Gary G. Podboy, James E. Bridges, and Brenda S. Henderson

Glenn Research Center, Cleveland, Ohio

Phased Array Noise Source Localization

Measurements of an F404 Nozzle Plume

at Both Full and Model Scale

NASA/TM—2010-216366

August 2010

GT2010–22601
NASA STI Program . . . in Profi le

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA Scientifi c and Technical Information (STI) 

program plays a key part in helping NASA maintain 

this important role.

The NASA STI Program operates under the auspices 

of the Agency Chief Information Offi cer. It collects, 

organizes, provides for archiving, and disseminates 

NASA’s STI. The NASA STI program provides access 

to the NASA Aeronautics and Space Database and 

its public interface, the NASA Technical Reports 

Server, thus providing one of the largest collections 

of aeronautical and space science STI in the world. 

Results are published in both non-NASA channels 

and by NASA in the NASA STI Report Series, which 

includes the following report types:

• TECHNICAL PUBLICATION. Reports of 

completed research or a major signifi cant phase 

of research that present the results of NASA 

programs and include extensive data or theoretical 

analysis. Includes compilations of signifi cant 

scientifi c and technical data and information 

deemed to be of continuing reference value. 

NASA counterpart of peer-reviewed formal 

professional papers but has less stringent 

limitations on manuscript length and extent of 

graphic presentations.

• TECHNICAL MEMORANDUM. Scientifi c 

and technical fi ndings that are preliminary or 

of specialized interest, e.g., quick release 

reports, working papers, and bibliographies that 

contain minimal annotation. Does not contain 

extensive analysis.

• CONTRACTOR REPORT. Scientifi c and 

technical fi ndings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientifi c and technical 

conferences, symposia, seminars, or other 

meetings sponsored or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientifi c, 

technical, or historical information from 

NASA programs, projects, and missions, often 

concerned with subjects having substantial 

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientifi c and 

technical material pertinent to NASA’s mission.

Specialized services also include creating custom 

thesauri, building customized databases, organizing 

and publishing research results.

For more information about the NASA STI 

program, see the following:

• Access the NASA STI program home page at 

http://www.sti.nasa.gov

• E-mail your question via the Internet to help@

sti.nasa.gov

• Fax your question to the NASA STI Help Desk 

at 443–757–5803

• Telephone the NASA STI Help Desk at

 443–757–5802

• Write to:

 NASA Center for AeroSpace Information (CASI)

 7115 Standard Drive

 Hanover, MD 21076–1320
Gary G. Podboy, James E. Bridges, and Brenda S. Henderson

Glenn Research Center, Cleveland, Ohio

Phased Array Noise Source Localization

Measurements of an F404 Nozzle Plume

at Both Full and Model Scale

NASA/TM—2010-216366

August 2010

GT2010–22601

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

Turbo Expo 2010

sponsored by the American Society of Mechanical Engineers

Glasgow, Scotland, United Kingdom, June 14–18, 2010
Acknowledgments

The authors wish to thank Dr. John T. Spyropoulos of NAVAIR for inviting them to the Naval Air Engineering Station to 

acquire the full-scale engine source location data presented in this report. This work was supported by the NASA

Fundamental Aeronautics Program, Supersonics Project.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

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

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.

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2010-216366 1

Phased Array Noise Source Localization Measurements

of an F404 Nozzle Plume at Both Full and Model Scale

Gary G. Podboy, James E. Bridges, and Brenda S. Henderson

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

A 48-microphone planar phased array system was used to 

acquire jet noise source localization data on both a full-scale 

F404-GE-F400 engine and on a 1/4th scale model of a F400 

series nozzle. The full-scale engine test data show the location 

of the dominant noise sources in the jet plume as a function of 

frequency for the engine in both baseline (no chevron) and

chevron configurations. Data are presented for the engine 

operating both with and without afterburners. Based on 

lessons learned during this test, a set of recommendations are 

provided regarding how the phased array measurement system 

could be modified in order to obtain more useful acoustic 

source localization data on high-performance military engines 

in the future. The data obtained on the 1/4th scale F400 series 

nozzle provide useful insights regarding the full-scale engine 

jet noise source mechanisms, and document some of the 

differences associated with testing at model-scale versus fullscale.

Introduction

The very high levels of jet noise produced by highperformance military aircraft during take-off and landing 

operations pose serious health risks to personnel working on 

aircraft carriers. While the United States Navy considers 

prolonged exposure to sounds above 84 dB as having the 

potential to cause hearing loss, noise levels on a carrier flight 

deck can often exceed 145 dBA (Ref. 1). Depending on the 

susceptibility of the exposed individual, a single exposure to 

impulse noise with a peak level of 140 dB SPL may be 

hazardous (Ref. 2).

The unsafe noise levels are not confined to the carrier flight 

deck. Typically, 1400 sailors live and work on the gallery 

deck, where noise levels often exceed 100 dBA. These high 

noise levels impair the ability to communicate, reduce 

cognitive skill levels, and can result in permanent hearing loss. 

Noise induced hearing loss is the leading occupational 

health expense for the US Navy. In 2004, the Veterans 

Administration paid $108 million dollars in disability 

payments to 15,800 former Navy personnel for hearing loss –

an increase of $65 million over 1999 levels (Ref. 3). 

The severity of this problem has prompted a number of 

studies aimed at characterizing the jet noise produced by 

military aircraft engines. By analyzing F-15 flyover data,

Norum, et al. (Ref. 4) investigated the effects of high speed 

flight on the exhaust noise produced from overexpanded jets. 

Gee, et al. (Ref. 5) examined F/A-18E acoustic data in order to 

determine the role that nonlinear noise propagation plays in 

spectral broadening. Schlinker et al. (Ref. 6) conducted an 

extensive study of data obtained during a ground test of an 

unspecified military engine. They obtained acoustic data in the 

near field, the far field and the very far field in an effort to 

characterize the propagation of high intensity, impulsive jet 

exhaust noise. They also measured the axial distribution of 

noise sources in the jet plume using a linear phased array.

Besides merely characterizing the noise, the US Navy has 

also begun to investigate noise reduction technologies aimed 

at lessening the adverse effects of jet noise. The objective is to 

develop technologies which reduce the noise at the source. 

Adding chevrons to the exhaust nozzles of commercial, 

subsonic aircraft has proven to be an effective noise reduction 

technology for medium- and high-bypass ratio turbofan 

engines (Refs. 10 and 11). Chevrons are triangular serrations 

in the nozzle trailing edge which induce streamwise vorticity 

into the shear layer. This added streamwise vorticity increases 

turbulent mixing which, in turn, tends to reduce jet noise. The 

application of chevrons to military aircraft is particularly 

attractive since existing engines can be retrofitted with these 

noise reduction devices. The exhaust jets of high performance 

tactical aircraft at takeoff, however, are typically 

overexpanded. Consequently, chevrons designed for subsonic 

jet noise reduction may not be effective in military engine 

applications.

