. NASA Technical Memorandum 

AVSCOM Technical Memorandum 89-6-005 

I 0 1 6 24 

REDUCTION OF BLADE-VORTEX INTERACTION NOISE 

USING HIGHER HARMONIC PITCH CONTROL 

THOMAS F. BROOKS 

EARL R. BOOTH, JR. 

J. RALPH JOLLY, JR. 

WILLIAM T. YEAGER, JR. 

MATTHEW L. WILBUR 

[q&SA-TH- ta8626) REDUCZIOY OF BLADE- VORTEX 889-27465 IFTEBACZIOM l!JOISt USIlG Iff6HBR EAUOBIC 

PZTCE COIT%OL (BAS&.< Langley Research 

center) 20 -p CSCL 20a Dlrclaii 

6301 0223909 

JULY 1989 

National Aeronautics and 

Space Administration 

Langley Research Center 

Hampton, Virginia 23665-5225 

US ARMY 

AVIATION 

SYSTEMS COMMAND 

AvmnoN R~T ACTIVITY 
REDUCTION OF BLADE-VORTEX INTERACTION NOISE USING HIGHER 

HARMONIC PITCH CONTROL 

Thomas F. Brooks and Earl R. Booth, Jr. 

NASA Langley Research Center 

Hampton, Virginia 

J. Ralph Jolly, Jr. 

Planning Research Corporation 

Hampton, Virginia 

William T. Yeager, Jr. and Matthew L. Wilbur 

Aerostructures Directorate 

US Army ARTA, AVSCOM 

Hampton, Virginia 

ABSTRACT 

An acoustics test using an aeroelastically scaled rotor was conducted to 

examine the effectiveness of higher harmonic blade pitch control for the reduction of 

impulsive blade-vortex interaction (BVI) noise. A four-bladed, 1 10 in. diameter, 

articulated rotor model was tested in a heavy gas (Freon-12) medium in Langley's 

Transonic Dynamics Tunnel. Noise and vibration measurements were made for a 

range of matched flight conditions, where prescribed (open-loop) higher harmonic 

pitch was superimposed on the normal (baseline) collective and cyclic trim pitch. For 

the inflow-microphone noise measurements, advantage was taken of the 

reverberance in the hard walled tunnel by using a sound power determination 

approach. In the paper, initial findings from on-line data processing for three of the 

test microphones are reported for a 4hev (4P) collective pitch control for a range of 

input amplitudes and phases. By comparing these results to corresponding baseline 

(no control) conditions, significant noise reductions (4-5 dB) were found for low-speed 

descent conditions, where helicopter BVI noise is most intense. For other rotor flight 

conditions, the overall noise was found to increase. All cases show increased 

vibration levels. 

1 
SYMBOLS 

a0 

CT 

MT 

nP 

R 

a 

a' 

r 

8 

e 

eC 

P 

P 

Y 

YC 

n 

speed of sound in test medium, ft/sec 

rotor thrust coefficient, t h rust/pnR2( QR)* 

hover tip Mach number, QWaO 

n'th harmonic of rotor rotational period 

rotor radius, ft 

rotor tip path plane angle referenced to tunnel streamwise axis, deg 

effective a corrected for closed-wall wind tunnel effect, deg 

tip vortex strength, ft2/sec 

calculated "full-scale helicopter" flight descent angle, positive in descent, deg 

pitch angle of blade at Y, deg 

amplitude of higher harmonic pitch at Yc, deg 

advance ratio, tunnel flow velocity/QR 

density of test medium, slug/ft3 

blade azimuth angle, deg 

blade azimuth angle selected for Bc, see Fig. 3, deg 

rotor rotation frequency, rad/sec 

I NTRODU CTlON 

Impulsive blade-vortex interaction (BVI) noise, due to blade interaction with 

shed vortices of preceding blades, has been a major topic of rotorcraft acoustics 

research for a number of years. One noise reduction concept' is that decreases in 

blade lift and/or vortex strength at the blade-vortex encounters should reduce the 

intensity of the interactions and thus noise. An application of this idea is illustrated in 

Fig. 1 showing the rotor blades undergoing higher harmonic pitch angle variations. It 

is quite apparent that pitch control would not only modify pitch but would also modify 

the strengths of the shed vortices as well as possibly the interaction locations. The 

amplitude and phasing of such pitch controls may be expected to be important to the 

noise problem, since the strongest BVI occurrences tend to be located within a limited 

rotor azimuth angle range of the first rotor quadrant2 (roughly between Y = 45' and 

75"). It appears therefore that although, historically, higher harmonic control of pitch 

2 
has been studied as a means to reduce helicopter vibration levels (for example see 

reference 3), its use to reduce noise offers potential. 

