NASA/CR--2002-210823/VOL1 EDR No. 18580

Lobed Mixer Design for Noise Suppression

Acoustic and Aerodynamic Test Data Analysis

Vinod G. Mengle and William N. Dalton

Rolls Royce Allison, Indianapolis, Indiana

July 2002
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NASA/CR--2002-210823/VOL1 EDR No. 18580

Lobed Mixer Design for Noise Suppression

Acoustic and Aerodynamic Test Data Analysis

Vinod G. Mengle and William N. Dalton

Rolls Royce Allison, Indianapolis, Indiana

Prepared under Contract NAS3-27394, Task Order No. 6

National Aeronautics and

Space Administration

Glenn Research Center

July 2002
Acknowledgments

We would like to acknowledge the efforts of several people at NASA Glenn Research Center, Cleveland, Ohio and

Allison Advanced Development Company (AADC), Indianapolis, Indiana: Dr. James Bridges and Ms. Kathleen Boyd

of NASA Glenn for overseeing and running the acoustic tests at NASA, and for processing all the acoustic data

obtained there; Mr. Eugene Krejsa of NASA Glenn for encouragement to do the mixer tests and theoretical modeling

and for several discussions during the annual NASA Advanced Subsonic Technology Engine Noise workshops;

Mr. Allen Barta, Mr. David Baker, and Dr. Bailey Vittal of AADC for help in the design of lobe mixers and the

analysis of plume data; Mr. Frederick Smith and Dr. John Kingsley of AADC for help in mixer geometry definitions

using CAD and CFD analysis. The mixer/nozzle models and the static thrust tests were made by Aero Systems

Engineering's FluiDyne Laboratories at St. Paul, Minnesota, for which we would like to thank Mr. Randy Merchen

and Dr. Dean Long. The first author, Dr. Vinod G. Mengle, was the principal scientific investigator responsible for the

concept and design of the new mixers ("boomerang scallops" and the "tongue mixer"), and analysis of all the acoustic

and aerodynamic data. The second author, Mr. William N. Dalton, was the Rolls-Royce Allison program manager for

this task. For most of the task-order period (from April 1995 to December 1997) Ms. Kathleen Boyd was the NASA

technical monitor and, thereafter (from January 1998 onwards), Dr. James Bridges was the technical monitor.

Note that at the time of research, the NASA Lewis Research Center

was undergoing a name change to the

NASA John H. Glenn Research Center at Lewis Field.

Both names may appear in this report.

Contents were reproduced from the best available copy

as provided by the authors.

Trade names or manufacturers' names are used in this report for

identification only. This usage does not constitute an official

endorsement, either expressed or implied, by the National

Aeronautics and Space Administration.

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

Available from

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

Available electronically at http: / / gltrs.grc._aasa.gov/GLTRS
Preface

The final report is written in two volumes.

In the first volume we present the design

philosophy of the new lobe mixers tested,

and then analyze the results of various

acoustic and aerodynamic tests done at

NASA Lewis Research Center, Cleveland,

Ohio and Aero Systems Engineering

FluiDyne Laboratories, St. Paul, Minnesota

over a period of three years (1995-1997).

The second volume is a compilation of the

plume survey data, the aerodynamic data for

the acoustic tests and the acoustic data.

NASA/CR 2002-210823/VOL1 iii

Table :of Contents

Volume I

Pret_ce

Summary

1, Introduction

1.I Backgound and Objectives

1.2 Overview

2. Mixe>Nozzle Designs and CFD Analysis

2,1

2,2

2.3

2,4

General Model Assembly and Common Flow Lines

Previous Existing Mixers

New Mixers

2,3,1 Conceptual Selectmn

2.3.1, I Scalloped Family of Mixers

2,3,1,2 The Tongue Mixer

2,3,2 Design of the New Mixers

2.3,2A Unscatloped 20 Lobe Mixer

2.3.2,2 Implementation of the New Rules for Scalloping

2,3.2,3 Design Guidelines for the Tongue Mixer

CFD Analysis of New Mixers

2,4,1 Code, Grids and Boundmy Conditions

2.4.2 ResuIts

in

ix

1

1

2

5

5

7

9

9

9

14

18

18

18

28

31

35

38

55

NASA/C_2002-210823/VOL 1 Vi
3.1 Acoustic Tests

3.2 Aerodynamic Tests

3.3 Plume Surveys

4. Aerodynamic Results and Analysis

4.1 Aerodynamic Performance

4.1.1 Definition of Various Terms

4.1.2 Thrust Coefficients

4.1.3 Nozzle Discharge Coefficients

4.1.4 Bypass Ratios

4.2 Typical Plume Survey Results

4.2.1 Nozzle Exit Plane Survey

4.2.2 The Tongue Mixer and Mixer with Cutout Lobes

4.2.3 Plume Downstream Evolution

4.2.4 Effect of Scalloping

4.2.5 Effect of Nozzle-Length

4.2.6 Center-line Velocity Comparisons

5. Acoustic Results and Analysis

5.1 Overview

5.2 Mixers with Reference Nozzle Length at Baseline Take-Off Condition

5.2.1 Static Free-Jet Data

5.2.2

5.2.1.1

5.2.1.2

5.2.1.3

5.2.1.4

12 Lobed Mixers

Lobe Number Effect

Effect of Scalloping

Probable Causes of Noise Reduction

Flight Effect: Free-Jet Mach No. 0.2

5.2.2.1

5.2.2.2

5.2.2.3

5.2.2.4

12-Lobed Mixers

Lobe Number Effect

Effect of Scalloping

Confluent Mixer Benefits

55

62

68

75

75

75

77

81

83

87

87

93

96

96

100

105

113

113

114

115

115

121

126

130

133

133

137

141

145

NASA/CR 2002-210823/VOL1 vi
5.3

5.4

5.2.3 Flight-Effect:OtherFreeJetMachNumbers

MixerswithReferenceNozzle-LengthatHigherNozzle

PressureRatioConditions

5.3.1 TO# 2 and# 3with FreeJetMachNo.0.2

5.3.1.1

5.3.1.2

5.3.1.3

12-LobedMixers

LobeNumberEffect

EffectofScalloping

5.3.2 EffectofFree-JetMachNumberatTO# 3

DiagnosingHigh-FrequencySourceLocationswithChangesin

Free-JetSpeeds

5.4.1

5.4.2

5.4.3

RayTheoryApplication

TheNewDiagnosticMethodfor"Excess"NoiseSources

Applicationto OperatingConditionTO# 3

5.5

5.6

Effectof Nozzle-Length

5.5.1 OverallAcousticMetrics

5.5.2 PolarSPL-Spectra

5.5.3 CompetingMechanismsfor LocalMinimumof Noise

FlyoverEPNL

5.6.1 Effectof Free-JetMachNumber

5.6.1.1 Effectof ObservableAngularLimits onEPNL

5.6.1.2 CollapseofEPNLPlots

5.6.2 Effectof NozzleLength

5.6.3 ComparisonofEPNLBetweenMixers

5.6.4 QuietestMixer-NozzleCombinations

6.ConcludingRemarks

146

157

157

157

160

160

165

170

170

177

180

186

187

195

206

208

209

209

211

219

226

232

239

NASA/CR 2002-210823/VOL1 vii
7. References

Appendices:

A

B

C

D

Geometrical Details of Mixers and Nozzles

Refraction Correction Procedure for Free-Jet/Ambient Shear-Layer

and Angular Interpolation Scheme

Refraction Angle, Convection Angle and Transmission Coefficient

Changes Due to Free-Jet Speed Changes

Comparison of SPL's at Different Free-Jet Mach Numbers and

"Shifted" Angles at TO # 3

243

245

269

271

277

NASA/C_2002-210823/VOL 1 viii
Summar_

A comprehensive database for the acoustic

and aerodynamic characteristics of several

model-scale forced lobe mixers of bypass

ratio 5 to 6 has been created for mixed jet

speeds up to 1080 ft/s at typical take-off

conditions of small turbofan engines. The

flight effect is simulated with a free jet

surrounding the model nozzle with Mach

numbers ranging from 0.0 to 0.3. The static

thrust performance and jet plume flow is also

examined at typical take-off and cruise

conditions. This data is scaled for a nozzle

with 29 in. diameter, and the effect of several

lobe-mixer and nozzle parameters is

examined in terms of flyover noise at

constant altitude and, also, noise in the

reference frame of the nozzle to understand

the changes in the noise-source

characteri sties.

Several new concepts, mechanisms, methods

and findings are reported here for the first

time regarding such mixers. The new

concepts include a rational method for

scalloping lobe sidewalls leading to

"boomerang" scallops, and an extreme limit

of scalloping leading to the "tongue" mixer.

A new diagnostic method to detect "excess"

internal noise sources due to fan/core mixing

is given; it uses appropriate "shifted" angles

for comparison of non-Doppler-shifted noise

at several free jet speeds. Another new

method is found to extrapolate known flyover

noise for lobed mixers from one flight speed

to different flight speeds.

The effects of scalloping, number of lobes

and mixing-length on noise are

systematically examined. Compound effects

of lobe penetration and fan-to-core area ratio

are also examined. For all forced mixers, the

flyover effective perceived noise level

(EPNL) is found to scale with "net" thrust

when tested at different free jet speeds.

Different types of lobe mixers are found to be

most effective at different thrust levels to

provide maximum noise benefit over a

coaxial nozzle. At about 9500 lb net-thrust a

maximum of 3 EPNdB noise suppression is

accrued with a deeply scalloped lobe mixer

having higher number of lobes (20) and

higher lobe penetration. However, at a netthrust of 5000 lb, a maximum of only 1

EPNdB benefit is obtained from a mixer with

fewer lobes (12), lower lobe penetration and

lobe sidewall cutouts. These benefits

increase by about 1 EPNdB when

comparison is made on the basis of same

mixed jet velocity rather than same net thrust.

Thus, for example, at approximately 1060

ft/s, corresponding to the highest thrust levels

tested, a maximum noise suppression of 4

EPNdB is achieved by the 20-lobe deeply

scalloped mixer. And, in general, the noise

benefits increase at higher and higher thrust

levels or mixed jet speeds.

Over most of the range of thrusts or jet

speeds tested none of the unscalloped lobe

mixers are quieter than the coaxial jet. Thus

"scalloping" is the most important mixer

geometrical parameter identified for reducing

noise. Deep scalloping reduces the typical

low frequency mixing noise without

increasing the annoying mid-to-high

frequency noise when compared to the

unscalloped mixer, and gives a benefit of 1.8

EPNdB to 3 EPNdB depending on the thrust

level. The "gradual" introduction of

streamwise vorticity by appropriate

scalloping is found to bc the key to this

behavior. However, scalloping also increases

NASA/CR 2002-210823/VOL1 ix
thebypassratio.Hence,theeffectivelobefan

and core areas should be resized

appropriatelyfor desiredbypassratios in

applications to specific engine cycles.

Scalloping also introduces slight thrustcoefficientlossattypicalcruiseconditionsas

comparedto unscallopedmixers,especially,

for mixers with higher number of lobes.

Hence,the amountof scallopingshouldbe

usedas a tradingparameterbetweennoise

reductionand thrust loss. However, with

fewerlobesandlowerpenetration,evenwith

lobe cutouts (which act like scallops),an

improved cruise thrust coefficient is

obtained.

One surprising new result regardingthe

effectof nozzlelengthis alsofound: for a

givenmixer,thereexistsan"optimal"mixing

lengthor nozzlelengththat createsa local

minimum in noise. A reduction in the

mixing-lengthby 25% givesa noisebenefit

of asmuchas3 EPNdBfor theunscalloped

mixer comparedto the baselinemixing

length (which is 1.5 times the nozzleexit

diameter). A competing mechanismof

acoustic unshielding vs aerodynamic

unshielding is proposed for this behavior.

Using optimal nozzle length can be very

helpful in reducing noise with unscalloped

mixers if thrust loss due to scalloping is

unacceptable. However, its effect on boat-tail

angle and cruise thrust coefficient is not

examined. Also, the implications of all these

lobe and nozzle geometric changes on thrust

specific fuel consumption are not examined.

NASA/CR 2002-210823/VOL1 x
Chapter I

Introduction

1.1 Background and Objectives

Future commercial applications of the

turbofan engine will be required to meet

increasingly stringent noise abatement

criteria, primarily, due to an increase in

aircraft traffic near airports. Many of the

current turbofan engines are based on

thermodynamic cycles with a bypass ratio of

five to six as opposed to the lower bypass

ratio of 1.5 or so in older engines. With very

high mixed jet velocities for older lower

bypass engines, jet-noise is a dominant

contributor to total engine generated noise,

especially at take-off (TO). However, even

with the current higher bypass ratios, and

consequently reduced mixed jet velocities,

jet noise continues to be a significant

contributor.

Reduction of jet noise has been sought

earlier by mixing the hot core flow and the

cooler fan flow before they exit through the

nozzle. A more uniform flow at the nozzle

exit plane is supposed to lead to reduced

noise levels. Uniform flow at the nozzle exit

plane is also known to yield better cruise

thrust efficiency thermodynamically than

partially mixed flow or separate unmixed

flow nozzle systems. In fact, that is the

primary reason for mixing the two flows

internally. However, the overall noise

benefit or penalty incurred due to internal

mixing to achieve the uniform exit flow is

not well understood. The actual level of

noise abatement realized in a specific

application must be critically related to the

manner and extent to which internal mixing

is achieved. Traditionally, the mixing

between the hot core flow and the cold fan

flow has been achieved with so-called

"lobed" mixers which are essentially

convolutions at the end of the splitter plate

that separates the two flows. However, there

does not exist a significant acoustic database

for such lobed mixers at higher bypass

ratios. In addition, a good correlation

between the aerodynamic thrust performance

and acoustic performance for such mixers

does not exist. This task was conceived to

address these two needs as part of NASA's

Advanced Subsonic Technology (AST)

Program for Engine Noise Reduction,

thereby permitting the design of thrustefficient noise suppression mixers for

modern turbofan engines.

Lobed mixers were studied extensively

during the mid-seventies and early eighties

as a means for improving thrust efficiency,

for example, under NASA's Energy

Efficient Engine (E 3) Program. Both far

field noise data (1) and detailed measurement

of fluid-dynamic and aerodynamic

properties (2-4) for lobed mixers have been

reported in the literature. Previously

published noise data (1,5) is typically for low

bypass ratio engine cycles (around 1.5) with

high ideally-expanded mixed jet velocities

of 1330 ft/s or so. Hence, a principal

objective of this task was to extend the

mixer acoustic data base to higher bypass

ratios and operating conditions typical at

take-off for modern turbofan engines. At the

same time, we also wanted to study the

aerodynamic thrust performance at cruise

conditions for the same mixers so that the

trade-off between take-off noise-suppression

NASA/CR 2002-210823/VOL1 1
and cruise thrust efficiency loss, if any, can

be quantified.

Before entering this program, Rolls-Royce

Allison had already designed, fabricated and

aerodynamically tested four mixer

configurations. These included three lobe

mixers and one baseline annular or confluent

configuration. Three of these configurations

were originally developed under company

funding, while the fourth was completed

under Task 15 of NASA Contract NAS3-

25950. The geometric variations in these

mixers were not defined by parametrically

varying only one geometric variable to

examine its incremental effect on the

aerodynamic or acoustic results. However

these designs are representative of current

technology levels and were included in the

current program for their intrinsic value in

expanding the experimental database.

Early Rolls-Royce Allison experience with

lobed mixers with cutouts in the lobe

sidewalls, showed that these cutouts may be

acoustically beneficial. This is in line with

work done on scalloping of lobes (partial

removal of lobe sidewalls) by Pratt &

Whitney and General Electric Aircraft

Engines. However, no systematic study has

been published examining the effect of

scalloping on both acoustics and

aerodynamic thrust performance especially

at high bypass ratios. Hence, one of the

objectives of this task was to systematically

study the effect of scalloping of lobe

sidewalls on noise suppression and thrust

performance. This also led us to designing

and testing a unique new mixer concept

called the "tongue" mixer.

Since the distance between the mixer exit

plane and the nozzle exit plane, called the

"mixing- length," governs the axial

evolution of flow and the consequent noise

distribution inside and outside the nozzle it

is expected to be an important parameter.

Hence, another objective of the program was

to examine the relationship between nozzle

length and the jet noise generated by the

various lobed mixers.

1.2 Overview

The tests in the program were conducted

over a period of three years from 1995 to

1997. The acoustic tests were performed in

two phases at NASA Lewis Research

Center's (now Glenn Research Center)

Aeroacoustic Propulsion Laboratory (APL)

Nozzle Acoustic Test Rig (NATR) facility.

The aerodynamic tests were conducted at

Aero Systems Engineering's (ASE)

FluiDyne Aerotest Laboratory static thrust

stand in St. Paul, Minnesota and were also

done in two phases. This report summarizes

the principal results of the voluminous data

collected from these four different tests and

the lessons learned.

For the acoustic tests, data was obtained

over a range of free-jet Mach numbers,

which simulates the forward motion of the

aircraft at take-off. This not only allows us

to compare noise benefit at take-off speeds,

as opposed to static conditions, but also to

approximately locate the different noise

sources associated with the different

frequency bands, as explained in later

chapters. The operating conditions were

chosen to cover typical take-off operating

conditions for modem turbofan engines. All

aerodynamic thrust tests were conducted

under static (or no free-jet) conditions, and

included both cruise conditions and most of

the take-off operating points. Plume surveys

outside the nozzle, including total pressure,

total temperature and static pressure were

also conducted so that possible trends could

NASA/CR 2002-210823/VOL1 2
be obtained between the acoustic

characteristicsandtheplumeevolution.

This volume summarizesthe designof a

seriesof lobedmixersdevelopedspecifically

for this program and the results of the

acousticandaerodynamictestsdonefor all

the mixers. The secondvolume collates

detailed plots obtained from the plume

surveysandtheaerodynamicdatacollected

duringtheacoustictests.

Chapter 2 of this volume discussesthe

conceptualdevelopmentand new design

rulesfor the scallopedmixersandthe new

tonguemixer. It alsobriefly describesthe

computationalfluid dynamic(CFD) results

for someof the new mixers and givesthe

geometricalpropertiesof all the models

tested.

Chapter 3 summarizesthe experimental

facilities for the acoustic and the

aerodynamictests,andthemannerin which

the datawasprocessed.It alsocollatesall

the test matricesfrom the different test

phases.

Chapter4 summarizesthe aerodynamictest

resultsobtainedfrom the staticthruststand

at ASEFluiDyne,aswell asthoseobtained

during the acoustictestsat NASA Lewis.

Some selectedplume surveys are also

discussedto facilitate the acoustic data

analysis,andthefull jet plumedatais given

inVolume2.

Chapter 5 summarizesthe acoustictest

resultsfor all themixerstested.It discusses

the parametric effects of operating

conditions, surrounding free-jet Mach

number,and nozzlelengthon eachmixer.

Comparisonbetweendifferentmixersat the

sameoperatingconditionsis alsopresented

here to bring out the effect of different

geometricparametersuchasscallopingand

numberof lobes. We also presenta new

way of processingthe acoustic data to

deducethe presenceof excess"internal"

mixing noise,that is, noiseproducedinside

thenozzleductdueto themixingof fanand

coreflow, asdistinguishedfromtheclassical

jet mixingnoisedueto sourcesin theplume

outsidethe nozzle. Certainrulesof thumb

for improving noise suppressionwhile

minimizing the adverseeffect on thrust

performanceare also summarizedin this

chapter.

Finally, in Chapter6, we summarizethe

main conclusions from these tests.

Recommendationsfor implementationin

full-scaleenginesandfurtherusefultestsare

alsogiven.

Theappendicesattheendof this volume are

intended to supplement the main body of the

text by providing additional detailed

derivations, data, figures or tables.

NASA/C_2002-210823/VOL 1 3

Chapter 2

Mixer-Nozzle Designs and CFD Analysis

Under this program, a total of eight mixers

were tested. Four of the configurations were

existing designs from previous tests, while

four new mixers were designed, fabricated

and tested. All testing was carried out at

model scale. This allowed utilization of an

existing test rig for performance

measurements at the ASE FluiDyne facility.

Additional hardware was fabricated as part

of the program to allow adaptation of the

mixed nozzle flow rig to the NASA NATR

facility.

In this chapter, we briefly outline the history

and geometry of the four existing mixers,

describe the concepts that led to new rational

design rules for scalloping and the unique

tongue mixer, and finally present the

geometrical details of the four new mixers

which were fabricated. The main body of

this chapter has the schematic figures (or

photos), whereas all detailed geometrical,

dimensional figures are collated in Appendix

A. The results of a CFD analysis for one of

the new scalloped mixers and the tongue

mixer are also briefly described to validate

the conceptual picture and to help analyze

the acoustic data later.

2.1 General Model Assembly and

Common Flow Lines

Figure 2.1 shows the general arrangement of

the mixer-nozzle configurations and

geometrical definitions. All the mixernozzle configurations had common inner

flow lines for the fan nozzle or shroud.

However the outer (external) surface of the

fan-nozzle was different for the models

tested in the acoustic rig (NATR) and the

static thrust stand. This is due to the

requirement of providing flight simulation

effects via a free-jet surrounding the nozzle

in the NATR which demanded a smoother

outer surface. Figure A.1 and A.2 show the

detailed model assemblies for the acoustic

rig and the thrust-stand respectively. Both

figures also show the location of the

pressure and temperature rakes used at the

charging stations and Figure A.2 further

shows the additional rake used to measure

total pressure (PT) near the mixer exit plane.

The exit P-r rake was removed during thrust

measurements.

Figures A.3 and A.4 show the detailed inner

flow lines of the baseline or reference fan

nozzle downstream of the reference station 0

depicted in Figure A.2. In general, it forms

a converging area from about 10.290 in. in

diameter at station 0 to 7.245 in. in diameter

at the nozzle exit plane (station 57), with a

nominal mixing length, L, of 11 in. and

gives a baseline mixing length-to-mixing

plane diameter ratio (L/Drop) of about 1.10.

Three additional fan nozzles, which were

compatible with all the mixers, were also

fabricated. Each of these nozzles

maintained a constant exit diameter but

varied the mixing length (see figure A.5);

this varies the axial rate of contraction for

each nozzle. The additional mixing lengths

tested were 1.363, 0.818, and 0.545, when

normalized to the mixing-plane diameter.

This corresponds to nominal variations of

+25%, -25%, and -50% from the baseline

mixing length. The coordinates of the

additional nozzles are provided in Figures

A.6, A.7 and A.8.

NASA/CR 2002-210823/VOL1 5
Ref.

 o z,e

Lobed Mixer D/2

Center-Cone _ 1

C.L. --" ,r, \\'.'_ _

Figure 2.1 Schematic of mixer-nozzle configuration and geometrical definitions.

