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V ;i (NASA CR O R T M X O R AD N U M B E R ) PmLIMINARY FiESULTS O F BOUNDARY-LAYER NOISE M E A S U ~ ON THE x-15 AIRPLANE S. Tanner By Eldon E. Kordes and Carole N A S A Flight Research Center Edwards, Calif.
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ff 653 July 65 Second I n t e r n a t i o n a l Conference on Acoustical Fatigue Dayton, Ohio
A p r i l 29 t o May 1, 1964
PRELIMINARY RESULTS OF BOUNDARY-LAYER NOISE MEASURED ON THE x-15 AIRPLANE By Eldon E. Kordes and Carole S. Tanner NASA Flight Research Center Edwards, Calif.
INTRODUCTCON The f l u c t u a t i n g pressures that generate noise i n the boundary layer and the e f f e c t s of t h i s acoustic energy on panel response, panel fatigue, and i n t e r n a l sound l e v e l s are of continuing concern t o the designers of high-speed f l i g h t vehicles. Theoretical analysis and experimental studies using model j e t s , a i r c r a f t , and missiles have been conducted t o define the physical quantities governing the i n t e n s i t y and frequency content of boundary-layer noise ( r e f s . 1 t o 4, f o r example). Most of the f l i g h t - t e s t data have been limited t o Mach numbers l e s s than 2.5 ( r e f s . 1 and 2 ) , although the Scout missile has
gathered boundary-layer-noise data at Mach numbers up t o 4 ( r e f . 3)
and t h e Project Mercury vehicles (ref. 4) up t o Mach 5.7. These f l i g h t - t e s t data have been extremely valuable; however, a l l data obtained on f l i g h t vehicles at Mach numbers of 3 and above have been f o r highly t r a n s i e n t f l i g h t conditions which requires compromise i n the analysis f o r t h e e f f e c t s of various f l i g h t parameters.
In order t o provide d e t a i l e d information on boundary-layer noise over a wide range of controlled f l i g h t conditions, t h e NASA Flight Research Center i s conducting a boundary-layer-noise research program with the X - 1 5 airplane. This paper describes the program and presents some of the preliminary r e s u l t s .
The X-15 airplane has performance c a p a b i l i t i e s t h a t make it useful i n obtaining boundary-layer-noise data f o r Mach numbers up t o 6 and dynamic pressures up t o 1,500 psf. During f l i g h t , many of the conditions can be held e s s e n t i a l l y constant long enough t o obtain excellent data samples, and d a t a can be obtained i n both the acceleration and the The X-15 instrument package i s completely recover- deceleration phases.
able and can be calibrated before and a f t e r each f l i g h t o r modified as required.
The shaded area i n figure 1 shows the X-15 f l i g h t envelope Only t h e portion of the available f o r boundary-layer-noise studies.
X - 1 5 f l i g h t envelope f o r dynamic pressures above 100 psf and a l t i t u d e s below l 5 0 , O O O f e e t i s shown. For dynamic pressures of 1,500 psf, the f l i g h t - t e s t Mach number range extends t o s l i g h t l y greater than 5. For dynamic pressures of 5 O O psf or l e s s , the vehicle i s capable of f l i g h t t o a Mach number of 6.
INSTRUMENTATION It would be desirable t o measure boundary-layer noise at many locations i n order t o study the e f f e c t s of l o c a l flow f o r d i f f e r e n t
. boundary conditions. O n the X-15, however, only a limited number of
a r e a s are available without modification t o the basic airframe. The X - 1 5 side f a i r i n g s were constructed of access panels f o r servicing wiring, hydraulic l i n e s , and control cables. Four s p e c i f i c panels on the o r i g i n a l X - 1 5 side-fairing surfaces have been modified f o r the
boundary-layer-noise studies ( f i g . 2) . The preliminary data presented
i n t h i s paper were obtained from the lower t e s t panel on t h e r i g h t side of the a i r c r a f t , j u s t behind the wing leading edge.
