James E. Bridges and Mark P. Wernet

Glenn Research Center, Cleveland, Ohio

PIV Measurements of Supersonic 

Internally-Mixed Dual-Stream Jets

NASA/TM—2012-217250

March 2012

AIAA–2011–2786
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James E. Bridges and Mark P. Wernet

Glenn Research Center, Cleveland, Ohio

PIV Measurements of Supersonic 

Internally-Mixed Dual-Stream Jets

NASA/TM—2012-217250

March 2012

AIAA–2011–2786

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

This work was supported by the Supersonics Project of the NASA Fundamental Aeronautics Program. The authors 

acknowledge the able efforts of their colleagues at the NASA AeroAcoustic Propulsion Lab. 

Cliff Brown and Nick Georgiadis provided very helpful reviews of the work.

Available from

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Trade names and trademarks are used in this report for identification 

only. Their usage does not constitute an official endorsement, 

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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—2012-217250 1

PIV Measurements of Supersonic 

Internally-Mixed Dual-Stream Jets

James E. Bridges and Mark P. Wernet

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

While externally mixed, or separate flow, nozzle systems are most common in high bypass-ratio 

aircraft, they are not as attractive for use in lower bypass-ratio systems and on aircraft that will fly 

supersonically. The noise of such propulsion systems is also dominated by jet noise, making the study and 

noise reduction of these exhaust systems very important, both for military aircraft and future civilian 

supersonic aircraft. This paper presents particle image velocimetry of internally mixed nozzle with 

different area ratios between core and bypass, and nozzles that are ideally expanded and convergent. Such 

configurations independently control the geometry of the internal mixing layer and of the external shock 

structure. These allow exploration of the impact of shocks on the turbulent mixing layers, the impact of 

bypass ratio on broadband shock noise and mixing noise, and the impact of temperature on the turbulent 

flow field. At the 2009 AIAA/CEAS Aeroacoustics Conference the authors presented data and analysis 

from a series of tests that looked at the acoustics of supersonic jets from internally mixed nozzles. In that 

paper the broadband shock and mixing noise components of the jet noise were independently manipulated 

by holding Mach number constant while varying bypass ratio and jet temperature. Significant portions of 

that analysis was predicated on assumptions regarding the flow fields of these jets, both shock structure 

and turbulence. In this paper we add to that analysis by presenting particle image velocimetry 

measurements of the flow fields of many of those jets. In addition, the turbulent velocity data documented 

here will be very useful for validation of computational flow codes that are being developed to design 

advanced nozzles for future aircraft.

Introduction

The Supersonics Project at NASA has identified Airport Noise as one of the Technical Challenge 

areas for future civilian supersonic vehicles. The charter for this research thrust is to identify and mature 

relevant noise reduction technologies and engineering tools that will be required to create a commercially 

viable supersonic aircraft. Given that most paper studies point to moderate bypass ratio propulsion 

systems as being likely for such aircraft, one activity of the program has been to assure that appropriate 

flow and noise databases exist for this type of jet flow. While several acoustic datasets exist for single 

flow supersonic jets and many more for dual stream subsonic jets, there was a lack of data for supersonic

dual stream jets. There was especially a dearth of flow field data, a big hole given the driving need for 

detailed data for validating advanced computational methods. Having been identified, a campaign was 

mounted to fill this hole in research data by building an internally mixed nozzle system with variable 

bypass area and multiple convergent-divergent nozzles, and making extensive acoustic and particle image 

velocimetry (PIV) measurements over a range of flow conditions. Along with a series of externally mixed 

nozzle flows this test was referred to as the Supersonic Dual Flow test, and it required multiple test entries 

over a three-year span to complete. The acoustic data were analyzed and presented in Reference 1; this 

paper primarily presents and analyzes the particle image velocimetry data, although corresponding 

acoustic results are given alongside for convenient reference.

For many years researchers have compared convergent-divergent and convergent nozzles to isolate 

broadband shock noise (BBSN), under the assumption that BBSN is independent of jet mixing noise 
NASA/TM—2012-217250 2

(Refs. 2 and 3). That seems consistent with observations that low frequency noise stays the same for a 

given ideally expanded Mach number. The physical mechanism of BBSN seems well understood in that it 

involves the interaction of the standing pattern of flow inhomogeneities (shocks) with convecting 

turbulence. The fully expanded Mach number, M, give the shock strengths and spacings. The strength and 

spectral content of the turbulence in the shear layer is given by velocity gradient, ∂U/∂r. Further, Tam’s 

wavy wall analogy (Ref. 4) that successfully predicts the frequency shift of the peak BBSN, implicates 

the convection speed of the turbulence as having a major role in the source mechanism of BBSN. 

Consequently, BBSN is a function of Mach number M and jet velocity U. Conventionally, BBSN has 

been assumed to be independent of jet temperature (Ref. 2), although recent papers have found some 

dependence (Ref. 3). Whether this is sensitivity to temperature or to shear velocity is not possible to 

ascertain since the two are tied for a given Mach number M.

The test matrix for the flow measurement portion of the Supersonic Dual Flow (SDF) test was created 

to study the following issues as they relate to the flow and noise of dual-stream jet flows:

1) Impact of splitter area ratio (Bypass Ratio)

2) Impact of shocks (Convergent versus Ideally Expanded Nozzles)

Objective 1 was addressed simply by changing the area ratio of the upstream splitter and running the 

same core and bypass pressure and temperature ratios. It was expected that there would be a major change 

in the structure of the turbulence associated with the difference in geometry of the shear layers. An 

alternative approach would have been to keep the mixed velocity constant, but that was not pursued since 

jet noise does not simply scale on mixed jet velocity. Of special interest was how splitter area ratio

changed the turbulence in the shock cells and indeed the strength of the shocks themselves.

Objective 2 was addressed by testing convergent nozzles and ideally expanded nozzles at the same 

bypass ratio and flow condition. This was an important set of data because it documents how turbulence 

was affected by the shocks for a given flow condition. It has been observed that jet mixing noise is 

roughly the same with or without shocks, so the question was raised: is the turbulence also the same?

