David B. Stephens and Edmane Envia

Glenn Research Center, Cleveland, Ohio

Acoustic Shielding for a Model Scale

Counter-rotation Open Rotor

NASA/TM—2012-217227

January 2012

AIAA–2011–2940
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David B. Stephens and Edmane Envia

Glenn Research Center, Cleveland, Ohio

Acoustic Shielding for a Model Scale

Counter-rotation Open Rotor

NASA/TM—2012-217227

January 2012

AIAA–2011–2940

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

17th Aeroacoustics Conference

sponsored by the American Institute of Aeronautics and Astronautics and the Confederation of European 

Aerospace Societies

Portland, Oregon, June 5–8, 2011
Acknowledgments

This paper was funded by the NASA Environmentally Responsible Aviation project with Dr. Fay Collier as the project 

manager, and by the NASA Subsonic Fixed Wing project with Dr. Ruben Del Rosario as the project manager. The diffraction 

shielding calculations were performed using code written by Bruce Walker from Channel Island Acoustics. NASA researchers 

Dale Van Zante and John Gazzaniga were responsible for managing the open rotor test and the 9 by 15 Wind Tunnel crew 

assembled and operated the test hardware.

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Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

Level of Review: This material has been technically reviewed by technical management. 
Acoustic Shielding for a Model Scale

Counter-rotation Open Rotor

David B. Stephens and Edmane Envia

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, OH 44135

Abstract

The noise shielding benefit of installing an open rotor above a simplified wing or tail is explored

experimentally. The test results provide both a benchmark data set for validating shielding prediction

tools and an opportunity for a system level evaluation of the noise reduction potential of propulsion noise

shielding by an airframe component. A short barrier near the open rotor was found to provide up to

8.5 dB of attenuation at some directivity angles, with tonal sound particularly well shielded. Predictions

from two simple shielding theories were found to overestimate the shielding benefit.

Nomenclature

f Frequency, Hz

Bf Number of blades on the front rotor

Ba Number of blades on the aft rotor

Nf Front rotor shaft rotation frequency, Hz

Na Aft rotor shaft rotation frequency, Hz

ps Complex acoustic pressure when the acoustic shield is in place, Pa

pu Complex acoustic pressure in the unshielded configuration, Pa

p ̃s Measured RMS acoustic pressure spectrum with acoustic shield installed, Pa

p ̃u Measured RMS acoustic pressure spectrum in the unshielded configuration, Pa

SO Propulsor Shaft Order, 2f / (Nf + Na)

∆SPL Decibel reduction in sound pressure level due to the acoustic shield

1 Introduction

From its inception, the growth in the popularity of jet-powered commercial transportation has been

followed by the demand to make such aircraft quieter. A wide variety of aircraft propulsion noise reduction technologies have been explored in the last forty years. Most of the success enjoyed in this

endeavor has been due to the adoption of high bypass ratio engines with increasingly large diameter fans. This has reduced both fan noise and jet noise while increasing propulsive efficiency. Further increases in bypass ratio are expected, but a number of technological challenges are making this

difficult. These challenges range from the weight of a large fan cowl, to the installation challenges

of under-wing mounting, to flutter of large radius fan blades. Additional noise reduction technologies that have seen widespread adoption are acoustic liners for engine nacelles and chevrons to reduce jet exhaust noise. Acoustic shielding is a fundamentally different noise mitigation technique than

those mentioned above, in that the noise generated by the engines is not actually reduced, but rather

blocked from reaching the observer. In order to meet the aggressive noise reduction goals established

by NASA (http://www.aeronautics.nasa.gov/nra_pdf/sfw_proposal_c1.pdf), acoustic shielding of

engine noise by the airframe is essential. For example, the difference between the NASA N+1 and

N+2 noise goals is assumed to come primarily from switching from conventional to advanced airframes

that incorporate shielding. There are many technical challenges to implementing noise shielding by the

airframe, including the need for designing an aircraft from ground up to utilize shielding.

NASA/TM—2012-217227 1
The objective of the present paper is to describe the results from a recent acoustic shielding test that

utilized a model scale counter-rotating open rotor as the sound source. The engine simulator provides

a realistic representation of what a modern open rotor may sound like, while the shielding geometry in

the test was chosen to be simple enough to provide a reference database for noise attenuation theories

and code development. Additionally, we have included two simple shielding models to serve as reference

prediction capabilities.

As with many topics of applied engineering research, interest in acoustic shielding has waxed and

waned over in the last few decades. Early work was mostly analytical and experimental, while recently

computational efforts have been brought to bear on the subject. Acoustic shielding predictions for

realistic problems are challenging for a number of reasons such as complex airframe geometries, existence

of a wide range of length and frequency scales, and complicated source distributions. This combination

of issues has resulted in a topic that has a lot of room for development. Below is but a sampling of the

body of work dedicated to airframe shielding studies in the last several decades.

In 1975, Conticelli et al.[1] applied a formulation for spherical waves intersecting a semi-infinite plate.

They performed a wide range of shielding experiments for both jet and fan noise sources. The location,

distribution, and directionality of the sources were found to have a considerable affect on the amount

of shielding observed in the far field. For example, they found that while aft radiating fan noise was

attenuated considerably, jet noise was not as well attenuated.

