Acoustic results of the Boeing model-360 whirl tower test

ACOUSTIC RESULTS OF THE BOEING MODEL 360 WHIRL TOWER TEST Michael E. Watts

Rotorcraft Technology Branch NASA Ames Research Center

Moffett Field, CAUSA David ,Jordan Flight Experiments Branch NASA Ames Research Center

Moffett Field, CA USA

Abstract

In January of 1987, a whirl tower test of the Boeing Helicopter Model 360 rotor system was performed at Philadelphia, Pennsylvania. The Model 360 advanced high performance 4-bladed rotor system is composed of an articulated hub, blades that employ advanced airfoils, and a 3:1 taper on the outboard 10% of the rotor span. The tandem rotor Model 360 is designed to be a 200+ knot helicopter of primarily composite material construction.

In addition to the whirl-tower performance meas-urements, acoustic data were acquired from seven microphones. Three of the microphone locations were carefully selected to correspond with three geometrically similar locations of microphones from the Model360 1/5 scale wind tunnel test conducted in The Netherlands. Two of these microphones were located 3.0 rotor diameters from the rotor hub, one 6° and one 15° down from the rotor disk. The third similar position microphone was at 4.6 rotor diameters and 15° down from the rotor disk. Comparison of the results of the whirl-tower test with theory show that the theoretical predic-tion accuracy varies with microphone posipredic-tion and the inclusion of ground reflection. The prediction error, with ground reflection included, ranged from 0 to 40% of the measured signal peak-to-peak am-plitude. This peak-to-peak accuracy is on the order ofthat obtained by previous anechoic facility model scale comparisons of experiment and theory. Introduction

Acoustic prediction has become important in the design of rotorcraft as civil environments and battlefield detectability become increasingly im-portant issues. Just as stealth technology is incor-porated into aircraft at the design level in the form of reduced radar cross section, electronic emis-sions, and infrared signatures, reduced acoustic emissions have become a major design goal in the reduction of rotorcraft detectability and

commu-nity annoyance. To accomplish this task, predic-tive codes have become more sophisticated in recent years. However, to prove the validity and worth of these codes, accurate experimental data is essential. In many cases, rotor tests are de-signed with the primary goal of obtaining per-formance data with acoustics data being obtained as a target of opportunity. In these cases, the experimental setup is usually not optimized for acoustics data acquisitions to minimize back-ground noise and reduce reflected signals. Even when the effort is made to acoustically treat the experimental test stand, the ground surface still provides a reflection plane which distorts the data. To use such data for detailed waveform shape and amplitude research, the contribution of reflected signals must be addressed in theoretical predictions.

The Boeing Helicopter Model 360 front rotor was tested on the Boeing Helicopter Engineering Whirl Tower (BHEWT, Fig 1) located next to the Philadelphia International Airport. The primary purpose of the test was obtaining hover perform-ance data and an acceptperform-ance endurperform-ance test. Acoustics data were acquired during the whirl tower test as a target of opportunity.

Eotor Description

The 4-bladed rotors for the Model 360 (Ref l) are 49.7 ft (15.15 m) in diameter. The rotor blades incorporate airfoils from the VR-12 and -15 family of airfoil sections, the newest representatives of Boeing Helicopter's in-house airfoil design evolu-tion. The VR-12 (10.6% thick) section is used from the root cut-out to 85% radius, at which point the transition begins to the VR-15 (8% thick) section at the blade tips. A constant chord is used from 28% radius to 90% radius at which point a 3:1 taper begins and is continued to the blade tips. The twist distribution is +6.5 degrees at 23% radius, -1.28 degrees at 86% radius, and -3.7 degrees at the blade tips, with linear transitions between these stations. The concept of the ad-vanced rotor design was to utilize optimized

val-ues of blade taper and twist to improve hover per-formance while using the advanced VR-12 and -15 airfoil sections to improve cruise performance. J~xperimental Setun.

The rotor hub is mounted at a height of 50ft (15.24 m) above ground level on the BHEWT which has a conical structure. Electric motors power a spiral bevel gear and a four spf,ed gear box. '!'he gear box powers a drive shaft which in turn powers the rotor hub. For the whirl-test, rotor thrust, torque and speed were ·measured as woll as trim actuator loads. Rotor blade instrumentation con-sisted of 6 f1ap, 4 chord and 3 torsion strain gage bridges. In addition, blade pitch, flap and lead/lag angles, and pitch link loads were measured.