In July 2009, a test was conducted at the Outdoor Test 

Stand at the Naval Air Engineering Station at Lakehurst, New 

Jersey, USA, in order to determine the extent to which 

chevrons can reduce the jet noise generated by a F404-GEF400 engine. The F404 is a low-bypass, high specific thrust 

engine which powers the F/A-18 aircraft. The primary 

objective of this test was to obtain both near-field and far-field 

acoustic data which could be used to determine the extent to 

which the chevrons reduce the noise produced by the engine 

relative to a baseline (no chevron) configuration, especially 

when the engine was operated with afterburners on. This test 

represents the first known use of chevrons on a military engine 

which employed the use of afterburners. The acoustic data 

obtained during this test indicate that the chevrons were 

effective at reducing the jet noise in the peak noise direction 

when the engine was operating with afterburners on. Details 

regarding these near-field and far-field acoustic results are 

presented by Martens, et al. (Ref. 7).
NASA/TM—2010-216366 2

A secondary test objective was to obtain phased array 

source localization data on the noise sources within the jet 

plume and to determine the effect that chevrons have on these 

noise sources. Such data could help guide the development of 

future noise reduction technologies for military engine 

applications. The test was also conducted to evaluate the 

suitability of a commercially-available OptiNav Inc. Array48 

phased array system for this large-sized, high-sound-level 

application and to determine what changes to the measurement 

system would be required in order to obtain more useful 

acoustic source localization data on military engines in the 

future. The purpose of this paper is to present phased array 

data acquired during this full-scale engine test that 1) provide 

information regarding the location of the noise sources within 

the jet plume for both chevron and no chevron configurations, 

and 2) demonstrate the suitability of the Array48 phased array 

system for this application. In addition to these full-scale data, 

an extensive set of noise source localization data obtained with 

the Array48 system on a 1/4th scale model of an F400 series 

nozzle are also provided. These model scale data provide 

useful insights regarding the full-scale engine jet noise source 

mechanisms, and document some of the differences associated 

with testing at model-scale versus full-scale. The data obtained 

during these tests supports milestones of the NASA 

Fundamental Aeronautics Program, Supersonics Project to 1) 

demonstrate that phased arrays can locate noise sources in 3D 

shocked jets, 2) establish a supersonic aeroacoustic database 

for 3D dual-flow supersonic jets, and 3) deliver noise data 

from military engines to NASA partners.

Nomenclature

alocal local speed of sound

aamb ambient speed of sound

b beamforming result

BBSN broadband shock noise

c the speed of sound

C cross spectral matrix

g steering vector

IRP intermediate rated (max. nonafterburning) power 

Ma acoustic Mach number = Ujet/aamb

Mjet jet Mach number = Ujet/alocal

PLA power lever angle, deg

Pamb ambient pressure

Pt bypass bypass total pressure

Pt core core total pressure

Tamb ambient temperature

Tt bypass bypass total temperature

Tt core core total temperature

Ts jet static temperature

Ujet jet velocity

x

 grid point location

y

 microphone location

ω angular frequency

w normalized steering vector

Research Instrumentation

Test Hardware

The F404-GE-F400 is a two-spool, low-bypass axial flow 

turbofan which employs afterburners for thrust augmentation

(Ref. 9). It contains a three-stage fan driven by a single-stage 

low pressure turbine and a seven-stage high-pressure 

compressor driven by a single-stage, high-pressure turbine. 

Approximately 1/4th of the air passing through the fan stage is 

bypassed to the afterburner for combustion and cooling. The 

area of the exhaust nozzle is controlled using 12 sectored 

hinge seals and 12 forward primary flap hinges. The nozzleexit to nozzle-throat area ratio is a function of the power 

setting, and is not set to provide ideal expansion at most 

operating conditions. More information on the F404 engine is 

provided by Walton, et al. (Ref. 9).

The F404 engine is controlled by adjusting a single throttle 

control, the power level angle (PLA). A PLA setting of 35° 

corresponds to engine idle, while 102° represents the 

maximum nonafterburning power setting, referred to as the 

intermediate rated power (IRP). The maximum afterburning 

power setting occurs at PLA = 130°. This report will present 

source localization results for PLA settings of 83°, 102°, and 

130°. Changing the power lever angle results in changes to the 

exhaust-nozzle area, turbine speeds, internal temperatures, and 

fuel flow rate. 

The F404 engine test was conducted at the Outdoor Test 

Stand at the Naval Air Engineering Station at Lakehurst, New 

Jersey, USA. The engine was mounted approximately 2.9 m 

above the surrounding concrete ground plane. A photo of the 

engine on the test stand is shown in Figure 1.

The test involving the 1/4th scale model of the F400 series 

nozzle took place at the Nozzle Acoustic Test Rig (NATR) at 

the NASA John H. Glenn Research Center (GRC) in 

Cleveland, Ohio, USA. The NATR is a 1.35 m diameter freejet used to simulate forward flight conditions during the nozzle 

tests. It is located within the Aeroacoustic Propulsion 

Laboratory (AAPL), a 19.8 m radius anechoic geodesic dome. 

The interior surface of the dome is covered with sound 

absorbing acoustic wedges. Nozzles tested at NATR are 

mounted on the aft end of the High Flow Jet Exit Rig (HFJER) 

that is located at the exit of the NATR duct. Heated air is 

delivered from the facility compressed air system to the test 

nozzle via the HFJER. A dual flow capability allows both the 

primary and bypass streams of the full-scale engine to be 

simulated at model-scale. More information on the NATR is 

available in Soeder, et al. (Ref. 12).

The 1/4th scale baseline nozzle used during the test is shown 

mounted to the High Flow Jet Exit Rig in Figure 2. Although 

this is a fixed-area nozzle, it is similar to the full-scale nozzle 

in that: 1) the convergent-divergent section is made up of two 

conic sections, and 2) the internal surface is faceted to 

simulate the flaps and seals which provide the area change 

capability in the full-scale nozzle. The nozzle exit and nozzle 

throat diameters are fixed at 12.87 and 11.43 cm, respectively, 
NASA/TM—2010-216366 3

and the design Mach number is 1.65. Flow conditions 

corresponding to test points for which data are presented later 

in this report are presented in Table 1. Values are provided in 

the table for the core and bypass flows and also for the mixed 

stream in the exhaust jet downstream of the nozzle exit. The 

mixed stream conditions provided in columns 2 through 4 

were determined assuming that the core and bypass flows mix 

completely. The results presented later in this report will be 

identified by these mixed stream conditions. Note that none of 

the hot conditions listed in this table mimic any of the engine 

operating conditions set during the full-scale test. At the 

corresponding Mach numbers, the full-scale jet flow was 

always hotter than that produced at model-scale.

The chevrons used during the model-scale test (referred to 

in other publications as the P03L08W10 chevrons) were 

mounted at the nozzle exit as shown in Figure 3. Twelve 

chevrons were used, one on each facet. As shown in Figure 4 

the chevron penetration was defined as the distance from the 

line extended along the inner conical surface to the chevron 

tip, the length was defined at the curved distance along the 

chevron surface from the nozzle trailing edge to the chevron 

tip, and the width was given by the width of the chevron base 

relative to the nozzle facet width. The chevrons used during 

the model-scale test had a penetration of 7.62 mm, a length of

19.1 mm, and a width of 100 percent.

The Phased Array System

Figure 5 shows both front and rear views of the OptiNav 

Array48 phased array system used during this test. This system 

consists of 48 Earthworks M30 microphones flush-mounted to a 

1 by 1 m aluminum plate. The microphones are arranged in a 

series of log spirals in an effort to reduce sidelobes (errors in the 

phased array data). The microphones have 1/4th in. diameter 

diaphragms and a flat frequency response over a frequency 

range of 5 Hz to 30 kHz. They can be used to measure sound 

pressure levels as high as 142 dB before they begin to saturate 

(136 dB when flush mounted in the array plate).

The photo at the right in Figure 5 shows the microphones 

mounted to the back of the array plate and a camera located at 

the center of the plate. There is a hole in the center of the plate 

through which the camera can be used to take a photo of the 

“field of view” of the phased array system. The phased array 

data reduction software superimposes the acoustic source 

localization data on top of the image taken with the phased 

array camera. 

