M. Nallasamy

QSS Group, Inc., Cleveland, Ohio

E. Envia and S.A. Thorp

Glenn Research Center, Cleveland, Ohio

A. Shabbir

University of Toledo, Toledo, Ohio

Fan Noise Source Diagnostic Test Computation

of Rotor Wake Turbulence Noise

NASA/TM—2002-211798

August 2002

AIAA–2002–2489
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M. Nallasamy

QSS Group, Inc., Cleveland, Ohio

E. Envia and S.A. Thorp

Glenn Research Center, Cleveland, Ohio

A. Shabbir

University of Toledo, Toledo, Ohio

Fan Noise Source Diagnostic Test Computation

of Rotor Wake Turbulence Noise

NASA/TM—2002-211798

August 2002

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

Eighth Aeroacoustics Conference

cosponsored by the American Institute of Aeronautics and Astronautics

and the Confederation of European Aerospace Societies

Breckenridge, Colorado, June 17–19, 2002

AIAA–2002–2489
Acknowledgments

The authors wish to acknowledge the help and support of many colleagues who contributed to this work.

In particular, we would like to thank Tim Beach, John Gallagher, Mark Celestina, Richard Mulac, Vijay Garg,

 John Admaczyk, and Richard Woodward.

Available from

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

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Available electronically at http://gltrs.grc.nasa.gov
NASA/TM2002-211798 1 

FAN NOISE SOURCE DIAGNOSTIC TEST

COMPUTATION OF ROTOR WAKE TURBULENCE NOISE

M. Nallasamy 

QSS Group, Inc. 

Cleveland, Ohio 44135 

E. Envia and S.A. Thorp 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

A. Shabbir 

University of Toledo 

Toledo, Ohio 43606

ABSTRACT

An important source mechanism of fan broadband noise is the interaction of rotor 

wake turbulence with the fan outlet guide vanes. A broadband noise model that utilizes 

computed rotor flow turbulence from a RANS code is used to predict fan broadband 

noise spectra. The noise model is employed to examine the broadband noise 

characteristics of the 22-inch Source Diagnostic Test fan rig for which broadband noise 

data were obtained in wind tunnel tests at the NASA Glenn Research Center. A 9-case 

matrix of three outlet guide vane configurations at three representative fan tip speeds are 

considered. For all cases inlet and exhaust acoustic power spectra are computed and 

compared with the measured spectra where possible. In general, the acoustic power levels 

and shape of the predicted spectra are in good agreement with the measured data. The 

predicted spectra show the experimentally observed trends with fan tip speed, vane count 

and vane sweep. The results also demonstrate the validity of using CFD-based turbulence 

information for fan broadband noise calculations. 

INTRODUCTION

Owing to the success of fan tone noise 

reduction techniques [1], fan broadband noise is 

now widely acknowledged as the next major 

obstacle to be overcome in attempts to reduce 

aircraft engine noise. But, unlike tone noise, the 

sources of fan broadband noise are many and most 

are not as well understood. In response to this 

challenge a number of theories [2–5] have been 

devised to help elucidate the nature and 

characteristics of fan broadband noise. These 

theories require flow turbulence characteristics 

(i.e., spectrum, intensity, length scales, etc.) as 

input for computing noise spectra. Typically, such 

information is available from an experiment [6] or 

is simply guessed at to provide the best fit of the 

predicted noise spectra to the measured ones. An 

alternative approach is to compute the turbulence 

characteristics using a Reynolds-Averaged NavierStokes (RANS) type calculation, since that would 

provide a true prediction capability. 

This paper presents one such attempt that 

combines predicted flow turbulence characteristics 

obtained from a turbomachinery CFD code with a 
NASA/TM2002-211798 2 

fan broadband noise theory to compute fan noise 

spectra. The flow code, called APNASA [7], is a 

three-dimensional Navier-Stokes solver with an 

improved k − ε turbulence model [8]. The noise 

code is a revised and enhanced version of an 

analytical noise prediction theory originally 

developed by Ventres et al. [9]. The testbed is a 

1/5-scale model of a representative high bypass 

ratio turbofan engine for which extensive 

aerodynamic and acoustic data were obtained [10–

12]. Only the broadband noise produced by the 

interaction of fan wake turbulence with the stator 

outlet guide vanes (OGV) will be considered here. 

