Wind tunnel measurements of forward speed effects on jet noise from suppressor nozzles and comparison with flight test data

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NASA TECHNICAL NASA TM X62,449

MEMORANDUM

N75-25946 MIND TUNNEL 1EASUREt1$NTS

X (NASA-TM-X-62449)

ET NOISE FROM .

2 OF FOPWARD SPF.FD EFFECTS ON .

► N WITH M PARIS ESSOF NOZZLES AND CO B Q SUL " Unclds P HC $3.25 N FLIGHT TEST DATA (NASA) CSCL 20A G3/07 2b6b9 Q

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WIND TUNNEL MEASUREMENTS OF FORWARD SPEED EFFECTS ON JET NOISE FROM SUPPRESSOR NOZZLES AND COMPARISON WITH FLIGHT TEST DATA Adolph Atencio, Jr.

Ames Research Center and U.S. Army Air Mobility R&D Laboratory +4 4FC Moffett Field, Calif. 94W'S

/A^y ` s jP

a I k eele'e May 1975

.

3, R5lplent's Catalog No, 2. Government Accndon No, 1, Report No.

NASA TM X-62-449 6. Report Date btitle 4. d Tltl Tunnel Measurements of Forward Speed Effects ^indw

on Jet Noise from Suppressor Nozzles and S. Performing Organization Code

Comparison with Flight Test Data 0, Performing Organization Report No, 7, Author(s) Adolph Atencio, Jr. A-6054 10. Work Unit No, 9, Performing Organization Name andAddrea 743-03-11-00-21

Ames Research Center and U.S, Army Air

11, Contract or Grant No, Mobility RPD Laboratory, Moffett Field, Calif.

94035 13, Type of Report and Period Covered 12, Sponsoring Agency Name and Address Technical Memorandum

National Aeronautics and Space Administration

14. Sponsoring Agency Code Washington, D. C..20546 ementery Notes 16, Sup Presented at AIAA 8th Fluid and Plasma Dynamics Conference, Hartford, Conn., June 16-18, 1975.

19. Abs;rprs The results of a test program conducted in the NASA Ames 40- by 80-Foot Wind Tunnel to determine the effect of forward speed on the noise levels emanating from a conical ejector nozzle, a 32-spoke suppresso nozzle, and a 104-e',liptical-tube suppressor nozzle are reported. It is shown that noise levels are reduced as forward speed is increased and that for one suppressor configuration, forward speed enhances suppression, Comparisons of noise measurements made in the wind tunnel with those obtained in fl.ght tests show good agreement. It, is concluded that wind an effective means of measuring the effect of forward tunnels ,provide speed on aircraft noise.

17. Key Words (Suggested by Author(s)) 19. Distribution Statement Unlimited Aircraft Noise, Jet Noise, Wind-Tunnel Noise Measurements, Acoustics, Suppressor Nozzles, STAR Categories - 01, 07, 45, 71 Flight Noise Data 22. Price' 19, Security Claul4 lot this report) 20. Security Clauif. (of this pagal 21. No. of Pages 10 $3.25 Unclassified Unclassified 'For sale by the Natloha2 Technical Information Service, Springfield, Virginia 22161

WIND TUNNEL MEASUREMENTS OF FORWARD SPEED EFFECTS ON JET NOISE FROM SUPPRESSOR NOZZLES AND COMPARISON WITH FLIGHT TEST DATA Adolph Atencio, Jr,* U.S. Army ASr Mobility Research and Development Laboratory and Ames Research Center, NASA, Moffett Field, Calif, Abstract under the right wing of an aircraft research model as shown in Fig. 2. Both arrangements were used The results of a teat program conducted in the for studies on the static test stand and in the NASA Ames 40- by 80-Foot Wind Tunnel to determine wind tunnel. The nozzles studied were Lite Imes- the effect of forward speed on the noise levels line conical ejector nozzle, n 32-spoke nozzle, and emanating from a conical ejector nozzle, a 32-spoke a 104-elliptical-tube nozzle with an acounticallY suppressor nozzle, and n 104-elliptical-tube sup- treatedShroud. The nozzles are Shown in Figs, I pressor nozzle are reported. It is Shown that through 5.

