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AEDC-TR-79-45 C?_

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O~P J 0

Boundary-Layer Transition on Large-Scale CMT Graphite Nosetips

at Reentry Conditions

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R. M. Raper ARO, Inc.

January 1981

Final Report for Period January 26, 1978 - August 10, 1978

I Approved for public release; distribution unlimited, l

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P J ' ( : p z , £ . . . . . ." , .~ ~.,.~

ARNOLD ENGINEERING DEVELOPMENT CENTER ARNOLD AIR FORCE STATION, TENNESSEE

AIR FORCE SYSTEMS COMMAND UNITED STATES AIR FORCE

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NOTICES

When U. S. Government drawings, specifications, or other data are used for any purpose other than a def'mitely related Govemment procurement operation, the Government thereby incurs no responsibility nor any obl/gation 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 be regarded by impl/cation or otherwise, or 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.

Qualified users may obtain copies of this report from the Defense Technical Information Center.

References to named commercial products in this report are not to be considered in any sense as an indorsement of the product by the United States Air Force or the Government.

This report has been reviewed by the Office of Public Affairs (PA) and is releasable to the National Technical Information Service (NTIS). At NTIS, it will be available to the general public, including foreign nations.

APPROVAL STATEMENT

This report has been reviewed and approved.

JOSEPH F. PAWLICK, JR., Lt Colonel, USAF Test Director, VKF Division Directorate of Test Operations

Approved for publication:

FOR THE COMMANDER

Technical Director Deputy for Operations

UNCLASSIFIED "

REPORTDOCUMEHTATIOH PAGE I REPORT NUMBER ~2 GOVT ACCESSION NO,

t AEDC-TR-79-45 4 T I T L E ( ~ d Subt i t le)

BOUNDARY-LAYER TRANSITION ON LARGE-SCALE CMT GRAPHITE NOSETIPS AT REENTRY CONDITIONS

? AU THORrs)

R. M. R a p e r , ARO, I n c . , a S v e r d r u p C o r p o r a t i o n Company

9 PERFORMING ORGANIZATION NAME AND ADDRESS

A r n o l d E n g i n e e r i n g D e v e l o p m e n t Cente r /DOOV A i r F o r c e S y s t e m s Command A r n o l d A i r F o r c e S t a t i o n , T e n n e s s e e 37389

I I, CONTROLLING OFFICE NAME AND ADDRESS

Naval Surface Weapons Center/WOL N a v a l O r d n a n c e L a b o r a t o r y S i l v e r S p r i n g . M a r y l a n d 20910

14 MONITORING AGENCY NAME & ADDRESS(If dil/erenJ IrGm ConJtoIWin80fhce.}

R E A D I N S T R U C T I O N S B E F O R E C O M P L E T I N G FORM

3 RECIPlENY'S CATALOG NUMBER

S TYPE OF REPORT & PERIOD COVERED

F i n a l R e p o r t - J a n u a r y 26 - A u g u s t 10 , 1978 6 PERFORMING ORG. REPORT N.JMBER

S. CONTRACT OR GRANT NUMBER,is)

10 PROGRAM ELEMENT. PROJECT, TASK AREA & WORK UNIT NUMBERS

P r o g r a m E l e m e n t 921B03

12 REPORT DATE

J a n u a r y 1981 IS NUMBER OF PAGES

42 1S SECURITY CLASS (o ! rhea report)

UNCLASSIFIED

ISe DECL ASSI FICATION ' DOWNGRADING SCHEDULE

'N/A 16 DISTRIBUTION STATEMENT (o f th le R e p ~ l )

A p p r o v e d f o r p u b l i c r e l e a s e ; d i s t r i b u t i o n u n l i m i t e d .

17 DISTRIBUTION STATEMENT (of the ebet roc l en lered in BJock 20, I ! d l f le ren t from Report)

18 SUPPLEMENTARY NOTES

A v a i l a b l e i n D e f e n s e T e c h n i c a l I n f o r m a t i o n C e n t e r (DTIC)

19 KEY WORDS ~ConJmue ~ reverse aide i f neceano~ ~ d I d e n t i ~ ~ b iock n ~ b e ~

b o u n d a r y l a y e r t r a n s i t i o n n o s e t i p s l a m i n a r f l o w g r a p h i t e t u r b u l e n t f l o w f r e e s t r e a m h y p e r v e l o c i t y r a d i a t i o n p y r o m e t e r s t r a c k s s t a g n a t i o n p o i n t

20 ABST RACT (Cont inue ~ reverae aide I f n e c e e a o ~ and I d e n l l ~ ~ b lock numbe~

A s e r i e s o f t e s t s d e s i g n e d t o i n v e s t i g a t e t h e b o u n d a r y - l a y e r t r a n s i t i o n c h a r a c t e r i s t i c s o f p r e a b l a t e d CMT g r a p h i t e n o s e t i p s was c o n d u c t e d i n a h y p e r v e l o c i t y t r a c k f a c i l i t y . The l o c a t i o n o f t h e z o n e a t w h i c h t r a n s i t i o n f r o m l a m i n a r t o t u r b u l e n t f l o w o c c u r r e d was m e a s u r e d on a number o f n o s e t i p s h a v i n g r a d i i o f 0 . 4 0 , 0 . 8 0 , a n d 1 . 1 2 5 i n . The n o s e t i p s w e r e t e s t e d a t , n o m i n a l l y 16 k f p s o v e r a f r e e - s t r e a m p r e s s u r e r a n g e o f 0 . 1 3 t o 0 . 4 0 a t m . T r a n s i t i o n f r o n t

F O R " 1473 EDITION OF I NOV 65 IS OBSOLETE D D , JAN 73 J

UNCLASSIFIED

UNCLASSWIED

2 0 . ABSTRACT ( C o n t i n u e d )

l o c a t i o n s w e r e i n f e r r e d f r o m n o s e t t p s u r f a c e t e m p e r a t u r e d i s t r i b u - t i o n s o b t a i n e d u s i n g a p h o t o g r a p h i c p y r o m e t r y t e c h n i q u e , and f r o n t s w e r e o b s e r v e d t o r a n g e f r o m t h e p r o x i m i t y o f t h e s t a g n a t i o n p o i n t t o t h e s o n i c p o i n t . I t was o b s e r v e d t h a t t h e mean t r a n s i - t i o n l o c a t i o n moved f o r w a r d on t h e n o s e t i p a s t e s t p r e s s u r e i n c r e a s e d . The c i r c u m f e r e n t i a l v a r i a t i o n a b o u t t h e mean t r a n s i t i o [ ] l o c a t i o n o v e r a g i v e n n o s e t i p d e c r e a s e d a s t h e mean l o c a t i o n moved n e a r e r t o t h e s t a g n a t i o n , p o i n t .

