NASA T E C H N I C A L

M E M O R A N D U M

e h

N A S A TM X-1574

WIND TUNNEL INVESTIGATION OF FLUCTUATING PRESSURES ON

AT TRANSONIC SPEEDS

by Bernard J. Bl~zha Lewis Research Center CZeveZand, Ohio

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N A T I O N A L A E R O N A U T I C S A N D SPACE A D M I N I S T R A T I O N W A S H I N G T O N , D. C. MAY 1968 ~

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NASA TM X-1574

WIND TUNNEL INVESTIGATION OF FLUCTUATING PRESSURES

ON A l/lO-SCALE CENTAUR MODEL AT TRANSONIC SPEEDS

By Bernard J. Blaha

Lewis Research Center Cleveland, Ohio

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

For sole by the Clearinghouse for Federal Scientific ond Technical Information Springfield, Virginio 22151 - CFSTI price $3.00

WIND TUNNEL INVESTIGATION OF FLUCTUATING PRESSURES

ON A 1110-SCALE CENTAUR MODEL AT TRANSONIC SPEEDS

Wind tunnel tests of a l/lO-scale Centaur model with a Surveyor nose cone were con- I

ducted over a range from Mach 0.55 to 1.95 to determine amplitudes and frequencies of fluctuating pressures that exist on the model surface at these speeds. Tests were con- ducted for roll angles of Oo, 90°, and 180' at 0' angle of attack. Peak-to-peak fluctua- ting pressure levels up to 30 percent of tunnel dynamic pressure were measured immedi-

by Bernard J. B laha

Lewis Research Center

SUMMARY

INTRODUCTION

Several previous investigations showed that surface pressure fluctuations of consid- erable amplitude occur on missiles in the transonic speed range (e. g., refs. 1 to 5). These fluctuations are particularly severe just aft of the cone-cylinder juncture of the payload shroud and also upstream of protuberances as a result of shock and boundary layer interactions. They are also observed downstream of protuberances where the sep- arated flow reattaches to the surface. In order to establish the fluctuating pressure dis- tribution on a Centaur stage a t transonic and low supersonic speeds, a series of tests were conducted in the Lewis 8- by 6-foot Supersonic Wind Tunnel on a 1/10-scale model with a Surveyor nose cone. Fluctuating pressures were investigated over the length of the model and in local regions of particular interest, such as the various model protu- berances. Investigations were made over a range from Mach 0.55 to 1.95 a t roll angles of Oo, 90°, and 180' at 0' angle of attack. The results of these investigations are pre- sented in this report.

SYMBOLS

Pmax - Pmin q" ACp maximum fluctuating peak-to-peak pressure coefficient,

D

f

M

P

q

V

1

e

(P

- nominal full-scale Centaur diameter, 120 in. (304.5 cm)

frequency, Hz

Mach number

static pressure

dynamic pressure

velocity

axial distance along vehicle measured from cone shoulder

angular coordinate of transducer locations

model roll angle

Subscripts:

0 f ree stream

max maximum value

min minimum value

APPARATUS AND PROCEDURE

The model w a s 62.02 inches (152.5 cm) long and 12 inches (30.5 cm) in diameter and was installed in the transonic portion of the Lewis 8- by 6-foot Supersonic Wind Tunnel (fig. 1). It was comprised of three major sections: the Surveyor nose cone, the Centaur stage, and the interstage adapter. Thirty piezoelectric pressure transducers, having a rated frequency response of over 20 kilohertz, were installed on the model (fig. 2). These locations were selected to yield the general distribution of fluctuating pressures over the length of the stage downstream of the cone shoulder (fig. 2(a)) and in local regions of particular interest such as the cone shoulder (fig. 2(b)), the hydrogen vent stack (fig. 2(c)), and the umbilical island and boost pump fairing (fig. 2(d)). Table I shows the locations of the transducers in t e rms of 1/D, which is the ratio of axial dis- tance measured along the vehicle from the cone shoulder to the vehicle diameter. The transducers (0.25-in. (0.63-cm) diam) were mounted flush to the model surface. Dynamic pressure fluctuations were recorded on FM magnetic tape through an electrical

2

I -

62.02 in. (152.5 cm)

_ _ I_

(30.5 cm)

/

adapter LCentaur Transonic

I I Lsurveyor // Linterstage

nosecone 1

stage

E) Air flow

\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\\ Tunnel floor

TT iTS

CD-9237

Figure 1. - Installation of the 1/10 scale Centaur dynamic pressure model in the Lewis 8- by 6-foot Supersonic Wind Tunnel (not to scale).