The data theory comparisons will be done on an 

acoustic power basis since the noise theory 

furnishes acoustic power predictions only. 

The 22-inch diameter model fan was tested in 

the NASA Glenn Research Center’s 9′×15′

acoustic wind tunnel. The fan has 22 blades and a 

design tip speed of 1215 ft/sec. The baseline OGV 

has 54 radial vanes and represents a conventional 

“cut-off” stator design (i.e., the blade passing 

frequency is cut off). The low-count OGV is a 

“cut-on” concept which has only 26 radial vanes 

but has the same solidity as the baseline stator. The 

smaller vane count was selected to reduce 

broadband noise [13]. The swept OGV has 26 

swept vanes (with 30 degree sweep) and was 

designed to minimize the blade passing frequency 

tone penalty associated with a cut-on design. 

The aerodynamic and acoustic tests were 

conducted over a wide range of fan operating 

conditions for each of the three OGV packs. For 

the purposes of the current work, however, only 

the speeds corresponding to the noise certification 

points will be discussed. The relevant fan operating 

conditions are listed in Table 1. 

Power 

Setting 

Corrected 

Fan rpm 

Corrected Fan 

Tip Speed, fps 

Approach 7,808 750 

Cutback 11,074 1,060 

Takeoff 12,656 1,215 

 

Table 1: Fan operating conditions of interest. 

In what follows, a summary of the analysis 

tools used in this work will be presented followed 

by a discussion of the theoretical results and their 

comparison with measured data (where possible). 

The paper concludes with a summary of important 

results. 

ANALYSIS

The input for the noise prediction was 

generated by means of a series of CFD simulations 

of the fan rotor coupled with the fan outlet guide 

vane. The code used to generate the fan flow field 

is APNASA (see Ref. 7), which solves for the time 

average flow field in a typical passage of a blade 

row embedded within a multistage turbomachine. 

The simulations are three-dimensional and include 

the effect of viscosity. A modified k − ε

turbulence model (see Ref. 8) was used to account 

for the effect of turbulence mixing. However, 

before these CFD simulation could be undertaken 

it was necessary to establish the fan rotor geometry 

at tests conditions. 

Fan Running (Hot) Shape

Fan rotors un-twist as they are run-up in 

rotational speed. It is well known that the effect of 

rotation on fan rotor shape has a non-trivial impact 

on aerodynamic performance. Thus, any CFD 

simulation of the fan rotor based on fan geometry 

other than the running geometry would be 

pointless. For the present study the chosen fan 

operating conditions included off-design 

conditions. Thus, a structural analysis was 

undertaken to determine the fan shape at the 

conditions corresponding to those of the test 

matrix. This analysis was conducted using the 

SABER [14] coupled aero/structural analysis 

system. An outline of the SABER system is 

presented in Figure 1. Starting with the fan rotor 

shape as manufactured (CAD geometry definition 

of the cold shape) the SABER system generates 

input for a structural analysis code. Included in the 

input is the rotational speed of the fan rotor. 

Output from the structural analysis code is 

post-processed to generate a flow field grid for a 

CFD simulation. The output of the CFD simulation 

is post-processed by SABER to yield the pressure 

and temperature distribution on the fan rotor 

surface. This information is supplied to the 
NASA/TM2002-211798 3 

structural analysis pre-processor, which then 

updates the input information to the structural 

analysis code and the cycle repeated until 

convergence (i.e., no significant change in fan 

rotor geometry) is obtained. 

The transfer of information between the 

various elements of SABER is by means of NIGES 

and GGA files. The surface definition of fan rotor 

geometry, pressure and temperature distribution is 

in the form of NURBS [15–17]. 

The fan rotor geometry is established at 

running conditions (hot shape) along with its flow 

field upon convergence of this iterative process. 