noise levels are reducud as forward speed is in- The static portion of Lite testing was Aunt it creased and that, for one suppressor configuration, the Ames Static Test Facilit y . The model was forward speed enhances suppression. Comparisons mounted on a test stand so that the centerline of of noise measurements made in the wind tunnel with the engine axis was 6,1 m (20 ft) above Lite ground those obtained in flight tests show good agreement. surface. Microphones were placed to duplicate Lite It is concluded that wind tunnels provide an effec- wind tunnel positions and were also placed on a tive means of measuring the effect of forward speed 30.5 m (100 ft) arc, referenced tr, the nozzle exit on aircraft noise.

centerline, to make far-field measurements. The wind tunnel microphone positions were lee m (6 ft) above the surface while the Far-field microphone Nomenclature height was 6.1 m (20ft). The microphone Setup was similar for both of the nacelle mounting liz cycles per second arrangements. Schematics of the microphone setup OAS?L over all auund pressure level, dB are shown in Fig. 6. The installation In the wind SPL sound pressure level, de tunnel for both the isolated nacelle and nacelle Vj jet exit velocity, m/sec (ftlace) mounted to the model was w:.tli the centerline of the Vo wind tunnel velocity, m/sec (£t/sec) nozzle exit 6,1 m (20 ft) above the tunnel floor.

VR Vj - Vs, m/sec (ft/sec) Figure 7 shown the isolated nacelle inbtalled In O1 acoustic angle measured from inlet axis the wind tunnel, introduction Data Acquisition and Corrections Use of a wind tunnel facility to measure the Data from the static test facility were effect of forward speed on aircraft noise offers obtained for several jet velocities; wind tunnel several advantages over flight tests in terms of noise measurements were obtained at several combi- costs, test condition control, model configuration nations of forwarc speed and jet velocity. The control,. and data sample time. The staff of NASA static tests provided data to establish: (1) near- Ames Research Center has used Ames 40- by 00-Foot field to far-field d-irectivity difference for data Wind Tunnel, to arudy such effects on propeller air- measured in the wind. tunnel and extrapolated to craft, large. Seale STOL models, and jet engines.)-s flight distances; (2) a comparison of data for the This paper summarizes a test program conducted in isolated nacelle and the nacelle under a wing; the 40• by 80-Foot Wind Tunnel to determine the (3) far-field directivity for each nozzle; and (4).

effect of forward speed on noise from a conical free-field data for wind tunnel microphone positions ejector nozzle, a 32-Spoke suppressor nozzle, and a to determine the reverberation corrections for wind 104-elliptical-tube suppressor nozzle with and with- tunnel data.

out a treated ejector shroud. Wind tunnel data The reverberation corrections s ' s were estab- showing the .effect of forward speed on noise for lished on a 1/3-octave bond SPL spectrum basis by each nozzle are presented and comparisons of wind comparing the outdoor data (corrected for ground tunnel data with flight test data from NASA Lewis reflections) to similar spectrums measured in the Research Center F106H fly-over tests are shown for wind tunnel at zero forward speed. The differences the conical ejector and the 104-elliptical-tube between the wind tunnel data and free-field outdoor nozzles. Wind tunnel data corrections and flight data, at ee.,ch 1/3••octavx center frequency, were test data corrections necessary to make comparisons used as the reverberation corrections. It was are summarized.

found that the reverberation corrections established were independent of nozzle type and power siting of the engine. Addit.lonal details are given 4,n Model Description and Teat Setup Ref. 7.

