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ASPECTS OF INVESTIGATING STOL NOISE USING, LARGE-SCALE

WIND-TUNNEL MODELS

Michael D. Falarski, David G. Koenig, and Paul T. Soderman

Ames Research Center

and

U.S. Army Air Mobility R&D LaboratoryMoffett Field, Ca. 94035

(NLASA-TM-X-62164') ASPECTS OF INVESTIGATINGSTOL NOISE.- USING LARGE SCALE WIND TUNNEL.MODELS M.DF.: .Falarski, %et al (NASA) E Jun.1972 35 p CSCL 01B

June 1972

-ieproduced by ' , i -

ji ATIONAL TECHNICAL.'* IlNFORMATION SERVICE

U S Department of CommerceSpringfield VA 22151

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https://ntrs.nasa.gov/search.jsp?R=19720018358 2020-03-28T23:47:35+00:00Z

ASPECTS OF INVESTIGATING STOL NOISE USING LARGE-SCALE WIND-TUNNEL MODELS

Michael D. Falarski, David G. Koenig and Paul T. Soderman

Ames Research Centerand

U.S. Army Air Mobility Research & Development Laboratory

SUMMARY

The applicability of the NASA Ames 40- by 80-Foot Wind Tunnel for",,acoustic research on STOL concepts has been investigated. The acoustic

characteristics of the wind tunnel test section has been studied with

calibrated acoustic sources. Acoustic characteristics of several large-scale STOL models have been studied both in the free-field and wind tunnel

acoustic environments. The results of these studies indicate that theacoustic characteristics of large-scale STOL models can be measured in

the wind tunnel if the test section acoustic environment and model acoustic

similitude are taken into consideration. The reverberant field of the

test section must be determined with an acoustically similar noise source.Directional microphone and extrapolation of near-field data to far-field

are some of the techniques being explored as possible solutions to thedirectivity loss in a reverberant field. The model sound pressure levels

must be of sufficient magnitude to be discernable from the wind tunnelbackground noise.

etaiAi o ilus~rations inthis d0cum nf may be better

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-2 -

INTRODUCTION

Acceptance of a commercial STOL aircraft will be very dependent onits ability to operate at low noise levels. The primary reason for thisis the close proximity of STOL (short takeoff and landing) ports toareas of high population density. The proposed 95 EDNdB noise limit at152 m (500 ft) from runway centerline is substantially lower than theexisting FAR 36 limits on commercial CTOL aircraft. As the concern overnoise pollution grows and STOL operation becomes more widespread, it isquite possible that the acceptable noise level will be reduced evenfurther. Extensive research will have to be conducted to achieve noisereductions throughout the STOL flight regime.

The staff of the Ames 40- by 80-Foot Wind Tunnel has conductedextensive research in STOL aerodynamics (see reference 1). Because ofthe wind tunnels large size (see figure 1), the models tested incorporateflight-worthy propulsion systems. Aside from the aerodynamic systembenefits of this capability, the large-scale models provide realisticacoustic sources. With this equipment available the desirability ofacoustic investigation is evident. The development of the wind tunnelas an acoustic research facility permits the analysis of fly over noiselong before an aircraft is available. This paper will present theresults of work to develop the Ames 40- by 80-Foot Wind Tunnel into an '

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acoustic measurement range.

- 3-

NOTATION

OASPL overall sound pressure level, dB ref. .0002 microbar

PNL perceived noise level, PNdB

PTAUG augmentor total pressure, cm. hg(in. hg)

q free-stream dynamic pressure, n/m2(psf)

SPL sound pressure level, dB ref. .0002 microbar

Vi fan exit jet velocity, m/sec (ft/sec)

V. free-stream velocity, m/sec (ft/sec)

6f flap deflection, deg

Ac/4 wing sweep at the quarter chord, deg

-4-

ACOUSTIC MEASUREMENTS IN THE WIND TUNNEL

The acoustic characteristics of the wind tunnel were investigatedwith several acoustic sources (see reference 2). These studies haveshown that reverberations, background noise levels, and standing wavespresent problems in measuring free-field noise in the wind tunnel.These problems, as well as possible solutions, will be discussed inthis section.

Test Section Reverberation

The acoustic studies have shown the test section to be quitereverberant. This can be seen in the rise above free-field levels ofthe test section measurements as the microphone is moved away from thesource (see figure 2). The test section Hall Radius, which is thedistance from the source where the reverberant field and the directfield are equal, is approximately 4.6 m (15 ft). Because of the modelshape, location, and size, it is usually not possible to place micro-phones within the Hall Radius. Therefore, to determine free-field noiselevels the data must be corrected for the reverberation. This requirescalibration of the wind tunnel with an acoustic source which is similarto the acoustics of the model.