Rotor hover performance data was acquired during rotor speed sweeps at constant collective actuator settings of approximately 0, 8, 10, 11, 12, 13, 14, and H5 degrees, and a collective actuator sweep at a constant nominal rotor speed of 256 rpm. The thrust and torque data obtained indi-cated a maximum F'igme of Merit of 0. 758. De·· tailed results of the performance measurernents are presented in reference 2.

A seven~1nicrophone array was used for the acoustic portion of the test. The positions for these microphones (#1121 and i3) wnre geometrical1y sirnilar to positions used dur--ing the Duits-Neder--landse Windtunncl (DNW) test of a model scale 360 rotor (B.ef J) and two microphones (#4 and 6) matched positions ust~d chn·ing tests ofothnr rotors on the BHE\VT. 'I'he remaining microphones (#[) ancl7) wcTe set in an inverted position over ahnni--num ground p1anes. These _microphone positions are shown in figure 2 and the relation of the micro-phones to the DN\V positions -is shO\vn in table L The microphone~_; u:;ed \Vcre pre..::;.mpli.fler pow-creel, 1/2 inch condenser -Lypc1 cac"\1 covered w·iLh n

foarn w_indscreen. /\_ -~ 1i_-·Lrnck, J·inch ern) Lape

recorder was used nt a tape ~;.peed of 30 i.n/scc (0. 762 rn/scc) to record the rnlerophonc signals. 'J'hc antp!ificrs in the reeordt:r were E>ct. for Inter-· Hange Instrumentation Grou.p (HUG) ·intcrrncch~ ate band FM recording./\ frequency rnsponse ca.li-bration of the tape recorder produced a curve that was flat to within !.: 1 dB from {) .. ]{)kHz. Single frequency ealibrations at 1,000 Hz, 114 dB were performed at the beg·inning and end of the test. Along with acoustic data, a rotor 1/rev signal, voice inputs, and a time--code signal were recorded.

Ambient conditions during the acquisition of acoustic data consisted of low winds and an aver .. age temperature of33 F (0.6 C) with snow covering the ground to a depth of approximately 1ft (0.3 m).

Theoretical Modeling

The theoretical code used for the prediction of generated acoustics for this paper is called Rotor Noise (RTN) and was developed by Aggarwal and Schmitz (Ref 4). '!'he theory used in RTN is derived from first principles based on the Ffowcs Williams and Haw kings equations as found in reference 5. (See reference 6 for details of this derivation and implementation into RTN.)The RTN code as devel-oped in reference 6 had no capacity for the inclu.· sion of the effects of reflected \Vaves on the resul-tant wave as seen by the observer or microphone. The correb.tion of ln'N with the DNW Model 360 hover data performed in reference 4 showed good correlation of theory and experiment for low- to medium- hover tip Mach numbers.

Acoustics prediction depends on the quality of the input lift and drag distributions. The Analyti·· cal Methods Incorporated (AMI) free-wake lifting surface code (Ref 7) was used for generating the input distributions since no experimental values were available. '!'he lift coefficients were obtained by adjusting the input collective angle to match the measured thrust. The integrated drag distri-bution was corrected to match the measured torque and this correction was applied to the drag distri-bution in HTN.

Acoustic data acquired in non-anechoic facili-ties such as the BH!~WTwill have reflections which modify the measured acoustic wave. Even near-anechoic facilities such as the DNW will have reflections in certain microphone locations, as noted in reference 4. Because of these considera-tions, the HTN code was modified to include re-flected paths. Normally, attenuation of the rc·· fleeted signal must be included to propedy model the absorption characteristics of the reflecting surface. As noted previously, the test was per·· formed w_ith one foot of snow cover over frozen ground. Albert and Orcutt. have shown, in refer .. cnce 81 that th1s ground cover condition will result ·in no signal attenuation below 35Hz at the micro-phone distances used during the whirl test. Since th(~ ptlmary contributors to the hovering acoustic wave shape and amplitude are the 4- and 8- per rev frequencies, nominally 17.6 and 35.2 Hz, no at.tmmal;ion ofthe reflected signal was included in the results used for this paper.

