Jeffrey J. Berton

Glenn Research Center, Cleveland, Ohio

Empennage Noise Shielding Benefits 

for an Open Rotor Transport

NASA/TM—2012-217218

January 2012

AIAA–2011–2764
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Jeffrey J. Berton

Glenn Research Center, Cleveland, Ohio

Empennage Noise Shielding Benefits 

for an Open Rotor Transport

NASA/TM—2012-217218

January 2012

AIAA–2011–2764

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

17th Aeroacoustics Conference

cosponsored by the American Institute of Aeronautics and Astronautics and 

the Confederation of European Aerospace Societies

Portland, Oregon, June 5–8, 2011
Acknowledgments

Thanks go to David Stephens, David Elliott, and Dale Van Zante of NASA’s Environmentally Responsible Aviation Project for 

collecting and providing the open rotor acoustic data used in this study. Thanks also go to Edmane Envia and the Subsonic Fixed 

Wing Project for providing the opportunity and resources to study the data.

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NASA/TM—2012-217218 1

Empennage Noise Shielding Benefits for an Open Rotor Transport

Jeffrey J. Berton

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

NASA sets aggressive, strategic, civil aircraft performance and environmental goals and develops 

ambitious technology roadmaps to guide its research efforts. NASA has adopted a phased approach for 

community noise reduction of civil aircraft. While the goal of the near-term first phase focuses primarily 

on source noise reduction, the goal of the second phase relies heavily on presumed architecture changes

of future aircraft. The departure from conventional airplane configurations to designs that incorporate 

some type of propulsion noise shielding is anticipated to provide an additional 10 cumulative EPNdB of 

noise reduction. One candidate propulsion system for these advanced aircraft is the open rotor engine. In 

some planned applications, twin open rotor propulsors are located on the aft fuselage, with the vehicle’s 

empennage shielding some of their acoustic signature from observers on the ground. This study focuses 

on predicting the noise certification benefits of a notional open rotor aircraft with tail structures shielding 

a portion of the rotor noise. The measured noise of an open rotor test article—collected with and without

an acoustic barrier wall—is the basis of the prediction. The results are used to help validate NASA’s 

reliance on acoustic shielding to achieve the second phase of its community noise reduction goals. The 

noise measurements are also compared to a popular empirical diffraction correlation often used at NASA 

to predict acoustic shielding.

Nomenclature

ai calibration constants introduced in Maekawa diffraction model

c speed of sound, ft/s

f frequency, Hz

IL insertion loss (barrier-shielding attenuation), dB

MFlight flight Mach number

MTunnel wind tunnel Mach number

NF Fresnel number, 2f∆/c

∆ characteristic length of Fresnel number

θE emission yaw angle

θG geometric yaw angle

1.0 Introduction

Rising, unstable fuel prices are spurring research into advanced, energy-efficient propulsion concepts 

for transport aircraft. As a result, rekindled attention is being given to open rotor propulsion systems. 

Once hailed as an innovative response to the sharp increases in aviation fuel cost beginning in 1973, 

interest in open rotors waned in the face of falling oil prices starting in 1986. Current energy concerns are 

reviving development efforts into open rotor propulsors (alternately known as advanced turboprops, 

propfans, or unducted fans).

Counter-rotating open rotor propulsion systems—with highly-swept, contoured, wide-chord rotor 

blades—combine the fuel efficiency of traditional turboprops with the high cruising airspeed of turbofan 

engines. Without inlet and bypass exhaust ducts, however, an unfortunate disadvantage of the open rotor 

relative to ducted turbofans may be higher community noise. Although promising low-noise rotor designs 
NASA/TM—2012-217218 2

Figure 1.—Proposed Airbus A30X and easyJet ecoJet open rotor airplane candidates having potential empennage 

shielding benefits (Refs. 6, 7).

exist, and although any eventual open rotor aircraft will necessarily have to meet all noise certification 

requirements, it is likely that open rotor propulsion systems will be louder than modern, high-bypass ratio 

turbofans of comparable technology. People living near airports have the expectation of continuing the 

historical downward trend of aircraft noise.

NASA sets aggressive, strategic civil aircraft performance and environmental goals and develops 

ambitious technology roadmaps to guide its research efforts. Under NASA’s Fundamental Aeronautics 

and Integrated Systems Research Programs, the Subsonic Fixed Wing and Environmentally Responsible 

Aviation Projects, respectively, have adopted community noise goals for the new, subsonic, single-aisle, 

civil aircraft expected to replace the current Boeing 737 and Airbus A320 families of airplanes. In NASA 

vernacular, these aircraft are designated as “N+1” vehicles since they will follow the current, in-service, 

“N” airplanes. NASA’s N+1 community noise reduction goal calls for certification noise levels of 32 

cumulative EPNdB under current FAA Part 36 Stage 4 (Ref. 1) and ICAO Chapter 4 (Ref. 2) noise limits.

A more challenging noise goal is set for the following generation of N+2 aircraft. This goal calls for a 

cumulative Stage 4 margin of 42 EPNdB. The additional 10 EPNdB margin is very ambitious, and it is 

thought by NASA not to be achievable by further reductions in engine and airframe noise sources alone. 

Instead, some sort of unconventional change in aircraft architecture, where propulsion noise sources are 

shielded from observers on the ground, is thought to be necessary to achieve the additional noise margin. 

The additional 10 EPNdB noise margin of the N+2 goal was set based on several preliminary studies of 

shielding benefits (e.g., Refs. 3, 4, and 5).

Noise shielding (also referred to as barrier attenuation or insertion loss) is an acoustic diffraction 

phenomenon where acoustic waves are attenuated when propagated past an impermeable barrier placed 

between the noise source and an observer. Shielding is particularly efficient when the observer is located 

in the “shadow region” where the noise source is obscured. An obvious candidate for effective propulsion 

noise shielding is the proposed Blended Wing Body (BWB) aircraft. When turbofan engines are located

above its large wing planform, a great deal of their fan and core noise would be shielded. Only jet noise—

a distributed source generated downstream throughout the axial exhaust plume—would be difficult to 

effectively shield. Open rotor propulsion has been proposed for the BWB. Having little distributed jet 

noise,* the relatively compact noise signature of an open rotor propulsor and its gas generator core may be 

more effectively shielded by a barrier than a turbofan.