The phased array data were reduced using classical 

beamforming in the frequency domain. Classical beamforming 

assumes that the noise sources are incoherent, stationary point 

sources. The first step in the data reduction process was to 

compute the cross spectral matrix, C, from the array data using 

the periodogram method with Hann windowing functions and 

50 percent block overlap. The diagonal elements of the cross 

spectral matrix are then deleted (yielding C ), and the 

beamforming result, b, at a given grid point, k, is computed 

using the classical beamforming expression

wkCwk = ′ k b

where wk is a normalized version of the steering vector gk. The 

individual elements of gk represent the Green’s function for a 

monopole located at grid point k as observed by microphone i. 

In free space with no flow

i k

y x cj

ki

y x

e

g

i k

 

 

− =

− ω

where k x

 and i y denote the locations of grid point k and 

microphone i, respectively. A beamforming result is computed 

for each point in a beamforming grid. The results 

corresponding to an entire grid are then displayed as color 

contour maps (referred to as beamform maps).

The output of the phased array data reduction routine is a 

stack of beamform maps. Each beamform map in the stack 

corresponds to a different frequency band. The maps are 

arranged in the stack such that the lowest frequency band is at 

the top of the stack, and the highest frequency band is at the 

bottom. Each beamform map shows the location of the 

dominant noise source or sources in the corresponding

frequency band as a 2D color contour map overlaid on top of a 

photograph taken with the phased array camera. The color 

contours correspond to the location and strength of the noise 

sources found within an image plane (the beamform grid) at 

some specified location away from the microphone array. The 

user selects the location of the image plane by specifying a 

microphone-array-to-image-plane distance. Usually this image 

plane is parallel to the microphone array and usually its 

location is selected to coincide with a region where dominant 

noise sources are suspected, such as along the centerline of a 

jet flow. The dynamic range of the color contours (max value 

minus min value) is the same for each beamform map in the 

stack, but the peak value varies from map to map. The number 

of maps in the stack and the bandwidth corresponding to each 

map depends on how the data are processed. The data can be 

processed into 1/12th octave, 1/3th octave, octave, narrowband, 

or kilohertz bands or into one OASPL band.

Phased Array Deployment

During the full-scale test, data were acquired with the 

microphone array at four different locations relative to the 

engine. These four locations are depicted in the schematic 

diagrams of Figure 6 and in the photographs of Figure 7. As

indicated in these figures, three of these positions are referred 

to as “broadside” positions; the fourth is a “downstream” 

position. 
NASA/TM—2010-216366 4

In the three broadside positions the array plate was parallel 

to the jet centerline. In the “broadside/far/low” position the 

array was located 11.28 m (approx. 22 diameters) from the jet 

centerline; whereas in the “broadside/close/low” and 

“broadside/close/high” positions the array was located 8.1 m 

(approx. 15.5 diameters) from the jet centerline. In the two 

“low” positions the array center was 1.35 m above the ground; 

while in the “high” position it was 2.52 m above the ground. 

In each of the three broadside positions the array was located 

such that a normal to the center of the array intersected the jet 

axis 2.74 m (approx. 5.4 nozzle diameters) downstream of the 

nozzle exit.

In the downstream position, the array was rotated so that it 

could sense the jet from a downstream location. As in the 

broadside locations, the array was located such that a normal 

to the array intersected the jet centerline 2.74 m (approx. 5.4 

nozzle diameters) downstream of the nozzle exit. The angle 

between the normal to the array plate and the upstream jet 

centerline was 126°. As such, the array was close to the peak 

directivity direction. Post-test analysis revealed that the 

microphone signals were saturated at all but the lowest power 

engine operating conditions when the array was at this 

downstream position. Consequently, data obtained with the 

array at this position will not be presented in this report.

During the model scale test the microphone array was 

positioned to maintain geometric similarity with the 

broadside/close/high position of the full-scale test. The array 

was located 15.5 diameters (2 m at model scale) from the jet 

centerline and a normal to the array center intersected the jet 

centerline 5.4 diameters (0.69 m at model scale) downstream 

of the nozzle exit. During the model-scale test the array was 

positioned so that the center of the array was at the same 

height as the jet centerline, 3.05 m above the floor. Figure 8 

shows both a schematic representation and a photograph of 

this arrangement. Since the nozzle diameter of the model was 

1/4th the size of that of the full-scale engine, maintaining 

geometric similarity of the array/nozzle position meant that 

the dimensions of the array were effectively four times bigger 

relative to the size of the jet during the model scale test vis-àvis the engine test. This relative size difference can be seen in 

the figures by comparing the size of the array and the nozzle 

in the broadside/close position depicted in Figure 6 with the 

array and nozzle size shown Figure 8. Due to this larger 

effective size, the array was much better suited for identifying 

details in the jet plume during the model-scale test. 

Another significant difference between the two tests is that 

sound-absorbing wedges were located on the ground between 

the array and the jet during the model-scale test, but not during 

the full-scale test. The lack of wedges in the engine test 

proved detrimental to the full-scale phased array results, 

especially to those obtained at low frequency. Some of the 

spectra generated from the full-scale phased array results show 

low frequency ripples caused by constructive and destructive 

interference between sound waves coming directly to the array 

from the jet plume with sound waves reflecting off of the hard 

ground plane. The ground plane reflection was also 

responsible for a vertical displacement of the beamform maps 

about the jet centerline that varied with frequency. Both of 

these effects – the ripples in the spectra and the vertical 

displacement of the beamform maps – were less pronounced 

in the data obtained with the array at the broadside/close/high 

position. Despite these ground reflections, all three broadside 

positions provided essentially identical results regarding how 

the axial location of the peak jet noise source varied with 

frequency at a given engine operating condition. 

Consequently, only full-scale data obtained with the array at 

the broadside/close/high position will be presented in this 

report.

Unfortunately, the microphone signals obtained with the 

array at the broadside/close/high position exhibited some 

evidence of saturation at PLA settings of 102° and higher. 

This microphone saturation, however, is not thought to have 

had a significant impact on the ability of the phased array 

system to determine the axial location of the sources within 

the jet plumes. This is based on the observation that the noise 

source locations changed very little between the 

broadside/close and the broadside/far array positions even 

though the degree to which the microphones were saturated 

did change between these two positions. 

Results

Full-Scale Results

Figure 9 shows baseline (no chevrons) data obtained with 

the phased array at the broadside/close/high position at a 

power lever angle setting of 83°. The data were processed in 

such a way that the data reduction routine searched for noise 

sources in a plane which was parallel to the array plate and 

which passed through the jet axis. The data were reduced into 

1/12th octave bands over a frequency range from 300 to

20,000 Hz. The complete stack of beamform maps 

corresponding to this frequency range contained 73 maps. Six 

beamform maps selected from the complete stack are shown at 

the right in Figure 9. Beamform maps are shown for 1/12th

octave bands centered at 500, 1000, 1600, 2500, 5000, and 

10000 Hz. The center frequency is shown in the upper left 

corner of each map. The dynamic range of the contours 

provided in these and in all the other beamform maps provided 

in this report is 7 dB, i.e., the peak contour level is 7 dB higher 

than the level corresponding to where the contours fade into 

the background. The plot in the upper left of the figure shows 

spectra computed from the microphone in the array which 

measured the median sound level of all the microphones in the 

array. This median microphone spectrum does not show 

absolute sound levels along the ordinate. However, the 

horizontal grid lines in this plot are separated by 5 dB. The 

plot at the lower left in the figure shows the distance between 

the peak in the beamform maps and the nozzle exit as a 

function of frequency, in units of nozzle diameters. This latter 

line plot will be referred to herein as a peak location plot. Red 

asterisks have been superimposed on each of these two line 
NASA/TM—2010-216366 5

plots at the frequencies represented by the six beamform maps 

at the right in the figure. 