Acoustic directivity.- The major difficulty a reverberant fieldpresents is the masking of directivity. This can be handled by timeconsuming calibration procedures for broad-band noise. However, othersolutions to this problem which are being explored include a directionalmicrophone, a phased array of microphones, and calculation of far-fieldlevels from near-field measurements. An example of the type of direc-tional microphones being investigated is the porous pipe microphone shownin figure 3. The porous pipe allows weak coupling between the acousticpressures inside and outside the pipe. A ½ inch omnidirectional micro-phone is located at the end of the porous pipe. The solid cone insidethe pipe produces an anechoic termination of all waves except thosetraveling towards the microphone. In the frequency range for which it isdesigned the microphone increases the Hall Radius of the wind tunnel to9.15 m (30 ft). Its use is limited at present because it must be orientedparallel to the wind stream to avoid high wind noise measured whenoriented in any other plane (see figure 4). This flow noise can possiblybe reduced by designing the porous pipe in the form of a porous stripconformal with the upper and lower surface of an airfoil. The otherlimiting factor is the microphone length. Its directivity is a function

-5-

of its length and to achieve reasonable directivity at 500 Hz, themicrophone must be 1.22 m (4 ft) long.

Another possible method of measuring directivity in a reverberantfield is an electronically phased array of microphones. The microphonescan be arranged in either an end-on array or a broadside array. The.end-on array consists of 4 or 5 microphones spaced along a line radiatingfrom the noise source. The broadside array has a row of 4 or 5 micro-phones spaced along a line parallel to the source. In either case thesignals from each microphone are electronically delayed such that thedirect signal is reinforced while the wave fronts from reverberation orbackground sources are cancelled. The frequency response and directivityof the array are a function of the delay time, microphone spacing, andsystem rigidity.

By taking measurements in the near-field the direct radiated signalis made much larger than the reverberant signal. Far-field levels com-puted from these near-field measurements offer another solution to thewind tunnel reverberant field problem. This approach is difficult toimplement experimentally because of the large number and accuracies ofthe measurements required.- Analyses and experiments are underway todefine the experimental problem and devise means of practicalimplementation.

Acoustic power.- The investigation of the test section inlet anddiffuser acoustic environment with a source in the test section, indicatesthe reverberant field in these two areas is fairly diffuse and that theacoustic energy tends to flow out of the test section. This means thatthe sound waves are traveling equally in all directions except toward thesource and the sound pressure for a given cross section does not varyfrom point to point due to reflection or interference. This allows themeasurement of the total sound power radiated in the test section to beestimated from single measurements of the sound pressure in the inletand diffuser of the wind tunnel.

Wind Tunnel Background Noise

Operation of the wind tunnel creates a background noise floor, whichis established by the noise of the tunnel drive system and wind noise onthe microphones. A frequency spectrum of this noise (see figure 5) showsit to be fairly low frequency and dominated by the fan tones and broad-band noise at q > 958 n/m2 (20 psf). At q < 958 n/m2 (20 psf) the fanmotor-generator noise is no longer negligible. It consists of three tonesat 125, 250, and 500 Hz, which do not vary in frequency with tunnel speed.

-6-

The overall level of the noise floor increases with forward speed as isshown in figure 6. The acoustic source being investigated must besufficiently loud and have a frequency spectrum such that its soundlevels are discernible from the background noise. At present this isnot a significant problem, but it will be in the future when acousti-cally treated models having lower noise levels are studied. Because ofthe large size of the test section and the dominant low frequencies,any significant reduction in the noise floor with acoustic treatment ofthe wind tunnel is quite difficult. The solutions for the reverberantfield problem are also directly applicable to the background noiseproblem in that they allow discrimination between the noise sources inthe test section.

Placing a microphone in a moving air stream creates local turbulencewhich the microphone senses as radiated acoustic pressures. This con-tributes a high frequency sound to the background noise floor. Whilethis is not a problem with existing models, future quiet propulsion sys-tems will approach this noise floor. At present aerodynamically shapednose cones, which substantially diminish the response to turbulencewhile having little effect on the microphones acoustic properties, aremounted on microphones used in the wind tunnel. Different types offairings are being studied experimentally to reduce this wind noise evenfurther.

Acoustic Focusing

Reverberant enclosures such as the wind tunnel test section willreflect and focus sound, creating standing waves. To study thisphenomenon an optical model of the wind tunnel test section was con-structed. The internal surface of the model was covered with a highlyreflective foil. A small light source was placed in the center tosimulate an acoustic source. The results of the experiment are shownin figure 7. Planes and areas of acoustic focusing were found in themodel. These are areas in the wind tunnel which should be avoided.Because this experiment simulated sources of very high frequency orsound after several reflections, further studies must be conducted withmore representative acoustic sources to completely understand thisproblem.

