L. Danielle Koch

Glenn Research Center, Cleveland, Ohio

An Experimental Study of Fan 

Infl ow Distortion Tone Noise

NASA/TM—2010-215844

January 2010

AIAA–2009–3290
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L. Danielle Koch

Glenn Research Center, Cleveland, Ohio

An Experimental Study of Fan 

Infl ow Distortion Tone Noise

NASA/TM—2010-215844

January 2010

AIAA–2009–3290

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

15th Aeroacoustics Conference (30th Aeroacoustics Conference)

sponsored by the American Institute of Aeronautics and Astronautics and 

Council of European Aerospace Societies

Miami, Florida, May 11–13, 2009
Acknowledgments

I sincerely appreciate the patient guidance of Ed Envia and Dan Sutliff throughout the course of this work. I would also like to thank 

Joel Lauer, Ray Loew, and the team at the AeroAcoustic Propulsion Laboratory for making the experiment a reality. This work 

was supported by the NASA Fundamental Aeronautics Program, Subsonic Fixed Wing Project in the area of noise reduction.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

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Available electronically at http://gltrs.grc.nasa.gov

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2010-215844 1 

An Experimental Study of Fan Inflow Distortion Tone Noise 

L. Danielle Koch 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

The tone noise generated when a fan ingests circumferentially distorted flow was studied by an 

experiment conducted with the Advanced Noise Control Fan at the NASA Glenn Research Center. The 

inflow was distorted by inserting cylindrical rods radially into the duct. The rods were arranged in 

circumferentially irregular patterns in three of the five configurations tested. Rods were held in place 

using a mounting ring with 30 equally spaced holes placed at an axial location one rotor chordlength 

upstream of the fan. Acoustic pressure was measured in the inlet and exhaust duct of the fan using the 

Rotating Rake fan tone measurement system. Sound power levels, calculated from the measured data, 

were plotted as a function of circumferential mode. An analytic description of the unsteady pressure 

distribution at the interaction plane between the stationary rods and the fan rotor is presented in a form 

suitable for representing the circumferentially irregularly placed rods. Terms in the analytical description 

for sound power were proven to be useful in determining the dominant circumferential modes measured 

in the experiment and the differences in mode power level between the configurations tested. Insight 

gained through this work will be useful in the development of tools to compute fan inflow distortion tone 

noise. 

Nomenclature 

a speed of sound 

amnμ axial direction of cosine of m, μ wave, n

th harmonic, =

Amnμ power-effective area for m, μ mode, n

th harmonic 

am,n direction cosine of n

th-harmonic, m-lobe mode, = 

B number of rotor blades 

cm,n amplitude of n

th harmonic, mth mode due to rotor interaction with single vane 

Cm,n amplitude of n

th harmonic, mth mode due to rotor interaction with stator assembly 

d rod diameter, ft 

f frequency, cps 

i , also used as index 

j index 

k index 

M bulk inlet Mach number 

m circumferential mode index 
NASA/TM—2010-215844 2 

n harmonic index 

N rotor shaft speed, revolutions/sec 

p local duct pressure due to rotor interaction with stator assembly 

p’ local duct pressure due to rotor interaction with single stator vane 

q stator vane index 

Re real part of operand 

St Strouhal number, 

t time coordinate 

U normalized power, = 

V number of stator vanes 

W acoustic power 

W’ approximation for acoustic power 

q (nB – m)q

qj q – j

 pressure field angular coordinate 

mn phase of m

th mode 

μ radial mode index 

mnμ cutoff ratio of m, μ mode, n

th harmonic 

mn cutoff ratio of m-lobe mode, n

th harmonic 

 circular frequency, 

 rotor shaft angular velocity 

* complex conjugate 

Introduction 

Meeting the growing demand for air transportation poses challenges to aircraft engine manufacturers 

and airlines. As our airports operate closer to their maximum capacity and more stringent community 

noise limits are imposed, it becomes more important to design quieter aircraft and follow operational 

procedures designed to reduce noise exposure in surrounding neighborhoods (Refs. 2 and 3). Fans have 

been known to be a dominant source of aircraft engine noise. Progress has been made in measuring, 

predicting, and reducing tones generated by rotor-stator interaction, which is an important contributor to 

total fan noise emissions (Ref. 4). Less attention has been given to understanding the noise generated by 

fans ingesting circumferentially non-uniform flow. This is an area of concern for those who are 

conceptualizing the next generation of quiet commercial aircraft that may use embedded engines (Ref. 5). 