The test-panel instrumentation i s i l l u s t r a t e d i n figure 3. Two c r y s t a l microphones are flush-mounted i n a block a t the f r o n t of the panel. The sectional view shows the microphones i n t h e i r mounting block and a uniaxial accelerometer attached t o the back of t h e mounting block. The noise-measuring microphone i s shown on the r i g h t . The d m microphone on the l e f t measures the temperature environment of t h e diaphragm and the c r y s t a l . This dummy microphone i s required f o r defining t h e temperatures, since it i s not possible t o instrument the active microphone. I m e d i a t e l y behind the microphones i s a t e s t area with a removable panel. After t h e boundary l a y e r has been defined, a study of the response of d i f f e r e n t structures t o boundary-layer noise i s planned. A f t of t h e response panel i s a boundary-layer r&e with 12 total-pressure-measuring s t a t i o n s .
A l l of t h e data, except temperature, from the microphones, accelero- meter, and boundary-layer rake are recorded on an onboard tape recorder.
Figure 4 i l l u s t r a t e s the schematic hookup of the data-recording system.
The microphone and accelerometer signal outputs a r e amplified and recorded on separate channels of the tape recorder. The rake pressure transducer s i g n a l outputs are each fed t o a voltage control o s c i l l a t o r box where they are multiplexed and recorded on the tape recorder. The dummy micro- phone thermocouple s i g n a l outputs are recorded on the onboard oscillograph.
DATA mDUCTION The methods of data analysis used i n t h i s program a r e i l l u s t r a t e d For t h e noise analysis, the f l i g h t tape by t h e block diagram of figure 5 .
i s played back and e i t h e r a time history of o v e r a l l sound pressure l e v e l s i s obtained, or t h e noise i s fed t o a one-third octave-band (OASF'L) analyzer and a time h i s t o r y of a l l the t h i r d octave bands from 50 cps t o 10,000 cps i s obtained. Specific portions of t h e f l i g h t data can be recorded on a tape loop and analyzed i n third-octave-band spectrograms.
The f l i g h t data on the tape loop vary from 3 seconds t o 30 seconds, depending on the f l i g h t parameters selected. The pressure data a r e played back through a discriminator and recorded on oscillograph paper. Accelero- meter data a r e played back d i r e c t l y t o obtain t h e acceleration time h i s t o r y and then played through a frequency analyzer t o obtain g-levels as a -2- function of frequency. The accelerometer data a r e used as a check on the microphone acceleration environment. The thermocouple data from t h e dummy .
microphone are read from the oscillograph f i l m using standard film-reading equipment t o obtain t h e temperature (OF) t i m e h i s t o r y of t h e microphone diaphragm and c r y s t a l .
SYS'IIEM RESPONSE A s might be expected, one of the major problems has been t o obtain a microphone and a recorder t h a t are capable of predictable response over a w i d e range of temperature, a l t i t u d e , and accelerations. On t h e X-15 the e f f e c t s of vibratory acceleration on the microphone output have not been a problem. The c h a r a c t e r i s t i c s of systems used i n t h i s program a r e summarized i n figure 6 i n which t y p i c a l pressure response of the microphone and recorder system i s shown f o r several environmental conditions.
A s shown i n t h e t o p p l o t the complete system has an essentially constant response of -5 db* over t h e frequency range from 50 cps t o 10,000 cps a t room temperature and l o c a l a l t i t u d e (2,500 f t ) . I n order t o i l l u s t r a t e the of t h e microphone with e f f e c t s of t r a n s i e n t heating, the s e n s i t i v i t y change temperature i s shown i n t h e middle plot f o r a constant frequency of 1,000 cps.
The microphone diaphragm was heated over t h e temperature p r o f i l e from -40" F t o 325" F i n 5 minutes, then allowed t o cool. This temperature h i s t o r y i s t y p i c a l of those experienced on an X-15 f l i g h t . During the i n i t i a l portion of heating shown i n the figure, t h e e r r a t i c behavior of t h e microphone i s a t t r i b u t e d t o thermal buckling of t h e diaphragm. Once t h e diaphragm and c r y s t a l a r r i v e a t a stable configuration, the microphone s e n s i t i v i t y change approaches zero with a d d i t i o n a l change i n temperature.