The microphone spectrum provided in Figure 9 indicates 

that at this engine power setting the dominant noise source 

occurred at about 2300 Hz, and the peak location plot shows 

that this source was located between 1 and 2 nozzle diameters 

downstream of the nozzle exit. The peak location plot also 

indicates that the peak source is located about 5 nozzle 

diameters downstream of the nozzle exit at the lowest 

processed frequency (300 Hz), and that the peak source 

continuously moves closer to the nozzle exit with increasing 

frequency. For frequencies above about 1500 Hz, the peak 

source location is always within two nozzle diameters of the 

nozzle exit. At this engine operating condition the dominant 

noise sources do not appear to extend significantly in the axial 

direction. Instead, the noise sources associated with each 

frequency band tend to be centered about one axial location in 

the jet plume, although this location does vary with 

frequency– it gets closer to the nozzle exit with increasing 

frequency.

Figure 10 shows baseline data processed in the same 

manner as that provided in Figure 9, but for a higher power 

lever angle setting of 102°. The PLA = 102° results presented 

in Figure 10 are very similar to the PLA = 83° results of 

Figure 9. The peak location plots for these two operating 

conditions are almost identical, indicating virtually no change 

in how the peak noise location varies with frequency as the 

engine power is increased between these two operating 

conditions. In order to provide some indication of the relative 

strength of the noise sources in the beamform maps, a median 

microphone spectrum is presented in Figure 10, but the plotted 

levels are thought to be biased low due to the aforementioned 

microphone saturation which occurred at PLA settings of 102°

and higher. The main difference between the PLA = 83° and 

PLA = 102° results is that the peak in the spectrum shifts to a 

lower frequency for the higher power condition (from about 

2300 to 1500 Hz). Another difference is that the 1600 Hz 

beamform map for the higher power condition shows some 

evidence of noise coming from both an upstream location 

(centered about 1.5 nozzle diameters downstream of the 

nozzle exit) and a downstream location (about 5.5 nozzle 

diameters downstream of the nozzle). The noise generated 

downstream in the plume is thought to be broadband shock 

noise. At 1600 Hz this broadband shock noise source is about 

3 dB lower than the dominant source closer to the nozzle exit.

The PLA = 83° and the PLA = 102° baseline results 

presented in Figures 9 and 10 were acquired with the engine 

operating with afterburners off. The results presented in

Figure 11 were acquired with the engine operating at the 

maximum afterburner condition, PLA = 130°. A comparison 

of the max afterburner results of Figure 11 with the no 

afterburner results of Figure 10 indicates very little change in 

how the peak source location varies with frequency between 

these two conditions. The peak location plot still shows the 

peak noise source moving toward the nozzle exit as frequency 

increases, and for frequencies above about 1500 Hz the peak 

source is still always within two diameters of the nozzle exit. 

The next three figures illustrate how the noise source 

location results discussed above for the baseline nozzle 

changed when chevrons were added to the nozzle. Figures 12, 

13, and 14 show a comparison between baseline and chevron 

data acquired at PLA settings of 83°, 102°, and 130°, 

respectively. Spectra are only provided for the 83° PLA 

setting since the spectra corresponding to the two higher 

power settings are thought to be biased due to microphone 

saturation. The spectra provided for the 83° setting indicate a 

modest broadband reduction (<1 dB) in the output of the 

median microphone when chevrons were added to the nozzle 

as measured at this broadside array location. The peak location 

plots indicate that for frequencies below 1500 Hz the chevrons 

tend to move the peak source location downstream slightly 

when the engine was operated at the two lower power settings, 

but show very little change in the noise source location at max 

power. Overall, these data indicate that the chevrons had a 

slight impact on the noise source locations when the engine 

was operated with the afterburners off, but almost no change 

when the afterburners were on.

The baseline results presented in this section for the three 

different engine operating conditions indicate that at low 

frequency (300 Hz) the peak noise source was located about 

five nozzle diameters downstream of the nozzle exit and that it 

moved upstream toward the nozzle as frequency increased. 

The short distance between the peak source location and the 

nozzle exit measured here agrees with the full-scale military 

engine results of Schlinker et al. (Ref. 6), although they found 

that the peak source remained about 4 to 5 nozzle diameters 

downstream of the nozzle exit irrespective of frequency. 

Despite this agreement with the other full-scale engine test, 

the results presented above were unexpected based on 

previous model-scale supersonic jet experiments conducted 

with the Array48 system at the Small Hot Jet Acoustic Rig 

(SHJAR) at NASA GRC. Some results from these model-scale 

tests have been presented previously (Ref. 8). The vast 

majority of these previous studies involved hot and cold 

subsonic jets and cold supersonic jets. The data obtained 

during these studies on cold, shock-containing jets tend to 

exhibit much more character than the full-scale data presented 

above. In particular, they often show a more complicated 

behavior regarding how the peak source location in the jet 

plume varies with frequency. 

Model-Scale Results

In order to better understand the origin of the differences 

between the full-scale data presented above and the previous 

model-scale data, another model-scale test was conducted 

after the completion of the full-scale test. During this test the 

Array48 system was used to obtain data on the jet flow 

produced by a 1/4th scale model of the F400 series nozzle at 

the Nozzle Acoustic Test Rig (NATR) at NASA GRC. This 
NASA/TM—2010-216366 6

nozzle was tested in both baseline (no chevrons) and chevron 

configurations, and the jet was operated at both cold and hot 

conditions. Table 1 lists the operating conditions set during 

this test. The hot conditions listed in this table do not mimic 

any of the engine operating conditions set during the full-scale 

test. At the corresponding Mach numbers, the full-scale jet 

flow was always hotter than that produced at model-scale.

As indicated in Table 1, the subscale nozzle was only tested 

at supersonic conditions. In order to illustrate the differences 

between a subsonic and a shock-containing jet, Figures 15 and 

16 show data obtained during an earlier test at SHJAR 

involving the Array48 system (Ref. 8). These results were 

acquired on a 50.8 mm diameter, round convergent (RC) 

nozzle (referred to in previous publications as the smc000 

nozzle). The array plate was parallel to and 2 m from the jet 

axis and the array was centered 10 nozzle diameters 

downstream of the nozzle exit. The data were processed into 

1/12th octave bands over a frequency range from 300 to

45000 Hz, yielding stacks containing 87 beamform maps.

Figure 15 shows seven beamform maps selected from a 

complete stack obtained with the jet operating at Mjet = 0.89, 

Ma = 1.46 and Ts/Tamb = 2.7 (referred to in other publications 

as set point 49). At this subsonic jet Mach number the jet noise 

is produced solely by the turbulent mixing occurring between 

the jet and the ambient air. The spectrum at the upper left in 

the figure shows the broad, haystack-like character expected 

for turbulent mixing noise in a subsonic jet. Both the peak 

location plot in the lower left and the beamform maps at the 

right show the peak noise source moving steadily toward the 

nozzle exit with increasing frequency. Also notable is that the 

jet shear layer is visible in the two beamform maps shown at 

the lower right in the figure; these correspond to the 32 and

40 kHz bands, respectively.

Figure 16 shows 1/12th octave band beamform maps 

computed from data obtained with this RC nozzle generating 

an underexpanded, cold supersonic jet with Mjet = 1.4, Ma = 

1.4, and Ts/Tamb = 1.0 (set point 9020). The spectrum and the 

peak location plots provided at the left show distributions for 

both this supersonic condition (in black) and for the subsonic, 

Mjet = 0.9, Ma = 1.46, and Ts/Tamb = 2.7 condition (in blue) 

which was presented above in Figure 15. The 10 points 

labeled along the black lines correspond to the ten beamform 

maps provided at the right. The small red squares overlaid on 

top of the beamform map contours show the approximate axial 

location of the peak noise source in the jet plume. At low 

frequency (<2800 Hz, near point 1) the spectrum for the 

supersonic jet is very similar to that of the subsonic jet, 

indicating that turbulent mixing noise is dominant in this 

frequency range. A fundamental screech tone occurs at

2800 Hz (point 2) and appears in the beamform maps 

predominantly as a reflection off of the upstream nozzle 

hardware. At 5600 Hz (point 4) the first harmonic of the 

fundamental screech tone comes from a location in the jet 

plume about 5.5 diameters downstream of the nozzle exit.