Around the same time, a series of studies were carried at NASA Lewis Research Center (now NASA

Glenn), concerned with reducing aft engine noise by mounting the engines over a conventional wing, as

reported on by von Glahn et al.[2], Reshotko et al.[3] and Bloomer[4]. The culmination of the engineover-the-wing research performed by NASA was the Quiet Short-Haul Research Aircraft[5].

Examples of more recent, computationally based work, include Agarwal et al.[6], who proposed a

ray-tracing approach based on geometrical optics solutions including diffracted rays. Their work was

compared to a validation experiment performed in an anechoic chamber, using a compression driver

coupled to a pipe and a two-dimensional profile of the Silent Aircraft geometry. A method, suggested

by Ng and Spakovszky[7], is based on the Kirchhoff diffraction theory. Their implementation considers

only the outline of an object and assume only edge diffraction rays. This results in a “medium fidelity”

computation that may be a good compromise in predictive quality. An experimental investigation by

Czech et al.[8] has been conducted at the Boeing Low Speed Aeroacoustic Facility (LSAF). A planform

geometry based on Boeing’s Blended Wing Body (BWB) concept has been used along with Jet Noise

Rig Simulator to experimentally study shielding effectiveness. The results document up to 20 dB of

noise reduction at some directivity angles. A recent paper by Turkdogru et al.[9] has been specifically

concerned with the validity of shielding research performed with simplified sources. The current NASA

computational baseline capability is the Fast Scattering Code (FSC) [10] which is a computer-intensive

tool.

The present paper describes a recent NASA wind tunnel test that incorporates a model scale open

rotor, along with a simplified shielding geometry. The results in this paper are intended to serve as benchmarks for code validation as well as a data set for generating system-level airframe shielding estimates.

The next section describes the open rotor shielding experiment and the operating conditions that were

run. A section on the experimental results is presented and is followed by a section with predictions from

two basic shielding methods that can serve as a benchmark for the more sophisticated codes required to

handle realistic, three-dimensional shielding geometries.

2 Experiment

A series of acoustic shielding tests was conducted as part of the ongoing wind tunnel testing of a scale

model counter-rotating open rotor engine simulator. The open rotor simulator provides scaled aerodynamic performance, structurally responsive blades, and realistic acoustic signatures. In other words, the

aeroacoustic, aeroelastic, and performance data from this test rig can be scaled to flight sized engines.

Comparisons of model scale acoustic data to the full-scale counter rotating engine flight test in the 1980’s

were found to be in good agreement.[11]

2.1 Model Description

2.1.1 Drive Rig

The NASA counter rotating Open Rotor Propulsion Rig (ORPR) was used to power the open rotor

blades in this shielding experiment. The rig consists of two counter rotating spools with a non-rotating

NASA/TM—2012-217227 2
center shaft. Each of the two counter rotating spools is attached to a two-stage air turbine on the aft

end that can produce up to 560 kW (750 HP). The two air turbines are supplied by 2 MPa (300 psi) air

heated to 90◦C (200◦F), and can be independently controlled using a feed-back system to drive the rotor

speeds independently up to 8,000 RPM with an accuracy of ± 15 RPM. This drive rig was last operated

in the late 1980’s, and was refurbished in order to correct any detrimental effects of long-term storage.

The refurbishment activity consisted of a general inspection and rebuild of the mechanical components

and construction of a new swept pylon. All instrumentation was inspected and replaced as necessary,

including refurbishment of the forward and aft rotating force balances and installation of a telemetry

system. The refurbished ORPR installed in the 9x15 Wind Tunnel (described in Section 2.2) is shown

in Figure 1, and the microphone instrumentation (described in Section 2.4) is visible in the foreground.

Figure 1: Open Rotor Propulsion Rig in 9x15 Wind Tunnel at NASA Glenn Research Center. NASA Image

C-2010-3454.

2.1.2 Historical Baseline Blades (F31 / A31)

The model test hardware simulates a counter rotating open rotor unducted fan engine design. New hubs

were manufactured as part of the current test campaign, with support for 12 blades for the forward rotor

and a 10 blades for the aft rotor. The pitch angle setting for both blade rows are manually adjustable.

The blade set used for the data in the present paper was designed and manufactured by General Electric

and is known as F31/A31. This blade set represents one of the better early 1990’s aerodynamic designs,

but is not acoustically optimized. The F31/A31 blade set was manufactured with modern materials and

techniques for this test, and will be used to provide a baseline to anchor new data. The forward propulsor

is 65.18 cm (25.66 inches) in diameter with 12 composite fan blades and the aft rotor is 62.97 cm (24.79

inches) in diameter with 10 blades. Axial spacing between the two rotors is variable, but was set at 19.91

cm (7.84 inches) for this test. The center of the aft rotor at the blade pitch change axis location is the

origin of the coordinate system for all results shown in this paper.

2.2 9x15 Wind Tunnel

The 9x15 Wind Tunnel at Glenn Research Center (GRC) is a low speed wind tunnel with acoustic

treatment in the test section. The test section is 2.74 m high by 4.67 m wide (9 feet by 15 feet) and

8.72 m (28.6 feet) long. The test section walls have four 10.2 cm (4 inch) wide slots that run the length

NASA/TM—2012-217227 3
of the test section, which are designed to reduce tunnel wall effects. The flow is driven by three electric

motors that can produce up to 65 MW (87,000 HP) and drive the tunnel velocity up to Mach 0.22. A

flow conditioner provides a very low turbulence level in the center part of the test section. A description

of the 9x15 Wind Tunnel complex is given by Soeder[12]. Background noise levels were reported on

by Woodward and Dittmar[13] and the acoustic qualities of the facility were documented by Dahl and

Woodward[14].