The suppressor nozzles and base line conical Results ejector nozzles were tested with a GE 85 turbo- jet engine mounted in a flight nacelle identical to Static Testa that used during the F106B flight tests at Lewis Research Center. The nacelle was mounted in two Figure. 8 shows the noise measured at the static ways, first isolated as shown in Fig. 1 and then test facility at Several microphone locations with the conical nozzle on the isolated nacelle.. Data from the near-microphone positions were corrected *Reuearch Engineer -I- U G ^ OF OOR Q

M I for ground reflection and extrapolated to the far- Figure 14 shown the relative velocity effect -ft) arc, Com- field positions on the 30,5 m ' on a 1/3 octave spectrum basis for the horxles parison of the data for different; microphone tested. Tito affect in all canes in a reduction In locations indicates that for ti;e close-in micro- the spectrum levels for increasing forward npeed, phones there appears to be a shift in tite directiv- ity pattern when compared with far-field, However, Flight Comparison for the maximum noise angle, the near-field and far-field sound pressure levels were in good Wind tunnel data for the conical nozzle and agreement, F1'ure 9 compares the far-field direc- for the 104-tube nozzle were compared with tent tivity with tGe isolated nacalla to the directivity data obtained in flight teats of an Fio6h nirerafc.

with the naville mounted under the wing of a model, The comparison of flight data with wind tunnel as measured at the static test facility. There data requires that both sets of data be froo-flcld, was a change in directivity in the far-field for that the same distance be used for the comparison, each of the nozzles. The 32-spr' • e nozzle allows a and that flight test tints and wind tunnel data be alight increase in noise in the .:ward quadrant compared at the noise angle at the time of noise and little or no change in the aft quadrant when emission. The flight dnta were reduced and the the angina is mounted on the model. The conical emission angle was accounted for by using retardod ejector nozzle showed an increase in for% 4c time. The emission angles in tics wind tunnel were quadrant noise of from 1 to 3 PNdB when the engine corrected for flow convection, The flight test was mounted on the modal, and a decrease of 1 to 2 data were corrected to frge-field by first correct- PNdB for most of the aft quadrant r.Dlae. The effect ing for ground reflectians using the procedures observed for the 104-tube nozzle is that, the model outlined in Ref, 10, These procedures are similar caused forward quadrw.t noise to increase 1 PNdB to those of Refs. 5 and 6 except for modificntfono t 3 PNdB. This and the peak noise to decrease by 2 to make tie suppress pz nozzle correction more suggests that installation effects can alter the realistic, Tile flight data were also corrected observed effectiveness of suppressors.

for Doppler .shift and to 1;a1dard da,? conditions (59°F and 70% relative hu- 1 0lty).. 'rhe wind tunnel Wind Tunnel Tests data were corrected to fr!,c- , field by applylag the reverberation corrections determined from Oe The effect of forward speed on noise was static tests and ware then extrapolated to the observed in the wind tunnel for the various nozzles flight measurement distances by using aphericnl studied. In general, the effect of forward speed, attenuation and applying the near-field to far- for a given jet velocity, was a reduction in noise field corrections established from the static teats.

with increasing forward speed. The effect for the The wind tunnel data were also corrected for conical nozzle was similar to that observed Stati- standard day atmospheric attenuation by using cally when the jet velocity was reduced by changing Ref. 11, Slight differences in relative velocity eagir.o power, Figure 10 shows OASPL as a function between fly-over data and wind tunnel data were of relative velocity (Vj-Vo) at the Peak noise accounted for by correcting the wind tunnel data angle. The wind tunnel data for the conical nozzle to the Same relative velocity as flight by using at var.ioua forward speeds all fall on the static Figs. 10, 11, and 12. The actual corrections (zero forward speed) line; thus relative jet velo- from this Source were less than 1.5 dB.

city adequately defines the variation of peak angle The resulting comparisons of data are shown noise with forward speed for the conical nozzle.

in Figs. 15 through 19. Two different nozzle com- Similar data are shown for the 101,-tube-mixer sup- parisons are shownt a comparison of data for the pressor nozzle in Figs. 11 and 12. The figures conical ejector nozzle and a comparison of data show that as forward speed is increased the noise for the 104-elliptical-tube nozzle, both with and at the peak angle both with and without the acous- without the treated ejector shroud, Figures 15, tic Shroud is reduced more than would be predicted 16, and 17 show perceived noise level versus by relative velocity alone. This is especially acoustic angle from the illet. The flight data true for a tunnel Speed of 91 m/sec (300 ft/sec).

are from PNL time historic ,, of fly-overs with the The reason for this excess attenuation is unknown.

same relative velocities ,s the wind tunnel data.