Mixer Mixer

Code

Confluent CONF

12 Lobe with Cutouts/Low 12CL

12 Lobe Unscalloped/High 12UH

16 Lobe Unscalloped/High 16UH

Lobe

No.

12

12

16

Scallop Lobe Lobe

Penetration Length

(Hn_rlmp) (L,_/Dmp)

Triangular 0.48 0.33

Cutouts

None 0.68 0.34

None 0.72 0.34

Table 2.1 Previous Existing Mixer Parameters

Area Mixing

Ratio Length

(AtIA¢) (I-JDmp)

2.554 1.15

2.637 1.13

2.637 1.09

3.199 1.09

NASA/CR 2002-210823/VOL1 6
Two center tailcone configurations were

employed. The first shown in Figure A.9,

had a length of 5.935 in. and was used with

all the lobed mixers. The second one,

shown in Figure A.10, was shorter and used

only with the confluent nozzle

configuration. This tailcone had a length of

4.258 in. and slightly different shape. The

two tailcones are shown side-by-side in

Figure A.2. The resulting confluent mixer

configuration, described below, had a

slightly smaller fan-to-core area ratio but

was designed such that, after accounting for

the boundary-layer thickness, the mass-flow

rates would match with the original lobed

mixers. From a viewpoint of systematic

variation of geometric parameters, it would

have been preferable not to have varied the

tailcone at all for the confluent mixer, but

rather to have varied the confluent mixer

geometry itself to match the mass-flow rates

or area ratios of the lobed mixers. However,

due to the availability of previous

aerodynamic, as well as, full-scale acoustic

data (6-8) we continued using the old

confluent-mixer/tailcone combination

instead of building a new one. This fact

needs to be borne in mind when comparing

acoustic or aerodynamic data of the lobed

mixers with that of the confluent mixer.

[Note that the uniquely truncated tailcone

shown in Figure A.2 (one with cross-hatched

cross-section) was a carry-over from an old

program and was not used in this task.]

2.2 Previous Existing Mixers

Four existing mixer configurations from

previous programs were available. These

were:

1. Annular or confluent mixer which acted

as the reference configuration.

2. 12-lobed mixer with cutouts in the lobe

sidewalls and low penetration (that is,

low lobe-height-to-maximum possible

,

4.

lobe height in the duct (Hm]Hmp in Figure

2.1)).

12-lobed, unscalloped, high-penetration

mixer.

16-lobed, unscalloped, high-penetration

mixer.

Figure 2.2 shows the relative shapes of the

mixers and Table 2.1 lists their nondimensional geometric properties as defined

in Figure 2.1. As shown in Table 2.1, we

will use certain symbols for lobed mixers for

convenience: ab c d where ab represents the

number of lobes, c represents the type of

scalloping or cutouts (U for unscalloped, C

for cutouts) and d represents relative lobepenetration (L for low, H for high lobe

penetration).

The two 12-lobed mixers, 12CL and 12UH,

differ in two aspects: presence of sidewall

cutouts and lobe penetration with fan-to-core

area ratio, lobe number, and lobe length held

constant. The last two configurations in

Table 2.1, 12UH and 16UH, differ in terms

of both number of lobes and fan-to-core area

ratios, however, they are similar in terms of

lobe penetration and lobe length. Between

12UH and 16UH mixer at least two

parameters vary from one mixer to the other.

As a result, it is not possible to isolate the

effect of a particular geometric parameter on

the resulting aerodynamic and acoustic data.

The detailed geometrical characteristics of

these four mixers is provided in Figures

A. 11 through A. 17.

Before discussing the new mixers, let us

make a note of some of the characteristics of

the original mixers. Firstly, the confluent

mixer, CONF, has its own tailcone different

from that for all other mixers, as mentioned

before. Secondly, only the cut-out version

of the 12 lobed, low-penetration mixer

(12CL), shown in Figure A.14, was used in

NASA/C_2002-210823/VOL 1 7
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this program. It was made from a previous

uncut version whose flow-lines are given in

Figure A.13. The approximate triangular

cutouts in the lobe sidewalls of 12CL were

made where the sidewalls were flattest. The

fillet radii at its 3 apexes were made simply

to reduce stress concentrations in the

sidewalls. The high-penetration 12-lobe

unscalloped mixer, 12UH, was designed

earlier to improve thrust mixing efficiency

over the uncut, low-penetration 12-lobe

mixer. The design philosophy, some CFD

results and aerodynamic test results for these

three mixers can be found in Booher et al (6)

where they are referred to as confluent,

baseline (or conventional mixer), and

advanced mixer, respectively. The last

mixer in Table 2.1, the 16 lobe mixer

(16UH), was designed for a different

operating point at lower pressure ratios and

had a much larger fan-to-core area ratio;

consequently, 16UH stands apart from the

previous mixers. This should be borne in

mind when acoustic comparisons are made

later. Results from CFD analysis and

comparison with aerodynamic test data for

all these four mixers can be found in Barta et

al (7) where the 16-lobe mixer is referred to as

the acoustic mixer.

2.3 New Mixers

Acoustic data obtained from the four

previous existing mixer configurations, in

the first phase of this task, confirmed

previous observations that aggressive, highpenetration, unscalloped mixer

configurations do suppress the low

frequency spectrum which is characteristic

of unmixed, coaxial turbofan exhausts, but

produce secondary spectral peaks at higher

frequencies which are heavily weighted in

the Perceived Noise Level (PNL) metric.

However, the moderate penetration 12 lobe

configuration with cutouts in the sidewalls

(12CL) produced the low frequency

suppression of the more aggressive designs

without incurring a penalty at the higher

frequencies. Since both the radial lobe

penetration and sidewall cutouts were

different from the other designs, it was not

possible to determine directly from the

available data which change was most

responsible for the acoustic characteristics.

However, data previously presented in the

open literature (t) for "scalloped" sidewall

mixers for a lower bypass ratio, higher

pressure ratio cycle showed similar

behavior. From these two results, it is

inferred that the sidewall scalloping is the

controlling parameter. However, existing

data (1'5) defining the acoustic impact of

sidewall scalloping was not obtained by

parametric variation of a single parameter.

For example, the data reported in Reference

1 was obtained on a series of mixers with

varying scalloping. However, the variation

in scalloping was achieved by cutting back

the mixer exit plane, introducing an

additional variation in the mixer cant angle.

To address this shortcoming, three

additional mixers were designed and tested

in this program which parametrically varied

the sidewall scalloping while holding all

other parameters fixed. A fourth mixer

represents a unique new concept referred to

as a "tongue mixer." As part of this effort, a

systematic approach to developing the

scalloped curves has been defined and rules

of thumb for its application developed. Both

the process and the resulting designs are

described in the following section.

2.3.1 Conceptual Selection

2.3.1,1 Scalloped Family of Mixers

The idea here is to test a family of mixers

which have exactly the same geometric

properties, except for the amount of

scalloping or cutouts in the lobe sidewalls.

NASA/CR 2002-210823/VOL1 9
This would entail a minimum of two mixers

an unscalloped and a scalloped one.

However, to form a trend a third mixer with

an "in-between" scallop is required. Then

the questions to be answered from an

acoustic point of view are:

i) Should we select scallops or

cutouts?

ii) Are there preferred shapes, from an

acoustic standpoint, for the scallops or

cutouts?

iii) How many lobes should we select

for this family of mixers?

We invoke certain fluid-dynamic and

aeroacoustic principles to answer these

questions qualitatively. The precise

mechanism of how scalloping may help

suppression of noise is not known but what

happens to the two flows in the scalloped

region must be quintessential to its

downstream evolution and noise generation.

We present certain new hypotheses which

will be used as guidelines to shape the

scallops or cutouts.

(i) As compared to an unscalloped forced

mixer, it is clear that scalloping, of any

type, allows the two streams (fan and

core) to interact with each other "earlier"

that is, upstream of the corresponding

unscalloped mixer exit plane, so that the

flows are already partially mixed by the

time they reach the original mixer exit

plane. Hence, if the two streams are not

parallel near the lobe sidewall, then with

scalloping, their radial velocity

components will give rise to axial or

streamwise vorticity right from the start

of the scallop. It is well known (9-11)that

streamwise vorticity helps enhance

mixing of these two flows as opposed to

mixing only due to Kelvin-Helmholtz

type of vortex-sheet instability. What

should be done differently with

scalloping is the manner in which

vorticity is initially introduced into the

flow - earlier and also "gradually". By

gradual initial mixing, we mean the axial

gradient at which net vorticity is

introduced into the flow should be

smooth and gradually increasing. This

should presumably reduce the relatively

high-frequency noise sources, so

important to perceived noise level

metrics, and which were indeed, found

in the unscalloped lobe mixer data of

Reference 1, as well as, in phase I of this

program. In unscalloped mixers, since

the two streams "see" each other for the

first time across the "full height" of the

mixer exit plane, it is presumably the

"sudden" interaction there that shears a

larger area of fluid generating the highfrequency noise.

This control over the rate of introduction

of vorticity in scalloped lobes can be

obtained by gradually increasing the

radial height of the scallop starting from

zero. This can be verified by referring to

the circulation around a loop in the

transverse plane enclosing the radial

height of the scallop (see Figure 2.3);

such a circulation is an integral measure

of the axial vorticity at each axial

station. The fact that we seek to mix the

flows gradually in the beginning may be

at odds with the desire that the flows

should be fully mixed by the end of the

nozzle exit plane because the mixinglength may not be enough. Achievement

of a fully mixed flow is desirable from a

thermodynamic thrust efficiency point of

view and also provides a reduction of

low-frequency classical jet mixing noise

from the far downstream plume.

Obviously, the distance from the mixer

exit plane to the nozzle exit plane - the

mixing-length - will also play a role in

NASA/CR 2002-210823/VOL1 10
Circulation around this loop, C,in transverse plane

Figure 2.3 Loop in transverse plane for finding circulation at an axial station inside scallops.

"Leading" Edge "Trailing" Edge

FLOW

LO_

Edge of Scalloped d J/

/.///I

,/__Flow Velocity

Edge of Scalloped ___._Flo

Lobe Wail -7.___..---

w Velocity.

-.

Figure 2.4 Definition of "leading" edge and "trailing" edge of scallop.

NASA/C_2002-210823/VOL 1 11
(ii)

howwell-mixedtheflow is atthenozzle

end and how muchthe low frequency

noiseis suppressedby ascallopedmixer.

So,in fact,weshouldalsotesttheeffect

of nozzle length. We also expect a

difference in the total pressureloss

betweenscallopedlobesandunscalloped

lobedesigns.

Thenoisemechanismdealtwithaboveis

theresultof themixingof twoflowsand

is best representedby the Lighthill

"quadrupole"type of noise sources.

Whenwe considerflows nearscalloped

edgeswe alsoneedto worry aboutthe

formation of so-called"dipole" noise

sources,which are evenmoreefficient

sound generators than quadrupoles.

Thesearethesametypeof sourcesthat

areresponsiblein fan noisefor creating

toneswhentheblade-wakesperiodically

hit the leading edge of downstream

vanes.To besure,thetrailingedgeof a

splitterplate separatingtwo flows can

alsoact asa dipole sourceof noise,but

of much less magnitude than the

correspondingleading edge source.

While designingthe scallopswe must

minimize the formation of "dipole"

noisesources.Basedon theanalogyof

leadingedgedipole sourcesin the fan

blade-vaneinteractionproblem,a design

rule is developed for scalloping to

minimizedipoleformationwhichwill be

calledthe"trailing edgerule."

First a definitionfor the "leadingedge"

in a scallop must be developed.

Analogous to wing-aerodynamics, we

define that part of the scalloped edge as a

"leading edge" with respect to a stream

if the velocity vector of that stream has

the component orthogonal to the edge

going "towards" the edge, as shown in

Figure 2.4. Conversely, if this

orthogonal component is going "away"

from the edge then we will define it as a

"trailing-edge". To minimize the dipole

noise source generation, the scalloped

edge should be shaped such that it acts

as a trailing edge throughout its length

with respect to both the streams (fan and

core) around it. This, we will call "th___ee

trailing-edge rule for scalloping." In

adopting this rule of thumb, it must be

remembered that the relative strength of

the dipole source relative to the usual

quadrupole sources has not been

established. It is possible that the dipole

source incurred due to scalloping may

not be a strong noise source. By

avoiding the formation of the dipole, the

intent is to simply avoid the formation of

additional sources. Based on this

guideline, several unacceptable

configurations can be immediately

identified and are shown in Figure 2.5.

For example, the vertical strip at the

downstream edge of mixer 12CL

violates this rule of thumb.

(iii) Selection of Mixer Lobe Number: In

considering the number of lobes to be

used in a mixer, it should be

remembered that lobe number

establishes the wetted perimeter, By

increasing the number of lobes in the

mixer, the interface area between the two

flow streams is increased. This

increased area of interaction leads to an

overall increase in turbulent mixing.

However this process is not entirely

straightforward. Since all mixers must

fit within the same duct cross sectional

area, increasing the number of lobes

produces a corresponding decrease in

lobe width and the resultant diameter of

the axial vortex shed from each lobe

sidewall. The resulting changes in

vortex growth, diffusion, and interaction

NASA/C_2002-210823/VOL 1 12
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NASA/C_2002-210823/VOL 1 13
will also substantially alter and

complicate the mixing process.

Unpublished data exists which

establishes that an increase in lobe

number can be acoustically beneficial

under certain operating conditions.

However, this acoustic benefit is offset

by an increase in skin friction and total

pressure loss which will adversely affect

thrust production.

It is not easy to place a limit on how high

the number of lobes must be from these

considerations but other properties like

weight, blockage due to lobe metal

thickness, and manufacturability of the

lobes do come into the picture. Having

previously tested mixers with 12 and 16

lobes in the first phase and in order to

strike a balance between noise

suppression and thrust loss, we selected

20 lobed mixers for the second phase

which would be similar to the highpenetration 12 lobe mixers (12UH) in

other respects.

2.3.1.2 The Tongue Mixer

As discussed earlier, scallops in forced lobe

mixers may be acoustically advantageous if

designed properly. Most probably, it is the

earlier and more gradual initiation of

streamwise vorticity due to the scallops that

is helpful in reducing high-frequency noise.

Suppression of the low frequency portion of

the spectrum is enhanced as the flow

approaches a uniform state at the nozzle

exit. By letting the scallops become deeper

and deeper both these processes can be

accentuated, although we then lose control

of the cross-sectional areas for the two

streams and, hence the flow properties

between the lobe sidewalls. It is also

possible that for a given shape and number

of lobes in a lobe mixer, there exists an

optimal depth of scallop from a noise

suppression point of view. If we take this

scalloping to its ultimate limit, that is,

scallop all of the lobe sidewall, only the lobe

crests and valleys remain. This bare

skeleton of fully scalloped lobes will look

like multiple tongues of metal deflected

towards the outer nozzle wall and the nozzle

center-line hence, the name "tongue"

mixer. If the two streams are subsonic and

if the two streams flowing along their

respective tongue surfaces still do not

separate then the process of formation of a

streamwise vortex between the two tongues

will still exist and enhance mixing.

However, without the lobe sidewalls to

guide the two streams the tongue surfaces

will now have to be modified from the

corresponding unscalloped lobed mixer

geometry. In fact, the older rules of

designing forced lobe mixers may not be

applicable to the tongue mixer.

This concept of the tongue mixer as an

extreme limit of scalloping actually can be

viewed as one of the steps bridging the gap

between the confluent mixer (splitter plate)

and a scalloped lobe mixer, as shown in

Figure 2.6. Here, the conceptual variations

between the tongue mixer and the confluent

mixer, called the "comb" mixer and the

"screen" mixer are obtained by reducing the

width of the tongues and then adding a

transverse grid. The comb and the screen

mixers will simply generate small-scale

eddies downstream, where the mixing

between the two streams will proceed due to

turbulent diffusion alone and will not be

efficient. However, a tongue mixer, if

designed properly, may stand to benefit

acoustically. It may also have advantages in

terms of manufacturability, weight and cost.

The principle of operation for the tongue

mixer, thus, still remains the earlier and

gradual generation of axial vorticity by

NASA/CR 2002-210823/VOL1 14
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NASA/C_2002-210823/VOL 1 15
guiding each stream in its own tongue such

that the mixing between them occurs across

the vortex sheet which replaces the lobe

sidewalls. This vortex sheet has both axial

and radial vorticity components due to

differences in, respectively, the radial

velocity components and the axial velocity

components of the two streams. The axial

vorticity component is known to enhance

mixing in lobed mixers by increasing the

interface area between the two flows by

curling the vortex-sheet, and thus

introducing the "engulfing" process which is

far superior to mixing due to viscous

diffusion occurring in a confluent nozzle.

This same enhanced mixing process will

now occur in the region between any two

adjacent tongue edges and continue

downstream with interaction between the

adjacent vortex sheets. With axial vorticity

first ingested at the "root" of the tongues

(that is, where the two neighboring tongues

first start diverging from each other) and

with axial vorticity at downstream stations

being possibly strongest midway between

the two tongues, this vortex-sheet will curl

at mid-height and the central vortices will

grow in size as they convect downstream. In

this regard it is interesting to note that Elliott

et al (9) at M.I.T. have suggested through

their CFD studies that for lobed mixers

maximum mixing contribution due to axial

vorticity can be obtained if the axial

vorticity is highest or concentrated at the

mid-height of the lobes. The hypothesized

vorticity dynamics for the tongue mixer is

shown in Figure 2.7.

We will embody these considerations as

three basic principles, to be stated below,

around which we can base our design

guidelines for tongue mixers. First, since we

are talking about a new concept, let us

clearly define some of the new terminology

we have adopted for the tongue mixer:

(a)

(b)

Top tongue or core tongue refers to the

radially outward tongue (as we go

downstream); bottom tongue or fan

tongue for the radially inward one.

Internal stream or surface, for a given

tongue, refers to the flow or surface that

would be inside the unscalloped lobe

from which this tongue can be thought to

be made of; the external stream or

surface would be the one outside that

lobe. Thus for a conventional twostream nozzle with fan flow surrounding

the core flow, for the top tongue (which

can be thought of as being made up of

the crest of a core lobe) the internal

stream is the core stream and the

external stream is the fan stream;

whereas, for the bottom tongue the

internal stream is the fan stream and the

external stream is the core stream.

The three basic fluid-dynamics and acoustic

principles for the tongue mixer to work are:

P1. The internal flow on the tongue should

no____t separate.

P2. The external flow over the tongue should

no___!t go around it towards its internal side

displacing the internal flow and forming

wake like eddies.

P3. The dipole noise-source due to the

interaction of tongue and external flow

should be minimized while achieving the

desired goals of mixing performance.

Thus, for any tongue geometric parameter

under consideration it should be designed

such that the above three principles are

satisfied. We will refer to these principles as

P1, P2 and P3. There may be other rules of

thumb which may improve the non-acoustic

NASA/C_2002-210823/VOL 1 16
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NASA/C_2002-210823/VOL 1 17
performance, such as aerodynamic thrust,

and should be incorporated, if possible,

without violating the above principles.

2.3.2 Design of the New Mixers

2.3.2.1 Unscalloped 20-lobe mixer

The 12 lobe, high penetration mixer (12UH),

12 lobe low penetration mixer (12CL), and

the confluent mixer (CONF) configurations

of Table 2.1 were designed to have equal

bypass ratio. For the parametric study of the

effects of scalloping, a 20 lobe configuration

was selected. The unsealloped 20 lobe

baseline was designed to have an identical

area ratio, lobe penetration and lobe length

as the 12UH configuration. In this way, it

would be possible to isolate the effect of

lobe number on acoustics and aerodynamics.

The angular coordinates of the new 20 lobe

mixer (denoted 20UH) were initially derived

by symmetrically compressing the 12UH

coordinates on either side of the crest

centerline by the lobe ratio (12/20). This

initial profile was then modified by

increasing the radius of curvature of the

crest and decreasing the radius of curvature

of the keel, as viewed in the transverse

plane. In addition, it was necessary to recontour the crest surface in the meridional

plane to suppress a small region of separated

flow. Some CFD results are discussed later

in section 2.4.2. The final configuration is

shown in Figure A.18.

2.3.2.2 Implementation of the New Rules

for Scalloping

In the previous section, we developed two

general guidelines for designing scalloped

edges for a given lobe mixer:

(i) the radial height of the scallop should

gradually increase from zero to some value.

(ii) the scalloped edge should be a "trailingedge" all along its length with respect to

both fan and core flows.

These guidelines do not address the

selection of the axial depth of the scallop or

the specific curve defining the edge of the

scallop. Additional criteria for selecting

these parameters must be defined.

Since the flow inside the lobes and close to

the lobe surface closely follow the shape of

the crest or keel (depending on whether it is

core flow or fan flow respectively) it

appears, at first sight, that a simple V-shaped

deep scallop in the lobe sidewall (with the V

pointing upstream) will satisfy both the

geometrical principles mentioned above.

This appears satisfactory for lobes with

mo.notonic radially diverging crest and keel

lines, or scallops with relatively shallow

axial depth. However it can fail in cases

where either the crest or the keel line has an

inflection point, especially, when the scallop

is axially deep. We will show how that

limitation can be removed and give a general

method to obtain a first-order scalloped edge

satisfying the two rules. This will be done

by building approximate flow lines purely

from the given geometry of the unscalloped

lobe and without the use of any CFD tools.

Any further refinements will of course

necessitate CFD tools; however, in most

cases, the first order estimate of the

scalloped edge should be adequate.

While discussing the geometry of a

scalloped edge on the lobe wall it is

convenient to consider a reference

meridional plane rx (r being the radial

direction and x being the axial flow

direction) in the case of a circular nozzle and

rotate all the corresponding lines (such as,

the lobe crest or keel lines) from meridional

planes at different azimuthal angles on this

NASA/CR 2002-210823/VOL1 18
one referenceplane. We will call the

straightlinejoining thetop of thelobecrest

and the bottom of the lobe keel in this

reference meridional plane as the

"unscalloped"trailingedgeof thelobe. This

line may be canted and not necessarily

radial. Any deviationof the actualtrailing

edgeof the lobefrom this straightline will

beconsideredasscalloping.

It is conveniento uselocalmixeraxes(_,rl)

as parallel and orthogonal to this

"unscalloped"lobetrailing edge.Figure2.8

showsone selectionof theseaxes (with

origin O) with G-axisalongthe unscalloped

edge and q-axis passing through the

intersectionof thecrestandkeellines(point

C). Some of the geometrical terms we have

introduced to describe the scalloped edge

characteristics are also shown. Thus the

depth, height and offset of the scallop

broadly describe its geometry and the type of

the curve tells its shape. Notice that the

depth FH is defined here to be in the qdirection which may not be necessarily in

the axial direction due to a possible cant

angle of the unscalloped edge. The scallop

depth, FH, can be expressed as a percentage

of the maximum depth, CO, that can be

achieved for the given lobe; thus,

100*FH/CO is the percentage depth. The

offset of the scallop apex F is defined from a

point G which in some way defines the

"mid-height point" at that depth. With the

local axes defined as in Figure 2.8, it is

natural to define the mid-height point G as

the mid-point of a line parallel to the

unscalloped edge and passing through the

apex F. The offset can then be quantified as

a percentage of the maximum offset

achievable at that depth as 100*FG/GI.

