NBSIR 77-1258

An Experimental Study of

Transition on Blunt Bodies

(Part I)

G. Christoph

Naval Undersea Systems Center

Newport, Rhode Island 02840

K. D. Tidstrom

Fluid Mechanics Section

Mechanics Division

Institute for Basic Standards

National Bureau of Standards

Washington, D.C. 20234

July 1977

Final

Prepared for Commanding Officer

Naval Underwater Systems Center

Newport, Rhode Island 02840

NBSIR 77-1258

AN EXPERIMENTAL STUDY OF

TRANSITION ON BLUNT BODIES

(PART I)

G. Christoph

Naval Undersea Systems Center

Newport, Rhode Island 02840

K. D. Tidstrom

Fluid Mechanics Section

Mechanics Division

Institute for Basic Standards

National Bureau of Standards

Washington, D C. 20234

July 1977

Final

Prepared for Commanding Officer

Naval Underwater Systems Center

Newport, Rhode Island 02840

U S. DEPARTMENT OF COMMERCE, Juanita M. Kreps, Secretary

Dr. Sidney Harman, Under Secretary

Jordan J. Baruch, Assistant Secretary for Science and Technology

NATIONAL BUREAU OF STANDARDS. Ernest Ambler, Acting Director

LIST OF SYMBOLS

u' root-mean-square value of longitudinal component of velocity

fluctuation

v’ root-mean-square value of component of velocity fluctuation

lateral to the mean flow direction

w' root-mean-square value of component of velocity fluctuation

vertical to mean flow direction

U mean velocity of flow approaching model

o x axial distance downstream from nose face

x axial distance downstream from nose face at which transition

t ,begins

y nose radius at position x

R maximum nose radius

P surface static pressure

s

P static pressure in flow approaching model

o

P total pressure from surface tube

T

p

p air density

q free-stream dynamic pressure

Cp pressure coefficient, (P

g - P

Q )/q

n frequency

F(n) normalized spectral function
" ' '
EXPERIMENTAL STUDY OF TRANSITION ON BLUNT BODIES

(PART I)

G. Christoph 1 and K. D. Tidstrom

ABSTRACT

Experimental results are presented for the effects of

Reynolds number and angle-of-at tack on the static pressure

distribution and the location of laminar to turbulent transition for the flow over several blunt noseforms; comparisons

are made with analytical methods.

1.

INTRODUCTION

It was the purpose of this experiment to gain further knowledge of the

flow processes over blunt-nosed bodies of revolution. An experimental program

was outlined which was intended to discern the effects of Reynolds number,

background turbulence, and angle-of-attack on transition for several blunt

noseforms. It was felt that the program objectives would be accomplished by

taking the following data:

1. Static pressure profiles.

2. Dependence of transition region on Reynolds number, freestream

turbulence, and angle-of-attack.

3. Relative spectral content of disturbances, at a fixed position

from the surface for before, during, and end of transition.

4. Spatial uniformity of transition around noses.

Relating analytics and the above data would allow for an accurate prediction

of the location of transition on blunt axisymmetric bodies, and would serve

as a starting point for modeling transition as a noise source.

The transition results presented in this report, except for the effect

of free-stream turbulence, are for two different blunt-nosed bodies of

revolution identified as A and B. A subsequent report will present data

on transition for the blunt-nose body of revolution identified as W, as well as data on the effect of free-stream turbulence on body B.

i 2. NBS WIND TUNNEL

The NBS wind tunnel used for this experiment is a closed circuit, low

turbulence tunnel. The test section has a 1.22 x 1.52 meters cross-section.

Previously at Naval Undersea Systems Center, Newport, RI ; presently at

Sun Ship, Chester, PA.
rind has a maximum air speed of 76 m/sec.

A hemispherical nose Pitot-static, tube was used to make longitudinal

static pressure and velocity measurements in the test section of the wind

tunnel. The blunt-nosed body was positioned in the test section in a region where the longitudinal pressure gradient was approximately zero.

In this region, the velocity did not vary by more than 0.7 percent across

the tunnel

.

The turbulence level of the wind tunnel was measured by constant

current hot-wire anemometry using a special vibration isolation mount.