How can we predict tones produced by a fan ingesting distorted inflow? To begin to answer this 

question, an experiment has been conducted at the NASA Glenn Research Center using the Advanced 

Noise Control Fan (ANCF). The purpose of the test was to measure the tone noise produced by the rotor 
NASA/TM—2010-215844 3 

blades when the inflow was distorted by a set of rods placed upstream of the rotor in circumferentially 

irregular patterns. Measurements of the acoustic pressures in the inlet and exhaust duct were obtained 

from the Rotating Rake are presented here. 

In order to explain features of the measurements, Sofrin and Mathews’ (Ref. 1) analytical description 

of acoustic pressure at the interaction plane was recast in a form suitable for representing the 

circumferentially irregularly placed rods that were tested. Terms in the analytic sound power expression 

were proven to be useful in predicting the dominant modes measured in the experiment and the 

differences in sound power levels between the configurations tested. 

Experiment 

The experiment was conducted using the NASA Glenn Advanced Noise Control Fan that is housed 

within AeroAcoustic Propulsion Laboratory (AAPL). The AeroAcoustic Propulsion Laboratory (Ref. 6) 

is a 39.6 m (130 ft) diameter geodesic dome that serves to both minimize testing noise heard by 

neighboring communities and provide an anechoic (above 125 Hz) environment for fan and jet noise 

research. 

ANCF, shown in Figure 1, is a modified commercial ventilation fan valued for being highly 

configurable and proven to be useful in furthering our understanding of fan and duct acoustics and noise 

control techniques. Originally built in the early 1990s, the fan has been primarily used in the investigation 

of rotor-stator interaction. Data from these experiments are often used to validate noise prediction codes 

under development (Ref. 7). 

The ANCF is approximately 1.2 m (48 in.) in diameter and the centerline of the rig is 3.0 m (10 ft) 

above the floor. For these tests, no stator vanes were installed downstream of the rotor. The rotor had 16 

blades (Ref. 8) set at a 28° pitch angle for all test conditions. (The coordinates for the blades are available 

upon request.) A nominal rotor blade tip clearance gap of 0.7 mm (0.03 in.) is made possible by the use of 

an abradable material built into the rotor shroud. 

An Inflow Control Device (ICD) is used to condition the flow entering the fan, removing large-scale 

turbulence and ground vortices. The ICD, shown in Figures 1 and 2, has 20 thin metal ribs and is 

constructed from 0.61 m (2 in.) thick metal honeycomb. Hub diameter was 0.46 m (18 in.) at the fan 

leading edge yielding a rotor hub-to-tip ratio,  = 0.375. The inlet plane was located 0.92 m (36.2 in.) or 

approximately 0.75 L/D ahead of the leading edge of the fan at the tip location. The exhaust plane was 

located 1.2 m (45.44 in.) or approximately 1.0 L/D downstream of the trailing edge of the fan at the tip 

location. Downstream of the test section, centerbody diameter increases to 0.71 m (24 in.) to provide the 

necessary nozzle area contraction, yielding a hub-to-tip ratio,  = 0.500. 

The inflow to the rotor was distorted by installing smooth-surfaced cylindrical rods upstream of the 

rotor. Four distortion patterns were tested, as shown in Figure 3, using the 30-rod ring. The circumferential 

positions of the rods for the various configurations are given in Table 1. Angles indicated are referenced to 

the center of the circumferential slot cut through the shroud for hot wire probe access for aerodynamic 

measurements not reported here. Angles indicated on Figure 3 increase from top dead center in the 

clockwise direction, aft looking forward. The rods were 1.27 cm (0.5 in.) in diameter and were 31.75 cm 

(12.5 in.) long, resulting in a 6.35 cm (2.5 in.) gap between the bottom of the rods and the rotor hub. The 

centerline of the rods was 14.29 cm (5.625 in.) upstream of the rotor leading edge at the tip, or 

approximately one rotor chord length upstream. Unused holes in the rod-mounting ring were covered with 

aluminum tape during testing. 