Upon convergence of the iterative cycle within 

SABER the output takes full account of centrifugal 

loading, loading due to aerodynamic forces, and 

the temperature imposed on the fan rotor by the 

flow field. 

The execution of SABER parallels the process 

used in a fan rotor mechanical design. A total of 

three coupled structural aero analysis were 

executed corresponding to the tip speeds of interest 

(i.e., the approach, cutback and takeoff conditions). 

Using the computed fan hot shapes, fan speed 

lines (constant corrected speed) were established 

using APNASA for the nine (3 speeds × 3 OGV) 

test cases. 

Rotor Flowfield

The CFD grid has 407 grid cells in the axial 

direction, 51 grid cells in the radial direction and 

51 grid cells in the azimuthal direction. Two views 

of the grid used in the flowfield simulations are 

shown in Figures 2 and 3. Figure 2 shows the 

meridional grid definition where locations of the 

fan rotor and outlet guide vane are noted. Figure 3 

shows the fan rotor blade-to-blade grid layout at 

midspan. The streamline grid plane was designed 

to nearly coincide with the trajectory of the rotor 

wake in order to reduce numerical diffusion of the 

fan wakes. 

Figure 4 shows the computed axial velocity 

field associated with the rotor at 75% of span at the 

takeoff condition. The wakes are well defined all 

the way to the leading edge plane of the outlet 

guide vane. Figure 5 shows the corresponding 

distribution of turbulent kinetic energy. It too is 

well defined to the leading edge plane of the outlet 

guide vane. 

Figure 6 shows a comparison of the CFD 

results for the baseline configuration compared to 

data at the takeoff condition (see Ref. 6). The plots 

show the comparisons for an axial location 2 

inches downstream of the rotor trailing edge (at the 

tip). Shown are the time averaged axial velocity, 

absolute tangential velocity and a measure of the 

intensity of the turbulent field. The agreement 

between the computed results and the 

corresponding measured results is good. The 

agreement extends to the end-wall region and 

includes the region of flow associated with the fan 

rotor tip clearance flow. 

The turbulence intensity and length scale 

derived from the CFD simulations at the stator 

leading edge were used to estimate the frequency 

spectrum of the turbulent field impinging on the 

outlet guide vane. This information and the time 

average flow field of the outlet guide vane formed 

the input to the acoustic predictions. It should be 

noted that the input was used “as is” and no 

adjustments were made to the input to achieve a 

better agreement with data. 

Rotor Wake Turbulence - Stator Noise Model

The fan broadband noise prediction model 

used for this study is based on a theory developed 

by Ventres et al. (see Ref. 9). However, significant 

corrections, revisions and enhancement were made 

to the theory and a new code was written to take 

full advantage of the flow turbulence information 

available from the APNASA solutions. Details of 

the modified theory and the associated computer 

code will be reported in a future paper. Here, only 

salient features of the noise model are presented. 

Predicated on a modal description of the noise 

field inside the fan duct, the theoretical model 

relates the expected values of modal acoustic 

power levels to the expected values of the incident 

turbulence. The turbulence is assumed locally 

isotropic at each radius with its statistics 

parametrically dependent on the radius. The 

unsteady aerodynamic response is based on a 
NASA/TM2002-211798 4 

strip-wise description that relates the OGV 

unsteady pressure distribution at each radial 

location to the incident wake turbulence. The 

underlying unsteady response is the same as that 

used for tone calculation [18]. The resulting Quasi3D distribution of the expected values of vane 

unsteady pressure field are then related to the 

estimates of the modal power via the classical 

Green’s function formulation. The end results is 

the mode power level spectral density, Pmn ( ) ω , 

which in short-hand notation may be written as 

() ( ) () ( ) ∫ = Ψ

T

H

r

r

mn mn mn P C r, r r, dr 2 2

ω ω φ ω (1) 

where Cmn is the chordwise integral of the 

unsteady pressure distribution at each radius, φmn

the annular duct mode shape, and the function Ψ

is related to spectrum of the incident turbulence. 