For microphone temperatures below 75" F, the data are considered t o be unreliable. The bottom curve shows the e f f e c t of a l t i t u d e on t h e sensi- t i v i t y change of t h e microphone f o r a constant frequency of 1,000 cps. The s e n s i t i v i t y change i s 2 db or less f o r a l t i t u d e s up t o 100,000 feet.
Similar data have been obtained a t other frequencies and heating rates.
Results of t h i s type a r e used t o adjust the f l i g h t data and t o indicate where improvements i n the system are required.
RESULTS AND DISCUSSION No f l i g h t s f o r the s p e c i f i c purpose of obtaining boundary-layer-noise Equipment has been checked data have been made thus far i n t h e program.
out and gross measurements of boundary-layer noise t o be expected have been During t h i s obtained on f l i g h t s made f o r other general research purposes.
preliminary portion of the program, acceptable boundary-layer-noise data have been obtained on only one f l i g h t . A time h i s t o r y of t h i s f l i g h t i s shown i n A s can be seen, the four f l i g h t parameters are highly t r a n s i e n t ; f i g u r e 7.
*Referred t o 0.0002 dynes/cm2.
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only 5 t o 10 seconds of near-steady-state conditions occurred i n a l t i t u d e , dynamic pressure, or Mach number. The angle of a t t a c k f l u c t u a t e s g r e a t l y ' throughout t h e e n t i r e f l i g h t ; however, f l i g h t s made p r i n c i p a l l y t o obtain boundary-layer-noise measurements would be flown with much b e t t e r control of angle of attack. The noise data measured on t h i s f l i g h t were used t o obtain a t i m e h i s t o r y of o v e r a l l sound pressure l e v e l and one-third octave band spectrograms f o r the near-steady-state Mach number a t 210 t o 220 seconds and f o r two values of dynamic pressure a t 280 t o 300 seconds and 320 t o 335 seconds.
The o v e r a l l sound pressure l e v e l measured on t h i s f l i g h t i s presented i n f i g u r e 8 as a function of f l i g h t t i m e . Data before t h e 170-second point have been omitted because of the e r r a t i c behavior of the microphones a t t h e low temperature during t h e i n i t i a l portion of the f l i g h t (see f i g . 6 ) . The m a x i m u m value of the o v e r a l l sound pressure l e v e l measured on t h i s f l i g h t
i s 147 db. The o v e r a l l sound pressure l e v e l tends t o follow the free-stream
dynamic pressure except f o r a r a t h e r abrupt change i n t h e sound pressure l e v e l a t 250 seconds, when a l a r g e change occurred i n airplane angle of a t t a c k . Although the angle-of-attack change does not a f f e c t the o v e r a l l sound pressure l e v e l d i r e c t l y , an increase i n angle of a t t a c k tends t o compress the boundary layer on t h e l o w e r surface of the vehicle and t o introduce a crossfI~=w arcxd t h e c y l i n d r i c a l fuselage.
These changes i n t h e a i r f l o w at the measuring s t a t i o n would be expected t o a f f e c t the o v e r a l l The data indicate t h a t small v a r i a t i o n s i n the angle sound pressure l e v e l .
of a t t a c k about an average value do not a f f e c t t h e noise l e v e l s . Changes i n t h e o v e r a l l sound pressure l e v e l a t '270 seconds and a t 365 seconds can be correlated with gross changes i n the average angle of a t t a c k ; however, no such c o r r e l a t i o n i s evident a t subsonic conditions.
Spectrograms of the boundary-layer noise a t m a x i m u m Mach number and a t t h e two portions of the f l i g h t with nearly constant dynamic pressure are shown i n figure 9. The two upper curves are f o r d i f f e r e n t ranges of Mach number but f o r t h e same average angle of a t t a c k and approximtely the same dynamic pressure. These two curves show about the same trends between 125 cps and 2,000 cps; t h e e f f e c t of Mach number appears t o be evident only The f a c t t h a t the a t t h e frequencies above 2,000 cps and below 125 cps.
s o l i d curve l i e s below t h e other two curves at a l l frequencies i s a t t r i b u t e d I t o differences i n the l o c a l f l o w conditions.