Noise shielding is certainly not limited to BWB architectures, however. More conventional tube-andwing aircraft may also provide some amount of noise shielding to open rotor propulsion systems. For 

example, one proposal for the Airbus A30X concept (Ref. 6, Fig. 1, left) has twin open rotor engines 

mounted to the aft fuselage, with the horizontal tail extending below and aft of counter-rotating 

propellers. The vertical twin tails, attached to the tailplane tips, are also likely to offer additional shielding 

benefit to lateral observers. Another notional concept is easyJet’s ecoJet airplane (Ref. 7, Fig. 1, right). Its 

empennage is located below and forward of two open rotor propulsors. This “U-tail” empennage type of 

 * Most of the core stream enthalpy is used to drive the power turbine, resulting in a low-energy exhaust.
NASA/TM—2012-217218 3

noise shielding is, most likely, less effective than the shielding provided by the BWB because of the 

relatively smaller blocking area.

Testing of various open rotor designs is underway in NASA Glenn’s experimental facilities as part of 

an ongoing NASA-General Electric collaborative partnership (Ref. 8). Low-speed aerodynamic and 

acoustic tests were conducted in the Glenn 9- by 15-ft Low Speed Wind Tunnel, while (at this writing) 

high-speed aerodynamic testing is underway in the Glenn 8- by 6-ft tunnel. Noise data were collected 

during low-speed testing from an open rotor test article in a simulated flight environment with and 

without the presence of an acoustic barrier wall placed between the rotor and the microphone array. These 

shielded and unshielded acoustic data form the basis of this study.

The focus of this study is to predict the certification noise shielding system benefits that a U-tail

empennage is likely to provide aircraft with twin open rotor engines mounted in the aft fuselage position. 

The additional cumulative margin of 10 EPNdB for NASA’s N+2 noise goal was set largely via analytical 

means with relatively simple, first-order, low-fidelity tools. With the use of realistic, measured, acoustic 

shielding data, the results of this assessment will help in determining whether the additional 10 EPNdB of 

cumulative margin is reasonable.

The noise measurements are also compared to the simple, popular Maekawa empirical diffraction 

correlation often used at NASA to predict noise shielding (see Sections 2.4 and 3.3).

2.0 Method of Analysis

In October, 2010, the acoustic data used in this study were collected from a counter-rotating, open 

rotor scale model in NASA Glenn’s 9- by 15-ft Low Speed Wind Tunnel (Ref. 9). NASA’s open rotor 

test rig was configured in these tests with General Electric’s F31/A31 rotor blade set, with 12 forward 

rotor blades and 10 aft blades. The approximately one-fifth scale F31/A31 rotor set—often referred to as 

the “historical blade set”—dates to the Advanced Turboprop Project (Refs. 10 and 11) of the 1980s. It is 

an unducted fan representative of an early 1990s technology level. The acoustic data collected from it is 

non-proprietary and may be published in the open literature. Additional details of the rotor design may be 

found in Reference 12. Other, advanced-design, proprietary rotor sets incorporating innovative low-noise 

features and performance shaping of the airfoil stack derived through three-dimensional aerodynamic 

analysis have also been tested at NASA Glenn. Although the advanced rotors sets may be more 

representative of what may one day fly, the F31/A31 historical blade set serves as an approximate 

representation of a modern open rotor noise signature.

The tests at NASA Glenn were performed shortly before the joint NASA-Boeing propulsion-airframe 

aeroacoustic tests were completed at Boeing’s Low-Speed Aeroacoustic Facility in November 2010. The 

tests at Boeing used another historical open rotor test model: the F7/A7 (at approximately one-tenth

scale), placed above a simulated BWB airframe and also at the rear of a simulated conventional fuselage. 

All of these tests are part of NASA’s partnerships with U.S. industry to develop open rotor propulsion 

systems, but only the F31/A31 data collected at Glenn are used in this study.

TABLE 1.—LOCATIONS OF ACOUSTIC BARRIER

WALL RELATIVE TO OPEN ROTOR MODEL

Wall 

description

Wall 

position

Wall length,

in.

Location of leading edge of wall

Long Forward 174.0 40.375 in. upstream of center of aft rotor

Long Aft 126.0 Even with center of aft rotor

Short Forward 28.625 40.375 in. upstream of center of aft rotor

Short Aft 28.625 Even with center of aft rotor
NASA/TM—2012-217218 4

2.1 Data Collection

A long and a short acoustic barrier wall (first reported in 1997 in Ref. 13) were placed between the 

F31/A31 open rotor source and the linear microphone array as shown in Figure 2. The impermeable 

barrier walls have been used in the facility to separate and mask noise sources radiating from targeted 

emission angles. For example, the long wall has successfully been used to virtually remove the discharge 

noise of shrouded fan models by shielding it from the microphones so that forward-radiated fan noise may 

be more closely examined in isolation.

In the tests of the F31/A31 open rotor, the barrier wall was configured in four different ways as 

shown in Figure 2. The zero reference of the geometric microphone yaw angle in the figure is directly 

upstream. The barrier in the “long wall” configuration approximates a theoretical semi-infinite shielding 

plane. Unfortunately, the long wall geometry is not a realistic representation of an open rotor in a BWB 

installation. The test article is located near the leading edge of the wall, and BWBs are not envisioned to 

use open rotors in a tractor implementation. Therefore, data with the long wall in place are not used in this 

assessment. The long wall data, however, are still of value to validate and verify analytical shielding 

prediction tools that predict diffraction in the presence of a semi-infinite wall.

The “short wall” configuration better represents the kind of empennage shielding expected of the 

aircraft illustrated in Figure 1. The short barrier wall located in the aft position may represent the 

installation on the proposed open rotor version of the A30X, while the short wall in the forward position 

may represent the open rotor installation of the notional ecoJet. A photograph of the installation in the 

tunnel is shown in Figure 3.

The barrier wall is 2.75 in. thick with an elliptical leading edge and a round trailing edge. The 

thickness of the wall and its rounded leading edge (in contrast to a sharp-edged, thin, flat plate, where 

diffraction behavior can be much different) contribute to the realism of the arrangement by roughly 

approximating an actual tailplane. The short barrier wall is 28.625 in. long and is located 26 in. from the 

centerline of the rotor model. The forward and aft rotor diameters are 25.7- and 24.8 -in., respectively. As 

a point of reference, the short barrier wall is approximately 12 percent greater than one forward rotor 

diameter in length. The microphone traverse array is located 60 in. from the centerline of the rotor model. 

In the forward and aft wall positions, the wall leading edge is located 40.375 in. ahead of, and even with,

the center of the aft rotor, respectively (see Table 1.).

Figure 2.—F31/A31 open rotor test article and barrier wall positions in the NASA Glenn 9- by 15-ft Low 

Speed Wind Tunnel. (Diagram courtesy of Edmane Envia, NASA Glenn Research Center.)
NASA/TM—2012-217218 5

Figure 3.—F31/A31 open rotor test article and “long” barrier 

wall in the NASA Glenn 9- by 15-ft Low Speed Wind 

Tunnel. (Photograph courtesy of David Stephens, NASA 

Glenn Research Center.)

