A preliminary comparison between the SR-6 propeller noise in flight and in a wind tunnel

A/ A,. z _ f ,s _- 2, 3 _ 9/

NASA Technical Memorandum 83341

N A SA-TM-83341 19830016016

+ A Preliminary Comparis o n Between the

SR-6 Pr o peller N o ise in Flight

and in a Wind Tunnel

Jams H . D itt mar

L ewis R es earch C enter

Cl ev elan d, Ohio

and

Paul L . Lasagna and Karen G. Mackall

Hugh L . Dryden Flight R es earch C ent er

Edwards, California

P rep a red for the

! One-hundred - fifth Meet i ng of the Acoustical Soc i ety of Amer i ca

' Cincinnati, Oh i o, Ma y 9-13 , 1983

:-,'a.:,. : 5 . , : '. . _.,- t d? U .-, UENTEt{ +,_,- , _ r ,v NASA L , --,_, 1_t I _:_- " 't, ' , VIRGINIA "\ A PRELIMINARY COMPARISON BETWEEN THE SR-6 PROPELLER NOISE IN

FLIGHT AND IN A WIND TUNNEL

James H. Dittmar National Aeronautics and Space Administration Lewis Research Center Cleveland, Ohio and Paul L. Lasagna and Karen G. Mackall Hugh L. Dryden Flight Research Center Edwards, California SUMMARY High speed turboprops offer an attractive candidate for future aircraft because of their high propulsive efficiency. However, one of the possible problems associated with these propellers is their high noise level at cruise condition that may create a cabin environment problem. Models of these pro- pellers have been tested for acoustics in the Lewis 8-by-6-foot wind tunnel and on the Dryden JetStar airplane. This paper shows comparisons between the airplane and wind tunnel data for the SR-6 propeller.

The comparison of maximum blade passing tone variation with helical tip Mach number between the tunnel and flight data was good when corrected to the same test conditions. Directivity comparisons also showed fairly good agree- ment. These good comparisons indicate that the wind tunnel is a viable location for measuring the blase passage tone noise of these propellers.

INTRODUCTION One of the possible propulsive systems for a future energy-efficient airplane is a high-tip-speed turboprop. When the turboprop airplane is at cruise, the combination of airplane forward speed and the propeller rotational speed results in supersonic velocities over the outer portions of the pro- peller blades which may create a cabin noise problem. Models of this type of propeller have been previously tested for noise in the NASA Lewis 8-by-6-foot wind tunnel (refs. 1 to 4). This wind tunnel does not have acoustic damping material on its walls and there has been a concern that this lack of acoustic material may have compromised the noise data.

As part of the program to evaluate the noise of these propellers the NASA Dryden JetStar airplane was modified to test them in flight. A previous com- parison between wind tunnel and flight data, for the SR-3 propeller, showed good agreement in the maximum blade passing tone variation with helical tip Mach number (refs. 5 and 6). Another propeller, SR-6, has now been tested both in a wind tunnel, reference 4, and in flight. The intent of this paper is to make a preliminary comparison between the SR-6 propeller noise measured on the JetStar airplane and that previously obtained in the 8- by 6-foot wind tunnel.

APPARATUS AND PROCEDURE The SR-6 propeller model used in these noise comparisons has ten blades with a 40 degree tip sweep and is nominally 0.696 meter (27.4 in.) in dia- I .

meter. Table I shows some of its design characteristics and more information • is availablein reference7. A picture of the propeller in the 8- by 6-foot wind tunnel can be seen in figure l(a) and on the JetStar airplane in figure l(b).

To measure the propellernoise, pressure transducerswere installedin the wind tunnel bleed holes visible in figure l(a). The locationsof these transducersare shown in figure 2(a). The positionsare identifiedas A thru E, as in reference 3, and are locatedon the top wall of the wind tunnel which is approximatelyone and one-half propellerdiametersabove the propeller tip. On the JetStar airplane,microphoneswere installedflush in the air- plane fuselage which is approximatelyeight-tenths of a propellerdiameter from the propellertip. The propeller axis is tipped three degrees downward on the airplane to align it with the flow. The microphoneson the airplane were installedalong a line directly underneaththe propellercenterline. The locationsof the microphonesreported herein are numbered 1 thru 9 as can be seen in figure 2(b).