2.3 Geometry of Shielding Test

The geometry of the shielding test was a compromise between several objectives. First, the rotor sound

source should be unaffected by the presence of the wall. Second, the test should be useful for validation

of prediction methods. Third, the results should be useful for estimating the total noise reduction in

a system level analysis. The present shielding experiment utilized a simplified barrier geometry that

approximates a wing or horizontal stabilizer on a conventional “tube and wing” airframe. A paper by

Berton[15] utilizes the measurements from the present test to perform a system level noise study of

shielding benefits of a suitable tail on a full sized aircraft.

It should be noted that more complex installations, such as those involving open rotors installed on

top of a lifting body were not part of the objective of the present study. Such geometries would introduce

additional complexities that would complicate a direct assessment of basic shielding effects. For example,

streamlines ingested into the rotor would be distorted by the lifting body, which would change the rotor

loading and hence the sound source itself. Additionally, turbulence from the fuselage surface or boundary

layer may be ingested into the rotor, which would further complicate the sound producing mechanisms.

A consistent rotor sound source was desired, so the barrier wall was located at positions where the flow

into the rotor would not be disrupted.

The barrier wall was positioned relative to the open rotor in four different configurations, shown as

a sketch in Figure 2. All four were along a 66 cm (26 inch) sideline distance from the centerline of the

model. The barrier wall was 7 cm (2.75 inches) thick with a 2:1 ellipse on the leading edge and a circular

trailing edge. The wall was positioned in two streamwise locations, and both “long” and “short” wall

lengths were used. Note that the two “long” configurations were slightly different in length. The forward

wall position places the leading edge of the barrier wall slightly upstream of the model forebody. The aft

wall position places the nose of the barrier wall even with the center of the aft rotor. A schematic of the

two wall location is shown in Figure 2. A table with the relevant geometry details is given in Table 1.

Figure 2: Schematic of Shielding Test.

2.4 Acoustic Instrumentation

Acoustic instrumentation was a single Br ̈uel and Kjær 4939 6.35 mm (1/4 inch) microphone with the

standard bullet nose for making in-flow microphone measurements. This microphone was mounted on

NASA/TM—2012-217227 4
Table 1: Shielding Configurations

Wall Length Wall Position Location of Wall Leading Edge

Long 4.42 m (174 inches) Forward 102.4 cm (40.3 inches) Upstream of

Center of Aft Rotor

Long 3.20 m (126 inches) Aft Even with Center of Aft Rotor

Short 72.6 cm (28.6 inches) Forward 102.4 cm (40.3 inches) Upstream of

Center of Aft Rotor

Short 72.6 cm (28.6 inches) Aft Even with Center of Aft Rotor

a linear traverse offset on a 1.52 m (5 foot) sideline from the centerline of the model. The sideline

measurement covers the observation angles between 17.6 and 140 degrees from the upstream axis of the

fan, at 18 angles. An RC Electronics DataMAX II data recorder was used to record 15-second samples

at 200 kHz, along with the once-per-revolution signals from the front and aft rotors. The recorded

time series of pressures was processed into sound pressure spectra with 12.2 Hz frequency bins, using

the Digital Acoustics Data System (DADS) codes developed at GRC. In addition, this software applied

corrections for the microphone and bullet-nose sensitivity and directivity, and corrected for atmospheric

attenuation, resulting in 30.5 cm (1 foot) lossless pressure spectral densities. Background acoustic levels

were documented as part of the model checkout process. The wind tunnel was run with the test rig

installed, but without rotor blades, i.e., a tare run. This gives a measure of the background noise level in

the wind tunnel including the scrubbing noise caused by airflow over the drive rig. The tare configuration

noise was found to be at least 10 dB lower than the fan data presented in this paper. A muffler was

installed on the drive rig exhaust where the airflow through the drive turbines is expelled, and was shown

to mitigate noise from the drive rig air turbines. The muffler was used for all measurements presented

in this paper. Frequencies below 500 Hz have been omitted from results shown in this paper, due to

background noise in the facility and limitations of the acoustic treatment below 350 Hz.

2.5 Operating Conditions

The variable pitch blades were set to nominal take-off pitch angles for this test. The test matrix included

two barrier wall locations, two wall lengths, two tunnel Mach numbers, and three rotor RPMs including

the rotor at “windmill” condition. The two rotors were driven at equal rotation rates during the powered

runs described in this paper. Test parameters are given in Table 2. The complete open rotor diagnostics

test matrix includes many more conditions, but the listed subset was deemed adequate for testing acoustic

shielding. The operating condition could affect the sound source distribution on the rotor. The diameter

of the rotor is roughly the same as the distance between the rotor and the acoustic barrier, and therefore

also large with respect to the length of the short barrier. A significant change in the sound source

distribution may affect the acoustic shielding, but it is likely a secondary effect compared with the

geometry of the system.