Each tunnel “reading” consists of narrowband spectral density levels taken at 18 sideline microphone 

locations with frequencies ranging from zero to 100 kHz at frequency intervals of 12.2 Hz. Data with the 

open rotor model operating at corrected shaft speeds of 5550 and 6300 rpm (85 and 96 percent of design 

speed) were collected. Both rotors always operated at identical shaft speeds. Tests were also made to 

estimate the facility’s background noise levels with no power supplied to the open rotor test article and 

with the blades windmilling. 

The spectral density levels are corrected by the facility’s data acquisition software to lossless, 

freefield conditions using an atmospheric attenuation model. The spectral density levels are further 

corrected for spherical spreading from the microphone’s constant sideline distance to a 1-ft constant 

radius arc.

Emission yaw angles (θE) are computed from the geometric sideline microphone array angles (θG) 

using the relation θE = θG – sin–1

[MTunnel sin θG], where the tunnel Mach number MTunnel in the experiments 

is 0.20. Data were also collected for MTunnel = 0.22, but they are not used in this assessment. 

Insertion losses are calculated by Stephens in Reference 9 by subtracting the measured spectral 

density levels of the open rotor tests conducted with and without the short barrier wall in place. These 

insertion losses are shown in Figure 4 as a function of narrowband frequency and geometric yaw angle for 

the short barrier wall in the forward position. Note that positive values of insertion loss represent noise 

reductions perceived by the microphones. Also note that the data are plotted to frequencies of only 

10 kHz. The insertion losses are shown for both tonal (left) and broadband (right) components. The 

shielding is the most effective in the deepest shadow zone of the wall (at about 51° from the inlet), where 

the attenuation is 5 dB or greater. The barrier wall is partially effective (less than 5 dB) in the transition 

zone, and it is not effective at all in the bright zone. There is, interestingly, some amplification of the 

noise—particularly of the rotor-rotor interaction tones—in the aft quadrant. In this area, the microphones 
NASA/TM—2012-217218 6

perceive two constructively-interfering sources: from the source itself as well as from the noise diffracted 

at the edge of the wall. This noise amplification can be seen in the blue areas in Figure 4. Similar insertion 

losses are shown in Figure 5 for the short wall in the aft position. The corrected rotor shaft speeds in these 

tests are 6300 rpm.

2.2 Data Manipulation

The measured, model-scale, narrowband data must be manipulated in several ways in order to 

properly compute certification noise shielding benefits for an actual, full-scale aircraft. The spectra have 

already been corrected by the facility’s data acquisition software to freefield, lossless spectral density 

levels on a 1-ft arc. The spectra are modified further by applying the following corrections.

Figure 4.—Open rotor insertion loss at 6300 rpm with the short wall in the forward position; tones (left), 

and broadband (right). Plots courtesy of David Stephens, NASA Glenn Research Center (Ref. 9).

Figure 5.—Open rotor insertion loss at 6300 rpm with the short wall in the aft position; tones (left), and 

broadband (right). Plots courtesy of David Stephens, NASA Glenn Research Center (Ref. 9).

Insertion Loss, dB

-15

-10

-5

0

5

10

15

Frequency, kHz

10

8

6

4

2

0

Geometric Microphone Yaw Angle, deg.

20 40 60 80 100 120 20 40 60 80 100 120

Insertion Loss, dB

-15

-10

-5

0

5

10

15

Frequency, kHz

10

8

6

4

2

0

Geometric Microphone Yaw Angle, deg.

20 40 60 80 100 120 20 40 60 80 100 120
NASA/TM—2012-217218 7

2.2.1 Removal of Background Facility Noise

Noise measurements with the F31/A31 test article in an unpowered windmilling mode and with the 

tunnel in operation were conducted for each of the barrier wall configurations. The spectral density levels 

of the unpowered tests are subtracted from each noise spectrum of interest in the powered tests for

frequencies less than 1000 Hz, where background facility noise is problematic and its removal is required.

The spectral densities are then converted to sound pressure levels with a reference pressure of 20 μPa.

2.2.2 Conversion to Static Conditions

The sound pressure levels are corrected for convection effects to static conditions by adding an 

amplitude correction of 10 log10[1 – MTunnel cos θE]

SME, where the source motion exponent (SME) is taken 

to be 2.0. No corrections are made to the frequency scale, since the microphones are “traveling with” the 

source at the speed of the tunnel Mach number, which in these cases is 0.20.

2.2.3 Conversion to Flight Conditions

The sound pressure levels are corrected to realistic takeoff and approach flight conditions by 

subtracting an amplitude correction of 10 log10[1 – MFlight cos θE]

SME. The flight Mach number, 

MFlight, depends on airplane trajectory calculations for approach and departure, and are discussed in 

Sections 2.3.1 and 3.1. The source motion exponent is again taken to be 2.0. A frequency shift of 

1 – MFlight cos θE is applied to the frequency scale. An ambient source strength correction relative to 

international standard atmospheric (ISA) conditions of 10 log10[(ρ/ρISA)

2 (c/cISA)4

)] is also made at this 

time, where ρ and c are the density and the speed of sound at the flight condition of interest, respectively.

Standard acoustic day (ISA+18 °F) conditions are used for this adjustment.

2.2.4 Conversion to Full Scale

The sound pressure levels are corrected from model scale to full scale using the linear scale factor 

(LSF) and the area scale factor (ASF). Amplitudes are adjusted by adding a correction of 10 log10[ASF], 

and frequencies are shifted to lower values by a factor of LSF. Based on current, parallel NASA open 

rotor system studies (Refs. 14 and 15) values for LSF and ASF are taken to be 5.27 and 27.8, respectively.

2.2.5 Conversion to 1/3rd Octave Band Center Frequencies

The frequency basis for FAA and ICAO aircraft noise certification is the traditional 1/3rd octave band 

center frequency spectrum, represented by 24 discrete sound pressure levels ranging from 50 Hz to 

10,000 Hz. The acoustic energy contained in the intervals on either side of each center frequency (i.e., 

from 2–1/6 times the center frequency to 21/6 times the center frequency) are summed to form each of the 

24 standard 1/3rd octave band sound pressure levels. 

Several of the above processes are illustrated in Figure 6 using a sample spectrum taken from a 

microphone reading in the forward quadrant. A 1-ft-arc, lossless, narrowband spectrum with facility noise 

removed is shown by the black line in Figure 6 (left). This spectrum has an emission yaw angle of 24°

from the inlet, the tunnel Mach number is 0.20, and the F31/A31 open rotor test article is operating at a 

corrected shaft speed of 6300 rpm. The short barrier wall is not present. Although noise measurements are 

available up to 100 kHz, only lower-frequency data are plotted.