The intent of this paper was to comparedata taken at the same test points in the wind tunnel and in flight but difficultiesin the blade setting angle mechanism for the wind tunnel tests (ref.4) and drive system limita- tions on the airplaneforced somewhat differenttest conditions. Experiments in the wind tunnel were performedwith blade settingangles (62° and 64°) near the design blade setting angle of the propeller (63") and at blade setting angles (59° and 61°) near to the blade setting angle (60 °) which would give performanceclose to the design performanceof the previouslytested SR-3 propeller.

The airplanetests were performedat 62.4 and 58.9 degrees. The intent was to test the propellersat constant advanceratios J (see appendix)at different axial Mach numbers. However,for the airplanetests, an available power limitationpreventedthe tests from being performedat constant advance ratio and at the higher Mach numbers this resulted in higher advanceratios than desired.

The signals from the pressure transducersin the tunnel and microphones on the airplanewere recordedon magnetic tape and narrowbandspectra from 0 to 10,000 Hz, with a bandwidthof approximately26 Hz, were taken for each of the test points. The blade passing tone levelwas read from each of these spectra and is presented,along with the test conditions,in table II for the tunnel data and in table Ill for the airplanedata.

RESULTSAND DISCUSSIONS In order to make comparisonplots betweenthe wind tunnel and airplane data it is first necessaryto bring the two sets of data to the same experi- mental conditions. There are two primarydifferences: atmosphericconditions and geometry. The airplanetests were performedat 9.1 km (30,000ft) where the air is less dense than in the wind tunnel tests. As shown in reference8, to correct the wind tunnel data to flight conditions,the wind tunnel sound pressure levels are changed by 20 times the log of the ratio of the atmos- pheric pressure in flight to that in the tunnel. Since at each axial Mach number in the tunnel the pressure is different,this correctionresults in reductionsin the tunnel noise of 9.51 dB at an axial Mach number of 0.6, 9.02 dB at 0.7, 8.56 dB at 0.75, 7.97 dB at 0.8 and 7.45 dB at 0.85. (These correctionsare the same as those used previouslyin ref. 5).

The geometricconditionsare also differentbetween the tunnel and flight tests with the propellerbeing closer to the airplanefuselage than to the wind tunnel wall. Because, in both cases, the measurementlocationsare so close to the source,the noise does not necessarilymeet far-fieldcriteria.

Therefore,the proper correctionfor distance is somewhatuncertain. Refer- ence 9 has suggestedthat a distance correctionof 15 times the log of the distance ratio be used in this near field insteadof the normal far-field correctionof 20 times the log. The 15 log correctionis used here and was used previouslyin reference5. It should be noted, however, that, because of the small distance,the standard20 log correctiongives a value less than one decibel differentand its use would not materiallyalter the comparison.

Taking the distancefrom the measurementlocationto the tangent point on the propellertip circle gives a distance correctionof 3.2 dB. (This is a slightly differentcorrectionfrom that used for SR-3 in reference5 because of the differentdiametersof the two propellers.) No correctionwas made for the slightly differentdistancesto each microphonelocation that result from the three degree tilt of the propeller. The combinationof altitude and dis- tance correctionsreduces the wind tunnel data by 6.3 dB at an axial Mach number of 0.6, 5.8 dB at 0.70, 5.4 dB at 0.75, 4.8 dB at 0.80 and 4.3 dB at 0.85. When these correctionsare applied to the tunnel data (table II) the data are therebycorrectedto flight conditionsand presentedin table IV.

Variationwith Helical Tip Mach Number The maximum measured blade passing tone levels on the airplane fuselage and on the tunnel wall, correctedto flight, are plotted as a function of helical tip Mach number,MH, (vectorsum of axial and rotationalMach num- bers) in figure 3. Figure 3(a) is for the SR-6 propelleroperated near its design blade setting angle and figure 3(b) is for the SR-6 propelleroperated near its design blade setting angle which would give the same performance (j and CD) as the SR-3 design. In general the comparisonsbetween the wind tunnel and flight data are very good. The slightly lower sound pressure levelsof the airplane data in figure 3(a) are probably the result of the higher advance ratios for the airplane tests necessitatedby the power limita- tions. These good comparisonsindicatethat the wind tunnel is a viable loca- tion for determiningthe maximum blade passage tone levels of these types of propellers.