The barrier wall measurements with the long wall installed (RDGs 504-529) were compared with

the isolated measurements taken the prior day (RDGs 466-496). The short wall measurements (RDGs

1056-1075) were compared against earlier readings (RDGs 4235 and 4241). This is because the blade

pitch angles were set slightly improperly for the RDG values 466-529, causing a small shift in the acoustic

spectra. Nonetheless, the comparison between the two measurements of attenuation should still be valid.

3 Results and Discussion

This section presents selected acoustic measurements from the open rotor with and without shielding.

The first section will describe the isolated open rotor, and briefly discuss the character of the radiated

sound. The second section will consider the effect of shielding on the sound spectra. An algorithm for

quantifying the tone and broadband levels will be presented, and finally the measured attenuation will

be reported.

NASA/TM—2012-217227 5
Table 2: Operating Conditions

Percent Corrected Total Shaft

ESCORT Wall Wall Tunnel Rotor Design Total Thrust Power

RDG Length Location Mach RPM Speed N (lbf) kW (HP)

466 - - 0.20 5710 85.2% 1681 (378) 179 (240)

470 - - 0.20 6486 96.7% 2535 (570) 303 (406)

494 - - 0.22 5969 88.7% 1668 (375) 185 (248)

496 - - 0.22 6732 100% 2518 (566) 311 (417)

504 Long FWD 0.20 5685 85.2% 1641 (369) 176 (236)

506 Long FWD 0.20 6461 96.7% 2495 (561) 298 (399)

510 Long FWD 0.22 5945 88.7% 1624 (365) 181 (243)

512 Long FWD 0.22 6708 100% 2473 (556) 306 (410)

521 Long AFT 0.20 5692 85.2% 1655 (372) 177 (237)

523 Long AFT 0.20 6468 96.7% 2522 (567) 299 (401)

527 Long AFT 0.22 5949 88.7% 1637 (368) 181 (243)

529 Long AFT 0.22 6712 100% 2487 (559) 306 (410)

1056 Short FWD 0.20 5678 85.2% 1677 (377) 188 (252)

1058 Short FWD 0.20 6457 96.7% 2540 (571) 315 (423)

1073 Short AFT 0.20 5699 85.2% 1708 (384) 189 (254)

1075 Short AFT 0.20 6476 96.7% 2513 (565) 316 (424)

4235 - - 0.20 5675 85.2% 1788 (402) 189 (253)

4241 - - 0.20 6456 96.7% 2682 (603) 318 (427)

3.1 Isolated Sound Measurements

The counter-rotating propellers produce substantial tonal sound with many harmonics. Most of these

tones are related to the interaction of the wakes from the forward rotor with the aft rotor blades. Numerous studies have already commented on the sound radiated from counter-rotating propellers(

[16]−[19]),

but a small sample of results will be presented here.

The present open rotor featured blade counts of 12 on the forward rotor and 10 on the aft rotor. This

mismatch results in a large number of “interaction tones.” The blade rate tones occur at multiples of 12

and 10 shaft orders, while interaction tones can be approximated as

Shaft Order Interaction Tones = mBfNf + nBaNa

.5 (Nf + Na)

, (1)

when m 6= 0 or n 6= 0 and when Nf and Na are nearly equal, and shaft order is defined as SO =

2f /(Nf + Na). For all of the data presented in this paper, the two rotors were set to the same rotation

speed, Nf = Na. In this case, the first few of these tones are shown in Table 3.

Table 3: Table of shaft orders for counter rotation propeller tones.

n = 0 n = 1 n = 2 n = 3 n = 4 n = 5 n = 6

m = 0 10 20 30 40 50 60

m = 1 12 22 32 42 52 62 72

m = 2 24 34 44 54 64 74 84

m = 3 36 46 56 66 76 86 96

m = 4 48 58 68 78 88 98 108

m = 5 60 70 80 90 100 110 120

A example directivity is shown in Figure 3. The highest tone levels are at shaft orders 42, 32, 34

and 10, respectively. Two very similar frequencies (3460 and 3514 Hz) are compared corresponding

to SO = 34.0 and fractional shaft order 34.5 respectively. The wavelengths of the two frequencies are

similar, but their radiated sound directivity is very different. The tone at shaft order 34 exhibits a

complicated directivity, while the sound at fractional shaft order 34.5 is effectively broadband, and has

a relatively weak dependence on radiation angle. This plot is qualitatively typical of the directivity

of specific frequencies. Tones were found to have dense and complicated radiation patterns that were

under-resolved by our microphone measurements while broadband noise was a much weaker function of

radiation angle.

NASA/TM—2012-217227 6
20 40 60 80 100 120 140

70

75

80

85

90

95

100

105

110

115

SPL, dB

Directivity Angle, Degrees

Unshielded, SO = 34.0

Unshielded, SO = 34.5

Long Shield Aft, SO = 34.0

Long Shield Aft, SO = 34.5

(a) Long barrier wall in aft position

20 40 60 80 100 120 140

70

75

80

85

90

95

100

105

110

115

SPL, dB

Directivity Angle, Degrees

Unshielded, SO = 34.0

Unshielded, SO = 34.5

Short Shield Aft, SO = 34.0

Short Shield Aft, SO = 34.5

(b) Short wall in aft position

Figure 3: Example directivity of a tone and broadband at nearly the same frequency of ≈ 3500 Hz. 96.7%

design speed, M=0.20.