When flight condition scaling is applied (scaling to a flight Mach of 0.26 at 1000 ft is shown in the 

figure) the spectrum shifts to higher amplitudes and frequencies to the condition represented by the red 

line in the figure. Opposite shifting behavior occurs for spectra in the aft quadrant. Scaling from 

experimental tunnel conditions to appropriate, simulated flight conditions can result in changes of several 

decibels in the fore or aft quadrants and is thus an important processing step.
NASA/TM—2012-217218 8

Figure 6.—Sample open rotor scale-model narrowband spectrum conversion to full size (left); and 

conversion from full-scale narrowband spectrum to standard 1/3rd octave band center frequencies (right).

When the linear and area scale factors are applied, the spectrum shifts to much higher amplitudes and 

lower frequencies represented by the blue lines in both sides of the figure. At one-fifth scale, the inaudible 

sound pressure levels measured at frequencies of about 50 kHz are shifted to the audible, regulated range 

used to estimate noise certification levels.

The 1/3rd octave band sound pressure levels are computed as described above and are plotted in 

Figure 6 (right). The fundamental blade passage frequency tones for the front and aft rotors are labeled 

“1f” and “1a” in the figure, respectively. Several higher-order rotor-rotor discrete interaction tones are 

also noted in the figure. Seldom do any of the interaction tones above the 2f+2a tone contribute 

significantly to the overall sound pressure level.

The reasoning behind the frequency “coarsening” from the narrowband to the 1/3rd octave band is that 

the resulting spectra become much easier to manipulate and store. Many of the rotor-rotor interaction 

tones lie within a single frequency interval, such that a single 1/3rd octave band sound pressure level may 

incorporate two (or more) tones, particularly at higher frequencies since the scale is logarithmic. In an 

open rotor aircraft noise certification test (or a simulation of one), the acoustic spectra observed by the 

microphone will fluctuate significantly as tones move from one frequency interval into another as the

aircraft-observer geometry changes, and convective amplification and Doppler effects enter play. This 

will lead to irregular tone-weighted, perceived noise level (PNLT) time histories as the airplane traverses 

a path relative to its certification observers. This is an effect observed in large turboprops in the transport 

category as well, but it should be even more pronounced for an open rotor airplane with its much more 

complex tonal content.

The strategy of using unequal blade counts is apparent. With 12 blades on the front rotor and 10 

blades on the aft rotor, and with both shafts turning at the same rate, the shaft tones are distributed more 

broadly throughout the spectrum. And when the narrowband spectra are “coarsened” to the 1/3rd octave 

band, much tone detail is lost and, in some cases, tones may not even be noticeable. This is a noise 

advantage not typically seen in conventional general aviation propellers, or even in civil aviation 

turboprop aircraft, where single-rotation propellers having more than six blades are uncommon. Indeed, at 

some emission angles, scale factors and flight conditions, the tonal penalty may be insignificant. For 

example, for the case illustrated in Figure 7, the tone penalty of the PNLT metric is only 0.2 PNdB! The 

PNLT metric is of particular importance since it is the basis of Part 36 noise certification.

70

80

90

100

110

120

130

140

150

10 100 1000 10000

SPL, dB

Frequency, Hz

Processed and Scaled Narrowband

Processed and Scaled 1/3rd Octave Band

1a

1f

1f+1a

1f+2a

2f+2a 5 dB

70

80

90

100

110

120

130

140

150

0 1000 2000 3000 4000 5000

SPL, dB

Frequency, Hz

Data with Facility Noise Removed

Brought to Simulated Flight Conditions

Scaled to Full Size
NASA/TM—2012-217218 9

Figure 7.—Sample open rotor full-scale spectrum showing 

very little tonal penalty.

2.3 Computation of System Shielding Benefits

2.3.1 Takeoff and Approach Trajectories

An analytical aircraft model with twin open rotor propulsion systems mounted on the aft fuselage is 

used in this assessment. Its general configuration is similar to the U-tail aircraft shown in Figure 1. This 

notional airplane is used for the parallel NASA effort and is described in greater detail in Refs. 14 and 15. 

An airplane model with appropriate low-speed aerodynamics, weight, and propulsion performance data is 

necessary to model the takeoff and approach trajectories needed for aircraft certification predictions. The 

airplane is modeled using NASA’s Flight Optimization System (FLOPS (Ref. 16)) computer program. 

FLOPS computes detailed, low-speed takeoff and landing profiles using a built-in, time-stepping 

trajectory analysis module. Compliance with the airworthiness requirements described in Parts 36 and 25 

of the Federal Aviation Regulations (Refs. 1 and 17, respectively) is observed.

2.3.2 System Noise Prediction

The Aircraft Noise Prediction Program (ANOPP, Release Level 29 (Refs. 18 and 19)) is a systemslevel code used in this study to compute certification noise for the open rotor airplanes. The center 

frequency basis for ANOPP is the 1/3rd octave band. The scaled 1/3rd octave band open rotor spectra 

described above are fed into ANOPP via its Acoustics Data Module (ACD (Ref. 20)). ACD is an ANOPP

utility that allows user-supplied spectra to be fed into a certification simulation in lieu of using ANOPP’s 

own, built-in, source noise prediction modules.

Extreme care must be taken by users of ANOPP’s ACD module, especially when attempting to model 

noise sources with tone content. As of ANOPP Level 29, the ACD module does not attempt to modify the 

user-input spectral content to account for convection effects as the airplane’s airspeed and altitude vary 

throughout its trajectory. The onus therefore is on the user to supply proper spectra to ACD with a 

sufficient quantity of flight Mach numbers, and with the convection effects for each Mach number 

properly modeled and accounted for offline. Furthermore, the ACD module linearly interpolates for sound 

pressure level data between these flight Mach numbers, as well as across engine power setting, emission 

yaw angle, and emission roll angle. Quite often, a tone is improperly “washed out” by artifacts of 

interpolation. And perhaps even less obvious, a tone may be artificially diminished by interpolation error 

from simple convection effects. As Doppler shifts occur when the aircraft-observer geometry changes, a 

tone may move from one frequency interval into another and cause an interpolation error. Broadband 

noise is of course less prone to this effect. The solution to this problem is to eliminate interpolation error 

70

80

90

100

110

120

130

140

150

10 100 1000 10000

SPL, dB

Frequency, Hz

Processed and Scaled Narrowband

Processed and Scaled 1/3rd Octave Band

5 dB
NASA/TM—2012-217218 10

whenever possible by specifying data at the exact power settings required for the trajectory, and to 

minimize interpolation error by supplying data in sufficient quantity for the remainder of dependencies 

(i.e., flight Mach number and emission angles). 