Broadband shock noise (BBSN) shows up clearly in the 

supersonic spectra shown in Figure 16 as the elevated region 

above 6300 Hz (to the right of point 5). A dominant BBSN 

hump is visible in the spectrum between 6300 and 11000 Hz 

(between points 5 and 8). BBSN humps such as this are 

thought to be created by large scale turbulent structures 

passing through the shocks. Note that the location of the peak 

noise source gradually moves downstream as frequency 

increases through the BBSN hump (between points 5 and 8). 

This is visible in both the peak location plot and in the 

beamform maps. Once the frequency is increased beyond the 

BBSN hump (between points 8 and 9), the peak location 

jumps back upstream. In a cold supersonic jet emanating from 

a nozzle which does not contain chevrons, the peak location 

tends to remain near this upstream location over a wide 

frequency range (to the right of point 9), unless this region 

contains another BBSN hump. The peak location plot 

presented in Figure 16 for the supersonic jet suggests that 

there may be a second BBSN hump at about 15000 Hz.

The BBSN data presented in Figure 16 indicate that there is 

a distinct difference in behavior regarding how the location of 

the peak noise source varies with increasing frequency 

depending on whether or not this variation occurs through a 

BBSN hump. The peak source location moves downstream 

when the frequency is increased through a hump; it does not 

move downstream when the frequency increase occurs in the 

elevated region of BBSN outside the humps. This difference 

in behavior suggests that there may be two different 

mechanisms creating the elevated region of BBSN. The BBSN 

humps are thought to be created by turbulent eddies which are 

large enough to span more than one shock cell. The interaction 

of these large eddies with the shocks generates highly 

correlated noise radiating from multiple shocks 

simultaneously. The BBSN associated with the regions outside 

the humps appears to be generated by turbulent eddies which 

are smaller than the shock spacing. This turbulence/shock cell 

interaction produces uncorrelated noise radiating from the 

individual shock cells. The beamform map at the lower right 

of Figure 16 (#10), shows high frequency noise coming from 

individual shock cells in the jet plume.

The remainder of this section will focus on results obtained 

from the test of the 1/4th scale model of the F400 nozzle that 

was conducted at GRC after the completion of the full-scale 

engine test. Figure 17 shows source location data obtained 

from two cold supersonic jets produced by this nozzle. Data 

are provided for two conditions, Mjet = 1.22, Ts/Tamb = 0.77 (in 

blue on the line plots and in the left column of beamform 

maps) and Mjet = 1.47, Ts/Tamb = 0.73 (in black on the line 

plots and in the right column of beamform maps). Both of 

these represent overexpanded conditions. As might be 

expected, the peak location plots show that the dominant low 

frequency turbulent mixing noise is produced further 

downstream in the higher speed jet vis-à-vis the lower speed 

jet (points 1 to 3). Like the supersonic data shown above for 

the RC nozzle, each of the microphone spectra show a 

dominant BBSN hump, and the corresponding line plots and 

beamform maps show the peak source moving downstream as 

the frequency increases through the humps (points 7 to 9 in the 
NASA/TM—2010-216366 7

Mach 1.22 jet, and points 5 to 7 in the Mach 1.47 jet). Also 

like the RC nozzle data, the peak source location in each jet 

jumps back upstream when the frequency is increased beyond 

the hump. At frequencies higher than the humps the peak 

source location in each jet jumps back and forth between 

individual shock cells, but remains close to the same axial 

location (about 5 nozzle diameters downstream) over a wide 

frequency range.

Figure 18 shows source location results for two Mach 1.22 

jets, one cold (Ts/Tamb = 0.77) and one hot (Ts/Tamb = 2.16). 

This cold condition data was also presented above in

Figure 17. The peak location plots for these two different 

temperatures are nearly identical at low frequency (points 1 to 

6), indicating that the location of the source of low frequency 

turbulent mixing noise changes very little with increasing jet 

temperature. Unlike with the cold jet, however, the spectrum 

for the hot jet does not show a BBSN hump, and neither the 

peak location plot nor the beamform maps provided for the hot 

jet show the peak source location moving downstream with 

increasing frequency over any part of the frequency range. 

Instead, both the spectrum and the peak location plots of the 

hot jet resemble those of a subsonic jet in which only turbulent 

mixing noise is present. These results suggest that either heat 

has been added to the jet to the point that the BBSN hump has 

been masked by turbulent mixing noise, or that the mechanism 

responsible for generating the hump (large scale turbulent 

structures crossing the shocks) has been disrupted. Particle 

Image Velocimetry data obtained on similar jets (cold vs. hot 

shock-containing jets of the same Mach number) indicate that 

the addition of heat has very little effect on shock structure

(Ref. 13). This suggests that the first scenario is more likely –

that the BBSN hump is being masked by turbulent mixing 

noise. However, it is still possible that the added heat altered 

the turbulent structures in such a way that the eddy/shock 

interaction produced less BBSN.

Figure 19 shows the effect of adding heat to the cold

Mach = 1.47 jet produced by the 1/4th scale nozzle. This figure 

shows results for both the cold (Ts/Tamb = 0.73) condition 

presented above in Figure 17 and for a hot condition (Ts/Tamb

= 2.45). Like the results presented above in Figure 18 for the 

lower Mach number, the location of the source of the low 

frequency turbulent mixing noise changes very little with 

increasing jet temperature (indicated by the overlap of the 

peak location plots at points 1 and 3). Unlike the lower Mach 

number results, however, both the cold and hot data obtained 

at this higher Mach number show a dominant BBSN hump 

and the movement of the peak source location downstream as 

frequency increases through the humps. In this case it appears 

that the addition of heat did not increase the turbulent mixing 

noise to the point that it masked the BBSN hump. Despite this 

common feature of the two spectra, these data do show a 

noticeable difference in where the high frequency noise to the 

right of the humps comes from inside the plume. In the cold 

jet it is produced by turbulence passing through shocks located 

relatively far (about 5 nozzle diameters) downstream. In the 

hot jet it is also produced by turbulence passing through 

shocks, but the peak source rapidly approaches the nozzle exit 

as the frequency is increased, jumping from one shock to 

another along the way.

Figure 20 shows the effect of adding chevrons to the 1/4th

scale CD nozzle for a cold (Ts/Tamb = 0.77), overexpanded jet 

at Mach 1.22. This is the same baseline (no chevron) data that 

was shown above in Figures 17 and 18. As shown in

Figure 20, the spectra, the peak location plots and the 

beamform maps for the baseline and chevron cases are almost 

identical, indicating that the chevrons have almost no effect on 

either the noise levels or on the noise source locations at this 

combination of jet operating condition and array position 

(broadside to the jet).

Figure 21 shows another no chevron/chevron comparison 

for two cold jets (Ts/Tamb = 0.70), but at a higher Mach 

number, Mjet = 1.47. A comparison of these results with those 

obtained at the lower Mach number (Mjet = 1.22 shown in 

Figure 20) indicates that the chevrons had a much bigger 

impact on the noise produced by the jet at this higher Mach 

number. The spectra show a low frequency broadband noise 

benefit associated with using the chevrons and a reduction in 

amplitude of the screech tones, but the peak of the BBSN 

hump increases in amplitude and shifts to a higher frequency 

when the chevrons are added. The source location data show 

the peak source location slightly further downstream (11 

nozzle diameters versus 10) at the high frequency edge of the 

hump (just to the right of point 7 in the line plots). There is 

also a noticeable difference regarding where the high 

frequency noise just to the right of the humps comes from in 

the plume (points 8 and 9). The peak source of this noise is 

located much further downstream in the chevron case (8 to 9 

nozzle diameters downstream vs. 5). The beamform maps 

shown at the bottom of the Figures 9 to 11 indicate that the 

high frequency BBSN is more evenly distributed across the 

shocks in the jet plume in the chevron case, despite the fact 

that the noise levels are almost identical in the two cases 

(points 8 to 11 in the spectra).