In discussion of the results, three jet noise components will be frequently mentioned: jet mixing noise, 

broadband shock noise (BBSN), and screech. Notionally, these are defined as shown in Figure 1. Another 

source of jet noise is nominally found in the cases presented in this paper, often referred to as Mach wave 

emission. This sound is associated with the large-scale instability waves or convecting eddies moving at 

supersonic speeds and producing sound that has strong directivity in the downstream direction. 

Figure 1.—Notional definition of jet noise sources as 

identified in acoustic spectra acquired at broadside 

angles.
NASA/TM—2012-217250 3

Facility, Model Hardware, Instrumentation, and Flow Conditions

Model Hardware

The test was conducted in the NASA Glenn Research Center AeroAcoustic Propulsion Laboratory 

(AAPL). The AAPL is a 65-ft radius anechoic geodesic hemispherical dome. Acoustic wedges covered 

the walls of the dome and the floor area of the test arena. The ambient temperature, pressure, and relative 

humidity were recorded within the dome.

The Nozzle Acoustic Test Rig (NATR) is a large, ejector-driven freejet used to generate forward 

flight airflow through its 53 in. (1.35 m) exit. Simulated flight speeds up to Mach 0.35 were produced 

through the acoustically lined tunnel. 

The High Flow Jet Exit Rig (HFJER) was positioned at the end of the NATR, with the nozzle exit 

roughly 60 in. (1.5 m) downstream of the NATR exit plane. The HFJER delivered high-pressure air from 

the Center’s compressed air system to the test article. Flow rates up to 25 lbm/sec are available in each of 

two streams. One of the streams passes through a natural gas-fired combustor that could heat the flow to 

approximately 1800R. The rig is instrumented with a 40-element charging station rakes in both core and 

bypass streams. The control system allows instantaneous calculation of total pressure ratio and total 

temperature ratio of both streams along with the deviation from the researcher-input conditions. In this 

way the flow conditions were held to keep jet velocities within 0.5 percent of the intended condition, 

including the simulated flight stream.

The test nozzles were attached to the aft end of the HFJER and are shown in Figure 2. The model 

system consisted of an interchangeable set of four axisymmetric splitters and four nozzles, all of exit 

diameter 4 in. (101.6 mm). The axisymmetric splitters yielded area ratios of 0.2 (S1), 1.0 (S2), 2.0 (S3)

and 3.0 (S4), and terminate 6.5 in. (166 mm) upstream of the nozzle exit. The convergent-divergent 

nozzles were designed using the method of characteristics and tuned with RANS CFD to produce nearly 

perfectly expanded flows of Mach 1.18 (M2), 1.4 (M4), and 1.5 (M5). A center plug, required to achieve 

the highest area ratio, was varied in length and found to be inconsequential to the sound produced by the 

nozzles. 

Flow Conditions

A series of flow conditions were defined that matched the design Mach numbers Md of the 

convergent-divergent nozzles while varying the temperature of the core to hit a common set of velocity 

ratios. (Even in the case of convergent nozzle flows, the fully expanded Mach number M will be used.) In 

doing this the shock structure and turbulence intensities were to be independently varied. For this report a 

subset of these flow conditions are used and are documented in Table 1. Actual test conditions were 

specified in terms of nozzle pressure ratio (NPR) and total temperature ratios on core (Tt,c/T∞) and bypass

(Tt,b/T∞) streams, respectively. Note that since both streams emerged inside the common nozzle, the 

pressure was the same on both streams for all setpoints. The table also lists ideally expanded velocities of 

the core (Uic), bypass (Uib), and mass-weighted fully mixed streams (Uim).
NASA/TM—2012-217250 4

C4-S1 C4-S3

C4-S4

M2-S1 M4-S1

M5-S1

Figure 2.—Nozzles used in Supersonic Aeroacoustic Database.
NASA/TM—2012-217250 5

TABLE 1.—TEST MATRIX WITH NOZZLE ID AND FLOW CONDITIONS.

Config Setpoint M NPR Tt,c/T∞ Tt,b/T∞ BPR Uic, 

ft/s

Uib, 

ft/s

Uim, 

ft/s

C4S1 6010 1.0 1.89 0.99 0.99 0.21 1014.48 1016 1015

C4S1 6040 1.0 1.89 1.65 1.00 0.27 1311.86 1026 1256

C4S1 8020 1.2 2.43 1.23 1.17 0.20 1316.53 1287 1312

C4S1 8060 1.2 2.40 2.76 1.03 0.34 1973.04 1208 1810

C4S1 9020 1.4 2.42 1.23 1.19 0.20 1314.22 1294 1311

C4S1 9050 1.4 3.17 2.21 1.02 0.30 1971.44 1336 1845

C4S1 10020 1.5 3.67 1.23 1.19 0.20 1534.71 1514 1531

C4S1 10050 1.5 3.67 1.97 1.12 0.26 1971.04 1483 1880

C4S3 6040 1.0 1.88 1.67 0.99 2.96 1309.8 1011 1094

C4S3 8060 1.2 2.41 2.73 1.09 4.01 1977.61 1251 1425

C4S3 10060 1.5 3.66 2.48 1.16 3.36 2192.97 1497 1681

C4S4 6040 1.0 1.88 1.63 0.98 4.82 1311.8 1020 1076

C4S4 8060 1.2 2.40 2.71 1.10 6.29 1974.97 1260 1380

C4S4 9050 1.4 3.17 2.15 1.10 4.83 1977.1 1413 1524

C4S4 10050 1.5 3.67 2.00 1.15 4.37 1972.67 1492 1592

M2S1 8120 1.2 2.42 1.21 1.17 0.21 1314.39 1292 1311

M2S1 8130 1.2 2.43 1.64 1.03 0.25 1537.23 1220 1479

M2S1 8160 1.2 2.39 2.73 1.04 0.35 1969.93 1214 1804

M2S3 8130 1.2 2.42 1.64 1.09 2.70 1530.39 1247 1329

M2S3 8160 1.2 2.40 2.74 1.13 4.08 1970.83 1269 1433

M2S4 8130 1.2 2.43 1.68 1.13 4.05 1534.45 1259 1318

M2S4 8160 1.2 2.40 2.73 1.10 6.93 1974.89 1259 1370

M4S1 9120 1.4 3.17 1.34 1.28 0.20 1533.55 1495 1527

M4S1 9130 1.4 3.18 1.55 1.13 0.23 1645.27 1403 1603

M5S1 10120 1.5 3.68 1.21 1.16 0.21 1535.67 1504 1530

M5S1 10150 1.5 3.67 1.99 1.12 0.27 1974.27 1482 1879

M5S1 10170 1.5 3.64 2.96 1.12 0.35 2417.65 1483 2212

M5S4 10150 1.5 3.66 2.02 1.16 4.59 1973.6 1494 1590

M5S4 10170 1.5 3.64 3.01 1.16 6.63 2419.92 1498 1645

Far-Field Acoustic Instrumentation

This report will cover both far-field acoustic spectral directivity and plume measurements using rakes 