 

It should be pointed out that in deriving 

Eq. (1), the OGV is modeled as an ensemble of 

zero thickness flat plates in an infinite hard-wall 

annular duct containing a uniform axial mean flow. 

The parameters controlling the aerodynamic 

response are the absolute Mach number of the 

mean flow and vane chord, vane count and stagger 

angles at each radius (i.e., strip). In particular, in 

this work, the vane cascade stagger angle at each 

radius is defined as a weighted average of the 

airfoil section camber line with the leading edge 

receiving more weight as compared to the trailing 

edge in recognition of the fact that high-frequency 

response of the vane is concentrated near the 

leading edge. In addition to the aero input, the 

noise calculations also require the r.m.s. amplitude 

of the upwash turbulent fluctuations and integral 

length scales of turbulence in the axial, tangential 

and radial directions. Implicit in the derivation of 

Eq. (1) is the additional assumption that the radial 

integral scale is small compared with the span of 

the vane. Furthermore, the circumferential 

distribution of the turbulence intensity is taken to 

be of the form 

( ) () () 2 2 /

b , Lw

w u r u r u r e ξ ξ − ′ = ′ + ′ (2) 

where b u′ and w u′ are the background and wake 

turbulence intensities at each radial strip, and ξ is 

the distance across one blade pitch measured in the 

direction normal to the mean wake centerline. The 

parameter Lw will be defined shortly. b u′ and w u′

are computed from the CFD solution via 

 1/ 2 u′ = (2k /3) (3) 

where k is the computed local turbulence kinetic 

energy of turbulence. Eq. (2) is in excellent 

agreement with the pitch-wise distribution of 

computed intensity. Lw is chosen to be an integral 

length which multiplied by w u′ equal to the area 

contained under the computed profile (Figure 7). 

The distribution described by Eq. (2) is assumed to 

apply to all fan blade passages so that the pattern 

repeats around the annulus. Note that, this does not 

mean that the instantaneous turbulent fluctuations 

are the same for corresponding points in passages, 

but only their statistics. The same argument of 

course applies to the length scales. The length 

scales are estimated from the computed RANS 

flowfield solutions via 

/ ε 3 / 2 A = k (4) 

where k and ε are the circumferentially averaged 

turbulence kinetic energy and dissipation rate. 

Defined in this manner, A is the average size of the 

largest energy containing eddies in the flow. The 

parameter A represents the longitudinal and A / 2

the lateral integral length scales of the incident 

turbulence (i.e., the isotropic assumption). 

The variations (as a functions of radius) of the 

local Mach number, vane stagger angle, incident 

turbulence intensity and integral length scales, and 

the effective width of the intensity profiles are 

inputs for noise spectra computations. For each 

case considered in this work, this information was 

extracted from the corresponding CFD solution. 

Measured Noise Spectra for Comparisons

Farfield noise spectra were measured for all 

three OGV configurations as a function of fan tip 

speed (see Ref. 12). Using an acoustically treated 
NASA/TM2002-211798 5 

“barrier” wall, the contributions of the inlet and 

exhaust noise to the farfield spectra could be 

separated experimentally. 

Additionally, since the model fan could be run 

in a “rotor-alone” configuration (see Ref. 11), it is 

also possible to isolate the contribution of OGV 

noise by subtracting the rotor-alone contribution 

from the rotor plus stator contribution on a power 

basis (see Figure 8). Figure 8a shows measured 

spectrum of total acoustic power (inlet plus 

exhaust) for rotor-stator combination from which 

rotor-alone contribution is subtracted to get the 

OGV power level. Figure 8b shows the use of the 

measurements with the barrier wall (which blocks 

the exhaust radiation reaching the inlet). Here 

again, rotor-stator combination and rotor-alone 

spectra are available. This enables the separation of 

OGV inlet noise spectrum. Subtracting the inlet 

contribution from the inlet plus exhaust spectrum 

for OGV in Figure 8a, the exhaust spectrum of 

OGV alone is obtained (see Figure 8c). In practice, 

however, that proved problematic for all but the 

approach condition, since at high fan tip speeds, 

the rotor-alone levels tend to be higher than the 

stage (i.e., rotor plus stator) levels. Furthermore, 

since the current noise model does not account for 

the swirl between the rotor and stator or the 

transmission loss effects through the rotor, inlet 

noise predictions do not reflect the correct physics 

and as such are not suitable direct for comparisons 

with data. 