Upgraded in 2007, the rotor is driven by a 200 hp motor and speed is controllable within ±2 rpm. 

Rotor speed can vary from 0 to 2500 rpm and is detected by a 128 pulse/rev optical quadrature encoder. 

Direction of rotation is counterclockwise, forward looking aft. Data was recorded at three speed settings 

given in Table 2. 
NASA/TM—2010-215844 4 

TABLE 1.—ANGULAR POSITIONS OF RODS 

Configuration Rods Rod circumferential locations 

1 6 0°, 60°, 120°, 180°, 240°, 300°

2 6 0°, 12°, 24°, 36°, 336°, 348°

3 15 0°, 12°, 24°, 36°, 48°, 60°, 72°, 84°, 276°, 288°, 300°, 312°, 324°, 336°, 348°

4 8 0°, 12°, 24°, 108°, 120°, 192°, 216°, 348°

5 0 N/A 

TABLE 2.—TEST CONDITIONS 

Rotor rotational speed, 

rpm 

Rotor tip speed 

m/s (ft/s) 

Rotor BPF, 

Hz 

Bulk inlet 

Mach number 

1400 89 (293) 373 0.11 

1800 115 (377) 480 0.14 

2000 128 (419) 533 0.15 

Two 15-microphone arrays were used to record the farfield sound distribution from the Advanced 

Noise Control Fan. The origin of the inlet microphone array was at the centerline of the inlet plane, and 

the origin of the exhaust array was at the centerline of the exit plane. The 0° and 180° position of the 

farfield microphones are indicated on Figure 2. Microphones were placed 3.7 m (12 ft) from their 

respective origins. Angular locations for the microphones are given in Table 3. Farfield microphone 

measurements were acquired synchronously with shaft speed at a rate of 256 samples per revolution, 

which allows for analysis up to the 128th harmonic of the shaft frequency, or equivalently, up to the 8th

harmonic of the blade passing frequency (BPF) for this test (Ref. 7). 

In-duct acoustic pressure measurements upstream and downstream of the rotor were also acquired 

with the Rotating Rake (Ref. 9). The Rotating Rake system is a continuously rotating radial microphone 

rake that is inserted into the duct, generally at either the inlet entrance or exhaust exit plane. The system 

utilizes spinning mode theory and Doppler-shift physics to separate circumferential modes and the 

cylindrical wave equation solution to reduce the radial modes. It provides a complete map of the acoustic 

duct modal magnitudes and phases present in the fan duct. Details of the operation and data reduction are 

presented in Reference 9. 

 

TABLE 3.—FARFIELD MICROPHONE LOCATIONS 

Microphone 

number 

Inlet microphone 

array locations, 

deg. 

Exhaust microphone 

array locations, 

deg. 

1 0.0 90.0 

2 6.4 95.0 

3 12.9 100.0 

4 19.3 105.0 

5 25.7 110.0 

6 32.1 115.0 

7 38.6 120.0 

8 45.0 125.0 

9 51.4 130.0 

10 57.9 135.0 

11 64.3 140.0 

12 70.7 145.0 

13 77.1 150.0 

14 83.6 155.0 

15 90.0 160.0 
NASA/TM—2010-215844 5 

Analysis 

In their paper, Tyler and Sofrin (Ref. 10) analytically described the unsteady pressure distribution at 

the interaction plane between a rotor and stator. In their formulation, the blades and vanes were idealized 

by assuming equal circumferential spacing. Later, while investigating tones present in cut-off fan stage 

designs, Sofrin and Mathews (Ref. 1) analytically described the pressure distribution at the interaction 

plane between a rotor and a stator, assuming the stator vane circumferential locations deviated a small 

amount from uniform spacing. The present work extends Sofrin and Mathews’ (Ref. 1) relations in a form 

more convenient for representing irregularly spaced vanes without the small deviation angle assumption. 

