Document
NASA TECHNICAL NOTE
N A S A TN # -- D-3364
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FLIGHT MEASUREMENTS
OF BOUNDARY-LAYER NOISE
ON THE X-15
by Thomus L. Lewis and Normun J. McLeod
Flight Reseurch Center
Edwurds, Cu l$
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 ,
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TECH LIBRARY KAFB, NM
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0330555 NASA TN D-3364 FLIGHT MEASUREMENTS O F BOUNDARY-LAYER NOISE ON THE X-15 By T h o m a s L. Lewis and Norman J. McLeod Flight R e s e a r c h Center Edwards, Calif.
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - Price $0.20 FLIGHT MEASUREMENTS OF BOUNDARY-IAmR NOISE ON THE X-l?* By Thomas L. Lewis and Norman J. McLeod F l i g h t Research Center SUMMARY Boundary-layer-noise d a t a measured i n f l i g h t with the X-15 airplane over a Mach number range from 1.0 t o 5.4 and at a l t i t u d e s from 45,000 feet t o 105,000 f e e t are presented. The data were obtained a t four locations on t h e airplane, selected t o provide varied boundary-layer conditions. The highest recorded noise l e v e l w a s 150 decibels. kundary-layer parameters w e r e measured a t one location and a r e used t o present t h e noise data i n a nondimensional form f o r comparison with d a t a from f l a t - p l a t e wind-tunnel studies by other experimenters.
An appendix i s included which i l l u s t r a t e s t h e s e n s i t i v i t y of the micro- phones used t o obtain t h e data as they are affected by a l t i t u d e , vibration, and t r a n s i e n t heating.
INTRODUCTION An important aspect of research on boundary-layer flow i s the noise gener- ated i n t h e boundary layer by fluctuating pressures. One of the f a c t o r s affected by the fluctuating pressures i s s t r u c t u r a l weight. This e f f e c t i s demonstrated by use of a simplified structure, as i l l u s t r a t e d i n figure 1 (taken from r e f . 1). I n t h i s example, the skin thickness required f o r adequate fatigue l i f e i n an acoustic environment i s shown. The schematic diagram a t the lower r i g h t i l l u s t r a t e s t h e structure. The r e s u l t s presented show t h a t an increase of sound l e v e l from 115 dB t o 125 dB would require an increase i n skin thickness from 0.03 inch t o about 0.06 inch. An increase i n r i b thickness i s a l s o required t o withstand the more severe environment. This increase of 10 decibels i n sound l e v e l would r e s u l t i n a s t r u c t u r a l weight increase of a t l e a s t 75 percent.
Noise l e v e l s shown i n t h i s example are lower than may be encountered i n f l i g h t , but t h e curve c l e a r l y indicates t h a t , f o r higher noise levels, t h e increase i n weight w i l l be greater. Also, acoustic insulation necessary f o r passenger comfort w i l l increase t h e weight.
Graphs of t h i s type established on t h e b a s i s of ground t e s t s might be used i n t h e design of a i r c r a f t structures i f t h e noise l e v e l s could be predicted f o r - - - . - .~ . . - *This paper w a s included i n a c l a s s i f i e d report e n t i t l e d "Fourth Confer- ence on Progress of the X - 1 5 Research Airplane Program," F l i g h t Research Center, Oct. 7, 1965. N A S A SP-90, 1965. An appendix has been added t o describe t h e instrumentation, and i t s frequency response, t h a t w a s used i n obtaining the data.
new flight regimes. Unfortunately, the state of the art leaves much to be desired in this respect, and the designer must rely on measured noise data. In the past, only a few in-flight measurements of boundary-layer noise over a sig- nificant range of flight conditions have been taken.
The X-15 offers the capability of obtaining noise data over a wide range of dynamic pressure and Mach number that exceeds those available from the flight envelopes of present and proposed aircraft up to Mach 6.
