I NASA 

Technical Memorandum 4024 

AVSCOM 

Technical Report 87-B-4 

Acoustic Measurements From a 

Rotor Blade-Vortex Interaction 

Noise Experiment in the 

German-Dutch Wind Tunnel (DNW) 

R. M. Martin 

Langley Research Center 

Hampton, Virginia 

W. R. Splettstoesser 

Deutsche Forschungs- und 

Versuchsanstalt fur 

Luft- und Raumfabrt 

Braunschweig, West Germany 

National Aeronautics 

and Space Administration 

J. W. Elliott 

Aerostructures Directorate 

Langley Research Center 

Hampton, Virginia 

USAARTA-A VSCOM 

K.-J. Schultz 

Deutsche Forschungs- una? 

Versuchsanstalt fur 

Luft- und Raumfabrt 

Braunschweig, West Germany 

Scientific and Technical 

Information Division 

1988 
Contents 

Abstract ................................... 1 

Introduction ................................. 1 

Symbols ................................... 1 

Experimental Apparatus and Procedures ..................... 2 

WindTunnel ................................ 2 

Model Rotor Test Apparatus ......................... 2 

Model Test Stand and Support ....................... 2 

ModelRotor ............................... 3 

Rotor Performance Data Acquisition and Reduction .............. 3 

Matrix of Test Conditions .......................... 3 

Acoustic Instrumentation .......................... 4 

Microphone Array Traverse System ...................... 4 

Fuselage-Mounted Microphones ........................ 4 

Acoustic Data Reduction and Analysis ..................... 5 

Data Quality ................................ 5 

Tunnel Flow Quality ........................... 5 

Rotor Performance Data .......................... 5 

Background Noise ............................. 5 

Acoustic Reflections ............................ 6 

Acoustic Shielding by Fuselage ....................... 7 

Steadiness and Repeatability of Rotor Acoustic Data ............. 7 

Testing Procedures and Measurement Accuracy ................. 7 

Presentation of Data ............................. 8 

References .................................. 8 

Tables .................................... 9 

Figures ................................... 11 

Averaged Acoustic Signals ........................... 23 

PRECEDING PAGE BLANK NOT ME0 

iii 
Abstract 

Acoustic data are presented from a 40-percentscale model of the four-bladed BO-105 helicopter 

main rotor, tested in a large aeroacoustic wind tunnel. Rotor blade-vortex interaction (BVI) noise data 

in the low-speed flight range were acquired using a 

traversing in-flow microphone array. The experimental apparatus, testing procedures, calibration results, 

and experimental objectives are fully described. A 

large representative set of averaged acoustic signals 

are presented. 

Introduction 

The primary objectives of the test program reported here were (1) to acquire a comprehensive data 

base to improve the definition of the directivity patterns of rotor blade-vortex interaction (BVI) impulsive noise, (2) to investigate more fully retreatingside BVI noise, and (3) to improve knowledge of the 

locations of BVI noise sources. Secondary objectives 

of the program were (1) to access the far field of 

rotor BVI noise, (2) to extend the investigation of 

reference 1 into the scalability from model- to fullscale of the BVI signal at high advance ratio, (3) to 

explore the parametric effects of advance ratio, tippath-plane angle, and thrust coefficient on BVI noise, 

and (4) to acquire a noise data base for future development of a BVI noise prediction methodology. 

The test was performed in the open test section of the Duits-Nederlandse Windtunnel (DNW), 

a large aeroacoustic test facility, utilizing a fourbladed, 40-percent dynamically scaled model of 

the Messerchmitt-Bolkow-Blohm GmbH BO- 105 helicopter main rotor. The experiment was performed jointly with research personnel of the German 

Aerospace Research Establishment, the Deutsche 

Forschungs- und Versuchsanstalt fur Luft- und 

Raumfahrt (DFVLR). The model rotor and the rotor 

test stand were provided and operated by research 

personnel of the Rotory Wing Aircraft Branch of 

the DFVLR Institute of Flight Mechanics. Acoustic instrumentation support was provided by research 

personnel of the Technical Acoustics Division of the 

DFVLR Institute of Design Aerodynamics. 

High quality acoustic data were acquired due to 

the excellent fluid dynamic and acoustic properties 

of the DNW. In particular, the tunnel has quite low 

in-flow background noise and turbulence levels and a 

sizable anechoic open test section. The importance 

of low turbulence airflow on rotor acoustic testing 

is stressed in reference 2, a comparison of acoustic 

results from a model rotor tested in both the DNW 

and the French CEPRA-19 tunnel. 

~ ~~ 

An extensive range of in-flow acoustic measurements were acquired in a plane underneath and upstream of the rotor model using a traversing ninemicrophone array. A digital acoustic data acquisition system was employed to guide the choice of the 

test matrix and identify optimum measurement locations by using on-line identification of strong BVI 

noise test conditions. The BVI directivity was investigated by using an array of microphones traversing 

in a large plane under and upstream of the rotor. A 

high quality data base for acoustic localization techniques was acquired through a large range of measurement locations. The occurrence of BVI acoustic 

signals created on the retreating side of the rotor was 

pursued by positioning the microphone array under 

and downstream of the rotor, in positions previously 

found to best measure the phenomenon (ref. 3). 

This paper addresses the general philosophy and 

experimental approach of this program. Test results 

on acoustic directivity, retreating-side BVI, and BVI 

source locations have been presented in references 4 

to 6. In the present paper, the model test rig, instrumentation, data acquisition and reduction, acoustic 

calibrations, and experimental procedures are fully 

documented. A large representative set of the averaged acoustic signals are presented. 

Symbols 

CT 

fc 

T 

MH 

R 

T 

V 

XW 

YW 

ZW 

ashaft 

aTPP 

rotor thrust coefficient, T/.rrpwR4 

filter cutoff frequency, Hz 

radial distance from rotor hub to 

microphone (see fig. l), m 

hover tip Mach number 

rotor radius, 2.0 m 

rotor thrust, N 

flow speed, m/sec 

streamwise location of traversing array 

relative to hub, positive upstream 

(see fig. l), corrected for streamwise 

movement of hub due to model pitch, 

m 

cross stream location of traversing 

array microphones relative to hub, 

positive under advancing side (see 

vertical location of traversing array, 

positive above hub (see fig. l), m 

angle of rotor shaft from vertical, deg 

rotor tip-path-plane angle, positive 

nose up (see fig. l), deg 

fig. 11, m 
I Af spectral bandwidth, Hz 

8 polar angle, angle centered at hub 

from tip-path plane to microphone 

location, positive down (see fig. l), deg 

P advance ratio, VIwR 

P density, kg/m3 

1ct azimuth angle in tip-path plane from 

rotor hub to microphone location, zero 

over tail, 90' over rotor advancing side 

rotational speed of rotor, cycles/sec 

I (see fig. l), deg 

W 

Abbreviations: 

ATD analog- to-digi tal 

BVI blade-vortex interaction 

DFVLR Deutsche Forschungs- und Versuchsanstalt fur Luft- und Raumfahrt 

DNW Duits-Nederlandse Windtunnel 

(German-Dutch Wind Tunnel) 

I DTA digital- to-analog 

I Mic microphone 

rms root-mean-square quantity 

r Pm revolutions per minute 

SPL sound pressure level, dB referenced to 

2 x Pa rms 

Experimental Apparatus and Procedures 

Wind Tunnel 

The test was performed in the open test section of 

the Duits-Nederlandse Windtunnel (DNW), located 

in the North East Polder, The Netherlands. The 

DNW is a subsonic, atmospheric, closed circuit wind 

tunnel with three interchangeable, closed test section 

configurations and one open configuration. The open 

configuration employs an 8 x 6 m contraction section, 

shown in figure 2, and a 19-m-long test section, surrounded by a large anechoic hall of about 30000 m3 

lined with absorptive acoustic wedges. The tunnel 

has excellent fluid dynamic qualities, fully described 

in references 7 to 10. 