The acoustics data acquired was analyzed us-ingl:.he Macintosh-basedAcoustic Laboratory Data Acquisition/ Analysis System (ALDAS, Ref9). This system has one 12- and two 16- bit analog-to-digital cards, which are controlled by the ALDAS

program written at NASA Ames Research Center. The data were sampled at 2048 samples per revo-lution which resulted in sample rates of approxi-mately 9000 samples per second. All data were analog filtered at 2500 Hz before sampling to prevent aliasing.

Figure 3 (a and b) shows the variation in the wave shape of the raw data for one test condition (M,,P=0.631 and C,Jcr=0.0797). The variation of the signal is due to the unsteadiness of the flow field and the high background noise of the BHEWT. Since the hover condition is primarily a low fre-quency phenomenon, the acoustic data was fil-tered at a relatively low frequency. The cutoff fre-quency of this digital low-pass filter was deter-mined by first averaging a data trace for 64 cycles, and then performing a spectral analysis of the resultant cycle. As can be seen from figure 4, the point where the curve-fit to the first 24 harmonics intersects the medium line for the signal out to 1000 Hz is at 115 Hz. All data was digitally low-pass filtered at this frequency before cycle averag-ing was performed. Data was then cycle averaged for at least 32 cycles and usually 64 cycles. Figure 5 shows the effect ofthis process on the same data as shown in figure 3.

Test points in the acoustics portion of the whirl test fell within the range of0.354 to 0.665 hover tip Mach number and 0.0469 to 0.1223 C,Jcr. Efforts

were made to match test parameters from the

Army/Boeing/NASA DNW test (Ref 3). Figure 6 shows the test matrix as a function of hover tip Mach number versus C,Jcr. The solid line indicates the bounds of the data taken in the DNWhovertest and the diamonds show the actual test points ofthe whirl tower test. Data indicated by solid diamonds are discussed in detail below.

Discussion of Results

This paper will concentrate on data for a range of thrust conditions at two hover tip Mach num-bers, 0.543±0.004 and 0.630-±0.001, which are indicated by the solid diamonds in figure 6. The bandingofthe Mach numbers presented is because ofthe courseness ofthe speed control oftheBHEWT. In addition, the three microphones which closely matched locations from the DNW test will be used to provide commonality. Both wave shape and peak-to-peak amplitude comparisons of experi-ment with theory will be made.

Averaged experimental and predicted waveforms are presented in figures 7 thru 12. Both the direct path and direct plus reflected (dual) paths are also presented to show the effects of the reflection on

the wave shape.

The reflected wave causes the calculated wave-form to more closely match the positive and nega-tive slopes of the experimental waves for micro-phones two and three. This effect is evident for all conditions with the most benefit being seen in the lower thrust conditions, and degrading slightly for the higher thrusts. Higher frequencies not predicted in the theory are evident in the funda-mental wave shape of the experifunda-mental data. These frequencies tend to mask the wave shape for the higher thrust conditions. Figure 13 shows a typical comparison of the spectra between a high and low thrust case. The low thrust spectrum dies out at about 80Hz and the high thrust spectrum continues to above the filter frequency (frequency contents above 115 Hz are filter roll-off effects). Figure 14 shows a typical effect of filtering the experimental data at the 8/rev frequency (35 Hz) for microphone 3, M" of0.630 and C,Jcr of0.1103. Even though this filtering level shows a closer match of the theoretical and experimental curves, it is not applied to the data since it masks the higher frequency effects contained in the data and can be considered "tailoring" the results.

The phase of the reflected wave for microphone one forces a double-humped shape which is com-pletely different from that of microphones two and three. This hump is not as obvious for the lower as for the higher Mach number cases. The correla-tion with theory ranges from good, as seen in figure 10, to evident but obscure as seen in figure 12. It appears that the data for this microphone have high frequency data overlaying the signal as discussed above. However, filtering the data at 35 Hz does not improve the correlation significantly. The other consideration in comparing theory and experiment is the peak-to-peak amplitude of the signals. Figure 15 presents the peak pressure differences for the experimental and the dLwl-path theoretical data for microphones 1 thru 3. In examining this figure, similar results are to be noted as with the wave shape comparisons. In other words, microphones two and three had very good correlation at the low Mach number and good agreement at the higher Mach number. Micro-phone 1 had fair agreement for both Mach number conditions.