Angular coordinate, 8 =-I 1 - P I 1 l l I l I I 2 3 4 5 6 7 8 9

II'I, I 1 1 10 \ 13 14 15 Reference

transducer

CD-9238 180"

Normal f l ight position looking downstream

(a) Longitudinal distribution; angular coordinate, 90".

Figure 2. - Location of instrumentation on Centaur 1110 scale model. Model rol l angle, 0"

3

! Transducer 16 17 18 19

(b) Downstream of cone shoulder; angular coordinate, 0"

( c ) Hydrogen vent stack; angular coordinate, 245"

1ID

Transducer 23 24 25 26 27 28 '-29 30 CD-9239

(d) Umbilical island and boost pump fairing; angular coordinate, 270". Figure 2. - Concluded.

system that had good recording characteristics up to 10 kilohertz (described in ref. 5).

and 180' about the model centerline at 0' angle of attack to the stream flow. A s seen in 4 2 figure 3(a), free-stream dynamic pressure qo varied from 3.2 psi (2.21X10 N/m ) a t

Mach 0.55 to 9 .0 psi (6.21X10 N/m2) at Mach 1.95. These pressures were from 2 1 to 25 percent higher than those expected in flight over the Mach number range fo r a typical high q trajectory. Figure 3(b) shows that, although the dynamic pressures significantly differed, the tunnel stream velocities were within 8 percent of those estimated for flight.

The model w a s tested over a range from Mach 0.55 to 1.95 at roll angles of Oo, 90°,

4

4

I .

TABLE I. - TRANSDUCER LOCATIONS

'ransdu c e r number

1 2 3 4 5

6 7 8 9

10

11 12 13 14 15

16 17 i6 19 2 0

2 1 22 23 24 2 5

26 27 28 29 3 0

Model location and angular coordinate,

0

Longitudinal distribution a t 90'

Cone shoulder at Oo

I Hydrogen veyt stack at 245'

Umbilical island at 270' I

Boost pump fa i r ing and inter-

stage at

Length-to- l i ameter ra t io ,

1 /D

0.029 .217 .455 .584 .650

.816 1.020 1.270 1.640 2. 150

2.180 2.230 2.420 2.880 3.380

.029

.217

.455

. 650

. 550

.633

.716

.354

. 500

.633

2.320 2.480 2.640 2.840 3.240

5

>-

._ 2 10 a

L 2

m z CL c

5 2

“ 2 : E E c

1 = 2 m

(a) Dynamic pressure. 6x102

i

Y

. 4 . 6 . 8 1.0 1.2 1 4 1.6 1.8 2.0 Free-stream Mach number, Mo

(b) Velocity.

Figure 3. - Comparison of 8- by 6-foot Supersonic Wind Tunnel dynamic pressure and velocity with typical Atlas-Centaur flight trajectory.

RES U LT S

Background Noise

In the early phases of the test, i t was observed that discrete frequencies existed in the background noise of the wind tunnel which were not related to the flow characteristics of the test model and, hence, may not occur in flight. To further investigate the tunnel background noise, empty tunnel tests were made, and transducer measurements were obtained in the tunnel bellmouth, on the test section walls, and in the subsonic diffuser. A power spectral density analysis of a transducer located on the wall of the empty tunnel indicated the predominant frequencies and amplitudes of the tunnel-induced noise. The

value of 0.03 sound pressure levels were highest at Mach 0.775 and had a AC (fig. 4(a)). The most predominant frequencies at this Mach number were between 400 and 500 hertz and were estimated to contain about 75 percent of the r m s noise signal. This noise appeared to originate from the subsonic diffuser of the wind tunnel. The lower noise signal a t 820 hertz can be traced to the compressor speed. The r m s noise levels were considerably reduced a t Mach numbers greater than 1.0. At Mach 1.10

P, rms

6

3

- (a) Free-stream Mach number, 0.775; rms fluctuating pressure coefficient, - L -

0.030. >; s 5 .&lo7? c

01 V

m

c al V

m L c - -

a u ;## L a L a, 2

. 1

0

(b) Free-stream Mach number, 1.10; rms fluctuating pressure coefficient,

Figure 4. - Power spectral density analysis of transducer located on sidewall of

0.0084.

empty 8- by 6-foot Supersonic Wind Tunnel.