The spectra are propagated to the three certification observers on the ground in accordance with the 

specifications for certification measurements. Noise propagation effects accounted for include spherical 

spreading, Doppler shift and convective amplification, atmospheric attenuation within a layered 

atmosphere, ground reflections based on data for grass-covered ground, and extra ground attenuation. 

More complex propagation phenomena such as scattering, weather effects, and terrain are not modeled. 

The airplane trajectory, computed by FLOPS as described above, is fed into the ANOPP simulation. 

Vector geometry analyses for the airplane relative to the three certification microphone measurement 

locations—shown in Figure 8—are performed within ANOPP as functions of source time. Note that in 

the interest of international rulemaking terminology harmonization, the former “sideline” certification 

location term has been deprecated in favor of “lateral,” as has “takeoff” to “flyover.” In any event, the 

propagated acoustic spectra are predicted at half-second intervals at each of the three certification 

locations on the ground. From these propagated spectra, ANOPP computes several noise metrics of 

interest as functions of observer time. The Effective Perceived Noise Level (EPNL) certification noise 

metric is computed from the noise-time history at each observer as prescribed in Reference 1. In noise 

certification parlance, the cumulative, or algebraic, sum of the three certification EPNLs is often used to 

capture the range of operating conditions.

Frequently, artifacts resulting from ground reflections can cause irregularities in the spectra measured 

at the certification microphone. These spectral irregularities—called pseudotones—typically occur at 

frequencies below 800 Hz where ground reflections are most pronounced. This phenomenon is often 

observed in ANOPP when ground reflections are analytically predicted. Part 36 regulations permit the 

tone correction penalty of the PNLT to be ignored for pseudotones, provided they can be shown to be 

unrelated to engine source noise. Tones under 800 Hz, however, are very common for an open rotor 

propulsor (see, e.g., Fig. 6, right.). Thus, unlike most turbofan engine certification modeling, tone

correction penalties under 800 Hz are not ignored in this simulation and they do contribute to the PNLT.

Figure 8.—Noise certification observer arrangement relative to hypothetical combined takeoff 

and landing aircraft trajectories.

2000 m

(6562 ft)

Flyover (with cutback)

reference

Lateral

reference

Approach 

reference

6500 m

(21 325 ft)

450 m

(1476 ft)
NASA/TM—2012-217218 11

2.4 Comparison of Data to Maekawa Diffraction Model

The open rotor insertion loss data measured with the short barrier wall in place is compared to a simple

empirical diffraction model based on asymptotic results of optical diffraction theory, originally proposed by 

Maekawa (Ref. 21) and reproduced in many foundational acoustic textbooks (e.g., Ref. 22). The analytic 

treatment of diffraction effects in this manner is common in aeroacoustic applications. Being reliable, fast, 

and easy to implement, it has been coded into aircraft noise system prediction programs such as the General 

Aviation Synthesis Program (Ref. 23) in 1982 and later into ANOPP (Ref. 24). Thus, any validation of the 

Maekawa model with respect to new, measured aeroacoustic data is a valuable undertaking.

Maekawa proposed the shadow zone barrier attenuation relation:

IL = 20 log10 ( 2π NF tanh 2π NF )+ 5 ,

in dB, where NF is the frequency-dependent Fresnel number (2f∆/c), whose characteristic length ∆ is the 

difference between the shortest path around the barrier between the source and the observer and the 

source-observer distance directly through the barrier. 

For observers in the bright zone (NF < –0.192), the attenuation is neglected, and for observers in the 

transition zone (–0.192 < NF < 0), it is appropriate to replace the hyperbolic tangent with the 

trigonometric tangent. Although the above relation is intended for use with semi-infinite barriers, 

Maekawa suggested that superposition may be used for barriers of finite length and width.

The Maekawa diffraction model is intended for insertion loss predictions of acoustically compact 

sources, where the geometry-dependent ∆ may be determined using distances relative to a radiating point 

source. When the Maekawa relation is used in this investigation, the entire open rotor test article is assumed 

to be a compact point source radiating from the center of the aft rotor. The open rotor model, however, is not 

at all compact. Very efficient noise generation occurs nearer the rotor tips, at more than 12 in. from the rotor 

axis. Using the Maekawa relation in this way pushes it beyond its intended capability.

A modified relation is proposed here:

1 10 ( 2 ) 5 2 IL = 20 a log 2π NF tanh π NF + + a ,

where a1 and a2 are proposed new calibration constants intended to improve the prediction with respect to 

the open rotor data measured in this experiment. The influence of a1 on IL is shown in Figure 9. It has the 

effect of reducing the magnitude of IL in the shadow zone while having a lesser impact in the transition 

zone. a2, of course, simply shifts the entire relation. In the figure, a2 is zero. A value of unity for a1 yields 

Maekawa’s original relation.

Figure 9.—Modified Maekawa insertion loss relation 

showing influence of calibration constant a1.

0

5

10

15

20

25

-1 0 1 2 3 4 5 6 7 8 9 10

Insertion Loss IL, dB

Fresnel Number, N

1.2

1.0

0.8

0.6

0.4
NASA/TM—2012-217218 12

3.0 Results and Discussion

3.1 Trajectory Results

Takeoff and approach trajectories and throttle settings must be modeled properly to correctly compute 

certification noise. An open rotor propulsor’s thrust characteristics differ from the performance of 

turbofans, and thus a trajectory calculation is necessary to capture these effects. This is often in contrast to 

many other certification noise estimates made during the conceptual phase of aircraft design, where fixed 

trajectories and throttle settings are often assumed based on previously-collected data from other 

representative aircraft.

Trajectory data computed for the open rotor airplane for altitude above field elevation (AFE), 

airspeed, and throttle setting are shown in Figure 10, 11, and 12, respectively. The trajectory shown is

presented as an operation with both takeoff and landing data shown simultaneously. For presentation 

purposes, the touchdown point on landing is coincident with the point of brake release on takeoff. 

Calculations are made for a sea level runway at standard acoustic day (ISA+18°F) conditions. The 

triangular markers on each plot denote the three noise certification measurement locations.

Figure 10.—Open rotor airplane altitude. Figure 11.—Open rotor airplane airspeed.

Figure 12.—Open rotor airplane throttle setting.