Therefore, the data-theory comparisons are 

restricted to the exhaust noise at the approach 

speed only. However, predicted trends with the fan 

tip speed for both the inlet and exhaust at the three 

speeds are presented. 

RESULTS AND DISCUSSION

For all nine cases the inlet and exhaust duct 

acoustic power levels were computed for 

frequencies up to 50 KHz. This range corresponds 

approximately from 0.5BPF to 11BPF for takeoff, 

0.5BPF to 13BPF for cutback, and 0.5BPF to 

18BPF for approach. 

The spectral calculations were typically done 

at 0.5BPF increments, since the number of cut-on 

modes that must be computed increases so rapidly 

that smaller frequency increments would have 

significantly increased the computational time. In 

the following figures, the symbols denote the 

frequency points at which the spectral calculations 

were actually performed and the lines through the 

symbols are curve fits to provide an idea of the 

spectral shapes. 

It should be pointed out that, occasionally, the 

computed amplitude of the aerodynamic response 

of one of the strips would be inordinately large due 

to local cascade resonance condition. This, in turn, 

would lead to a “wild” point in the associated 

acoustic response. By slightly adjusting the 

aerodynamic input parameters, the resonance point 

could be avoided. However, in this paper the 

occasional wild point has been eliminated from the 

computed spectra. In a 3D response, such local 

resonance conditions would not occur. 

 a. Comparisons with the measured spectra

First, the data-theory comparison is presented 

in terms of the acoustic power spectra for the three 

stator configurations. Figures 9 through 11 show 

the predicted and measured spectra for the baseline 

54 radial vanes, 26 radial vanes, and 26 swept 

vanes configuration, respectively. Note that the 

measured spectra contain blade passing frequency 

(BPF) tone harmonics. While the BPF tone is 

cutoff (absent) for the baseline case, it is cut-on 

(present) for the other two cases. The current 

predictions have only the broadband noise 

component. The predicted shape of the spectra is in 

fairly good agreement with the measured spectra 

including the high frequency fall off. However, 

there is an under-prediction of levels, which may 

arise from several factors. The main one is the fact 

that the measurements may have contributions 

from additional noise sources while the predictions 

account for only the wake turbulence. Also, at 

high frequencies quadrupole noise may become 

important, as discussed by Mani et al. [19], and 

should be included to improve the prediction. 

b. Effect of fan tip speed on acoustic power spectra

The variation of acoustic power with fan tip 

speed is shown for the 54 radial vane configuration 

in Figures 12 and 13. The figures show inlet and 
NASA/TM2002-211798 6 

exhaust acoustic power spectra for the three tip 

speeds of interest. The spectra show the expected 

trend with the tip speed. 

c. Inlet and exhaust noise spectra

The computed inlet and exhaust acoustic 

power spectra are compared in Figure 14 for the 54 

radial stator vane configuration at takeoff 

condition. The inlet noise is substantially lower 

than that of the exhaust as observed in the wind 

tunnel measurements. 

d. Variation of noise with vane number

As the vane count is reduced, the rotor-wake 

turbulence generated broadband noise is expected 

to reduce substantially. The cut-on fan design is 

expected to produce less broadband noise than the 

cut-off design to counteract the cut-on tone levels. 