-7-

Other Considerations

All extraneous noise sources on the models must be suppressed iftheir levels are less than 10 dB below the source of interest and theirfrequency spectrum is coincident with the model source of interest.This is the case with an internally-blown flap (IBF) model such as theaugmentor wing, where the high pressure air is provided by an internalair compressor which masks the noise of the augmentor. During a recenttest of the augmentor wing model, the compressor inlet noise was sup-pressed with acoustically lined inlets (see figure 8). In an externally-blown flap (EBF) model a turbofan engine, which provides a good aerodynamicsimulation, may not be acceptable for an acoustic simulation. The internalengine noise is very much a function of the engine design. It may have tobe suppressed with an acoustically treated tail pipe to allow investigationof jet scrubbing noise.

As discussed previously, the model noise must be of sufficient am-plitude to be discernable from the wind tunnel background noise. Thiscan be accomplished by mounting the microphones closer to the model,although not close enough to be in the acoustic near-field.

TESTS

To gain more understanding of the wind tunnelacoustic environment,several models were investigated both in the free-field and in the windtunnel. The configurations listed below will be discussed (the aero-dynamic and acoustic characteristics of the models are presented inreferences 3 thru 8):

1. Lift fan semispan wing

2. OV-10 rotating cylinder flap aircraft

3. Externally-blown jet flap semispan wing

4. Swept augmentor wing-internal flow jet flap model

For the first three tests, wind tunnel reverberation corrections werederived from the point source calibration of the test section reportedin reference 2. For all the tests,½ inch B. and K. omnidirectionalcondenser microphones, equipped with windshield nose cones, were used

to measure the sound pressure levels.

-8-

RESULTS AND DISCUSSION

Lift Fan Semispan Wing

This configuration was a swept, tapered semispan wing with asingle, .93 m (3 ft) diameter, 1.3 pressure ratio, lift fan mounted inthe wing. The model installed in the wind tunnel is shown in figure 9.The geometric details of the wing are presented in figure 10. The fanwas tip driven by the exhaust gas of a J-85 turbo jet with an acousti-cally treated inlet to suppress the engine compressor noise. A free-field investigation was made with the isolated fan. The results ofthis study were corrected for ground and atmospheric attenuation. Ascan be seen in figure 11, the sound pressure level and directivity atthe blade passage frequency for the free-field and wind tunnel studiesshow good agreement except for the 200 near the plane of the fan. Thelower noise in that direction is contrary to expectations because thereverberant field would increase noise; however, the wind tunnel modelhad a relatively large chord surface in the plane of the fan which mayhave reduced the inlet to-exhaust noise propagation.

OV-1OA Rotating Cylinder Flap STOL Aircraft

An OV-1OA aircraft has been modified to incorporate a rotatingcylinder flap. Geometric details of the flight test vehicle and theflap are presented in figure 12. The rotating cylinder located at thewing-flap junction induces flow over the flap, creating high circulationlift. The aircraft noise was measured both in actual flyover tests aswell as in the wind tunnel. The wind tunnel data was corrected forreverberation and the flyover data was adjusted for ground reflectionand ground and atmospheric attenuation. The one-third octave bandfrequency-spectrums of the two studies are compared in Figure 13. Theagreement is very good, although the wind tunnel levels tend to beslightly lower than the flight results. There is also no tone at the100 Hz blade passage frequency in the flight data. This may be due to

the 1/8 of a second sampling time for the flight test data analysis.

-9-

Externally-Blown Jet Flap Semispan Wing

The externally-blown jet flap model is shown installed in thewind tunnel in figure 14. The microphones were mounted in the planeof the fan on a 6.1 m (20 ft) semicircle. The geometric details ofthe model are also presented in figure 14. The .76 m (30 in) diameter,1.03 pressure ratio, ducted fan is mounted below the wing so that itsslipstream impinges on the large chord flap system. Because the modelcould not be tested in free-field, it was installed in a large, enclosedpreparation area. The acoustic response of this area was calibratedusing the dodecahedron acoustic source described in reference 2. Itproved to be much less reverberant at higher frequencies than the windtunnel, as can be seen in figure 15. The results of the two independenttests show good agreement. As shown in figure 16, the perceived noiselevels agree within 1 PNdB and there is no significant loss in directivity.The sound frequency spectrums compare favorably above 500 Hz (see figure17). Below 500 Hz the wind tunnel sound pressure levels are 2-3 dB abovethose of the static test. This is only. partially due to the wind tunnelbackground noise.