SYMBOLS Cfi incompressible skin-friction coefficient
-
a = 4.177 x 10-7 pounds per
a decibels (dB = 20 loglo &F where
IF
square foot) acceleration due to gravity, feet per second2 g altitude, feet hP 2 rib spacing, inches M Mach number root-mean-square value of pressure fluctuations (sound level), pounds per square foot mean-square pressure fluctuation per unit bandwidth, defined by P2(4
- O D -
p2 = 4 $(cu)dcu, lb2sec/ft4
q dynamic pressure, pounds per square foot Reynolds number based on momentum thickness Re tS skin thickness, inches U local velocity outside boundary layer, feet per second B constant of proportionality 6* boundary-layer displacement thickness, feet cu frequency, radians per second Subscript : co refers to free-stream conditions I INSTRUMENTAT I O N For the boundary-layer-noise experiments on the X-15 airplane, five loca- tions were selected to obtain data. These locations are shown in figure 2 Location 5 was recently chosen for investigation because a (from ref. 2).
large amount of boundary-layer information on heat transfer and skin friction has been obtained at this location as reported in reference 3. As yet, however, no noise data have been obtained here.
A At present, flow measurements have been taken only at location 1.
On the for- sketch of the test panel used for the study is shown in figure 3.
ward end of the panel is a mounting block which holds the microphone and temperature-sensing elements. These temperatures are used for microphone calibration. An accelerometer that was attached to the back of the mounting block was used to verify that no corrections were necessary to the acoustic data because of vibrations. To the rear of the microphone are a boundary-layer rake and a static-pressure orifice to determine local flow conditions.
All noise, acceleration, and pressure data were recorded on an onboard tape recorder, and temperature was recorded on an oscillograph.
The microphones were calibrated for linearity, frequency response, the effect of altitude, steady-state temperatures, and transient temperatures.
These calibrations resulted in an estimated accuracy of rt3 d B over a frequency The range from 50 to 10,000 cycles per second for all flight conditions.
frequency response of the microphones and the results of the transient heating tests are presented in the appendix.
RFSULTS AND DISCUSSION Test conditions for a typical X-15 flight made primarily to obtain boundary-layer noise data are shown in figure 4. The vehicle was launched at an altitude of 45,000 feet, was accelerated to a Mach number of 5 . 4 , and 105,000 feet. The maximum free-stream dynamic attained a maximum altitude of pressure is shown to be approximately 1000 pounds per square foot. It should also be mentioned that during the portion of the flight where Mach number and dynamic pressure are changing slowly, the angle of attack was relatively con- stant.
The purpose of this flight profile was to keep the environment from changing too rapidly.
Data illustrating how the noise varied at different points on the airplane during this flight are presented in figure 5. Time histories of the sound level in decibels are shown only for locations 1 to 4 (indicated in fig. 2).
The times for launch, engine burnout, and Mach 1 during deceleration are marked on the abscissa in figure 5 .
Noise levels for the two midfuselage stations (locations 1 and 2) are illustrated by the solid and dashed lines, for the aft fuselage station (loca- tion 3) by the upper line, and for the base region (location 4) by the lower line. For the most part, there is a spread of only 20 decibels in the levels-- but this spread means that the highest sound levels recorded are ten times the lowest, which illustrates the range of the designers' problems. Note also the
sudden jump in the traces for locations 3 and 4 when the speed brakes were
extended. This result shows that protuberances on the surface of the vehicle can cause large increases in the noise level.
Although the overall sound levels are important, the distribution of the For these sound in frequency bands is also important, as is shown in figure 6.
data, the free-stream Mach number was 3.2 and the dynamic pressure was The sound level is shown in decibels for the over- 380 pounds per square foot.
all noise and for each octave band (center frequencies from 63 to 8000 cps).
Locations 1 and 3, which are on the lower surface of the side fairings, have the highest sound levels, particularly at the higher frequencies. Location 2 on the upper side fairing has its highest noise level predominantly at the
lower frequencies. The spectrum for location 4 on the left blowout panel is
lower in each octave band and is approximately flat.