However, note that it may be possible to

define the locus of "mid-height points" as

the locus of centers of circles inscribed

between the lobe crest and keel lines (see

Figure 2.9(a)). In this latter sense the point

p in Figure 2.9(a) is equi-distant from the

two neighboring crest and keel surface lines

(pq = qr); whereas, in the previous sense (as

shown explicitly in Figure 2.9(b)) it still is

equidistant from these two surface lines but

on a given canted line pqr parallel to the

unscalloped edge. We define the zero offset

line as the simpler-to-draw mid-height line

CC' of Figure 2.9(b). Lastly the scallop

height DE can also be defined as a

percentage of the maximum scallop height

possible AB as 100*DE/AB.

When constructing the scalloped edge one

can ask the broadest question: "What is the

scallop depth, offset, height and type of

curve that should be selected to reduce

noise?" We do not know the answer to this

question and how indeed noise varies with

these parameters. It does not appear that

such a question has ever been posed at all or

answered in the open literature to our

knowledge. It appears that the most

dominant parameters amongst these may be

the depth of the scalloped curve. The effect

of scallop depth is what will be studied here.

Thus a reasonable question that we propose

to answer is "Given the scallop apex

location F (that is depth and offset) how do

you construct the scalloped edge so as to

satisfy the two rules (given at the beginning

of this sub-section)?"

The "trailing-edge (t.e.) rule" requires an

estimate of the flow direction on both sides

of the corresponding unscalloped lobe

sidewall, ff we can obtain a simple

approximation to the flow velocity direction

near the crest and keel lines then that would

indeed help in constructing at least deep

scallops whose edges are close to the crest

and keel lines and which satisfy the t.e. rule.

From two-dimensional inviscid flow

analysis results we know that, for example,

NASA/CR 2002-210823/VOL1 19
Fan Flow

Core Row

Lobe

Offest

FG

Keel

J

_,calloped

Edge

Depth FH"

Unscalloped

Edge

/

Height

DE

B

Center-Une

of nozzle duct

Figure 2.8 Geometrical definitions for describing a scalloped edge on a lobe mixer sidewall.

C'

C

C

7"

(a) Locus of centers of .... inscribed circles (b) Mid-heights of lines parallel to (canted)

unscalloped edge

Figure 2.9 Two possible methods of prescribing the mid-height locus for a lobed mixer.

NASA/C_2002-210823/VOL 1 20
the steady streamlines near a surface and

away from stagnation points and separated

regions are very nearly "locally parallel" to

that surface. But there is an ambiguity when

we say that two curves are locally parallel

because we can translate a given curve

horizontally or vertically or in any other

direction and consider any of these

translated curves to be "parallel" to the

curve in its original location. So we need to

be more precise. We assume, as a first order

estimate for obtaining the flow velocity

direction at a point, P, near a solid line that

the flow velocity there is influenced most by

the point, Q, close to P on that line and the

slope of the line at the point Q. The closest

point Q to a line from a point P in the flow is

of course obtained by dropping a

perpendicular from P to that line as shown in

Figure 2.10. Then we assume that the flow

velocity direction at P is the same as the

tangent to the solid line at Q. Of course, this

procedure fails when point P in the flow

region is equidistant from two or more solid

lines or in general is influenced more or less

equally by two or more nearby surfaces.

However, in our application if we consider

the core flow velocity component in the

meridional plane, say, at the azimuthal

center of the core-lobe, the two influencing

solid surfaces of interest are the lobe crest

and the plug surface, with the plug exerting

little influence on the velocity direction near

the crest line. Similarly, the fan flow

velocity component in the meridional plane

at the center of the fan lobe, which is

bounded by the keel and the nozzle wall is

influenced more by the keel surface than the

nozzle wall when P is near the keel surface.

So it appears reasonable to apply this

approximation to the flow velocity direction

near the crest and keel surfaces which is all

that is needed for satisfying the t.e. rule.

Once the local flow velocity direction is

known at a point then it is easy to verify

whether a proposed scalloped edge near that

point P is a "trailing" or "leading" edge, as

shown in Figure 2.10.

Using this construction we can easily find

the upper and lower radial bounds on the

scalloped edge corresponding to a give apex

location which does not violate the t.e. rule.

First note that if we draw a circle with

diameter PQ (see Figure 2.10) then the

tangents to it at P and Q give the direction of

the flow velocity at P and Q. So if we roll

this circle on the crest line then the envelope

of the rolling circles will correspond to the

streamline or the pathline of the particle

passing through point P in this

approximation. Figure 2.11 shows these

"rolling-ball envelopes" corresponding to a

given scallop apex location F. There are two

of them, one corresponding to the core flow

and the other to the fan flow. Consider for

the moment the upper core flow particle

path passing through F. If the downstream

portion of the tangent to the scalloped edge

at F lies below the rolling ball envelope then

obviously it would be a "leading" edge.

Hence, any scalloped edge which lies wholly

below this rolling ball envelope will violate

the t.e. rule at least in regions close to F. In

this sense the rolling ball envelope from F

gives a "'lower" radial bound for the

scalloped edge on the upper side. Similarly

the lower rolling ball envelope from F gives

the "upper" radial bound for the scalloped

edge on the lower side. Thus, for example,

even a straight scalloped edge (which can be

considered as upper part of a V-scallop)

from F which lies wholly below the upper

rolling ball envelope will not satisfy the t.e.

rule. Further, even if a portion of the upper

straight scalloped edge near F is below the

upper envelope then it will still violate the

t.e. rule.

NASA/CR 2002-210823/VOL1 21
A

C

Figure 2.10 Approximate estimate of the direction of flow velocity vector at point P near the crest-line

and the relative location of a trailing or a leading edge of a scallop near P.

Figure 2.11 Rolling ball envelopes as approximate particle paths corresponding to the apex, F, &the

scallop on two sides of a lobe wall and the corresponding bounds for the scalloped edge.

NASA/C_2002-210823/VOL 1 22
Sincetherolling ball envelopegivesoneof

the bounds for the scallopededge it is

pertinento askif any curve which is wholly

above that envelope (we are referring to the

upper scalloped edge here) satisfies the t.e.

rule. The answer is no because if, for

example, the scalloped edge "approaches"

the crest line and then "recedes" from it, it is

possible that the t.e. rule is locally violated

somewhere in-between. However, one

family of curves which is wholly above the

rolling ball envelope appears to be

promising, namely, the envelope formed by

circles of monotonically decreasing

diameters and touching the crest line, as

shown in Figure 2.12(a). In general, with

the chords between the tangency points of

consecutive circles decreasing

monotonically (P1P1'>P2Pz'>P3P3' etc.), the

two curves (the crest line and the scalloped

edge) can be defined to be converging. In

the special case of a straight crest line and

another converging straight, concave or

convex line (which can be thought as the

envelope of circles with monotonically

decreasing diameters with linear, more than

linear and less than linear rate of decrease

respectively) we can easily see from Figure

2.12(b) that the converging scalloped edge

line indeed satisfies the t.e. rule. We also

observe from Figure 2.12.(b) that for such a

special case not only the chords but also the

perpendiculars drawn from the scalloped

edge on the crest line are monotonically

decreasing. Such perpendiculars are easy to

construct for two given curves on an

interactive CAD program. Without having

proven it for the general case of converging

curves, as in Figure 2.12(a), we will take it

as a rule of thumb that if the lengths of the

perpendiculars from the given scalloped

edge on the crest or keel lines form a

monotonically decreasing sequence as we go

downstream then it will very likely satisfy

the t.e. rule.

In practice it may be easier not only to check

for this decreasing sequence of

perpendiculars but also to simply double

check at several points whether the t.e. rule

is satisfied or not by drawing the

approximate flow velocity directions as in

Figure 2.10 and then iterating on the

scalloped curve if needed. It turns out that

this sort of iterative geometrical method is

extremely fast when implemented on an

interactive computer program and is the

method of choice. We further note here that

if the flow separates from the surface, say, at

the crest-line then the application of the

above rule of thumb becomes even more

conservative because in that case the flow

velocity near the scalloped edge will diverge

away from the crest surface.

Figure 2.13 shows one such construction for

the scalloped edge of the 20 lobed mixer.

Here we have arbitrarily selected the

location of the scalloped-apex as one with

about 58% depth and 0% offset. This gives

us a sufficiently deep scallop (much deeper

than conventional scallops) without any

danger of a structurally weak lobe sidewall

and with a reference offset of zero. In order

to obtain "converging" curves the endpoints

of the scalloped height, namely, D and E are

originally selected well above the limits

imposed by the rolling ball envelopes from

the apex F. The scallop height DE is also

chosen initially to be large so that the two

flows can interact in the scalloped area as

much as possible forming nascent axial

vortices inside the scalloped region before

going downstream of the unscalloped edge

AB. The initial scalloped height chosen was

about 92% with AD = BE. Having thus

located the points F, D and E the problem is

of obtaining a scalloped edge passing

through these points and satisfying the t.e.

rule. One other condition imposed is the

avoidance of sharp corners on the scalloped

NASA/CR 2002-210823/VOL1 23
Figure 2.12(a) Envelope fol_ed by circles of monotonically decreasing diameters and touching

_the crest line. The chords between points oftangency also decrease leading to

"converging" curves.

[_EbT CRESI CRbST

(a) Straight edge Co) Concave edge (e) Convex edge

Figure 2.12(b) Satisfaction of the traume,-edge rule" when the crest is a straight line and the

scalloped edge is converging whether straight, concave or convex.

NASA/C_2002-210823/VOL 1 24
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NASA/C_2002-210823/VOL 1 25
edgesoasto avoid points of high stresses on

the lobe skin and also to avoid secondary

comer vortices which can lead to total

pressure losses.

This process of constructing the scalloped

edge is broken into two pieces - the upper

and the lower edges with respect to the apex.

Let's focus on the upper portion. First note

that a simple straight edge FD will indeed

violate the t.e. rule near F because it is

below the corresponding upper rolling ball

curve as mentioned before. The two

portions will also have to meet smoothly at

the apex which implies that both the

portions have a common tangent at the apex

F which is parallel to the unscalloped edge

AB. With three properties of the upper

curve fixed, namely, two points (F and D)

and the tangent at one of those points (F) we

can let a unique three parameter curve pass

through those points. However, we have

more freedom for the shape of the curve if

we use a four parameter curve with one

parameter left free. B-splines with four

parameters (poles) are easily available and

easy to construct in typical CAD

applications and, hence, were selected. The

fourth parameter, it turns out, controls the

angle of the tangent at the remaining point

D. After selecting this fourth parameter

such that the slope at D is slightly higher

than at A one can draw the proposed

scalloped edge passing through F and D, and

use the rule of thumb of decreasing

perpendiculars (which can be equally spaced

on the scalloped edge) to check if the t.e.

rule is satisfied. After going through this

iteration a couple of times and satisfying the

t.e. rule of thumb we further pulled the point

D upward to point 10 (see Figure 2.13) to

make the curves more converging and

increase the margin of safety. The

decreasing perpendiculars on this final curve

are shown in Figure 2.13 and so is the

double verification of the t.e. rule by

drawing the approximate local velocities on

the scalloped edge as lines orthogonal to

these perpendiculars. A similar process was

used to draw the lower portion of the

scalloped edge, FE. In order to avoid the

sharp comers at D and E large fillet radii

were used there. Since the crest and keel

lines of the corresponding 20-lobe

unscalloped mixer are diverging radially

from each other, the scalloped edge,

constructed as above, also automatically

satisfies the first design rule of gradually

increasing radial distance between the upper

and the lower edges of scallop. Projection of

these curves back on to the threedimensional lobe surface gives the final

scalloped lobe. Scallops obtained in this

manner are seen to have a "boomerang"

shape. Hence, we will call them

"boomerang scallops." Figure A.19 shows

the details of the final geometry of the 20

lobe deeply scalloped mixer of highpenetration (called 20DH).

We have also applied this process to obtain a

family of moderately deep scallops as shown

in Figure 2.14. For a systematic study of the

effect of the depth of scallop on noise we

have selected a scallop of depth equal to half

of the previous one but of the same scallop

height. As before with B-splines we still

have a non-unique choice of several curves.

Six of these are shown in Figure 2.14

ranging from a V-cut (curve 1) to a very

concave line 6 where the fourth poles of the

B-spline are equally spaced on the tangent

line at the apex F. Since the depth is

deemed shallow it appears that all of these

will easily satisfy the t.e. rule and no check

is needed in this case. Each curve, however,

gives a different scalloped area. Notice that

if we had V-cut scallops in both cases then

with half-depth and same height the

moderate scallop would have half of the

NASA/CR 2002-210823/VOL1 26
c

/

/

Curve # % Area

1 39.4

2 41.1

3 42.9

4 44.7

5 46.5

6 48.5

/

/

Figure 2.14. Family of scalloped edges with depth equal to half of the deeply scalloped mixer (20DH)

but same scallop height. The scalloped areas differ and are shown as percentages of the

deeply scalloped one in the inset table. Curve 6 was selected for 20MH mixer.

Table 2.2 New Mixer Parameters

Mixer Code Lobe

No.

20 Lobe Unscalloped 20UH

20 Lobe Moderately Scalloped 20MH

20 Lobe Deeply Scalloped 20DH

Tongue 12TH

Scallop Lobe

Penetration

(Hm/Hmp)

20 None 0.67

20 Moderate 0.67 0.34

20 Deep 0.67 0.34

12 pairs Tongue type

Lobe

Length

(Lm/Dmp)

0.34

0.56 0.34

Area Mixing

Ratio Length

(A_/A_) (LlDmp)

2.637 1.10

2.637 1.10

2.637 1.10

1.765 1.10

NASA/C_2002-210823/VOL 1 27
scalloped area as the deeper one. Hence, we

have finally selected the curve which gives

approximately half of the scalloped area as

the previous deep scallop, namely, curve 6

which has a scalloped area equal to 48.5% of

that for the deeply scalloped case. Figure

A.20 shows the 20 lobe, moderately

scalloped, high-penetration mixer (termed

20MH). Table 2.2 lists the non-dimensional

geometric properties of these 20 lobe

mixers.

2.3.2.3 Design Guidelines for the Tongue

Mixer

The design guidelines given here are those

which were used more in the spirit of

building a proof-of-concept tongue mixer

model adhering to the three principles, P1,

P2 and P3 discussed in the previous section

2.3.1.2. Time constraints did not permit

optimization of the tongue contours through

the use of CFD analysis. Primarily, basic

fluid dynamic principles and engineering

judgment were used to pin down the flow

lines for a proof-of-concept tongue mixer.

We give below a description of the selection

for its major parameters, namely, the number

of tongues, the internal and external flow

lines in profile, and the transverse crosssections.

(i) Number of Tongues

As a baseline design we start with all top

tongues of one shape and all bottom tongues

of another shape. Principle P3 suggests that

we should not have too many tongues lest

there be considerable dipole type noise from

the tongues themselves. The root-widths of

the tongues, although independent

parameters, go hand-in-hand with the

number of tongues and are considered

together. Principle P2 suggests that the

tongue widths should not be too thin, that is,

the number of tongues should not be too

large. This selection can be partially

resolved by using the arguments developed

from vorticity dynamics as described earlier.

The interface area for mixing due to

engulfment depends on the largest diameter

that can be achieved by the two adjacent

vortex structures shown in Figure 2.7 and

that is limited by the smallest tongue width.

This width will govern when adjacent

vortices touch each other and begin

interacting. Hence, if we have unequal

tongue widths then the vortices will grow to

a smaller size than if the tongue-widths were

equal. This would suggest equal tongue

widths.

Examination of the three original lobed

mixers shows that the only mixer with equal

lobe-width at the root was the conventional

mixer with cutouts (12CL). This has 12

lobes. The 12CL mixer may serve as a

comparison point if we also select 12 pairs

of tongues for the tongue-mixer. Any larger

number will make the widths thinner with

more chance of violating Principle P2. Also

we already have a deeply scalloped mixer

with 20 lobes (20DH) which will provide

some indication as to what happens to a

tongue mixer with thinner widths. Hence, in

order to conservatively satisfy P2 for a

proof-of-concept study, we selected 12 pairs

of tongues, with both top and bottom

tongues of equal widths. With an average

radius of about 3.284 in. at the adapter ring

where the roots of these tongues will be

located, this leads tO about 0.86 in. width at

the root of the 24 tongues. This is deemed

conservatively wide for the core and fan

flow velocities to be considered, so that

Principle P2 is not violated.

(ii) Internal Flow Lines in Profile

Consider the tongues in a longitudinal crosssection view along their center-lines in a

meridional plane. Clearly avoidance of

NASA/CR 2002-210823/VOL1 28
separation (Principle P1) governs the

internalflow-lines. It is possiblethat asa

first stepof iterationonecouldstartwiththe

correspondingflow linesof thekeelandthe

crestlinesof a comparablelobedmixerfor

the bottomandthe top tonguerespectively.

However,without thesidewallsthe internal

flow will probablyhavea largertendencyto

separateforthesamesurfaceangle.

Considerfirst the flow line of the bottom

tongue over which the fan-stream flows.

The axial static pressure distribution along

the keel line of a lobed mixer shows that

after a short initial rise it generally

decreases. The pressure distribution along

the internal side of the bottom tongue may

not be so favorable. Its maximum angle

should be somewhere between that used for

tailcones (which is an axi-symmetric body)

and that used for advanced mixer lobes,

perhaps even more on the conservative

lower side, so as to satisfy Principle P1.

This implies that 22 ° < max(0bot) <30 °. The

flow line itself can be made up of a simple

circular arc and straight line combination

with the fan flow exiting at the maximum

angle at its tip which may be made almost

parallel to the tailcone. An angular value

lower than that for the corresponding lobed

mixer with the same length would now

imply that the distance between the tongue's

tip and the tailcone has increased or, in other

words, that the corresponding penetration of

the tongue has decreased. Increased

penetration increases the height of the vortex

sheet or the interfacial mixing area and is

deemed beneficial from a mixing

effectiveness point of view. (Increased

mixing effectiveness is believed to increase

thrust coefficient and reduce low frequency

noise contribution.) This can be achieved by

either increasing the length of the tongue or

being more aggressive on the maximum

angle. This first choice would go against

Principle P3 and must be weighted against

satisfaction of Principle P1.

Now consider the top tongue flow-line

internal to which the higher speed core

stream flows. The GEE 3 mixer lobe data

shows that the axial static pressure

distribution on the crown's core side

decreases as we go downstream (favorable

pressure gradient). For the top tongue we

still need to be cautious so as not to violate

Principle P1. We again conservatively

establish the tongue angles to be below the

maximum angles typically used for the

crown side of lobes. As before, smooth

circular-arc/straight-line combinations can

be used with the exit angle for the core-flow

being kept axial so as to minimize thrust

losses due to non-axial discharge. With

smaller maximum angles we again face the

issue of reduced penetration, which can be

resolved as suggested above.

Overall it is believed that a longer tongue

may be preferable to a higher angle in order

to increase penetration. That is, it is more

important to follow P1 than P3 for the

concept to work at all. Also it should be

noted that for the tongue mixer the angle

between the line joining the tips of adjacent

tongues and the vertical (or radial) line - the

scarf or cant angle in lobe mixers - is not so

relevant because the flows have already

started mixing. So the length of the tongues

should not be adjusted according to this

scarf angle.

(iii) External Flow Lines in Profile

Once the internal flow lines are defined,

then the external flow lines for the center

cross-sectional view of the tongues in the

meridional plane is not so much an acoustic

issue as a structural and aerodynamic issue.

Aerodynamically, the space between the

external and the internal flow lines of a

NASA/CR 2002-210823/VOL1 29
given tongue defines the blockage created by

the bottom and the top tongue respectively

for the core and the fan flow. This blockage

thickness should be minimized. The

minimum thickness at any axial station, on

the other hand, is governed by the structural

stiffness required to withstand the applied

stresses and assure that fatigue failures due

to vibration do not occur.

For the proof-of-concept model, individual

tongues were machined from bar stock and

attached to an adapter ring to form an

assembly. To reduce base drag at the tongue

tip, it must be as thin as possible. A simple

thickness distribution shape which tapers

monotonically from that prescribed at the

root by the mixer adapter ring thickness

(0.298 in.) to a thin machinable edge (0.010

in.) at the tip of the tongue was deemed

adequate.

(iv) Transverse Cross-sections

Transverse cross sections were developed by

prescribing an internal arc and an external

arc at a given axial station once its center

thickness had been decided from previous

rules. Aerodynamically, the tongues should

not act as blunt bodies to the external flow

over them lest they increase the ram drag.

Also for them to function as envisaged

earlier through vorticity dynamics they

should not violate Principle P2. Thus on a

cylindrical surface of the streamlines of the

external flow (that is, whose normal is more

or less in the radial direction and not in the

axial one) their external surface should

appear streamlined like the leading edge of

an airfoil. If we further prescribe that the

external flow on the tongue should leave its

edges axially and, hence, that the tangent to

the external arc at the edge should be axial,

then this would demand a deep curve for the

external arc and unless matched similarly on

the internal side with another deep curve

will lead to very thick tongues which are

undesirable from blockage and weight points

of view. Both these criteria can be met by

something like a super ellipse but it is

obviously more blunt and perhaps difficult

to machine than a simple circular arc. Also

the angle at which the external flow will

leave the tongue edge is determined more

directly by the shape the vortex sheet takes

due to balance of pressure in the two

streams and not so much from the metal

cross-sectional angle. Thus if we remove

the condition that the external arc should

have an axial tangent at its edge this leaves a

wide variety of arc shapes at our disposal to

satisfy the streamlined criterion. For

simplicity of machining, circular arcs were

selected for both the external and the

internal curves such that the thickness at the

edge is again as thin as possible and then

chamfered. (Knife edges are not preferred

from machining viewpoint although they are

preferred aerodynamically.)