This paper reports the initial finding of a rotor test designed to evaluate the 

noise reduction benefit of higher harmonic pitch control. The test approach involves 

the measurement of noise and vibration with and without prescribed higher harmonic 

pitch inputs superimposed on the normal cyclic trim pitch. Uniquely, the acoustic 

testing was conducted in a heavy gas (Freon-12) flow medium, rather than air, and 

the reverberant field of the hard wall tunnel test section was used to advantage by 

making acoustic measurements using a sound power determination approach. 

EXPERIMENT 

The test was conducted using the Aeroelastic Rotor Experimental System 

(ARES) in the Langley Transonic Dynamics Tunnel (TDT). The test setup, with in-flow 

microphones mounted upstream and downstream of the rotor model in the test 

section, is shown in Fig. 2. The tunnel test section is 16 ft. square with cropped 

corners. In the TDT, either air or Freon-1 2 gas can be used as the test medium. The 

advantages of using Freon-12 for aeroelastic testing of scale model rotors are 

discussed in reference 4. For this test , Freon-12 at a nominal density p of .0046 

slug/ft3 (nominal tunnel pressure of .45 atmosphere) was used. The 11 0 in. diameter 

dynamically-scaled rotor has untwisted NACA 0012 section blades with a 4.24 in. 

chord. The articulated flap and lead-lag hinges are offset 3 in. from the center of 

rotation. 

Because this was the first aeroacoustic test to be conducted in a heavy gas 

medium, detail flow-noise calibrations were performed in the TDT for both air and 

Freon-1 2. The results reinforced the conclusions of a scaling law analysis, using 

fundamental aeroacoustic equations, that acoustic pressures are readily scaled 

between test media. To address microphone sensitivity questions for the Freon-1 2 

medium, a special calibration was performed prior to the tunnel test. It was found that 

3 
for a medium pressure of 950 psf (corresponding to tunnel test conditions), the 

microphones had almost the same diaphragm sensitivity (within 0.2 dB) as air at one 

atmospheric pressure. As for microphone body diffraction effects on sensitivity, the 

acoustic wavelengths for identical frequencies in Freon-1 2 are smaller by the ratio of 

the speed of sound for Freon-1 2 and air, (aO)freon/(aO)air = 500/1130 = .43. However, 

for matched Mach number conditions, the test speeds (rotor and tunnel) are reduced 

by the same ratio thereby reducing frequency and rendering the same wavelengths 

for the same rotor harmonics for both air and Freon-1 2. Therefore the microphone 

response at specific harmonics of the blade passage frequency is the same in 

Freon-12 as if the test on this model had been conducted in air. With the noise data 

interpretation being straight forward, a net test advantage is found for using a heavy 

gas medium for this aeroacoustic rotor test because the rotor is dynamically scaled 

and the Reynolds numbers are higher (by 17 percent for this test) compared to air. 

Twelve one-quarter inch diameter B&K pressure type microphones, six 

upstream and six downstream of the rotor model, were used to make the noise 

measurements. Figure 2 shows the microphones fitted with nose cones and mounted 

in vibration isolated streamlined microphone stands. Because of the reverberant 

character of the TDT test section, it was decided not to attempt directivity 

measurements but to employ the microphone distribution shown and special noise 

field calibrations to determine sound power spectra. Figure 2 shows that the 

microphones are placed away from the nearfield of the rotor BVI noise source region. 

The present report includes data from the indicated microphones at 16 and 13 feet 

upstream and a microphone at 13 feet downstream from the rotor model center. The 

normally open slots in the tunnel wall were covered to further enhance test section 

reverberance, thereby reducing statistical variance of noise measured between 

microphones. 

, 

4 
Blade pitch motion is input to the rotor by moving the swashplate with three 

hydraulic actuators. For this four-bladed rotor, the higher harmonic pitch is achieved 

by superimposing 4hev (4P) swashplate motion upon basic fixed swashplate 

collective and cyclic (1 P) flight control inputs. Fourhev collective pitch motion (all 4 

blades pitching the same way simultaneously), as well as pitch schedules containing 

3P, 4P, and 5P pitch harmonic components, are possible by phasing the 4P inputs315. 