A F S C A~o td AFS T~lm

UNCLASSIFIED

A EDC-T R-79-45

PREFACE

The work reported herein was conducted by the Arnold Engineering Development Center (AEDC), Air Force Systems Command (AFSC) at the request of the Naval Surface Weapons Center (NSWC/WOL). The NSWC project manager was Dr. Daniel C. Reda. The

results of the test were obtained by ARO, Inc., AEDC Group (a Sverdrup Corporation Company), operating contractor for AEDC, AFSC, Arnold Air Force Station, Tennessee, under ARO Project Number V41G-61. The data analysis was completed on October 16, 1978, and the manuscript was submitted for publication on May 7, 1979.

A E D C-T R - 7 9 - 4 5

CONTENTS

Page

1.0 I N T R O D U C T I O N ........................................................ 5

2.0 A P P A R A T U S

2.1 Test Unit ............................................................. 6

2.2 Instrumentation ....................................................... 7

2.3 Model Description ..................................................... 9

3.0 TEST P R O C E D U R E S

3.1 Test Technique ....................................................... 10

3.2 Data Reduction ....................................................... 11 3.3 Data Uncertainty ...................................................... 11 3.4 Aerothermal Environment .............................................. 12

4.0 E X P E R I M E N T A L RESULTS

4.1 In-Flight Laser-Illuminated Photographs ................................. 13

4.2 Image Intensifier Camera Photographs ................................... 14

4.3 Thermal Contours ..................................................... 15

5.0 SUMMARY OF RESULTS ................................................ 16

REFERENCES ........................................................... 17

ILLUSTRATIONS

Figure

1. AEDC 1,000-ft Hypervelocity Range/Track (G) ............................... 19

2. Track G Photopyrometer System ............................................ 20 3. Track G Front-Light/Back-Light Laser Photography System . . . . . . . . . . . . . . . . . . . . 21

4. NosetipTransi t ion Models ................................................. 22

5. Surface Temperature Data Reduction Flow Chart . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

6. Photopyrometer Temperature Measurement Uncertainty . . . . . . . . . . . . . . . . . . . . . . . 26

7. Velocity Decay for Large Nosetip Model ...................................... 27

8. In-flight Photographs of Nosetips ........................................... 28

9. Typical image-Converter Camera Photographs ................................ 29

10. Example of Temperature 13ata in Spherical Coordinates . . . . . . . . . . . . . . . . . . . . . . . . 31

11. Tabulated Temperature Data in Spherical Coordinates . . . . . . . . . . . . . . . . . . . . . . . . . 34

12. Example of Temperature Distribution and Definition of Transition Location . . . . . . 35

13. Planar Projection of Transition Front Location ............................... 36

14. Variation of Mean Transition Location with Range Pressure . . . . . . . . . . . . . . . . . . . . . 37

15. Measurements of Transition Zone Asymmetry ................................. 38

D

A E D C - T R - 7 9 - 4 5

TABLES

Page

!. Nominal Measurement Ranges for Photopyrometers . . . . . . . . . . . . . . . . . . . . . . . . . . . 39

2. Test Summary ............................................................ 40

3. Model Angle Data ........................................................ 41

N O M E N C L A T U R E ..................................................... 42

A E D C-T R-79-45

!.0 INTRODUCTION

The impetus for studying nosetip boundary-layer transition is. provided by ablation rates

and shape change histories for currently employed reentry vehicle (RV) nosetip materials (e.g., graphite and other carbonaceous materials) dependent pri.marily on a complex interaction between intrinsic surface roughness and the boundary-layer flow. At high altitudes the boundary layer over the full-scale nosetip is laminar, and ablation acts to

increase bluntness. Also, preferential ablation" of binder elements in the surface material causes the surface to de,,.eiop microroughness related to the grain size or, in the case of carbon/carbon type materials, the preform geometry. It is this microroughness that eventually causes boundary-layer transition to occur and at ahigher altitude than at which it would have occurred for a very smooth surface. Boundary-layer transition results in much

more severe heat-transfer rates over the region between the stagnation point and the sonic point than occur for laminar flow. The increased sidewall heat flux in the regions behind the

stagnation point results in the formation of the concave nosetip shape associated with transitional and/or turbulent flows. At still lower altitudes the concave shape gives way to a sharp biconic as transition moves toward the stagnation point. Since the nosetip changes shape as the transition point moves forward, any circumferential variation in transition location can result in the formation of an asymmetrical geometry and trajectory dispersion

may occur.

Experimental information required for the correct modeling of such complex phenomena is often generated in wind tunnel or arc jet environments that are incapable of

achieving complete simultaneous simulation of reentry conditions in terms of Mach number, stagnation pressure, and total enthalpy. Semiempirical correlations based on such data must therefore be verified through actual flight tests. Because of the cost of full-scale testing, the

inherent lack of control over the experiment, and the difficulties involved in data acquisition and evaluation, techniques that meet the verification criteria for doing subscale testing in the ballistics range have been developed. The advantage of the ballistics, range in this regard is

that it has the capability to duplicate actual reentry conditions in terms of simultaneous duplication of Mach number, pressure, and enthaipy for any point on a reentry trajectory. The major potential disadvantages presented by the ballistics range are limited test time and

limited model scale for tests of the type described in this report.

Nosetip boundary-layer transition tests in the free-flight Range G began in 1972 in a

program involving AXF-SQ graphite; a companion program for ATJ-S graphite was conducted in 1973 (Ref. I). These tests were handicapped by shock cap radiation that

obscured the low temperature laminar region, and means were subsequently devised to provide a short length of helium environment along the range centerline in which more

A E D C - T R - 7 9 - 4 5

meaningful measurements could be obtained. Extensive tests of CMT graphite using this

technique were conducted with both smooth and preablated nosetips. The preablated

specimens were produced by exposing the test articles in the AEDC 5-MW Arc Jet for

several seconds to establish a characteristic surface roughness before launching the test

articles in the range. The very different results (not available in the open literature) that were

obtained for the two types of nosetips indicated that, at least for CMT graphite, apparently

very different results were usually produced for nonpreablated surfaces. The initially

smooth-surface response was generally characterized by hot spots and a transition front that

was grossly irregular and did not necessarily extend completely around the nosetip. The

preablated nosetips produced the more symmetrical patterns one would expect.

The objective of the present test program was to obtain more and higher-quality

transition data for CMT nosetips with radii approaching those of full-scale RV's. Because

of the previous results for this material, only preablated test specimens were tested. In

addition, the recent availability of track guidance for test models has relieved the

aerodynamic stability constraint on model design and permits much larger nosetips to be tested. Other advantages offered by model guidance are that (l) the test specimens may be

recovered for posttest evaluation of surface roughness and (2) the elimination of trajectory

dispersion makes it possible to obtain higher resolution photographic pyrometry data. These

advances in test technique have greatly enhanced the capabilities of the range in nosetip

transition testing.