Unless the width of the tongue is defined at

each axial station this still does not uniquely

define the circular arcs. The consideration

for width-distribution of the tongue can be

made by looking at its top-view (or actually

radially towards the center from outside the

nozzle) which will also be reflected in its

ALF (aft-looking-forward) view. Each

tongue can be made of constant width

throughout its length or tapered (convergent

or divergent). Here the principle that applies

is P2. Constant width is probably adequate,

however, it will lead to more blockage and

ram drag (clear in ALF view). Tapering will

reduce some of this blockage but the

minimum width at the tip should again be

governed by Principle P2. Again with

simplicity in mind, we decided on the width

distribution that follows from selecting

radial lines for the tongue edges when seen

in ALF view. The final selection gives a

NASA/CR 2002-210823/VOL1 30
convergently tapered bottom tongue and a

slightly divergent top tongue.

This still leaves a wide variety of shapes for

both the arcs of the transverse cross-section.

Approached sequentially the arcs can be

constructed as follows:

• known values are center-line thickness,

tongue-width and edge thickness

• the arbitrariness is then in only one of

the arcs

• select some "appropriate" circular radius

for the external arc which will pass

through the external surface point

• the tongue-width then automatically

defines the edges for that external arc

• with prescribed tongue edge thickness

and the internal point (due to the centerline tongue thickness) the internal arc is

now uniquely prescribed by the three

points

• determination of the arbitrary external

curve radius should also consider that

the internal arc needs to be concave for

following Principles P1 and P2 except

perhaps for some transition region near

the root of the tongue, in addition to the

previous aerodynamic discussion.

• another aerodynamic consideration,

which would be apparent from the side

view, is that any sharp comers should be

avoided, lest they introduce secondary

corner vortices which may increase

losses.

It is possible that there are other machining

issues that can make the cross-section shape

simpler and should be incorporated after

keeping all the above principles in mind.

This procedure basically defines all the flow

lines for the tongues and should be a good

place to start for CFD iterations.

The tongue mixer penetration, as defined in

Table 2.2 and Figure 2.1, was set at 0.56,

which is between the corresponding values

for the lobed mixer configurations 12CL and

12UH. This was selected to provide an

improvement in mixing effectiveness

compared to 12CL. As previously

mentioned the root width of both the inward

and outward projecting tongues were set

equal. This results in a considerable

reduction in the area ratio (Ae/Ah), projected

on the transverse plane, compared to either

the 12 or 20 lobe mixer designs (see Table

2.2). This selection was expected to result

in a considerable bypass ratio reduction

compared to the 12 or 20 lobe mixers, but

was made purposely to satisfy the vortex

dynamics arguments discussed earlier

Figure A.21(a) through A.21(c) show the

geometrical details of the final tongue mixer

as fabricated. It's non-dimensional

geometric properties are included in Table

2.2 Figure 2.15 shows the relative locations

and shapes of the tongue mixer and the 20

lobe deeply scalloped mixer (20DH) for

comparison. Note that attaching the tongue

mixer to the upstream cylinder required an

extra adapter ring. In order to maintain the

relative location of the tongue mixer inside

the nozzle, we have had to fabricate

additional adapter rings for the outer nozzle

and the center-cone, as shown in Figure

2.15.

Figures 2.16(a) and (b) are photos of all

eight mixers included in the test program.

2.4 CFD Analysis of New Mixers

Computational fluid dynamics (CFD)

analysis was done on three of the new

mixers as part of this task, namely the 20

lobe unscalloped (20UH), 20 lobe deeply

scalloped (20DH) and the tongue mixer

(12TH). The mixers were analyzed with the

NASA/CR 2002-210823/VOL 1 31
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NASA/C_2002-210823/VOL1 34
purpose of qualitatively validating our fluiddynamic concepts and to help explain some

of the aerodynamic and acoustic data. We

only briefly touch upon some of the results

from the analysis for the traditional

unscalloped 20UH mixer but expand more

on the CFD results for the other two nontraditional mixers.

2.4.1 Code, Grids and Boundary

Conditions

Numerical Modeling

A 3D, viscous CFD analysis was conducted

using the NPARC analysis code. The

NPARC code, Version 3.0, (Ref. 12) solves

the full three-dimensional Reynoldsaveraged Navier-Stokes equations in strong

conversation form using the Beam and

Warming approximate factorization scheme

to obtain a block tri-diagonal system of

equations. Pulliam's scalar penta-diagonal

transformation provides for an efficient

solver. The code has several turbulence

models available including: the BaldwinLomax, RNG, Baldwin-Barth, SpalartAlmaras, and k-s turbulence models. The

calculations presented in this study use the

Chien low Reynolds number k-s turbulence

model. The implicit scheme uses centraldifferencing with artificial dissipation to

eliminate oscillations associated with the use

of central differences. The code allows the

use of structured, multiple grid blocks. Trilinear interpolation is used to transfer

information at the grid block interfaces. The

NPARC code has been used previously to

predict the internal and external flows

related to other mixer/nozzle exhaust

systems.

Grid Generation

The mixers are composed of identical pairs

of lobes which are spaced at equal angular

intervals. This symmetric geometry is

exploited to reduce the computational

requirements resulting in a grid which

extends circumferentially between the

centerline of one lobe pair. The GRIDGEN

code, Version 11.2 (Reference 13), was used

to generate the computational grid. The

distance of the first grid point off a viscous

surface is 0.001. The blocks are made with

contiguous interfaces and extend from a

station far upstream of the mixer exit plane

to the nozzle exit plane. In order to save

time, simulation outside the nozzle (for the

jet plume) was not done but the downstream

boundary conditions were adjusted

appropriately, as discussed later.

The computational grid for the 20 lobe

unscalloped mixer (20UH) is shown in

Figure 2.17. The grid consists of 6 blocks

and approximately 500,000 points. The

highlighted grid contains the mixer region of

the grid. The grid for the 20 lobe deepscallop mixer (20DH) is shown in Figure

2.18. The grid consists of 5 blocks and

approximately 1,000,000 points. The

additional points were needed to resolve the

scalloped region of the mixer. The

computational grid for the 12 pair internal

tongue mixer is shown in Figure 2.19 and

consists of 6 blocks and approximately

1,000,000 points.

Boundary Conditions

Inside the nozzle, the fan and core upstream

total pressures and temperatures are

specified. The walls are modeled as no-slip

adiabatic surfaces. The downstream

boundary at the nozzle exit plane has an

"extrapolated static pressure gradient." The

centerline is modeled as an axis of

symmetry. The flow variables are massaveraged along this axis. Symmetry

conditions were used along the sides of the

computational domain which are the

meridional planes in the center of each lobe.

NASA/CR 2002-210823/VOL1 35
Mixer

Figure 2.17 Partiat grid for the 20-lobe unscalloped mixer (20UH) inside the nozzle.

NASA/C_2002-210823/VOL 1 36
Scallop

Figure 2.18 Partial grid for the 20-lobe deeply scalloped mixer (20DH) inside the nozzle.

Mixer

Figure 2.19 Partial grid for the tongue mixer (12TH) inside the nozzle.

NASA/C_2002-210823/VOL 1 37
The extrapolatedstaticpressuregradientis

implementedby extrapolatingall primitive

variables(density,velocities,energy). The

staticpressuregradientwhichis determined

from the total energydistributionis scaled

by a userprovidedpressurewhich controls

themass-flowrate. Thisboundarycondition

allowedavortexto existthroughtheexit of

the computationaldomain and improved

convergenceof the solutionascomparedto

using a constantexit static pressurefield

boundarycondition.Theconvergedsolution

producesglobal massconservationwithin

thenozzleto within 1%of theinletvalue.A

slight radial static pressure gradient

remainedat theexit planein the converged

solution.This is not unusualand similar

behavior was observed in the plume

experimentalsurvey.

2.4.2 Results

(a) 20 Lobe Unscalloped Mixer (20UH)

CFD runs on the first version of the 20-

lobed mixer, which was simply an angularly

compressed version of the previous 12-lobed

high penetration (12UH) mixer, were made.

It was found that in the upper half region of

the underside of the lobe crest a strong

adverse pressure gradient existed which led

to separation of the core flow. A refined

analysis of 12-lobed unscalloped mixer

(12UH) (designed some six years ago) also

revealed an adverse pressure gradient and

some separation in the same region which

was not resolved by the grid used during the

original effort. Although a comparison of

the measured pressure on this mixer (12UH)

at the core crest centerline showed this

feature, the location of the beginning of the

adverse pressure gradient was further

downstream, thus confining the separation to

a small region at the core crest tip. This is

possibly due to a very aggressive core flow

line deflection angle at the inflection point

of the lobe crest line - which was done to

maintain a high penetration (0.68) in a small

length of the mixer. Hence, to alleviate

separation in the 20-lobed mixer the core

crest line maximum angle was decreased

while maintaining the lobe penetration by

slightly increasing the lobe length and thus

introducing a slightly larger cant angle of the

mixer exit plane.

As an example, Figure 2.20 shows the flow

velocity vectors just outside the boundary

layers on two sides of the lobe wall. First,

observe that there is now no separation of

the core flow at the crest, or of the fan flow

at the keel. Secondly, note that the core

flow close to the crest and the fan flow close

to the keel on the lobe sidewall are "almost

parallel" to the crest and keel lines

respectively, as assumed in Section 2.3.2.2

(see Figure 2.10). Moreover, slightly away

from the crest and the keel lines, the core

and the fan flow direction on the lobe

sidewall are such that the rules of thumb

given in Section 2.3.2.2 for satisfying the t.e.

rule become even more conservative. This

puts the previous design rules for scalloping

on a firmer ground.

(b) 20 Lobe Deeply-Scalloped Mixer

(20DH)

Typical cross-sectional results for the 20

lobe deep scalloped mixer (20DH) are given

in Figures 2.21 to 2.23 for the total

temperature (TT), axial vorticity (_) and the

turbulent kinetic energy (k) in terms of nondimensionalized quantities for axial

locations both inside the scallop and outside

the mixer up to the nozzle exit plane. In

these figures, X is measured positive

downstream of the mixer exit plane, Ls is

the axial length of the scallop and L is the

mixing length (from the mixer exit plane to

NASA/CR 2002-210823/VOL1 38
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the nozzle exit plane). The operating

conditionis NPRf= 1.457,NPRo- 1.375,

Ttc/Ttf= 2.296 with free-jetMachnumber,

Mfj = 0. The datapresentedin theplotsis

non-dimensionalizedin the following

manner:

lengtharegivenin Figures2.24-2.26.This

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done.

Tnon = T/(Yref/Tred,

_'_"non = _Xref/a ref,

knon = k/a ref 2

where Yref= 1.4, Tref = 533 °R, xref = 1/12 ft,

a ref = _/'YrefRTrefand R = 1716 ft-lb/(slug-°R).

From these figures, we observe that strong

nascent vortices begin to form immediately

inside the scallop, as expected, and continue

to form a tighter vortex downstream of the

mixer which quickly migrates to the nozzle

wall (within the first one-third distance of

the mixing-length), and finally interacts and

merges with the neighboring similar vortex

by the end of the nozzle exit plane. By the

time the flow has reached the mixer exit

plane the interface between the core flow

and the fan flow has already been increased

tremendously as compared to the

unscalloped case, which leads to better

mixing. The turbulent kinetic energy

magnitude does not appear to be significant

and is fairly uniform in the cross-sectional

plane close to the nozzle exit. All these

features appear to be in the direction

expected earlier and validates the rules of

thumb used in designing the scallops, at

least from the point of view of flow physics.

It should also be observed that no

indications of separation were noted on the

inside of the lobes with the introduction of

the deep scallop.

(c) The Tongue Mixer

Similar results from the CFD analysis for the

tongue mixer (12TH) with baseline nozzleFrom these figures it can be seen that:

(i) Mixing does indeed begin vigorously

right from the beginning at the first station

shown (X/L_ = -0.66) where the tongues

have already diverged from each other

(Figure 2.25(a)). This is as expected and is

similar to the behavior in 20DH mixer at the

start of the scallop.

(ii) By the time the core and fan streams

have reached the tongue-tip exit plane (x/L_

= x/L = 0) considerable axial vorticity is

already formed, similar in intensity to that

form 20DH but a with larger area of

coverage due to the broader tongue widths.

(iii) Downstream of the tongues, an isolated

vortex is immediately formed similar to that

in 20DH. However, unlike 20DH, it lingers

at the same radial height for a longer axial

distance and does not immediately convect

to the outer duct wall. The radial convection

velocity of this vortex is governed by the

average radial velocity components of the

core and the fan streams in the tongues. This

is probably much smaller than in 20DH due

to smaller angles of divergence of the

tongues as compared to the aggressive lobe

crown- and keel-angles of 20DH.

(iv) At the nozzle exit plane, (x/L = 1), the

total temperature (Figure 2.24(b))is fairly

uniform except for the annular hot ring near

the central axis. This shows that the core

flow over the center-cone is not as well

mixed as in mixer 20DH. The vortex

(Figure 2.25(b))is less intense and further

NASA/CR 2002-210823/VOL1 46
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away from the nozzle wall than in 20DH but

the turbulence intensity (Figure 2.26(b))

suddenly seems to have grown inside this

vortex. It appears that in 12TH the

boundary layer on the nozzle wall is growing

much faster than in 20DH (compare

turbulence intensity contours at x/L = 0.66

in Figures 2.26(b) and 2.23(b)) and the axial

vortex has sucked into itself a large part of

this wall boundary layer by induction by the

time it reaches the nozzle exit plane.

Whether this is an artifice of the turbulence

modeling used in the CFD code or a real

physical effect is not known. One acoustic

effect that such higher turbulence intensity

can have is an increase in amplitude of the

higher frequency portion of the sound

spectrum due to the small length scales of

these axial vortices. The relative amplitude

of the hot spots and the hotter central core

should also be observable in the plume

survey. This verification will be done later.

One of the important items verified by the

CFD analysis was the absence of flow

separation on the inside surface of both the

tongues as required by the design principles

in Section 2.3.2.2. We will keep these CFD

observations in mind when we compare the

acoustic results for the tongue mixer later.

Finally, Figure 2.27 shows a comparison of

the non-dimensional axial velocity contours

at the nozzle exit plan for all the three

mixers analyzed in this CFD study: 20UH,

20DH and 12TH (normalization is with

respect to the reference speed of sound, aref,

mentioned earlier). This figure shows that,

overall, the unscalloped mixer, 20UH, has a

much higher axial velocity throughout,

whereas, both the deeply scalloped 20DH

mixer and the tongue mixer (12TH) have

lower overall velocity but are not as

uniformly mixed as the unscalloped mixer.

NASA/CR 2002-210823/VOL1 53
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Chapter 3

Experimental Program

A comprehensive experimental program was

carried out which included measurement of

far-field acoustic levels, nozzle aerodynamic

performance, and jet plume aerodynamic

surveys. Due to the specialized

requirements for obtaining some of this data,

two separate facilities were used. A

description of these facilities and the

experimental apparatus employed are

provided in the following sections. In

addition, the test matrix for each phase of

the program is presented along with a

description of the experimental procedure

and data reduction methods employed.

3.1 Acoustic Tests

Test Facility

All acoustic tests were carried out in

NASA's Aerocoustic Propulsion Laboratory

(APL) at Lewis Research Center. This

facility is a 65 ft. radius, acoustically treated,

hemispherical geodesic dome (see Figure

3.1). The walls of the dome are treated with

acoustic wedges as is approximately half of

the floor area. Within the confines of the

dome is the Nozzle Acoustic Test Rig

(NATR). The schematic floor plan of APL

with NATR is shown in Figure 3.2. The

NATR duct work is acoustically lined

(inside and outside) and extends from an

annular air ejector system to a plenum and

bellmouth transition section which is an

ASME long-radius nozzle followed by a

free-jet nozzle duct having an exit internal

diameter of 53 in. and a nozzle centerline

approximately l0 ft. above the floor. This

arrangement provides a free-jet Mach

number capability up to 0.3 to simulate

forward flight effects on the noise source.

An acoustically treated wall installed in the

APL near the NATR exit plane and

extending aftward along the length of the jet

exit rig and test article shields the test

vehicle noise source from the untreated floor

region to prevent reflections from unrelated

test rig equipment residing in this vicinity.

Downstream of the NATR is the Jet Exit Rig

(JER). Test models are installed on the aft

end of the JER, and the movable JER is

positioned axially relative to the NATR exit

plane at the desired location (generally a

distance which aligns the test model nozzle

exit approximately 24 in. downstream of the

NATR exit plane) to appropriately utilize

the 48 ft. arc microphone array of the APL.

The JER is the structure through which

airflow is delivered to the test article via

connections to facility compressed air

supplies. Exhaust gases from the

JER/NATR are expelled through the 43 ft.

high by 55 ft. wide APL exhaust door.

More detailed information relative to the

APL facility, its test rigs and support system

is available in Reference 14.

The jet rig provides two stream flows whose

flow rates are measured by choked Venturimeters. The total pressure and total

temperature are monitored in the model

nozzle at a charging station just upstream of

the mixer exit plane (see Figure A.1). In the

first phase of this program, in November

1995, it was found that the desired high

nozzle pressure ratios in the secondary (fan)

stream could not be achieved. Hence, a year

later, the jet rig was modified to provide

supplementary air and the desired nozzle

pressure ratios acquired in the second phase

of this task.

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Figures 3.3 and 3.4 show the mixer-nozzle

models installed in the jet exit rig. With a

nozzle exit diameter of 7.25 in., the ratio of

the diameter of the free-jet nozzle to the

model nozzle is 7.31 and is deemed

appropriate for the range of frequencies of

interest. The merging of the freejet/ambient shear layer with the modeljet/free-jet shear layer occurs several

diameters downstream of the noise

producing region of interest in the model

and is not a concern in this test. Narrow

band acoustic data was acquired using _Ain.

Bruel & Kjaer microphones positioned on a

48 ft. radius from the nozzle exit center in a

horizontal plane through the nozzle axis.

Twenty-five microphones were positioned in

the upstream and downstream quadrants of

the jet ranging from an angle of 0 = 45 ° with

respect to the inlet axis to 0 = 165 °. [Note

that the angle 0 used in this report is with

respect to the jet inlet axis, rather than the jet

exit axis.] This is deemed to be the range of

observable angles away from the

diffraction/shadow lip effect of the free-jet

nozzle in the upstream quadrant and outside

the free-jet in the downstream quadrant.

Test Procedure and Data Processing

Acoustic testing was conducted by

establishing the initial desired free-jet Mach

number in the NATR. An acoustic test for

the background noise with free-jet flow was

done. Following this, generally, the test

point conditions for the lowest pressure ratio

and temperature were fixed. When

conditions stabilized, acoustic data was

acquired. Fan and core flows were then

adjusted to the pressure and temperature

conditions corresponding to the power

setting for the next highest pressure ratio test

point. After conditions stabilized, acoustic

data was again acquired. This procedure

was repeated in the order of increasing

pressure ratio until acoustic data was

acquired for all power setting simulations at

the first free-jet Mach number. Once this

was accomplished, the free-jet Mach number

was changed and the procedures described

above repeated in reverse order (decreasing

pressure ratio) until acoustic data was

acquired at all desired power settings

associated with the second Mach number.

This cycle was repeated for all free-jet Mach

numbers of interest.

The acoustic and aerodynamic performance

data for the test program was provided by

NASA LeRC to Allison Engine Company in

electronic data base format for further

analysis. The acoustic data processing

scheme used at NASA LeRC is outlined in

Figure 3.5. Its final goal is to produce noise

data on the ground below the flight path for

a flyover at an altitude of 1500 ft. This

method takes into account, broadly

speaking, microphone calibrations, free-jet

shear layer refraction, atmospheric and

spherical spreading attenuation, data scaling,

Doppler shifting for flyover data and

standard day meteorological corrections.

The free-jet background noise was

subtracted from the measured acoustic data

for test points simulating flight conditions.

Some details regarding the specific method

used for free-jet shear layer refraction are

given in Appendix B. It is essentially an

adaptation of Amiet's method described in

Ahuja et al (15) and extended to nat-row-band

spectra. With no prior knowledge of the

location of sources in the different frequency

band, they are all assumed to be at the

nozzle exit plane center. This assumption is

valid at higher frequencies but may

introduce some error at relatively low

frequencies, which are hopefully below the

full-scale low frequency range.

NASA/CR 2002-210823/VOL1 58
NASA/C_2002-210823/VOL1 59
NASA/C_2002-210823/VOL 1 60
[ Digitize sound signals I

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Figure 3.5 NASA LeRC acoustic data processing scheme.

NASA/C_2002-210823/VOL 1 61
The final acoustic data provided by NASA

LeRC to Rolls-Royce Allison was in the

following form, scaled to a nozzle with

660.5 in 2 throat, at standard day conditions

(77°F and 70% relative humidity):

(1) 150 ft. polar data: 1/3 ra octave-band

sound pressure level (SPL re. 2"104

N/m 2) spectra (80 hz to 10,000 hz) and

overall SPL (OASPL) at angles ranging

from 55 ° to 165 ° in increments of 5° in

the reference frame of the nozzle; 1/3 rd

octave-band sound power spectrum

(PWL re. 10 "12 W) and overall power

(OAPWL) were also provided. All of

this data was found for ambient static

pressure of 14.3 psi.

(2) 1500 ft. flyover data: Doppler shifted

1/3 ra octave-band SPL spectra, OASPL

directivity, PWL spectra, OAPWL, Tone

Corrected Perceived Noise Level

(PNLT) and the Effective Perceived

Noise Level (EPNL). The flyover data

was found for an ambient static pressure

of 14.7 psi.

In particular, note that the 150 ft. polar data

in the reference frame of the nozzle is not

Doppler-shifted, whereas, the 1500 ft.

flyover data is, indeed, Doppler-shifted

using the free-jet Mach no. as the flyover

Mach number and both are in the free-field.

There are no corrections for extra ground

attenuation, multiple jet shielding or

airframe shielding/reflections.

The aerodynamic data provided by NASA

LeRC to Rolls-Royce Allison for these

acoustic tests included the actual operating

conditions (that is, fan and core nozzle

pressure ratios and total temperature ratio),

the free-jet Mach number, the test-cell

ambient conditions and measured mass-flow

rates in the two streams.

Test Matrix

The test conditions at which data were

obtained are listed in Table 3.1. Data was

obtained over a range of fan and core

pressure ratios typical of those observed in

current engines at approach and take-off.

Initial acoustic testing, carried out in 1995,

revealed that fan pressure ratios above 1.44

could not be reached due to JER air-supply

limitations. During the 1995 test program,

data was acquired at four values of free-jet

Mach numbers (0, 0.1, 0.2, and 0.3).

Following modification of the rig air

delivery system, a second test program was

carried out in 1996. During this test, data

was acquired at all intended nozzle

operating conditions but measurements at

the free-jet Mach number of 0.1 were

deleted from the test matrix, as shown in

Table 3.1. Thus, there is significant overlap

of operating conditions between the two

tests done a year apart and, generally

speaking, the results of the second test were

considered to have superceded those of the

first test. It is the results of the 1996 test

that are reported and analyzed here. Table

3.2 lists the mixer-nozzle configurations

tested in 1995 and 1996. Note that only the

four original mixers were tested in 1995;

whereas, in 1996, all the eight mixers were

tested but no data was obtained for the

nozzle with 25% increased length.