To show more c l e a r l y t h e differences i n t h e boundary l a y e r f o r the three portions of the f l i g h t analyzed f o r t h e data of figure 9, boundary-layer p r o f i l e s obtained from the rake pressures are presented i n f i g u r e 10.
A s can be seen, t h e region of maximum pressure gradient f o r the curve at M = 3.3 and c x = 0" i s f a r t h e r from the surface than the corresponding region at t h e lower Mach numbers and higher angles of attack. The r e s u l t s indicate that t h e nearer t h e region of m a x i m u m pressure gradient i s t o the surface, the higher t h e sound pressure l e v e l w i l l be ( r e f . 5). Future e f f o r t s w i l l be directed toward a b e t t e r understanding of the f a c t o r s a f f e c t i n g the l o c a l flow and t h e i r c o n t r i - bution t o the boundary-layer noise.
I n figure 8 there i s an apparent r e l a t i o n s h i p between sound--pressure l e v e l and dynamic pressure. This relationship has been noted previously by other i n v e s t i g a t o r s ( r e f . 3, f o r example). The r a t i o of surface pressure
fl
-4- I , t o t’he free-stream dynamic pressure q i s plotted as a function of Mach number i n f i g u r e 11. The s o l i d curve f o r t h e X-15 i s f o r both the climb * and the descent. For comparison, data obtained on a Scout missile during t h e climb (ref. 3) are shown also. The angle of a t t a c k i s e s s e n t i a l l y zero f o r t h e Scout during the e n t i r e f l i g h t and f o r the X-15 during climb from a Mach number of 3 t o 5.3. For the Mach number range from 3.5 t o 4 , both vehicles have values of free-stream dynamic pressure i n excess of 1,000 psf. A s shown, t h e agreement of the data over t h i s range of Mach numbers i s excellent. The higher values of t h e surface-pressure c o e f f i c i e n t f o r the X-15 during descent m y be a t t r i b u t e d t o the e f f e c t of angle of attack, as mentioned previously. For Mach numbers from about 3.5 t o 1.75, t h e average angle of a t t a c k was approximately 4 O , and the values of the surface-pressure coefficient f o r the X - 1 9 a r e approximtely twice t h e values f o r the Scout. Even with the differences i n t h e a c t u a l values of t h e pressure r a t i o , t h e trend with Mach number i s similar f o r is, t h e surface-pressure c o e f f i c i e n t decreases t h e two vehicles, t h a t with an increase i n Mach number up t o 3.5.
CONCLUDING REMARKS Preliminary data on boundary-layer noise obtained from an experimental program with the X - 1 5 airplane show trends s i m i l a r t o data obtained on t h e Scout m i s s i l e f o r comprable f l i g h t conditions. I n addition, these data show t h a t changes i n the noise l e v e l s occur as a r e s u l t of vehicle maneuvers a t supersonic speeds. The X-15 program includes instrumentation f o r obtaining real-time i n f o r m t i o n on the boundary-layer flow as w e l l as w i l l be directed toward a b e t t e r boundary-layer noise, Future e f f o r t s understanding of t h e f a c t o r s affecting t h e l o c a l flow and t h e i r contribu- t i o n s t o the boundary-layer noise.
SYMBOLS M free-stream Mach number P t o t a l pressure, psf
@ root-mean-square surface pressure, psf
q free-stream dynamic pressure, psf a angle of attack, deg Subscript: 12 rake s t a t i o n 12 -5- REFERFNCES 1. McLeod, Norman J., and Jordan, Gareth H.: Preliminary F l i g h t Survey of Fuselage and Boundary-Layer Sound-Pressure Levels. NACA RM ~ 5 8 ~ 1 1 , 1958.