For this test, a specially developed computer-based open-loop control system was 

used to superimpose the higher harmonic pitch signals on the ARES control system. 

The pitch motion achieved, as well as the test procedure, can be described with the 

aid of Fig. 3 which shows blade pitch angle data versus blade azimuth angle for a 

specific flight condition. For a given advance ratio and tip path plane angle, the mean 

collective (6.5', for the case shown) required to achieve the prescribed CT and the 

basic 1 P (3.8") pitch control for zero flapping trim, with respect to the rotor shaft, were 

attained. Once aeroelastic and acoustic data were taken for this baseline case, 

prescribed higher harmonic pitch was superimposed to obtain a deflection of €Ic at 

azimuth angle Yc and data again taken. For some rotor conditions, small adjustments 

were necessary in the mean collective and cyclic to maintain identical CT and trim 

flight conditions, although none were needed for the case of Fig. 3. The 4P higher 

harmonic pitch portion (obtained by subtraction of the total from the baseline case) is 

seen at the bottom of the figure. The net pitch is seen not to be purely a 4P collective, 

but contains other harmonics due to normally occurring pitch-flap and pitch-lag 

couplings. For the 4P collective noise data shown in this report, the higher harmonic 

collective pitch amplitude eC at azimuthal angle Yc in the first quadrant (0 5 Yc < 90') 

is defined in the manner shown in Fig. 3. 

The rotor was tested over a broad range of operating conditions where the rotor 

thrust coefficient CT was maintained at 0.005. Rotor advance ratios p less than 0.1 1 

5 
were not possible due to wind tunnel minimum operating speed limitations. The rotor 

rotational speed was i2 = 650 rpm (the hover tip Mach number was nominally 

MT = 0.62). Specific test flight conditions were defined based on the tunnel 

referenced tip path plane angle a and the advance ratio p at the specified CT. For the 

data presented, the tip path angles were corrected6 to account for the closed wall 

wind tunnel effects to obtain equivalent freestream a' values. Also, in order to 

interpret the noise results in terms of full scale flight conditions, equivalent flyover 

descent angles €3 were calculated7 based on fuselage-rotor drag of a MBB 60-105. 

A portion of the test was concerned with a specific pitch control containing 3P, 4P, and 

5P components, but this paper only deals with results for the 4P collective pitch 

control. 

RESULTS AND DISCUSSION 

The noise data presented were obtained on-line during the test from the three 

microphones mentioned earlier. The microphone signals were analog band-passfiltered between 200 to 1600 Hz (3 dB down at 4.5 and 37 blade passage harmonics) 

to emphasize the impulsive BVI dominated portion of the noise. The sound pressure 

levels for each microphone were averaged to obtain a single dB value for each test 

point. Although little meaning is attached to the absolute dB values for present 

purposes, the relative levels and trends should follow that of more detailed analyses. 

A detailed sound power analysis, using all 12 microphones for all the data, has not 

been completed. 

In the results to follow, the levels include contributions from not only impulsive 

BVI noise, of particular concern here, but also from other harmonic and broadband 

noise sources.7 Harmonic noise from unsteady, but non-impulsive, loading can be 

expected to be significant for operating conditions where BVI noise is diminished. 

This would be especially true with the increased blade motion unsteadiness 

associated with the pitch controls tested. Also, broadband noise from blade-turbulent 

6 
. 

wake interactions (BWI) will contribute. For climb conditions, broadband self noise 

from blade boundary layer sources will be important. 

Noise level results are presented in Fig. 4 for ten different flight conditions, 

where the rotor operated at baseline (without higher harmonic pitch) and also where 

4P pitch was used at different amplitudes and phases. Fig. 4(a)-(c) are for steep 

descent angles where the rotor wakes are primarily above the plane of the rotor. Part 

(a) is for a advance ratio of p = 0.17 and helicopter descent angle of 8 = 9.1 O 

(respective tip path plane angles a and a' are shown in parentheses). The noise 

levels are plotted versus the input azimuthal angle Yc corresponding to the amplitude 

€Ic. The amplitudes tested of eC = -0.5" and -1 .Oo are indicated by the symbols. The 

baseline case (ec = 0") is shown positioned at the Yc = 0" plot location, for 

convenience. The noise results represented by solid symbols are repeat test points to 

be subsequently discussed. It is seen, for the Fig. 4(a) flight case, that the noise level 

increases above the baseline condition for all 4P cases shown, especially near 

Yc = 0" and 60". The larger control pitch of eC = -1 .Oo produced the larger noise 

increases. Similar trends are seen for Fig. 4(b) and (c), where the descent angles are 

also steep. The noise character at these angles was not substantially impulsive, with 

or without pitch controls, indicating a less than dominant role for BVI noise in the 

trends observed. 