3.2 Aerodynamic Tests

Test Facility

Static thrust and nozzle performance

measurements were obtained in the

blowdown facilities at ASE FluiDyne Test

Group's St. Paul, Minnesota Laboratory.

Tests were carried out in 1995 on the four

original mixers in FluiDyne's Channel 14.

Selected results from this sequence are

presented in this report. A second series of

tests were carried out in 1997 in Channel 11.

NASA/CR 2002-210823/VOL1 62
Condition

Code

A

B

C

D

E

F

G

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Mq

(Free-JetMach

No.)

0.0

0.0

0.0

0.0

0.2

0.2

0.2

0.2

0.3

0.3

0.3

0.3

NPRr

(PreP,,,0

(fan)

1.21

1.44

1.61

1.82

1.21

1.44

1.61

1.82

1.21

1.44

1.61

1.82

(PtdPa_0

(core)

1.17

1.39

1.54

1.74

1.17

1.39

1.54

1.74

1.17

1.39

1.54

1.74

TTR

(Ttc/Ttf)

2.21

2.34

2.62

2.79

2.21

2.34

2.62

2.79

2.21

2.34

2.62

2.79

Operating

Condition

Approach

T.O. #1

T.O. 4_2

T.O. #3

Approach

T.O. #1

T.O. #2

T.O. #3

Approach

T.O. #1

T.O. #2

T.O. #3

Table 3.1 Acoustic Test Matrix

Mixing-Length Change (AL/L)

Mixer Code 0 % -25 % -50 % +25 %

co_w ox o o o

12CL OX O OX O

12UH OX OX

12TH X X X

16UH OX O OX O

20UH X X X

20MH X X

20DH X X X

Table 3.2 Mixer Nozzle Configurations for Acoustic Tests

(O = 1995, X = 1996)

NASA/CR 2002-210823/VOL1 63
Hardware in the second test program

included the 4 new mixer designs and

selected repeat runs for some of the original

mixers. Both these channels are twotemperature flow static thrust stands in

which the nozzle thrust is determined from

force measurement with a strain-gage force

balance. The general arrangement of

Channel 14 and Channel 11, as well as, the

station notations are shown in Figure 3.6,

3.7 and 3.8 respectively.

The airflows for both the cold and hot

passages of a test nozzle are obtained from

the facility 500 psi dry air storage system.

Air for the cold passage is throttled, metered

through a long-radius ASME nozzle, ducted

to the cold passage of the test nozzle, and

finally exhausted to atmosphere. Air for the

hot passage is throttled, passed through a

regenerative storage heater, mixed with

unheated bypass flow to achieve a desired

temperature, metered through a long-radius

ASME nozzle, ducted to the hot passage of

the test nozzle, and finally exhausted to

atmosphere.

The model assembly is supported by a 3-

component strain-gage force balance and is

isolated from the facility piping by two

elastic seals (see schematic in Figure 3.8)

The ASME meter at Station 1 is watercooled to protect the elastic seal from

thermal effects. Since the cooling water is

confined to the upstream (i.e. non-metric)

hardware only, no tare forces are introduced

by the water supply lines. Facility

implementation was provided to calculate

mass flow rates at Stations 1 and 4 (Figure

3.8) and to calculate the exit thrust produced

by the test nozzle. The data were recorded

with the laboratory digital data acquisition

system. For the 1997 test in Channel 11,

charging station instrumentation in the core

duct consisted of four 5-probe total pressure

rakes (Pt8), two 4-probe and two 2-probe

total temperature rakes (Tt_). Also

associated with each total pressure rake was

on outer wall static pressure tap (P_8).

Charging station instrumentation for the fan

passage consisted of four 12-probe areaweighted rakes (Pt7) and three 5-probe

thermocouple rakes (Tt7). An inner and

outer wall static pressure tap (P_7) was also

associated with each total pressure rake in

the fan duct.

Test Procedures

Two types of tests were done: "hot" when

the core flow is heated up to approximately

860°F and "cold" when the core flow was

nominally at 70°F, same as the fan flow

temperature for all tests. For all hot nozzle

tests, Tt7 and Tt8 were calculated as an

average from thermocouples at each

charging station (12 for the core, 15 for the

fan). For all cold nozzle tests Tt7 and Tt8

were calculated form Tt4 and Ttl,

respectively, by subtracting the temperature

drop due to adiabatic throttling of flow

between the meter station and the nozzle

charging station. The charging station total

pressures were defined as the area-weighted

average from all the available probes in each

duct (20 in core and 48 in fan).

The fan and core mass flow rates through the

test nozzle were determined using choked

ASME long-radius metering nozzles. These

values were used to determine the effective

thrust areas (discharge coefficient, reference

duct-area) for each duct and the whole

nozzle.

To evaluate the axial and vertical thrust

components, the force balance was first

calibrated. The force balance calibration

determined the output characteristics of the

three force balance flexures and the two

elastic seals between the metric model

NASA/CR 2002-210823/VOL1 64
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NASA/C_2002-210823/VOL 1 67
assembly and the non-metric facility

structure. The elastic seals produce a small

tare force, largely due to radial seal

deflections necessary to support the static

pressure differential across the seal. The

seal and balance assembly were calibrated

under simulated operating conditions of load

and seal differential pressures. This

procedure yields the thrust vector

(magnitude, angle and location), thrust

coefficient, fan and core effective throat

areas, fan and core mass flow rates, and

overall nozzle discharge coefficient.

Standard ASME long-radius flow nozzles

were tested before and after each test

program to demonstrate facility accuracy in

determining the discharge coefficient and

the thrust coefficient of test nozzles. The

test results were compared with predicted

(target) values which are based on semiempirical equations and were found to be

very accurate in general.

Test Matrices

The purpose of these tests was to compare

the aerodynamic performance of the mixers

at cruise and take-off conditions. Since the

facility had a limit on the maximum

allowable temperature in the core flow of

860 °, the desired "cruise" total temperature

ratio of 2.79 could not be achieved, but a

maximum of 2.50 was used instead. Also,

in order to assist the evaluation of thrust

mixing efficiency, which effectively

captures the effect of temperature on thrust,

additional "cold" tests were performed

where the core flow was kept at the same

temperature as the fan-flow. Aerodynamic

tests were done only on selective mixers in

1997 after results from the acoustic tests

were known. Tables 3.3 and 3.4 give the

nominal test matrices for which data was

obtained in the 1995 and 1997 tests

respectively for the various mixer

configurations with baseline nozzle-length

and no free-jet.

3.3 Plume Surveys

Limited jet plume aerodynamic surveys were

carried out in 1995 and 1996 in the NASA

APL/NATR facility following the acoustic

tests, with the data used to correlate noise

with flow properties.

A unique multi-sensor rake assembly, shown

schematically in Figure 3.9, was used to

measure total pressure (Pt), static pressure

(P0, and total temperature (Tt). The rake

assembly was mounted in a frame which

allowed horizontal traversing of the plume

in the axial and the transverse directions.

This assembly was positioned at a series of

axial locations downstream of the nozzle

plane, providing a cross sectional survey of

the aerodynamic properties in the plume at

each location.

The axial distance X (downstream of the

nozzle exit plane), for which good data

could be taken varied from 0.2D to 10D (D

= nozzle exit diameter) A "full" axial

survey consisted of X/D = 0.2, 0.5, 1.0, 3.0,

5.0, 7.5 and 10. At each X/D location a

lateral sweep of the rectangular array across

the plume was made in 'A in. steps in the Y

direction (see Figure 3.9 for definition of X,

Y, Z). The data was acquired in an overlapping manner such that a corresponding

measured Pt Tt and Ps value was obtained for

each data point location in the transverse XY

measurement plane. A "full" transverse

survey typically swept from Y = -4.75 in. to

Y = +4.75 in., whereas a "center" transverse

survey, done to conserve testing time,

typically varied from Y = -0.75 in. to Y =

+0.75 in.. The extent of the vertical, "Z",

sweep was constant from Z = -5.0 in. to Z =

+5.0 in. with 1,4in. spacing between adjacent

NASA/CR 2002-210823/VOL1 68
Table3.3 1995NominalAerodynamicTestMatrix

Mixer NPRr NPRc TTR

CONF 1.44 1.40 1.0/2.34

1.80 1.76 1.0/2.50

2.38 2.30 1.0/2.50

12CL 1.44 1.40 1.0/2.34

2.38 2.30 1.0/2.50

16UH 1.44 1.40 1.0/2.34

2.38 2.30 1.0/2.50

Table 3.4 1997 Nominal Aerodynamic Test Matrix

Mixer TTR

12CL

12UH

12TH

20UH

20DH

1.44 1.39

1.62 1.54

1.82 1.74

2.40 2.32

1.44 1.39

1.62 1.54

2.40 2.32

1.44 1.39

1.62 1.54

1.44 1.39

1.62 1.54

2.40 2.32

1.44 1.39

1.62 1.54

2.40 2.32

2.34

2.50

2.50

2.50

1.0/2.50

1.0/2.50

1.0/2.50

2.34

2.50

2.50

2.50

1.0/2.50

2.50

2.50

1.0/2.50

NASA/CR 2002-210823/VOL1 69
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NASA/C_2002-210823/VOL 1 70
probes. The total and staticpressuresand

the total temperatureat a givenpoint were

usedto calculatethelocalplumevelocityby

using the isentropicrelations. This data

could then be used to examine the

distribution of the various aerodynamic

quantitieswithin a slice of the plume or

track the axial evolutionof a parameterat

variouslocationsrelativeto the centerline.

Table 3.5 showsthe testmatrix for which

plumesurveydatawasobtained. The test

sequenceemployedfour of the available

mixers, the baseline length nozzle, fan

pressureratiosof 1.44and1.6,andfree-jet

Mach number of 0.2. A few selected

additionalpoints were also consideredas

timepermitted.

Duringthe 1995aerodynamicperformance

tests at FluiDyne Laboratories,exit rake

surveysfortotalpressure(Pt) obtained at the

mixing plane and Tt at the nozzle exit plane

were obtained using a rotating (gear-driven)

survey rake assembly (see Figure A.2).

Figure 3.10 and 3.11 show the details of the

15 probe Pt-rake and Tt-rake respectively.

Typically, data was taken radially at angles

ranging from 0 ° to 25 ° in increments of 2.5 °

after the blow-down facility had stabilized

for each angle. This test employed the four

original mixer configurations (CONF, 12CL,

12UH, 16UH) and the baseline length

nozzle.

The test-matrix covered the reference point

of NPRf = 1.44, NPRc = 1.39, TTR = 2.35

and the "hot" cruise point of NPRf = 2.4,

NPR_ = 2.32, TTR - 2.50. These

correspond to the operating points for which

full plume data was also obtained (see Table

3.5).

NASA/CR 2002-210823/VOL1 71
Mixer

CONF

12CL

12UH

12TH

16UH

20UH

20MH

20DH

NozzleLength

Change

0%

0%, -50%

0%

0%

0%

0%

0%

0%

Operating Condition*

M0=0

B C D

X X X

O

O

O

O

M_ = 0.2

F G

X X X

X X X

(x)

X X X

H

m,ll

* See Table 3.1 for definition of operating condition codes B, C, etc.

** BF is same NPR, TTR as B or F but with Mfj = 0.1

O = 1995, X = 1997, (X) = Partial Survey

Table 3.5 NASA Plume Survey Test Matrix

NASA/C_2002-210823/VOL 1 72
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NASA/C_2002-210823/VOL 1 74
Chapter 4

Aerodynamic Results and Analysis

In this chapter, we present the aerodynamic

performance data and plume survey results

for some of the mixer-nozzle configurations

tested. Aerodynamic performance is

presented in terms of mass-flow rates,

bypass ratio, nozzle-discharge coefficients,

and the thrust coefficient at various

operating conditions. The plume survey

data is presented in terms of total

temperature and local velocity distribution

across transverse cross-sections of the jet

from the nozzle exit plane to ten diameters

downstream and also along the center-line of

the jet. The plume surveys help bridge the

gap later between the plume flow physics

and the acoustic properties of the jet. The

aerodynamic performance data, on the other

hand, provides the metric for thrust

efficiency, both at take-off and cruise, to

compare against the noise suppression

ability of the mixers.

4.1 Aerodynamic Performance

Table 4.1 summarizes the aerodynamic

performance data collected for the various

mixer-nozzle configurations from the static

thrust-stand tests at FluiDyne. We will first

define the various terms appearing in this

table and then discuss the results.

4.1.1 Definition of Various Terms

For a given operating condition, the data is

divided into three broad categories:

(i) Mass-Flow Rates: These are the actual

values, as obtained at test-day

conditions, for the fan passage (mr),

(ii)

(iii)

core passage (me), the total mass-flow

rate (mtotal = mf + me) and the bypass

ratio (BPR, defined as mf/mc).

Nozzle Discharge Coefficients: The

discharge coefficients for the fan and

the core nozzles (Co), are defined as

the ratio of actual mass-flow rate to the

ideal isentropic mass-flow rate

discharging to the ambient test-cell

conditions for the given nozzle

pressure ratio (NPR). These

coefficients are given in terms of the

"effective throat area" for the

individual nozzles, CBeAf and CDcAo,

respectively. Individual flow

discharge coefficients cannot be found

because the individual throat areas Af

and Ao are not known for all the

mixers, especially, mixers with

scalloped or cutout lobes (20MH,

20DH and 12CL) and the tongue mixer

(12TH). The ratio of these two terms

(CDfAf)/(CDcAc), provides the

"effective area ratio" between the fan

and the core streams. In addition, an

"overall" nozzle discharge coefficient,

CDnozzle, defined as the ratio of [total

effective throat-area of fan and corenozzles] to [the nozzle-throat-area] or

(CofAf + CDcAe) /Anozzle-t_roat, is also

presented. The inspected cold nozzle

exit area, Anozzle-throat, of 41.2233 in z

was used.

Thrust Related Terms: The measured

model-scale value of net static axial

component of thrust, Hx, is presented

first. This thrust component is defined

NASA/CR 2002-210823/VOL1 75
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NASA/C_2002-210823/VOL 1 76
as the axial exit momentum of the exhaust

flow, plus the excess of exit pressure over

ambient pressure times the projected nozzle

exit area. Next, the corresponding "fullscale" value (with scale-factor of 4) of this

axial thrust component, normalized to a

reference ambient pressure of Pref = 14.7 psi,

and defined as Hx* = 16Hx(Prcf/Pamb) is

given. The effective mixed velocity, defined

as the ratio of measured axial thrust to

measure total mass-flow rate and given by

Vmix = Hx] mtotal, and sometimes used for

comparisons between mixer acoustic results,

is presented next. Finally we present the

thrust coefficient, Ctr, defined as the ratio of

measured total thrust to the ideal separate

flow or unmixed thrust. The latter is defined

as the sum of ideal thrusts obtained from the

fan and core streams when each stream, with

the measured mass-flow rate, is ideally

expanded to the ambient pressure from the

given total pressure and temperature without

mixing with the other stream. In addition, to

understand how the thrust-coefficient

changes with core-to-fan temperature ratio

we also present a measure of percentage

thrust-mixing efficiency, eta, defined by

eta =100(Ctrho t - Ctrcold )[ Ctrcold .

Here "hot" stands for hot core flow and

"cold" stands for cold core flow at the same

temperature as the fan stream. Recall that

there was a limit on the highest temperature

that the core-stream could achieve at

FluiDyne which prevented it from achieving

the core-to-fan associated with some of the

operating points. Hence, a knowledge of the

thrust-mixing efficiency allows us to

extrapolate, to the first order, the value of

the thrust-coefficient to higher core

temperatures than could be achieved in the

test facility.

Some of the important terms used in Table

4.1 are also defined in the footnotes below it

for convenience. This table includes the

results of two tests done at FluiDyne. The

first, completed in 1995, included only the

original four mixers (CONF, 12CL, 12UH

and 16UH) marked with an asterisk in the

NPRf column. The second was completed in

1997 and used a mix of the original and the

new mixers. All results are for the baseline

or 100% nozzle-length, and under static

conditions (M 0 = 0). Also note that the

"cruise" conditions differ slightly between

the 1995 and the 1997 tests.

4.1.2 Thrust Coefficients

Figure 4.1 shows a plot of the static thrustcoefficient, Ctr, as a function of NPRf for all

mixers tested at all operating conditions.

The margin of error in Ctr, from repeatability

tests, is around 0.15%. In general, we

observe that Ct_ increases with nozzle

pressure ratio, as expected. A better

perspective is obtained for comparison

between different mixers and between "hot"

and "cold" results for a given mixer by

focusing individually on different operating

conditions. Figures 4.2(a) and (b) show Ct_

in bar-chart form for the cruise condition

and the reference takeoff condition, TO #1

(see Table 3.1 for definition of TO

conditions).

We first note that at "hot cruise" (Figure

4.2(b)) all the lobed mixers produce a higher

thrust coefficient than the confluent mixer

(CONF), which has a Ct_ = 1.0000. The

CONF-mixer also has the lowest thrust

mixing efficiency, as measured by eta in

Table 4.1. This increase in thrust coefficient

for lobed mixers at cruise is, of course, the

primary reason for using them rather than

the confluent mixer for internal exhaust

nozzles.

NASA/CR 2002-210823/VOL1 77
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At cruise,the 12-lobemixer with cutouts

(12CL) showsone of the highestthmstcoefficients, 1.0055, in the 1995 tests.

However,notefromTable4.1thatwhenthis

samemixer was tested in 1997, its Ctr

reducedto 1.0013. This apparentdecrease

in Ct,,whichis abovethemarginof error,is

probably due to the "test-hardships"that

mixer 12CL wentthroughduringthe twoyearperiodbetweenthetwo FluiDynetests.

Somenicksanddentson the thin walls of

12CL were, indeed,observedduring the

1997testsatFluiDyne.

The unscalloped,high-penetration20-lobe

mixer (20UH) has the next highestcruise

thrust coefficient of 1.0048 which is

significantlyhigherthanits deeplyscalloped

counterpart(20DH)whichhasCtr= 1.0010.

This shows that deep scallopingimplies

thrustandtotal pressurelossesat cruiseas

comparedto unscallopedlobe walls and

shouldbekeptin mind whencomparingits

noisebenefits,if any,at take-off. Both the

20-lobemixers (20UH, 20DH) havefairly

highthrust-mixingefficienciesasmeasured

by eta (last column in Table 4.1). This

means the increase in their thrust

coefficients with an increase in core

temperature, say, to the actual core

temperaturedemandedat cruiseconditions,

will also be higherthanothermixerswith

loweretavalues.

The 12-lobeunscalloped,high-penetration

mixer (12UH) producesa cruise Ctr of

1.0042,which is comparableto 20UH. It

alsohasthehighesthrust-mixingefficiency,

eta,whichiscomparableto thatof 20DH.

The cruise thrust coefficient of the

remainingunscallopedmixerwith 16lobes,

16UH, is 1.0043 and is comparableto

12UH,but it hasa muchlower eta value.

Recall that 16UH has a very different

(larger)fan-to-corearearatiothantherestof

thelobedmixers(seeTables2.1and2.2),so

that its mass-flow rates would be very

different leading to a different (larger)

bypassratio, actualthrustsetc,eventhough

it is operatingat the samepressureand

temperatureratioastheothermixers. Thus,

whencomparisonsaremadelaterfor noise

characteristicsat constanthrust,weneedto

considernot only the thrust-coefficientof

the mixer-nozzleconfigurationbut also its

differentmass-flowratecharacteristics.

Most of the observationsmadeabovefor

cruiseconditions also hold good at lower

pressure ratios, like TO #1, shown in Figure

4.2(a), with a few exceptions. At this

condition, the confluent mixer (CONF) has

one of the highest Ct_ = 0.9951, but still the

lowest thrust mixing efficiency. The tongue

mixer (12TH) has the lowest Ct_ = 0.9818

implying larger total pressure losses due to

the strong mixing produced by its

streamwise vortices found earlier in the CFD

simulations (see Figure 2.25). However,

like 16UH, the tongue mixer (12TH) also

has a very different projected fan-to-core

area ratio (see Table 2.2), much lower than

the other mixers. This would lead to

different mass-flow rates (lower bypass

ratio) and different thrust at the same

operating conditions. Comparing the lobed

mixers tested at TO #1, the relative values

of Ctr remain similar to those at cruise:

12CL has the highest Ctr, 20UH has the next

highest, 20DH is lower than 20UH, and so

on. However, it should be noted that at TO

#1, a slight difference in the total

temperature ratio exists between the tests

done in 1995 and those done in 1997. For

configurations CONF, 12CL, 16UH, the

1995 tests set Ttc/Ttf at 2.34. For

configurations 12UH, 12TH, 20UH and

20DH, the 1995 tests were conducted at

Ttc/Ttf = 2.51. Also note that the reduction

NASA/CR 2002-210823/VOL1 80
in Cufor mixer 12CLobservedatcruisewas

also observedat the lower NPR when

comparingthe 1997and1995results(Ctr=

0.9932in 1997ascomparedto Ctr= 0.9952

in 1995).

In summary,amongthelobedmixers,12CL

consistentlyhas the highest static thrust

coefficient,with 20UHaclosesecond.The

unscalloped lobe mixers with high

penetration(12UH,16UH)havecomparable

performanceat cruise with referenceto

20UH. The deep scallopingpresent in

20DH introduces significant thrust

coefficient lossesat both the cruise and

takeoff conditions. The tongue mixer

(12TH) suffers the strongestlosses at

pressureratiostypicalattake-off.

4.1.3 Nozzle Discharge Coefficients

Figures 4.3(a) and (b) show that the static

nozzle discharge coefficients, CDnozzle, for

the various mixers at TO #1 and cruise

conditions respectively. The discharge

coefficient at "hot" conditions is lower than

"cold" conditions, as usual, for all the

mixers.

With no metal blockage, the confluent mixer

(CONF) predictably shows the highest

discharge coefficient. Mixer 12CL which

has the fewest number of lobes and lowest

penetration with cutout lobes, shows the

next highest discharge coefficient, as might

be expected from a metal blockage point of

view. Mixer 20UH, which has the highest

number of lobes, shows the lowest value of

discharge coefficient of all the unscalloped

configurations. However, mixer 16UH, with

large fan-to-core area ratio, shows the

highest value. Deep scalloping (20DH)

affects the discharge coefficient adversely

both at cruise and TO The tongue mixer

(12TH), with the lowest fan-to-core area

ratio, also has one of the lowest discharge

coefficient for both TO #1 and TO #2 (see

Table 4.1).