0.0

0.5

1.0

1.5

2.0

2.5

-15 -10 -5 0 5 10 15 20 25 30

Altitude, 1000 ft AFE

Distance from Brake Release, 1000 ft

0

50

100

150

200

250

-15 -10 -5 0 5 10 15 20 25 30

True Airspeed, ktas

Distance from Brake Release, 1000 ft

0.0

0.2

0.4

0.6

0.8

1.0

1.2

-15 -10 -5 0 5 10 15 20 25 30

Thrust / Maximum Thrust

Distance from Brake Release, 1000 ft
NASA/TM—2012-217218 13

The approach microphone markers are shown in the figures at 6562 ft (2000 m) behind the runway 

threshold (i.e., behind the location of the 50 ft landing obstacle), and 7518 ft from the touchdown point on 

the runway centerline. The lateral microphone markers are shown in the figures at 13,000 ft from brake 

release, on a lateral sideline displacement distance of 1476 ft (450 m) from the runway centerline. The 

13,000 ft location corresponds to the distance from brake release where the airplane has reached 984 ft 

(300 m) AFE, as permitted—when approved—by regulation B36.3.a.1 (Ref. 1). A check of other lateral 

locations less than and greater than 13,000 ft from brake release confirms that peak lateral EPNL does in 

fact occur approximately at the 13,000 ft location. The flyover microphone markers are shown in the 

figures at 21,325 ft (6500 m) from brake release on the runway centerline. A noise abatement throttle 

cutback is used. The engine climb thrust at this point is reduced to the minimum level permitted by 

regulation (i.e., Ref. 1 requires a minimum climb gradient of 4 percent with both engines operating, or 

level flight with one engine inoperative.). The throttle cutback takes place between 16,000 and 17,000 ft 

from brake release. Note the power cutback takes place at approximately 1300 ft AFE. This is above the 

minimum altitude permitted (i.e., 984 ft, or 300 m AFE for a twin-engine airplane), in an attempt to gain 

additional altitude and to minimize the noise at the flyover observer. 

The computed trajectories and throttle settings are used to guide the open rotor data scaling process 

described in Section 2.2.3. For the barrier wall tests conducted in October, 2010, the F31/A31 test article 

was operated at only two shaft speeds. Based on guidance from the open rotor propulsion system model

analyzed by Hendricks (Ref. 14), the acoustic data collected at the corrected shaft speed of 5550 rpm is 

used to predict the approach observer noise, while the data collected at the corrected shaft speed of 

6300 rpm is used for the both lateral and flyover noise predictions. These shaft speeds are equivalent to 

1050 and 1200 rpm after scaling to full size with a linear scale factor of 5.27, respectively.

3.2 Certification Noise Results

The variation of PNLT with emission yaw angle is shown in Figure 13 for the unshielded source

(solid black lines) along with PNLTs of the source shielded by the short wall in the forward (top) and aft

(bottom) positions. The lossless, freefield data are shown on a constant, 1-ft arc. They have been scaled 

using values for LSF and ASF of 5.27 and 27.8, respectively, using a flight Mach number of 0.30. The test 

article corrected shaft speed is 6300 rpm, with the effects of scaling simulating a speed of 1200 rpm. The 

shielding benefits of the short wall are apparent: insertion losses as high as 8.8 PNdB are present with the 

wall in either position. 

Figure 13.—Variation of PNLT with emission yaw angle, 

with and without barrier walls. (Lossless data on 1-ft arc, 

scaled to full size and to flight Mach 0.30. Note vertical 

offset shift for clarity.)

120

125

130

135

140

145

150

155

160

0 20 40 60 80 100 120 140

Change in PNLT, PNdB

Emission Yaw Angle, deg.

5 dB

Withoutwall

With short wall 

in forward 

position

With short 

wall in aft

position
NASA/TM—2012-217218 14

The PNLT scale is not shown in the figure since only the differences in PNLT (and EPNL) are 

important in this study of noise shielding. Emphasis is not placed on the absolute predicted levels of the 

EPNLs, but rather on the differences between them. The F31/A31 open rotor (an early 1990s design) is 

also not quite representative of a modern open rotor design that may one day fly. Advanced open rotor 

propulsors will be designed using 3D aerodynamic analytical tools and will employ advanced, low-noise 

features: some of which have been made public and others that are still company-proprietary. Noise 

mitigation strategies that have been made public (Ref. 25) include aeroacoustic blade shaping, blade pitch 

angle and rotational speed optimization, increased blade counts, low disk loading, rotor-rotor spacing and 

pylon-rotor spacing optimization, aft rotor clipping, and pylon wake reduction. Moreover, the data 

collected with the barrier wall in place did not have a simulated pylon or fuselage installed, nor was there 

any variation in angle of attack.* The aerodynamic installation is known to have a significant effect on 

open rotor noise (see, e.g., Ref. 26).

The next step in assessing the system-level benefit of the insertion loss is to compute the certification 

EPNLs. The overall noise benefit is not readily apparent from Figure 13 alone, since trajectory effects 

come into play. The PNLT-time noise history naturally rises and falls as the source first approaches—and 

then recedes from—a ground-based observer. The source’s directivity pattern and source-motion 

convection effects also play important roles in this noise system problem. 

The benefit of the insertion loss may be squandered if the peak PNLT perceived by an observer 

occurs at an angle where there is no shielding benefit. Correct placement of the propulsors relative to the 

U-tail is important if any system noise shielding benefit is expected.

The scaled acoustic data are used to generate spectra at half-second time intervals along the 

trajectory. The spectra are propagated to each observer, and the overall sound pressure level (OASPL)

and PNLT noise-time histories are computed and plotted in Figures 14, 15, and 16 for the approach, 

lateral, and flyover observers, respectively. Observer time relative to the point of brake release (or 

touchdown) is used as the independent parameter in each figure. The OASPL metric is shown on the left 

in each figure because of its simplicity and its ability to clearly show the relatively smooth rise and fall of 

each noise source over time. The PNLT metric—shown on the right in each figure—has qualities that 

capture level, frequency weighting, and tone annoyance penalties. Its time histories are therefore much 

more irregular than the OASPL histories: as the airplane approaches and recedes, Doppler and convective 

amplification effects influence the PNLT metric’s frequency-weighting and tone penalties. Allowing sub800 Hz tones to contribute to the PNLT’s tone correction penalty (as noted in Section 2.3) adds to the 

irregularities. In some instances, ground effects and tone corrections may cause the PNLT to actually 

increase as the airplane recedes from the observer. Note that only the PNLT-time histories within an

integration region of 10 PNdB from the maximum PNLT are used to compute EPNLs.

Figure 14.—Approach observer OASPL noise-time histories (left), and PNLT noise-time histories (right).

 *

These influence of these effects were measured in other tests at the NASA Glenn facility, but not in the tests made 

with the barrier wall.