Model Rotor Test Apparatus 

Model Test Stand and Support 

The DFVLR rotor test stand (refs. 11-13) is 

shown installed in the DNW open test section in figure 3. Details of the rotor test stand are given in 

figure 4 and table I. The stand consists of three major subsystems: the hydraulic drive system, the rotor 

balance system, and the rotor control system. During this test, the stand was housed within an acoustically insulated fiberglass shell and was attached 

to the computer-controlled, hydraulic sting support 

mechanism. 

The rotor drive system consists of a seven-piston 

axial drive motor connected by hydraulic lines to 

a remotely located electrically driven pump. The 

axial hydraulic motor is connected directly to the 

rotor shaft and has a power capability of 100 kW at 

1050 rpm. 

The rotor balance system is a six-component 

balance containing separate measuring elements for 

static and dynamic load components. The balance 

consists of a lower and an upper plate connected 

through transmission rods. Seven load cells on the 

lower plate measure the static components. The dynamic components are measured by means of piezoelectrical transducers connected to the transmission 

rods. A separate load cell is provided to measure 

the rotor torque and is located between the upper 

and lower balance plates. This ensures that no undesired coupling between the rotor balance and the 

drive shaft occurs. 

The rotor control system is of a swashplate system consisting of three electric actuators attached 

to the upper plate, the swashplate, and the rotating 

blade pitch control rods. The actuators provide collective and cyclic blade pitch control by moving the 

nonrotating part of the swashplate and are remotely 

controlled. A potentiometer at the root of one of the 

blades measures the blade angle of incidence during 

rotation. 

The fuselage is constructed of a 4-mm-thick reinforced fiberglass shell covered on the inside with 

a heavy 3.5-mm-thick layer of sound-damping . material and a 30-mm-thick layer of open cell soundabsorbing foam. The outer side of the fuselage is 

lined with 50-mm-thick open-cell foam to minimize 

acoustic reflections. The overall sound pressure level 

transmission losses through the fuselage were 27 dB 

for simulated hydraulic drive noise and 30 dB for 

white noise. To maximize the transmission loss, special care was taken to shield both the gap between 

the rotating rotor head and the top of the fuselage 

and the gap between the fuselage and sting. 

The model rotor rig was supported by the 

computer-controlled, hydraulically actuated sting. 

Prior to the test, a thorough analysis was made of 

the dynamic behavior of the rotor test rig and sting 

combination to ensure that a ground resonance problem would not be encountered. The sting was covered with a streamlined sound absorptive lining to 
minimize acoustic reflections. The sting control system was programmed to keep a model reference point 

at a preselected height in the test section, for the 

range of angle of attack (ashaft). However, it was 

not possible to maintain the reference point at a ked 

streamwise location for the range of angle of attack. 

Model Rotor 

The rotor is a 40-percent, dynamically scaled 

model of a four-bladed, hingeless BO-105 rotor 

(fig. 5). The rotor has a diameter of 4 m with a root 

cutout of 0.350 m and a chord length of 0.121 m. 

The rotor blade is constructed of an NACA 23012 

airfoil, with the trailing edge modified to form a 

5-mm-long tab, to match the geometry of the fullscale rotor. The rotor blades have -8' of linear twist, 

a standard square tip, and a solidity of 0.077. Further details are given in table I and reference 11. 

The nominal rotor operating speed was 1040 rpm, 

giving an acoustic blade-passage frequency of about 

70 Hz. The nominal hover tip Mach number was 0.64. 

The rotor blades are made of glass-fiber-reinforced 

plastic and have essentially the same mass and stiffness distributions as the full-scale rotor. However, 

the blade chord length (0.061R) is a slightly larger 

scale than the full-scale chord (0.054R) in order to 

maintain the proper Locke number scaling (ratio between the aerodynamic forces and the mass and elastic forces, ref. 11). Each blade is equipped with a 

small tab located at 70 percent of the radius, measuring 60 x 10 mm (shown in fig. 5). These tabs, in addition to the pitch rods, are used to adjust the track of 

each blade in the rotor system. Strain gages installed 

near the root of each blade are used to measure the 

blade flapping, lagging, and torsional moments. 

Rotor Performance Data Acquisition and 

Reduction 

The rotor performance data were obtained from 

the rotating and fixed systems of the DFVLR test 

stand control system (ref. 12). The rotating system 

acquired the signals of the strain gages on the rotor 

blades (flapping, lagging, and torsional moments), 

on the control rods (pitch link forces), on the rotor shaft (bending moment), and on the blade root 

potentiometer (blade pitch angle). 

The rotating signals were transmitted by two 16- 

channel pulse-code-modulated (PCM) encoders on 

the rotating rotor head. The data were multiplexed, 

converted to a digital form, and then formatted into 

two serial data streams for transmission through slip 

rings. Subsequently, both serial data streams were 

reformatted into parallel signals by two decoders 

followed by digital-to-analog (DTA) conversion and 

filtering. The rotating data were then merged with 

the fixed system data via a 64-channel encoder. 

The ked system also has a 32-channel capacity. Signals from the displacement transducers, the 

static and dynamic balance transducers, the shaft 

torque transducer, the shaft rpm encoder (360-perrevolution), and the shaft position encoders (onceper-revolution and 512-per-revolution) were transmitted to a PCM data acquisition system. The data 

acquisition system consists of a central processor 

unit, a coprocessor, a random access memory unit, 

and a magnetic tape recorder and has a useful frequency range up to 250 Hz. 

Details of the rotor data processing are given in 

reference 13 and are summarized here. For all signals, 

20 revolutions of time domain data were written to 

digital tape. For on-line analysis and control the 

central processor performed a Fourier analysis of the 

data of a single revolution. A second processor fed 

with selected analog input of the ked and rotating 

system provided continuous on-line display of the 

rotor control parameters: hover tip Mach number, 

thrust, hub moments, blade bending, and pitch link 

loads. The final processing of the rotor performance 

data was done off-line by averaging the results of 

the 20 revolutions of stored data. These results 

include such parameters as collective pitch, cyclic 

pitch, coning angle, rotor power and thrust, and 

both the steady and the dynamic components of the 

rotor forces, moments, and longitudinal and lateral 

flapping angles. 

Matrix of Test Conditions 

A nominal test plan was developed to investigate 

the three primary test objectives and concentrated 

on two thrust coefficients, 0.0044 and 0.0030, and 

the forward speed range of 20 to 40 m/sec (p = 0.09 

to 0.18) at a range of tip-path-plane angles for constant nominal hover tip Mach number of 0.64. The 

value of 0.0044 is the nominal thrust coefficient for 

the full-scale BO-105, while the value of 0.0030 was 

chosen as a lower limit of thrust level. To investigate 

the secondary test objectives, a smaller set of conditions included velocities of 50 and 60 m/sec, thrust 

coefficients ranging from 0.0036 to 0.0056, and hover 

tip Mach numbers ranging from 0.56 to 0.72. 