An examination of the effective area ratios

for the different mixers (see Table 4.1) also

reveals a correspondence with their

geometric area ratios (see Tables 2.1, 2.2).

For example, all mixers with similar

geometric area ratios (12CL, 12UH, 20UH)

also show similar "effective area ratios".

The effective area ratio does not change

much with nozzle pressure ratio.

On the other hand, deep scalloping (20DH)

definitely increases the effective area ratio.

This is in spite of the fact that the individual

fan and core flow areas, as axially projected

on a transverse plane at the mixer-exit, are

the same as its unscalloped counterpart,

20UH. A closer examination of the

individual effective areas for 20DH and

20UH shows that the increase in effective

area ratio is due to a larger decrease in

effective core flow area compared to the

smaller increase in effective fan flow area.

For example, at cruise, with deep-scalloping,

the effective core flow area is reduce by

11.6%, whereas, the effective fan-flow area

increases by only 3.5%. Recall that although

the static pressure at the scalloped edge of a

lobe must be balanced between the fan and

the core flows, the ratio of their total

pressures, Ptfan/Ptcor_ is greater than 1 (about

1.036). The fan stream, with higher total

pressure head, appears to "pinch" or narrow

down the flow area available for the core

stream, as is also seen in the CFDsimulations of the total temperature contours

in Figure 2.21(a). Although the absolute

changes in the effective areas of the core and

fan streams due to scalloping are of the same

order of magnitude, for a high bypass ratio

mixer-nozzle such as these, the change in

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core area on a percentage basis is larger than

that in fan area.

a brief look at least at their bypass ratios is

useful.

4.1.4 Bypass Ratios

Static Case

Figure 4.4 shows the measured bypass ratio

(BPR) for all mixers tested at FluiDyne

under all operating conditions. For all

mixers, the change in BPR due to operating

condition is not large. However, for the

lobed mixers bypass ratio does peak at an

intermediate nozzle pressure ratio and then

decrease to a minimum at the cruise

condition. Comparing the various lobed

mixers, configurations 12UH and 20UH,

which have the same fan-to-core area ratio,

have comparable BPR's at all pressure

ratios. Mixer 12CL, which has cutouts in

the lobes but the same fan-to-core area ratio

as 12UIt, has a slightly higher BPR than

12UH. By comparison, mixer 16UH, with

the highest fan-to-core area ratio, has the

highest BPR. Similarly, mixer 12TH, with

the smallest fan-to-core area ratio, has the

smallest BPR. Deep scalloping (20DH)

increases the BPR compared to 20UH at all

operating conditions, as expected from

previous discussions.

All the tests at FluiDyne were performed

under static conditions, that is with a flee-jet

Mach number of 0.0 and with the baseline

(100%L) nozzle-length. Hence, the effect of

Mcj or nozzle-length on the aerodynamic

performance parameters cannot be

determined from the FluiDyne test data.

However, the acoustic tests done at NASA's

APL did vary Mr] and nozzle-length, but

were restricted to the measurement of the

mass-flow rate and not the thrust or the

discharge coefficients. Since both M 0 and

nozzle-length turn out to have significant

effects on the noise characteristics of mixers,

Effect of Free-Jet Mach No.

Figure 4.5 shows the effect of free-jet Mach

number, Mfj, on BPR at performance

condition TO #3 for mixers with 100%

nozzle-length using data obtained during

acoustic tests done at NASA. The overall

trends in terms of bypass ratio at static

conditions, amongst the different mixers is

the same as found in the FluiDyne tests

(Figure 4.4) That is, the static BPR's of

12UH, 20UIt are comparable to each other

while 12CL is slightly higher than both. As

before, 16UH has the highest BPR while

12TH has the lowest. The addition of

scalloping increases the BPR, with the

additional information from the 20MH

mixer that BPR is in some sense also

proportional to the scalloped area.

(Compare BPR's of 20UH, 20MI-I and

20DH.) However, the absolute values of

BPR are slightly different between the

NASA and FluiDyne tests at the same

operating conditions, and with the NASA

tests consistently showing a slightly larger

BPR value than the corresponding FluiDyne

data. The difference is most likely

attributable to the differences in the

measuring instruments and the data

processing methods at the two facilities.

Since we are interested in studying the effect

of M_j on BPR from tests done at one

facility, the difference in BPR-values are of

primary importance rather than their

absolute values. From Figure 4.5 we can see

that the effect of changing M 0 from 0.0 to

0.3 on the bypass ratio is generally not

much. Two exceptions are noted: a

significant increase in bypass ratio is

observed in the confluent configuration as

the free jet Mach number increases;

however, mixer 12CL experiences a

NASA/CR 2002-210823/VOL1 83
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reduction in bypass ratio between free jet

Mach numbers of 0.2 and 0.3.

Effect of Nozzle Length

Figures 4.6(a) and (b) show the effect of

decreasing nozzle length (100%L to 50%L)

on BPR for the mixers at M 0 = 0.2 and

performance conditions TO #1 and #3,

respectively. Three of the lobed mixers

(12TH, 20UH and 20DH) also had data

taken with the 75%L nozzle. In general, it is

observed that bypass ratio decreases with

decreasing nozzle length, for both

performance conditions. This trend is

monotonic over the range of nozzle lengths

tested.

In order to understand the reasons for this

monotonic decrease in BPR with decreasing

nozzle length, the changes in the individual

fan and core flow-rates and total flow rates

need to be scrutinized. These comparisons

will be made at standard day conditions to

remove the effects of changes in ambient

conditions through the use of "corrected"

flow rates. The corrected mass flow rates

are defined, as in the turbomachinery

literature:

mcorr =

where

rrei

Pre 

mmeasured _Tt / Zref

= 518.7 ° R

= 14.7psi.

Figure 4.7 first shows the bypass ratio

defined in terms of individual corrected

flow. The previously observed trend does

not change when corrected flow is used to

determine bypass ratio: the corrected bypass

ratio also decreases as the nozzle-length is

decreased for all mixers. Figure 4.8 shows

the individual corrected mass-flow rates for

fan and core streams in a comparative barchart form for several mixers. A general

trend for all the mixers is that the fan flow

rate decreases when the nozzle-length

decreases and, with the exception of 12UH,

the core flow rate increases simultaneously

so as to approximately maintain the same

total flow rate.

We will attempt to list the possible

ingredients of the mechanism to explain the

above mass-flow rate data, however, the

complete aerodynamic mechanism is not

fully understood at this point. Recall that the

reduction in mixing length was achieved by

preserving both the nozzle throat area (or the

nozzle exit diameter) and the mixing plane

diameter. This reduction in mixing-length

has the effect of "squeezing" the nozzle duct

flow in a shorter axial distance and

"releasing" it to ambient pressure at the

nozzle lip "earlier" than a longer nozzle.

This reduction in mixing-length would

imply less mixed flow at the nozzle exit

plane. "Squeezing and early releasing"

would imply faster axial expansion rate and

higher acceleration of the subsonic mixed

flow inside the shorter nozzle than that in

the longer nozzle. For the shorter nozzle, the

aft half of the flow outside the nozzle

dominated by axial vorticity, is now in an

"unbounded' region rather than in the

"bounded" duct environment. All of these

processes downstream of the nozzle must

affect the axial evolution of fan/core flow

mixing upstream at the mixer exit plane.

However, both the subsonic streams must

have the same static pressure at the trailing

edge of the mixer (Kutta condition). This

common static pressure can be different for

the two nozzles and can change the effective

throat areas for the fan flow and the core

flow while keeping the sum of their areas

the same due to the same duct crosssectional area near the mixer exit plane. This

NASA/CR 2002-210823/VOL1 85
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NASA/C_2002-210823FVOL l 86
Effect of Nozzle Length on Corrected Bypass

Ratio at T.O. # 3 and M(fj) = 0.2

5.0

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Figure 4.7 Effect of nozzle length on

corrected bypass ratio

can explain the simultaneous rise in one

mass-flow rate (core flow) and decrease in

the other (fan flow) with the total mass-flow

rate remaining the same. Also based on one

dimensional ideal-mixing arguments, which

is independent of the mixing length, the total

flow rate of two streams should not change

with nozzle length when their pressure ratios

and temperature ratio are not changed. This

is partly what is observed in Figure 4.8,

except for the 12UH mixer. However the

internal mixing is by no means ideal and

both streams suffer different total pressure

losses when the nozzle-length is decreased.

A possible mechanism for the decrease of

fan mass flow rate when the nozzle length is

decreased is that the axial vortices (which

migrate radially outward inside the nozzle as

seen in the CFD simulation in Figures 2.21,

2.22 or 2.24, 2.25) close to the aft

convergent part of the nozzle partly block

the fan flow near the duct wall. How far

radially these axial vortices move depends

on the lobe penetration and their radial

convection velocity. As a result, the higher

the penetration the more the fan blockage

due to these axial vortices. This is partly

borne out in the observed differences in fan

flow changes between 12CL and 12UH or

20UH. In all mixers, it turns out that all

these mechanisms collude to decrease the

fan flow rate when the nozzle-length is

decreased. It appears then, that to maintain

approximately the same total mass flow rate,

as demanded form the 1-d ideal mixing

analysis, the core flow rate must increase.

Further deeper investigation is needed to

fully understand this aerodynamic effect of

shortening the nozzle length.

4.2 Typical Plume Survey Results

In this section, typical plume survey results

are presented as obtained in static tests at

FluiDyne and in tests done at the NASA

APL with a free jet flow. Several of these

plume surveys are collated in Volume 2 as

plots of total temperature or local velocity

distribution, and Table 4.2 gives overall

guidance to those figures for each mixer

tested. The purpose of this section is to

selectively discuss some of these plume

survey plots so that a better understanding

can be obtained regarding the plume flow

physics and its connection to the far field

noise which is presented in the next chapter.

4.2.1 Nozzle Exit Plane Survey

Figure 4.9(a) shows the total temperature

contours for a section at the nozzle exit

plane for the four original mixers (CONF,

12CL, 12UH 16UH) with the baseline

nozzle-length (100%L), at performance

condition TO #1 and Mfj = 0 (static case).

Figure 4.9(b) shows the radial temperature

profiles for the same case and compares

NASA/CR 2002-210823/VOL1 87
12CL

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NASA/C_2002-210823/VOL 1 88
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NASA/CR2002-210823/VOL1 91
them to the ideally mixed temperatures

foundfrom the knownmassflow ratesand

theenthalpyconservationequationfor each

mixer. Thetotal temperatureis shownin a

non-dimensional form defined by

r?

=1 implies core temperature and r?

= 0 corresponds to the fan temperature.

We make the following observations:

(i) The confluent mixer (CONF) shows

hardly any mixing between the core

and the fan flow, as expected. Viscous

mixing is the dominant mechanism

available in this configuration and it is

obviously not effective in the available

mixing-length. The Kelvin-Helmholtz

instability which occurs in vortex

sheets or thin shear-layers and which

increases the interface area between

the two streams, is apparently not a

strong mixing mechanism for the

confluent mixer which has a fairly high

fan-to-core velocity ratio (Vf/Vc= 0.7).

(ii) The forced lobe mixers, on the other

hand, produce much better mixing than

the confluent, as can be seen from the

smaller deviation of the local

temperature from the ideally mixed

temperature (Figure 4.9(b)). One can

also qualitatively say from these

figures that 16UH is much better

mixed than 12UH, which in turn is

better mixed than 12CL. The

mechanism for enhanced mixing in

forced lobe mixers is by now well

known: It is largely attributable to the

streamwise or axial vortieity generated

by the difference in radial velocity

components of the core and the fan

flows near each lobe sidewall. This

(iii)

(iv)

axial vorticity generated downstream

of the lobe sidewalls rotates the two

flows around each other in tight

spirals, increasing the interface area,

and producing better mixing though

the "engulfment' process. The relative

strength of the vortices generated and

their convection in the radial direction

depends on the lobe geometry as given

by parameters such as the angle

between the two flows at the mixer

exit plane, lobe penetration, and

number of lobes.

The effect of lobe penetration on the

radial location of these vortices at the

nozzle exit plane is clearly captured in

these figures by the location of "hot

spots". The local peaks in temperature

are created by the partially mixed hot

core stream and are typically found in

the central core region of an axial

vortex as seen in the previous CFDsimulations (see Figures 2.21, 2.22 or

2.24, 2.25). Thus, we observe that

12CL, with low lobe penetration, has

its hot spots and, hence, axial vortices

closer to the central axis than 12UH

and 16UH mixers which have higher

lobe penetration. Since the nozzle exit

radius is also smaller than the radial

height of the lobe crests in 12UH or

16UH, it appears that these axial

vortices may even be interacting with

the nozzle duct wall at the very aft end.

This data also gives some credence to

the fan-flow blockage mechanism

discussed earlier in Section 4.1.4 for

the shorter nozzle.

Figure 4.9(b) shows that, even in the

lobed mixers, the region near the

central axis is not well mixed.

However, the mixing improves as we

go from 12CL to 12UH to 16UH.

NASA/CR 2002-210823/VOL1 92
Mixing in thecentralregionis believed

to becontrolledbythegapatthemixer

exit planebetweenthe lowestpartof

the lobe andthe tail-conesurface,as

well ashow closelythe fan flow exit

anglematcheswith thetail-coneangle.

An improvementin mixing can be

obtainedin this regionby decreasing

the gap as much as practical and

aligningthefanparalleltothetail-cone

surfaceatthemixerexitplane.Further

confirmationof the relative mixing

efficienciesbetweenthesefour mixers

is alsoobtainedfromFigures8 through

11in Volume2, pt. 2 wherethefreejet Machnumberis setto M_= 0.1for

the sameTO #1 operatingcondition.

Theseplots tendto imply that higher

lobenumberandhigherpenetrationare

conduciveto fastermixing. Boththese

geometrical features increase the

interfaceareabetweenthefan andthe

core streamsand enhancesmixing.

However,weshouldalsobearin mind

that16UHhasalargerfan-to-corearea

ratiothanthe othermixers,and12CL

has cutouts, so that there may be

secondarycompound effects when

comparedwith 12UH.

4.2.2 The Tongue Mixer and Mixer with

Cutout Lobes

The effect of the tongue mixer (12TH) on

total temperature mixing near the nozzle exit

plane (X/D = 0.2 where X = axial distance

from nozzle exit plane and D = nozzle exit

diameter) is shown in Figure 4.10(a) where

it is also compared with 12CL (Figure

4.10(b)) at the higher NPR of TO #3 and Mfj

= 0.2. The total temperature is given in

degrees Fahrenheit. Here again we observe

that the hot-spots for the tongue mixer,

which has high penetration, have migrated to

the outermost extremities of the nozzle

circumference and even beyond, since at

X/D = 0.2 the hot spots can penetrate the

nozzle-lip shear layer. For the tongue mixer

the central region remains unmixed as in the

previous mixers and, at this higher

temperature ratio condition, the "stems" of

the mushroom-type axial vortices are still

visible, denoting less azimuthal mixing

there. However the temperature of the

peripheral hot-spots in the tongue-mixer are

less than half of that in the 12CL mixer

(compare Figures 4.10(a) and 4.10(b)). This

means that the tongue-mixer has much more

vigorous mixing and dissipation than 12CL.

We can also compare this data with the

CFD-simulation of the tongue-mixer at a

lower nozzle pressure and temperature ratio

in Figure 2.24(b) at the nozzle-exit ("X"/L =

1.0 where "X" is from the mixer exit plane

and L is the mixing length). Some

similarities and some dissimilarities are

observed. For example, the size and

location of the outer hot-spot and its "stem"

are well-captured -- the hot-spot is expected

to radially migrate further when it reaches

X/D = 0.2, as in Figure 4.10(a). However,

the central hot-spot is not captured in the

CFD-simulation, where, instead, CFDresults show a thin hot annular ring and a

cold spot at the center. This is probably due

to the problems in CFD simulation near the

central axis of the cylindrical domain of

computation.

Figure 4.10(b) for mixer 12CL can also be

compared to the static case for 12CL at

lower pressure ratios (Figure 4.9(a)). This

comparison shows that, even at higher

nozzle pressure and temperature ratios the

radial location of the axial vortices has not

changed much, implying that the vortex

position is strongly dictated by the lobe

penetration at all operating conditions. In

addition, the central hot-spot has not

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diffused at all at this axial station. Note that

in both mixers (12CL and 12TH), since a

"core" lobe is at the 12 o'clock position and

there are 12 hot-spots in Figure 4.10, the

axial vortices from the neighboring lobe

sidewalls of a "core" lobe, although of

opposite signs, coalesce first rather than

those from a "fan" lobe. It is only further

downstream that all these 12 hot-spots

merge azimuthally to form an annular ring

of higher temperature implying that the jets

will quickly become axisymmetric. This

axisymmetry will be more evident in the

next section. The radial temperature

gradients, however, are also unsteady

entropy regions and, thus, generators of

extra noise beyond that created by radial

gradients in velocity. It should also be noted

that although a free jet flow is present (Mq =

0.2) the overall hot-spot structure has not

changed much from the static case (Figure

4.9(a)), presumably because with low lobe

penetration the axial vortices are still away

from the growing nozzle-lip shear layer.

4.2.3 Plume Downstream Evolution

The downstream evolution of the plume

velocity profile for mixer 12CL at operating

condition TO #3 and Mej = 0.2 is shown in

Figure 4.11 at Y = 0.0 and several

downstream stations. [Note that the

measurements are obtained on a rectangular

Cartesian grid with Y in the horizontal

direction in the transverse plane and Z in the

vertical direction in this plane. Hence, plots

for constant Y-values are along vertical

"chords" lines and not radial lines. Y = 0

with variable Z values represents

measurements along a vertical diameter

which passes through the top and bottom

"core" lobes.] The plume velocity is

calculated from the measured Pt, Ps and Tt

values and the isentropic relations.

Some general observations follow from

Figure 4.11:

(i) The initial complicated structure of the

velocity profiles at lower X/D-values

quickly gives way to a simpler plume

further downstream. Although not

shown here but evident from the

detailed profiles shown at various Yvalues in Vol. 2, part 1, figures 40 - 46,

the plume quickly becomes

axisymmetric further downstream.

(ii) The radial gradients in axial velocity

govern part of the turbulence intensity

and, hence, are strong sources of noise.

Thus, the plume will generate noise

not only from the radial gradient in

velocity at the nozzle-lip shear-layer

but also from these axial vortex

structures, hot-spots and the central

velocity peak. These are "excess"

noise sources, in the sense, that they

will not occur in a jet with an

equivalent uniform velocity and

temperature profile at the nozzle-exit

plane.

(iii) The thickening of the lip shear-layer

downstream is evident from the

decreasing radial slope of the velocity

profiles near Z = + 3.625 in.

The center-line velocity decay will be

examined in a later section where it will be

compared with other mixers.

4.2.4 Effect of Scalloping

Figures 4.12(a) and (b) show the typical

effect of scalloping on the total temperature

profiles at X/D = 0.2 for the lower nozzle

pressure ratio condition of TO #1 with M 0 =

0.2. The peripheral hot-spots in the

unscalloped mixer (20UH) diffuse out into

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two smallerhumpsfor thedeeplyscalloped

mixer (20DH). Comparingthe data of

Figure4.12(b)with theCFD simulationfor

the 20DH mixer configurationin Figure

2.21(b),it appearsthatthehot-spotof 20UH

atZ= _+3.0in. hasactuallymigratedradially

outwardin the20DHmixer andis reduced

in strength. At the sametime, the second

inwardhot-spotin 20DH at Z = _+1.75in.

with strengthEta = 0.2 is from the "stem" of

the mushroom vortex seen in Figure 2.21(b).

In any case, although scalloping appears to

weaken the outer hot-spots, the central hotspot is not affected. The central hot-spot, as

mentioned earlier, is primarily a function of

the gap between the lowest portion of the

lobes and the tailcone surface and which is

the same in 20UH and 20DH.

Figure 4.13 shows the velocity profile for

the deeply scalloped mixer (20DH) at the

axial location X/D = 0.2. The peripheral

velocity is very well-mixed as compared to

that of mixer 12CL (Figure 4.11(a)). The

unscalloped mixer (20UH), with larger

peripheral humps in the temperature, is not

expected to be as uniform as 20DH. The

central region in Figure 4.13, on the other

hand, still shows a small hump in the

velocity profile of 20DH which would,

indeed, act as an "excess" noise source due

to shear.

4.2.5 Effect of Nozzle-Length

We compare the effect of halving the

mixing-length using data for the 12CL

mixer. Since reducing the nozzle-length

from 100%L to 50%L for 12CL also

surprisingly reduced noise (Section 5.6.2)

we will study the plume profiles for these

two nozzle lengths in more detail.

Figure 4.14 shows the total temperature

contours at X/D = 0.2 for the 12CL mixer

with a 50% nozzle length at performance

condition TO #3 and Mfj = 0.2. Compare

this with 100%L at same X/D = 0.2 (Figure

4.10(b)). It appears that 50%L is not as

well-mixed as 100%L as expected. In the

50%L case, the "stems" of the mushroom

axial vortices are still visible, the hot-spots

on the periphery are still well separated

azimuthally, and the central hot-spot is

slightly larger. Further downstream at

X/D=0.5 the same differences between

100%L and 50%L are still observed (see

Figures 5, 6 in Volume 2, part 2). Thus,

with a shorter mixing-length we can safely

conclude that the flow is not as well mixed

as with the longer mixing-length at the same

distance from the nozzle exit-plane.

Comparison of velocity profiles for the two

nozzle lengths at X/D = 0.2 for operating

condition TO #3 is shown in Figures 4.15 (a)

and (b). The local velocity peaks at the

periphery for the 50% nozzle-length appears

higher than for 100%L (compare Z = -2.25

in.) implying less mixing with the shorter

nozzle-length.

The above comparisons were done at the

same distance from the nozzle exit plane.

Since mixing between the fan and the core

streams begins at the mixer exit plane as

opposed to the nozzle-exit plane, it is more

instructive to compare the flow profiles

between the two nozzle lengths at same

distance from the mixer exit plane. This

will also bring out the effect of removing the

nozzle-duct in the aft portion of the mixinglength and exposing the partially mixed flow

to the ambient pressure sooner. In both

nozzle-length cases, since data was taken

only at discrete axial locations (X/D-values

of 0.2, 0.5, 1.0, 3.0, 5.0, 7.5 and 10.0) and

L/D = 1.555 for the baseline nozzle, it turns

out that there is only one axial location

where data exists at approximately the same

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distance from the mixer-exit plane for the

100%L and 50%L configurations. An axial

location of X/D = 0.2 for the 100%L nozzle

corresponds approximately to X/D = 1.0 for

the 50%L nozzle. Figures 4.16 (a) and (b)

show a comparison of the velocities at these

two axial locations at different values of Y

in the transverse plane for the two nozzlelength configurations with mixer 12CL

installed. It is very interesting to see that the

mean velocity profiles, at all Y-locations

shown, match very well indeed - except for a

small shift in the Z-direction for local peaks.