It is the purpose of this report to present new and improved-quality transition data that have been obtained for near full-scale CMT graphite nosetips as described above. These

results are interpreted in terms of mean transition location. The asymmetric nature of the

transition process is described statistically.

2.0 APPARATUS

2.1 TEST UNIT

The present test program was conducted in the AEDC Hypervelocity Range/Track (G)

facility. This test unit is basically the free-flight Range G (Ref. 2) that has been equipped

with a removable track model guidance and recovery system (see Fig. l). Hence, the facility

can be operated in either a free-flight mode or track mode. The primary components of the

facility in the track mode are (I) a model launcher that consists of a 2.5-in.-caliber, two-

stage light-gas (H2) gun approximately 150 fl long, (2) a model guidance system that consists

of ninety-one 10-ft-long track sections, (3) a model recovery device that is a 500-ft-long tube

in which the model is aerodynamically decelerated, (4) the 1,000-ft range tank that encloses

6

A E D C - T R - 7 9 - 4 §

the track and provides the desired test environment, and (5) an instrumentation system capable of in-flight data acquisition. Bulkheads are located in the range tank at 85 and 390 ft

from the uprange end; openings in the bulkheads through which the track passes are equipped with quick-opening valves. These allow staged test pressures along the flight path

if desired. Other quick-opening valves are installed along the the recovery tube to maintain its helium and nitrogen gases at the required pressure before model arrival. A more complete description of the track system and its operation is given in Ref. 3.

2.2 INSTRUMENTATION

2.2.1 Photographic Pyrometry

In-flight nosetip surface temperature measurements, the primary data in this test program, were obtained by means of photographic pyrometry. This temperature

measurement technique involves basically the following sequence: (l) a high-speed, image- converter camera system is used to obtain a siop-motion, self-luminosity photograph of the test model in flight, (2) calibration data from a carbon arc (or tungsten) reference source are recorded on identical film and are processed simultaneously with the model photograph,

and (3) densities on the film image of the model surface are measured and converted to

temperatures using the calibration data.

Temperatures measured by this method are brightness temperatures if the radiation observed is strictly incandescent radiation. In general, true surface temperatures may be calculated if the emissivity of the test material is known, in most instances, test materials of current interest are carbon and graphite compounds with emissivities very near unity. In these cases, temperatures measured are surface temperatures accurate to within the

measurement uncertainty of the systems.

The manner in which photographic pyrometry systems are employed in the track mode of testing for in-flight model surface temperature measurements is shown schematically in Fig. 2. The near head-on view angle is highly advantageous for transition testing because it

offers the best possible coverage of the subsonic portion of the flow field. Two types of proximity-focused image-converter cameras are used. One employs fiber-optic coupling

from phosphor anode to the recording film (Generation I, or Gen-l, image intensifiers), and

the other type features a microchannel plate amplifier between photocathode and phosphor

anode (Generation II, or Gen-II, image intensifiers). Both types of intensifiers have S-20R spectral, response: They respond to visible and near-infrared radiation to a wavelength of alJproximately 0.93 p.m. In terms of applications, the most important difference between the two types of intensifiers is that the Gen-ll type is more sensitive than the Gen-I type (see

A EDC-T R-79-45

Table 1). For instance, for an exposure duration'Of 1,000 nsec, the Gen-li system threshold

is 1,250°K, whereas for the Gen-I system it .is 1,600°K. However, the dynamic range of the

Gen-I is smaller than that of the Gen-ll. For the present test program, the wide range of nosetip sizes and test conditions that were used resulted in a wide range of temperature levels

at a given station location. For this reason, the five range locations available for pyrometry

instrumentation (Stations 4, 11, 20, 29, and 41, which are 100, 238, 398, 596, and 836 ft

from range entrance, respectively) were equipped with different systems consistent with the

expected noselip temperature range. Hov,.ever, for shots utilizing the largest nosetips, the temperature levels generally did not reach the threshold of even the Gen-II systems until the

model had reached station 20. Consequently, for most of those shots, only three Gen-ll systems could be operated.

For ahnost all shots in the present test program, the brightness temperature of the air

shock cap exceeded the nosetip surface temperature over the full flight distance. When this

condition exists, use of the measurement technique herein described requires elimination of

shock cap radiation. This problem was solved in previous free-flight tests by providing a

small chamber filled with helimn at each measurement station. Each chamber was opened

immediately preceding model arrival so that measurements could be obtained in the helium

environment. The advantage of a helium environment is that its radiation intensity is much

lower than that for air for the velocity and pressure conditions desired for the presenl test.

The shock cap quench technique developed for track testing uses a plenum that is suspended

from the track tube and that vents a low-velocity column of helium vertically through the

slots in the track tube. This system provides a helium environment of 95 + percent purity

over a flight distance of 10 in. The model is photographed near the end of the helium-rich

region. Such systems were used at stations 11 and 29 for all the present tests. For stations 4,

20, and 41, the track tube slots were lengthened so that four of the I-ft plenums could be

utilized end to end. The purpose of this was to provide greater flow-field quench distances to

minimize extraneous model and wake luminosity, a condition that was required for another

test program. The I-ft-sweep distance was sufficient to quench noselip flow fields. Evidence

of this is provided by stagnation point temperatures down to the threshold of the pyrometry

systems being measured at some test conditions in which the air shock cap temperature with no quench would have been much higher.

2.2.2 l.aser Photography

Laser-lighted photography was used as part of the standard instrumentation to obtain

high-resolution in-flight photographs of test articles. The intensity of the light source

permits extremely brief exposure times {20 nsec) and effectively eliminates motion blur. For

ablation and erosion tests, laser photography provides the means to obtain in-flight surface

8

A E D C - T R - 7 9 - 4 5

recession data. For the present test program, the role of laser photography was to evaluate

the in-flight quality of the model nose test surface.

A typical laser photographic system for track applications is shown schematically in Fig.

3. This optical arrangement provides a combination of diffuse front' and back lighting

through appropriate slots in the track tube. This system produces a magnification of

approximately 0.7. Peak photographic resolvability has been shown to be approximately 25

p.m. In contrast, the resolution of the laser photography systems configured for the free-

flight mode of testing is 200 #m.

2.3 MODEL DESCRIPTION

The three model configurations shown in Fig. 4 were required to accommodate the

different sizes of nosetips tested. Only one shot was made with the 0.40-in. nos~tip (Fig. 4a),

and most of the present shots used the model configuration shown in Fig. 4c. This configuration features a hemispherical nosetip (rn = I. 125 in.) nearly as large in diameter as

the launch tube, therefore representing the maximum model size. Such a model was not

suitable for free-flight testing because the aerodynamic stability requirement for free-flight

models could not be satisfied within the launcher-imposed weight limit.

