» m o n , --I '0 ;;0 , 0-~ , w ......

HB-2 Model » Sym Moo d, in. Facility m

0.51

I I I I [ I I~ I 0 n ·

VKF (A) -I

0 1.50 4.00 0 ;;0

0.4 · 0-

.I:>-0.3 · w 'I

0.4

1 I I I I [f I o I 0 2.00 4.00 VKF (A) i 0.3

.!:::: 0.2 --Q)

0

0.3 [ u

I I I lEI I a I a I Q)

U L-0

0 3.00 4.00 VKF (A) I.L. , (t:I

0.2

~ 'x 0.1 m -«: Q) VI (t:I

0.3 [ co

I I I If I ~ J I --:::i ,

0 4.02 4.00 VKF (A) 0 L-Q)

0.2 N

0.1

::: t I- I I I lett I~ 16 0 5.08 4.00 VKF (A) V 5.00 2.50 VKF (E) £I 5.30 2.00 Sandia

11018111 O.

4.0 X 106 0.1 0.2 0.4 0.6 1.0 2.0

Reynolds Number, Red

a. Mach Numbers 1.5, 2, 3, 4, and 5

Fig. 14 Zero-Lift Base Axial.IForce Coefficient versus Reynolds Number for Model HB-2

..c -=c u

-' c Q) :§ :::: Q) 0 u Q) u ~

&£ I

n5i 'x <: Q)

CJJ VI nll

-J co =: ::i , 0 ~ Q)

N

HB-2 Model 0.2

0.1

0 I IIIJ1f I w I 0.2

0.1

0 I I I II mi-----i....--.ll 0.1

0

-0.1 1-"1 "II ffJTn f I 0.1

0

-0 1 I f I II t,~--J-..I.--I 0 ~ . 6

O. 1 O. 2 O. 4 O. 6 1. 0 2. 0 4. 0 x 10

Reynolds Number, Red

b. Mach Numbers 6, 7, 8, and 10

Fig. 14 Concluded

Sym

v

TV

o III TV a

<>

Moo

6.06

7.07-7.20

8,07 8.07

7.80-8.06 7.50

10.15

d, in.

2.50

2.50

7,50 4.00 2.50 2.00

7.50

Facility

VKF (E)

VKF (E)

VKF (B) VKF (B) VKF (E) Sandia

VKF (C)

[1018121

>­m o () , -I o ;;0 , 0-~ , w .....

0.10 C') Q)

"'C 10-Q)

Q. 0.08 .. t:i

Z U

~ 0.06 ..2 V')

Q)

> 10-

0.04 :::l u Q) U 10-

~ W I 0.02 00 co

E 10-0

Z o

Modified -Newtonian

~ Theory !

I ~ ~ lrHB-2 I

~ / lU I

..........-.,.. Ir-....

'I~

/ r "'\ r""'--"",-~ P--....

(..., I..n

"-I'- HB-1 ~,

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

IJ Mach Number, Moo I

I I i I

2.0 2.6 2.6 1.3 2.2 1.4 11018131

Red x 10-6 1.4 1. 7 2.5

a. Initial Slope of Normal.Force Curve

Fig. 15 Compariscln of Mach )\4umber Effect on Models HB-l and HB·2

> m o () · -I o :;0 · '" .I:>. · W -.J

~ Q..

u x . c: 0 .-~ u .9 CII) I-::I II) II) CII) l-e..

I

'0 I CJ:)

co I-CII) -c: CII)

U

1.0

0.8

0.6

0.4

0.2

o o I

1/ I

) 1

~ y--- ~ I.-.

~

-- h

~ r Ir-. L-..

~

-2 3 4 5 6 7 8

Mach INumber'l Moo

1.4 1. 7 2.5 2:.0 2.6 2.6 1. 3 2.2

Red'( 10-6

b. Center of Pressure

Fig. 15 Continued

Modified Newtonian

Theory

/ ~HB-2

~

" ~

"'- r-HB-1 c~

9 10 11

1. 4 \101814\ » m o n , -I o Al . ()'o

:.. . w ......

~ 0

.c <C 0.9

U "C c: co «

u 0.7 V)~

<C -c: U Q)

.~ 0.5 --Q) 0 u Q)

~a.