7 0

7 5

8 0

8 5

9 0

9 5

-60 -55 -50 -45 -40 -35 -30

Approach OASPL, dB

Observer Time, s

Unshielded

Fwd. wall shielding

Aft wall shielding

5 dB

8 0

8 5

9 0

9 5

100

105

110

-60 -55 -50 -45 -40 -35 -30

Approach PNLT, PNdB

Observer Time, s

Unshielded

Fwd. wall shielding

Aft wall shielding

5 dB
NASA/TM—2012-217218 15

Figure 15.—Lateral observer OASPL noise-time histories (left), and PNLT noise-time histories (right).

Figure 16.—Flyover observer OASPL noise-time histories (left), and PNLT noise-time histories (right).

From the appearance of the figures, the short barrier wall in the aft position seems to be nearly ideally 

placed: the maximum insertion loss approximately coincides with the peak PNLT of the unshielded 

source. The EPNL shielding benefits of the barrier wall in the forward and aft positions are shown in 

Table 2. The maximum shielding benefit is approximately 9 cumulative EPNdB. 

TABLE 2.—EPNL BENEFIT DUE TO SHIELDING FOR EACH OBSERVER (EPNDB)

Wall position Lateral Flyover Approach Cumulative

Forward –0.9 –0.6 –1.3 –2.8

Aft –2.7 –2.9 –3.3 –8.9

Although 9 EPNdB falls slightly short of the 10 EPNdB that is required for the NASA goal, there is 

some reason to believe that the shielding benefits calculated here may be somewhat conservative. 

Certainly a blended wing body airplane, with far more shielding surface area than a conventional airframe 

configuration, is likely to provide a greater shielding benefit. Forward-radiated propulsor noise may be 

virtually eliminated on such an architecture. But even for the tube-and-wing airplane assumed here, where 

the shielding surface is small relative to a BWB planform, the shielding benefit may be underestimated. 

For a lateral observer, there would undoubtedly be additional shielding of one of the open rotor 

propulsors by the rear fuselage that is not accounted for in this assessment.

On the other hand, there is also reason to believe that the insertion losses measured in the Glenn 

facility are greater than what would be observed in actual flight. In an in-flight application, ground 

observers would be located much further away from the source and the barrier than the microphones are 

6 0

6 5

7 0

7 5

8 0

8 5

9 0

40 45 50 55 60 65 70 75 80 85 90

Lateral OASPL, dB

Observer Time, s

Unshielded

Fwd. wall shielding

Aft wall shielding

5 dB

6 5

7 0

7 5

8 0

8 5

9 0

9 5

40 45 50 55 60 65 70 75 80 85 90

Lateral PNLT, PNdB

Observer Time, s

Unshielded

Fwd. wall shielding

Aft wall shielding

5 dB

6 0

6 5

7 0

7 5

8 0

8 5

9 0

60 65 70 75 80 85 90 95 100 105 110

Flyover OASPL, dB

Observer Time, s

Unshielded

Fwd. wall shielding

Aft wall shielding

5 dB

7 0

7 5

8 0

8 5

9 0

9 5

100

60 65 70 75 80 85 90 95 100 105 110

Flyover PNLT, PNdB

Observer Time, s

Unshielded

Fwd. wall shielding

Aft wall shielding

5 dB
NASA/TM—2012-217218 16

Figure 17.—Variation of leading-edge Fresnel number with 

source-to-observer distance (frequency 1000 Hz, emission 

yaw angle 102°, barrier wall in aft position).

in the Glenn facility. In the experiment, the microphone array is only 60 in. away from the centerline of 

the open rotor rig, and just 31 in. away from the acoustic barrier wall. This geometry results in Fresnel 

numbers that are higher than they would be in an actual flight test. In an airplane flyover condition, the 

source-to-observer distances are very large relative to the source-to-shield distance. In a flyover, the 

characteristic lengths of the Fresnel number (∆) would approach their asymptotic limits. This effect is 

shown graphically in Figure 17. In the figure, the Fresnel number (at 1000 Hz) is computed for varying 

source-to-microphone distances for an emission yaw angle of 102° with the short barrier wall located in 

the aft position. A figure like this could be drawn for any geometry, but this emission angle is chosen in 

particular because it lies is the deepest part of the shadow zone. The barrier wall and open rotor test article 

geometry remains fixed. The Fresnel number for the geometry in the facility is 1.7 (marked by the symbol 

in the figure), while its asymptotic limit is about 1.0. For any given observer location, the Fresnel 

numbers of the experiment are higher than they would be in an actual flight test. The actual noise 

shielding benefit in flight therefore may be expected to be less than the benefit indicated by the 

experimental data.

Other sources of noise would also exist. In addition to the open rotor propulsor noise, core noise 

radiating from the gas generator, turbulent flow noise sources radiating from parts of the airframe, and 

even some jet noise from the exhaust will combine to form the overall noise signature of the airplane.

Except for the gas generator core noise, the U-tail would provide little or no shielding, and the overall 

shielding benefit predicted here would shrink.

Given these issues, an uncertainty analysis of the EPNL calculations presented here is difficult to 

perform with the data available here alone. It should be noted, however, that noise shielding data 

collected within the confines of any relatively small facility would be prone to the same limitations. It is 

therefore advisable, perhaps, that any experimental shielding data collected in a facility be used to 

validate and verify more accurate analytical acoustic diffraction tools, which in turn could be used to 

more accurately predict the certification noise benefits of shielding.

3.3 Comparison with Maekawa Diffraction Model

The insertion losses are shown in Figure 18 for the short barrier wall in the forward position. The 

measured insertion loss (indicated by the blue line) is determined by subtracting the spectra recorded with 

the barrier wall in place from the spectra recorded without the wall. No scaling of the data is performed. 

In other words, the sound spectra are left unadjusted for size and convection effects, with LSF = ASF = 1 

and MTunnel = 0.20. The data are presented on a constant 1-ft arc. All insertion losses are summed to 

0.0

0.5

1.0

1.5

2.0

2.5

0 100 200 300 400

Fresnel Number, 

NF, at 1000 Hz

Source-Observer Distance, in.

Asymptotic limit

Facility installation
NASA/TM—2012-217218 17

OASPLs so that their variations with emission angle can be more easily shown in the figure. The 

maximum OASPL reduction (8 dB) occurs in the shadow zone, at an emission yaw angle of 43°.

The Maekawa diffraction model described in Section 2.4 is used to predict the insertion loss. For this 

purpose, the open rotor test article is assumed to be an acoustically-radiating point source. The source 

location used to compute the characteristic lengths (∆) is taken to be at the center of the aft rotor. The loss 

predicted by the unmodified, original Maekawa relation—that is, with the calibration constant a1 set to 

unity and a2 set to zero—is indicated by the red line in the figure.