Figures 15 and 16 show inlet and exhaust 

power spectra for 54 radial vane and 26 radial vane 

configurations at approach. The acoustic power for 

26 radial vanes is substantially lower than for the 

54 radial vanes configuration. The power is 

expected to vary as 2 10log NV (see Ref. 9) where 

NV is the vane count. This should result in 6.3 dB 

reduction in power levels for the 26-vane OGV 

compared with the 54-vane OGV. The predicted 

reduction in power due to the reduction in vane 

number is higher than expected at high 

frequencies. At frequencies lower than 5.5 kHz, the 

noise levels produced by the two configurations 

are not significantly different probably because at 

low frequencies there are additional noise sources 

present, which are not modeled in the present 

theory. A similar trend was observed in Boeing 

broadband noise experiments (see Ref. 13). 

e. Effect of vane sweep

Vane sweep has been found to result in 

substantial reduction in tone noise levels compared 

to the radial vane configuration [20]. Woodward 

(see Ref. 12) found that the vane sweep reduces, in 

addition, the broadband noise by one to two 

decibels at certain speeds. The measured spectra 

for the approach exhaust radiation show (see 

Figure 17) a small reduction in noise level in the 

low frequency range (<10 kHz) due to vane sweep. 

It should be emphasized that if tones are removed 

from the measured spectra the reduction will be 

even smaller. The predictions do not show any 

decrease in acoustic power levels for the 26 swept 

geometry compared to the 26 vane radial (see 

Figure 18). 

 

f. Acoustic power as a function of mode order m

Figure 19 shows the inlet and exhaust acoustic 

power as function of circumferential mode m order 

for frequencies from 3BPF, 5BPF and 7BPF. The 

number of cut-on circumferential mode orders (m)

and the number of cut-on radial mode orders (n) in 

each m-order increase linearly with frequency (see 

Figure 20). In general, the peak acoustic power 

level reduces with increases in frequency. At 

frequencies higher than 5BPF, the fall off with 

increasing m order is higher for the positive m 

orders than for negative m orders. 

g. Cutoff ratio and acoustic power

Sometimes it is helpful to look at the acoustic 

power radiated as a function of cutoff ratio and that 

is particularly done in design and tradeoff studies 

[21]. Figure 21 shows the variation of acoustic 

power as a function of cutoff ratio for exhaust 

radiation at approach for two frequencies, 4BPF 

and 6BPF. The symbols denote different m orders 

and five n orders (n = 0 – 4) are included in the 

plot. Since no effect of flow swirl is included in the 

current model, the same number of co-rotating and 

contra-rotating modes are cut-on at each 

frequency. The acoustic power decreases sharply 

as the cutoff ratio is increased and eventually 

reaches nearly a plateau beyond a cutoff ratio of 

three for both frequencies. 

CONCLUDING REMARKS

The rotor wake turbulence stator interaction 

broadband noise has been computed for 

representative fan operating conditions for three 

fan stage configurations. The computations employ 

the wake flow turbulence information from 

accurate computational fluid dynamic solutions. 

The predicted noise spectra show the observed 

trends with fan tip speed and vane count. The 
NASA/TM2002-211798 7 

predicted acoustic power levels and shape of the 

spectra show reasonable agreement with the 

measured spectra for the exhaust noise at approach 

condition. 

REFERENCES

1. Envia, E., “Fan Noise Reduction: An 

Overview,” International Journal of 

Aeroacoustics, Vol. 1 No. 1, pp. 43–65, 2002. 

2. Glegg, S.A.L. and Jochault C., “Broadband 

Self Noise from a Ducted Fan,” AIAA Paper 97–

1612, May 1997. 

3. Gouville, B., Roger, M., Cailleau, J.M., 

“Prediction of Fan Broadband Noise,” AIAA Paper 

98–2317, 1998. 

4. Morin, B.L., “Broadband Fan Noise Prediction 

System for Gas Turbine Engines,” AIAA paper 

99–1889, 1999. 

5. Hanson, D.B., “Broadband Noise of Fans – 

With Unsteady Coupling Theory to Account Rotor 

and Stator Reflection/Transmission Effects, 

NASA/CR2001-21136, November 2000. 

6. Podboy, G.G., Krupar, M.J., Helland, S.M., 

and Hughes, C.E., “Steady and Unsteady Flow 

Field Measurements within a NASA 22 Inch Fan 

Model,” AIAA Paper 2002-1033, January 2002. 