(fig. 4(b)), the predominant frequencies varied from about 730 to 1000 hertz, with the peak amplitude at 840 hertz. Some smaller amplitude noise was present at frequencies below 200 hertz and between 1500 and 2000 hertz. The predominant frequencies and their corresponding peak-to-peak fluctuating pressure coefficient are summarized in f igure 5 over the Mach number range for the empty tunnel. In general, low-frequency, high- amplitude noise signals were obtained at subsonic Mach numbers from about 0.65 to 0.90 while high-frequency, low-amplitude noise was measured at the supersonic Mach numbers.

amplitudes may have differed considerably from those measured on the tunnel wall because of differences in local flow conditions. Additional disturbances also occurred on the model as a result of the local flow characteristics created by the model geometry which presumably would also occur in flight. Because of the background noise, a de- tailed frequency analysis of the model data w a s not attempted. Also, because of the

Disturbances occurred on the test models at the same frequencies; however, the

7

6

5

4

3

2

1 I L 1 I

8 1.0 1.2 1.4 1.6 1.8 Free-stream Mach number, Mo

(b) Frequencies.

Figure 5. -Typical amplitude and frequency data obtained i n empty 8- by 6-foot Supersonic Wind Tunnel.

8

.

0

background noise, the average fluctuating pressure level on transducer 15 was used as a reference level for all the other transducers. It was located downstream on the model inter stage away from protuberances and areas where unstable and dynamic flow conditions were expected. The resulting fluctuations seen on this transducer were relatively low over the Mach number range (fig. 5(a)) and were comparable to levels expected in a tur- bulent boundary layer (ref. 6).

In the present report, peak-to-peak variation of the local pressure coefficient AC is presented without correction for tunnel background noise. The relative trends and distributions were presumed valid and indicated those portions of the vehicle which would be subjected to large amplitudes of fluctuating pressure in flight.

P

I i .

Longitudinal Distribution

Fluctuating pressures indicated by the 15 transducers mounted longitudinally on the model side at angular coordinate 8=90° a r e presented in figure 6 for each Mach number tested. The AC profiles indicate that large peak-to-peak pressure fluctuations (up to 28 percent of 907 were experienced on the Surveyor shroud at low transonic speeds, with the amplitudes decreasing to lower levels at speeds above Mo = 0.85. These large fluc- tuations could have been caused by the unstable flow conditions that exist behind a cone shoulder at transonic Mach numbers; namely, the alternate separation and attachment of the f low and eventual formation and movement of the terminal shock downstream of the

AC = 0.02 which is below the level of pressure fluctuations, especially in the transonic P regime. At the high supersonic Mach numbers, the data approach the level of electrical noise which indicates little o r no fluctuating pressure activity. The peak amplitudes of fluctuating pressure in the area of the cone shoulder range up to 19 percent of qo larger than those experienced by the reference transducer 15. A comparison with data presented in reference 1 for the longitudinal dynamic pressure distribution over a clean configura- tion (no protuberances) of a 0.074 scale Centaur model shows similar trends and ampli- tudes over the same Mach number range, especially in the area of the cone shoulder. Substantial variations in peak- to-peak pressure fluctuations existed for different model roll angles indicating an inadvertent misalinement of the model axis with the free-stream flow direction and also a significant sensitivity of the data to small angles of attack.

Figure 7 presents the data as a function of Mach number for each transducer. Again, the data show that large peak-to-peak pressure fluctuations existed on the model in the low transonic regime, with the highest being 28 percent of qo at Mo = 0.7.

shroud at e=Oo are presented in figure 8. A comparison with four transducers in similar positions for the complete longitudinal row (transducers 1 to 4 on fig. 7) shows similar magnitudes and trends with Mach number. The maximum AC was about 30 percent of 90 at Mach number 0.75. The Mach number of maximum activity for the two sets of transducers differs only by about 0.05. The transducers located on and near protuber- ances of the Surveyor shroud (transducers 3 and 19) show relatively low pressure amPli- tudes in comparison with those existing near the cone shoulder. The maximum value was about 20 percent of 90 at Mo = 0.75.

shou:der* Electrjcd letvels for a:: the traiis&Ucera ai -upper liiiiit of ?&o-ut

Data for four other transducers downstream of the cone shoulder on the Surveyor

P

9

.20

0

I I

(a) Free-stream Mach number, 0.55.

I I I I I I

.a

.M

n u n a - 3 (b) Free-stream Mach number, 0.65.