Figure 16 is presented to illustrate the im-provement in the peak-to-peak amplitudes with the addition of the reflected wave. This figure shows the percent error of the direct-path and dual-path predictions with the measured value.

P ercent Error = (Experiment- Theory) x 100

In this figure, a positive percent error represents an under prediction of amplitude. Notice that the change from the open (direct path) symbols to the closed (dual path) symbols is towards zero percent error for microphones two and three. Even though the change in curves may go through zero to the over-prediction side, this trend indicates that the addition of the reflection has a beneficial effect on the prediction ofpeak-to-peak amplitudes fm·these locations. Microphone one has the opposite trend indicating that the reflection is driving the ampli-tude away from the correct value. This inconsis .. tency of microphone 1 trends with those of micro-phones two and three was also seen in the wave shape comparisons.

As mentioned previously, the microphone 3 po-sition closely matches the 3.0 diameter, 15° down microphone from the DNW test. Two papers have previously compared acoustic theory to DNW 115

scale model data for this microphone location (Ref 4 and Ref 10). Both papers used the AMI code to generate the loads distribution for input into the theoretical codes. Reference 4 used the RTN code and reference 10 used the Rotor Acoustics Predic--tive Program (RAPP). Neither application included reflection effects. The DNW test Mach number used was 0.636 which is slightly different than the full scale 0.63 from the whirl test. To eliminate the effects ofthis difference, the results are presented in figure 17 as percent error of the peak-to-peak values. The measured value in this figure is the ex-perimental value from the test being predicted. This figure shows that the addition of the reflected wave in the whirl tower data brings the accuracy of the peak-to-peak prediction into the same error Tange as that for ETN and RAPP compared with data from the anechoic DNW.

'l'he results pres<mted show that the addition of a ground reflection into the predicted acoustics

curves can have a sig-nificant effect on the results.

For a normal hover tip Mach number, the addition

of a ground reflection yields accuracies for the non~ anechoic whirl--tower test data on the order of those seen for the anechoic DNWtest. The common level of error could be the result of incorrectly predicted loads or limitations in the theoretical acoustics model. Application of the reflection cor-rection must be used carefully as its effects arc dependent on the phase of the reflection and the damp-ing of the reflecting surface (Ref 8).

References

1. Hartman, L. J., Mecklin, R., and Wiesner, R., "Boeing Model 360 Advanced Technology Helicopter Design Features and Flight Test Update", Presented at the 44th Annual Na-tional Forum ofthe American Helicopter So-ciety, June 1988.

2. ''Whirl Tower Test of an Advanced High Pm·-fonnance (Model 360) Rotor", Boeing Heli--copter Final Report, NASA Contract NAS 2-11966, Ames Research Center, Moffett Field, CA, 1988.

3. Dadone, L., Dawson, S., and Ekquist, D. "Model 360 Hotor Test at DNW -Review of Performance and Blade Air load Data", Pre-sented at the 43rd Annual Forum of the American Helicopter Society, 1987.

4. Aggarwa.l, H.R., Schmitz, F.H., and Boxwell, D.A., "Prediction and Measurement of Low-Frequency Harmonic Noise of a Hovering Model Helicopter Rotor", Presented at the 45th Annual National Forum of the Ameri-can Helicopter Society, May 1989.

5. Ffowcs Williams, J.E., and Hawkings, D.L., "Sound GemJration by Turbulence and Sur-faces in Arbitrary Motion", Philosophical I'rmLOill..tio.!l.!l..Q[J;b.D_l~G.i.\l.ty of Lon don,

SIT.k.:LA:_l\1a.tl.lilJmltiw.ll ....

umL..PhY~fui:

=<>

4 Vol. 264, 1969, pp. 321-342.

6. Aggarwal, H.R., "Low-Frequency Helicopter Acoustic Prediction", Final H.eport, Aerof-lightdynamics Directorate Contract NAS 2-12239, Ames Research Center, Moffett Field, CA, 1988.