3.2 Effect of Shielding

The rotor acoustics measured in the shadow region of the barrier wall is significantly different from that

for the isolated (i.e., unshielded) rotor. The simplest configuration to consider may be that of the long

wall in the aft position. In this case, the upstream edge of the barrier is at the same streamwise location

as the aft rotor and the barrier is long enough to be considered semi-infinite in the downstream direction.

Measurements from these cases are shown in Figure 4, with a schematic as Figure 4 (a). These figures are

ordered from forward (no direct shielding) to aft (most shielded region). At 45◦from upstream, shown

in Figure 4 (b), the radiated sound is dominated by shaft order tones 22 and 32, followed by 42 and 44

and dozens more tones. The barrier appears to have a small effect on some tones. The subfigures (c), (d)

and (e) show the emission angles around 90◦, which is the angle tangent to the nose of the barrier wall.

When the shield is in place, nearly all the tones are reduced, typically more than 10 dB, which is an

order of magnitude or more. This clearly demonstrates the benefits of a large shield blocking the source

from the observer. As the observer location moves from behind the barrier wall, the spectra quickly fill

with tones and the spectral levels approach the isolated measurement.

The effect of a short wall is shown in Figure 5 for contrast. The short wall has the additional

complications of diffraction around both edges of the wall, as well as scattering through the shear layer

in the wake of the barrier. A more extensive measurement suite would be required to confirm the effect

of the shear layer. Considerable noise reduction was found directly behind the barrier, shown in Figure

5 (e).

NASA/TM—2012-217227 7
(a) Schematic

0 20 40 60 80 100

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Shaft Orders

SPL, dB

(b) 45◦

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SPL, dB

(c) 67.5◦

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(d) 90◦

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SPL, dB

(e) 112.5◦

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Shaft Orders

SPL, dB

Unshielded

Shielded

(f) 135◦

Figure 4: Comparison of measured sound with long wall aft vs isolated rotor. 85.2% design speed,, M=0.20.

NASA/TM—2012-217227 8
(a) Schematic

0 20 40 60 80 100

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SPL, dB

(b) 45◦

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(c) 67.5◦

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(d) 90◦

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(e) 112.5◦

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SPL, dB

Unshielded

Shielded

(f) 135◦

Figure 5: Comparison of measured sound with short wall aft vs isolated rotor. 85.2% design speed, M=0.20.

NASA/TM—2012-217227 9
3.3 Tone vs Broadband

A peak-finding algorithm was used to identify the tones in the open rotor sound spectra, defined as levels

above broadband of 1 dB or greater. Each such peak was “chopped off” and replaced with a level that

was an interpolated value between the neighboring points. In that way, a reasonably smooth spectrum

was created that was devoid of the forest of tones typically observed in open rotor spectra. This curve

may be approximately described as the broadband level of the spectrum. This “broadband” spectrum

was subtracted (linearly) from the total spectrum to give a spectrum of tones.

Figure 6 shows all three spectra (tone, broadband and total), with tones less than 1 dB above the

broadband hidden. The broadband curves returned by the algorithm were all found to be well-behaved,

and relatively weak functions of frequency.

0 20 40 60 80 100

65

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SPL, dB

(a) 45◦

0 20 40 60 80 100

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SPL, dB

Measured SPL

Broadband

Tone

(b) 135◦

Figure 6: Tone and broadband separation results for long wall aft, 85.2% design speed, M=0.20.

3.3.1 Net Attenuation

The net attenuation at each angle is defined as the difference between the shielded and unshielded levels

after their respective spectrum is integrated over the frequency range from 1 kHz to 50 kHz. The tone and

broadband spectra as defined in Section 3.3 were integrated at each directivity angle to get an estimate

for the tonal and broadband over-all sound pressure level (OASPL). Note that this definition results in

OASPL values for tones and broadband that add up to the total OASPL. Simply put:

Attenuation (dB) = Unhielded OASPL (dB) − Shielded OASPL (dB). (2)

Results from attenuation calculations are shown in Figure 7. For nearly all combinations of wall configuration, operating condition and directivity angles, the tones were found to be attenuated more than

the broadband. This is most dramatic when comparing Figure 7 (c) and Figure 7 (d), which feature the

exact same geometry, but different operating conditions. The increase in rotor RPM raises the tone level

more than the broadband. Consequently, more tones remain above the broadband level at 96.7% design

speed, than do at 85.2% design speed, when the barrier is introduced.

4 Shielding Predictions

For the sake of comparison with these results, predications from two simple acoustic shielding theories

were calculated for the shielding geometries investigated here. Neither theory incorporates convective

effects of the wind tunnel flow.

4.1 Barrier Theory

Standard barrier theory is the type of prediction method usually used for roadside acoustic barriers.