Fig. 4(d)-(g) are for descent angles and speeds where the rotor generally 

operates in or about its own wake. BVI noise would be expected to be most intense 

for these cases and, indeed, the baseline cases have higher levels than those at 

steeper angles, especially at lower p. Subjectively, the impulsive character of the 

noise was quite noticeable. The use of 4P collective control is seen to reduce the 

noise for a range of azimuth control angles for these lower p values. The greater pitch 

amplitude of eC = -1 .O" is seen to be most effective at the lowest p values of 0.1 4 and 

0.20, while the smaller eC = -0.5" is more effective at the somewhat higher p = 0.266 

7 
value. The net reductions were due to substantial impulsive BVI noise reductions 

along with some increase, particularly in the low frequencies, in the noise of other 

noise components. Note that the azimuthal range where reductions occur, roughly 

between Y = 45' and 75', correspond to the expected BVI locations in reference 7. 

This is consistent with the concept that reductions in blade loading and vortex 

strength in the vicinity of BVI occurrences should reduce BVI noise. 

Fig. 4 (h)-(j) are for mild descent angles where the wake generally lies below 

the rotor. At the lower p value of Fig. 4(h), the 4P pitch is seen to reduce noise using 

both eC = -0.5' and -1 .Oo. For the higher advance ratio of Fig. 4(i), €Ic = -0.5' is more 

effective. No net benefit is seen for pitch control in Fig. 4(j) for p = 0.30. 

A portion of the test was directed at more clearly defining flight regimes where 

higher harmonic pitch control can be used to reduce BVI noise. Fig. 5 shows, for the 

baseline (no control) case, a contour map of noise levels for a broad range of "full 

scale helicopter" descent angles 8 and advance ratios p. A contouring program was 

used with measured levels at the test grid points indicated. Some test grid points are 

seen to be overlaid by letters which correspond to the parts of Fig. 4. For reference, 

the noise levels determined during this part of the test are shown in Fig. 4 by the solid 

symbols. These are seen to be matched within one dB to the corresponding open 

symbols which demonstrates the degree of repeatability. The BVI noise is seen to be 

most intense at lower speed and descent angles corresponding to normal landing 

approach for helicopters. The tunnel limitation, which prevented acquisition of data at 

advance ratios below p = 0.1 1, is unfortunate because of the importance of BVI noise 

at low p. The intense BVI impulsive noise lies in a region which is approximately 

centered about 8 = 9' at p = 0.1 1 and ranging to 8 = 6' at almost p = 0.3. The flight 

conditions of Figs 4(d)-(f) are positioned in this region, whereas the other points are 

seen to border it. The level and climb flight regimes are dominated by non-impulsive 

loading and broadband noise. 

8 
The flight matrix of Fig. 5 was also conducted for a 4P pitch control of eC = -1 " 

and Yc = 60". While this pitch is seen in Fig. 4 to not always be optimum, it appears to 

give representative noise reductions for flight conditions where reductions were 

found. Fig. 6 shows the contour plot for the resultant levels. The effect on the noise is 

dramatic since the particularly intense BVI noise region is eliminated. Fig. 7 shows 

the relative change between the levels of Fig. 6 and that of Fig. 5. Noise reduction 

(negative level change) is seen limited to the landing approach flight regime where 

BVI noise is most important. The maximum net reduction found was 4.7 dB, at 

€3 = 8.5" and p = .11. Noise tends to increase where BVI noise is not dominant for 

baseline conditions; that is, for climb, level flight, steep descent, and high speed flight 

for all angles. As mentioned, the particular 4P pitch control amplitude and phase 

used is not always optimum. Based on the discussion of Fig. 4, the noise reduction 

region could be expanded for 4P control over that shown by employing less 

amplitude in the other fringes of the region. 

The practicality of implementing specific higher harmonic pitch control for noise 

reduction will depend in part on accompanying vibratory loads. Early analysis 

reveals that the vibratory forces and moments, as measured by the six-component 

balance mounted below the model base, consistently experienced increases 

(primarily in the 4P components) with the application of the pitch control over that 

measured at baseline conditions. Of particular concern is that the baseline and pitch 

control loads were phased such that the resultant 4P loads tended to be maximized at 

minimum noise (azimuth control angle Yc = 60" to 75") and minimized at maximum 

noise (Yc = 15" to 30"). This trend was noted throughout the test matrix. 