The Maekawa relation appears to overpredict the maximum insertion loss by about 13 dB. The 

method, however, is being used beyond its original scope. With forward and aft rotor diameters of 25.7 in. 

and 24.8 in., respectively, and with efficient noise generation occurring near the rotor tips, the open rotor 

model is certainly not a compact source. Computing Fresnel numbers in this manner introduces error.

The method can make an improved prediction by altering the calibration constants a1 and a2 from 

their neutral values of unity and zero, respectively. Setting a1 = 0.3 with a2 = 0, or setting a2 = –14 with 

a1 = 1.0 provides a better fit of the data. The predicted insertion loss using a1 = 0.3 and a2 = 0 is indicated 

by the green line in the figure. These calibration values match the shielded data equally well when the 

wall is in the aft position.

From an ANOPP user’s perspective, however, it is not possible to implement alternative values of 

a1 or a2; at least via the ordinary use of the WING module. The Maekawa method, as coded into 

WING, does not allow Maekawa’s relation to be customized via user input. Furthermore, the output of 

the WING module (a data table that is also named “WING”) is not defined to the database manager, so 

that any other modules that are designed to alter data tables (e.g., the General Suppression Module

(Ref. 27)), cannot be used.

Possibly the only way to alter the values in the WING data table would be to unload the member data 

into an external file, manipulate the data by means of some offline analysis, and load the revised member 

data back into the user’s simulation. This approach would be very tedious. Modifying the source code

with new features may be the only practical option.

Another alternative is to wait for a higher-order diffraction method to be implemented into ANOPP.

The Maggi-Rubinowicz formulation of Kirchhoff’s diffraction theory (Ref. 28) is anticipated to be 

programmed into a release of ANOPP later in 2011.

Figure 18.—Measured and predicted insertion losses with 

the short barrier wall in the forward position. (Data on 1-

ft arc, tunnel conditions, no size scaling).

-25

-20

-15

-10

-5

0

5

0 20 40 60 80 100 120 140

Insertion Loss OASPL, dB

Emission Yaw Angle, deg.

Measured loss

Predicted loss 

(Maekawa)

Predicted loss 

(Modified Maekawa)
NASA/TM—2012-217218 18

4.0 Conclusions

The first phase of NASA’s noise reduction goal calls for aircraft with cumulative Stage 4 margins of 

32 EPNdB. The second phase requires an additional margin of 10 EPNdB. The second phase is thought 

by NASA not to be achievable by reductions in engine and airframe noise sources alone. A change in 

aircraft architecture, where propulsion noise sources are shielded from observers on the ground, is thought 

to be necessary to achieve the additional noise margin. 

This study focuses on predicting the noise certification benefits of a notional open rotor aircraft with 

empennage shielding a portion of the rotor noise. Twin, aft-fuselage-mounted, open rotor propulsors are 

partially shielded by a U-tail structure. The basis of the prediction are noise measurements of an open 

rotor test article collected with and without an acoustic barrier wall that simulates the U-tail. The chord 

length of the simulated U-tail is 12 percent greater than the forward rotor diameter. Two tail positions are 

analyzed. When the tail is located in the aft position, the cumulative system noise benefit is predicted to 

be 9 EPNdB. The tail in the aft position appears to be ideally located to reduce noise perceived by noise 

certification ground observers. With the tail in the forward position, the cumulative noise reduction is 

predicted to be only 3 EPNdB.

The predicted 9 EPNdB margin is encouraging for the tube-and-wing airplane assumed here. A

blended wing body airplane, with far more shielding surface area than a conventional airframe 

configuration, is likely to provide an even greater shielding benefit. Forward-radiated propulsor noise 

may virtually be eliminated on such an architecture.

However encouraging these results may be, the acoustic data used in this analysis may not be ideally 

suited to make a systems noise prediction. The measurements were collected within the confines of a 

relatively small facility, where the insertion loss benefits are likely to be exaggerated relative to those that 

would be experienced during an actual airplane flyover. However, the experimental shielding data 

collected in this study could be used to help validate and verify more accurate analytical acoustic 

diffraction tools. These, in turn, could be used to more accurately predict the certification noise benefits 

of shielding.

References

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Airworthiness Certification.

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Engines,” ISABE XIV Paper 99-7254, 1999, and NASA/TM—2000-210025, 2000.

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NASA/TM—2012-217218 19

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Noise Shielding,” 14th AIAA/CEAS Aeroacoustics Conference (29nd AIAA Aeroacoustics 

Conference), AIAA paper 2008-3050, 2008.
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PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 

1. REPORT DATE (DD-MM-YYYY)

01-01-2012 

2. REPORT TYPE

Technical Memorandum 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

Empennage Noise Shielding Benefits for an Open Rotor Transport 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Berton, Jeffrey, J. 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 561581.02.08.03.18 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

8. PERFORMING ORGANIZATION

 REPORT NUMBER

E-17899 

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

Washington, DC 20546-0001 

10. SPONSORING/MONITOR'S

 ACRONYM(S)

NASA 

11. SPONSORING/MONITORING

 REPORT NUMBER

NASA/TM-2012-217218 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Categories: 05 and 07 

Available electronically at http://www.sti.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

NASA sets aggressive, strategic, civil aircraft performance and environmental goals and develops ambitious technology roadmaps to guide 

its research efforts. NASA has adopted a phased approach for community noise reduction of civil aircraft. While the goal of the near-term 

first phase focuses primarily on source noise reduction, the goal of the second phase relies heavily on presumed architecture changes of 

future aircraft. The departure from conventional airplane configurations to designs that incorporate some type of propulsion noise shielding 

is anticipated to provide an additional 10 cumulative EPNdB of noise reduction. One candidate propulsion system for these advanced 

aircraft is the open rotor engine. In some planned applications, twin open rotor propulsors are located on the aft fuselage, with the vehicle’s 

empennage shielding some of their acoustic signature from observers on the ground. This study focuses on predicting the noise certification 

benefits of a notional open rotor aircraft with tail structures shielding a portion of the rotor noise. The measured noise of an open rotor test 

article--collected with and without an acoustic barrier wall--is the basis of the prediction. The results are used to help validate NASA’s 

reliance on acoustic shielding to achieve the second phase of its community noise reduction goals. The noise measurements are also 

compared to a popular empirical diffraction correlation often used at NASA to predict acoustic shielding. 

15. SUBJECT TERMS

Aircraft engines; Aircraft noise; Contrarotating propellers; Propeller noise 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

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26 

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

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