In order to examine the effect of compressibility on boundary-layer noise, figure 7 shows a comparison of X-15 data with subsonic data obtained by other experimenters (refs. 4 , 3 , and 6). For this comparison the data from location 1 are examined. For the X-13 data, an equivalent-flat-plate incompressible skin-friction coefficient was calculated by using Blasius' formula with the measured-momentum-thickness Reynolds number.
According to reference 4, Kraichnan's analysis of the boundary layer on
a flat plate indicated the proportionality constant B to be in the range from 2 to 12. Wind-tunnel investigations at low speed have given values from
2 . 3 to 4.8. The X-15 data indicate values from 2 to 4 . 6 and show a tendency
to decrease slightly with increasing Mach number. (Because this is an equivalent-flat-plate comparison, only X-13 data which closely approximate these conditions are shown.) The results indicate that the general levels of this dimensionless coefficient, at this location, are changed very little by effects of compressibility.
Since the comparison between data from wind tunnel at subsonic speeds and from flights at high supersonic speeds shown in this figure deals only with the overall sound pressure, it is of interest to compare the frequency content.
Willmarth, Serafini, and others (refs. 4 to 7 ) have shown that the noise power
in the frequency bands can be normalized in terms of the boundary-layer displacement thickness and local free-stream velocity. This procedure has been applied to some of the X-13 data. The results, presented in figure 8 , show that the frequency content of the noise measured in flight is signifi- cantly different from the wind-tunnel noise. This difference lies in how the energy is distributed over the frequency range. The flight data have more energy at the low frequencies and less energy at the high frequencies than the wind-tunnel data. Even though the data are for two widely different Mach numbers, it cannot be concluded at this time that the difference can be attri- buted entirely to Mach number, since it is known that the X-15 surface has considerable roughness. In boundary-layer flow, the effects of surface rough- ness, typical of flight vehicles, perturbs the boundary-layer pressure fluctua- tions in the lower frequencies even at subsonic speeds.
The results in figures 7 and 8 , although somewhat preliminary, show that
the theory of Kraichnan (as given in ref. 4), supported by wind-tunnel measurements at subsonic speeds, can be used to estimate the total noise expected on a hypersonic flight vehicle but it does not give a true picture of the energy distribution that is important for design.
CONCLUDING REMARKS Flight results from boundary-layer noise measurements at specific loca- tions on the X-15 have been presented and a limited comparison with analytical equivalent flat-plate results made. These results indicate that more exten- sive flight measurements are required to establish a guide for the theoretical studies and laboratory tests that are necessary to define the characteristics of boundary-layer noise.
Continued effort in the X-15 boundary-layer-noise program is necessary to provide for the extension of the range of the present data, as well as to permit a comprehensive analysis of the noise at other locations, particularly on the vertical tail where data can be obtained that can be favorably compared with flat-plate experiments of other investigations.
Flight Research Center National Aeronautics and Space Administration Edwards, Calif ., October 7, 1965.
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BOUNDARY-LAYER-NOISE INSTRUMENTATIOIV The instrumentation panel used f o r obtaining surface-pressure fluctuations and local-flow measurements a t location 1 i s shown i n figure 3. The noise data and the local-flow data were recorded i n t h e vehicle on a miniature airborne Noise data w e r e tape recorder using 1.5 m i l , 1-inch tape a t 15 in./sec.
recorded on d i r e c t record, and the flow data were FM multiplexed. The temper- a t u r e data were recorded on an oscillograph.
A l/2-inch Photocon microphone (model 504), of the condenser type with a diaphragm recessed below a perforated cap, w a s used. The microphone w a s mounted i n a heavy mounting block with the cap f l u s h t o the surface of the vehicle. This method of mounting r e s u l t s i n an orientation such t h a t the microphone diaphragm i s generally p a r a l l e l t o the t h r u s t a x i s of the vehicle for locations 1, 2, and 3. The microphone diaphragm f o r location 4 w a s per- pendicular t o the t h r u s t axis. The response of the microphones a t high a l t i t u d e was determined by an e l e c t r o s t a t i c actuator i n a vacuum chamber. The r e s u l t s of t h i s t e s t , obtained by methods presented i n reference 8 , are shown i n figure 9.