The method of model recovery in the track guidance system imposes a model design

constraint: The test article must retain sufficient structural integrity at the end of its flight to

withstand deceleration loading. Initially, most of the track models used in materials testing

were similar to that shown in Fig. 4a. With this model design, the test material configuration

is a reverse-taper plug that is swaged in place in the model forebody. Larger nosetips, which were required for expanding the range of diameters, could not survive launch loading in a

plug configuration because of the low tensile strength of graphitic materials. Thus, large

nosetips (r,1 = 1 in.) were constrained to follow a thin shell design. The difficulty posed by

this configuration is that the only practical means by which a shell may be attached to a

substrate material is a chemical bond (epoxy), which can be expected to weaken at about

200 ° F. This problem was addressed by parametric calculations of the back-face temperature

of the shell for a range of thicknesses and exposure times. A wall thickness of 200 mils was

sufficient to contain the back-face temperature below 200°F for a 60-msec test duration and

a 30-msec recovery cycle duration. The thermal response of the surface for this thickness was

essentially the same as that for a semi-infinite wall thickness. Small carbon-carbon pins were

included in the model design to hold the cap in place after model motion had been arrested and after the back-face temperature had been driven by conduction of heat from the hot

surface region to beyond the bondline limit. Another purpose of the pins was to provide hot

spots at known surface locations for data reduction purposes. This is discussed further in

Section 4.2.

.

A ED C-TR-79-45

The ballistic characteristics of the three model configurations (nose radius, rn) utilized in this test program are summarized below:

rn, in. l;eight, gm CD* 6 *, ib/ft 2

0.40 510 0.43 76

0.80 532 0.60 57

1.125 619 0.84 48

*Includes track friction drag (CD obtained using velocity measuremenls from X-ray stations)

The CMT graphite nosetip test specimens utilized in the present test series were

fabricated at AEDC from billet Nos. 104C-04, 103J-15, 104K-08, and 104K-15 supplied by

the Navy.

3.0 TEST PROCEDURE

3.1 TEST TECHNIQUE

The basic test technique used in the present program was to launch the model on the

track at a specified range pressure and velocity and to obtain brightness photographs of the nosetip at selected locations along its flight path. As the nosetip moved down the range, its

surface temperature--and brightness--increased in accordance with the distribution of heat

input over its surface. The temperature distribution data were obtained directly from the

brightness photographs. Transient temperature measurements plus calorimetric properties

of the test material provide sufficient data for inferring the actual convective heat-flux

distribution.

All the nosetips tested during this program were preablated in the I-mw APG Vacuum

Facility of the Acurex Corporation. These runs were made at a centerline enthalpy of about

43,000 btu/Ib and a reservoir pressure of 0.06 arm. The exposure durations and resulting

surface recession levels varied with model geometry as follows:

Exposure r n - Initial, r n - After Arc Time, Stagnation Point

in. Run~ in. sec Recessionp in.

0.40 0.43 20 0.035

0.80 0.90 61 0.063

1.125 1.25 72 0.062

l0

A E D C - T R - 7 9 - 4 5

Considerable care was taken to preserve the surface roughness features developed in the arc

jet. To assure that each nosetip surface was undamaged, each was inspected and photographed before being loaded in the launcher. The nosetips that were recovered intact were photographed and then shipped to the NSWC as directed.

3.2 DATA REDUCTION

A self-luminosity photograph was obtained for each model as it passed through the field-

of-view of each image-converter camera. This photographic negative and a calibration negative--produced by photographing with the same camera using either a carbon-arc light source or a tungsten source, depending on the desired calibration range--were simultaneously developed. These negative pairs were then processed on a high-speed,

digital-scanning microdensitometer that stored on magnetic tape the film density

distribution for the nosetip image and the density-versus-temperature data from the calibration images. These data were then computer processed to produce a thermal contour map of the nosetip. Next, a computer drawing of the noset!p shape as it would be viewed by an image-intensifier camera was superimposed on the contour map so that the hot spots caused by the marker pins coincided with their known locations. Aligning these known surface points established the geometric center of the nosetip. The contour map

representation is in film plane coordinates; however, a procedure has been recently implemented to transform the film plane data to spherical coordinates. Thus, the final temperature data consisted of a computer-generated tabulation of the temperature

distribution along each of 120 rays spaced at 3-deg intervals around the nosetip for each

photographic pyrometry station. A flow chart of this data reduction procedure is shown in Fig. 5.

3.3 DATA UNCERTAINTY

The uncertainty model employed to evaluate the experimental results subsequently presented consisted of two parts: bias error and precision error. Their sum defines the total measurement uncertainty. The following are estimates for measurements of free-stream pressure, temperature, and velocity:

Est imated Uncer ta in ty~ 2o

P r e c i s i o n T o t a l Bias , E r ro r U n c e r t a i n t y ,

Parameter percent percent percent

Free-stream p r e s s u r e , p® 0.7 0.3 1.0

Free-stream t e m p e r a t u r e , T® 0.1 0 .2 0.3

Free-sCream velocity, V Negligible 0.3 0.5

Stagnation enthalpy," H ° Negligible 1.0 1.0

Stagnation pressure, p; 0.8 1.2 2.0

l l

A E D C - T R -79 -45

The uncertainty involved in measuring the surface temperature at a specific location on a

nosetip has also been assessed. This uncertainty arises primarily from (l) relating points

viewed in the film plane to actual spherical nosetip surface locations and (2) measuring Ihe

temperature at a given point in the fihn plane. The model image in Ihe film plane contains

tour known surface locations; these are the previously described hot spots caused by marker pins. The location uncertainty occurs in aligning these marker spots with their known

locations. Uncertainty in this alignment is estimated to result in a stagnation point location

uncertainty of about 0.04 in. A location bias error of up to 1.7 deg in the flow direction

could be incurred in transforming the fihn plane data to spherical coordinates.

Measurements of surface temperature at a given 'film plane location are derived using a

densitometer with a 100-/~m aperture. Near the stagnation point, where the nosetip surface is

most nearly parallel to the fihn plane, this aperture size corresponds to a streamwise angle

increment of 0.4 deg; the temperature is an average value over this region. The magnitude of

the resulting precision error is dependent on the temperature gradient.

The precision with which the photographic pyrometry systems can measure a given

uniform temperature was determined from static measurements of a standard carbon-arc

light source. The system precision error associated with temperature measurements, shown in Fig. 6, varied with temperature level. The main sources of uncertainty are gradients in

gain across the face of the image intensifier and microdensitometer imprecision (_+0.01 D).

As shown in Fig. 6, uncertainty increases toward the high-temperature end of each dynamic

range for each photopyrometer. This is a consequence of the particular relationship of the

reciprocal of temperature with fihn density (approximately linear). For a given uncertainty

in film density, the corresponding uncertainty in temperature is considerably greater for the

higher temperatures (lower densities).