Swept Augmentor Wing

The augmentor flap is a form of internally-blown flap incorporatingan ejector system. Through a cooperative program with the CanadianDefense Research Council and DeHavilland Aircraft of Canada, Ltd., theaerodynamic and acoustic characteristics of a swept augmentor wing modelare being investigated. The model is shown installed in the wind tunnelin figure 18. The augmentor geometry is described in figure 19.

For the acoustic investigation, the inlets of the air compressorwhich provides the high pressure primary air and the tail pipes thatexhausted surplus air from the compressor were equipped with soundsuppressors. This eliminated the inlet tones and reduced the frequencyof the tail pipe noise, thus making the augmentor flap the dominantacoustic source.

The augmentor wing model is a much more distributed acoustic sourcethan any of the previously discussed models, and there was some reserva-tion about the applicability of the previously described point sourcecalibrations. Therefore, reverberation corrections were derived from acomparison of the free-field and wind tunnel data for the take-off flap

- 10 -

setting at several power settings. As indicated in figure 20, there isa significant difference between the two results, especially in the highfrequency range. Above 500 Hz the point source shows a correction of6 dB, while the actual data would indicate little or no correction. Also,the omnidirectional point source calibration indicates that, for the samedistance from the source, the reverberation is not a function of locationin the test section. This is not the case with the augmentor wing, whichshows large variations in reverberation in the middle frequency range withlocation.

The reverberation corrections derived from the test comparison wereapplied to the other flap and power settings. Deflecting the flap to thelanding configuration changes the acoustic directivity, but the correc-tions are still applicable as indicated by the close agreement in figure21. This is also true of the frequency spectrum presented in figure 22.

This data was scaled to a 91,000 Kilogram (200,000 pound) grossweight aircraft. This can be done simply and accurately by operatingthe model at full scale values of jet flow (total pressure and temperature).The acoustic scale factor is then the square root of the nozzle area ratio.The full scale distance is the product of actual measurement distance andthe scale factor and the frequency is reduced by the scale factor.

The results of the augmentor wing investigation leads to the con-clusion that the point source calibrations are not adequate for correctingacoustic measurements from a large diffused noise source. An investigationof a 3-dimensional externally-blown jet flap model is planned for the nearfuture. This will also be tested in free-field and the wind tunnel. Thereverberation field determined from this test will be compared with theaugmentor wing results. From this comparison it is hoped that some generalguidelines for formulating reverberant field corrections can be determined.

- 11 -

CONCLUDING REMARKS

The results of several model investigations indicate that theacoustic characteristics of STOL models can be measured in the Ames40- by 80-Foot Wind Tunnel if the test section acoustic environment andmodel acoustic similitude are taken into consideration. The test sec-tion reverberant field must be determined with an acoustically similarnoise source. Directional microphones, phased microphone arrays, andthe calculations of far-field noise from the far-field measurements arebeing explored as possible solutions to the problem of measuring direc-tional noise sources in a reverberant field. The model source pressurelevels must be of sufficient magnitude to be discernible from the windtunnel background noise.

- 12 -

REFERENCES

1. Anon: Conference on V/STOL and STOL Aircraft. NASA SP-116, 1966.

2. Bies, David Alan: Investigation of Making Model AcousticMeasurements in the NASA Ames 40- by 80-Foot Wind Tunnel.NASA CR-114352, 1971.

3. Kirk, Jerry V., Hall, Leo P., and Hodder, Brent K.: Aerodynamicsof Lift Fan V/STOL Aircraft. NASA TMX-62086, 1971.

4. Kazin, S. B., and Volk, L. J.: LF-336 Lift' Fan Modification andAcoustic Test Program. NASA CR-1934, 1971.

5. Weiberg, James A., and Gamse, Berl: Large-Scale Wind Tunnel Testsof an Airplane Model with Two Propellers and Rotating CylinderFlaps. NASA TND-4489, 1968.

6. Falarski, Michael D.,-and Aiken, Thomas N.: Large-Scale Wind-Tunnel Investigation of a Semispan Wing Equipped with anExternally-Blown Jet Flap. NASA TMX-62079, 1971.

7. Falarski, Michael D.: Large-Scale Wind-Tunnel Investigation ofthe Noise Characteristics of a Semispan Wing Equipped with anExternally-Blown Jet Flap. NASA TMX-621

8. Falarski, Michael D., and Koenig, David G.: AerodynamicCharacteristics of a Large-Scale Model with a Swept Wing andAugmented Jet Flap. NASA TMX-62029, 1971.

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