Figure 22 shows a chevron/no chevron comparison for two 

hot jets (Ts/Ta = 1.86) at Mach 1.22. Like the results presented 

in Figure 20 which showed a no chevron/chevron comparison 

at this jet Mach number at cold conditions, the Figure 22 

results show that the chevrons had very little effect on the 

noise produced by the jet as measured with the array at this 

broadside location. Overall, the spectrum is slightly lower 

(about 1 dB) for the chevron case, but there is almost no shift 

in where the noise sources are located in the jet plume.

Figure 23 shows the effect of adding chevrons to a hot jet 

(Ts/Ta = 1.80) at a higher Mach number (Mjet = 1.47). Like the 

cold jet results presented above, a comparison of the data 

obtained at the two Mach numbers (Figs. 22 and 23) indicates 

that the chevrons have a bigger impact on the peak noise 

source locations at the higher Mach number. As shown in 

Figure 23 both the baseline and no chevron spectra show what 

appear to be small BBSN humps, but only the baseline source 

location data show the peak noise source location moving 

downstream as frequency increases through the hump. The 
NASA/TM—2010-216366 8

beamform maps provided for the baseline nozzle indicate that 

the BBSN hump is created by noise generated between 7 and 

10 nozzle diameters downstream of the nozzle exit (points 5 to 

8). The corresponding maps for the chevron nozzle suggest 

that the BBSN generated in this part of the jet is being masked 

by a stronger source close to the nozzle exit. This upstream 

source may be stronger in the chevron case due to turbulence 

introduced into the flow just downstream of the nozzle exit by 

the chevrons. This added turbulence also seems to be 

responsible for the increased levels of high frequency noise

evident in the microphone spectra computed from the data 

obtained with the chevrons installed (points 9 to 11). 

Interpretation of the Full-Scale Results Based on 

the Model-Scale Test

As mentioned above, the test involving the 1/4th scale 

model of the F400 nozzle was conducted in order to determine 

the origin of the differences between the full-scale results and 

previous model-scale results obtained on cold, shockcontaining jets. The spectra of cold, shock-containing jets tend 

to show separate regions dominated by turbulent mixing noise, 

screech and broadband shock noise, and the corresponding 

source location data tend to show a rather complicated 

behavior in which the peak source location moves upstream, 

then downstream, then upstream again as frequency increases. 

In contrast, the full-scale results show little evidence of shock 

noise and the peak source location always moves upstream 

toward the nozzle exit as frequency increases. Instead of 

looking like the results of a cold, shock-containing jet, the 

full-scale results tend to resemble a simple subsonic jet in 

which only turbulent mixing noise is present. 

The results presented above from the test at the NATR 

involving the 1/4th scale nozzle provide some insights 

regarding why the full-scale results have this simple character. 

These model-scale results show that it is possible to start with 

a cold shock-containing jet and add heat and turbulence until 

the resulting spectra and source location data begin to 

resemble those of the full-scale engine. In particular, the hot, 

high Mach number, chevron results presented in Figure 23 are 

quite similar to the full-scale data. They show the downstream 

sources of BBSN (which must also be present in the full-scale 

overexpanded jets) masked by a stronger source closer to the 

nozzle exit. At model scale it appears that this upstream source 

is dominant because 1) the added heat increases the jet 

velocity which leads to an increase in the jet turbulence kinetic 

energy, which, in turn, tends to increase the turbulent mixing 

noise, and 2) the chevrons add turbulence to the flow just 

downstream of the nozzle exit. At full-scale the upstream

source dominates even when chevrons are not installed on the 

nozzle. As mentioned above, at similar Mach numbers the 

full-scale jets were always hotter than those tested at modelscale. Apparently, the upstream source dominates in the fullscale baseline application because the flow coming out of the 

engine is hotter and, hence, has a higher velocity and a higher 

turbulence kinetic energy than the flows tested at model scale. 

Recommendations

The test on the F404 engine provided some valuable 

insights regarding how future source localization tests of fullscale military engines could be improved. The following 

recommendations should be implemented before another fullscale test of this type is conducted: 

1) Efforts should be made to reduce the adverse effects of 

the ground reflections. This can be achieved by installing 

sound absorbing acoustic wedges between the jet plume 

and the array or by positioning the array on the ground 

either under or off to the side of the jet.

2) A bigger array should be used in order to improve the 

ability of the array to identify details in the jet plume. At 

the broadside/close/high position for which most of the 

full-scale engine data were acquired, the Array48 system 

did not offer sufficient spatial resolution to identify the 

noise produced by individual shock cells. Assuming a 

new array placed at the same location (15.5 nozzle 

diameters away from the jet centerline), its dimensions 

would have to be four times greater than those of the 

Array48 system in order to provide the same spatial 

resolution as that offered by the experimental setup used 

during the 1/4th scale model test.

3) The array microphones should be replaced with less 

sensitive piezoresistive transducers that will not saturate 

when the array is located close to the jet plume and the 

engine is operating at high power conditions.

4) The solid array plate should be replaced with an open 

framework of thin, connected beams which would hold 

the transducers firmly in place. This open framework will 

make the array less susceptible to vibration due to the 

high acoustic loads encountered when the array is located 

near the jet.

A final recommendation concerns the need to document the 

characteristics of the turbulence produced by full-scale 

military engines. The results presented above suggest that the 

flow exiting the full-scale nozzle was more turbulent than the 

model-scale flow. Higher turbulence levels could be expected 

due to the higher jet temperatures at full-scale. It is also 

possible that the full-scale jet flow was more turbulent due to 

the more complicated flow path inside the engine. In order to 

develop noise reduction technologies for full-scale engines 

during model-scale tests it may be necessary for the model 

scale flow characteristics to more closely match those of the 

full-scale engine. At present, however, the characteristics of 

the turbulent flow exiting the full-scale engine are unknown. 

Therefore, they should be measured using a technique such as 

Particle Image Velocimetry.
NASA/TM—2010-216366 9

Summary

1) Data were acquired with the Array48 phased array system 

on a F404 engine over a wide range of operating 

conditions, including with afterburners on. The full-scale 

engine source localization results tend to have a simple 

character. The data obtained at the different operating 

conditions all show the peak source located about five 

nozzle diameters downstream of the nozzle exit at low 

frequency (300 Hz), and moving continuously toward the 

nozzle exit as frequency increases. The full-scale engine 

data obtained with the phased array show little evidence 

of broadband shock noise.

2) A comparison of full-scale engine data obtained with and 

without chevrons installed on the nozzle indicates that for 

frequencies below 1500 Hz the chevrons moved the peak 

source location slightly further downstream when the 

engine was operated at the two lower power settings, but 

had very little effect at full power. The data obtained with 

the array broadside to the jet also indicate that the 

chevrons reduced the broadband jet noise slightly

(<1 dB) at the PLA = 83° setting. The noise reduction 

potential of the chevrons could not be determined using 

the array microphones at the two higher power settings 

due to microphone saturation.

3) The model-scale results tend to exhibit much more 

character than the full-scale results. The spectra of cold, 

shock-containing jets tend to show separate regions 

dominated by turbulent mixing noise, screech and 

broadband shock noise, and the corresponding source 

location data tend to show a rather complicated behavior 

in which the peak source location moves upstream, then 

downstream, then upstream again as frequency increases. 