I n order t o assess the vibration s e n s i t i v i t y of t h e microphones, they were placed i n an electromagnetic shaker and vibrated perpendicular t o the diaphragm from 0.23g t o 3g i n O.23g steps over t h e e n t i r e response range. The output f o r a vibration t e s t a t 3g i s shown i n the l e f t p l o t of f i g u r e 10, and the output a t 500 cps over the e n t i r e range of vibration i s shown i n t h e r i g h t p l o t .
The e f f e c t of temperature on the microphones has a l s o been determined.
The response change of the microphones a f t e r being soaked f o r 30 minutes at constant temperatures i s shown i n figure 11. To determine the e f f e c t of t r a n s i e n t heating, t h e microphones were a l s o heated using a temperature time history, taken from data recorded i n f l i g h t on a thermocouple located d i r e c t l y below the perforated cap of the dummy microphone. Using a 1000-cps, 120-dB sound source ( 3 0 dB above ambient and without anechoic chamber), t h e response w a s found t o d r i f t l e s s than -+1 dB a s the temperature changed from -40" F t o The dummy-microphone thermocouples were 320' F, as i s shown i n figure 12.
used t o insure the same heating r a t e s as recorded i n f l i g h t during t h i s t e s t .
It can be seen t h a t t r a n s i e n t heating causes a smaller change at the 1000-cps frequency than t h e constant temperature frequency response shows.
This difference i s due t o expansion e f f e c t s on the diaphragm while heat i s applied. It i s believed t h a t a t r a n s i e n t heating response best applies t o these X - l 5 data, thus no correction w a s applied t o t h e data presented.
However, the noted accuracy of 23 dB includes t h e e f f e c t of t r a n s i e n t heating.
1. Hubbard, Harvey H.; Edge, Philip M., Jr.; and Modlin, Clarence T., Jr. : Design Considerations for Minimizing Acoustical Fatigue in Aircraft Structures. WADC-University of Minnesota Conference on Acoustical Fatigue, W. J. Trapp and D. M. Forney, Jr., eds., WADC Tech Rep. 59-676, U.S. Air Force, March 1961, pp. 321-338.
2. Kordes, Eldon E.; and Tanner, Carole S.: Preliminary Results of Boundary Layer Noise Measured on the X-15 Airplane. Acoustical Fatigue in Aero- space Structures. Walter J. Trapp and Donald M. Forney, Jr., eds., Syracuse University Press, May 1964, pp. 85-96.
3 . Banner, Richard D.; and Kuhl, Albert E.: A Summary of X-15 Heat-Transfer and Skin-Friction Measurements. NASA TM X-1210, 1966.
4. Serafini, John S.: Wall-Pressure Fluctuations and Pressure-Velocity Correlations in a Turbulent Boundary Layer. NASA TR R-165, 1963.
5. Willmarth, William W.: Wall Pressure Fluctuations in a Turbulent Boundary Layer. NACA TN 4139, 1958.
Space-Time Correlations and Spectra of Wall Pressure in 6. Willmarth, W. W.: a Turbulent Boundary Layer. NASA MEMO 3-17-59WJ 1959.
7. Lilley, G. M.; and Hodgson, T. H . : On Surface Pressure Fluctuations in Turbulent Boundary Layers. AGARD Rep. 276, April 1960.
8. Beranek, Leo L.: Acoustic Measurements. John Wiley & Sons, Inc., July, 1962.
EFFECT O F N O I S E ON S K I N T H I C K N E S S
E X A M P L E F O R TYPICAL SKIN-STRINGER C O N S T R U C T I O N .06 .05 .04 t,, in.