7. Adamczyk, J.J., Celestina, M.L., Beach, T.A., 

and Barnett, M., 1990, “Simulation of Threedimensional Viscous Flow within a Multistage 

Turbine,” Tran. ASME, 112, 370–376. 

8. Shih, T. H., Liou, W.W., Shabbir, A., Zhu, J., 

and Yang, Z., 1995, “A New k − ε Eddy 

Viscosity Model for High Reynolds Number 

Turbulent Flows,” Computers in Fluids, 24, 3, 

227–238. 

9. Ventres, C.S, Theobald, M.A., and Mark, W.D., 

“Turbofan Noise Generation Vol. 1: Analysis,” 

NASA CR–167951, July 1982. 

10. Hughes, C.E., “Aerodynamic Performance of 

Scale-Model Turbofan Outlet Guide Vanes 

Designed for Low Noise,” AIAA Paper 2002-

0374, January 2002. 

11. Hughes, C.E., Jeracki, R.J., and Miller, C.J., 

“Fan Noise Source Diagnostic Test – Rotor-Alone 

Nacelle Aerodynamic Performance Results,” 

AIAA–Paper 2002–2426, June 2002. 

12. Woodward, R.P., Hughes, C.E., Jeracki, R.J., 

and Miller, C.J., “Fan Noise Source Diagnostic 

Test – Farfield Acoustic Results,” AIAA–Paper 

2002–2427, June 2002. 

13. Ganz, U.W., Joppa, P.D, Patten, T.J., and 

Scharpf, D.F., “Boeing 18-Inch Fan Rig 

Broadband Noise Test”, NASA/CR1998-

208704, September 1998. 

14. Thorp, S.A., Trowbridge D., Bartos K.F., 

SABER - Structural Airfoil and Blade Engineering 

Routines, User Manual, June 1998, 104 pgs. 

15. Theory Document, DT_NURBS Spline 

Geometry Subprogram Library Version 3.6, 

Craderock Division, Naval Surface Warfare 

Center, October 1998. 

16. Users Manual, DT_NURBS Spline Geometry 

Subprogram Library, Version 3.6, Craderock 

Division, Naval Surface Warfare Center, October 

1998. 

17. Reference Manual, DT_NURBS Spline 

Geometry Subprogram Library, Version 3.6, 

Craderock Division, Naval Surface Warfare 

Center, October 1998. 

18. Meyer, H.D. and Envia E., “Aeroacoustic 

Analysis of Turbofan Noise Generation,” 

NASA/CR–4715, March 1996. 

19. Mani, R, Gliebe, P.R., and Ho, P.Y., “Fan 

Broadband Noise Model Development,” NASA 

CR–198457, December 1997. 

20. Envia, E and Nallasamy, M., “Design 

Selection and Analysis of Swept and Leaned Stator 

Concept,” Journal of Sound and Vibration, Vol. 

228, pp. 793–836, 1999. 
NASA/TM2002-211798 8 

21. Rice, E.J., “Broadband Noise Radiation 

Models for Aircraft Engines”, Final Report, May 

1998. 

Figure 1. Outline of SABER system.

Figure 2. A meridional view of the 

computational mesh showing the 

rotor and baseline stator. 

Figure 3. Blade-to-blade view of the mesh at 

midspan. 

Figure 4. Contour lines of the axial velocity 

showing the blade wake evolution in 

the blade-to-blade plane at 75% span. 

Figure 5. Contour lines of the turbulence 

kinetic energy at 75% span. 
NASA/TM2002-211798 9 

Figure 6. Comparison of experimental (top) and 

CFD (bottom) results for the baseline 

configuration at takeoff at 2 inches 

behind the rotor. 