4) The model-scale results indicate a distinct difference in 

behavior regarding how the location of the peak source 

varies as frequency increases depending on whether or not 

the frequency variation occurs within a BBSN hump. The 

peak source location tends to move downstream when the 

frequency increase occurs through a BBSN hump. It does 

not move downstream if the frequency increase is not 

through a hump. This difference in behavior suggests that 

there may be two different mechanisms responsible for 

producing the BBSN.

5) The model-scale results obtained on the 1/4th scale model 

F400 nozzle show that it is possible to start with a cold, 

shock-containing jet and add heat and turbulence until the 

resulting spectra and source location data begin to 

resemble those of the full-scale engine. This suggests that 

the differences between the full-scale and previous model 

scale results are due to the hotter nozzle exit flows in the 

full-scale test.

6) The inability of the model-scale tests to mimic the very 

hot flows of the full-scale nozzle flow may be hindering 

efforts to develop noise reduction concepts for military 

engines at model scale.

References

1. “Acquisition Safety—Noise Control Aboard Navy 

Ships,” Naval Safety Center, 26 August 2009. Web. 10 

November 2009.

2. Humes, L.E., Joellenbeck, L.M., Durch, J.S., (ed.) Noise 

and Hearing Loss: Implications for Hearing Loss and 

Tinnitus, Washington, DC. National Academies Press, 

2006.

3. Shaw, G.B., and Trost, R.P., “Statistical Analysis of 

Hearing Loss Among Navy Personnel,” CRM 

D0011228.A2/Final, February 2005.

4. Norum, T.D., Garber, D.P., Golub, R.A., Santa Maria, 

O.L., Orme, J.S., “Supersonic Jet Exhaust Noise at High 

Subsonic Flight Speed,” NASA/TP—2004-212686, 

January 2004.

5. Gee, K.L., Gabrielson, T.B., Atchley, A.A., Sparrow, 

V.W., “Preliminary Analysis of Nonlinearity in Military 

Jet Aircraft Noise Propagation,” AIAA Journal, Vol. 43, 

No. 6, June 2005.

6. Schlinker, R.H., Liljenberg, S.A., Polak, D.R., Post, K.A., 

Chipman, C.T., Stern, A.M., “Supersonic Jet Noise Source 

Characteristics & Propagation: Engine and Model Scale,” 

AIAA paper 2007-3623, May 2007.

7. Martens, S., Spyropoulos, J.T., “Practical Jet Noise 

Reduction for Tactical Aircraft,” ASME paper GT2010-

23699, June 2010.

8. Dougherty, R.P., Podboy, G.G. “Improved Phased Array 

Imaging of a Model Jet,” AIAA paper 2009-3186, May 

2008.

9. Walton, J.T., Burcham, F.W., “Exhaust-Gas Pressure and 

Temperature Survey of F404-GE-400 Turbofan Engine,” 

NASA TM-88273, December 1986.

10. Huber, J., Vincent Fleury, J.B., Britchford, K., 

Laurendeau, E., Long, D., “Understanding and Reduction 

of Cruise Jet Noise at Model and Full Scale,” AIAA paper 

2009-3382, May 2009.

11. Mengle, V.G., Ganz, U.W., Nesbitt, E., Bultemeier, E.J., 

Thomas, R.H., “Flight Test Results for Uniquely Tailored 

Propulsion-Airframe Aeroacoustic Chevrons: Shockcell 

Noise,” AIAA paper 2006-2439, May 2006.

12. Soeder, R.H., Wnuk, S.P., Loew, R.A., “Nozzle Acoustic 

Test Rig User Manual,” NASA/TM—2006-212939, 

November 2006.

13. Bridges, James “Broadband Shock Noise in InternallyMixed Dual-Stream Jets,” AIAA Paper AIAA–2009–

3210, May 2009.
NASA/TM—2010-216366 10

TABLE 1.—FLOW CONDITIONS DURING THE 1/4TH SCALE MODEL TEST

Configuration Mjet Ts/Tamb Ma Pt core/Pamb Pt bypass/Pamb Tt core/Tamb Tt bypass/Tamb

Baseline 1.22 0.77 1.07 2.50 2.51 0.99 1.00

Baseline 1.47 0.73 1.23 3.52 3.52 1.00 0.99

Baseline 1.23 1.86 1.67 2.50 2.50 2.80 1.04

Baseline 1.48 1.80 1.97 3.51 3.51 3.00 1.05

Chevron 1.22 0.79 1.09 2.50 2.50 1.03 1.00

Chevron 1.47 0.70 1.23 3.50 3.59 1.01 1.00

Chevron 1.23 1.87 1.67 2.49 2.49 2.80 1.05

Chevron 1.48 1.80 1.97 3.51 3.50 3.00 1.05

Figure 1.—Photo of the full-scale F404-GE-F400 engine mounted 

at the Outdoor Test Stand at the Naval Air Engineering Station.
NASA/TM—2010-216366 11

Figure 2.—Photo of the 1/4th scale F400 nozzle mounted to the 

High Flow Jet Exit Rig at the Nozzle Acoustic Test Rig at 

NASA GRC. The nozzle is shown in the baseline (no chevron) 

configuration.

Figure 3.—Photo showing chevrons attached to the 1/4th scale F400 nozzle.
NASA/TM—2010-216366 12

Figure 4.—Schematics illustrating definitions of 

chevron penetration, length and width.

Figure 5.—Photo showing front (left) and rear (right) views of Array48 phased array system.
NASA/TM—2010-216366 13

Figure 6.—Schematic diagrams showing phased array positions during the full-scale engine test. 

2.74 m 2.74 m

2.74 m

8.1 m

11.28 m

14.22 m

126° 

Array48

Array48

Array48
NASA/TM—2010-216366 14

 

 Broadside far low position Broadside close low position

 (array center 1.35 m above the ground) (array center 1.35 m above the ground)

 

 

 Broadside close high position Downstream position

 (array center 2.52 m above the ground)

Figure 7.—Photographs of Array48 at the four test positions during the full-scale engine test. 
NASA/TM—2010-216366 15

Figure 8.—Schematic (top) and photograph (bottom) of microphone array deployment during the modelscale test at the Nozzle Acoustic Test Rig.

0.69 m

2.0 m

Array48
NASA/TM—2010-216366 16

Figure 9.—Median mic spectra (upper left), peak location plot (lower left), and beamform maps for the full-scale 

engine operating at PLA = 83°.

Figure 10.—Median mic spectra (upper left), peak location plot (lower left), and beamform maps for the

full-scale engine operating at PLA = 102° (max non afterburning condition).

5000 Hz 

10000 Hz 

500 Hz 

1000 Hz 

1600 Hz 

2500 Hz 

2500 Hz 

Microphone saturation

could be affecting spectra

500 Hz 

1000 Hz 

1600 Hz 

5000 Hz 

10000 Hz 
NASA/TM—2010-216366 17

Figure 11.—Median mic spectra (upper left), peak location plot (lower left), and beamform maps for the full-scale engine 

operating at max afterburner condition.

Microphone saturation

could be affecting spectra 500 Hz 

1000 Hz 

1600 Hz 

2500 Hz 

5000 Hz 

10000 Hz 
NASA/TM—2010-216366 18

Figure 12.—No chevron (left) versus chevron (right) comparison for the full-scale engine operating with PLA = 83°.

500 Hz 500 Hz 

1000 Hz 1000 Hz 

1600 Hz 1600 Hz 

2500 Hz 2500 Hz 

5000 Hz 5000 Hz 

10000 Hz 10000 Hz 
NASA/TM—2010-216366 19

Figure 13.—No chevron (left) versus chevron (right) comparison for the full-scale engine operating at PLA = 102°.