A test-related source of error not previously discussed is an image-intensifier

characteristic called flare. Light reflections inside the intensifier create ,,.ery bright image

• areas that contribute noise illumination to adjacent areas of the displayed image; i.e., a

bright spot will have a halo around it. Troublesome flare effects can produce significant

uncertainty in surface temperature measurements. These flare problems arise from test-

related sources--excessively hot surface areas on the nosetip, heat shield luminosity, and

wake radiation, in flare-dominated photographs, it is not possible to extract certain low

temperature dath, so it is very important to minimize flare sources. Figure 6 characterizes the

overall uncertainty associated with surface temperature measurements made in the absence

of significant flare.

12

A'E D C -T R -7 9 -4 5

3.4 AEROTHERMAL ENVIRONMENT

The aerothermal environment for a ballistics range test article is specified by the test gas

composition and its static temperalure, its static pressure, and the model velocity. For all but one of the shots reported herein, the test gas was air (nominally at 76°F). That one shot was made with nitrogen as the test gas at the same nominal static temperature as the air shots.

The nominal launch velocity for all shots was 16 kfps, and the actual velocity ranged from

15.8 to 16.2 kfps (expect as noted in Table 2). The corresponding nominal stagnation I

enthalpy was 5,100 Btu/Ib, and the stagnation pressure ranged from 33.atm for p** = 100

torr to 100 arm for p** = 300 torr. The velocity of the test model at any flight distance may be calculated from the ballistics characteristics of the configuration and the test conditions

using the following equation:

V,,,= V**. e( -4'87 x 10-5 p~x/fl) 1

with p® in mm-Hg, x in ft, and 8 in lb/ft 2. With this equation, the velocity decays for three

test pressures for the large nosetip model were calculated and are shown in Fig. 7.

4.0 EXPERIMENTAL RESULTS"

Twenty-seven shots were fired during the present test program. The test summary, shown in Table 2, indicates that "18 shots produced data of sufficiently high quality to be considered for analysis. The experimental results that are discussed in this section for these shots are in-flight laser photographs, image-intensifier camera photographs, thermal

contours, and mean surface temperature distributions.

4.1 IN-FLIGHT LASER-ILLUMINATED PHOTOGRAPHS

During this test program the in-flight laser photography capability described earlier was utilized to determine whether the test specimens were damaged during launch. Because of the sensitivity of boundary-layer transition to surface imperfections, the preablated surface

characteristics had to be preserved except for ablation effects. The small polyethylene particles, present in the track tube after each shot, that were swept up by the launcher muzzle blast and consequently impacted on the nosetip were the main cause of damage to the test surface. The muzzle blast preceded these models for about 30 ft following launcher exit for most test conditions; therefore, it was only over this 30-ft interval that damage

occurred. Though the entire track housing was routinely cleaned before each shot, special emphasis was given to the uprange end once the cause of the damage was discovered.

Examples of surface damage are shown in Fig. 8. For two of the'shots (5,014 and 5,043) the

13

A E D C - T R - 7 9 - 4 5

damage was so extensive (Fig. 8a) that the data were judged to be useless. For several other

shots (5,038, 5,044, and 5,063) the damage was minimal (Fig. 8b), and the data were judged to be acceptable although data for surface rays containing craters are questionable. Most of the shots were undamaged; their surfaces resembled the example in Fig. 8c. The small size of

the damaging particles and the fact that they were generated during every launch by the destruction of the pump tube piston combined to make it impossible to establish an

absolutely debris-free condition in the track tube.

4.2 IMAGE-INTENSIFIER CAMERA PHOTOGRAPHS

Some information concerning the presence and location of transition fronts for CMT

graphite nosetips is apparent in the image-converter camera photographs, examples of which are shown in Fig. 9. The most significant feature of these photographs is the bright ring surrounding the darker central region of each nosetip. The relatively dark (cool) center of each image corresponds to a region of laminar flow, whereas the relatively bright (hot)

region encircling it defines a zone of transitional/turbulent flow. This is in accordance with

the well-known laminar/turbulent heat-transfer distribution over a hemisphere. Only when a boundary layer is entirely laminar and, hence, there is no transition zone on the nosetip,

does the ring not occur. The point at which the brightness distributions change from dark to light in each of a series of photographs may be somewhat misleading. The response characteristics of the individual tubes and the modes in which they are,operated give rise to some of the apparent differences. However, for a given photograph, the circumferential variation in the "brightness onset" location is sign!ficant since it occurs at approximately the same temperature around a nosetip.

For the single 0.40-in. nosetip that was tested (Shot No. 5,000), the transition zone was near the stagnation point initially, but then it gradually disappeared with time. For the

other, larger nosetips, such dramatic changes in transition zone location were not observed. At a pressure level of 100 torr (Shot No. 5,021) a transition zone was only barely established

near the sonic point. All other shots demonstrated a well-established transition zone at the first measurement station, and its location tended to move toward the stagnation point as the test pressure was increased.

To derive spatially resolved temperature distributions from these photographs, a provision was made for producing bright spots at known surface locations directly on each nosetip photograph. This was accomplished by having the four nosetip anchor pins--spaced

90 deg apart around the nosetip and located 60 deg off its centerline in the flow direction-- protrude 0.050 in. above the surface. Extending outside the boundary-layer thickness, the

pins were subjected to severe heating and were consequently much hotter than the surrounding surface. These pins were clearly visible in all the image-intensifier camera

14

A E D C - T R - 7 9 - 4 5

photographs* and provided the reference points necessary to determine the geometric center

of the nosetip image.

It is evident in some of the photographs, particularly ihose from Shot Nos. 5,038 and 5,044, that the transition zone was not very symmetrical relative to the stagnation point.

This could possibly be a consequence of angular motion of the model. It is unfortunate that, because the model was surrounded by the track tube, the 53 orthogonal shadowgraph

systems used in free-flight tests to record model motion were not usable. However, during these tests the model was photographed through the slots in the track tube at ten range

locations using X-ray systems. Seven of these systems--positioned laterally--viewed the model from a horizontal direction, thus recording angles in the pitch plane. The other

three--positioned superiorlymviewed the model from above, thereby observing yaw plane motion. Though orthogonal views would have been desirable, the data shown in Table 3 are sufficient to indicate that the model angle of attack was usually less than I deg. However, this level was exceeded at some time during every shot. The nature of this motion is presumed to be random wander of the model back and forth between the rails. The mean of all the angles measured at different locations indicates that angular displacement may have

peaked somewhere near midrange.