Also shown in . Fig. 12 are the peak angle noise In Fig. 15, the compares-a of data for the conical.

levels at zero wind tunnel speed for the conical ejector nozzle shows Lilt; the data are within ejector nozzle and for the 104-tube nozzle without 2 PHdB except at the 15,1, 1 position where the dif- the acoustic Shroud. The figure Shows that for the ference is 6 M g . FiglAte 16 shows the comparison zero speed case, the noise from the 104-tube nozzle for the 104-elliptical-tube nozzle without the without the shroud is 4-5 dB higher over the range acoustically treated Shroud; the data agree within of relative velocities shown and that the conical ±1,5 PNdB. The angles not Shown for the 104-tube ejector is from 9 dB to 16 dB higher in noise level nozzle, but shown for the conical ejector nozzle, than the 104-tube nozzle wish the acoumbic shroud.

were influenced by the high background noise level The effect of forward speed on the directivity of the tunnel and model supports at those nngles.

of the conical. nozzle is shown in Fig. 13. There Figure 17 shows that the wind tunnel and flight was a-reduction in noise at all angles for increas- data. .:or the 104-elliptical-tube nozzle with the ing forward speed. The amount of reduction for acoustically treated .shroud agreed within 12 PNdB.

each forw,,cd dpeed change, however, is different Figures 18 and 19show comparisons of flight for different angles. Recently obtained data from data and wind tunnel data on a 1/3-octave spectrum Bi9 have ahoun iiy. over tents conducted in England basis, Figure 18 shows a comparison of spectral a measured increase in noise in the forward quad- data for the conical ejector nozzle, The flight rant angles (reference to inlet) and little or no dats are greet Ref, 12, which included only peak change at 90°; these effects were not observed in noise angle 1/3-octave spectrums for the flight the data presented here. However, as is shown in data. The actual correspondence of data from the next section, comparisons of wind tunnel data flight and wind tunnel tests is not exact but with F106B flight data show good agreement.

ORIGINAL PAGE,

i

Falatski, M.D., Koenig, D,G., and Soderman, within 6' of acoustic angle. The 3.

wi`1d tunnel data P.T., "Aspects of Investigating STOL Nolac have been extrapolated to the flight distance and Using Large Scnlo Wind Tunnel Modolu," NASA corrected for atmospheric attenuation, Doppler IMS-62, 164, June 1972.

shift, and near-field to far-field difference. The 4. Wilcox, Fred A., "Comparison of Ground and difference in acoustic angle and large atmospheric Flight Test Results Using A Modified 11106D attenuation corrections probably accounts for the Aircraft," NASA TMK-71, 439, November, '974.

difference in the. spectrums at high frequencies.

(touch, R. and Thomas, P., "The Influence rf Figure 19 shows 1/3-octavo spectrum compari- 5.

Reflections an the Sound-Pressure Spectra al mono for the 104-alliptical-tube nozzle, both with Jets," Societe Nationale d'Etudo et do and without the acoustic shroud. The flight data Moteurs d'Aviatlon Toulouse, 6, 7, and 8 were supplied by Lewis Research Center and are March 1968, within 3' of acoustic angle of the wind tunnel data..

6. Howes, Walton L., "Ground Reflection of .let The wind tunnel data were extrapolated to the flight distance and corrections applied as before; there Noise," NASA TR-35, 1959.

is good agreement over most of the spectrums. The 7. Beulke, M.R „ at al., "A Forward Speed Effects Study on Jet Noise From Several Suppressor disagreements in the lower frequencies are due Nozzles in the NASA/AMES 40- by 80-Pout Wlnd primarily to the inexact corrections for reflections Tunnel," The General Electric Company, NASA in both the wind tunnel and flight data. The reason for high frequency disagreement in the spec- CR 114741.

trums is not known. The agreement of the spectrums 8, Bushell, K,W„ "Measurerent and Prediction of Jet Noise in Flight," AIAA Paper No, 75-461, is considered good considering the magnitude and March 1975.

accuracy of the corrections applied to the data.