The root-mean-square values of the longitudinal and transverse components

of the turbulent fluctuations, u', v' and w’, respectively, as a function

of air speed are shown in figure 1. Such low values of free-stream turbulence are exemplary for wind tunnel studies of transition [1,2]. 2

Noise measurements of the NBS wind tunnel were made with a General

Radio type 1560P5 microphone 3 which was fed to a General Radio type 1551-C

sound level meter. The microphone was mounted through a hole in the wall

of the wind tunnel in such a manner that the face of the microphone was

parallel to the tunnel wall. In addition, a wind screen was placed over

the microphone face and flush with the tunnel wall.

The noise data were collected on tape at NBS for various air speeds

and two angl es-of-at tack (0° and +3°) of the model. The one-third octave

spectra were obtained at the Naval Undersea Systems Center (NUSC) by playing back the recorded signals to a General Radio type 1521, real time,

one-third octave analyzer. Figure 2 gives the sound pressure level versus

frequency for several air speeds with the model at 0° angle-of-attack

.

No noticeable differences were obtained in the sound pressure levels

when the model was at 3° angle-of-attack.

3. TEST MODELS

Three aluminum, flat-faced axisymmetric models, 20.32 cm in diameter,

were machined at NUSC, Newport. The models are about 30 cm long, and screw into a long cylinder, 20.32 cm in diameter and 1.5 m long. The

assembled configuration is held by a sting and mounted to the tunnel

floor as shown in figure 3. Each model has 24-30 pressure taps placed

so as to provide static pressure measurements from the stagnation region

to an x/R = 1.8, where x is the longitudinal distance from the nose, and

R is the radius of the cylinder. Pressure taps were also provided to

align model at zero angle-of-attack. On Model A, except for the alignment taps, all pressure taps were located along the top generator along

the plane of symmetry. The pressure taps on Model B were located as for

Model A except that additional taps were located near the nose at generators 90° apart and so spaced longitudinally as to give a greater density

Figures in brackets represent literature references at the end of this

paper

.

3 Brand names of equipment are used solely to provide a reference for

performance characteristics and do not represent an endorsement.

2
of measurement. The three models are designated A, B, and W. As seen from

figure 4, all model contours are quite similar, but as seen from figure 5,

the theoretical static pressure distributions are markedly different. Model

A has a broad pressure minimum. Model B has a sharper pressure minimum, and

Model W has two pressure minima. The pressure distribution curves shown

in figure 5 were generated by the Douglas-Neumann potential flow computer

program (developed by the Douglas Aircraft Company)

.

4. MEASURED STATIC PRESSURE DISTRIBUTIONS

Plastic tubing connected the pressure taps on the model to a scanivalve

mounted inside the model. A single tube connected the scanivalve to an inclined manometer. A hemispherical nose Pitot-static tube with a right-anglestem was used to obtain the reference static pressure at a position about

0.5 meters downstream from the flat face of the models. Measured pressure

distributions at a speed of 38 m/sec are compared in figures 6 and 7 to the

theoretical distributions for Model A at 0° angle-of-at tack and for Model

B at -3°, 0°, and +3° angle-of-attack . Measurements at a +3° angle-ofattack correspond to measurements on the leeward side of the model. The

agreement is very good. Pressure distributions were also measured at a speed ot 76 m/sec for Model A, and at speeds of 58 m/sec and 76 m/sec for

Model B. These pressure distributions, within the experimental accuracy

were essentially identical to those shown in figures 6 and 7. The small

differences observed between the theoretical and experimental pressure

distributions may be due to the possibility that the coordinates of the

machined models are not precisely the coordinates that were initially

specified and very slight differences in the computer input coordinates

will affect the theoretical predictions. Another possible reason may be

the effect that the boundary layers on the wind tunnel walls have on the

model pressure distribution. It should be noted that the theoretical calculations do not account for boundary layer growth.

5. LOCATING TRANSITION

Inasmuch as the detection of transition generally requires some judgment

in interpretation, it was desirable to use several methods in order to assess

the accuracy in locating transition. Hot-wire anemometry and a total-head

surface tube were both used to detect transition. The hot-wire probe was constructed of two needles cemented to a lucite sled. The sled was 1.59 cm

long, 1.11 cm wide, and 0.32 cm thick. A platinum wire, 1.27 ym in diameter

and 0.058 cm long was attached across the tips of the needles. Spring tension

held the sled against the model nose, and the distance of the wire above the

surface was adjusted by repositioning the needles. Usually, the distance

above the surface was 0.025 cm. The hot-wire probe and sled are shown in

figure 8. The total-head surface tube consisted of a nickel tube, 1 mm

outside diameter, flattened at the end to form a 0.0127 cm opening. A

static tube, also made of 1 mm nickel tubing was mounted 0.635 cm from the

surface tube. The surface tube arrangement is shown in figure 9.