The NASA Aircraft Noise Prediction Program (ANOPP) uses this type of method to estimate shielding

NASA/TM—2012-217227 10
0 20 40 60 80 100 120 140

0

5

10

15

20

Angles, Degrees

∆OASPL Attenuation, dB

Total

Tones

Broadband

(a) Long wall in aft position, 85.2% design speed, M=0.20

0 20 40 60 80 100 120 140

0

5

10

15

20

Angles, Degrees

∆OASPL Attenuation, dB

Total

Tones

Broadband

(b) Long wall in aft position 96.7% design speed, M=0.20

0 20 40 60 80 100 120 140

0

5

10

15

20

Angles, Degrees

∆OASPL Attenuation, dB

Total

Tones

Broadband

(c) Short wall in aft position, 85.2% design speed, M=0.20

0 20 40 60 80 100 120 140

0

5

10

15

20

Angles, Degrees

∆OASPL Attenuation, dB

Total

Tones

Broadband

(d) Short wall in aft position, 96.7% design speed, M=0.20

Figure 7: Integrated OASPL results.

effects. Li and Wong[20] provide a review paper that summarizes a number of empirical expressions for

barrier theory predictions.

The geometry used for the prediction is shown in Figure 8. The Fresnel number is given as,

N =

k

π

(rs + rr − R1), (3)

where rs is the distance from the sound source to the edge of the semi-infinite wall, rr is the distance

between the edge of the wall and the receiver, R1 is the direct distance (in the absence of the barrier)

and k is the acoustic wavenumber. It should be noted that, for the purposes of these calculations, the

open rotor was taken to be a point source located at the center of the aft rotor.

The Fresnel number was used to calculate the acoustic insertion loss with the empirical expression,

∆SPL = 10 log10 (3 + 20N), (4)

where the definition of attenuation is the ratio of unshielded sound pressure pu to shielded sound pressure

ps,

∆SPL = 10 log10 

pu

ps



. (5)

Li and Wong provide a large number of other empirical expressions that may be slightly more accurate

than Eq (4), but all essentially give the same results for the geometry and frequency range presented

here. These experimental fits were made to data with Fresnel numbers up to N = 10, while for the

present geometry frequencies of 10 kHz result in some values of Fresnel number over 100, which is a

NASA/TM—2012-217227 11
significant extrapolation from the curve fit. These high frequencies are not normally associated with

roadside barriers. To calculate the attenuation for the short barrier wall, two semi-infinite calculations

were carried out, one to account for each end of the barrier. The results were added as

∆SPL = 20 log10 

pu

ps1

+

pu

ps1



, (6)

which assumes the sound passing around each side of the barrier is coherent.

rs

rr

R1

Source

Receiver

Wall

Figure 8: Geometry for barrier wall theory.

The predicted acoustic shielding for the barrier geometry used in the wind tunnel test is shown in

Figure 9, where a point at 1 kHz was placed at the location of the aft rotor pitch change axis. The result

is essentially a shadow region. For the short wall, the greatest attenuation is immediately behind the

barrier.

Source

Wall

Spanwise Direction, m

Streamwise Direction, m

−2 −1 0 1 2 3 4

−2

−1

0

1

2

3

4

Attenuation, dB

−15

−10

−5

0

5

10

15

(a) Long Wall Aft.

Source

Wall

Spanwise Direction, m

Streamwise Direction, m

−2 −1 0 1 2 3 4

−2

−1

0

1

2

3

4

Attenuation, dB

−15

−10

−5

0

5

10

15

(b) Short Wall Aft.

Figure 9: Barrier theory predictions for acoustic insertion loss at 1 kHz.

4.2 Half-Plane Diffraction Method

The diffraction of a wave by a finite edge has been presented by Conticelli et al.[1] and Ahtye and

McCulley[21], among others. This method requires a Fresnel solution (the corresponding equations will

not be reproduced here, for the sake of brevity). The method calculates the complex pressure as a

sum of direct and diffracted waves, and offers the potential for solving more sophisticated problems

than the barrier theory method. For example, sound due to multiple point sources can be added,

possibly simulating one or more propeller disks. Finite edges can be simulated by summing the results

from multiple semi-infinite solutions. Good comparisons with carefully performed experiments were

demonstrated by Ahtye and McCulley.

NASA/TM—2012-217227 12
A simulation of geometry for the open rotor shielding experiment was performed. For simplicity, a

point source was again used for this result. An adaptation of the method to handle a distribution of

sources is currently being developed. Since the rotor diameter, barrier length (for the short wall) and

separation distance between the two are all roughly the same, a distribution of sources would be expected

to show less total attenuation. The axial separation of the two rotor planes is essentially 1/4 the short

barrier length, sound sources would be closer to both edges of the shield, so the attenuation would be

less. The distribution of sound sources opens additional complexities that are outside the scope of the

present paper, including magnitude, phase and modal content of the distributed source. It is clear that

strong tonal content is caused by the rotor-rotor interaction, which will likely have significant implications

for shielding of rotor tones. Existing experimental data and propeller noise theories may be tapped to

suggest what sort of source distribution would be appropriate to use.

The attenuation at 1 kHz in a two-dimensional plane is shown in Figure 10 for both long and short

barriers. In contrast to Figure 9 shown above, the constructive and destructive interference can be easily

seen for the short wall case. At 1 kHz, the wavelength is roughly 1/3 of the barrier length.

Streamwise Direction, m

Spanwise Direction, m

Wall

Source

−2 −1 0 1 2 3 4

−2

−1

0

1

2

3

4

Attenuation, dB

−15

−10

−5

0

5

10

15

(a) Long Wall Aft.

Streamwise Direction, m

Spanwise Direction, m

Wall

Source

−2 −1 0 1 2 3 4

−2

−1

0

1

2

3

4

Attenuation, dB

−15

−10

−5

0

5

10

15

(b) Short Wall Aft.