Beginning with Sofrin and Mathews’ (Ref. 1) description of the pressure field at a radial and axial 

location in a duct resulting from the interaction of a rotor and a single vane, we have 

p ( ) ,t = cmn cos( ) m  nt +mn = Re

m=



cmn expi m( )   nt +mn

m=



n=1



n=1



(1) 

Recognizing that the complete pressure field is found by the superposition of the effects of V vanes 

located at varying angular distances from the reference vane we have 

p(,t) = cmn cos m( )   q  n t  t ( ) ( ) q +mn

q=0

V 1

m=



n=1



(2) 

 = 2f = 2BN = B(3) 

tq = q

 (4) 

q = 2q

V

 for stators with V equally spaced vanes (5) 

q = q for stators with V irregularly spaced vanes (6) 

p(,t) = cmn cos( ) m  n+mn  mq + nBq

q=0

V 1

m=



n=1



(7) 

p(,t) = cmn expi m( )   n+mn expi nB [ ] ( )  m q )

q=0

V 1

m=



n=1



(8)

Recognizing that the first exponential term does not vary with q, we can define the relations: 

Cmn = cmn expi nB [ ] ( )  m q

q=0

V 1

 (9) 

Cmn

cmn

= expi nB [ ] ( )  m q

q=0

V 1

 (10)
NASA/TM—2010-215844 6 

Forming this quantity is important because it appears in the expression for the total acoustic power as 

described by Sofrin and Mathews (Ref. 1) 

W = 1

2a

amnμCmnμ

μ=



Cmnμ *

m=



n=1



Amnμ (11) 

Neglecting radial variations of pressure, the “normalized” expected power is defined as 

U = amn

Cmn

cmn

2

m=1



n=1



(12) 

Cmn

cmn

2

= Cmn

cmn











 Cmn

cmn













*

 (13) 

Sofrin and Mathews continue from this point to express this quantity in terms suitable for studying vanes 

whose angular positions deviate a small amount from a uniform spacing. Here, we will express this term 

in a general way suitable for representing the case of irregularly placed vanes. 

Cmn

cmn

2

= expi nB ( )  m q expi nB ( )  m j

j=0

V 1



q=0

V 1

 (14) 

Defining the following two quantities can simplify the expression 

q = ( ) nBm q

 (15) 

j = ( ) nBm j

 (16) 

Cmn

cmn

2

= exp( ) iq exp( ) ij

j=0

V 1



q=0

V 1



(17) 

When q = j, the summation is equal to V the number of vanes or pins, and when q is not equal to j, the 

terms in the summation appear in complex conjugate pairs resulting in the following relation: 

Cmn

cmn

2

=V + exp( ) i( ) q j +exp( ) i( ) q j

j>q

 (18) 

Cmn

cmn

2

=V +2 cos( ) q  j

j>q

 =V +2 cos( ) ( ) nB m ( ) ( ) q  j

j>q

 (19) 
NASA/TM—2010-215844 7 

Values of this “Sofrin-Mathews power term,” Cmn

cmn

2

, can be useful in predicting the circumferential 

modes generated by a rotor interacting with vanes with irregular circumferential spacing. The maximum 

and minimum values for this quantity are 

Cmn

cmn

2

= V +2 q

q=0

V 1



0











 (20) 

The maximum value occurs when the cosine is equal to 1 for all of the terms in the summation which 

happens when (nB–m) is an integer multiple of V. The minimum value occurs through cancellation of 

terms in the summation. 

Comparison of Measurements and Predictions 

The spectra at 2000 rpm from the farfield microphone at 35° are shown in Figure 5 for all 

configurations tested. A time history of acoustic pressure was acquired from each of the farfield 

microphones. Since this data is recorded synchronously with rotor shaft speed, tones at harmonics of 

blade passing frequency can be calculated precisely. Sound power levels (PWL) are calculated by 

squaring the sound pressure levels (SPL), multiplying by the appropriate area, then normalizing by the 

specific acoustic impedance. 

It is important to realize that the total sound power levels computed from the farfield microphone 

arrays include noise not synchronous with the shaft speed. In these tests, noise produced by Strouhal 

shedding downstream of the rods is clearly visible on the spectra, and included in the total sound power 

levels as computed from the farfield microphone arrays. While the spectra do show that Strouhal shedding 

from the rods contributes to the total noise produced (St  0.2, f = 128 Hz), the dominant source of noise 

is still attributed to the rotor-rod interaction that generates tones at the harmonics of blade passing 

frequency. 