To determine the optimum test conditions for 

strong BVI noise generation and the optimum measurement locations for the traversing microphone array, an initial exploratory approach was used. To 

minimize testing time during this exploratory phase, 

acoustic data were acquired with an on-line ATD system and were not recorded on analog tape. The 

on-line analysis performed during this exploratory 

3 
matrix was used to fine-tune the final, more extensive, test matrix, for which acoustic data were 

recorded on analog tape for postprocessing. Data 

from the initial exploratory matrix are presented in 

references 4 and 6, and the final expanded test matrix is presented in table 11. The data presented here 

are the postprocessed acoustic signals from the final 

expanded matrix. 

Acoustic Instrumentation 

The acoustic instrumentation consisted of a ninemicrophone in-flow array mounted on a traversing 

system and two in-flow microphones mounted on the 

rotor fuselage as shown in figure 3(b). The microphones were %-in. pressure-type condenser microphones equipped with standard “bullet” nose 

cones. Each complete microphone system (microphone, preamplifier, and adapter) was calibrated using the electrostatic actuator method to document 

its frequency response. This response was found to 

be flat between 5 and 5000 Hz, sufficient for the frequency range of the BVI noise (approximately 500 to 

4000 Hz for model scale). 

Signal conditioning of the microphones was performed by computer-controlled amplifier and filter 

systems. The amplitude and phase characteristics 

of the two-pole Buttenvorth filters were investigated 

before the test for high-pass filtering at three nominal cutoff frequencies. The amplitude response is 

flat in the passband. The phase shift is 180’ at the 

cutoff frequency fc, is approximately 40” at 5 fc, and 

negligible at approximately 10fc. The filters were 

found to be well matched in their phase characteristics, within 1’ in the passband. 

The acoustic data, the rotor once-per-revolution 

and 512-per-revolution signals, and a time code were 

recorded on a 14-track analog frequency-modulated 

recorder. The data were recorded at a tape speed 

30 inches per second, which yielded a useful frequency response to about 20 kHz. 

To verify the frequency response of the entire 

measurement system in situ (from preamplifier to 

tape recorder), a white noise signal was inserted 

simultaneously into each preamplifier and recorded 

on tape. The auto and cross power spectral densities 

of the white noise signal were checked to ensure 

flat frequency response and to document phase lag 

between channels. 

The digital amplifier and filter settings were 

transmitted to a second, larger computer. During 

testing, up to 10 microphone channels could be digitized using a 12-bit, high-speed simultaneous analogto-digital (ATD) converter. The computer was employed to control the ATD converter, to store calibration and amplification factors, and to perform on-line 

time domain and spectral analysis. A two-channel 

fast Fourier transform analyzer was also employed 

for quick-look analysis during testing. 

Microphone calibrations were performed using a 

sound level calibrator at the beginning and end of the 

test. All microphone gains were adjusted to produce 

500-mV rms output for an acoustic level of 114 dB at 

1000 Hz. In addition, a daily calibration of a 1000- 

Hz, 500-mV rms signal was recorded on all channels 

by signal insertion at the tape recorder inputs. 

Microphone Array Traverse System 

The microphone array traverse system consisted 

of a horizontal wing with its span normal to the flow 

direction and a traverse system with a total range 

in the flow direction of 8.2 m (7 m upstream, 1.2 m 

downstream of the hub, see fig. 3). The microphones 

were arranged symmetrically with respect to the 

tunnel centerline, spaced 540 mm apart. The array 

vertical position was nominally 2.1 m below the rotor 

hub, although a small set of data was acquired at 1.6 

and 2.6 m below the hub. 

The microphone holders employed a “soft” 

vibration-isolating mounting. A cross section of the 

wing is shown in figure 6. The total height of the 

entire structure was such that only the streamlined 

wing support struts crossed through the shear layer. 

The wing and support struts were covered with a 

25-mm-thick, open-cell foam airfoil section; and the 

supporting structure was covered with a 100-mmthick foam lining except at the base which was covered with 800-mm foam wedges. 

The traverse mechanism was powered by a 

variable-speed dc electrical motor. Control and positioning was obtained with a servo position controller. 

The positioning accuracy was 2 mm. Prior to data 

acquisition, the alignment of the traversing array was 

calibrated over its range of motion. The alignment 

in the cross flow direction was f2 mm; in the vertical direction, less than fl mm. The deflection of 

the traversing array due to aerodynamic drag was 

on the order of 9 mm in the streamwise direction 

at 60 m/sec. The accuracy of this correction is estimated to be f3 mm. These deflections were calculated off-line and applied to the measured array 

position data to correct the calculated microphone 

positions. 

Fuselage-Mounted Microphones 

Two microphones were mounted on the fuselage, 

one under each side of the rotor disk, at azimuth 

angles of approximately 90’ and 270’. The microphones were mounted in vibration isolation mounts 

I 4 
similar to the mounts for the traversing array microI phones (fig. 6). These microphones were mounted 

with a small negative angle (approximately 5') to 

the tunnel free stream. This was an attempt to minimize wind-induced background noise due to the freestream velocity, because most of the data were at 

positive rotor shaft angles. These microphones are 

still subjected to the influence of the induced flow of 

the rotor system, which cannot be calibrated. 

Acoustic Data Reduction and Analysis 

Before being recorded on analog tape, the microphone data were high-pass filtered at 4 Hz to remove 

very low-frequency content. The microphone and 

blade position reference signals (once-per-revolution 

and 512-per-revolution) were conditionally digitized 

using a sample rate keyed to the 512-per-revolution 

signal, which yielded 1024 data points per revolution, or a sample rate of about 18000 samples per 

second. An antialiasing filter was employed at 9 kHz. 

This conditional sampling technique is used to ensure 

that the acoustic samples are acquired when the rotor blades are in the same azimuthal position for every revolution sampled. The time histories were digitized in this manner, and then 40 revolutions were 

averaged to obtain an average acoustic waveform and 

standard deviation for each test condition. The average time history and its standard deviation were 

stored on computer disk. 

To obtain spectra, the data of four sequential rotor revolutions were transformed to the frequency 

domain using a 4096-point fast Fourier transform algorithm. This procedure was performed 40 times, 

requiring the data of 160 revolutions, and the results were averaged to obtain an average narrowband spectrum, also stored on computer disk. Considering each power spectrum as a chi-squared random variable, the statistical accuracy of the spectral 

results is 0.8 to -1.0 dB for an 80-percent confidence 

interval. 

A time domain window was not applied to these 

data. The conditional sampling approach employed 

minimizes leakage of the acoustic energy at the bladepassage frequency and its harmonics, and thus a 

time domain window is not required to improve the 

spectral estimates at those frequencies. 

Data Quality 

Tunnel Flow Quality 

The DNW has very good aerodynamic qualities 

when operated in the open throat mode (refs. 7-10). 

A potential core of uniform flow (mean flow velocity 

f0.5%) has been defined as a 5 x 3 m core within 

the 8 x 6 m nominal test section, at 7 m from the jet 

exit. Turbulence intensities in both the streamwise 

and the cross stream direction are less than 0.5 at 

this location. 