of static pressure probes and total temperature probes. Acoustic measurements were recorded by an array 

of 24 microphones placed on an arc with a radius of 45 ft (13.7 m) at 5° intervals from 45° to 160°. The 

microphones were 1⁄4 in. Bruel & Kjaer model 3949 free-field type, pointed at the nozzle exit with the grid 

caps removed. Bruel & Kjaer Nexus amplifiers provide the signal conditionings. Eight seconds of data 

are recorded at each point on a DataMAX Instrumentation Recorder (R.C. Electronics, Inc.) at 200 kHz 

sample rate using a 90 kHz low-pass filter to limit the bandwidth. After being corrected for amplifier 

gain, the prerecorded background sound levels were removed spectrally and actuator and free-field 

responses from the calibration facility were applied. Atmospheric attenuation was added back to the 

sound spectra using the formula of Shields and Bass (Ref. 5). In cases of simulated forward flight the

refraction of the sound through the freejet shear layer was removed and the data were transformed to a 

consistent 100 jet-diameter distance assuming lossless spherical spreading, and interpolated to a 

consistent set of observer angles. The narrowband power spectral densities were further normalized using 

jet diameter and ideally expanded, fully mixed velocity to be presented as power spectral density as a 

function of Strouhal number.
NASA/TM—2012-217250 6

Figure 3.—Schematic for PIV acquisition strategy.

PIV With Streamwise Light Sheet

The particle image velocity (PIV) equipment was mounted on a large traverse that provided a 

moveable, common base for light sheet optics and cameras. Al2O3 seeders were used for the core and 

bypass streams along with a Rosco Delta 3000 glycol fogger to seed the ambient air with particles of 

approximately 0.5 μm.

The flow field was imaged using dual cameras positioned side-by-side at right angles to the jet axis, 

offset in axial distance, at a distance of 63 in. (1.6 m) from the jet centerline. Each camera imaged a field 

of view of 12- by 8-in. (30- by 20-mm), which with 1 in. (25 mm) overlap, yielded an aggregate field of 

view of 12 in. (30 mm) vertical by 15 in. (38 mm) axial. This field of view was centered on the jet 

centerline. Acquisition went straight to disk in bursts of 400 frames at roughly 1Hz. The entire optical 

arrangement was traversed to 7 stations at a step size of 14 in. (35.5 mm) between acquisitions (84 in.

(2.13 m) of travel), delivering flow field data roughly 99 in. (2.51 m) of flow field in all. This strategy is 

depicted in Figure 3. The nozzle was not in first field of view, but the traverse was moved far enough 

forward before the laser was operated to allow the first camera to image the nozzle and establish the 

reference after the rig reached its full thermal growth.

Data Presentation

In the presentation of results, the parameters of nozzle type, splitter area ratio, pressure ratio, and 

temperature ratio were shuffled to look at the impact of the different parameters in isolation. A series of 

figures have been prepared which combine flow field and acoustic data. The flow fields are represented 

by contour plots of mean and rms velocities. The acoustic far field of the jets is represented by the 

OASPL directivity and spectrum at 90° to the jet axis.

For better cross-comparison of results, the flow field plots all used the same range in the contour 

levels. The mean axial velocity, U/Uic, ranged 0 to 1.2; the rms of axial velocity, u′/Uic, ranged 0 to

0.185. Both velocity quantities were normalized by ideally expanded core velocity. This normalization 

was chosen first to give a reasonable, consistent range for the mean velocity plots, and second in the 

belief that the TKE would scale on the mean gradients, starting with the peak velocity. This did seem to 

be a good choice as the peak TKE was fairly consistent for all setpoints and nozzle configurations. Other 

normalizations were explored, but for relating the flow to the sound for configurations that had a constant 

initial flow condition the ideally expanded core velocity seemed best.

Two plots of the acoustic far field of the jets are shown along with the mean and rms flow fields for 

each case. The spectra have been normalized to a 100 jet diameter distance, removing atmospheric 

attenuation, and expressed as normalized power spectral density in decibel, e.g., 20 log10((Pa/St)/20 μPa), 

plotted as a function of Strouhal number St = f D/Uic. Note that similar to the flow normalization, the 

Strouhal number used in the sound power normalization was the ideally expanded core velocity. Again, 

other choices of normalization could be used but as explained in the Discussion section this choice was 

most widely useful.
NASA/TM—2012-217250 7

Discussion accompanying the presentation of data attempts to relate the flow and acoustic data. To do 

this the approximation is made that the jet mixing noise at right angles to the jet scales roughly as U8 for 

jets of the speed range measured here. For subsonic jets acoustic analogy theory indicates that the noise 

also scales with turbulent kinetic energy to the 7/2 power, producing a scaling with rms velocity u′/U of 

70 log10(u′/U). No attempt is made to correlate the aft angle noise, which can be dominated by mach wave 

emission at the highest speeds shown here, with the flow data.

Results

Impact of Splitter Area Ratio (0.2 – 3.0)

In this section jets of the same nozzle and flow conditions, but different splitter area ratios, are 

compared. Bear in mind that with a common nozzle, the pressures on the two streams are necessarily 

equal, meaning that an increase in the core velocity temperature translates into higher velocity in the core 

stream. It is important to remember that the flow data have been normalized by the common core velocity

Uic. This makes the plots of mean flows have significant similitude near the nozzle, but means that farther 

downstream where the flows are dissimilar the dominant mean gradients are given more by equivalent 

mixed velocity. This scaling does not present the flows on an equal thrust basis—that would require 

normalizing by mass-averaged velocity instead. 