Since the flow at X/D = 0.2 for the 100%L

nozzle is still adjusting to the ambient

pressure, having only been recently released

from the nozzle, there may be statistical

variations even in the measured mean flow.

Hence, it may be better to compare the flow

profiles for these two configurations with

different nozzle-lengths further downstream

when flows from both the nozzleconfigurations have had sufficient time to

adjust to the surrounding environment.

However, as mentioned earlier, we do not

have comparable data for both cases further

downstream at the same distance from the

mixer exit plane. Hence, we will compare

only the center-line velocities for these two

cases after appropriately shifting their axial

locations so that now they are referenced

from the mixer exit plane. Figure 4.17

shows this comparison for the center-line

velocities where simple linear interpolation

is used between data points. Apart from the

initial small variations, the center-line

velocities for 100%L and 50%L match

rather well for Xmixerexitplane/D _" 4.0.

Figures 4.16 and 4.17 imply the following:

(i) The appropriate axial distance when

comparing flows from mixer-nozzle

configurations of different mixing-

(ii)

lengths is not relative to the nozzleexit plane, but is rather from the

mixer-exit plane.

Removing the aft-half of the duct

length in the internal mixing region for

mixer 12CL has little effect on the

mean velocity profiles at the same

distance from the mixer-exit plane

when compared downstream of the

longer nozzle. Hence the noise

radiating from the far downstream

portion of the two nozzles ought to be

similar, whereas the difference in the

noise spectra between the two nozzles

should be attributed to stations

upstream of at least X/D = 0.2 (with

reference to baseline nozzle), that is, to

stations close to the nozzle exit plane

both inside and outside the nozzle

duct.

As will be seen in the next chapter, this is

borne out for the two nozzles, respectively,

in the similarity of the low frequency spectra

radiating mainly from the large eddies in the

far downstream region and the dissimilarity

of the high frequency spectra corresponding

to the small-scale eddies near the nozzle exit

plane

4.2.6 Center-line Velocity Comparisons

In these lobed mixers we have seen that the

initial non-axisymmetric flow profiles

become axisymmetric a short distance

downstream of the nozzle exit due to the

good azimuthal mixing produced by the

axial vortices. Hence, a center-line velocity

decay comparison can give a fair idea of the

overall relative noise levels produced by

these mixers, especially the plume decay far

downstream where larger eddies govern the

low frequencies. With this in mind, we

compare the "center-line" velocities for a

few mixer-nozzle configurations in Figure

NASA/CR 2002-210823/VOL1 105
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1450

12CL Mixer with 100% and 50% Nozzle Length

T.O. # 3 and M(fj) = 0.2

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1350

1300

U

0

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1150

1100

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0 1 2 3 4 5 6 7 8 9 10 11 12

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Figure 4.17 Comparison of center-line velocities at same distance from the mixer exit plane

for nozzles with different lengths.

NASA/C_2002-210823/VOL 1 108
4.18 with semi-log axes for performance

condition TO #1 and Figure 4.19 for TO #2

and #3 at Mfj = 0.2.

From Figure 4.18 we observe that for all

mixers there is only a very short-distance

(X/D _< 1.0), beyond the nozzle exit plane,

where the center-line velocity is constant,

downstream of this location it starts

decaying. This does no___3trepresent a

"potential" core in the normal sense, as we

have seen in Figures 4.11, 4.13 or 4,15, but

simply a small portion of the hot core fluid

which flows over the tail-cone and which

quickly mixes with the surrounding fan flow

downstream. The center-line velocities

between the 12-lobed mixers and the 20-

lobed mixers are distinctly different in the

beginning with a reduction in magnitude of

about 50 ft/s for the 20-lobed mixers.

However, further downstream, around X/D

of 8, all the mixers seem to have similar

center-line velocities. It should be noted

that if the 100%L and 50%L configurations

for 12CL are shifted to correspond to the

distance from the mixer-exit plane, they

overlap as before in Figure 4.17.

Figure 4.19 for the higher nozzle pressure

ratios shows similar trends, however with

larger differences in the initial magnitudes

which decay to the same levels at X/D = 8.

We have discussed in this section only some

of the most pertinent plume-survey data.

Volume 2 contains the rest of the data

collected and should be referred to as the

need arises in understanding the acoustic

data discussed in the next chapter.

NASA/C_2002-210823/VOL 1 109
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NASA/C_2002-210823/VOL 1 111

Chapter 5

Acoustics Results & Analysis

5.1 Overview

The acoustic results for the various mixers

are presented and analyzed in this chapter.

Due to the vast amount of data collected,

representing variations in a number of

different parameters including both

operating conditions and mixer-nozzle

geometry, systematic presentation of results

to isolate parametric trends was a major

challenge.

In order to capture the effects of mixer

geometry between the two sets of lobed

mixers (the original four designs from 1995

and the new configurations from 1996),

results will be presented in terms of 3

general groupings (see Tables 2.1 and 2.2

for nomenclature and mixer properties):

Group 1. All 12 lobed configurations

(12CL, 12UH, 12TH).

Group 2. All unscalloped configurations

(12UH, 16UH, 20UH).

Group 3. All 20 lobed configurations

(20UH, 20MH, 20DH).

Group 1 examines the compound effects of

penetration and sidewall cutouts (with the

tongue mixer configuration (12TH)

representing, in some sense, the limiting

case of deep scalloping). Group 2 examines

the importance of lobe number, while Group

3 isolates the effect of sidewall scalloping.

In addition, comparisons between members

of different groupings are also used to

illustrate particular points. In all eases, the

confluent or coaxial mixer (CONF) is taken

as the baseline configuration.

Initially, comparisons between the mixer

configurations will be made with the

reference nozzle length (100%L) and freejet Mach number of 0.2. In the second

phase, the effects of free-jet speed on farfield noise will be examined. In that section,

a new method for diagnosing the location of

high-frequency noise sources is presented

and its importance examined. Finally, the

effect of nozzle length on the noise

produced by each of the mixer

configurations is examined, with some new

and interesting findings presented. All

results presented are for a nozzle with an

exit cross sectional area of 660.5 in 2 (or 29

inches diameter), representing a scale factor

of four.

While analyzing these acoustic results, two

questions arise: (a) How and why do the

"noise source characteristics" change with

the parameter being examined? (b) How

"important" is that parametric effect? The

first question is answered by examining the

far field noise characteristics in the

"reference frame of the nozzle". By this, we

mean the sound pressure or sound power

spectra whose directivity characteristics

have been adjusted to remove the refraction

effects which occur across the shear layer at

the free-jet/ambient interface. This data is

presented on a circle of 150 ft radius whose

center is coincident with the center of the

nozzle exit plane. All such polar data has

been corrected to reference conditions of

14.3 psi, 77°F and 70% relative humidity.

The polar spectrum has no Doppler-shift

corrections applied. The second question,

relating to the "importance" of a particular

parameter, is examined using perceived

NASA/CR 2002-210823/VOL1 113
noiselevel (PNL) and effectiveperceived

noiselevel (EPNL)and the corresponding

sound pressurelevel as observedby a

groundbasedobserverduring a simulated

flyoverof thetestnozzleat a fixed altitude

of 1500ft For this flyover simulation,the

SPL spectrais appropriatelyadjustedfor

slantdistancepropagationeffects,standard

day atmosphericattenuation,Dopplershift

due to aircraft motion relative to the

observer, and an ambient pressure at

standarddayof 14.7psia. For theDoppler

shift, the aircraft is assumedto fly at the

velocityof thefreejet appropriateto thetest

point considered. In addition,the ground

aroundthe observeris assumedto benonreflecting.

As mentionedearlier,the secondacoustic

testdoneatNASAin 1996repeatedmostof

the data-pointsdone in 1995 and added

manymore operatingpoints, and thus, in

many ways, supersededthe previous test

data. We presentandanalyzein thisreport

only the1996acoustictestresults.Someof

the 1995 acoustictest resultshave been

reportedin Mengleet al(16)and shouldbe

consultedforthesakeof completeness.

someglobal integralacousticmetric.

At the end of the chapter, a

comparisonis presentedof theirglobal

acousticmetrics,like flyover EPNL,

on a consistentbasissuchasconstant

thrust,sothatanobjectivemeasurecan

be obtained for their acoustic

suppressionbenefits. But thatcanbe

doneonlyif datais availableatseveral

operatingconditionsandinterpolation

of datais assume&Ontheotherhand,

if the mixers differ widely in their

geometricproperties,for instancearea

ratio (compare 12UH versus 12TH

versus 16UH in Tables 2.1, 2.2), then

their mass-flow rates and bypass-ratios

are widely different, and inter-mixer

comparisons, even though at the same

operating conditions, are not apt.

(ii) Effects of other "non-mixer"

parameters on a given mixer, such as,

nozzle-length or free-jet speed or

operating condition can, however, be

legitimately studied from an acoustics

point of view, although these

parameters may change the mass-flow

rates, etc.

A few general things to be kept in mind

while considering these data are:

5.2 Mixers with Reference Nozzle

Length at Baseline Take-off Condition

(i) When comparing acoustics of different

mixers at the same operating

conditions we may not necessarily

obtain the same mass-flow rates or

thrust. These aerodynamic quantities

are functions of the geometrical

parametric changes, as studied in the

previous chapter, and are themselves

independent objects of study. Hence,

the purpose of these inter-mixer

acoustic comparisons at the same

operating conditions is primarily to

establish the "relative shape" of their

acoustic/noise spectra, rather than

In this section we examine the noise

characteristics of the various mixer

configurations with the reference (100%)

nozzle-length (nominal mixing length L]Ornp

= 1.10) at a baseline operating condition, TO

#1 (nominal NPRf - 1.44, NPRc = 1.39 and

Ttc]Ttf = 2.34). Initially, we will concentrate

on the noise produced under static (no free

jet flow) conditions. At static conditions

there is no correction for free-jet/ambient

shear-layer refraction and the Doppler effect

is not present. Thus a "basic acoustic

datum" is created for later comparisons.

NASA/CR 2002-210823/VOL1 114
5.2.1 Static Free-Jet Data

5.2.1.1 12 Lobed Mixers

Figures 5.1 (a), (b), and (c) present,

respectively, (a) the 150 ft polar overall

sound pressure level directivity (OASPL)

and sound pressure level (SPL) spectra for

selected angles, (b) sound power spectra

(PWL) and sound pressure level directivity

for selected frequencies, and (c) PNL

directivity and flyover SPL spectra at

selected angles for mixer Group 1 (12 lobe

configurations).

The following characteristics stand out in

this group under static conditions:

(i) All lobed mixers are quieter for angles

between 125 ° and the jet axis than the

confluent (CONF) mixer, as seen in

the OASPL data (Figure 5. l(a)).

(ii) With the exception of 12CL, they do

this at the expense of increasing

OASPL in the upstream quadrant.

(iii) At shallow angles, like 150 °, the low

frequencies (100-300 Hz) dominate for

all mixers as seen in the polar SPL

spectra in Figure 5.1(a). This familiar

result is produced by the refraction of

sound waves away from the jet exit

axis as they propagate through the

negative radial gradient of the axial

velocity in the plume. This effect is

much stronger for higher frequencies

than for lower frequencies. This is a

feature typical of all such jets in the

similarity region far downstream of the

nozzle-exit plane as seen in Chapter 4

plume data. As a result, at 150 ° even

though 12UH and 12TH are louder for

mid-to-high frequencies (higher than,

say, 700 Hz), it is the large decrease in

low frequency SPL observed with

these mixers that drives their OASPL

below that of the confluent or 12CL

configuration.

(iv) "Low" frequencies are associated with

(v)

(vi)

(vii)

the strength of "large" eddies which

exist far downstream of the nozzle

exit. The observation that the SPL

spectra for 12UH and 12TH show a

reduction in low frequency amplitudes

for all angles (60 °, 90 °, 125 ° and 150 °)

implies that these two mixers have the

weakest large eddies amongst the

Group 1 mixers. This is intimately

related to the thoroughness of the

mixing achieved between the fan and

the core flow, and how that flow

subsequently mixes with the ambient.

This will be explored in somewhat

more detail later.

The reduction in low frequency noise

achieved in 12UH and 12TH is also

captured in the PWL-spectrum of

Figure 5. l(b).

The distinguishing feature of the SPL

spectra of configurations 12UH and

12TH for angles away from the jet

axis, as shown in Figure 5.1(a), is the

dramatic increase in noise level for

mid-range frequencies (approximately

1500-2000 Hz), producing a "hump"

in the spectra. In fact, at all these

angles (60 °, 90 °, 125 °) these humps

are also the "global" peaks or the most

dominant frequencies. Their peak SPL

frequencies shift slightly to higher

values as we go closer to the jet exit

axis. The high-frequency noise

contributions (5000 Hz and above) of

12UH and 12TH are also higher than

CONF or 12CL. Hence, it must be the

large mid-to-high frequency

contributions of 12UH and 12TH that

make their OASPL larger than CONF

or 12CL for angles upstream of 110 °.

"Mid-to-high" range frequency noise

is attributable to "smaller" eddies

which predominates near the nozzle

exit plane or even inside the nozzle.

These small scale eddies are produced

in lobed or tongue mixers by the

NASA/CR 2002-210823/VOL1 115
OASPL Directivity Comparison

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(_) = 0.O

106

104

102

_m" 100.

9B,

9B,

94.

92

90

88

55

--II-- 12CL I

_.o.. 12UH L._ ""

--_-- 12TH

*l*l';':" ;" :';':':':"

65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

A

m

.J

n

95 •

93.

91.

89,

87,

85,

a.

o_ 83,

81,,

79,.

77,.

75

10

150' SPL Spectra Comparison at 60 degs

NPR(f) = t.44, NPR(c) = t .39, 3"TR = 2.34, M(fj) = 0.0

85,

83,

81,

79,

77,

75,

73

71

69

67

65

• ...12UH F"'f

10 100 1000 10000

freq (hz)

150' SPL Spectra Comparison at 125 degs

NPR(f) = 1.44, NPR(c) -- 1.39, TTR = 2.34, M(_) = 0.0

---(P- CONF

---41-- 12CL

12UH

--X-- 12TH

/

a | i i .... .=

100 1000 10000

freq (hz)

.J

0.

u3

t 50' SPL Spectra Comparison at 90 degs

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(t]) = 0.0

90.

88.

86,

84,

_. 82,

" 80,

ol.

78,

76,

74.

72,

70

10

"-Ak--CONF

--11--12CL

_;12UH

...._--12TH

* | | .a|¢- = ........ . .......

100 1000

freq (hz)

150' SPL Spectra Comparison at 150 degs

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj) = 0.0

100

95 4;

90 4:

85 4,.

80 4:

75 4-

'° I

65 , " ........

CONF l

--II-- 12CL

• : 12UH

....-',K-- 12TFI

10000

10 100 1000 10000

freq (hz)

Figure 5.1 (a) OASPL directivity and SPL spectra for CONF, 12CL, 12UH and 12TH at T.O. # 1, M(t]) = 0.0

NASA/C_2002-210823/VOL 1 116
Power Spectrum Comparison

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2,34, M(f]) =

0.0

134-

132 I

1301__._ CONF

128 "1 --t-- 12CL . :- '.<_

| .. _:;.. 12UH X_ (;'"

126 l _ 12TH

t ---

o. 122._

120 t

118 1

10 100 1000 10000

freq (hz)

Directivity Comparison at 1600 hz

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj)

= 0.0

90 .

86 , ' ,

84

F

70 _ , I ' I , I , I , I , I , I ,, i , I , I ,

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

g

.J

u)

100

95

90

85

8O

75

7O

Directivity Comparison at 160 hz

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj)

= 0.0

--4-- CONF

;--I-- I:;LH /_

I ' I ' I " _ ..... I ' I"' I ' J i

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

Directivity Comparison at 5000 hz

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj)

= 0.0

81-

d_'75

73

71 I

69

67

65 , _

55 65 75 85 95 105 115 125 135 145 155 165

85

Angle from Inlet AXis (deg)

Figure 5.1 (b) Power spectrum and SPL directivity for CONF, 12CL, 12UH and 12TH at T.O. # l, M(fJ) = 0.0

NASA/C_2002-210823/VOL 1 117
80

75

7O

65

6O

55

1500' PNL Directivity Comparison

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj) = 0.0

J

--41'-- CONF "_ q_X

"" --B--- 12CL '_% '_'t

,

65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

1500' SPL Spectra at 60 deg

°55 I "

F

7O

65

lO loo lOOO 10000

freq (hz)

1500' SPL Spectra at 120 deg

_" 6O

-.I

Q,,

U) 55

i==.,

10000

5O

--4-- CONF

--II-- 12CL

.....,.,,-, 12UH

12TH

45 ' ' ...... : ...........

10 100 1000

freq (hz)

65

6O

_" 55

.,J

G,.

cn 5O

45

75,

70,

65,

40

10

60,

"O

_" 55,

¢_ 50

45

40

35

10

1500' SPL Spectra at 90 deg

• • .=..=_ . = • .==.•, • ._

100 1000 10000

freq (hz)

1500' SPL Spectra at 150 deg

+

100 1000

freq (hz) t

10000

Figure 5.1(c) Flyover PNL directivity and SPL spectra for CONF, 12CL, 12UH and 12TH

at T.O. # l, M(fj) = 0.0

NASA/C_2002-210823/VOL 1 118
(viii)

mixing of fan and core flows which

emerge from the lobes/tongues with

widths much "smaller" than the nozzle

diameter. This issue will also be

explored later in more depth.

The mid-frequency hump and larger

high-frequency output of 12UH and

12TH are also captured in PWLspectrum of Figure 5. l(b).

(ix) It appears from the PWL-spectrum of

Figure 5.1(b) that the 12UH and 12TH

mixers suppress low frequency sound

far better than a confluent

configuration, but at the expense of

increasing the mid-to-high frequency

spectrum. It is interesting to note that

although both have relatively high lobe

penetration (12UH has 0.68 and 12TH

has 0.56 compared to 12CL's 0.48),

12UH is unscalloped, whereas 12TH

represents the limit of deep scalloping.

The fact that they still appear to have

similar noise signatures may be related

to their widely differing area ratios

which have led to totally different

bypass ratios and thrusts, as seen in

Chapter 4.

(x) As opposed to the 12UH and 12TH

mixers, the acoustic signature of the

12-lobe low penetration mixer with

cutouts (12CL) is quite different. Its

OASPL is lower than the confluent

design at all angles. Although its highfrequency SPL is comparable to the

confluent mixer at all angles, the lowfrequency SPL is much lower than

CONF (although slightly higher than

12UH or 12TH). This is also reflected

in its power spectrum in Figure 5.1(b).

(xi) Thus, it appears that the compound

effect of cutouts and low lobe

penetration has benefited 12CL

tremendously in terms of reducing low

frequency sound without increasing

the high-frequency sound compared to

CONF. This must be a result of how

(xii)

the fan/core flow mixing progressed

inside the nozzle and outside it while

mixing with the ambient. This is an

issue which will be taken up later. It

also hints that in this case at least the

noise benefits of cutouts overcome any

excess noise produced by the dipoletype noise made by the leading edge of

the aft end of the cutouts mentioned in

Chapter 2.

Figure 5.1(b) also shows the directivity

of particular frequencies for all mixers:

160 Hz (corresponding to peak PWL

for configurations CONF and 12CL),

1600 Hz (where peak PWL of 12UH

and approximately of 12TH occurs)

and a relatively high frequency

component at 5000 Hz. For any of

these frequencies, the SPL directivity

for all mixers is similar in shape. The

low frequency SPL monotonically

increases as we approach the jet exit

axis (up to around 150 ° ) and then

decreases as has been observed in

single-stream nozzles. However, the

mid-to-high frequency SPL peak

occurs more upstream, at around 120 °-

125 ° . This is due to the refraction

effect of the sheared jet flow described

earlier which leaves a "relative zone of

silence" near the jet exit axis. On the

other hand, for any given lobed mixer,

the peak moves upstream with

increasing frequency. As in singlestream nozzles, we expect the different

frequency scales to originate at

different axial portions of the jet. The

higher frequencies, associated with

small-scale eddies, originate more

upstream; the lower frequencies,

associated with larger eddies, originate

further downstream in the plume. This

is, indeed, reflected in the coaxial jet

model of Fisher et al ¢17) and Balsa &

Gliebe _s). Hence, returning to the

SPL-directivities of Figure 5.1(b), the

NASA/CR 2002-210823/VOL1 119
relative amplitudes of the SPL's

betweendifferent mixers at a given

frequency tell us how the source

strengths, corresponding to these

frequenciesand located at similar

relativeaxialpositionsin thejet, differ

betweenmixers. However, without

specific source-locationdata,suchas

might be obtained with multimicrophonephasedarrays,we cannot

further pinpoint the source-locations

for a given frequency. In a later

section, however, we develop a

techniquewhichindicatesthelocation

of somehigh-frequencysourcesin an

approximatesense.

(xiii)The communitynoiseimpact of the

spectral and directivity differences

betweenthe variousmixersis clearly

shownby the 1500ft PNL directivity

of Figure 5.1(c). The increasein

frequencyof theSPLpeakobservedin

configurations12UHand12THis seen

to producea significantincreasein the

peakvalueof PNL comparedto either

theconfluentbaselineor 12CL. This

increasein peakPNL is totally the

result of the frequencyweighting of

the NOY scale, since little or no

difference in the amplitude of the

maximumvalue of SPL is observed

betweentheconfigurations.ThePNL

penalty incurred by configurations

12UHand12THpersistsfor all angles

upstreamof the peak. A different

trend is observed for mixer

configuration 12CL. This

configurationdisplaysthelowestpeak

PNL of anyof the configurationsand

this behavioris also observedfor all

anglesupstreamof the peak. This

resultagainties back to the spectral

characteristicsof this mixer which

shows a reduction in the low

frequencieswithout a corresponding

increasein thehigherfrequencies.The

spectralrangebetween1000Hz and

approximately4000 Hz controls the

observedPNL behavior.As alludedto

earlier,frequenciesbelow1000Hz are

not so heavilyweightedin the NOY

calculation. Although frequencies

above4000Hz areheavilyweightedin

the NOY calculation, atmospheric

attenuationof thesefrequenciesis very

high. TheSPLspectraat 1500ft show

a substantialfall off in the highest

frequenciesand this trend becomes

strongerfor anglesawayfrom 90° due

to the larger propagationdistances

involved.Forthisstaticcase(MI = 0),

the propagation and atmospheric

attenuationeffectsdisplaya symmetry

relative to the 90° directivity angle.