In the low-speed range, the open-jet configuration 

can exhibit a low-frequency, low-amplitude pulsation 

of the mean flow, similar to the pulsations previously 

observed in other open-jet tunnels (ref. 14). This pulsation occurred at certain velocities between 20 and 

28 m/sec, with a maximum amplitude of f0.75 m/sec 

and a period of about 10 seconds. In some cases this 

pulsation caused problems ,in achieving stable rotor 

performance parameters, so that some parts of the 

test matrix had to be redesigned or omitted. Overall, 

this pulsation was of sufficiently low frequency to be 

considered quasi-steady compared with the analysis 

time for both the rotor performance and the acoustic 

data. 

Rotor Performance Data 

The nominal thrust coefficients tested were 0.0030 

and 0.0044. The actual average thrust coefficients 

were 0.0029 and 0.00435 with standard deviations of 

3% and 1.4%, respectively. The nominal hover tip 

Mach number was 0.64, with average and standard 

deviation of 0.636 and 0.5%. The first harmonics of 

lateral and longitudinal flapping angles were generally less than 0.3O. During the pretest calibrations, 

a check was made to ensure that a change in the 

traversing array position in the test section did not 

affect the rotor performance by requiring changes to 

the rotor control inputs. 

Background Noise 

The background noise data were acquired with all 

testing and model hardware installed in the tunnel. 

The blades were not installed, but the hub was 

spinning at nominal speed (1040 rpm). 

Eflect of tunnel speed. Background noise was 

measured for a range of tunnel speeds at a fixed 

nominal traverse position 3 m upstream of the hub. 

Examples of the spectra for microphone 1 at 20, 40, 

and 60 m/sec are shown in figure 7(a). Each of 

the nine microphones located on the traversing array 

shows generally the same background noise levels as 

microphone 1 at this position (Xw = 3 m). The lowfrequency levels of the traverse microphones increase 

about 28 dB (from 82 to 110 dB) due to increasing 

the speed from 20 to 60 m/sec. Data for fuselage 

microphone 10 is shown in figure 7(b). The levels 

increase only about 15 dB (from 90 to 105 dB) with 

an increase in speed from 20 to 60 m/sec. The 

low-frequency background noise levels of the fuselage 

5 
microphones at the lower tunnel speeds are higher 

than those measured by the traversing microphones. 

The higher frequencies of the spectra are about the 

same for the fuselage and traverse microphones. The 

higher frequency noise levels for all microphones start 

at about 35 dB at 20 m/sec and increase to about 

55 dB at 60 m/sec. 

Effect of traversing array position. Background 

noise was measured at 40 m/sec for X, = 1, 3, 

and 5 m and at 60 m/sec for X, = -0.7 and 3 m. 

The spectral levels increase consistently (by about 

10 dB) as the traverse moves downwind and closer 

to the model (Xw = 5 to 1 m). This is illustrated 

by figure 8, which shows the background noise levels 

for X, = 1, 3, and 5 m at 40 m/sec from microphones 1 and 5 and by figure 9, which shows the 

data for X, = -0.7 and 3 m at 60 m/sec from microphones l and 5. This increase in level at the downstream positions is probably due to the increased 

proximity of the microphones to the bottom shear 

layer (see fig. 3(a)) and increased proximity to the 

model system. The levels measured at the fuselage 

microphones appear to be fairly insensitive to the 

traverse position, indicating that the traverse itself 

is not creating additional background noise. 

Background tones. Discrete tones were measured 

by some of the traverse microphones, particularly 

at tunnel speeds of 30, 40, and 50 m/sec, in the 

range of 2 to 5 kHz. The tone frequencies are not 

related to tunnel speed. An example can be seen 

in figure 8(b) for microphone 5 at 40 m/sec at two 

traverse positions. The highest levels of these tones 

appear to be about 70 dB, which is typically much 

lower than the rotor signals in that frequency range 

(see the following section on signal-to-noise ratio). 

The tones have the largest amplitude at upstream 

locations, although they are present at lower levels at 

Xw = 1 and -0.7 m. These tones are not measured 

by all the traverse microphones at once, indicating 

that the sources are localized. 

Effect of rotor shaft angle. Background noise 

levels were studied for X, = 3 and -0.7 m for rotor 

shaft angles of -5', Oo, 5', and 10'. No significant 

effect was seen in the traverse microphone data. 

The data for a fuselage microphone (microphone 10) 

exhibited slightly lower levels at -5', which was 

expected due to the preset microphone angle of 5' 

(see "Fuselage-Mounted Microphones"), 

Acoustic signal-to-noise ratio. From the above results, the highest background noise for the traverse 

microphones occurs when the traverse is the farthest 

downstream at the highest tunnel speed. The highest background noise for the fuselage-mounted microphones probably occurs at the largest shaft angle and 

highest tunnel speed. Several test conditions fitting 

these conditions were chosen and the rotor signals 

were compared with the background noise. The rotor 

spectra and the background spectra are almost identical at the first spectral line (20 Hz), but above this 

point the rotor spectrum is at least 5 dB higher than 

the background spectrum and typically 20 dB higher 

in the frequency range dominated by BVI noise, as 

shown in figure 10. The results indicate that the 

background noise levels are not a contaminant for 

any of the microphone measurements. 

The effect of the spurious tones observed by the 

traverse microphones at the higher tunnel speeds and 

upstream positions was also investigated. These data 

are seen in figure 10, which shows that although the 

background tones are just below the rotor signal, 

they do not appear to be a significant contaminant. 

Acoustic Reflections 

To measure the potential contamination of the rotor signal due to reflections, small explosive charges 

were specially mounted on the model rotor test rig 

and detonated to create an impulsive noise source. 

The explosives generate peak pressure levels of approximately 148 dB at a distance of 4 m. The charges 

were mounted in a three-rod mock-up rotor, with 

10 charges on each rod, mounted between 80 and 

100 percent of the rotor radius. The mock-up rotor 

was installed on the rotor head with the three rods 

located at 60°, 180', and 300' azimuth (q). Approximately 30 charges were ignited to assess the acoustic reflections for the following tunnel configurations: 

tunnel speeds of 0 and 60 m/sec, X, of 4, 2, 1, and 

-0.7 m, and 2, of -0.5, 0, 0.3, and 0.5 m. 

The acoustic signals were recorded on analog tape 

and analyzed on-site. For all the tested configurations, the reflected signals were generally less than 

10 percent of the direct blast signal, or 20 dB down 

from the direct signal, and in many cases were of 

negligible amplitude. The effect of increasing tunnel 

speed was found to be negligible, as was the effect of 

rotor hub vertical position. An example of the blast 

signals measured at 60 m/sec by microphones 1, 3, 5, 

7, and 9 at Xw = 4 m is shown in figure 11 for a blast 

located at 4 = 300'. The difference in the maximum 

pressure measured at microphones 1, 3, and 5 can be 

attributed to distance from the source. The signals 

at microphones 7 and 9 have been shielded by the 

fuselage, as discussed in the next section. 

6 
Acoustic Shielding by Fuselage 

Acoustic diffraction for the traversing array microphones was identified and is due to the presence 

of the fuselage. When the traversing array is moved 

close to the model, the microphones at one end of 

the traverse are shielded from the acoustic signal 

from the opposite side of the rotor disk. This was 

quantified by comparing the known propagation distances (from the source to each microphone) with the 

measured blast signal arrival times, amplitudes, and 

model geometry. 