For example, in Figure 4, the convergent nozzle is run at M = 1 with a warm core flow using three 

different splitters to achieve three different bypass ratios. The mean velocity fields show a normalized 

potential core velocity of approximately 1.0 near the nozzle exit. Recall that the splitter between hot and 

cold flows terminates approximately 1.5 nozzle diameters upstream of the nozzle exit. In the mean flow 

profiles at x/D = 0.25 the thin cold flow from the splitter C4S1 configuration has already merged with the 

hot core flow, the nozzle exit being 23 times the cold flow annulus radii downstream. In the C4S2 and 

C4S3 flow profiles the lower velocity cold flow is clearly discernable even at x/D = 2, although they are 

not easily distinguished from one another. Inspection of the turbulence profiles at x/D = 0.25 shows a hint 

of the inner shear layer just inside the common outer shear layer in the C4S1 profile.

Figure 4.—Transverse profiles of mean (top) and rms (bottom) velocities for convergent 

nozzle C4 with different splitter area ratios (setpoint 6040; M=1.0, Tt,c/T∞=1.65).
NASA/TM—2012-217250 8

Ideally Expanded Nozzles

Ideally expanded flows do not contain strong shocks making them better for exploring the impact that 

bypass ratio has on the flow and noise. To get a feel for the variations being made to the flow, consider 

the velocity profiles measured near the nozzle exit when the ideally expanded M = 1.19 jet is operated 

with three different core temperatures, shown in Figure 5. The increase in core temperature at a fixed 

pressure ratio results in an increase in core velocity, shown clearly in the flow profiles from the S4 splitter 

geometry. In the first case (setpoint 8020), the two streams are run at approximately the same temperature 

and there is only a shear layer on the nozzle lip line. In the second case (setpoint 8030), the core is heated 

slightly and a second shear layer emerges, much weaker that the one on the nozzle lip line. In the third 

case (setpoint 8060), the inner shear layer is over half the strength of the outer one.

The first ideally expanded flow to be examined is actually a subsonic jet from the conic nozzle C4 

operating at M = 1.0, shown in Figure 6. The mean flow shows no evidence of shock cells and has a 

potential core ending around x/D = 8. The rms of the axial velocity peaks near the lipline radially, broadly 

around the end of the potential core, at roughly u′/Uic = 0.15. This is in keeping with a wide range of 

subsonic jets measured previously (Ref. 6).

Now heat is added to the core stream, keeping the pressures of the two streams the same, set at M = 

1.0 (Figure 7). In the case of splitter area ratio = 0.2 (top plots of Figure 7) the two shear layers merge 

almost immediately downstream of the nozzle and the rms velocity levels are approaching 0.17. The jets 

with higher area ratios are roughly 5 percent longer, and have peak rms velocity levels that only reach 

0.14. In the jets with higher bypass area ratios the initial shear layers have much less turbulence intensity 

as each is driven by the velocity gradient between the cold flow and the ambient rather than the hot flow 

and ambient. Downstream of the merger between the hot and cold flows, where the turbulence is expected 

to scale on the fully mixed velocity, the low area ratio jet has higher fully mixed velocity and 

correspondingly higher turbulence levels, while the splitter area ratio 2 and 3 jets, being nearly matched in 

mixed velocity, have nearly matched turbulence levels.

Only the S1 and S3 splitter configurations were measured acoustically for this flow condition, and 

these are presented in Figure 7. The overall sound pressure level is approximately 4 dB louder for the jet 

with area ratio = 0.2 than the one with area ratio = 2 at polar angle of 90°. Given that the mass-averaged 

mixed velocities of the two jets were 1256 ft/s and 1095 ft/s respectively, a 5 dB difference would be 

expected around 90° using U8 scaling on the fully mixed velocities. This is also in line with what would 

be expected from the difference in peak u′/Uc of 0.17 and 0.14 using (u′/U)7 scaling.

(a) (b) (c)

Figure 5.—Mean velocity profiles, normalized by ambient speed of sound, at x/D = 0.2, for fully expanded M = 1.19 

jets with different core temperature ratios (setpoints) (a) Tt,c/T∞=1.23 (8020), (b) Tt,c/T∞=1.64 (8030), (c) 

Tt,c/T∞=2.76 (8060). 

C4S1

Figure 6.—Convergent nozzle C4S1 at cold subsonic flow (setpoint 6010; M=1.0, Tt,c/T∞= Tt,b/T∞=1.0).
NASA/TM—2012-217250 9

C4S1

C4S3

C4S4

Figure 7.—Convergent nozzle C4, various splitter area ratios: 0.2 (S1), 2 (S3), 3 (S4) (setpoint 6040; M=1.0, 

Tt,c/T∞=1.65).

M2S1

Figure 8.—Ideally expanded nozzle M2S1 (setpoint 8120; M=1.19, Tt,c/T∞=1.21).

In preparation for looking at velocity data corresponding to varying splitter area ratios at M = 1.19 

(setpoints 8130 and 8160), first consider the velocity fields from the M2S1 nozzle operating at this Mach 

number but with the same temperature on both stream (setpoint 8120) where the splitter should make no 

difference to the flow. In Figure 8 it is observed that even with this ideally expanded nozzle being 

operated at its design condition there are still slight shock cells present in the mean velocity contour plot. 

The end of the potential core occurs around x/D = 10. This is also shown in the plot of the turbulence 

intensity, where the peak turbulence level is around 16 percent.

Next, consider how the M = 1.19 flow changes when the core is moderately heated and how this 

change depends upon the splitter geometry, specifically the splitter area ratio. In Figure 9 the M2 nozzle 

again has barely perceptible shock cells in the mean velocity flows, and the potential core of the core 

stream is essentially the same for all splitter geometries. The rms velocity fields are also very similar in 

shape as the outer shear layer dominates the production of turbulence. While the turbulent velocity plots 

are very similar in shape, the peak turbulence intensity reduces with increase in splitter area ratio, from 

0.15 in the M2S1 flow to 0.13 in the M2S4 configurations. Note that this is a steady reduction in peak 

turbulence level with increase in the outer flow thickness, starting with the single-flow case (Figure 8).