7. Su.mnw1 tLM., and l'vlaskew, B., HA Surface

Singularity Method f(lr Rotors in Hover or Climb", US.M.VltADCOMTRSI-D-2:3, Decem-ber 1981.

8. Albert, D.G., and Orcutt, J.A., "Acoustic Pulse

Propagation Above Grassland and Snow:

Co·mparison ofTheoretical and

Experimen--tal Waveforms", Jgurna 1 ll.L..t!Jg__6£.QJ.!O..ti.c.&ll. S.Os.i.~.ty_of A!)leric.a, Vol. 87, No. l, January

1990, pp. 93-100.

9. Watts, M.R, "ALDAS Users Manual", NASA TM 102831, 1990.

10. Gallman, J.H., ''The Validation and Applica-tion of a Rotor Acoustic PredicApplica-tion Computer Program", Proceedings ofthe 1990 Army Sci-ence ConferSci-ence, June 1990.

Table I. Common whirl tower and DNW micro-phone locations.

BHEWTBV360 DNWBV360

Mic Radial Downward Mic Radial Downward

Location Angle Location Angle

(Dia.) (') (Dia.) (')

1 2.99 5.85 15 3.0 6

2 4.62 6.05 22 4.6 6

3 2.99 15.1 16 3.0 15

Fig 1. Boeing Helicopter Engineering Whirl Tower ROTOR HUB ~ -2.99Dia. _ ~ ~ 10 6

"'

~

2

"'

a._

-2 ..J Q.

en

-6 -10 10 6

"'

~

2

"'

"'

c.._ -2 ..J tl. (fJ -6 ·10 0 a) 0.2

0.4

0.6 O.B One Revolution

Fig 3. Examples of variation of raw data with time. Mlc1 1 2.99 Dia. _{-4.62 Dia._ Mlc 2} ~ ~ ~ Mlc3 ~ Mic ~ 1.2m 1.2m

90 (/) 80 (ii () 1/) 70

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0..

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60 '

w

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w

50

a:

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·6

(/) 0 50 100 150 200 250 300 Frequency, Hz ·1 0 L..l....J...Jc...J..-1--J.--'-l-J..-1-I-L-1-J--l-.-1...!.--'-!-.l 0 0.2 0.4 0.6 0.8 One Revolution

Fig 4. Spectrum used to determine filter fre-quency cutoff value.

Fig 5. Data after filtering and averaging

6 4 !!!. 2 0.14 0.12 b 0.1 j::: u 0.08 0.06 0.04

<>

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0.35 0.4 0.45 0.5 0.55 0.6 0.65 0.7

Hover Tip Mach Number

Fig 6. Data points taken for acoustics test.

--Experiment

- -Theory- Direct Path ---Theory- Dual Path

Microphone Positions (not to scale) Mlc 1, ·6° 3.0Dia.

~

"' 0 ~~-'"""~-=~"> 0.. -~- ·2 0.. Mlc 2, ·6° 4.6 Dla. (/) -4 •6 '-'--1-l....L-1-'--'---'--J..L-1-L.l.---'--.l...L-'--l-L--'

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Fig 7. Experiment and theoretical wave shapes, Mu, = 0.544, C,Jcr = 0.0531.

!1.3.2.6

0.8

1

6 4 !!l 2

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0..

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- -Theory- Direct Path

---Theory- Dual Path

0 One Revolution Microphone Posilions (not to scale) Mic 1,-6° 3.0 Dla. 0.2 0.4 0.6 One Revolution

Fig 8. Experiment and theoretical wave shapes, M"'

=

0.544, C₁/cr

=

0.0837.

--Experiment

- -Theory- Direct Path Microphone Positions _{(not to scale)}

----··Theory- Dual Path

Mic 1,-6°

·-~~~

-<----3.0Dia.

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-6 1-l...J-I-J..--l-!...L...l.-L....L--l-'--1.--'-'-'---'-'--'--'

t

I I I I I I I I I I I

0 0.2 0.4 0.6

0.8

1 0 0.2 0.4 0.6

One Revolution One Revolution

Fig 9. Experiment and theoretical wave shapes, M"'

=

0.544,

Cylcr

=

0.0951.

Mic 2,-6° 4.6Dia.