From the geometry of the test hardware, a free 

acoustic path to all the microphones exists at X, = 

2 m, and all the array microphones receive a directly 

propagated blast signal. However, the amplitudes 

measured at microphones 1, 2, and 3 when the source 

is located at .JI = 60’ are attenuated to a larger degree than can be attributed to spherical spreading. 

(This is also observed at microphones 7, 8, and 9 

when the source is located at .JI = 300O.) This indicates that the fuselage is causing diffraction of the 

acoustic signal for this particular geometry. For traverse positions that are close to and under the model, 

the blast signals received by the “shielded” microphones are not the direct signal but a combination of 

a reflected and a diffracted signal, with significantly 

lower amplitudes. This can be seen in the data of 

microphones 7 and 9 in figure 11, for X, = 1 m and 

the blast located at .JI = 300’. The delay time, attenuation, and distortion of the blast waveform can 

be seen in the shielded microphones. 

These results indicate that the directivity or absolute levels of advancing- or retreating-side BVI cannot be accurately determined from the shielded microphone locations. The conclusions can be summarized in the table below. 

azimuth, deg I Xw, m 

60 I 2.0 

I Source I I Shielded I 

microphones 

1, 2, 3, 10 

1, 2, 3, 4, 10 

300 1 ii 1 7, 8, 9, 11 1 

6, 7, 8, 9, 11 

-.7 6, 7, 8, 9, 11 

Steadiness and Repeatability of Rotor Acoustic 

Data 

To illustrate the steadiness of the rotor signals, 

several cases were chosen for comparison. Two of 

these cases, at p = 0.092 and p = 0.138, are shown 

in figure 12. At p = 0.092, figure 12 shows both 

the instantaneous data (fig. 12(a)) acquired using the 

on-line ATD system and the 40-average time history 

(fig. 12(b)) obtained from the postprocessing. The 

general features of the instantaneous data are consistently seen in the averaged waveform. Particular 

features to note are the small impulses approximately 

0.01 period after the major impulses of the first three 

blade passages and the large impulse 0.01 period before the third major impulse. An equally good comparison, particularly in the BVI characteristics, can 

be found for p = 0.138 in figures 12(c) and 12(d). 

The repeatability of tKe rotor signal was studied for several test conditions repeated during the 

experiment. An example of the instantaneous data 

acquired using the on-line system at different times 

during the experiment is given in figure 13, for 

p = 0.138 and a~pp = 1.8’. These data were filtered 

by a 500-3000 Hz band-pass filter (ref. 5). While 

there is considerable random noise evident in these 

data, the general amplitudes and features of the major impulses can be seen in the two measurements. 

Testing Procedures and Measurement 

Accuracy 

Corrections to the free-stream flow angle and 

speed due to the presence of the lifting rotor in the 

test section were calculated according to the method 

of Heyson (ref. 15) and applied to the measured data 

off-line to obtain the corrected tip-path-plane angles 

presented here. These corrections are considered reasonable estimates of the true flow angle at moderate 

tunnel speeds, but become less reliable as the correction increases at the lower tunnel speeds, less than 

20 m/sec. 

Because of the pivoting of the support sting to 

achieve a specified shaft angle, the center of the rotor 

hub experiences both a vertical and a streamwise 

translation. The sting controller was programmed 

to reposition the hub at the same vertical position 

for any shaft angle. The streamwise movement of 

the rotor hub could not be corrected by repositioning 

the sting, but is calculated off-line and is accounted 

for in the calculation of microphone locations. The 

streamwise position of the microphones X, is thus 

corrected for the movement of the hub due to model 

pitch angle. 

In addition to the kinematic motion of the sting 

support, the actual rotor hub position was also influenced by vertical deflections due to its weight, the 

rotor angle, the rotor thrust level, and aerodynamic 

drag. These effects were calibrated before data acquisition by measuring the rotor hub position for the 

ranges of shaft angle, rotor thrust, and tunnel speed 

contained in the test matrix. These corrections were 

7 
then applied to the data off-line. The accuracy of the 

rotor hub position measurements is f2 mm. 

The rotor parameters-tip-path-plane angle, 

hover tip Mach number, and thrust-were set using 

the digital displays at the rotor controls. The rotor operator set the rotor speed for the desired value 

of hover tip Mach number, simultaneously adjusting 

collective pitch to obtain the desired thrust. The rotor was operated so that the first harmonics of the 

shaft bending moments and of lateral and longitudinal flapping were as close to zero as possible. In this 

case, and neglecting blade bending, the measured rotor shaft angle was thus equivalent to the tip-pathplane angle. 

Presentation of Data 

The data presented in the plots at the end of this 

paper are a representative set of the postprocessed 

results of the final test matrix defined by table 11. 

Preceding these plots is a detailed table of test conditions corresponding to the plots. Each plot presents 

the averaged time histories of traversing array microphones l, 3, 5, 7, and 9. Spectral results are 

not presented here but averaged narrow-band spectra have been calculated for all test conditions. Both 

the averaged time histories and the spectral results 

are available to the research community for their use 

in rotor noise work and can be provided in the form 

of plots, computer listings, tapes, or computer disks 

as required. Interested persons should contact the 

principal author of this document. 

NASA Langley Research Center 

Hampton, VA 23665-5225 

December 23, 1987 

References 

1. Splettstoesser, W. R.; Schultz, K. J.; Boxwell, D. A.; 

and Schmitz, F. H.: Helicopter Model Rotor-Blade Vortex Interaction Impulsive Noise: Scalability and Parametric Variations. NASA TM-86007, USAAVSCOM 

2. Boxwell, D. A.; Schmitz, F. H.; Splettstoesser, W. R.; 

Schultz, K. J.; Lewy, S.; and Caplot, M.: A Comparison of the Acoustic and Aerodynamic Measurements of 

a Model Rotor Tested in Two Anechoic Wind lhnnels. 

TM-84-A-7, 1984. 

NASA TM-88364, USAAVSCOM TM-86-A-6, 1986. 

3. Leighton, Kenneth P.; and Harris, Wesley L.: On 

Mach Number Scaling of Blade/Vortex Noise Produced by 

Model Helicopter Rotors at Moderate Tip Speeds. FDRL 

Rep. 84-3, Dep. of Aeronautics and Astronautics, Massachusetts Inst. of Technology, Oct. 1984. 

4. Martin, Ruth M.; and Splettstoesser, Wolf R.: Acoustic 

Results of the Blade-Vortex Interaction Acoustic Test 

of a 40 Percent Model Rotor in the DNW. Proceedings 

of the National Specialists’ Meeting on Aerodynamics and 

Aeroacoustics, American Helicopter SOC., c.1987. 

5. Martin, R. M.; Splettstoesser, W. R.; Elliott, J. W.; 

and Schultz, K.-J.: Advancing-Side Directivity and 

Retreating-Side Interactions of Model Rotor BladeVortex Interaction Noise. NASA TP-2784, USA6. Splettstoesser, W. R.; Schultz, K. J.; and Martin, 

Ruth M.: Rotor Blade-Vortex Interaction Impulsive 

Noise Source Identification and Correlation With Rotor 

Wake Predictions. AIAA-87-2744, Oct. 1987. 

7. Van Ditshuizen, J. C. A.; Courage, G. D.; ROSS, R.; 

and Schultz, K.-J.: Acoustic Capabilities of the GermanDutch Wind Tunnel DNW. AIAA-83-0146, Jan. 1983. 