Directivity In the previous comparisonbetweenthe wind tunnel and flight data, reference5, the directivitiesat the lower Mach numbers agreed well, but at the higher Mach numbers the wind tunnel data fell off faster toward the front than did the airplanedata. These directivityplots for the SR-3 propeller from reference5 are repeated here in figure 4. Hanson, reference10, has suggestedthat significantreductionsin the forwardradiated noise measured at the wall may be caused by wall boundary layer refraction. The SR-3 noise at and behind the peak angle was not affected by this boundary layer refrac- tion. The amount of the refractionincreaseswith increasingMach number, increasingboundary layer thicknessand as the measuring positionmoves for- ward. Reference5 indicatedthat the more rapid forward falloff of the directivityin the wind tunnel may have been caused by a thicker boundary layer in the tunnel than on the airplane. The possibilityof boundary layer refractionprompted an investigationby the airplane test personnelof the boundary layer thicknesson the airplane. The data, taken with a 5 and an 8 inch rake, are shown in figure 5. The shape of this boundary layer profile is not typical and the bulge around a Y of 5 cm (2 in.) is indicativeof an energizationof the boundary layer which results in a thinner boundary layer.

The airplanewindshieldwipers and supportswere discoveredto be the source of the energizationand they were removedfrom the airplane. This resulted in the more typical boundary layer profile shown in figure 6. This thicker boundary layer was present on the airplane during the SR-6 tests since the wipers were removed.

Figure 7 shows the directivitiesof the blade passing tone obtainedwith the SR-6 propellernear its design conditionson the airplane and in the wind tunnel. As can be seen the directivitiescompare fairly well. In particular the large differencesin noise fall off toward the front which were observed in the comparisonof the SR-3 propellernoise in flight and in the tunnel at M = 0.75 and 0.80 (figs.4(c) and (d)) do not seem to occur here. At M = 0.75, for the SR-6 propeller (fig. 7(c)), the curves are at different levels because of the differentadvanceratios and consequentlydifferent helical tip Mach numbers of the test (see fig. 3(a)). They are almost the same curve displacedonly in level and have similarforward falloff in noise for both flight and wind tunnel tests. The tunnel curves do seem to fall off a little faster. At M = 0.80 the curves are very close to each other and the falloff toward the front is almost the same with the tunnel data falling of just a little bit faster toward the front than the airplane data. The one divergentpoint on the tunnel curve at positionE (solid symbol)appears to be an error in the original SR-6 tunnel data and may be caused by a malfunction- ing transduceror an improperlyrecorded amplifiergain setting. This error appearsto exist at the position for all of the data recorded after a certain time in the tunnel test program and representsan uncorrectableerror in the data of reference4.

The similarforward falloff of the data for SR-6 in the wind tunnel and in flight at M = 0.75 and 0.80 is probably the result of the boundary layers in the tunnel and on the airplanenow being closer to the same thickness. The directivitiesare not identicalwhich probablymeans the boundary layers are also not identical. Another possibility,although less likely,is that the flow around the windshieldwipers and supports presentedan inlet flow distor- tion to the SR-3 propellerand caused it to producemore forward radiated noise during the airplane tests. This possibililtyis less likely since the distortionfrom the wipers probably did not extend far enough above the air- plane fuselage to impact the propeller. In either case, because of the more nearly equivalentflow conditionsduring the airplane and wind tunnel tests of SR-6, the noise directivitiesare also more nearly equivalent. This provides further indicationthat the wind tunnel is a viabale locationfor measuring the blade passing tone of these propellers.

CONCLUDINGREMARKS Noise data taken with the SR-6 propellermodel flown on the NASA Dryden JetStar airplanewere compared with data taken previouslyin the NASA Lewis 8- by 6-foot wind tunnel. Comparisonsof the maximum blade passing tone variationwith helicaltip Mach number showed good agreementwhen the tunnel data were correctedto the flight test conditions. Directivitycomparisons also showed fairly good agreement. These good comparisonsindicatethat the wind tunnel is a viable location,having no more complicationthan the air- plane does, for measuringtheblade passage tone noise of these propellers.