Figure 10: Diffraction theory predictions for acoustic insertion loss at 1 kHz.

4.3 Comparison with Experimental Results

The theory presented in this section is for frequency-dependent attenuation along the sideline measured

in the wind tunnel. To that end, Figure 11 shows the total attenuation measured by the sideline microphone, as well as the attenuation predicted by barrier theory and diffraction theory. The microphone

measurements presented here are in 100 Hz frequency bins, rather than the 12.2 Hz bins presented previously. The measurements are made at only 18 sideline locations, so for the sake of simple comparisons,

the loss of measurement fidelity is balanced by improved visual comparisons. With the limited spatial

resolution in the experiment, however, there is no possibility of measuring the diffraction detail shown

in Figure 11 (f).

The results for both predictions and the experiment were averaged over frequency to estimate the

total attenuation. This result is shown in Figure 12. For the long wall, both theories agree quite closely,

but over predict the attenuation by more than 5 dB at 130 degrees. For the short wall, both theories

again over predict the attenuation, with the diffraction theory result off by over 10 dB at some angles.

The diffraction theory predicts nearly the same result for both walls up to 112.5 degrees, as shown in

Table 4. Deficiencies include the point-source approximation and the absence of convection effects.

NASA/TM—2012-217227 13
Measurement Angle, Degrees

Frequency, Hz

20 40 60 80 100 120

0

2000

4000

6000

8000

10000

 Attenuation (dB)

−15

−10

−5

0

5

10

15

(a) Long wall in aft position.

Measurement Angle, Degrees

Frequency, Hz

20 40 60 80 100 120

0

2000

4000

6000

8000

10000

 Attenuation (dB)

−15

−10

−5

0

5

10

15

(b) Short wall in aft position.

Measurement Angle, Degrees

Frequency, Hz

20 40 60 80 100 120

0

2000

4000

6000

8000

10000

 Attenuation (dB)

−15

−10

−5

0

5

10

15

(c) Long wall in aft position.

Measurement Angle, Degrees

Frequency, Hz

20 40 60 80 100 120

0

2000

4000

6000

8000

10000

 Attenuation (dB)

−15

−10

−5

0

5

10

15

(d) Short wall in aft position.

Measurement Angle, Degrees

Frequency, Hz

20 40 60 80 100 120

0

2000

4000

6000

8000

10000

 Attenuation (dB)

−15

−10

−5

0

5

10

15

(e) Long wall in aft position.

Measurement Angle, Degrees

Frequency, Hz

20 40 60 80 100 120

0

2000

4000

6000

8000

10000

 Attenuation (dB)

−15

−10

−5

0

5

10

15

(f) Short wall in aft position.

Figure 11: Frequency dependent barrier wall shielding. (a) & (b) Measured attenuation, 96.7% design

speed, M=0.20. (c) & (d) Barrier theory predictions for acoustic insertion loss. (e) & (f) Diffraction theory

predictions for insertion loss.

NASA/TM—2012-217227 14
0 20 40 60 80 100 120 140

−5

0

5

10

15

20

25

Angles, Degrees

∆OASPL Attenuation, dB

Diffraction Theory

Barrier Theory

Experiment

0 20 40 60 80 100 120 140

−5

0

5

10

15

20

25

Angles, Degrees

∆OASPL Attenuation, dB

Figure 12: Total frequency-averaged attenuation. a) Long wall in aft position, b) short wall in aft position.

Table 4: Net attenuation at 112.5 degrees.

Method Long Wall Aft Short Wall Aft

Experiment 12.6 dB 8.7 dB

Barrier Theory 16.6 dB 11.9 dB

Diffraction Theory 14.9 dB 14.1 dB

5 Conclusions

An experiment for investigating the fundamentals of acoustic shielding of an open rotor has been reported.

The results serve both as a benchmark for evaluating shielding prediction methods as well as providing

first order estimates of how much shielding may be expected from wing or tail shielding. The size of

the open rotor simulator, the size of the acoustic barrier used in the test, and the distance between the

open rotor simulator and the barrier are roughly the same making this an especially challenging case to

predict.

Two simple shielding methods were used to provide reference attenuation predictions. The datatheory comparisons indicate that these simple methods do not capture the nuances of shielding behavior

and tend to over predict the shielding benefits at high frequencies. It is expected that more advanced

prediction methods would result in more accurate predictions. The demonstrated shielding benefits from

a simple barrier (such as a wing or tail), together with results from more advanced tests performed elsewhere, strongly suggest the significant potential of shielding of propulsion noise by airframe components,

which should keep the interest in this technology high.

References

[1] Conticelli, V. M., Blasi, A. D., and O’Keefe, J. V., “Noise Shielding Effects for Engine-Over-Wing

Installations. AIAA 75-474,” 2nd AIAA Aeroacoustics Conference, Hampton, Virginia, USA, March

24-26, 1975.

[2] von Glahn, U., Groesbeck, D., and Reshotko, M., “Geometry Considerations for Jet Noise Shielding

with CTOL Engine-Over-the-Wing Concept. AIAA 74-568,” 7th AIAA Fluid and Plasma Dynamics

Conference, Palo Alto, California, USA, June 17-19, 1974.