Tones at harmonics of blade passing frequency (BPF) are evident in Figure 5 even for the baseline 

configuration without any rods (Configuration 5). The directivity plots of Figure 6 confirm that the 

amplitude of the tones for the baseline configuration are indeed lower than those for all the rod 

configurations tested at all angles. This allows us to isolate the effects of the various rod placements. The 

source of the tones for the baseline configuration (rotor alone, no rods or stators) is not presently well 

understood and is currently an area of investigation. 

For Configuration 1 (6 evenly spaced rods), the amplitude of the sound power level of the tone at 

blade passing frequency decreases as the harmonic index increases. This trend is not displayed for the 

asymmetric cases tested at farfield angle of 35°. If we examine the trends in the 1-BPF tone at the farfield 

angle of 35°, we see that the amplitude of this tone is greatest for Configuration 4 (8 asymmetric rods), 

followed by Configuration 1 (6 equally spaced rods), Configuration 2 (6 rods, 12° apart), and then 

Configuration 3 (15 rods, 12° apart). 

The directivity plots for the tones at 1-BPF, 2-BPF and 3-BPF are given in Figure 6. The directivity 

plots show the trends in the sound pressure levels of the tones as angular location varies. The trends in the 

1-BPF tone seen in the spectra at the farfield angle of 35° for the asymmetric cases are consistent at all 

microphone locations, though the trends in the 1-BPF tone for Configuration 1 shows some angular 

variation, particularly in the aft array. The trends seen for the 2-BPF and 3-BPF tones seen in the spectra 

at the 35° farfield microphone location are not seen consistently at all angles in the directivity plots of 

Figure 6. 
NASA/TM—2010-215844 8 

Figures 7 and 8 show the measured and predicted circumferential mode sound power levels (PWL). 

The reader is reminded that the predictions are not true power levels, but are simply an apparent dominant 

term in the sound power equation. To facilitate comparison of trends, all of the predicted values shown 

are anchored to the 1-BPF measured tone for Configuration 1, and scaled in the following way: 

dB Cmn

cmn

 2















scaled

=10log10

Cmn cmn

2

36

















+121.3 (21) 

where 121.3 dB is the sound power level measured for the 1-BPF Configuration 1 (6 equally spaced rods) 

and 36 is the value of the predicted power term for the 1-BPF Configuration 1 for mode number -2. 

Changes in the sound power levels of the 1-BPF and 2-BPF tones between the various configurations 

can also be calculated from the measurements and predictions, as shown in Table 5. All asymmetric 

configurations were compared to Configuration 1 (6 equally spaced rods). The good agreement seen 

between the predicted and measured deltas is attributed to the apparent little change in the value of amn in 

Equation (12) as the number of rods was varied for a single operating condition. The predicted changes in 

sound power level for the 1-BPF tone are closest to the difference calculated from the Inlet Rotating Rake 

measurements, while the 2-BPF predicted changes are closest to the differences calculated from the entire 

Farfield Array. The predicted changes in sound power level were calculated using the following relation: 

PWLpredicted,n =10 *log10

Cmn

cmn m



 2















Cmn

cmn

2

m



















Configuration1














	



































 (22)

TABLE 5.—A COMPARISON OF PREDICTED AND MEASURED CHANGES IN SOUND POWER LEVELS 

Configurations compared Inlet rotating rake, 

dB 

Exhaust rotating rake, 

dB 

Farfield array, 

dB 

Predictions, 

dB 

1 BPF 

C2 – C1 –12.6 –12.3 –10.5 –11.5 

C3 – C1 –12.0 –8.0 –18.0 –12.3 

C4 – C1 –3.5 –3.8 0.5 –3.5 

2 BPF 

C2 – C1 0.1 –2.8 –0.9 0.9 

C3 – C1 8.4 6.23 5.9 4.9 

C4 – C1 –5.0 0.8 1.5 1.5 

Conclusion 

Fan inflow distortion tone noise was studied experimentally and analytically. Cylindrical rods 

inserted radially into the duct distorted the flow entering the fan. The rods were arranged in four 

circumferentially asymmetric patterns. Farfield and in-duct acoustic pressure measurements were 

recorded. An analytic description of acoustic pressure resulting from the interaction of a rotor and an 

asymmetric stator was presented to represent the highly asymmetric configurations tested. Trends in 

circumferential mode power levels were predicted and shown to be in good agreement with 

measurements. This predictive capability is a beginning step in developing tools to estimate the mode 

power levels produced by fans ingesting distorted inflow, an area of concern for those developing aircraft 

with embedded engines. 
NASA/TM—2010-215844 9 

References 

1. Sofrin, T.G. and Mathews, D.C., “Asymmetric Stator Interaction Noise,” J. of Aircraft, Vol. 17, 

No. 8, 1979, pp. 554–560. 