This trend in turbulence fields is reflected in the far-field noise. The OASPL directivities of the 

configurations are very similar, with the M2S1 being louder than the M2S3 and M2S4 by 2 and 3 dB, 

respectively. The 3 and 4 dB decrease in OASPL at 90° for S3 and S4 relative to S1 respectively is a 
NASA/TM—2012-217250 10

M2S1

M2S3

M2S4

Figure 9.—Ideally expanded nozzle M2 various splitter area ratios: 0.2 (S1), 2 (S3), 3 (S4) (setpoint 8130; M=1.19, 

Tt,c/T∞=1.64).

bit less than the 8 and 9 dB expected from the fully mixed velocities, listed in Table 1. It is also a bit less 

than what would be expected by the small variation in peak u′/Uic. The spectra at 90° show very minimal 

BBSN, if any, and have the same shapes, varying only in amplitude.

The next dataset, in Figure 10, keeps the same Mach number, but increases the core temperature, and 

hence velocity. As seen in Figure 5, this also results in the inner shear layer becoming comparable to the 

outer shear layer in overall strength. 

The flows are again nearly free of shocks and now the increase in the gradient of the inner shear layer 

causes the jet with the low splitter area ratio to have a much stronger turbulent velocity field (peaking 

near u′/Uic = 0.16) than the jets with larger splitter area ratio (u′/Uic ~0.13). The turbulence distribution is 

very different as well, with the strong inner shear layer moving the peak turbulence toward the centerline 

with increase in splitter area ratio. 

For all the changes to the turbulence distribution, the OASPL directivities of the different splitter 

configurations are remarkably similar at the broadside angles, especially M2S1 and M2S3 which are 

virtually identical. To see differences in the acoustic fields at these angles corresponding to the 

differences in the turbulence distributions, one has to look at the spectra. There is a hint of BBSN in the 

spectra, with a slight hump at St = 1, most pronounced in the M2S3 configuration. At frequencies below 

the peak, frequencies presumed to come from the jet mixing noise component, the M2S1 splitter is 3 dB

louder than the M2S3 and M2S4 splitter configurations, correlating with the increased turbulence

measured. Near the spectral peak the M2S3 splitter configuration is loudest by 2 dB, while the M2S1 and 

M2S4 configurations are exactly the same. At the highest frequencies the M2S1 is again louder, perhaps 

correlating with the higher turbulence in the shear layer of that configuration in the first few jet diameters. 

These subtle variations in what appear to be mixing noise are not easily associated with the differences in 

the turbulent fields of the jets. Also not obvious from the contour plots of u′/Uic is the sharp spike in 

OASPL directivity at 130° for the M2S3 configuration. This strong directivity is what has been attributed 

to Mach wave emission and does not correlate with the turbulence directly.
NASA/TM—2012-217250 11

M2S1

M2S3

M2S4

Figure 10.—Ideally expanded nozzle M2, various splitter area ratios: 0.2 (S1), 2 (S3), 3 (S4) (setpoint 8160; M=1.19, 

Tt,c/T∞=2.73).

Figure 11.—Mean velocity along centerline plume of convergent nozzle C4 with 

splitter area ratios of 0.2 (S1) and 3 (S4) (setpoint 9050; M=1.4, Tt,c/T∞=2.21).

Convergent Nozzle

Now that the general trends of supersonic jets with varying bypass ratio have been shown, the 

complication of shocks is introduced. However, before diving into the flow and noise data in depth, one 

aspect of the underexpanded jets should be kept in mind: the shock cell lengths are only weakly 

dependent upon the temperature or the splitter geometry (Ref. 1). This is somewhat surprising given that 

the cells are established by reflection of the shocks from the velocity discontinuity at the shear layer, an 

aspect of the jet flow which is very strongly distorted by changes to the splitter geometry. Granted, the 

gradient from the outer shear layer was always greater than that from the inner in these experiments, but it 

seemed reasonable to expect some change in the shock cell geometry with changes in the inner shear 

layer strength and location. But this was not the case. Consider Figure 11, which shows centerline 

velocity profiles for the underexpanded convergent nozzle with the smallest and largest area ratio 

splitters. The flow is unchanged in the potential core, and the shock cell spacing only varies slightly after 

the mean flow begins to decay.
NASA/TM—2012-217250 12

Figure 12.—Contours of mean velocity, detailing shock structure of convergent 

nozzle C4 with splitter area ratios of 0.2 (S1) and 3 (S4) (setpoint 9050; 

M=1.4, Tt,c/T∞=2.21).

More detail of the shock structure in the presence of the inner shear layer is provided in Figure 12.

The interplay of velocity gradients between the expansion fans and the shear layers are surprisingly 

independent, accounting for the insensitivity of the shocks to the inner shear layer.

The same three configurations as were explored above in Figure 7 are now presented in Figure 13, 

only with an increase in pressure ratio that yields M = 1.19 flow. Shock cells appear in the mean velocity 

contours of the jet flows as a result of the underexpansion of the pressure at the nozzle exit. At this flow 

condition there is again enhanced turbulence in the jet with low area ratio splitter, as was seen in the 

ideally expanded jets, but the potential core of the low area ratio splitter is not much longer than the jets 

with high area ratio splitters. With the relatively large change in inner shear layer location there is a 

fundamental change in the distribution of the turbulence as the splitter ratio is changed, with very little

turbulence in the initial outer shear layers, and a much shorter region of strong turbulence. 

Acoustically, the OASPL at 90° does not reflect the change in the turbulence levels between the 

various splitters cases, having essentially only a 1 to 2 dB change in overall sound level. However, this is 

a case where the equalities are fortuitous. Spectrally, the three jets have almost identical BBSN spectra 

beyond the first peak, where the C4S3 nozzle has a significantly greater peak than the other two. The jet

mixing noise corresponds more with the levels of turbulence observed in the rms contour plots, the C4S1 

nozzle being 3 and 5 dB louder than the C4S3 and C4S4 configurations, respectively. The C4S3 reaches 

roughly the same level as the C4S1 by virtue of its high BBSN peak. The equality of the BBSN spectra 

are not understood in light of the strong difference in turbulence in the shear layers.