Brooks, J.R. and Woodrow, H.J., "The Erfeetn 9.

of Forward Speed on a Number of Turbojet Conclusions Exhaust Silencers,' AIM Paper No. 75-506, March 1975.

From the data presented in this paper, the 10. Burley, Richard R., Karabinus, R.J., and following conclusions can be made, Freedman, R.J., "Flight Investigation of 1. The forward speed effect .oscrved in the Acoustic and Thrust Characteristics of Several wind tunnel shows a decrease in noise Exhaust Nozzles InsSalled an Underwing Nacelles level as forward speed is increased for the on an F106B Airplane," NASA TMX-2854, August nozzles tested.

1973, a) Noise measured from the conical nozzle 11. Anon, "Standard Values of Atmospheric Absorp- at ttI peak noise angle showed the tion as a Function of TLpIperaturo and Bumidity classical relative velocity effect for Use in Evaluating Aircraft Fly Over Noise," with a predictable attenuation With Society of Automotive Engineers Inc., August forward speed.

31, 1964.

b) For the 104-tube nozzle, with and 12, Brousch, J.F., "Flight Velocity In f luence on without the acoustically treated Jet Noise of Conical Ejector, Annular Plug and shroud, the decrease in noise with Segmented Suppressor Nozzles," The General increased forward speed at the peak Electric Company, NASA CR-120961, August 1972.

noise angle in not predictable from the relative velocity; instead, at the higher forward speeds, more attenua- tion than would be predicted was observed.

1. Wind tunnel data compare favorably with flight teat data when appropriate correc- tions are made to both sets of data.

3. {easurements made in the near-field, as : squired in the wind tunnel, of the jet q..lae source show some differences when CL fared to measurements made in the far- fiE i for the same acoustic angle. These difi, cences should be accounte for when comparing wind tunnel data to flight data, III 4. The presence of the model changes the k directivity of the noise source in the i far-field as compared to the isolated nacelle.. The directivity change varies I with nozzle configurations.

i^ 1!

References i 1. Atencio, Adolph, Jr. and Soderman, Paul T., "Comparisons of Wind Tunnel and Fly Over Noise.

Measurements of the YOV-10A C TOL Aircraft," - NASA TMX-62166, June 197,2.

2. Atencio, Adolph, Jr., Kirk, Jerry V„ Soderman, Paul T., and Hall, Leo P., "Comparison of Flight

ORIGINAL PAGE IS

and Wind Tunnel Measurements of Jet Noise for

OF POOR QUALITY(

the XV-5B Aircraft," NASA TMX-62, 182,.

October 1972.

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F` F'1K. 1 Isol, tiivelly - static test installation.

32• ,. 1 Nacelle undcc tnstallatt,i. in wind ^41 44.4 Cm 63 5 Cm f I (j75,n dio) tunnel. 125+n dig ) 60 6 Cm (23 96 m. ) Fig. 3 Conical ejector nozzle.

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l2i.ia^+.doi 12i.^a^n.dal V-GUT TER DES, ^N 104 elliptical tube nozzle with Fig. S 39 spoke nozzle.

Fig. 4 acoustically treated shroud.

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I'll IJ , NOZZLE IIH/1 2m 17111 3 Tm 112 25f/) MICROPHONE REF PLANE NOIfZ .[ CONFIO' 'ATION CONICAL 32 SPOKE 104 TUBE WITH SHROUD ILIA TUBE WITHOUT SHROUD (a) Isolated nacelle microphone locations - wind tunnel set up.

Fig. 6 Microphone locations.

Fig. 1 In.,lated nacelle to wiml tunnel •O' I •0' CONICAL EJECTOR NOZZLE 30.Sm ARC (100111 VI•5e2M/wcovof1isec) 150' X, 15'n (10r' . t) 0 309111 f ARFIELD (1pOfil A S 41m NEARFIELD (left) `' 396m NFARF IELD 113ft) . W+ m 140 00^ J BS E..Gv* EXHAUST z CENTERLINE o ^ 1)0 AIRPL _ AN[f W ' > J \\ 1 W 1 V • W 6 J 1 10 20 40 fir, Bo 100 120 1.0 160 too ACOUSTIC ANGLE RfFFRENCED TO INLET, deg

(b) Microphone positions on

30.5 m

radius, arc.