Figure 7. A representative turbulence intensity 

profile. Dashed line denotes the CFD 

solution and the solid line the fit using 

Eq. (2). Figure 8. Separation of measured inlet noise 

from exhaust noise and rotor noise 

from stator (OGV) noise. Frames (a) 

through (c) show various stages of 

separation. Baseline: Approach. 
NASA/TM2002-211798 10 

Figure 9. Comparison of computed and 

measured spectra for the baseline 

stator (54 radial vanes) at approach in 

the exhaust. 

Figure 10. Comparison of computed and 

measured spectra for the low-count 

stator (26 radial vanes) at approach in 

the exhaust.

Figure 11. Comparison of computed and 

measured spectra for the low-noise 

stator (26 swept vanes) at approach in 

the exhaust. 

Figure 12. Effect of fan tip speed on inlet 

acoustic power spectrum for the 

baseline stator. Solid line: approach, 

dashed line: cutback, and dotted line: 

takeoff.
NASA/TM2002-211798 11 

Figure 13. Effect of fan tip speed on exhaust 

acoustic power spectrum for the 

baseline stator. Solid line: approach, 

dashed line: cutback, and dotted line: 

takeoff. 

Figure 14. Inlet and exhaust power spectra 

comparisons for the baseline stator at 

takeoff. Solid line: exhaust, dashed 

line: inlet. 

Figure 15. Effect of vane count on inlet acoustic 

power spectrum at approach. Solid 

line: 54-vane OGV, dashed line: 26-

vane OGV. 

Figure 16. Effect of vane count on exhaust 

acoustic power spectrum at approach. 

Solid line: 54-vane OGV, dashed line: 

26-vane OGV. 
NASA/TM2002-211798 12 

Figure 17. Measured effect of vane sweep on 

exhaust acoustic power spectrum at 

approach.

Figure 18. Predicted effect of vane sweep on 

exhaust acoustic power spectrum at 

approach. Solid line: 26-vane radial 

OGV, dashed line: 26-vane swept 

OGV.

Figure 19. Inlet acoustic power distribution as a 

function of spinning mode order (m) 

for the baseline stator at approach. 

Solid line: 3BPF, dashed line: 5BPF, 

dotted line: 7BPF. 

Figure 20. Variation of the number of cut-on 

modes with frequency. Circles: m

orders, triangles: n orders. 
NASA/TM2002-211798 13 

Figure 21. Inlet acoustic power as a function of 

cut-off ratio for the baseline stator at 

approach. 
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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Technical Memorandum

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National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

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19

Fan Noise Source Diagnostic Test Computation of Rotor Wake Turbulence Noise

M. Nallasamy, E. Envia, S.A. Thorp, and A. Shabbir

Acoustics; Fan noise; Broadband noise; Computational fluid dynamics

Unclassified -Unlimited

Subject Category: 71 Distribution: Nonstandard

Prepared for the Eighth Aeroacoustics Conference cosponsored by the American Institute of Aeronautics and Astronautics

and the Confederation of European Aerospace Societies, Breckenridge, Colorado, June 17–19, 2002. M. Nallasamy,

QSS Group, Inc., Cleveland, Ohio 44135; E. Envia and S.A. Thorp, NASA Glenn Research Center; A. Shabbir, University

of Toledo, Toledo, Ohio. Responsible person, M. Nallasamy, organization code 5940, 216–977–1287.

An important source mechanism of fan broadband noise is the interaction of rotor wake turbulence with the fan outlet guide

vanes. A broadband noise model that utilizes computed rotor flow turbulence from a RANS code is used to predict fan

broadband noise spectra. The noise model is employed to examine the broadband noise characteristics of the 22-inch

Source Diagnostic Test fan rig for which broadband noise data were obtained in wind tunnel tests at the NASA Glenn

Research Center. A 9-case matrix of three outlet guide vane configurations at three representative fan tip speeds are

considered. For all cases inlet and exhaust acoustic power spectra are computed and compared with the measured spectra

where possible. In general, the acoustic power levels and shape of the predicted spectra are in good agreement with the

measured data. The predicted spectra show the experimentally observed trends with fan tip speed, vane count, and vane

sweep. The results also demonstrate the validity of using CFD-based turbulence information for fan broadband noise

calculations.