There was no screech measured from any of these flow configurations.

Changing to setpoint 9050 brings about an increase in pressure ratio and a slight reduction in core 

temperature, increasing the cold flow velocity while maintaining the hot core velocity. As seen in Figure 

14, this increase in pressure ratio increases the strength of the shocks, but again the shock cells lengths are 

not changed by the change in splitter geometry. Also, the jet with low area ratio splitter has a longer 

potential core. With an increase in velocity gradient on the outer shear layer relative to the inner shear 

layer, the turbulence contour plots of the C4S1 (area ratio = 0.2) and C4S4 (area ratio = 3) are very 

similar, both driven by the outer shear layer. The jet from the high area ratio configuration has reduced 

turbulence in the initial shear layer region, but has only slightly less volume of high intensity turbulence 

and a peak level within 10 percent of the C4S1 configuration.

Looking at OASPL is very misleading as there is significant BBSN present in the acoustic far field, a 

fact made plain in plot of spectra in Figure 14. Again, the BBSN spectra seem to match pretty closely, 

including the peak. The jet mixing noise components differ by some 5 dB, a difference not expected 

based simply on the similarity of the turbulence levels in the two jets downstream of the potential core.

Neither configuration at this flow condition exhibited screech.
NASA/TM—2012-217250 13

C4S1

C4S3

C4S4

Figure 13.—Convergent nozzle C4, various splitter area ratios: 0.2 (S1), 2 (S3), 3 (S4) (setpoint 8060; M=1.19, 

Tt,c/T∞=2.73).

C4S1

C4S4

Figure 14.—Convergent nozzle C4, various splitter area ratios: 0.2 (S1), 2 (S3), 3 (S4) (setpoint 9050; M=1.4, 

Tt,c/T∞=2.21).
NASA/TM—2012-217250 14

Impact of Shock on Turbulence and Noise (Convergent Versus Ideally Expanded Nozzles)

During the presentation of the flow and noise from the ideally expanded (shock-free) and convergent 

nozzle systems there were several allusions to comparisons between two. This is natural since this would 

address the question of the impact of shocks on the flow and noise. To better facilitate this investigation, 

direct comparisons of the two nozzles will be made for the same splitter and flow condition. 

Splitter Area Ratio = 0.2

The comparison of shocked and shock-free flows will begin with the case of minimal bypass ratio, 

e.g. splitter area ratio = 0.2. Of all the ideally expanded nozzles the Md=1.19 convergent-divergent nozzle 

is the worst at being shock-free; hence there is a slight hint of shock cells in the velocity contours of the 

M2 nozzle shown in Figure 15. Of course, these shocks are very weak compared with those found in the 

convergent nozzle flow. Correspondingly, there is also a small amount of broadband shock noise (BBSN) 

that shows above the expected pure jet mixing noise in the spectra. Besides having a distinctive BBSN 

hump in its spectra there is a very small amount of screech in the convergent nozzle noise spectrum. 

Looking at the contour plots of u′/Uic, the distribution of turbulence at the end of potential core is

similar, but the shear layer is quite a bit thicker up front in the convergent, or shocked jet, and the 

potential core is slightly shortened by the enhanced mixing. The ideally expanded M2S1 nozzle has peak 

rms velocity levels of roughly 16 percent, compared to 15 percent for the convergent nozzle. 

The increase in turbulence is reflected in the spectra where the jet mixing component is 1 dB or so 

higher for the ideally expanded nozzle. At broadside angles the BBSN of the convergent nozzle 

dominates the acoustic spectra and makes the OASPL significantly higher. At aft angles the OASPL of 

the convergent-divergent nozzle is higher. It is tempting to say that the increase in turbulence and the 

corresponding jet mixing noise increase is responsible. Perhaps this is true for this relatively low Mach 

number, but in higher Mach number jets to be shown below the correlation between peak turbulence and 

aft angle noise does not hold.

C4S1

M2S1

Figure 15.—Splitter area ratio = 0.2, convergent C4 versus convergent-divergent M2 nozzles (setpoint 8020; 

M=1.19, Tt,c/T∞=1.23).
NASA/TM—2012-217250 15

C4S1

M2S1

Figure 16.—Splitter area ratio = 0.2, convergent C4 versus convergent-divergent M2 nozzles (setpoint 8060; 

M=1.19, Tt,c/T∞=2.76).

The hotter M=1.19 jets in Figure 16 have flow fields that are very similar to each other, e.g., the 

potential core lengths, peak turbulence levels, and the shear layer thicknesses. Their acoustic fields are 

also very similar, with only a slight amount of BBSN in the convergent nozzle showing above the jet 

mixing noise. Compared with the sound spectra in Figure 15 the BBSN of the convergent nozzle is 

roughly the same amplitude (peaking near 109 dB) but the jet mixing noise has increased to make it less 

noticeable. The lack of screech is understood to be due to the jet velocity being too high to support the 

feedback relationship required to match acoustic waves propagating upstream and instability waves

convecting downstream (Ref. 7). 

The isothermal M=1.4 jet (Figure 17) is considered next. The convergent nozzle has the expected 

shock cell pattern for this underexpanded flow condition, and has a slightly shorter potential core length 

than the ideally expanded jet. The ideally expanded nozzle has only the slightest shock cells discernable 

in the mean velocity contour plots. Looking at the turbulent velocity fields, the convergent nozzle again 

has an initial shear layer that is quite a bit thicker than the ideally expanded nozzle. The downstream 

turbulence levels are comparable at u′/Uic = 0.15. The thicker initial shear layers, and hence greater 

mixing, probably account for the reduced potential core length.

Acoustically, the ideally expanded nozzle shows little if any sign of BBSN, whereas the convergent 

nozzle shows both strong BBSN and screech. The jet mixing noise at polar angle 90° is a bit louder for 

the convergent nozzle, a result that is at odds with the similarity in the turbulence measured in the 

downstream portions of the two jets. More of an anomaly is the 4 dB increase in the OASPL of the 

convergent-divergent nozzle at the far aft angles. This was forewarned above, and is presumably caused 

by a difference in the convection velocity and hence strength of the Mach wave emission which begins to 

dominate the aft angle noise at this Mach number.