3
Locating transition by observing and interpreting the hot-wire signal

displayed on an oscilloscope proved to be very difficult. The laminar

boundary layer oscillation frequencies (in the adverse pressure gradient

region) were found to have very high frequencies and were difficult to

distinguish from the frequencies in the transition zone. The narrow extent

of the transition region (only 3 to 6 centimeters) further complicated the

situation. It was obvious when the boundary layer was fully laminar and

when the boundary layer was fully turbulent, but the location of the

transition from laminar to turbulent flow was not obvious. By filtering

out the lower frequencies, the interpretation was somewhat easier. Figure

10 displays oscillograms illustrating the presence of laminar boundary

layer oscillations and the transition from laminar to turbulent flow on Model B for an air speed of 38 m/sec. The signal amplitudes in figure 9 were adjusted for illustration and bear no relation to actual amplitudes.

Another method used to locate transition was to measure the spectra

of velocity fluctuations in the boundary layer. Spectra illustrative of

the spectral behavior of velocity fluctuations associated with transition

are shown in figures 11-14 for Model B at an air speed of 58 m/sec and a 0° angle-of-attack . The measurements shown were made in the boundary layer

at 0.025 cm from the surface, and are presented in terms of the spectral

function, F(n), the energy per unit cycle at frequency, n, normalized with

the total energy. The flow characterized by the spectrum at x/R = 1.71

is regarded as still being laminar with the peaked behavior of the spectrum

indicating the presence of laminar boundary layer oscillations as in the

oscillograms of figure 10. The laminar boundary layer oscillations are a product of the experimental environment coupled to the instability process

and more than one peak in the spectrum can occur, particularly in an adverse pressure gradient. For example, the frequencies of the laminar

boundary layer oscillations increase with speed and at 76 m/sec the

dominant frequency in the spectrum just before transition is 6200 Hz.

At 38 m/sec there are two dominant peaks in the spectrum, 1800 Hz and

2150 Hz, at the beginning of transition. At 76 m/sec it was observed

that the spectral content of the laminar boundary layer oscillations

also varies with angle-of-attack with higher frequencies occurring in the more severe pressure gradient. This was not evident at the lower speed

of 38 m/sec. Within the transition region the spectral content broadens

to include more energy at lower and higher frequencies and at x/R = 2.34

the flow is fully turbulent as characterized by the continuous spectrum

shown in figure 14. It is the initial broadening of the spectrum which

is considered to be indicative of the onset of transition, and for the case shown the onset of transition appears to be at x/R = 1.84.

The surface tube method of detecting transition is illustrated in

figure 15 in which results at three air speeds for Model B, at a +3°

angle-of-attack, are shown. It is seen that for a fixed position above

the model surface, the velocity decreases in the laminar region, increases

rapidly through the transition region, and then decreases, once again, due

to the growing layer in the turbulent region. There was no direct evidence

of a strong steady separation. However, the possibility of a transition

associated with an intermittent separation cannot be ruled out. In order

4
to determine whether the latter does indeed occur, a more extensive effort

would be required.

All three of the above methods were used to locate transition but not

all three were used for each test condition. The results obtained together

with the method of detecting transition are given in tables la and lb. It

is seen from tests made under the same conditions with different methods of

detection that the location of transition was reasonably well-determined to

within ± 0.1R. It should be noted that of the three methods the surface

tube was felt to be the most convenient and easiest to interpret. The

transition location for Models A and B for various wind speeds and at a 0° angle-of-attack were also calculated by the Mangier transformed Michel

method [3]. The results of these calculations are presented in tables 2a

and 2b

In order to assess the spatial uniformity of transition, total-head

surface tube measurements were made for Model B at the 90° (top) and 180°

(side) positions, as one views the model from the front. These measurements

were made at wind speeds of 38 m/sec and 76 m/sec for an angle-of-attack of

0° . The results are presented in figure 16. Transition appears to be

spatially uniform at the lower speed and slightly nonuniform at the higher

speed

6. CONCLUSIONS

1. The measured and theoretical static pressure distributions were in

excellent agreement. This means that one can machine blunt nosed

models accurately enough to give a predicted pressure distribution.