U :- o 0

I..L. I

CO .-x <C 0.4 Q) V) CO

!CO

-g .c 0.2 co <C ~u o ij 5 0 I..L.

;t:: :.::i b :­Q)

N

o

I

1

Modified

~ Newtonian

~ Theory ~ ~--..! 1- ;- HB-2

I,., II~ 1

0..._ ..... '.J' lo---__ ( 1"'\ I,..,.

\..- HB-l

~

~ t..-0.."""-1

........... {

~ I HB-2

I ?---- ~ -- h v:; -

'- HB-l '( -2 3 4 5 6 7 8 9 10 11

Mach Number, Moo I I _I

1.4 1.7 2.5 2.0 2.6 2.6 1.3 2.2 1.4 F~

Red x 10-6

c. Zero-L.ift Forebody and Base Axial.Force Coefficients

Fig. 15 Concluded

:> m o () , -l o ;0 . '" .l>-. w ......

fl:>. I-'

-60 r if -40

u I +

0-E

u -20 g

o 0.2

0-

~D 0

P-

,

0.4 O. 6 O. 8 1. 0 -6

Reynolds Number, Red' x 10

DO

"" 6.

2.0

M 8, deg wd/2V , rad ~ CD

2 ±1.5 0.0050

3 ±1.5 0.0041

4 ±1.7 0.0038

x I J, = 0.398 cg ,

3.0

[l{}~9431 Fig. 16 Damping.in.Pitch Derivatives versus Reynolds Number for Model HB-2 CIt Mach Numbers 2, 3, and 4

> m o <;1 -t o ;:c

0-.!>. •

Co) .....

1.0 0.8

0.6

0.4

0.2

0

..e-o.

.:2- 0.1

~ 0.08

'" ~ 0.06

'" ~

.t 0.04

~

0.02

0.01

0.008

0.006 0

Sym

0

• fj.

.II.

0 lIB

(Typical!

~ \-~B~ (~9~8~ ---! !

0.6 0.7 0.8 0.9 1.0 xlt

11009581

Fig. 17 Longitudinal Pressure Distributions at Zero Lift for Model HB-2 at Mach Numbers 1.5, 2, 3, 4, 5, 8, and 18

Moo Red x 10-6 d, in. Facility

1.48 1.41 4.00 VKF (A) 1.47 0.49 4.00 VKF (A)

1. 99 1.63 4.00 VKF (A) 1. 98 0.32 4.00 VKF (A)

2.99 2.02 4.00 VKF (A) 2.98 0.48 4.00 VKF{A)

4.00 2.00 4.00 VKF (A) 3.99 0.54 4.00 VKF (A)

5.11 2.32 4.00 VKF (A) 5.05 0.48 4.00 VKF (A)

8.09 1.06 4.00 VKF (B)

8.04 0.34 4.00,7.50 VKF (B)

17.8 0.11 4.00 VKF (H) 17.8 0.03 4.00 VKF (H)

» m o n . -I o ;:0

0. """ . w -..j

, 0

..e-o.. 0

:;:::: n:I 0:: Q) ~

::J

'" '" Q) ~ ~ c... uJ

1.0 0.8

0.61-1.

0.4 ~ ),

0.2

0.1 0.08

0.06

0.04

0.02

0.01

0.008

0.006

0.004

0.002

I I

o 2 4

Sym Mm Red x 10-6 Facility d, in. --

0 8.08 1.1 -2.2 VKF(B) 7.5

• 8.08 10.5 -1.0 VKF(B) 4.0 0 17.6,.20.10 10.03-0.12 VKFIH) 4.0

o

) 6 8 10 12 14 16

Surface Distance, Nose Radius, S/RN 11009561 Fig. 18 Longitudincil F'ressure Distributions at Zero Lift for Model HB-l

at Mach Nlullllbers 8 and 20

» m o () . -I o ;:c

0. .;... . W "'-I

J.!::> J.!::>

, 0 Q.

a.