A previous directivitycomparisonusing a differentpropeller (SR-3) showed that the tunnel directivitydata fell off more towards the front than did the airplane data at high axialMachnumbers. Thisprevious difference was attri- butedto the different boundary layerrefractions in the tunneland on the airplane probably a resultof the different boundary layerthicknesses in the two testsituations.It was foundthatthe airplane windhsield wipersand theirsupports were causingan energization of the airplane boundary layer resulting in an apparently thinnerairplane boundary layer. The windhsield wipersand supports weresubsequently removed for the SR-6 airplane tests, yielding a thickerboundary layerwhichwas probably closerto the tunnel boundary layer,and may haveresulted in the improved agreement betweenthe tunneland airplane noisedirectivities for the SR-6propeller model. In both of the datacomparisons (SR-6or SR-3)the noisepeak,whichliesbehindthe propeller plane,did not seemto be a function of boundary layerthickness.

APPENDIX _p powercoefficient, Cp : P / pN3D 5 propeller diameter J advanceratio,J = V / ND M axialMachnumber MH helicaltip Machnumber(vector sum of tip rotational and axialMach numbers) N propller rotational speed(revolutions / time) P shaftinputpower V axialvelocity Y distance awayfrom airplane fuelage Z axialdistance frompropeller plane(positive downstream) B bladesettingangleat 0.75radiuswith respectto planeof rotation e anglewith respectto propeller axis p density REFERENCES 1. J. H. Dittmar, B. J. Blaha, and R. J. Jeracki, "Tone Noise of Three Super- sonic Helical Tip Speed Propellers in a Wind Tunnel at 0.8 Mach Number," NASATM-79046 (December 1978).

2. J. H. Dittmar, R° J° Jeracki, and B. J. Blaha, "Tone Noise of Three Super- sonic Helical Tip Speed Propellers in a Wind Tunnel," NASATM-79167 (1979).

3. J. H. Dittmar, R. J. Jeracki, "Additional Noise Data on the SR-3 Pro- peller," NASATM-81736 (May 1981).

4. J. H. Dittmar, G. L. Stefko, and R. J. Jeracki, "Noise of the lO-Bladed, 40° Swept SR-6 Propeller in a Wind Tunnel," NASATM-82950 (September 1982).

5. J. H. Dittmar and P. L. Lasagna, "A Preliminary Comparison Between the SR-3 Propeller Noise in Flight and in a Wind Tunnel," NASATM-82805 (1982). _ 6. K. G. Mackall, P. L. Lasagna, K. Walsh, and J. H. Dittmar, "In-Flight \ Acoustic Results from an Advanced-Design Propeller at Mach Numbers to 0.8," AIAA Paper No. 82-1120 (June 1982).

7. R. J. Jeracki, D. C. Mikkelson, and B. J. Blaha, "Wind Tunnel Performance of Four Energy Efficient Propellers Designed for Mach 0.8 Cruise," NASA TM-79124 (1979).

8. J. D. Revel and R. H. Tullis, "Fuel Conservation Merits of Advanced Turbo- prop Transport Aircraft," LR-28283, Lockheed-California Company, Burbank, CA, (August 1977) (NASA CR-152096).

9. "Energy Consumption Characteristics of Transports Using the Prop-Fan Con- cept," DG-75780, Boeing Commercial Airplane Co., Seattle, WA, (October 1976) (NASA CR-137937).

10. D. B. Hanson, "Shielding of Prop-Fan Noise by the Fuselage Boundary Layer," HSER-8165, Hamilton Standard, Windsor Locks, CT (August 1981).

TABLE I. - SR-6 PROPELLERCHARACTERISTICS Design cruise tip speed, m / sec (ft / sec) ............. 213 (700) Design cruise power loading,KW / m2 / (shp / ft 2 . . . . . . . . . . . 30.0 (241) "deg" 40 Geometrictip sweep, ........................