.03 / +S .02 J I
. 01
110 11 5 120 125 130 NOISE LEVEL, dB Figure 1
MICROPHONE LOCATIONS ON THE X-15
Figure 2
BOUNDARY-LAYER-NOISE TEST PANELS
LOCATION 1 (RIGHT SIDE FAIRING) B O U N D A R Y - L A Y E R R A K E S T A T I C - P R E S S U R E M I C R O P H O N E Figure 3 TYPICAL FLIGHT CONDITIONS FOR MEASUREMENT OF X-15 BOUNDARY-LAYER NOISE 1200 I I 6 1 150X103 0 LAUNCH 100 200 300 400 506 TIME, sec Figure 4 BOUNDARY-LAYER-NOISE DATA FOR FOUR LOCATIONS LOCATION OVERALL 130 SOUND- PRESSURE - id; j -SPEED BRAKES LEVEL, dB 120 I / EXTENDED 'I 'lo LAUNCH BURNOUT M = l 1001 4 1 4 I I I 1 0 100 200 300 4 0 0 500 TIME, rec Figure 5 SOUND SPECTRUM FOR FOUR LOCATIONS M, = 5.2; q, = 380 psf LOCATION 0 1 0 2 1401 0 SOUND- 130 PRESSURE LEVEL, dB 120 A
110 I
A I I I I I I I I 100' ^" "63 125 2 5 0 500 1000 2000 4000 aooo OVERALL OCTAVE-BAND CENTER FREQUENCY, CPS Figure 6 1 0 EFFECT O F COMPRESSIBILITY ON OVERALL SOUND-PRESSURE LEVEL WIND TUNNEL FLIGHT 0 SERAFlNl O X-15 LOCATION 1 0 HARRISON , 0 2 5 6 A LILLEY AND HODGSON d WILLMARTH 0 1 2 3 4 5 6 LOCAL MACH NUMBER Figure 7 MEAN-SQUARE SPECTRA OF BOUNDARY-LAYER PRESSURE FLUCTUATIONS (X-15, 4 < M C 5 )
&
0 00
I I I I -70 I
.01 .1 1.0 10 .001 Figure 8 1 1
I I1 I 1 1 1 1 1 1 1 1 1 1 1 1 l l l llll11l
MICROPHONE-RESPONSE CHANGE AT VARIOUS ALTITUDES 8.4 mm Hg (99,500 ft) 4 0 21 mm Hg 3 0 5 4 mm Hg (60,100 ft) MICROPHONE 2o RESPONSE, d B 141 mm Hg (40,100 ft) 350 mm Hg (20,000 ft) I I I I -10 1 10’ 102 103 104 1 0 5 FREQUENCY, c p s Figure 9 MICROPHONE VIBRATION TESTS 500 cps
OUTPUT, dB l 1 : : [ l o o , ; 3 g ; ; , ~, , - 7rT,:j/
10’ 102 103 5 X 1 O 3 0 1 2 3 FREQUENCY, cpr VIBRATION LEVEL, g Figure 10
RESPONSE CHANGE AT
VARIOUS CONSTANT TEMPERATURES
l o r
- 6 0 " F
OUTPUT' -1 O 0 t
0" F dB 70" F
9 200" , F
1 0 ' 102 103 104 2x104 FREQUENCY, c p s Figure 11 R E S P O N S E CHANGE DURING TESTS TRANSIENT HEATING OUTPUT, dB -1 -2
TEMPERATURE, :!!
O F -100 I I I 0 100 200 300 T I M E , sec Figure 12 NASA-Langley, 1966 H-418 “ T h e aeronautical and space activities of the United States shall be conducted 50 as to cotitribrite . , . to the expansion of humatz Rtzowl- edge of phenometza i n the atmosphere and space. T h e Administration shall provide for the widest practicable aizd appropriate dissemination of information coizcertziiig its activities aiid the resdts thereof .” -NATIONAL AERONAUTICS AND SPACE ACT OF 1758
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