Compilation of Calibration Data of the German-Dutch 

Wind lhnnel. MP-82.01, Duits-Nederlandse Windtunnel, Mar. 13, 1982. 

9. ROSS, R.; Van Nunen, 3. W. G.; Young, K. J.; Allen, 

R. M.; and Van Ditshuizen, J. C. A.: Aero-Acoustic Calibration of DNW Open Jet. DNW TR 82.03, GermanDutch Windtunnel (DNW) (North East Polder, Netherlands), July 1982. (Also available as Boeing Document 

D6-51501, 1983.) 

10. Seidel, M.; and Maarsingh, R. A.: Test Capabilities 

of the German-Dutch Wind Tunnel DNW for Rotors, 

Helicopters and V/STOL Aircraft. Proceedings of the 

Fifth European Rotorcraft and Powered Lift Aircraft Forum, Sept. 1979, pp. 17-1-17-22. 

11. Langer, H. J. (SCITRAN, transl.): DFVLR RotorcraftConstruction and Engineering. NASA TM-77740, 1984. 

12. Langer, H.-J. (Leo Kanner Assoc., transl.): Data Input, 

Processing and Presentation. NASA TM-77739, 1984. 

13. Breustedt, Wolfgang (SCITRAN, transl.): Data Analysis 

on the Rotor Test Stand Program for Interactive Processing. NASA TM-77948, 1985. 

14. Applin, Zachary T.: Modification to the NASA Langley 

4- by 7-Meter Tunnel for Improved Rotorcraft Aerodynamics and Acoustics. NASA paper presented at the 

American Helicopter Society National Specialist’s Meeting on Helicopter Testing Technology (Williamsburg, 

VA), Oct. 2SNov. 1, 1984. 

15. Heyson, Harry H.: Use of Superposition in Digital Computers To Obtain Wind- Tunnel Interference Factors for 

Arbitrary Configurations, With Particular Reference to 

V/STOL Models. NASA TR R-302, 1969. 

AVSCOM TR 87-B-3 1988. 

8. 

8 
i Table I . Details of the BO-105 Model Main Rotor and Rotor Test Stand 

Main rotor: 

Rotor diameter. m ......... : ................... 4 

Blade airfoil section ...................... NACA 23012 

Number of blades ............................. 4 

Blade solidity ............................. 0.077 

Blade tip speed. m/sec .......................... 218 

Flapping frequency ratio ........................ 1.111 

Lagging frequency ratio ......................... 0.785 

Nominal rotational speed. rpm ...................... 1040 

Nominal rotor thrust. N ......................... 3600 

Nominal rotor thrust coefficient .................... 0.0044 

Shaft power. kW 100 

Rotor drive moment. N-m ................... 900 at 1050 rpm 

Power (electric) .................... 3 phase. 380 V at 300 A 

Balance load range: 

Axial force. N ............................. 1000 

Side force. N 2000 

Thrust force. N .................. 7000 static 43500 dynamic 

Rolling moment. N-m ........................... 700 

Pitching moment. N-m .......................... 700 

Blade setting angle range. deg .................... -4 to 14 

Drive system: 

............................. 

.............................. 

Control system: 

9 
Table 11. Final Test Matrix 

P sweep 0.0044 

xw sweep 0.0044 

a sweep 

U SWWD 

0.0044 

0.0044 

High p 0.0044 

MU sweeD 0.0044 

MU sweeD 0.0044 

Mu SW~D 0.0044 

38 0.175 

V, rn/sec 

17 

20 

25 

30 

32 

37 

"shaft, deg 

a7 to 10 

XW, m 

5.0 

2.5 

2.5 

1.0 

1 .o 

2.5 

MH 

0.636 

Zw,m 

-2.1 

P 

0.078 

.092 

.115 

.138 

,147 

.170 

6 to 10 

4 to 8 

2 to 6 

0 to 4 

-2 to 2 

b33 to 41 -2.1 0.636 

28 to 36 

25 to 35 

20 to 26 

17 to 22 

13 to 21 

38 

35 

31 

25 

20 

0.152 to 0.187 

0.129 to 0.166 

0.115 to 0.161 

0.092 to 0.120 

0.078 to 0.101 

0.060 to 0.097 

0.175 

,161 

.143 

,115 

.092 

.087 

0.138 

0.129 to 0.156 

0 

2.0 

4.0 

6.0 

8.0 

10.0 

0.5 

3.0 

5.0 

5.5 

8.0 

10.0 

a4 to 7 

5.0 

2.5 

1.0 

1 .o 

2.5 

2.5 

5.0 

C - 1.2 to 6.0 -2.1 0.636 

19 

30 

b28 to 34 

~ 0.636 

0.636 

-2.1 

-2.1 

-1.15 

-1.15 

0.0044 30 0.138 5.0 d - 1.2 to 4.4 -2.1 0.636 

30 0.138 

,184 

-1.15 -2.1 0.636 

40 

I O.OO3O 

b28 to 35 

36 to 44 

30 

40 

-1.15 -2.1 0.636 

0.636 

0.129 to 0.161 

0.166 to 0.202 

0.138 

.184 

2.8 

1.2 

2.8 

1.2 

d - 1.2 to 4.4 -2.1 

0.087 

.143 

,087 

.143 

10.0 

5.0 

10.0 

5.0 

e 1.5 to 6.0 

e1.5 to 6.0 

f 2.5 to 6.2 

f 2.5 to 6.2 

-1.6 

-1.6 

-2.6 

-2.6 

19 0.636 

31 

19 

31 

50 

data 

0.230 -2.1 

-2.1 

-2.6 

-2.6 

-2.1 

0.636 

h0.56 to 0.72 

-5 to 1 

-7 to 1 

-7 to 0 

-7 to 0 

5.0 

2.5 

2.5 

2.57 

3.68 

4.5 

60 .276 

60 1 .276 

927: 35 1- 

31 1 0.140 

g17.5 to 22.5 

-2.1 0.636 

-2.1 

5.0 

8.0 

4.5 

4.0 ~~ 

h0.56 to 0.72 

20 ~ 0.090 

933.3 to 42.8 1 0.175 

8.0 4.0 -2.1 0.636 

0.5 4.0 -2.1 h0.56 to 0.72 

CT sweep I 0.0036 to 0.0056 0.5 4.0 -2.1 0.636 

"Varied in increments of 0.5. 

bVaried in increments of 1. 

CVaried as follows: 6.0, 5.2, 4.4, 3.6, 2.8, 2.0, 1.5, 0.8, 0, -0.7, -1.2. 

%aried as follows: 4.4, 2.8, 2.0, 1.2, 0.4, 0, -0.4, -0.7, -1.2. 

eVaried as follows: 1.5, 2.1, 2.7, 3.3, 4.0, 4.6. 

fVaried as follows: 2.5, 3.5, 4.4, 5.4, 6.2. 

Waried in increments of 2. 

hVaried in increments of 0.04. 

10 
+ 

FLOW MIC 

I 

TOP VIEW SIDE VIEW 

Figure 1. Diagram of coordinate system and test geometry. 