Predicteddesign efficiency,percent ............... 81.9 Nominal diameter,D, cm (in) ................. 69_6 (27.4) TABLE II. - PROPELLERSR-6 BLADE PASSINGTONE MEASURED IN 8- BY 6-FOOT WIND TUNNEL Test conditions Transducerposition Blade Approx. Propeller Propeller A B i C D E t setting tunnel advance helical angle Mach ratio tip Mach Blade passing tone, _PL, dB, number number ref. 2xlO-D N / m_ 62 ° 0.85 3.5 1.149 (a) 138.5 142.0 143.5 141.5 .80 1.078 131.0 (a) (a) (a) 140.5 .75 1.008 137.0 146.0 142.5 137.5 128.5 .70 .937 129.5 129.5 130.5 135.0 126.0 .60 .807 119.0 119.0 118.5 115.0 112.5 64 ° .85 1.138 (a) 141.0 143.5 144.5 142.0 .80 ! 1.074 133.5 147.0 142.0 146.0 143.5 .75 I 1.009 137.5 147.5 143.5 139.5 129.0 .70 .943 134.5 138.0 137.5 136.5 126.5 .60 I ' .814 121.5 124.5 118.0 124.0 114.0 59 ° .85 3.06 1.222 (a) 140.0 147.0 147.5 145.5 .80 3.04 1.143 131.5 141.5 145.0 147.5 143.5 .75 3.06 1.074 137.5 145.5 142.5 143.5 142.5 .70 3.06 1.001 138.5 148.5 139.0 141.0 131.5 .60 3.06 .86 128.5 127.5 128.0 125.0 126.5 61 ° .85 3.04 1.220 (a) 140.0 145.0 149.5 149.0 .80 3.18 1.131 (a) 145.5 144.5 147.5 147.5 .75 3.09 1.068 138.0 146.5 145.0 148.5 147.5 .70 3.06 1.006 138.5 147.0 141.0 141.0 135.0 aData not available.

TABLE III. - PROPELLER SR-6 BLADEPASSINGTONEMEASURED ON JETSTARAIRPLANE Test conditions Microphone position Blade Airplane Propeller Propeller 1 2 1 3 I 4 5 ] 6 7 1 8 [ 9 setting Mach advance helical angle number ratio tip Mach Blade passing tone, SPL, dB, ref. 2x10 -5 N / m2 number 62.4 0.809 3.84 1.05 (b) 125.0 125.0 134.0 139.5 135.5 135.5 119.5 118.0 .754 3.78 .98 116.0 122.0 130.0 134.0 129.5 127.5 121.0 117.0 .714 3.61 .95 105.5 120.5 127.0 128.5 126.5 128.5 (a) 118.0 .614 3.60 .82 (a) 109.0 110.5 116.0 110.5 111.5 112.5 112.0 58.9 .805 3.29 1.11 106.0 120.5 135.0 143.0 141.0 142.5 132.5 (a) .753 3.31 1.04 108.0 128.0 136.5 140.0 136.0 135.0 125.0 113.5 .708 3.29 .98 (a) 126.0 132.0 135.5 131.5 127.0 120.0 113.5 .623 3.21 .87 _ , (a) 116.0 121.0 123.5 119.5 119.0 114.0 111.0 aData not available.

bNot operating.

TABLE IV. - PROPELLER SR-6 BLADEPASSINGTONE, MEASURED IN THE WINDTUNNEL,CORRECTED TO FLIGHT Test conditions Transducer position Blade Approx. Propeller Propeller A B C D E setting tunnel advance helical angle Mach ratio tip Mach Blade passing tgne,_PL, dB, number number ref. 2x10-_ N 62* 0.85 3.5 1.149 (a) 134.2 137.7 139.2 137.2 .80 1.078 126.2 (a) (a) (a) 135.7 .75 1.008 131.6 140.6 137.1 132.1 123.1 i • 70 .937 123.7 123.7 124.7 129.2 120.2 .60 .807 112.7 112.7 112.2 108.7 106.2 64* .85 1.138 (a) 136.7 139.2 140.2 137.7 .80 1.074 128.7 142.2 137.2 141.2 138.7 .75 1.009 132.1 142.1 138.1 134.1 123.6 .70 .943 128.7 132.2 131.7 130.7 120.7 .60 .814 115.2 118.2 111.7 117.7 107.7 59" .85 3.06 1.222 (a) 135.7 142.7 143.2 141.2 .80 3.04 1.143 126.7 136.7 140.2 142.7 138.7 .75 3.06 1.074 132.1 140.1 137.1 138.1 137.1 .70 3.06 1.001 132.7 142.7 133.2 135.2 125.7 .60 3.06 .86 122.2 121.2 121.7 118.7 120.2 .80 3.18 1.131 140.7 139.7 142.7 142.7 61" .85 3.04 1.220 lal 135.7 140.7 145.2 144.7 .75 3.09 1.068 132.6 141.1 139.6 143.1 142.1 .70 3.06 1.006 132.7 141.2 135.2 135.2 129.2 aData not available.