0.8

I I I I 0.8 1 I 1

6

--Experiment

Microphone Positions

4

_-

-Theory - Direct Path

(not to scale)

"'

2 -- -- - Theory - Dual Path

(;j _{Mlc 1,-6°} 0 3.0Dta.

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0

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if

-y

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One Revolution _{One Revolution}

Fig 10. Experiment and theoretical wave shapes, Mtip = 0.630, C.rfcr = 0.0469.

6 4 !!!. 2

~

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en

-4 --Experiment

- -Theory- Direct Path

- - - Theory - Dual Path

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2

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-6~~~~~-L~~~~~~~~ 0 0.2 0.4 0.6 0.8 1 One Revolution 0 Microphone Positions (not to scale) Mlc 1,-6° 3.0Dta. \ 0.2 \ 0.4 0.6 One Revolution

Fig 11. Experiment and theoretical wave shapes, Mtip

=

0.630, C.rfcr

=

0.0797.

II.3.2.8 Mtc 2,-6° 4.6 Dia. / 0.8 1 / 0.8 1

6 4

'*l

2

!il

0 ,/ :'

~

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-2 0..

en

-4 --Experiment - -Theory - One Path - - - Theory - Two Path

Microphone Positions (not to scale) Mlc 1, ·6° /. _{. , 3.0 Dla.}

v· ·r._

~---:~~~~~~~~~~~~~

4 :· _:'. !!l 2 \ Mlc 2,-6° 4.6Dia . / \'

j

0

\. \ I /. '-",

...r

·2 0..

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·4 -s~~~~~~~~~~~~ 0

0.2

0.4

0.6

0.8

1 One Revolution 0

0.4

0.6

0.8

One Revolution

Fig 12. Experiment and theoretical wave shapes, M,;. = 0.630, C.jcr = 0.1103.

1

0.8

_{" - - L o w Thrust} "

"'

0.6

iii : ---High Thrust

!il

"'

6 4 !!l

2

_' ' ' ' ' ' ' _•, \': --Experiment, 35 Hz Filter - -Theory • Direct Path • • • • • Theory • Dual Path

I:

1 0..

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0

0.. I _{~ /u'} . 1., /, 0..

en

0.2

.,

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50

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150

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-:=

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~ 2 ::I

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1 ~

c..

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₀

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(I) 0.04

c..

a) 3.0 Dia., 6° Down ...-e-Mach 0.544 Measured - - • - Mach 0.544 Dual Path -a-Mach 0.63 Measured - - • - Mach 0.63 Dual Path

...

·---~

~---·----·

b) 4.6 Dia., 6° Down

---·-·

... c) 3.0 Dia., 15° Down

••

0.06 0.1

•••

-··

•

0.12

Fig 15. Effect of reflected path on peak pressure differences.

60

40 ~ ₂₀

g

w

~

0 ~ -20 -40 60 40 ~ 20

g

w

~

0

1:!

8?.

·20 -40

60

40 ~ 20

g

w

E 0

~

8?.

-20 ·40

·60

0.04 a) 3. 0 Dia., 6° Down

·T·

...-e-Mach 0.544 Direct Path -- • - Mach 0.544 Dual Path -a-Mach 0.630 Direct Path -- • - Mach 0.630 Dual Path

b) 4.6 Dia., 6° Down 0 0 E)

·----·--

---

_-.-.

·---·----·

_·---.

c) 3.0 Dia., 15° Down

~/)

_EY

_EJ

·-·---.

0.06

-

.•

•·~-;

..

~

·-··

----·

0.08 CT/0' 0.1 0.12

Fig 16. Effect of reflected path on percent error.

100 75 50

g

25~~~~~--~---0

w

--- .,':,.

5l

0 ---·---· CJ

t

·25

D. -50 -75

-&-RTN, BHEWT, Direct Path - - . - RTN, BHEWT, Dual Path

~RTN, DNW, Direct Path (Ref 6)

ll RAPP, DNW, Direct Path (Ref 9)

·100

0.04 0.06 0.08

CT/cr

0.1 0.12

Fig 17. Comparison of percent error with previ-ous investigations for microphone 3 (3.0 Dia., 15° down) at M",~ 0.63.