500 Hz 500 Hz 

1000 Hz 1000 Hz 

1600 Hz 1600 Hz 

2500 Hz 2500 Hz 

5000 Hz 5000 Hz 

10000 Hz 10000 Hz 
NASA/TM—2010-216366 20

Figure 14.—No chevron (left) versus chevron (right) comparison for the full-scale engine operating with PLA = 130°.

500 Hz 

1000 Hz 

1600 Hz 

2500 Hz 

5000 Hz 

10000 Hz 
NASA/TM—2010-216366 21

Figure 15.—Median mic spectra (upper left), peak location plot (lower left), and beamform maps for a 50.8 mm diameter round 

convergent nozzle operating at Mjet = 0.9, Ma = 1.47, Ts/Tamb = 2.7.

8000 Hz 

2000 Hz 

4000 Hz 

1000 Hz 

16000 Hz 

32000 Hz 

40000 Hz 
NASA/TM—2010-216366 22

Figure 16.—Median mic spectra (upper left), peak location plot (lower left), and beamform maps for a 50.8 mm diameter round 

convergent nozzle operating at Mjet = 1.4, Ts/Tamb = 1.0.

1600 Hz 

2806 Hz 

4000 Hz 

5612 Hz 

6300 Hz 

7099 Hz 

8459 Hz 

11180 Hz

12500 Hz 

25000 Hz 
NASA/TM—2010-216366 23

Figure 17.—Median mic spectra (upper left), peak location plot (upper right), and beamform maps (left Mjet = 1.22 Ts/Ta = 0.77, 

right Mjet =1.47 Ts/Ta = 0.70) for the 1/4th scale model of the F400 nozzle.

754 Hz 

1118 Hz 

1789 Hz 

2236 Hz 

2649 Hz 

3150 Hz 

3768 Hz 

5297 Hz 

7536 Hz 

10000 Hz 

13296 Hz 

754 Hz 

1118 Hz 

1789 Hz 

2236 Hz 

2649 Hz 

3150 Hz 

3768 Hz 

5297 Hz 

7536 Hz 

10000 Hz 

13296 Hz 
NASA/TM—2010-216366 24

754 Hz 754 Hz 

1118 Hz 1118 Hz 

1789 Hz 1789 Hz 

2236 Hz 2236 Hz 

2649 Hz 2649 Hz 

3150 Hz 3150 Hz 

3768 Hz 3768 Hz 

5297 Hz 5297 Hz 

7536 Hz 7536 Hz 

10000 Hz 10000 Hz 

13296 Hz 13296 Hz 

Figure 18.—Median mic spectra (upper left), peak location plot (upper right), and beamform maps (left Mjet = 1.22 Ts/Ta = 0.77, 

right Mjet = 1.22 Ts/Ta = 1.86) for the 1/4th scale model of the F400 nozzle.
NASA/TM—2010-216366 25

Figure 19.—Median mic spectra (upper left), peak location plot (upper right), and beamform maps (left Mjet =1.47 Ts/Ta = 0.70, right 

Mjet = 1.47 Ts/Ta = 1.80) for the 1/4th scale CD nozzle.

754 Hz 754 Hz 

1118 Hz 1118 Hz 

1789 Hz 1789 Hz 

2236 Hz 2236 Hz 

2649 Hz 2649 Hz 

3150 Hz 3150 Hz 

3768 Hz 

5297 Hz 

10000 Hz 10000 Hz 

13296 Hz 13296 Hz 

3768 Hz 

5297 Hz 

7536 Hz 7536 Hz 
NASA/TM—2010-216366 26

Figure 20.—No chevron (left) versus chevron (right) comparison of cold jets (Ts/Ta = 0.77) at Mjet = 1.22 for the 1/4th scale CD nozzle.

754 Hz 754 Hz 

1118 Hz 1118 Hz 

1789 Hz 1789 Hz 

2236 Hz 2236 Hz 

2649 Hz 2649 Hz 

3150 Hz 3150 Hz 

3768 Hz 3768 Hz 

7536 Hz 7536 Hz 

10000 Hz 10000 Hz 

13296 Hz 13296 Hz 

5297 Hz 5297 Hz 
NASA/TM—2010-216366 27

Figure 21.—No chevron (left) versus chevron (right) comparison of cold jets (Ts/Ta = 0.77) at Mjet = 1.47 for the 1/4th scale CD nozzle.

754 Hz 754 Hz 

1118 Hz 1118 Hz 

2236 Hz 2236 Hz 

2649 Hz 2649 Hz 

3150 Hz 3150 Hz 

3768 Hz 3768 Hz 

5297 Hz 5297 Hz 

7536 Hz 7536 Hz 

10000 Hz 10000 Hz 

13296 Hz 13296 Hz 

1789 Hz 1789 Hz 
NASA/TM—2010-216366 28

Figure 22.—No chevron (left) versus chevron (right) comparison of hot jets (Ts/Ta = 1.86) at Mjet = 1.22 for the 1/4th scale CD nozzle.

754 Hz 754 Hz 

1118 Hz 1118 Hz 

1789 Hz 1789 Hz 

2236 Hz 2236 Hz 

2649 Hz 2649 Hz 

3150 Hz 3150 Hz 

3768 Hz 3768 Hz 

5297 Hz 5297 Hz 

7536 Hz 7536 Hz 

10000 Hz 10000 Hz 

13296 Hz 13296 Hz 
NASA/TM—2010-216366 29

Figure 23.—No chevron (left) versus chevron (right) comparison of hot jets (Ts/Ta = 1.86) at Mjet = 1.47 for the 1/4th scale CD nozzle.

754 Hz 754 Hz 

1789 Hz 1789 Hz 

2236 Hz 2236 Hz 

2649 Hz 2649 Hz 

3768 Hz 3768 Hz 

10000 Hz 10000 Hz 

13296 Hz 13296 Hz 

1118 Hz 1118 Hz 

3150 Hz 3150 Hz 

5297 Hz 5297 Hz 

7536 Hz 7536 Hz 
REPORT DOCUMENTATION PAGE Form Approved 

OMB No. 0704-0188 

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the 

data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this 

burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. 

Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB 

control number. 

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 

1. REPORT DATE (DD-MM-YYYY)

01-08-2010 

2. REPORT TYPE

Technical Memorandum 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

Phased Array Noise Source Localization Measurements of an F404 Nozzle Plume at Both 

Full and Model Scale 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Podboy, Gary, G.; Bridges, James, E.; Henderson, Brenda, S. 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 561581.02.08.03.18.02 

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-17302 

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-2010-216366 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Category: 71 

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

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

13. SUPPLEMENTARY NOTES

14. ABSTRACT

A 48-microphone planar phased array system was used to acquire jet noise source localization data on both a full-scale F404-GE-F400 

engine and on a 1/4th scale model of a F400 series nozzle. The full-scale engine test data show the location of the dominant noise sources in 

the jet plume as a function of frequency for the engine in both baseline (no chevron) and chevron configurations. Data are presented for the 

engine operating both with and without afterburners. Based on lessons learned during this test, a set of recommendations are provided 

regarding how the phased array measurement system could be modified in order to obtain more useful acoustic source localization data on 

high-performance military engines in the future. The data obtained on the 1/4th scale F400 series nozzle provide useful insights regarding 

the full-scale engine jet noise source mechanisms, and document some of the differences associated with testing at model-scale versus fullscale. 

15. SUBJECT TERMS

Array; Noise source; F404 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU 

18. NUMBER

 OF 

 PAGES

35 

19a. NAME OF RESPONSIBLE PERSON

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

a. REPORT

U 

b. ABSTRACT

U 

c. THIS 

PAGE

U 

19b. TELEPHONE NUMBER (include area code)

443-757-5802 

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18