CONCLUDING REMARKS 

The reduction of impulsive BVI noise by higher harmonic pitch control appears 

to be a viable concept. For landing approach conditions, where noise reductions 

occur, the impulsiveness of the noise was found to be diminished and replaced in 

9 
part, by noise of a more low frequency character. The impulsive noise reductions 

appear to correspond to reductions in blade pitch and vortex strength in the vicinity of 

BVI occurrences. All 4P pitch control cases tested showed increased vibration levels. 

The vibration concern, as well as a more detailed quantification of the noise results, 

must be addressed in subsequent analyses. Important questions of noise directivity 

effects are not readily addressable with the present data base. In the present study, 

aeroacoustic testing in heavy gas was demonstrated for the first time. 

REFERENCES 

1. J. C. Hardin, and S. L. Lamkin, "Concepts for Reduction of BladeNortex Interaction 

Noise," J. Aircraft, Vol. 24, No. 2, pp. 120-125, Feb 1987. 

2. W. R. Splettstoesser, K. J. Schultz, and Ruth M. Martin, "Rotor Blade-Vortex 

Interaction Noise Source Identification and Correlation with Rotor Wake Predictions," 

AlAA 1 1 th Aeroacoustics Conference, AIAA-87-2744, Oct. 1987. 

3. E. R. Wood, R. W. Powers, and C. E. Hammond, "On Methods for Application of 

Harmonic Control," Vertica, Vol. 4, pp. 43-60, 1980. 

4. W. R. Mantay, W. T. Yeager, Jr., M. N. Hamouda, R. G. Cramer, Jr., and C. W. 

Langston, "Aeroelastic Model Helicopter Rotor Testing in the Langley TDT," NASA 

TM-86440 (also USAAVSCOM TM-85-B-5), June 1 985. 

5. F. J. McHugh and J. Shaw, Jr., "Benefits of Higher-Harmonic Blade Pitch: Vibration 

Reduction, Blade-Load Reduction and Performance Improvement," Proceedings of 

the American Helicopter Society Mideast Region Symposium on Rotor Technology, 

August 1976. 

6. H. H. Heyson, "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. 

7. T. F. Brooks, J. R. Jolly, and M. A. Marcolini, "Helicopter Main Rotor 

Noise-Determination of Source Contributions Using Scaled Model Data," NASA TP2825, August 1988. 

10 
Reduced vortex / 

pitch 

Fig. 1. Illustration of noise reduction concept 

. 

11 
Fig. 2. Noise test set-up with ARES model in the TDT. Arrows show microphones of 

interest. 

12 
"E ,*-. 

I '. I' '\ 

i 

10: ._ 

Baseline 1P 

?Net higher 

harmonic pit 

0 

ch 

90 180 270 360 

wr Azimuth (deg) 

Fig. 3. Blade pitch angle 8 versus azimuth Y for p = 0.266 and a = 0". 

Pitch control is 4P collective with eC = -1 .Oo at Yc = 60". 

13 
t 

ao P a 

dB 

E t t t 

~5~0-0 

AZIMUTH .Dag 

750 45 

AZIMUTH ,DCg 

750 43 90 

AZIMUTH .Deg 

sor- E 

85 

E 0-0 

AZIMUTH ,Dag 

(d)p=O. 140,8=6.6 

(a=4.0.a'=4.9) 

t t 0-0 

AZIMUTH .Dag 

(e)&== 0.200.8 = 5.9. 

(a =2.0.a'- 2.4) 

t 0450 

AZIMUTH .Dag 

(f) p=o. 2 66, e = 6.4 

(a=O.O.a*=0.2) 

AZIMUTH .Dag 

(g)p = 0.3 3 3.8.. 9.4 

(a=O.O.a'=0.2) 

dE 8 0 ao IfL 

t75~ 45 GO 

(h) p=o. 170.e~3.2 (i)p=o.233.e=3.1 (j) p==o.3oo.e=4.o 

AZIMUTH .Dag AZIMUTH .Dag AZIMUTH ,Dcg 

(a=O.O.a'=0.6) (a=-2.0.a'-- 1.7) (a--4.0.a'=-3.8) 

Fig. 4. Noise level (dB) variations with different 4P pitch control amplitudes €Ic and 

azimuth Yc. Symbols are for eC=Oo (baseline), 0; for eC=-0.5O, A; and for 

€IF-1 .Op fr . Solid symbols are repeat cases. 