4.3 THERMAL CONTOURS

Following application of the microdensitometer processing technique earlier described, the photographic images presented in Fig. 9 were computer-plotted as thermal contour maps. Several examples of these maps are shown in Fig. 10. Since they are drawn in film plane coordinates, they are a quantitative representation of the photographs. The contour interval is 100°K; a heavy contour line denotes 500°K intervals. One purpose of the contour maps is to so quantify the temperature gradients that were apparent in the photographs that test results can be rapidly surveyed. In addition, the approximate location of the transition front is made readily apparent by the contour line spacing; i.e., the region of the closest spacing of the contour lines corresponds to the zone over which heating rate increases from

laminar to turbulent levels.

A refinement of the contour map presentation produced for each nosetip photograph was obtained by transforming the digital temperature data into spherical, surface coordinates. An example of the resulting data is shown in Fig. ! I for station 20 of Shot No. 5,045. Here the circumferential location (around the nosetip) is given in the first column, and listed in the first row is the angular location in the flow direction (3 to 75 deg). The

D

*Pins were installed on the nosetip for Shot No. 5,000 because of the small size.

15

A E D C - T R - 7 9 - 4 5

angular increment in either direction is 3 deg. Note that a single row of measurements

defines the temperature distribution along one ray. Finally, the mean temperature values

along with the standard deviation of the individual circumferential measurements (at each

location in the flow direction) are tabulated in the last two lines. These data were compiled

on magnetic tape in this same format and have been examined ray by ray using the

interactive graphics techniques described in Ref. 4. A transition location as defined in Fig.

12 for a typical temperature distribution ',,,.as tabulated for each ray of each noselip image.

The objective of this work, described by Reda and Raper in Ref. 5, was Io quantify the

variations in the transition front locations that are apparent from the contour maps. The

data from each photograph yielded 120 transition location measurements, example plots of

which are shown in Fig. 13, and from these measurements a mean transition location,

S/rnrR.,~ean was calculated. As shown in Fig. 14 (also Fig. 14 in Ref. 5), the mean transition

location progressed toward the stagnation point as the range pressure was increased. A

parameter used to further characterize the transition-front variations was the standard

deviation (olr) Of the 120 circumferential measurements about the mean transition location.

An asymmetric transition front location (characterized by at,) and the associated asymmetric

heating distribution can result, in flight, in a longer term, ablation-induced nosetip

asymmetry with potentially serious implications for vehicle accuracy. As shown in Fig. 15

(also Fig. 15 of Ref. 5), the mean of the o,, measurements decreased as the mean transition

location moved forward. The data shown in Figs. 14 and 15 were derived only from

photographs fo.r which the transition front had not yet begun to shift rearward (during a

flight). Thus, more emphasis was placed on uprange data than on that obtained later in a

flight.

5.0 SUMMARY OF RESUI.TS

A series of tests to in,..estigale the boundary-layer transition characteristics of preablated

CMT graphite nosetips was conducted in the AEDC Hypervelocity Range/Track (G). The

nominal test velocity was 16 kfps, and free-stream pressures ranged from 0.'!3 to 0.40 arm.

Three different sizes of nosetips were tested, r, = 0.40, 0.80, and 1.125 in. The results

indicate the following:

1. Measured trans,tion-front locations on the hemispherical nosetips varied from

the proximity of the stagnation point to the sonic point; transition moved

forward on a nosetip as the test pressure was increased.

2. The magnitude of deviations in the transition-front location data--as defined

by Otr measurements--decreased as transition moved forward on a nosetip.

16

• A E D C - T R - 7 9 - 4 5

3. Nosetips with radii approaching those of full-scale RV's can be successfully tested in the ballistics range at velocities and pressures duplicating reentry conditions, and they can be recovered intact for posttest examination.

4. Test articles can be preablated to improve the reentry simulation without significantly degrading their strength.

5. The addition of protuberances near the aft end of the test surface provided a valuable geometric reference that enhanced the data quality because the stagnation point location uncertainty ~,'as reduced.

REFERENCES

1. Miller, John T. "Nosetip Transition Measurements in a Reentry Test Environment." AEDC-TR-76-118 (ADB013270L), August 1976.

2. Test Facilities Handbook (Tenth Edition). "Von Kdrman Gas Dynamics Facility, Vol. 3." Arnold Engineering Development Center, Arnold Air Force Station, Tennessee, July 1971.

3. Norfleet, G. D., Hendrix, R. E., and Jackson, D. "Development of a Hypervelocity Track Facility at AEDC." American Institute of Aeronautics and Astronautics, Aerospace Sciences Meeting, 15th, Los Angeles, Calif., January 1977, AIAA Paper No.

77-151.

4. Reda, Daniel C. and Brown, Hensel, S. "Analysis of Nosetip Boundary-Layer Transition Data Utilizing Interactive Graphics." Instrumentation in the Aerospace Industrym Vol. 24~Advances in Test MeasurementmVol. 15, Part Two, Proceedings o f the 24th International Instrumentation Symposium--Albuquerque, N.M. Publications Department, Instrument Society of America, Pittsburgh, Penn., 1978, pp. 515-524.

5. Reda, Daniel C. and Raper, Richard M. "Measurements of Transition Front Asymmetries on Large-Scale, Ablating Graphite Nosetips on Hypersonic Flight." American Institute of Aeronautics and Astronautics, 17th Aerospace Sciences Meeting, New Orleans, La., January 1979, AIAA Paper No. 79-0268.

17

Recovery Tube m

Figure 1. AEDC, 1,000-ft Hypervelocity RangeITrack (GI.

7 Imare Intensifier ~iber-optics Coupling

Notes :

1. Distance a l o n g Optical Axis from Camera Lens to Focal Plane: 3 4 . 5 in.

2 . Distance from Camera Camera Lens - Lens to Mirror: -12 i n . 8 - i n . Focal

3 . Magnification: 0 . 3 Length, f/4.5

Image- Intensifier Camera

%

Elliptical,

r *lane Front-Surface Mirror

D i r e c t i o n

I

\

Figure 2. Track G photopyrorneter system.

Pulsed Ruby

M

B,

Collimating

4x5 Camera Narrow-

Track TubelRa I-~ Band Assembly-7 \ ~- Fi Iter

t~

B, ~ - " E -'J

I---33 cm-~

M 1 High-Energy Reflector M 2 Front-Surface Mirror (;1, (;2 (;round-(;lass Diffuser (;3 Mirrored Ground (;lass B 1 Beamsplitter, 50~TISO~R B 2 Beamsplitter, 80~T/20~R Camera Lens: 135-mm f. I., f116

Light Source Stand Range Tank Wall

Figure 3. Track G front-light/back-light laser photography system.

)> m 0 o -n ¢0 j= 01

L . d

I 4 . 9 2 i n .

C e n t e r b o d y

v

m

i

CMT G r a p h i t e

b . )

0 . 4 0 i n .

2.50 in.

L B e r y l l i u m - C o p p e r N o s e t i p H o l d e r

Figure 4.