1.0 0.8

0.6

0.4

0.2

~- 0.10 '" 0.08 '" ~ 0.06 e "'" 0.04

0.02

0.01

o

6 Moo = 4.00, Red = 2.00 x 10' 6 Moo = 5.11, Red· 2.32 x 10' 6 Moo = 8.09, Red' 1.06 x 10· 6 Moo = 17.8, Red· 0.11 x 10

0.04 0.08 0.12 0.16 0.2 0.3 0.4 0.5 0.6 10.7 10.8 0.9

xU

Sym ~ Red x 10-6 Facility d, in.

t::. 2.01

0 2.98 l1li 2.98

0 3.98

"V 4.95

D 8.08

16.5

1.0

11009571

3.8 DAL 3.75

7.1 DAL 3.75 3.8 DAL 3.75

3.8 DAL 3.75

4.4 DAL 3.75

1.1 DAL 3.75

0.16 ONERA, 2.34 ARC 1

Fig. 19 Comparison of Longitudinal Pressure Distributions at Zero Lift for Model IHB-2 from VKF, DAl, and molERA

)-

m 0 n · -I 0 ;:0 · 0-~ · w -..j

'0 .!2" 0-

0-;; '" 0:: ... Q)

'Vi I:::

'" .... l-

I

«i Q) ~ ::I: C)l

1.0 0.8

0.6 .- • Rem x 10-6 Sym Config. Moo Facility d, in. --

0.4 J- 0 HB-l 18.4-20.0 0.03-0.12 VKF(HI 4.0

0 HB-2 17.8 0.03-0.11 VKF!HI 4.0

I ~n 0.2

0.1 0.08

0.06

0.04

cr ..........

0.02 I ~ ~

Beginning of Flare, HB-2

0.01 B-U~L 0 0

0.008

0.006

0:004

0.002 I L.._,

--) o 2 4 6 8 10 12 14 16 18

Surface Distance, Nose Radius, S/RN 1100955 1

fig. 20 Longituciinal Heat-Transfer Distribution at Zero lift for Models HB·l anci HB-2 at MClch 18

> m o () · -I o AJ · 0-./>. · w ......

AEDC-TDR-64-137

p/p~

1.0

0.1

0.01

1.0

0.1

0.01

1.0

0.1

nm v. V.L

1.0

0.1

0.01

"-

I I I r-----r-----r-----

~ ,\.

\:

\

\" '\

L I

\\. '\

V o

I'.. .-<r ~'" .",.,..

"- ~.Q" zy<r

"[). .-q- ;...-" "<>---

Sym 0, deg ~

0 0 Red x 10-6 ,d, Moo Facility 0 90 in.

t:. 180 1.48 1.40 4.0 VKF (A)

a. Mach Number 1.5

----M Red x 10-6 d, in. Facility 00

~ 2.99 2.02 4.0 VKF (A) ~'\ '0-.-0-1.<>- vrv

[}-

..t>-~

b. Mach Number 3

Moo Red x 10-6 d, in. facility

"- 5.11 2.33 4.0 VKF (A) "\'\ -<>-'Xl

~

'\ ........... .-A-P' .r:?

I I I J I I I • --

c. Mach Number 5

Moo Red x 10-6 d, in. Facility

'\. 8.09 1.06 4.0 VKF (B) ~ '\ ..DO

\. .). ---."

'\. '0.... 4

~ -...0

" "

A 0.08 0.16 0.2 0.4 0.6 0.8 1. 0

xli-d. Mach Number 8

11018161

Fig. 21 Longitudinal Pressure Distributions for an Angle of Attack of 10 deg at Three Circumferential Positions for Model HB-2

46

AE DC- TDR-64-137

Sym Moo Red x 10 -6 d, in. Facility

--0 1.48 1.40 4.0 VKF (A) 0 2.99 2.02 " VKF (A) b. 5.11 2.33 " VKF (A) \l 8.09 1.06 II VKF (8)

0.16

0.12 cN

0.08 . -1 In.

0.04

0 a. local Normal.Force Coefficient Distribution

1.0 I I

0.8

r. " £ VA U.O I

. -1 In. 0.4

10 CA"\ '\ ~ ~

0.2

0

17\

~ 2 _\ -~

~ ..... I I I i V I I I I I

I I I I , , , I I I I' I I I I I I I I I I

o o. 1 o. 2 O. 3 0.4 O. 5 O. 6 O. 7 O. 8 O. 9 1. 0 xl £ 11018171

b. local Axial.Force Coefficient Distribution

Fig. 22 local Normal.Force and Axial.Force Coefficient Distributions on Model HB.2 at an Angle of Attack of 10 deg for Mach Numbers 1.5, 3, 5, and 8

47