[3] Reshotko, M., Goodykoontz, J. H., and Dorsch, R. G., “Engine-Over-the-Wing Noise Research,”

Journal of Aircraft, Vol. 11, No. 4, 1974, pp. 195–196.

[4] Bloomer, H. E., “Investigation of Wing Shielding Effects on CTOL Engine Noise,” 5th AIAA Aeroacoustics Conference, Seattle, Washington, USA, March 12-14, 1979.

NASA/TM—2012-217227 15
[5] Shovlin, M. D. and Cochrane, J. A., “An Overview of the Quiet Short-Haul Research Aircraft

Program,” NASA/TM-78545 , November 1978.

[6] Agarwal, A., Dowling, A. P., Shin, H.-C., Graham, W., and Sefi, S., “A Ray Tracing Approach to

Calculate Acoustic Shielding by the Silent Aircraft Airframe. AIAA 2006-2618,” 12th AIAA/CEAS

Aeroacoustics Conference, Cambridge, Massachusetts, USA, 8-10 May 2006.

[7] Ng., L. W. T. and Spakovszky, Z. S., “Turbomachinery Noise Shielding Assessment of Advanced

Aircraft Configurations. AIAA 2010-3914,” 16th AIAA/CEAS Aeroacoustics Conference, Stockholm,

Sweden, 7-9 June 2010.

[8] Czech, M. J., Thomas, R. H., and Elkoby, R., “Propulsion Airframe Aeroacoustic Integration Effects

for a Hybrid Wing Body Aircraft Configuration. AIAA 2010-3912,” 16th AIAA/CEAS Aeroacoustics

Conference, Stockholm, Sweden, 7-9 June 2010.

[9] Turkdogru, N., Ahuja, K. K., and Gaeta, R. J., “Validity of the Point Source Assumption in Rotor

Noise Measurements with Shielding. AIAA 2010-3853,” 16th AIAA/CEAS Aeroacoustics Conference, Stockholm, Sweden, 7-9 June 2010.

[10] Tinetti, A. F., Dunn, M. H., and Pope, D. S., “Fast Scattering Code (FSC) Users Manual, Version

2.0,” NASA/CR-2006-214510 , October 2006.

[11] Woodward, R. P., Loeffler, I. J., and Dittmar, J. H., “Measured Far-Field Flight Noise of a Counterrotation Turboprop at Cruise Conditions,” NASA/TM-101383 , January 1989.

[12] Soeder, R. H., “NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel User Maunal,” NASA/TM

106247 , August 1993.

[13] Woodward, R. P. and Dittmar, J. H., “Background Noise Levels Measured in the NASA Lewis 9-

by 15-Foot Low-Speed Wind Tunnel. AIAA 95-0720,” 33rd AIAA Aerospace Sciences Meeting and

Exhibit, Reno, Nevada, USA, January 9-12, 1995.

[14] Dahl, M. D. and Woodward, R. P., “Comparison Between Design and Installed Acoustic Characteristics of NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel Acoustic Treatment,” NASA/TP-2996 ,

April 1990.

[15] Berton, J. J., “Empennage Noise Shielding Benefits for an Open Rotor Transport,” 17th

AIAA/CEAS Aeroacoustics Conference, Portland, Oregon, USA, 5-8 June 2011.

[16] Janardan, B. A. and Gliebe, P. R., “Acoustic Characteristics of Counterrotating Fans from Model

Scale Tests. AIAA 1989-1142,” 12th AIAA Aeroacoustics Conference, San Antonio, Texas, USA,

April 10-12, 1989.

[17] Woodward, R. P., “Noise of Two High-Speed Model Counter-Rotation Propellers at Takeoff/Approach Conditions,” Journal of Aircraft, Vol. 29, No. 4, 1992, pp. 679–685.

[18] Magliozzi, B., Brown, P., and Parzych, D., “Acoustic Test and Analysis of a Counterrotating PropFan Model,” NASA/CR-179590 , October 1987.

[19] Elliot, D. M., “Initial Investigation of the Acoustics of a Counter Rotating Open Rotor Model

with Historical Baseline Blades in a Low Speed Wind Tunnel,” 17th AIAA/CEAS Aeroacoustics

Conference, Portland, Oregon, USA, 5-8 June 2011.

[20] Li, K. M. and Wong, H. Y., “A review of commonly used analytical and empirical formulae for

predicting sound diffracted by a thin screen,” Applied Acoustics, Vol. 66, 2005, pp. 45–76.

[21] Ahtye, W. F. and McCulley, G., “Evaluation of Approximate Methods for the Prediction of Noise

Shielding by Airframe Components,” NASA TP-1004 , January 1980.

NASA/TM—2012-217227 16
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Acoustic Shielding for a Model Scale Counter-rotation Open Rotor 

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14. ABSTRACT

The noise shielding benefit of installing an open rotor above a simplified wing or tail is explored experimentally. The test results provide 

both a benchmark data set for validating shielding prediction tools and an opportunity for a system level evaluation of the noise reduction 

potential of propulsion noise shielding by an airframe component. A short barrier near the open rotor was found to provide up to 8.5 dB of 

attenuation at some directivity angles, with tonal sound particularly well shielded. Predictions from two simple shielding theories were 

found to overestimate the shielding benefit. 

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Acoustics shielding; Open rotor 

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