2. Mckod, Norman J.: Flight-Determined Aerodynamic-Noise Environment of N A S A TN D-1160, 1962.
an Airplane Nose Cone Up t o a Mach Number of 2.
3. Hilton, David A., Bracalente, Fhedio M., and Hubbard, Harvey H.: In-Flight NASA TN D-1818, Aerodynamic Noise Measurements on a Scout Launch Vehicle.
1963.
4. Mayes, W i l l i a m H., Hilton, David A., and Hardesty, Charles A.: In-Flight NASA TN D-997, 1962.
Noise Measurements f o r Three Project Mercury Vehicles.
5 . L i g h t h i l l , M. J.: O n Sound Generated Aerodynamically. 11: Turbulence As a Source of Sound. Proc. Roy. SOC. (London), ser. A, vol. 222, 19-54, pp 1-32.
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X-15 FLIGHT ENVELOPE
BOUNDARY- L A Y ER-NO1 SE PROGRAM 175x lo3 9. P S F
I
I50 I25 A LT I TU D E, FT 0 1 2 3 4 5 6 7 MACH NUMBER Figure 1 X-I5 BOUNDARY-LAYER-NOISE TEST PANELS
r7 = PRESENT TEST PANEL
U FUTURE TEST M N E L S
rn
U Figure 2 BOUNDARY-LAYER-NOISE TEST PANEL BOUNDARY-LAYER RAKE A f MY MICROPHONE NOISE MICROPHON NOISE MICROPHONE CCELEROMETER SECTloN A-A F i g u r e 3 DATA- RECOR DING SYSTEM BOUNDARY-LAYER NOISE ACCELEROMETER ~ AMPLIFIER PRESSURE TRANSDUCERS DUMMY MICROPHONE + s c ~ L L o G ~ ~~ Figure 4 DATA-REDUCTION SYSTEM BOUNDARY-LIIYER NOISE - - - - - - - , - FLIGHTTAPE I OSCILUGRAPH ~~ ~ FILM CWliRIYj(rl TEMPERATLRE THERMWUWLE DATA PROFILES CRYSTALTEMPERATUREPROnLES Figure 5 TYPICAL RESPONSE CHARACTERISTICS SENSITIVITY CHANGE, DB -loo 40 100 200 5 0 0 lp00 2,000 5,000 0,000 FREQUENCY,GPS SENSITIVITY 0 CHANGE,DB (1,000 GPS) - 2000 I 2 3 4 5 6 7 8 9 - 2 0 TIME, MIN SENSITIVITY 4 r CHANGE, DB Cl,OOO CPS) 0 IO X, 30 40 50 60 70 80 90 100~10~ ALTITUDE, F T Figure 6 X-15 FLIGHT HISTORY BOUNDARY-LAYER NOISE - ALTITUDE DYNAMIC PRESSURE
MACH NUMBER - 1400
ANGLE OF ATTACK TIME, SEC F i g u r e 7 OVERALL-SOUND-PRESSURE-LEVEL TIME HISTORY 150r 1301 : ' Do 200 300 400 500 LAUNCH TIME, SEC F i g u r e 8 BOUNDARY- LAYER-NOISE SPECTRA M q, PSF a, DEG - 5.3 1.330 EO =4 _ _ _ _ _ _ 3.75 TO 3.3 660 2.8TO 2.35 730 '54
130 -
-
126 -
- - THIRD-OCTAVE- BAND SOUND PRESSURE
1 1 8 -
LEVEL, DB I 1 I L I I 1 1 0 $0 100 200 5 0 0 1000 3150 8000 FREQUENCY, CPS Figure 9 NORMALIZED BOUNDARY-LAYER PROFILES I 2 - M q, PSF a, DEG
IO -
5.3 1,330 Oo -_- 3.5 660 3.8O
a - 2.8 7 3 0 4.3O
i
SURFACE-PRESSURE COEFFICIENT AS
FUNCTION OF MACH NUMBER
- .02 X- 15 DESCENT
u
-
.01
-
.008
SURFACE -
- PRESSURE ,006 COEFFICIENT
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,004 LX-15 CLIMB I I I I I 1