.- (c) Free-stream Mach number, 0.7. c

I m Y (d) Free-stream Mach number, 0.75. n. .M

0 (e) Free-stream Mach number, 0.77.

.20

0 (f) Free-stream Mach number, 0.8.

.M

0 .5 1.0 1. 5 2.0 2. 5 3.0 3. 5 0

Ratio of distance from cone shoulder to vehicle diameter, ZID

(g) Free-stream Mach number, 0.85.

10

.20

0 (h) Free-stream Mach number, 0.9.

.20 n

V Q c- 0 c ~

W u (i) Free-stream Mach number, 1.0. .- .- - b .20

2 0 u

3

$ 0

s .20

2 0

9 .20

L

Ul (j) Free-stream Mach number 1.1. n

m 3

u 3 c

- m W n

Y m W a

0 (0 Free-stream Mach number, 1.46.

.20

0 0 . 5 .10 .15 .20 .25 .30 35

Ratio of distance from cone shoulder to vehicle diameter, [ I D

(m) Free-stream Mach number, 1.95.

I l l I I '7 Reference transducer Transducer 1 2 3 4 5 6 7 8 9 10-

{ - 1

I 1 ' Full scale diameter, 120 in. (344.5 cm)

Figure 6. -Concluded.

11

I ' Z,' I I Model rol l angle,

0 180 Reference transducer number 15 Upper l imi t to electrical noise leve .2

for a l l transducers

0 =/=4-4 (a) Transducer 1.

(b) Transducer 2

(cl Transducer 3.

(d) Transducer 4.

le) Transducer 5 . 2

0

( f ) Transducer 7 . 2

0 . 4 . 6 . 8 1.0 1.2 1.4 1.6 1.8 2.0

Mach number, M,

(gl Transducer 8

Figure 7. - Longitudinal pressure fluctuations as function of Mach number, Angular coordinate, 90".

2 Reference transducer

0

(h) Transducer 9. . 2

n 0 a -- 0

% . 2

2 0 E

E? z . 2

g o

c a V .- .- L

0 V

aJ L =I

n

m 3

V 3 - - c

m B (k) Transducer 12. 7 . 2 0

Y m a a

0

( I ) Transducer 13. 2

0 . 4 . 6 .8 1.0 1.2 1.4 1.6 1.8 2.0

Mach number, M,

(m) Transducer 15.

Figure 7. - Concluded.

13

I I I I I I I

Model roll angle, cp#

deg 0 0 0 90 0 180 . 4

. 2 CL

V -4 c- 0 c al

V (a) Transducer 16. .-

c V (b) Transducer 17. 3

." al (c) Transducer 18.

2

0 . 4 . 6 .8 1.0 1.2 JJl-LLlal~ 1.4 1.6 1.8 2.0 1 Mach number, Mo

(d) Transducer 19.

Figure 8. - Pressure fluctuations as function of Mach number for transducers downstream of cone shoulder angular coordinate, 0'.

Downstream of Hydrogen Vent Stack

The AC sented in figure 9. Fluctuating pressure levels behind the vent appear to be low except for the transducer located farthest downstream from the stack (transducer 22). This transducer showed peak-to-peak amplitudes up to 18 percent of s, at Mo = 0.75. This activity was probably influenced by a flow disturbance emanating from the adjacent umbilical island.

values for the transducers downstream of the hydrogen vent stack are pre- P

14

I

.- L L

a, 0 U

m L 3

Ztlookinq downstream -

Model ro l l angle, cp,

de9 0 0 0 90 0 180

0 VI Y) a, (a) Transducer 20.

(bl Transducer 21. a 0 . 2 b + x a,

. 4 . 6 . 8 1.0 1.2 1.4 1.6 1.8 20 Mach number, Mo

(c) Transducer 22.

Figure 9, - Pressure fluctuations as function of Mach number for transducers downstream of hydrogen vent stack. Angular coor- dinate, 245".

Umbil ical Is land Fa i r ing

Data from the umbilical island fairing (fig. 10) show strong pressure fluctuations ranging between 18 and 30 percent of go in the transonic speed range. These fluctua- tions are about 5 percent of go larger than those seen on other transducers in compa- rable positions on the model (transducers 3 to 5 on fig. 7, and 18 and 19 on fig. 8). This indicated that the umbilical island fairing geometry was contributing to the magni- tude of these disturbances.

15

c U

a3 0 U

.- al (a) Transducer 23.

.- - L .