I0 (a)PROPELLER IN 8- BY 6-FOOT WINDTUNNEL.

C-81-5895 (b) PROPELLER ONJETSTAR AIRPLANE.

FigureI. - SR-6propellerinstallation.

, - SUPPORT STRUT t P R O PELLER P LAN E-x / / _ . / _ CENTERLINEOF - . _"' / _Z r_ _l_"_'_"_ F LH_ % 2 > -- --"'- i / II ' , t CENTERLI N E OF

_ _ S _ DE W ALL -7

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I I. 8 _ _ POS I T I ON TRANSDU CER (II / 2 D I AMEIIZR FROMT I P)

A I , 3 I c i o I E

TRA N S D U CER PO SITI O N. cm ( in . ) Z - 3 3 . 0 ( 13 . 0) _ 0.9 5 3 ( 0 . 3 75 ) 2 3 . 9 ( 9 .4) 45. 2(1 7 . 8 ) 107 .4(42.3 ) LH 4. 83 (1. 9 )10.2(4.0) 2. 5 4(1.0 ) 7 . 6 2( 3 .0 ) 3 1.5(12.4) NOM I NAL ANGLE, 7 5 9 0 1 0 1 1 10 131 B, deg Fi gu re 2 . - Pr essure tra n sd uc er positio n s .

/" . --

_'--.__,_.._,_____ _ }._-.-.._,_..,______. PROPELLER 2 3 4 ._r,i S . 6 _.. 8 "-_ CENTE R L I NE PRO P E LLE R POS I T I ON M ICROPHONE (0. 8 D I A M ETER FRO M T i P) ] I 2 I 3 I 41 5 1 6 1 7 1 8 I 9 MICROPHONE POS I T I ON cm (in.)

- 66 .0 -4 6 . 7 - 3 0.5 -1 6 .0. 76 ]2.4 2.. 5 .1 42.4 7 1.1 Z ( -2 6 . 0) ( - 18 . 4 ) ( -1 2 . 0) ( - 6 .3 ) ( .3 )(4 . 9) (9 . 9) ( 1 6.7) ( 20.0 ) NO MI NAL AN G LE , 5 1 5 9 69 7 9 91 9 0 107 I1 7 1 30 0, d e g.

(b ) M icrop h o n es o n airp l an e.

Fi g ure 2. - Meas u r emen t lo c at i o n s .

1 4 0

15 ° I

I / / oW_N D TUN NEL C ORRECT ED TO _ L, GNT

120 [ -- _ - _ - _ / 6 2 ° B L A DE SETTING, J'3 . 5

Iu / .... []wNi o TU NNEL C O RRECTED T O nIG H T

I __6 ( _Ni N _; - ; J - 3.5

% _0 / A A I RP L AN E DATA A T 62 . 4 0 BLADE x 110I- - SETI I NG, " l "AS NWRKED I (a ) I I I I I " l O0 j 1501-- o _ o _ 140 -- N _L j . 3.31 3. 2 9 (a..

130- / i J- 3. 21 1 2 0 -- O W I N DT UNNEL CORREC T E D T OF L I G HT 5 9 o BL ADE SE I T IN G , J • 3 .06 O W I ND TU N NE L C ORR EC E D TOFLIG H T 6 1 o BL AD E SETTI N G , J - 3 .06U NL ESS M ARKED 1 1 0 -- D AIRP LANE DATA AT 58 .9o B LA DE S Ell ING ," J " A S MA R K ED 1oo (b) I I I I I .8 . 9 l . O 1 .1 1.2 1.3 HE L IC A L T IP MA CH NU M BER ( a ) S R- 6 o pe rat e d n e ar it s des i gnadv a nc e rati oa nd s etting a ngle.

(b)SR- 6 operated nearthe S R - 3 designadvance rat i o a nd set t ingangle.

Figure 3 . - Maximum blade p assag e t on e variationwith heli c altip M a t h number.