2. The energy associated with the laminar oscillations could be detected

well into the transition region.

3. Transition on the blunt bodies under test was not as easily identifiable as on a flat plate, that is, sudden large jumps in the mean

velocity close to the surface associated with turbulent spots were

not observed on the blunt bodies. If the fact that the transition

region is noisier than the turbulent region on a flat plate is due

to abrupt increases in the mean velocity, then perhaps on blunt

bodies the flow noise levels in the transition and turbulent regions

do not differ greatly.

4. At low Reynolds numbers, transition was uniform around the models.

At higher Reynolds numbers, transition appeared to be nonuniform.

This nonuniformity has also been observed in previous work at NBS

and in flow visualization tests conducted recently at the National

Physical Laboratory in Teddington, England, and appeared to be

associated with the formation of turbulent wedges.

5. Of all the transition prediction methods compared to the data, the

Mangier transformed Michel method was the only one which gave satisfactory agreement

5
7 . ACKNOWLEDGMENT

The authors would like to express their appreciation to D. Quadrini

of NUSC for his assistance in making the noise measurements, and L. M.

Sargent of NBS for his assistance in taking and compiling data. The

authors gratefully acknowledge the interest and support of P. S. Klebanoff.

8. REFERENCES

1. Schubauer, G. B. and Skramstad, H. K., Laminar Boundary Layer Oscillations and Transition on a Flat Plate, NACA Rept. 909, 1948.

2. Wells, C. S., Effects of Free Stream Turbulence on Boundary-Layer

Transition, AIAA Journal, vol. 5, no. 1, pp . 172-174, 1967.

3. Granville, P. S., Comparison of Existing Methods for Predicting

Transition from Laminar to Turbulent Flou on Bodies of Revolution,

NSRDC TN 111, August 1968.

6
Table 1(a)

Transition Location on Model A

U (m/s)

o

x

t /R Angle-of-Attack Detection Method

23 2,.5 30 2..2 38 2,. 1 46 1 ..9 76 1 .8

0

°

0

°

0

° 0

°

0

°

Oscillogram

Oscillogram

Osci 1 logram

Oscillogram

Osci 1 logram

Table 1(b)

I Transition Location on Model B

U (m/s)

0

x

t /R Angle-of-Attack Detection Method

38 2.1 0° Oscillogram

38 2.1 0° Spectra

38 2.0 0° Surface-tube

38 2.0 + 1° Oscillogram

38 1.8 + 3° Oscillogram

38 1.7 + 3° Surface-tube

38 1.6 + 3° Spectra

38 2.2 -1° Oscillogram

38 2.3 -3° Spectra

58 1.8 0° Spectra

58 1.8 0° Surface-tube

58 1.1 + 3° Surface-tube

76 1.3 0° Oscillogram

76 1.2 0° Spectra

76 1.2 0° Surface-tube

76 1.2 + 1° Oscillogram

76 0.6 + 3° Surface-tube

76 1.7 -1° Osci 1 logram

76 1.9 -3° Spectra
Tab 1 e 2(a)

Calculated Transition Location on Model A

U ( m/ s

)

o

v 7 x

t /R Angle-of-Attack

23 2.25 0°

30 2.00 0°

38 1.85 0°

46 1.75 0°

76 1.52 0°

Table 2(b)

Calculated Transition Location on Model B

! 0

(m/s) x

t /R Angle-of-Attack

38 1.52 0°

58 1.20 0°

76 0.95 0°
o

o

in

r~

o

in

inCN

O

r—\X

O

D

Uo

(m/s)

FIG.

1.

FREESTREAM

TURBULENCE

LEVELS

VS.

SPEED

IN

NBS

WIND

TUNNEL
120

N

E

4^

(awva %£z v ni ava * 3000*0//ap)

aaAaa aanssaad aunos

FIG.

2

.

SOUND

PRESSURE

LEVEL

VS.

FREQUENCY
V 11

FIGURE

3.

ASSEMBLED

MODEL

IN

WIND

TUNNEL
o

00 U3 CM

• • • o o o o

\

>1

12

FIG.

4.