P

c U (b) Transducer 24. = . 4

2i b

.- ..- m 3

3

Y m

2 . 2 m al a

0 . 4 . 6 . 8 1.0 1.2 1.4 1.6 1.8 2.0

Mach number, Mo

(c) Transducer 25.

Figure 10. - Pressure fluctuations as a function of Mach number for transducers on umbilical island fairing. Angular coordinate, 270".

Boost Pump Fa i r ing and lnterstage Adapter

Data from the boost pump fairing and the interstage adapter downstream of the boost pump are presented in figure 11. The transducer onthe boost pump (transducer 26) showed considerable activity over most of the Mach range. At supersonic speeds, the fluctuations averaged about 10 percent of 9,. The transducers downstream of the boost pump indicated that the wake of the fairing had a definite effect on the level of pressure fluctuations. When compared with transducers 12 to 15 (fig. 7), those behind the boost pump showed a n average increase in activity of about two percent of go. Transducer 27

showed a low level of activity because it was within the separated flow region close to the fairing, while transducer 28 showed pressure amplitudes in the transonic regime as high as 17 percent of go because it was probably near the point of flow reattachment.

16

7-1 1 - 7 Normal fl ight position looki ng downstream

. 2 ------- Reference transducer number 15 I

0 (a) Transducer 26. CL

I 0

V

P (b) Transducer 27. - 3

01 L a

8

. 2

m z o m 5 - U (c) Transducer 28. = . 2 3

Y m al

0

n c

5 0 - m L1 (d) Transducer 29.

2

0 . 4 . 6 . 8 1.0 1.2 1.4 1.6 1.8 2.0

Mach number, Mo

(e) Transducer 30

Figure 11. - Pressure fluctuations as funct ion of Mach number for transducers o n the boost pump fairing and interstaye adapter. Angular coorainate. iiu-.

SUMMARY OF RESULTS

Wind tunnel tes ts of a l / l0-scale Centaur model with a Surveyor nose cone were con-

of fluctuating pressures that exist on the model surface at these speeds. Tests were con- ducted for rol l angles of Oo, 90°, and 180' at 0' angle of attack. Because of noise in the tunnel, a detailed frequency analysis of the data was considered impractical. The fol- lowing resul ts were obtained:

1. Although the tunnel noise was present, those portions of the model subjected to

ducer measurements to one transducer located on the model away from protuberances

I

I ducted over a range of from Mach 0.55 to 1.95 to determine amplitudes and frequencies I

I large amplitude pressure fluctuations were determined by comparing the various trans- I

17

such that it should be subjected only to the noise level of the turbulent boundary layer. The areas of maximum activity were immediately downstream of the cone shoulder, the umbilical island, and the boost pump fairing.

2. Maximum pressure fluctuations usually occurred at Mach numbers from 0.70 to 0.90. Peak-to-peak fluctuating pressure levels up to 30 percent of free-stream dynamic pressure were measured.

I

I

Lewis Research Center, National Aeronautics and Space Administration,

Cleveland, Ohio, August 22, 1967, 120-03-01-08-22.

REFERENCES

1. Coe, Charles F. : Steady and Fluctuating Pressures a t Transonic Speeds on Two Space-Vehicle Payload Shapes. NASA TM X-503, 1961.

sure Fluctuations Measured on 15' and 20' Cone-Cylinders at Transonic Speeds. 2. Coe, Charles F. ; and Kaskey, Arthur J. : The Effects of Nose Bluntness on the Pres-

NASA TM X-779, 1963.

3. Chevalier, H. L. ; and Robertson, J. E. : Pressure Fluctuations Resulting from an Alternating Flow Separation and Attachment at Transonic Speeds. 204, DDC No. AD-423095), Aro, Inc., Nov. 1963.

4. Robertson, J. E. : Steady and Fluctuating Pressures on Cone-Cylinder Missile Con- figurations at Transonic Speeds. (AEDC-TR-65-269, DDC No. AD-477938). Aro, Inc., Feb. 1966.

(AEDC -TDR-63-

5. Nacht, Michael L. ; and Turk, Raymond A. : Origin of High-Intensity Noise in the Area of a Large Protuberance on a Launch Vehicle. NASA TM X-1261, 1966.

6. Jones, George W . , Jr. ; and Foughner, Jerome T . , Jr. : Investigation of Buffet Pressures on Models of Large Manned Launch Vehicle Configurations. NASA T N D-1633, 1963.

18 NASA-Langley, 1968 - 3 1 E-3965