11 
t- 47 

48 *- 

E 

14 

0 Prefabricated concrete circuit 

@ Eight-bladed fan with electric drive (nominal 12.5MW) 

@ Heat exchanger (1 1MW) with flow rectifier 

@ Anti-turbulence screens 

@ Exchangeable test sections: 9.5x9.5m2, 8x6m2, 6x6m2, open jet 

@ Air exchange hatches 

@ Throttling device 

@ Acoustically treated testing hall: 52x30x20m3 

@ Parking hall 

@ Control room 

0 Experiment hall 

@ Model assembly hall 

@ Calibration hall 

@ Office building 

Figure 2. Plan view of the DNW facility. 

12 
(a) Side view. 

1 

I 

I 

I 

I 

\I ,-Mic 11 

I 

7195 4 //Mic 10 I 

'I 

! 

Traversing Array 

_____ - ___ -_-- J 

(b) Plan view. 

Figure 3. Diagram of experimental apparatus installed in the DNW open test section. All dimensions are in 

millimeters. 

13 
Figure 4. Diagram of the 

(a) Side view. 

I -1 

650 1; 

(b) Front view. 

DFVLR rotor test stand. All dimensions are in millimeters. 

14 
w 350 4 

TRAILING EDGE 

Figure 5. Model BO-105 rotor blade geometry. 

7 

Acoustic lining 

\Micro p h o ne \ 

\Wing steel structure 150 x 75 

Figure 6. Cross section of traversing microphone array wing and microphone holder. All dimensions are in 

millimeters. 

15 
110 

loo[ 90 

v, 

80 

60 

50 

40 

Frequency, Hz 

(a) Microphone 1. 

100 i; 

90 

80 - 

J & 70 

60 

50 

40 

v, m/sec 

Frequency, Hz 

(b) Microphone 10. 

Figure 7. Background noise at X,,, = 3 m, as a function of tunnel speed. Af = 20 Hz. 

I 16 
110- 

look 90 ,<,, 

0 1000 2000 3000 4000 5000 

Frequency, Hz 

(a) Microphone 1. 

110r 

90 

80 

m 

U 

a 

v) 

70 

60 

50 

0 1000 2000 3000 4000 5000 

Frequency, Hz 

(b) Microphone 5. 

Figure 8. Background noise at V = 40 m/sec as a function of traversing array position. A f = 20 Hz. 

17 
120r 

70 

60 I 

-- 

0 1000 2000 3000 4000 5000 

120r 

Frequency, Hz 

(a) Microphone 1. 

j 90 

n 

60 

80 

70 

0 1000 2000 3000 4000 5000 

Frequency, Hz 

(b) Microphone 5. 

Figure 9. Background noise at V = 60 m/sec as a function of traversing array position. Af = 20 Hz. 

18 
100 

90 m 

TI. 

Q, 

> 

- 

2 80 

s 

rn 

rn 

Q, 

TI 

C 

3 

0 

cn 

h 70 

60 

50 

- Rotor spectrum 

clz.185 

CT= .0044 

aTpp= -0.9" 

X,= 2.5m 

Microphone 5 

----Background noise 

V=40 m/sec 

0 1000 2000 3000 4000 5000 

Frequency, Hz 

Figure 10. Comparison of typical rotor spectrum with background noise. Af = 20 Hz. 

Xw= 1.0m 

V = 60 mlsec 

Source at W= 300' 

s 

m 

f 

n 

m 

.- z 

X 

Q 

f 

f 

a 

m 

?! 

n 

m 

-1 e 

1 

0 7 

-1C 

0 lE 

-1 

0 lh 

0 l0 12 3 4 5 6 7 8 9 10 

-1 5 

Time. msec 

Max 

Pressure, 

Pa 

202 

194 

1302 

1657 

1580 

Figure 11. Example of blast calibration data, showing shielded microphones 7 and 9. 

19 
20 

15 

2 

$10 

3 

VI 

VI 

L5 

n 

0 

-5 

20 

15 

a 

n 

10 $ 

3 

v) 

;5 

n 

0 

r 

u= .092 

aTpp= 3.8" 

CT= .0046 

- X, = 6.0m 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 

Time/Period 

(a) Unaveraged on-line data at p = 0.092. 

I I I I 1 1 I I I I I -3 0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 

TimeIPeriod 

(b) Averaged postprocessed data at p = 0.092. 

Figure 12. Comparison of instantaneous on-line acoustic data with averaged postprocessed results 

(microphone 5). 

20 
pz.138 

QTPPZ2.2" 

25 X, =1 .Om 

20 

[ c~=.0044 

-10 LI - I I I 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1 .O 

Time/Period 

(c) Unaveraged on-line data at p = 0.138. 

6101 

2 $5 

5 

a 

0 

-5 

V 

-10 ' 

0 -1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 

Time/Period 

(d) Averaged postprocessed data at p = 0.138. 

Figure 12. Concluded. 

21 
I 

Time, msec 

(a) Data point 411. 

15 

10 

n a5 

2 

2 

30 

v) 

v) 

-5 

-10 

-15 

Y I 

0 10 20 30 40 50 60 

Time, msec 

(b) Data point 429. 

Figure 13. Repeatability of filtered on-line data. p = 0.138; a~pp = 1.8'; X, = 1.0 m; CT = 0.0030; 

i 
Averaged Acoustic Signals 

results of the final test matrix defined by table XI. Each plot presents the averaged time 

The data presented in the following plots are a representative set of tL,e postprocessef i 

I 

I 

I 

I 

histories of traversing array microphones 1, 3, 5, 7, and 9. The following table is an index 

of test conditions for the plots and page numbers. I 

-2.1 

” 

v, 

m/sec 

x17.0 

x17.0 

x17.0 

z20.0 

z20.0 

x20.0 

x25.0 

x25.0 

z25.0 

x30.0 

x30.0 

x30.0 

x32.0 

z32.0 

z32.0 

x32.0 

37.0 

37.0 

37.0 

33.0 

50.0 

39.0 

41.0 

27.5 

30.3 

32.0 

34.0 

35.9 

24.8 

26.5 

” 

Test Conditions for Plots of Averaged Acoustic Signals 

z0.636 

-.I 

P 

x0.078 

w.078 

x.078 

z.092 

z.092 

x.092 

x.115 

x.115 

w.115 

x.138 

x.138 

x.138 

w.147 

z.147 

x.147 

x.147 

.170 

.170 

.170 

.152 

.160 

.179 

.187 

.130 

.139 

.147 

.157 

.165 

.113 

.125 

aTPP 7 

deg 

b1.7 

b3.0 

b4.6 

b4.1 

b6. 1 

1.5 

3.5 

5.5 

.2 

2.2 

4.3 

-1.6 

-.l 

.9 

2.4 

-3.1 

-1.1 

.9 

-1.4 

-1.2 

-1.0 

-.8 

-.l 

.1 

.4 

.5 

.8 

1.6 

1.7 

b2.1 

“XW , m 

w5.3 

x5.3 

w5.3 

w2.8 

1 

1 

xl.1 

x2.5 

c 

xl.1 

1 

w1.2 

x1.2 

ZW , 

m Point 

2 10 

264 

267 

225 

221 

217 

226 

230 

234 

244 

239 

235 

253 

250 

248 

245 

262 

258 

254 

318 

320 

324 

326 

309 

311 

313 

315 

317 

297 

299 

- 

=ag 

27 

28 

29 

30 

31 

32 

33 

34 

35 

36 

37 

38 

39 

40 

41 

42 

43 

44 

45 

46 

47 

48 

49 

50 

51 

52 

53 

54 

55 

56 

- 

- 

“XW varies with a~pp (see section “Testing Procedures and Measurement Accuracy”). 

bAccuracy of estimated a~pp deteriorates as V decreases below 20 m/sec (see “Testing 

Procedures and Measurement Accuracy”). 