C a r b o n / E p o x y C l o t h Hea t S h i e l d

a. rn = 0.40 in. Nosetip transition models.

Lexan ® Base

i.mm

Marker/Anchor r Pins ( 4 ) ~ . ,

5 . 1 5 i n . d

2.50 in .

bO %0 0.80 i n . - - ~

CMT G r a p h i j - 0 . 2 0 in .

Aluminum Centerbody

/ Heat Shie ld

b. rn = 0.80 in. Figure 4. Continued.

Lexan ® Base m t~

& 01

L

F

0.20 in.

CMT G r a p h i t e

M a r k e r / A n c h o r P i n s (4

1.125 in.

4 . 4 2 i n .

--Alumlnum

C e n t e r b o d y

c. rn = 1.125 in. Figure 4. Concluded.

-I

2.50 in.

Lexan ® Base

m

v )

m

01

FJ

I m a g e - C o n v e r t e r

Camcra

Developed Nega t ives

(Data and P roces s ing Mic rodens i t ome te r C a l i b r a t i o n ) I

I Mag" Tape I

I

Dens i ty Contours

@

m l Computer

I .I

I C a l i b r a t i o n Curve

Temperature

I Temperature Contours

@

Tabula ted Data I in S p h e r i c a l Coord ina tes

Figure 5. Surface temperature data reduction flow chart.

m 0

q -n

Ol

t -J

+ 2 6 0 ,

±240

±220

±200 -

± 1 8 o

= + 1 6 0 _

: +140 -

± 1 2 0 -

o

±100 -

±80

± 6 0 -

±4O

±20 -

0 "

1 , 0 0 0

N o t e : Each p l o t t e d p o i n t r e p r e s e n t s t h e r e s u l t o f a s t a t i s t i c a l a n a l y s i s o f m u l t i p l e m e a s u r e m e n t s a m p l e s a t e a c h p a r t i c u l a r t e m p e r a t u r e ( s e e t e x t ) .

Gen I I 1 , 0 0 0 ( o r 300) n s e c

Gen I I 100 (o r 30 ) n s e c

Gen I 1 , 0 0 0 n s e c

!

1 , 4 0 0

Figure 6.

I I I I I I I I I I I

1 , 8 0 0 2 , 2 0 0 2 , 6 0 0 3 , 0 0 0 3 , 4 0 0 3 , 8 0 0

M e a s u r e d T e m p e r a t u r e , K

Photopyrometer temperature measurement uncertainW.

Gen I 100 n s e c

I I

4 , 2 0 0

3~ Ill c~ o

-n

tO

01

-..,]

16

14

q - I

12

10

. .

20O

- ~ - ~ 3 0 0 m

I I I , I I I I , I I 100 200 300 400 500 600 700 800 900

F l i g h t D i s t a n c e , f t

Figure 7. Velocity decay for large nosetip model.

m c ) c )

- n

Ol

t~J Oo

rn

c ) :4

~O

J= o i

a. Heavily damaged, Shot No. 5,015

b. Slightly damaged, Shot No. 5,038

c. Undamaged, Shot No. 5,045

Figure 8. In-flight photographs of nosetips.

STA 1 1 , 2 3 8 f t STA 2 0 , 3 9 8 f t STA 2 9 , 5 9 6 f t

N o t e : S T A = R a n g e S t a t i o n

a. Shot 5,014 Figure 9. Typical image-converter camera photographs.

STA 4 1 , 8 3 6 f t

m

¢3 :4 23

¢D

01

m O

o

A

L~

STA 2 0 , 3 9 8 f t

N o t e : STA = R a n g e S t a t i o n

STA 2 9 , 596 f t

b. Shot 5,027 Figure 9. Concluded.

STA 4 1 , 836 f t

" I//g~':~ff~'L-~,~ ~-~I

~'~~ ~ , ~ ~ , ~ . ~ !-

STA 20, 398 f t STA 2 9 , 596 f t STA 41 , 836 f t

a. Shot No. 5,024, contour interval IO0°K Figure 10. Example of temperature data in spherical coordinates.

m

t ~

- n

¢o & o l

m o t~

31

o l

• . . " . . ". • ~ . ~ .

STA 20 , 398 f t STA 2 9 , 596 f t STA 4 1 , 836 f t

b. Shot No. 5,037, contour interval 100°K Figure 10. Continued.

dA I//~'!-1,5ooOK~_-..~

STA 2 0 , 398 f t STA 2 9 , 596 f t STA 4 1 , 836 f t

C, Shot No. 5,045, contour interval 100°K Figure 10. Concluded.

~>

m

30

¢D

O1

I 8

(1

3) 1. I* "t

5. ,* *I .b# *I .I

I

I . .I S l 11 6-s TI

19s

::: III ,I*

111 11. 111 111

I>. I I< 111 111 1.1 I.. 1.1 111 1,) 1 1 1

I** 111

111 I., 111 17. 111 111 1.1 1.b I,* 111 l S l 1.I ZIY 29. ?.,P 1 1 )

1 1 1 I , . I , . 111 1>J ZZ*

:;i *I I I.*

1 . I I.* 11. nl

I , . 11 1 I * . I H 111 111 *I, I r a 7.e 111 1.. 1.1 a*. I.! l o o I* J 111 l l*

i t : 111 n l 111 I t ' 1 J* 111

11, 111 ,.I

8, deg -. I . * II 1% II I I IF l u 11 - R -I n .. %I r I r r r .I r *. r l rn

Shot No. 5,045, STA 20

Figure I T . Tabulated temperature data in spherical coordinates,

t*J

?-

20 -

18 - c~ I o r-4

o 16 -

4-)

14 1

~J

¢J

q~

12 -

0

) o o

/

/ /

/ ,o

/ /

/4)

9", o _o_~___ I

T r a n s i t i o n L o c a t i o n ( S / r n ) t r ,= 0 . 2 4

! I

I, I

S/R = 0 . 1 7 I 0.35 l I I • i

IO 20

Figure 12.

0 0

0 0 0 0

0

0

0 0

=0

P l a n a r P r o j e c t i o n o f Nos.et. ' ip

0.52 Shot No. 5,020, STA 20, ~ = 27 deg

I I , I I I

30 40 50 60 70

Body Angle, deg

Example of temperature distribution and definition of transition location.

m

0

- n

( D

j~ 01

A ED C - T R - 7 9 - 4 5

C i r c l e s E v e r y 5 deg t o 45,

P l u s 9 0 - d e g C i r c l e

a. Shot No. 5,021, STA 29

b. Shot No. 5,024, STA 20

c. Shot No. 5,037, STA 29

Figure 13. Planar projection of transition front location.