2. FAA, “FAA Aerospace Forecast Fiscal Years 2008–2025,” FAA HQ-027907.qxd, 2008. 

3. Donohue, George L., Shaver III, Russel D., Edwards, Eric (eds.), Terminal Chaos, American Institute 

of Aeronautics and Astronautics, New York, 2008. 

4. Bridges, J.E., Envia, E. and Huff, D., “Recent Developments in U.S. Engine Noise Reduction 

Research,” NASA/TM—2001-211083, 2001. 

5. Hall, C.A. and Crichton, D., “Engine and Installation Configurations for a Silent Aircraft,” ISABE2005-1164, 2005. 

6. Cooper, B., “A Large Hemi-Anechoic Enclosure for Community-Compatible Aeroacoustic Testing of 

Aircraft Propulsion Systems,” NASA TM–106015, 1993. 

7. Loew, R.A., Lauer, J.T., McAllister, J. and Sutliff, D.L., “The Advanced Noise Control Fan,” 

NASA/TM—2006-214368, 2006. 

8. Langford, M.D., Minton, C., Ng, W.F., Burdisso, R.A. and Halasz, C., “Fan Flow Control for Noise 

Reduction Part 3: Rig Testing of Optimal Design,” AIAA–2005–3027, 2005. 

9. Sutliff, D.L., “Rotating Rake Turbofan Duct Mode Measurement System,” NASA/TM—2005-

213828, 2005. 

10. Tyler, J.M. and Sofrin, T.G., “Axial Flow Compressor Noise Studies,” Journal of Aircraft, Vol. 70, 

No. 8, 1962, pp. 309–332. 
NASA/TM—2010-215844 10 

 Figure 1.—The Advanced Noise Control Fan Rig in the NASA Glenn AeroAcoustic Propulsion Lab. 

Figure 2.—Cross-sectional diagram of Advanced Noise Control Fan for the Inlet Distortion Tests. 
NASA/TM—2010-215844 11 

Figure 3.—Circumferential Locations of Rods Tested. 

Figure 4.—Photograph of Advanced Noise Control Fan Rig. 
NASA/TM—2010-215844 12 

Figure 5.—Spectra from Farfield Microphone at 35° in the Inlet Array. 
NASA/TM—2010-215844 13 

Figure 6.—Directivity Plots from the Farfield Microphone Arrays. 
NASA/TM—2010-215844 14 

Figure 7.—Measured and Predicted In-duct Circumferential Mode Power Levels—I-BPF. 
NASA/TM—2010-215844 15 

Figure 8.—Measured and Predicted In-duct Circumferential Mode Power Levels—2-BPF. 
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An Experimental Study of Fan Inflow Distortion Tone Noise 

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

The tone noise generated when a fan ingests circumferentially distorted flow was studied by an experiment conducted with the Advanced 

Noise Control Fan at the NASA Glenn Research Center. The inflow was distorted by inserting cylindrical rods radially into the duct. The 

rods were arranged in circumferentially irregular patterns in three of the five configurations tested. Rods were held in place using a 

mounting ring with 30 equally spaced holes placed at an axial location one rotor chordlength upstream of the fan. Acoustic pressure was 

measured in the inlet and exhaust duct of the fan using the Rotating Rake fan tone measurement system. Sound power levels, calculated 

from the measured data, were plotted as a function of circumferential mode. An analytic description of the unsteady pressure distribution at 

the interaction plane between the stationary rods and the fan rotor is presented in a form suitable for representing the circumferentially 

irregularly placed rods. Terms in the analytical description for sound power were proven to be useful in determining the dominant 

circumferential modes measured in the experiment and the differences in mode power level between the configurations tested. Insight 

gained through this work will be useful in the development of tools to compute fan inflow distortion tone noise. 

15. SUBJECT TERMS

Ducted fans; Noise 

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STI Help Desk (email:help@sti.nasa.gov) 

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