0 WIND TUN N EL (CORRECTED 0 WIND TUNNEL (CORRECTED TO TOFIGHT) _ , = 6 1 . 3 ° , J - 3 .06 , MH - .86 FLIGHT) _ , - 61. 3 o , J - 3 .06 , MH - 1.0 [] AIR PL A NE , 1 3 - 5 9 .3 o , J - 3. ]2,MH- . 87 [3 AI RP LANE, p - 5 9 . 3°J- 3.02 , _ - 1 . 03 A A IR PLA NE , I _= 61 . 5o , J " 3. 00 ,M H -. 8 3 A A IR PLA NE , 1 _ - 6L5°J =3. 0 5 , M H = 0 . 97 1 40 -- DO-- 11 0 % 1_--_ (a) (b) % i oo I I I I I I I I u ') 0 WIND TUNNEL (C O RRECTED TOFLIGHT) , 0 WINDTUNNEL (CORRECTED TOFLIGH T) , z _ - 61.3 ° J - 3.0 6 , M H - 1. 08 1 3 " 61.3° , J " 3.0 6 , MH - 1 .14 [] AIRPL A NE , _-5 9.3 o, J -3 . 0 6, M H - I. 0 8 [] A I RP LA NE , _-5 9.3 o , J - 3.13 , M H -l .14 < _ A AIRPLANE, _ '6L 5 ° , J=3.32 , MH-1.03 A AIRPLANE, _ -6L 5 ° , J-3.39, M H-1.09 (C) (d)

I I I I I I I I

110 5 0 " ?O 9 0 11 0 13 0 50 7 0 90 11 0 13 0 ANG L E FROM P RO P ELLER AXIS I I I I I I I I I I A B C D E A B C D E TUNNEL M EASURE M ENT LOCATIONS I I I I I I I I I I I I I I I I I 1 2 3 4 5 6 7 8 9 1 2 3 4 5 6 7 8 9 A IRP LA NE M EA S URE M ENT LOCATIONS (a)Nominal axial M achnumber - 0.6 . (b)Nominal a xial M achnumber - 0.7 .

(c)Nominal axial M achnumber - 0.7 5 . (d) N ominal axial M athnu m b e r- 0.80 .

F i gure4. - Blade p as si ng tone dir ect i vityf o r S R-3propeller ( re f. 5 ).

20 8 V O 5in. RAKE ?

• 15 >. . 10 _.

0 10_ .5 .6 . 7 .8 . 9 1 . 0 V (Y ) / VEDGE Figure 5. - Airplaneboundarylayerduring SR-3tests, with wipers; M=_ O.8.

8 -- 20 -- O 5 in. RAKE 7 -- O 8in. RAKE 15 -- 6-- 5-- S 1o-- 4-- > _- >_ - 5-- 2 O-- 0 I .5 . 6 .7 . 8 . 9 1.0 V (Y ) / VE D GE Figure6 . - Airplaneboundarylayerduring SR-6tests , withoutwipers ; M-- O.8.

1 40 -- m O WINDTUNNEL CORRECTED TO FL I GHT 13 - 6 2 ° , J - 3 . 5 , M H - O. 8 0 7 • 13 "64o , j. 3 . 5 ,M ! . I- 0 . 81 4 Z_A I RPLANE, _ - 6 2.4', J - 3 .60, MH - 0. 8 2 1 20 -- 130-- [] W IN D T U NNE L CORRECTED TO FLIGHT _t,_ / I_ - -_ , _ k I / O WI NDTUNNEL CORRECTED TO FL I GHT ... . _ / 1 3 " 6 2 °, J " 3. 5, M H " 0 . 93 7

_ 11 o - _.._ __ _ , ' _W I N D T UNNEL CORRECTE D TO FLIG ,T

..., I _' " 64° ' J " 3 . 5, M id - 0.9 4 3 ,_, /! , Z _ A I RPLA NE, 13 - 6 2.4 u, J - 3 .61 , M H " 0 .9 5 lOO(a ) I I I I I I I I I Lb) I I I I 1 I I 1 I J 150-- o_ _ 14o- < • _ 13o- / \ 1 20 -- , /Z _ _ Z _._ . _ \ O WIN D T UNNEL CORRECTE D _ r7 WIN D TU N N E L CORRECTED TOFLI G H: _ T O FL I G H T, F - 6 2 °, J - 3.5, M H - 1 .008 FLI G H T , _ - 64 °, J - 3 . 5, M H _A 1.0 7 4 110 -- r-I W I NDTUNNE L COR R EC T E D T O FL IG HT , _ _ A I R P LANE, I B- 6 2.4 ° , J - 3.84, l_ " 64 ° , J " 3 . 5, M H - I. 00 9 M H- I . 0 5 A A I RPLANE 13 - 62.4 ° ,J - 3 . 1 8 , M H - 0. 9 8 • SOL I D PO I NT A P PARENTL Y I N ERROR