Similar trends hold for the flow fields when the Mach number is increased to M = 1.5 (Figure 18) 

using a convergent-divergent nozzle designed for this Mach number (M5). The convergent nozzle has 

strong shock cells, BBSN, and screech. The convergent-divergent nozzle has only very faint shock cells 

and no significant BBSN. Again the initial shear layer of the convergent nozzle is thicker than the ideally 

expanded nozzle flow, and the potential core of the convergent nozzle is 10 to 15 percent shorter than that 

of the ideally expanded nozzle. The turbulence levels, and perhaps more notably the volume of 
NASA/TM—2012-217250 16

turbulence, at the end of the potential core are reduced somewhat in the M5 nozzle with a corresponding 

reduction in the jet mixing component of the acoustic spectrum. Interestingly, the OASPL of the 

convergent-divergent nozzle remains lower than that of the convergent nozzle at the far aft angles, unlike 

the M = 1.4 case above.

C4S1

M4S1

Figure 17.—Splitter area ratio = 0.2, convergent C4 versus convergent-divergent M4 nozzles (setpoint 9020; 

M=1.4, Tt,c/T∞=1.23).

C4S1

M5S1

Figure 18.—Splitter area ratio = 0.2, convergent C4 versus convergent-divergent M5 nozzles (setpoint 10020; 

M=1.5, Tt,c/T∞=1.23).
NASA/TM—2012-217250 17

C4S1

M5S1

Figure 19.—Splitter area ratio = 0.2, convergent C4 versus convergent-divergent M5 nozzles (setpoint 10050; 

M=1.5, Tt,c/T∞=1.97).

When the M = 1.5 jet is heated (Figure 19) the strong shock cells remain present in the convergent 

nozzle flow, but the shear layer is not as thick. In fact, it has become strongly distorted by the shock cells. 

The potential core difference has also disappeared. In the acoustic spectra, the convergent M = 1.5 hot jet 

has some screech, presumably because the temperature is not quite high enough to remove the feedback 

condition. The turbulence at the end of the potential core is just slightly less in the ideally expanded jet, 

which seems to correspond to the 1 dB reduction of the jet mixing noise component in the noise spectra at 

90°. Note how the OASPL in the aft angles are very strongly changed by the nozzle. There is very strong 

peak noise downstream of the convergent-divergent nozzle relative to the shocked jet.

Splitter Area Ratio = 2.0

When the splitter area ratio is increased from 0.2 to 2.0, the outer sheath of cold flow become much 

more substantial. There is a significant length of internal shear layer between the core and bypass flows, 

and this brings about a strong change in the shape of the turbulence in the jet plume. And this change is 

different for the convergent and ideally expanded nozzle flows.

This change can be seen right away (Figure 20) for the hot M = 1.19 jets from convergent C4 and 

ideally expanded M2 nozzles. Focusing on the differences between convergent and ideally expanded 

nozzles, we note a significant shift in the potential core and in the axial distribution of turbulence. The 

ideally expanded flow actually has slightly higher peak turbulence levels (u′/Uic = 0.15 versus 0.13 for the 

convergent nozzle), a fact reflected in the slightly elevated peak jet mixing noise of the broadside acoustic 

spectrum. This is different from the case of splitter area ratio = 0.2 in Figure 16.

In comparison with the splitter area ratio = 0.2 case, the convergent jet noise spectrum shows a bit 

more BBSN, mostly because the mixing noise, which presumably scales on the fully mixed velocity, is 

reduced with the increase in cold fan flow. There is still no sign of screech in this hot jet case. The 

OASPL directivity is somewhat complicated by the BBSN, for while the jet mixing noise is greater in the 

convergent-divergent nozzle at all angles, the BBSN elevates the OASPL at upstream and broadside polar 

angles.
NASA/TM—2012-217250 18

C4S3

M2S3

Figure 20.—Splitter area ratio = 2, convergent C4 versus convergent-divergent M2 nozzles (setpoint 8060; M=1.19, 

Tt,c/T∞=2.76).

C4S4

M2S4

Figure 21.—Splitter area ratio = 3, convergent C4 versus convergent-divergent M2 nozzles (setpoint 8060; M=1.19, 

Tt,c/T∞=2.76).

Splitter Area Ratio = 3.0

Moving on to even higher bypass ratios, the bypass sheath thickens a bit more, accentuating the effect 

seen going from 0.2 to 2.0. So in Figure 21 the potential core of the fully expanded nozzle ends even 

further upstream than the convergent nozzle, bringing with it the peak turbulence regions. The radial shift 

in the peak turbulence with increasing axial distance is very pronounced as the two shear layers merge at 
NASA/TM—2012-217250 19

almost the same location as the end of the potential core. Here the convergent nozzle has a much longer 

potential core than the C-D nozzle, but the peak turbulence levels are very comparable. 

In the noise spectrum, the mixing noise of the C-D nozzle is some 3 dB higher at 90° to the jet 

although the two jets have very similar peak noise in the aft angles. The BBSN of the convergent nozzle 

adds to the jet mixing noise at broadside angles to raise the OASPL above that of the convergentdivergent nozzle. The dominance of the jet mixing noise at the aft angle at this relatively low Mach 

number causes the OASPL of the convergent-divergent nozzle to be greater at these angles. 

Discussion

In comparisons of the flows from convergent nozzles with splitter area ratio = 0.2 at different Mach 

numbers and temperatures, there seems to be a correlation between screech and enhanced turbulence in 

the initial shear layer (see Figure 15 through Figure 18). This is reasonable since the screech entails 

amplification of large-scale vortical structures in the shear layer. It is unlikely that Reynolds-averaged 

simulations will reproduce this effect if it is indeed an effect of the unsteady flow amplification of 

screech.