23 
v, 

m/sec 

28.5 

31.2 

32.8 

35.0 

20.1 

22.0 

23.8 

25.6 

17.3 

19.1 

21.0 

13.1 

16.2 

18.3 

20.7 

38.0 

i 

I 

I 

I 

1 

35.0 

31.1 

24.8 

20.0 

18.9 

x.00435 

” 

,00476 

~~ 

Test Conditions for Plots of Averaged Acoustic Signals-Continued 

P 

0.132 

.142 

.150 

.159 

.092 

.lo1 

.110 

.120 

.079 

.087 

.095 

.059 

.073 

.083 

.093 

x.175 

I 

I 

I 

I 

1 

x.161 

x.143 

x.114 

x.092 

z.087 

ffTPP 7 

2.2 

2.3 

2.4 

2.7 

b2.1 

2.7 

3.2 

3.6 

b2.7 

b3.7 

4.4 

b1.4 

b4.0 

b5.2 

6.2 

x -.5 

deg 

I 

I 

I 

1 

1.7 

e3.3 

x2.9 

bx3.8 

bx3.8 

bx3.8 

x4.1 

x4.1 

bx5.7 

“XW 3 

m 

1.2 

I 

I 

i 

x2.8 

x2.9 

x2.9 

x2.9 

x5.5 

-.l 

.a 

2.0 

3.6 

5.2 

-1.0 

.5 

2.1 

3.7 

5.3 

-.9 

.6 

2.2 

3.8 

5.4 

-.9 

1 .o 

3.1 

4.7 

6.3 

-.8 

.8 

2.4 

4.0 

5.6 

- .6 

ZW 7 

m 

iPoint 

30 1 

303 

305 

308 

290 

292 

294 

296 

284 

286 

288 

275 

278 

280 

283 

400 

402 

404 

406 

408 

399 

397 

395 

393 

39 1 

380 

382 

384 

386 

388 

379 

377 

374 

372 

369 

358 

360 

362 

364 

366 

357 

- 

'age 

57 

58 

59 

60 

61 

62 

63 

64 

65 

66 

67 

68 

69 

70 

71 

72 

73 

74 

75 

76 

77 

78 

79 

80 

81 

82 

83 

84 

85 

86 

87 

88 

89 

90 

91 

92 

93 

94 

95 

96 

97 

- 

- 

“XW varies with o~pp (see section “Testing Procedures and Measurement Accuracy”). 

bAccuracy of estimated ~~pp deteriorates as V decreases below 20 m/sec (see “Testing 

Procedures and Measurement Accuracy”). 

24 
-2.1 

v 

Test Conditions for Plots of Averaged Acoustic Signals-Continued 

” 

v 

QTPP 9 

deg 

x5.7 

I 

2.3 

3.8 

5.3 

2.9 

3.2 

3.4 

3.6 

3.2 

1 

-.2 

1.3 

3.4 

5.9 

-1.6 

.4 

2.4 

1.4 

1.6 

1.9 

x.5 

1.6 

I 

1 

v 

“XW 7 

m 

” 

ZW , 

m CT 

x0.00435 

.00474 

x.00435 

290 

Point 

354 

352 

350 

348 

476 

473 

470 

512 

510 

508 

506 

536 

538 

540 

542 

544 

460 

457 

453 

448 

469 

465 

46 1 

489 

492 

496 

497 

499 

501 

503 

505 

515 

517 

519 

521 

523 

532 

530 

528 

527 

MH 

XO 36 

- 

Page 

98 

99 

100 

101 

102 

103 

104 

105 

106 

107 

108 

109 

110 

111 

112 

113 

114 

115 

116 

117 

118 

119 

120 

121 

122 

123 

124 

125 

126 

127 

128 

129 

130 

131 

132 

133 

134 

135 

136 

137 

- 

- 

“Xw varies with ~~pp (see section “Testing Procedures and Measurement Accuracy”). 

bAccuracy of estimated a~pp deteriorates as V decreases below 20 m/sec (see “Testing 

Procedures and Measurement Accuracy”). 

25 
v: 

m/sec 

40.0 

18.5 

18.5 

30.7 

30.7 

18.5 

18.5 

30.9 

30.9 

49.8 

49.8 

49.8 

49.8 

60.0 

60.0 

60.0 

60.0 

59.6 

59.6 

59.6 

59.6 

59.1 

59.1 

59.1 

59.1 

30.9 

x19.6 

I 

I 

I 

38.0 

x0.636 

Test Conditions for Plots of Averaged Acoustic Signals-Concluded 

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bAccuracy of estimated a~pp deteriorates as V decreases below 20 m/sec (see “Testing 

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178 
National AeiOnauliCS and 

Soace AdminisIralion 

9. Security Classif.(of this report) 

Unclassified 

Report Documentation Page 

20. Security Classif.(of this page) 21. No. of Pages 22. Price 

Unclassified 181 A05 

1. Report No. 

NASA TM-4024 

AVSCOM TR 87-B-4 

2. Government Accession No. 

4. Title and Subtitle 

Acoustic Measurements From a Rotor Blade-Vortex Interaction 

Noise Experiment in the German-Dutch Wind Tunnel (DNW) 

7. Author(s) 

R. M. Martin, W. R. Splettstoesser, J. W. Elliott, 

and K.-J. Schultz 

Aerostructures Directorate 

Langley Research Center 

Hampton, VA 23665-5225 

National Aeronautics and Space Administration 

Washington, DC 20546-0001 

U.S. Army Aviation Systems Command 

St. Louis, MO 63120-1798 

3. Performing Organization Name and Address 

USAARTA-AVSCOM 

12. Sponsoring Agency Name and Address 

and 

3. Recipient’s Catalog No. 

5. Report Date 

March 1988 

6. Performing Organization Code 

8. Performing Organization Report No. 

L-16368 

10. Work Unit No. 

505-61-51-06 

11. Contract or Grant No. 

13. Type pf Report and Period Covered 

Technical Memorandum 

14. Army Project No. 

lL161102AH45 

15. Supplementary Notes 

R. M. Martin: Langley Research Center, Hampton, Virginia. 

W. R. Splettstoesser and K.-J. Schultz: DFVLR, Braunschweig, West Germany. 

J. W. Elliott: Aerostructures Directorate, USAARTA-AVSCOM, Langley Research Center, 

Hampton, Virginia. 

Acoustic data are presented from a 40-percent-scale model of the four-bladed BO-105 helicopter 

main rotor, tested in a large aeroacoustic wind tunnel. Rotor blade-vortex interaction (BVI) noise 

data in the low-speed flight range were acquired using a traversing in-flow microphone array. The 

experimental apparatus, testing procedures, calibration results, and experimental objectives are fully 

described. A large representative set of averaged acoustic signals are presented. 

16. Abstract 

7. Key Words (Suggested by Authors(s)) 

Rotor acoustics 

Blade-vortex interaction noise 

Acoustic directivity 

18. Distribution Statement 

Unclassified-Unlimited 

NASA FORM 1626 OCT 86 NASA-Langley, 1988 

For sale by the National Technical Information Service, Springfield, Virginia 22161-2171