lool_I I 1 I I 1 I I I (_ )1 I I I I I I I I

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A N GLEFRO M PR O PELLER AX I S

i I I I I I i I I I

A B C D E A B C D E TUNNEL MEASUREMENT LOCAT I ONS

i I I I I I i I I I I I I i l l I I

1 2 3 4 5 6 7 8 9 1 2 3 4 5 6 7 8 9 AI R PLANE ME AS U R EMENT L OCAT I ON S (a) N o minal axialMachnumber-0.60. (b ) NominalaxialMachnumber-0.70 .

( c)NominalaxialMachnumber- 0. 7 5. ( d ) N o minal axialMachnumber- 0.80.

Figu re 7 . - Bl ad e p assi ng to n e d irec ti v i t y f o r SR - 6 pr op el l e r .

1. Report No. 2 . GovernmentAcce s sionNo . 3 . Reci p ient ' s Catalog No.

NASA T M- 833 41 4. Title and Subtitle 5. Re po rt Date A P R E LIMI N AR Y COMPA R IS O N BET W EEN T HE S R -6 PR O- PEL L ER N OI SE I N F LI G HT AND I N A WIN D TUNNE L 6. Pe rforming O rganiza t ion C ode 505- 3 1- 32 7 . Author( s ) 8 . P erformingOrga n izationReport No . : James H. Di t tmar , Paul L. Lasa gn a, and Ka r en G. M ackal l E- 1 5 96 10. Work Unit No.

9. PerformingOrganizationName and Addre ss Na t ion a l A e ron a u t i c s a ndSpa c eAd mi nist rat ion Lew is Res e a r chC e n t e r 11 . Contract o r Grant No.

C le v el and, Oh i o 44 1 35 13. T yp e o f Re po r t an d P eri od Co ve r e d 1 2. Spon so ringAg e ncy Name and Address Techn ic al M e m o randum Na t i o nal Aer ona utic s an d Sp ace Adm inis tr a t i o n Wa s h ington , D. C . 2 0 546 14. S pon s oring Agency Co de 15 . Su pp le m entary Note s Jame s H . Ditt m a r , NASA L e wi s Re s ear ch C e nter; Paul L. La s a gn a an d Ka ren G. M ac kall , Hu g h L. Dryde n Flight R esea r ch Cen t er, Edw ar ds , C alif ornia. P rep ar ed for the One-hu n d r ed- f ift h M eeti ng of the Acoustic al Socie ty of America, Cinci n nati, Ohio , May 9-13, 1 983.

16. Abstract Hig h s p ee d t u rb o pr o ps o f f er an attrac t iv e c an didate for f u tur e aircraft because o f t h e ir hig h " propuls i ve efficiency. However, one of the possible problems associated with these propellers is their hig h noise le v el at cruise con di tion that m a y create a ca b in enviro n ment problem.

Models of these propellers have bee n tested for acoustics in t he Le wi s 8-by-6-foot wind tunnel and o n the Dryde n JetS t a r ai rplane. This p a per shows co m p ar iso n s betwee n the airplane an d wind tun n el data for the S R -6 propeller. The co m p ar ison of ma x i m um blade passing to n e varia- tion with helical tip Mach n u mber between the tunnel and fli g ht data was g oo d when corrected to the s am e te s t co ndi tions. Di r ectivity comp ar isons also showed fairly g ood a g ree m ent. These g ood co m p ar isons i n dic a te that the wi nd tunnel is a viable location for measuri ng the blade pa ss a g e to n e n oise o f the se prop e ller s .

1 7. Key Words (Sugg es ted by Author( s )) 1 8. Distribution Statement Propell er no i s e Un c l a s sifie d - unl im ite d N oise STAR cat ego ry 71 Supersonic tip speed 19. Security Clas sif.(of thisre po rt) 20. SecurityCla s si f .(o f thi s p age) 2 1 . No . of Page s 22. P rice* Un cla s sifie d U n classifie d * For saleby t heNational Tec h nical I nfo r ma t ion Service Springfield , Vi r gin i a 2 2 161 " ) FOSTMASTER: If Undeliv e r a hle (S ,c t io n 15 _ Post a l Manual) IX) N u t R e turn I i I