36

A E D C - T R - 7 9 - 4 5

0

t~ V

1 . 0 0

O. 50

0 . I 0

0 . 0 5

m

m

m

D

m

B

0.I

r in. Sym n'

0 1 . 1 2 5 [] 0 . 8 0 & 0 . 4 0

Open = Air Solid = Nitrogen

[ ]

(2)

[:

[]

0

( S / r n) = 0 . 0 3 7 1 p : 1 . 4 1 2 / TR,Mean

I I I

Poo' a r m

I 0 . 5

Figure 14. Variation of mean transition location with range pressure.

I I I I I I

1 . 0

37

O0

0 . 0 8

0 . 0 6

0 . 0 4

0 02

0

0

Sym r n , i n .

0 1 . 1 2 5 [] 0 . 8 0

Open = A i r Solid = N i t r o g e n ------- Least-Squares Fit

0

o a o ~ . . ~

° - n ~ , , , 4 3 . ~ ' ~ 0 0 0 0

_ O 0

I I I I i I I 0 . 1 0 . 2 0 . 3 0 . 4 0 . 5 0 . 6 0 . 7

( S / r n ) T R , M e a n

m 0 (~

~u

~D & Ol

Figure 15. Measurements of transition zone asymmetry.

A E D C - T R - 7 9 - 4 5

Table 1. Nominal Measurement Ranges for Photopyrometers

System Number Exposure Dynamic Designation in Use Duration~ msec Range, °K

Gen-I 3 1,000 1,600 - 3,300

100 1,950 - 4 ,000

G e n - I I 2 1 , 0 0 0 1 , 2 5 0 - 1 , 9 0 0

100 1 , 5 0 0 - 2 , 4 0 0

Oen-ll I 300 1,250 - 1,900

30 1,500 - 2,400

39

A E DC-TR-79.45

Shot No.

4,975

4,985

4,995

5,000

5,001

5,002

5,014

5,015

5,016

5,017

5,019

5,021

5,024

5,027

5,028

5,031

5,032

5,033

5,037

5,038

5,043

5,044

5,045

5,054

5,055

5,056

5,063

rn, in.

1.125

1.125

0.80

0.43

1.125

1.125

1.125

1.125

1.125

1,125

1.125

1.125

10125

1.125

0.80

0.80

0.80

0.80

1.125

1.125

1.125

1.125

1.125

1.125

1.125

1.125

1.125

Table 2. Test Summary

p®, t o r t

200

200

200

200

300

300

300

200

150

200

150

100

200

125

200

150

250

300

125

125

150

150

200

150

250

115

115

Nosetip Recovery

No

Yes

No

Yes

No

Eo

No

No

No

Yes

No

Yes

Yes

No

No

Yes

Yes

No

~0

Ho

Yes

Yes

" No

Yes

Comments

Model development -- no data reported.

Model development -- no data reported.

Model development -- no data reported.

No data -- nosetip failed during launch.

No data -- nosetip failed during launch.

Low launch velocity (14,780 fps).

No data -- nosetlp damaged at range entrance.

Low launch velocity (14,460 fps).

No data -- nosetlp disintegrated in f light.

No data--nosetlp damaged at range entrance.

N 2 test gas.

No data--low launch velocity, temperatures below system thresholds.

40

A E D C-TR -79 -45

Shot No.

5,000

5,014

5,016

5,019

5,021

5,024

5,027

5,028

5,031

5,032

5,033

5,037

5,038

5,044

5,045

5,054

5,055

5,063

Mean

Table 3. Model Angle Data

Pi tch Plane Angle, deg

Dis tance Downrange, f~

Yaw Plane Angle, deg

7~ 156 226 38._..3.3 47.__8.8 58_..~3 70..._3.3 . 5~ 31..__88 823,

0.0 -0 .4 -0.1 -1 .6 -0 .5 -1 .0 -1 .5 0.0 . . . . 0.3

- 0 . 2 - 0 . 5 - 0 . 6 - 0 . 6 - 0 . 5 0 . 0 0 . 0 0 . 5 2 . 0 - 1 . 1

0 . 5 - 0 . 5 0.6 -0 .6 - - - 0.5 0 .0 1.0 1.6 -1 .6

0 . 4 - 0 . 8 1.4 -1 .0 -0 .4 0.9 0.9 0 .8 1.5 - 1 . 0

0 . 0 - 1 . 4 0.3 4 . 9 4 . 3 -0 .4 0.1 -0 .5 0 .9 -1 .0

0 . 4 - 1 . 0 -1 .2 - 3 . 3 ~ . 2 -0 .2 0 .3 -0 .5 0.5 - 0 . 3

0 . 0 - 0 . 4 -0 .5 -1 .6 4 . 7 -0 .5 0 .7 -0 .5 1.3 0 .9

. . . . 0 .7 -0 .6 0.2 -0 .3 0 .5 1.3 - 0 . 6 1.5 0 .3

0.5 . . . . 0 .9 0.0 0.3 0.9 0 .5 - - - 1.3 0.8

0.6 -0 .9 -1.1 -0 .1 1.1 1.9 -0.2- - 0 . 4 1.1 1.3

0.2 -0 .4 -0 .3 0.5 0 .0 0 .9 1.0 0.9 1.8 -0 .3

4 . 2 -0 .9 1.0 0.8 1.7 1.4 1.8 -0 .5 1.8 -0 .2

0.0 -0 .3 0.7 -1 .8 4 . 4 2.5 -1 .2 -0 .3 2.5 0.7

4 . 8 -1 .3 -0 .5 -2 .4 -2 .0 -1 .6 0.1 0.0 0.6 - 0 . 3

-1 .0 1.0 1.2 -2 .9 -0 .7 1.2 1.7 1.0 1.3 -1 .3

-0 .5 -0 .5 -1 .6 -2 .4 4 . 5 -0 .7 1.5 0.6 1.4 2.8

-1.1 -0 .2 1.0 -2 .3 -0 .5 -0 .7 1.0 1.O 1.1 1.1

0 . 1 - 0 . 8 0.9 -1.1 2.0 2.1 - 1 . 0 - 0 . 3 1.6 0.5

10"41 i 0"7 1 0 " 8 1 ,1"31 10"7 I 11"01 !0"81 10"51 11"41 10"91

41

A E D C -T R -7 9 -4 5

A

CD

Poo

rn

S

T

V

W

fl

0

O'tr

NOMENCLATURE

Model base area, ft 2

Drag coefficient

Free-stream pressure, tort

Nose radius, in.

Surface distance, in.

Temperature, °K

Velocity, fps

Model weight, Ib

Distance down range, ft

Ballistic coefficient, WcoAIb/ftz

Angular location on surface in flow direction, deg

Standard deviation of (S/rn)tr measurements (at 3-deg intervals around the nose)

Angular location in circumferential direction around nosetip, deg

S U B S C R I P T S

i Initial

tr Transition location

oo Free stream

42