Fig. 6 Continued.

Fig. 8

Isolated nacelle I .utdoor static test far-

field/ near . 6eld comparison, PNdB

directivity, conical effector nozzle.

MICS Iam 10.

ABOVE GROUND 1.00 20' 3O• •0' ENGINE CENTE RL -Nt 1 i6 1m ABOVE GROU40 Vi' "4miset 149. . Htfsecl 60' 120 - FO' 7 -, AIRPL ANt f +^ - _ --- -- Ile ^\ I,• BO' NOZZLE ISO .T[D NACELLE p IRPI ANE \\1I ^^ PIA( :LL[ UNDER ♦ WING MODEL MICROPHONE C REFERf NCF ;00. CONICAL o • LINE "0' 104 1 LiBE W/SHROUO • / 120• 32 SPOKE O • 130' I40, %CP I la 160• !K" (c) Microphone positions for wing and nacelle.

Fig. 9 Far-field directivity for iROlated nacelle Pig. 6 Concluded, versus nacelle/wing installation.

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150, WIND TUNNEL FORWARD SPEED O ZERO D 51.6 m/sec117O 11/seep Igor A702m/SeC(250f1/S0C) O 91.4 m/SeC 1300f 1/SeC) WIND TUNNEL FORWARD SPEED CONICAL FAZ24E m o ZERO IZERO SPEED) 076 m/uC (50011/980 140 j a 91 ml99c I50011p90 IOA TUBE NOZ Z LE 4 tp WITHOUT SHROUD ` E D) / ^I30^ (FERO SPEEDI Y] 104 TUBE NOZZLE % WITH SHROUD n

! (ZERO SPEED)

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- -1 __ n0 M 600 Soo 1000 m/SeC .— 0 4 2000 4 -.1 200 400 600 N,000 m/SeC t._.e. .. 30 3000 fl/5:,: 1 000 2000 000 2,000 4,000 N/SeC RELATIVE VELOCITY 1,000 JET RELATIVE VELOCITY Fig. 10 Static data and wind tunnel forward ap6ed data for conical ejector nozzle Fig, 12 Static data and wind tunnel forward speed data for tae 104-tube nnzzle (peak noise).

with acoustic shroud (peak noise )• V l *Slam/Set U7p0!!/xt) DATA CORRECTED FOR CONVECTION 150 WIND TUNNEL FORWARD SPEED 130x.

O ZERO *51.6 m/uc 170 N/Sec) ZERO SPEED A 76 M" 25011/91c) m 140 v 091 m/196130011/640 ^ 4f It d 91 m/sec WIND TUNNEL VELOCITY d • 0 0 /Sec (170f1/SeC)

y ^ A 13 766

A 76m/SeC (25011/Set) n120 4 91m/Sec 1301,11/sec) ..1. l.:>lx:l. I - 1 1 -.: -3 .

60 70 60 90 100 110 120 130 140 150 160 ACOUSTIC ANGLE REFERENCED Tn INLET, deg 200 400 600 1.000 m/sec.

500 it/see 1,000 2A00 4,000 JET RELATIVE VELOCITY Fig. 13 Wing nacelle wind tunnel test, OASFL directivity, conical effector nozzle.

Fig. 11 Static data and wind tunnel forward speed data for the 104-tube nozzle 'v ^a without shroud (peak noise ).

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ORIGINAL PAOU IS

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Bt 147' V1 • 518m/Sec 1170011/sect 3.96m SIDELINE 113111 VO m/3ec h/see VR .50Am/SCC 1165311ryeC) 0 0 0 > 130 ^^ 0 57 170 120,- ;,4 A m Sip ^. .. rr 110 1.

a 120 m W x { 100 . 41013000 WI N DATA ♦ 40 13 Y00 Y EI 110 WIND TUNNEL DATA 1 1_.1 1 1.. a._.l_ _1 _ 5 - J 60 70 80 90 100 I ED 140 150 160 ACOUSTIC ANGLE REFERENCED NCED TO INLET, -deg Er .—..+-.,.-.+ r 102 103 104 FREQUENCY, H2 Fig, 15 Comparison of wind tunnel data with (a) Conical ejector nozzle.

flight data for the conical ciet tot nozzle.