NOSE

CONTOURS

FOR

MODELS

A,

B,

AND
FIG.

5.

PRESSURE

DISTRIBUTIONS

FOR

MODELS

A,

B,

AND
00

a

u

x/R

FIG.

6.

PRESSURE

DISTRIBUTION

FOR

MODEL

A

AT

0

0

ANGLE-OFATTACK
o

PC \X

ro

r-~ O

M

Cm

a,

CJ

PRESSURE

DISTRIBUTION

FOR

B

NOSE

AT

-3°

ANGLE-OF-ATTACK
o

a

FIG.

7b.

PRESSURE

DISTRIBUTION

FOR

B

NOSE

AT

0°

ANGLE-OF-ATTACK
o

CM

00

KD

OS)

00o

>x> o

o

CNJ O

\

X

o

l o

I o

I

a

u

17

FIG.

7c.

PRESSURE

DISTRIBUTION

FOR

B

NOSE

AT

+3°

ANGLE-OF-ATTACK
18

FIGURE

8.

HOT-WIRE

PROBE

AND

SLED

FOR"UETE(JTING

TRANSITION
19

FIGURE

9.

SURFACE-TUBE

ARRANGEMENT

FOR

'DETECTING

TRANSITION
SWEEP

SPEED, 2

ms/division

20

FIGURE

10.

OSCILLOGRAMS

OF

VELOCITY

FLUCTUATIONS

ON

MODEL

B(U
F ( n ) xl 0“ 3

n (kHz)

FIGURE 11. Normalized Spectrum, Model B at 0° Angle-of- Attack

.
F ( n ) xl 0

3

FIGURE 12. Normalized Spectrum, Model B at 0° Anqle-of-Attack

.
F (n) xlO

FIGURE 13. Normalized Spectrum, Model B at 0° Angle-of-Attack

.
N

VC

5*5

CMO

CMO

OO

C o

r—

Cn

x

FIGURE

14.

Normalized

Spectrum,

Model B

of

0°

Angle-of-Attack.
o o o o o o

w

CL,

Itr*

CNJO

D

o. JC

FIG.

15.

DETERMINATION

OF

TRANSITION

WITH

SURFACE

TUBE

ON

MODEL

B

AT

3°

ANGLE-OF-ATTACK
38

m/s

in

o o o o

V- \X

w

0.

IH

CU

CMc> D

Q

Jf

FIG.

16.

UNIFORMITY

OF

TRANSITION

ON

MODEL

B

AT

0

0

ANGLE-OFATTACK
NBS-1 1 4 A (REV. 7-73)

U.S. DEPT. OF COMM. BIBLIOGRAPHIC DATA

SHEET

1. PUBl.K ATI ON OR Rl POR 1' NO. NBSIR 77-1258

2. C»ov 't An i s s ion No.3. Ree ipient \s Ai cession No.

4. TITLE AND SUBTITLE

EXPERIMENTAL STUDY OF TRANSITION ON BLUNT BODIES (PART I)

5. Publication Date July 1977

6. Performing Organization (.ode

7. AU'ITIOR(S) . G. Christoph and K. D. Tidstrom8. Performing Organ. Report No.

9. PER FORMING ORGANIZATION NAME AND ADDRESS

NATIONAL BUREAU OF STANDARDS

DEPARTMENT OF COMMERCE

WASHINGTON, D.C. 20234

10. Project /Task /Work Unit No. 2130487

11. Contract/Grant No. PO-4-OOll

12. Sponsoring Organization Name and Complete Address (Street, City, State, ZIP)

Naval Underwater Systems Center

Newport, Rhode Island 02840

13. Type of Report & Period Covered

Final

14. Sponsoring Agency (ode

15. SUPPLEMENTARY NOTES

16. ABSTRACT (A 200-word or less factual summary of most significant in formation. If document includes a significant

bibliography or literature survey, mention it here.)

Experimental results are presented for the effects of Reynolds number and angleof-attack on the static pressure distribution and the location of laminar to

turbulent transition for the flow over several blunt noseforms; comparisons

are made with analytical methods.

17. KEY WORDS (six to twelve entries; alphabetical order; capitalize only the first letter of the first key word unless a proper name; separated by semicolons) Axisymmetric ; boundary layer; blunt-nose; hot-wire anemometry;

laminar; pressure-distribution; transition; turbulent.

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