14 
14 

12 1. 

n 10 

6 

cn -4 

a, 

D -6 

I I I I 

0.20 0.30 

-10 -8 ti- I 

0.10 

advance ratio, JJ 

Fig. 5. Noise level (dB) contours versus flight condition for the baseline (no control) 

case. Contours based on values at grid points. Letters correspond to parls of 

Fig. 4. 

15 
14 

12 

- 10 

a8 

n 

6 CD 

75 

W 

"4 

a, 

0 

tn -4 

a, 

-6 

-8 t * 

-10 l-1 I I I I I 

0.10 0.20 0.30 

advance ratio, JJ 

Fig. 6. Noise level (dB) contours versus flight condition for the 4P pitch control 

eC = -1 .Oo and Yc = 60" case. Letters correspond to parts of figure 4. 

16 
n 

I 10 

6 CD 

-6 U 

l4 r 

12 

. 

I I I I I 

0.20 0.30 

-10 -8 t, * 

0.10 

advance ratio, p 

Fig. 7. Noise level (dB) differences between the pitch control case of Fig. 6 and the 

baseline case of Fig. 5. Negative values show noise reductions for control 

case. Letters correspond to parts of figure 4. 

17 
m woru i~,r~utcs arc Report Documentation Page SOKZ .‘o”lY’air 

1. Report No. 

NASA TM-101624 

AVSCOM TM-89-B-005 

2. Government Accession No. 

Reduction of Blade-Vortex Interaction Noise Using 

Higher Harmonic Pitch Control 

7. Author(s1 

rhomas F. Brooks, Earl R. Booth, Jr., J. Ralph Jolly,Jr. 

William T. Yeager, Jr. and Matthew L. Wilbur 

9. Performing Organization Name and Address 

NASA Langley Research Center 

and Aerostructures Directorate 

USAARTA- AVSCOM 

Hampton, VA 23665-5225 

12. Sponsoring Agency Name and Address 

National Aeronautics and Space Administration 

Washington, DC 20546-0001 and 

U.S. Army Aviation Systems Command 

St. Louis, MO 63120-1798 

5. Supplementary Notes 

3. Recipient’s Catalog No. 

5. Report Date 

July 1989 

6. Performing Organization Code 

8. Performing Organization Report No. 

10. Work Unit No. 

505-63-51-06 

11. Contract or Grant No. 

13. Type of Report and Period Covered 

Technical Memorandum 

14. Sponsoring bgency Code 

r. F. Brooks and E. R. Booth, Jr.: Langley Research Center, Hampton, Virginia. 

J. R. Jolly, Jr.: Planning Research Corporation, Hampton, Virginia. 

J. T. Yeager, Jr., and M. L. Wilbur: Aerostructures Directorate, USAARTA-AVSCOM, 

Langley Research Center, Hampton, Virginia. 

18. Abstract 

An acoustics test using an aeroelastically scaled rotor was conducted to examine the 

effectiveness of higher harmonic blade pitch control for the reduction of impulsive 

blade-vortex interaction (BVI) noise. A four-bladed, 110 in. diameter, articulated 

rotor model was tested in a heavy gas (Freon-12) medium in Langley’s Transonic 

Dynamics Tunnel. Noise and vibration measurements were made for a range of matched 

flight conditions, where prescribed (open-loop) higher harmonic pitch was superimposed on the normal (baseline) collective and cyclic trim pitch. For the inflowmicrophone noise measurements, advantage was taken of the reverberance in the hard 

walled tunnel by using a sound power determination approach. In the paper, initial 

findings from on-line data processing for three of the test microphones are reported 

for a 4/rev (4P) collective pitch control for a range of input amplitudes and phases 

By comparing these results to corresponding baseline (no control) conditions, signXicant noise reductions (4-5 dB) were found for low-speed descent conditions, 

where helicopter BVI noise is most intense. For other rotor flight conditions, the 

ovcrall noise was found to increase. All cases show increased vibration levels. 

17. Key Words (Suggested by Author(s)) 1 18. Distribution Statement 

Helicopter rotor acoustics 

Blade-vortex interaction noise 

Higher harmonic control 

Unclassified-Unlimited 

Subject Category 71 

I 

19. Security Classif. (Of this report) 22. Price 

Unclassified Unc las sif ied 18 A0 3 

20. Security Classf. (of this page) 21. No. of pages