Fig, 14 Wing nacelle wind tunnel test, 1/3 octave band spectra.

Bz .133' V) - 594mhec (195011hec) 120- 5.9m SIDELINE 118111 W 110- s m/:ee 11/CCC "— > ISOf W - O v 0 0 '• 0 52 170 T6 250 ° FIOG8 FLIGHT DATA e o 1 20 m 40 BY 80 WINO TUNNEL DATA o W 0 5D.

VR .. 534 m/see (1760(1/seC) U : 110 8090 IDO 110 A e ° 120 130 14n iSG ACOUSTIC ANGLE REFERENCED TO IN L ET, daq

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F - - ^- 162 103 104 FREQUENCY, Hz (b) 32-spoka area ratio nozzle, Fig, 16 Comparison of wind tunnel data with flight data for the 104-tube nozzle Fig. 14 Continued.

without acoustic shroud.

8,1127- VI • 533m/Sec 11750 HMO 3.96m SIDELINE (13h) VO 120 -- m/see 11/see 0 0 110- n 52 ITD o` 120 a a ^^ m1 a 100„ °^v m g 110 o F1060 FLIGHT DATA Woo ac • 40 BY 80 WINO TUNNEL DATA "^'A d 100 om VR • 536 m/WC (176011/Ser) ^a m m 90 --) X 90 100 110 x 120 130 140 130 1/.f e' 102 103 104 ACOUSTIC ANGLE REFERENCED TO INLET, deq FREOIIENCY, H2 (e) 104-tube nozzle without shroud.

Fig, 17 Comparison of wind tunnel data with flight data for the 104-tube nozzle Fig. 14 Concluded.

with acoustic shroud.

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P X[ I a NASA LEWIS non FLIGHT DATA A NASA AMES WIND TUNNEL DATA rt W ` ` ACOUS TI C e g ANGLE, d eg PNOLE, t l e 2 100 F +141 1 n a 90 P we S.

> ; DOL • u ^ 70!

¢ z Yp, 458 In/feC (1501 IIfSIC) s- 60, 100 1000 10000 FREOUENCY, H1 Fig, 1B Comparison of flight data with wind tunnel data, conical effector nozzle, flight acoustic angle 147 0 , wind tunnel acoustic 141°, 0 NASA LEWIS F1060 FLIGNT DATA 110 ' A NASA AMES WIND TUNNEL DATA 1Y 110' n NASA LEWIS FIOGe FLIGHT DATA ° ACOUSTIC A NASA AMES WIND TUNNEL DATA m c ANGLE,deg O ACOUSTIC 100 ANGLE,deg.

,x100- N 121, GO- O 1.,.Ad^ W Y , = B0 6G (a ). 1.1 1.....L- - 1. 1 1 1 l 1 1 L (C) 1 , , 1 1 1 , Y , WyJ J W W i 110 - 1 IW• ACOUSTIC Z ACOUSTIC Q ANGLE, deg a N i ANGLE, deg m 100 L. 140 p y1 r 14S r i H (d) GD (bI 1 I.l , 1 11 , 1 1 I • 80' 100 1,000 I0,OD0 loo 1.000 -?.c FREOUFNtY, NI FREQUENCY, H2 (a) Without acoustic shroud, flight (c) With acoustic shroud, fllghl acoustic angle 118°, wind tunnel acoustic angle In go , wind umLV, acoustic angle 121°. acoustic angle 1.21°.

(b) Without acoustic shroud, flight (d) With acoustic shroud, flight acoustic angle 140°, wind tunnel acoustic angle 109°, wind Luuurl acoustic angle 147°. acoustic angle 106°.

Fig. 19 Comparison of flight test data with Fig. 19 Concluded.

wind tunnel data, 104-elliptical-tube nozzle.

ORIGINAL PAGE M

OF POOR QUA=.

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