Implicit in the discussions of the jet mixing noise, as observed in the spectra at 90° to the jet, is that 

this noise is directly correlated with the turbulent velocity. This presumption comes from many 

theoretical developments, especially related to acoustic analogies that show how the noise is related to the 

turbulent kinetic energy to the 7/2 power. While this correlation held loosely for most of the cases studied 

here it clearly doesn’t hold for the noise in the aft directions at the highest jet speeds. For example, in 

Figure 19 the rms velocity contours of the two nozzles were very similar in level and area (hence 

volume), and the mixing noise component of the spectra at 90° were comparable. However, the 

convergent-divergent nozzle was over 10 dB louder at its peak angle than the convergent nozzle. This 

demonstrates the fact that just knowing the turbulence levels of the jets will not be adequate to predict the 

noise at these downstream angles for jets with Mach numbers greater than roughly 1.5.

The choice of core velocity for normalizing the flow and sound data was not arbitrary. While this is 

clearly a good choice for the mean velocity contours it is not such a clear choice for the turbulence plots. 

If anything one would expect that the peak levels in the downstream portions of the jet would depend on 

the local gradient that in turn is dominated by the fully mixed velocity. This was true. Compare the 

contour plots of rms turbulent velocities in Figure 22 with the same data in Figure 13. The peak 

turbulence levels in the jets with different splitter area ratios match well when normalized by massweighted velocities (Figure 22), whereas they were distinctly different when normalized by core velocity 

in Figure 13.

C4S1

C4S3

C4S4

Figure 22.—Turbulent velocity, u′/Uim, convergent nozzle C4, 

various splitter area ratios (setpoint 8060; M=1.19, 

Tt,c/T∞=2.76). Note normalization with fully mixed ideally 

expanded velocities.
NASA/TM—2012-217250 20

There was no similar good scaling of the acoustic data with velocity. Simple power laws for noise are 

very dependent on polar angle even at subsonic jet speeds, and become very erratic as convection speeds 

approach sonic. Toss in the low power scaling of the BBSN noise component and no simple collapse of 

the acoustic data could be achieved with a single velocity scaling law. Using the core velocity for 

Strouhal number at least gave a constant factor when comparing jets of the same core velocity. 

Summary

Both acoustic and particle image velocimetry measurements have been made on internally mixed 

exhaust systems running at Mach numbers between 1 and 1.5. The exhaust systems had different internal 

area ratios and used both convergent and ideally expanded nozzles. The experiment was designed to hold

the Mach number of the flow constant to keep the shock structure of the jet fixed. Changing the 

temperature of the core then changed the jet velocity and changing the internal splitter area ratio changed 

the turbulence distribution and the mixed velocity of the fully mixed jet. The flow measurements showed

that the common shock structure of the plume was barely affected by the changes in area ratio or velocity 

ratio for a given Mach number. The noise was more significantly affected. In the analysis of the data 

turbulence intensities were shown to scale on local mean gradients, as expected, and the jet mixing noise 

measured broadside to the jet generally scaled with the peak turbulence intensities, again as expected. At 

the high end of the Mach number range the independence of Mach wave emission noise sources from the 

turbulence was also captured. The documented changes in local gradients, and hence turbulence and noise 

source strength, caused by varying the area ratio of the dual stream flows, will be of use in jet noise 

modeling and in understanding how to manipulate jet velocity profiles to minimize noise production in 

future aircraft. The detailed flow data will also be useful in validating Reynolds-Averaged Navier-Stokes 

prediction methods, and Large Eddy Simulation codes.

References

1. Bridges, J. “Broadband Shock Noise in Internally-Mixed Dual-Stream Jets,” AIAA paper 2009-3210, 

(2009).

2. Tanna, H. K. “An Experimental Study of Jet Noise, Part II: Shock Associated Noise.” J. Sound and 

Vibration, 50(3), pp. 429-444, (1977).

3. Viswanathan, K., Alkislar, M., and Czech, M. “Characteristics of the Shock Noise Component of Jet 

Noise,” AIAA Paper 2008-2835, (2008).

4. Tam C. K. W. and Burton D. E. “Sound generated by instability waves of supersonic flows. Part 1. 

Two dimensional mixing layers,” J. Fluid Mech. 138, pp. 249–271, 1984.

5. Shields, F. D. and Bass, H. E. “Atmospheric Absorption of High Frequency Noise and Application to 

Fractional-Octave Band,” NASA-CR 2760, (1977).

6. Bridges, J. and Wernet, M.P. “Validating LES for Jet Aeroacoustics,” AIAA Paper 2011-0017 

(2011).

7. Tam, C.K.W., Chen, P., Seiner, J.M. “Relationship between instability waves and noise of high-speed 

jets.” AIAA J. 30, pp. 1747-52 (1993).

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

While externally mixed, or separate flow, nozzle systems are most common in high bypass-ratio aircraft, they are not as attractive for use in 

lower bypass-ratio systems and on aircraft that will fly supersonically. The noise of such propulsion systems is also dominated by jet noise, 

making the study and noise reduction of these exhaust systems very important, both for military aircraft and future civilian supersonic 

aircraft. This paper presents particle image velocimetry of internally mixed nozzle with different area ratios between core and bypass, and 

nozzles that are ideally expanded and convergent. Such configurations independently control the geometry of the internal mixing layer and 

of the external shock structure. These allow exploration of the impact of shocks on the turbulent mixing layers, the impact of bypass ratio on 

broadband shock noise and mixing noise, and the impact of temperature on the turbulent flow field. At the 2009 AIAA/CEAS Aeroacoustics 

Conference the authors presented data and analysis from a series of tests that looked at the acoustics of supersonic jets from internally mixed 

nozzles. In that paper the broadband shock and mixing noise components of the jet noise were independently manipulated by holding Mach 

number constant while varying bypass ratio and jet temperature. Significant portions of that analysis was predicated on assumptions 

regarding the flow fields of these jets, both shock structure and turbulence. In this paper we add to that analysis by presenting particle image 

velocimetry measurements of the flow fields of many of those jets. In addition, the turbulent velocity data documented here will be very 

useful for validation of computational flow codes that are being developed to design advanced nozzles for future aircraft. 

15. SUBJECT TERMS

Airport noise; Jet noise; Aeroacoustics; Large eddy simulation; Supersonic aircraft; Particle image velocimetry 

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