Sincethereis no Dopplershifting of

thefrequencybetweenthefront andaft

quadrants due to relative motion

betweenthe sourceandthe observer,

any asymmetryobservedin the PNL

directivity around 90° at static

conditions is solely due to the true

source asymmetry. When aircraft

motion is present, this result changes.

This effect will be examined later.

In conclusion, the 12 lobe mixer with

cutouts and low lobe penetration (12CL)

appears quieter than the other 12-lobed

mixers (12UH, 12TH) or the confluent

mixer because it reduces the low frequency

noise without raising the mid-to-high

frequencies which are heavily noy weighted.

The tongue mixer (12TH), although it

decreases the low frequency noise, is

heavily penalized due to the increase in midto-high frequency noise it creates. The

tongue mixer's low bypass ratio of 4.55

(compared to 4.95 of 12CL or 4.71 of

12UH) under static conditions and

consequent higher effective jet velocity

accentuates its noise too. Thus, the effects

of scalloping/sidewall-cutouts, which

NASA/CR 2002-210823/VOL1 120
produced a noise benefit for 12CL, were

overwhelmed by other effects in the tongue

mixer.

Mixer configuration 12UH produced the

second highest noise, as measured by PNL,

of all the configurations. This configuration

was designed to have an identical hot

stream/cold stream area ratio as compared to

12CL. However, configuration 12UH has a

higher penetration than 12CL and lacks the

sidewall cutouts. It is not possible to

determine the relative importance of these

two differences on the generation of jet

noise from the data presented in this section.

Although the 12-lobe unscalloped mixer

(12UH) proved to be noisier, the lack of

scalloping usually leads to a better thrust

coefficient at cruise. Hence, from the

viewpoint of keeping balance between

cruise thrust efficiency and takeoff noise

reduction, it is expedient to first study

unsealloped mixers but vary the number of

lobes with the hope of "improving" mixing

and noise characteristics. This is done next.

5.2.1.2 Lobe Number Effect

Figures 5.2(a), (b), (c) show, respectively,

the polar OASPL-directivity, PWL-spectra,

flyover PNL-directivity, and the

corresponding SPL-spectra or directivity for

12UH, 16UH and 20UH mixers, and the

baseline confluent (CONF) mixer. Since

12UH and 20UH have all lobe parameters

approximately equal, except for lobe

number, we will pay special attention to the

difference in noise signatures created by

these two mixers. The 16-lobe mixer,

16UH, is unique in that it has a much larger

fan to core area ratio than the other

configurations in the group. It has been

included in this group because of its

similarities in other parameters (lobe

penetration,, non-scalloped, lobe length) of

12UH and 20UH. General trends observed

within this group are:

(i) The 150 ft polar SPL spectra of the 20-

lobe mixer is lower in the mid-to-high

frequency range than the 12-lobe

mixer at all angles (Figure 5.2(a)).

This is also reflected in the power

spectrum (PWL of Figure 5.1(b) and

the flyover SPL's in Figure 5.1(c)).

However, 20UH does no___.2t suppress the

lower frequency SPL's as much as

12UH does; this can also be seen in

PWL and flyover SPL's.

(ii) The decrease in high frequency and

increase in low frequency of 20UH as

compared to 12UH is reflected in the

OASPL directivities as a relative

decrease for angles upstream of peak

PNL which occurs at 120 ° and a

relative increase in the angles

downstream of the peak PNL angle.

(iii) Although the OASPL of 20UH at

shallower angles (closer to the jet axis)

is higher than 12UH, the flyover PNL

(Figure 5.2(c)) for both the mixers at

those angles is the same. This is the

result of the difference in weighting

between the two metrics. Since each

frequency band is equally weighted,

the value of OASPL at a particular

angle is determined by the band with

the highest amplitude. For both 12UH

and 20UH, the lower frequencies are

dominant at the shallow angles. Since

the peak of the SPL spectrum is larger

for 20UH than 12UH, the OASPL of

20UH will also be larger. The

preferential NOY weighting of the

higher frequencies in computing PNL

changes this trend. Although 12UH

displays a larger peak SPL in the

important 1-4 kHz range in the 150 ft

polar than does 20UH, the atmospheric

NASA/C_2002-210823/VOL 1 121
O

104.

102,

100,

98,

96,

94,

92,

9ol

88 ";'

55

OASPL Directivity Comparison

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj) = 0.0

106,

CONF _ I

-=- 12UH ._u I

.... " 16UH __ J=

1 " ; " ; " ; " ; " ; " ; " : " I " " " , " .

65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

85

83

81

79

77

_. 75

73

71

69

6?

65

92,

90.

88,

86,

84,

¢Q

_2,

o.

78,

76,

74

72

10

150' SPL Spectra Comparison at 60 degs

NPR{f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj} = 0.0

--4k-- CONF

10 100 1000 10000

freq (hz)

150' SPL Spectra Comparison at 120 degs

NPR(f) = 1,44, NPR(c) = 1.39, TTR = 2.34, M(f]) = 0.0

CONF

--II-- 12UH

_,_ 16UH

_- 20UH

r

, ....... : ........ : ...... :

100 1000 10000

freq (hz)

150' SPL Spectra Comparison at 90 degs

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj) [] 0.0

8: ,

8

7!

7

n-J 71 _CONF

uJ 7: --B-- 12UH

7 ::: 16UH

6 _ 20UH

6

6 ...... _*1

10 100 1000

freq (h=)

100,

150' SPL Spectra Comparison at 150 degs

NPR(f) = 1.44, NPR(c) [] 1.39, "I'I'R = 2.34, M(f]) = 0.0

95,

90.

_-85

"o

75

70

65

-.I-. 12UH

......i:. 16UH

20U H

10000

10 100 1000 10000

freq (hz)

Figure 5.2 (a) Effect of lobe number on OASPL directivity and SPL spectra at T.O. # 1 and M(fj) = 0.0.

NASA/C_2002-210823/VOL 1 122
Power Spectrum Comparison

NPR(t} = 1.44, NPR(c) = 1.39, TrR = 2.34, M(fj) = 0.0

134,

132,

130, .--@.-- CONF

128, --II-- 12UH

;: 16UH

126, --X---2OUH

_124,

122.

120.

118,

116,

114

10

9O

88

86

B4

82

u)

. . . • ,,,. | ........ • .......

100 1000 10000

freq (hz)

Directivity Comparison at t300 hz

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj) =

0.0

78

76

74

72

70 , [ , E ' I _ I , I , E , I _ J -' 4 * I , ....

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

1O0

95

90

g

85

80

75

g

lO

DirecUvity Comparison at t60 hz

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj)

= 0.0

CONF

--IF-- 12UH

J

7O , I ' I , I , I , I , I , I ' I , I r I '

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (aeg)

Directivity Comparison at 5000 hz

NPR(f) = t.44, NPR(c) = 1.39, "R'R = 2.34, M(fj) =

0.0

82

80

78

76

,op,........ i;;iii .......... .

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

Figure 5.2 (b) Effect of lobe number on power spectrum and SPL directivity at T.O. # 1 and M(0) = 0.0.

NASA/C_2002-210823/VOL 1 123
85-

1500' PNL Directivity Comparison

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj) = 0.0

8o

= 75I;

z 70.

65.

60 , i , i , J ' 1 , I , I ' I . i ..... ; : ' i

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

60_

55.

g 50.

.d

a.

o,) 45,

40,

35

10

70.

65.

_" 60

.=1

co 55

5o

45

10

1500' SPL Spectra at 60 deg

=..=

l l l l I •

10o

---e--- CONF f

12UH

- .;,, 16UH

---7,F--20UH

1000 10000

freq (hz)

1500' SPL Spectra at t20 deg

..... . _ ..._ . ._._ .... 1

100 1000 10000

freq (hz)

.J

(O

65--

°°I

55

1500' SPL Spectra at 90 deg

50

45

--e-- CONF

12UH

16UH

i --N-- 20UH

40 ........ : ........ :

10 100 1000

freq (hz)

10000

75

1500' SPL Spectra at 150 deg

70

65

_.60

55

eL

50

45

40

35

10

-!--12UH I _X_

---_,_.- 20UH _ "_

_. _..... _. ......... "L_. .....

IO0 1000 10000

freq (hz)

Figure 5.2 (c) Effect of lobe number on flyover PNL and SPL spectra at T.O. # l and M(fj) = 0.0.

NASA/C_2002-210823/VOL l 124
(iv)

(v)

attenuation that occurs during the

propagation to the 1500 ft sideline is

sufficient to result in nearly equal PNL

for the two configurations at the

shallow angles.

For all angles less than or equal to the

peak PNL angle of 120 °, the 20 lobed

mixer is definitely quieter than the 12

lobed design. On an overall integral

basis this would make the mixer with

higher numbers of lobes (20UIt)

quieter than one with fewer lobes

(12UH). However, this does not

necessarily make the 20 lobe mixer

quieter than the confluent mixer on an

overall basis and we will have to await

for the EPNL comparisons (with the

nozzle moving) in a later section.

Turning our attention to the 16-lobe

mixer (16UH) with the higher fan-tocore area ratio, we find that it radiates

the least power at all frequencies,

produces the lowest OASPL at all

angles, the lowest peak PNL, and

generally the lowest PNL directivity of

all the mixers in this group. This

would, of course, make it the quietest

mixer in the group for a fixed

operating condition. As a result of the

difference in area ratio, the 16 lobe

mixer has a much larger bypass ratio

(7.64) than 12UH (4.95) or 20UH (4.8)

and consequently much lower effective

jet velocity. This change in bypass

ratio is accompanied by a change in

thrust for a fixed fan and core

operating conditions. Since a flight

application would require maintaining

a fixed thrust, the above comparisons

are not totally appropriate. On the

other hand, it also shows that a

decrease in noise can be made by

increasing the fan-to-core area ratio,

but a change in the cycle condition

(vi)

would be needed to maintain thrust.

It is noteworthy that for a given polar

angle, the frequency associated with

the maximum polar SPL value (Figure

5.2(a)) does not shift systematically

with lobe number for these

unscalloped mixer configurations,

although the use of one-third octave

bands may mask some of the details.

That is to say, the peak frequencies do

not appear to be correlated with the

widths of the mixer lobes which, in

turn, set the scale the diameter of the

axial vortices. However, the

maximum SPL amplitudes do decrease

with an increase in the number of

lobes.

(vii) The angle at which a particular

frequency band peaks is approximately

the same for the confluent

configuration and all the lobed

configurations. For example, the peak

amplitude for 1300 Hz occurs at

approximately 120 ° for all the mixers

discussed in this section. Of course,

the relative strength of a particular

frequency band does vary between the

mixers.

In conclusion, we see that an increase in the

number of lobes produces a decrease in the

overall noise at static conditions, This is

primarily achieved by reducing the mid-tohigh frequency noise in the angular range

upstream of the peak PNL angle while

changing the flyover noise at the shallower

angles only minimally. Note that the low

frequency polar SPL for the higher lobecount mixer is, however, slightly higher than

a low lobe-count mixer. This observation,

combined with a similar observation for the

previous group of mixers, implies that

reducing the mid-to-high frequency spectra

may be more important to overall noise

NASA/C_2002-210823/VOL 1 125
reduction than large reductions in low

frequency noise alone. The paradigm in

mixer-design in the past has been to mix fan

and core flows as fast as possible so that a

nearly uniform flow forms by the nozzle exit

plane. It was reasoned that this would lead

to faster plume decay and reduce low

frequency noise. The current data shows

that this strategy may not be the optimum

because the mid-to-high frequency noise

generated by the faster fan/core mixing may

upset the overall balance. Since the

presence of a moving free stream

surrounding the nozzle is known to suppress

the low frequency portions of the jet noise

spectra for a single round jet, these

observations may be modified when flight

effects are included. This will be examined

in detail in a later section. Prior to this, the

subtle effects of scalloping will be discussed

in the next section.

5.2.1.3 Effect of Scalloping

Figures 5.3(a), (b), (c) show the polar

OASPL directivity, the sound power spectra

(PWL), the flyover PNL-directivity, and the

associated SPL-spectrum or directivity for

the 20UH, 20MH and 20DH mixers. All

results are again presented for stationary

surroundings (Mrj = 0) and for a reference

operating point with a fan pressure ratio of

1.44, core pressure ratio of 1.39 and a coreto-fan total temperature ratio of 2.34. Recall

that the moderately scalloped mixer (20MH)

has half the scalloped area of the 20DH (see

Table 2.2).

(i) The primary difference between the

deeply scalloped mixer (20DH) and

the unscalloped design (20UH) is a

reduction in amplitude of the mid-tohigh frequencies (approximately 1-4

kHz) as shown in the 150 ft polar SPL

spectra of Figure 5.3(a). This

improvement is observed over all

(ii)

(iii)

directivity angles. At the same time,

the substantial low frequency

suppression produced by the

aggressive mixing of the 20 lobe

unscalloped configuration is not

reduced. A small increase in the

highest frequencies (greater than 4 kHz

at 90 °) is observed for 20DH

compared to 20UH, but the increase is

small enough to be of no practical

concern. The moderately scalloped

configuration (20MH) produces

similar trends, but the suppression of

the mid-range frequencies is

significantly less than observed for

20DH. The OASPL directivity, PWL

spectrum, and PNL directivity, as

shown in Figures 5.3 (a), (b) and (c),

for configuration 20MH are all

bounded above and below by the

corresponding results for 20UH and

20DH, respectively. However, the

correlation between noise reduction

and scalloped area, although

monotonic, is not simple.

All this is reflected in the lowest

OASPL directivity as well as the

lowest PNL directivity at all angles for

the deeply scalloped mixer, 20DH,

making it the quietest mixer in the

group.

As in the coaxial jet model of Fisher et

al (17) there appear to be at least two

dominant regions of frequency in

lobed mixers (Figure 5.3(a)). One is

the low-frequency peak governed by

the fully-mixed region far downstream

and the other is the mid-to-high

frequency peak governed by the shearlayer between the ambient and the

partially mixed fan/core flow close to

nozzle exit. The geometry of the lobed

mixer and changes to the mixing

process can be used beneficially to

NASA/CR 2002-210823/VOL1 126
O

OASPL Directivity Comparison

NPR(f) = 1.44, NPR(c) = 1.39, T'rR = 2.34, M(_) = 0.0

106

104

102

100

98

96

94

92

90

88 I '=

55

i '--_k--' CONF

i:._::u..

I ' ; " ; " ; " ; " ; " I ' I ' I " :

65 75 85 95 105 115 125 135 145 155

Angle from Inlet Axis (deg)

165

85.

83 •

81.

79 •

_. 77.

_" 75.

ID 73,

71,

69.

67,

65

10

I

90 °

85.

eel

.d

80,

75,

150' SPL Spectra Comparison at 60 degs

NPR(f) = 1.44, NPR(c) = 1.3g, TTR = 2.34, M(fj) = 0.0

-"0"-- CONF

20UH

20MH

.....<,_-'-"20DH

r "

100 1000 10000

freq (hz)

150' SPL Spectra Comparison at 125 dogs

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fi) = 0.0

CONF

20UH

20MH

• 20DH

Y'.... 2,_.; _:__"12:,,_ •

70 ........ ;

10 100 1000 10000

freq (hz)

85.

83.

81.

79.

77.

75,

o.

u)

73,

71,

d

69,

67,

65

10

150' SPL Spectra Comparison at 90 degs

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fi} = O.0

CONF

20UH

--IB-- 20MH

20DH

100 1000 10000

freq (hz)

100

150' SPL Spectra Comparison at 150 degs

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj) = 0.0

g5

go

85

80

75

7O

10

---4k-- CONF

20UH

-..41-- 20MH

:: 20DH

....: ......... . ',d 100 1000 10000

freq (hz)

Figure 5.3 (a) Effect of scalloping on OASPL directivity and SPL spectra at T.O. # 1, Mrj = 0.0

NASA/C_2002-210823/VOL 1 127
134

1o

5124

o-122

Power Spectrum Comparison

NPR{f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(rj) = 0.0

132 •

130 __i,_ CONF]

20UH )

128 --Ii 20MH

126 ...... 20DH J

120

118

116

114

10

88.

86

84

82

_'80

,,.J

u_ 78,

76

74

, ....... : ........ : ......

100 1000 10000

freq (hz)

Directivity Comparison at 1300 hz

NPR(f) = 1.44, NPR(c) = 1.39, "l-rR = 2.34, M(fj) =

0.0

20D

72

70

55 65 75

' I ' I , I , I r I ' I ' I ' I T I ' I ''

85 95 105 115 125 135 145 155 165

Angle from Inlet Axis {dog)

100 ,

Directivity Comparison at 160 hz

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj)

-- 0.0

--.--20M.

75

70

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

82

Directivity Comparison at 5000 hz

NPR(f) = 1.44, NPR(c) = t.39, TTR = 2.34, M(_)

--- 0.O

.- x,

80

78

[

76

74

72

55 65 75 85 95 105 115 125 135 145 155 t65

Angle from Inlet Axis (deg)

Figure 5.3 (b) Effect of scalloping on power spectrum and SPL directivity at T.O. # l, Mfj = 0.0

NASA/C_2002-210823/VOL 1 128
1500' PNL Directivity Comparison

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj) = 0.0

65

8 75

' 70 --t--20MH ' _'.

i:"i: J / 60 ......................... ' ', ' ', ' ', ,

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

60 _.

55 ,f

50

,.I

D.

,¢_ 45

40

35

10

1500' SPL Spectra at 60 deg

+7 H

, , , .....: ........ : - ._ ....

IOD I000 iO000

freq(hz)

70

1500' SPL Spectra at 120 deg

65

60

.J

50

45

40

35

10 100 1000 100O0

freq (hz)

A

rn

"O

.J

D.

(/}

65,

60,

55,

50.

45.

40

35

10

1500' SPL Spectra at 90 deg

, 20DH _,

|

....... : ........ , - .. _.,,.

100 1000 10000

freq (hz)

75,

70,

65,

A 60,

1:1

_" 55

m 5O

45

4O

35

10

1500' SPL Spectra at 150 deg

....X--- 20UH %-v

20MH

_ _,,_ 20DR

" = =.... 3 = ....... : "_--"

100 1000

freq (hz)

10000

Figure 5.3 (c) Effect of scalloping on flyover PNL directivity and SPL spectra at T.O. # 1, M 0 = 0.0

NASA/C_2002-210823/VOL 1 129
control one or the other of these peaks.

The unscalloped 20-lobe mixer

produces a substantial reduction in the

low frequency noise at the peak PNL

angle compared to the confluent

design, but a large increase in the midfrequencies is observed. Introducing

sidewall scallop improves the low

frequency suppression at tile peak PNL

angle and significantly reduces the

higher frequency penalty.

(iv) It also appears that the SPL-directivity

at 1300 Hz, (Figure 5.3(b)) for the

unscalloped and the moderately

scalloped mixers has a more rapid

decay away from the peak than the

deeply scalloped mixer ("beaming"

characteristics); whereas, the very high

frequency (5000 Hz) SPL directivity

for 20DH shows a slight increase over

20UH or 20MH for all angles

upstream of the peak PNL angle.

In conclusion, deep scalloping helps overall

noise suppression by reducing the mid-tohigh range frequencies without any increase

in the low frequency spectra and only a

small inconsequential increase in the very

high frequencies compared to unscalloped

mixer. However, we also saw in Chapter 4

how deep scalloping can cause higher thrust

losses at both takeoff and cruise conditions.

We also note here that noise suppression

does not appear to follow a linear relation

with the scalloped areas. It is possible that

there is some "'critical" scalloped area after

which one begins to reap noise benefits.

However, it is also possible that for a given

mixer there exists an "optimal" scalloped

area beyond which no additional noise

benefits occur. This second possibility can

not be proven, but is hinted at by the poor

acoustic performance of the tongue mixer in

§ 5.2.1.1, even after accounting for its area

ratio difference as compared to the other

mixers.

Before moving on to the case of non-zero

flee-jet speed, which represents the flight

effect, we digress briefly to the possible

causes of noise reduction observed in some

of these mixers even under static conditions.

The flight-effect will, of course, complicate

the matter further by changing some of the

sources discussed in the next section.

5.2.1.4 Probable Causes of Noise

Reduction

In the previous three sections we have seen

that noise can be reduced by the compound

effect of lobe cutouts and low penetration or

by increasing the number of lobes or by

deep scalloping, at least for static operation

and with the relatively low jet speed (830

ft/s) characteristic of the baseline operating

point. Out of these effects only the second

and third effects are parametrically captured

in our acoustic data base. Hence, we discuss

the scalloping effect first, which in a way

captures some of the "cutout" effects, and

then the lobe number effect.

Scalloping Effect

The probable mechanism for the scalloping

effect, from the fluid-dynamic point of view,

has already been discussed in Chapter 2

(Section 2.3.1) in fair amount of detail and

this understanding is embedded in our

scalloped lobe design philosophy. Briefly, it

was concluded in Section 2.3.1 that

"boomerang" type scalloping, as in 20DH,

will "introduce axial vorticity gradually"

into the flow beginning at an earlier axial

location so that the fan/core flow mixing

proceeds more "gently" than the unscalloped

mixer and the most intense turbulence spots

are acoustically shielded by the nozzle duct.

This may or may not lead to a more uniform

velocity/temperature profile by the time the

two flows reach the nozzle exit plane (that

NASA/C_2002-210823/VOL 1 130
depends on how long the mixing length is),

but it will tend to reduce the mid-to-high

frequency noise component generated by the

internal mixing and by the interaction of the

partially mixed flow with the ambient.

Some of the CFD results for 20DH mixer in

Figures 2.21-2.23 have captured the low

turbulent kinetic energy generated during

this internal mixing with axial vorticity. In a

later section, it is shown that the noise in this

frequency range for 20UH originates

primarily in the region close to the nozzle

exit plane.

It would have been useful to have

comparable plume survey data for 20UH

and 20DH at static baseline conditions but

from Table 3.5 we see that although data

exists for 20DH at Mfj = 0.2 only a partial

data set was collected for 20UH at the three

axial stations, X/D = 0.2, 0.5, 1.0. Although

a non-zero free-jet speed changes the shearlayer properties between the nozzle-jet/freejet, the arguments given here are

independent of it. Comparison of total

temperatures for 20UH and 20DH (see

Figures 4.12 (a) and (b)) in the transverse

plane at X/D = 0.2 shows that deep

scalloping caused the outer hot-spot of

20UH (at Z = + 3 in.) to reduce in intensity

and move outward, but at the same time

another less intense hot-plateau was formed

closer to the axis (at Z = + 1.75 in.).

Scalloping appears thus to encourage

azimuthal mixing of axial vortices and to

reduce the radial gradient of total

temperature in the outer periphery which

itself leads to another source of noise. The

central hot-spot intensity and size has not

changed much and the center-line traverse of

total temperature (see Figure 61 in Vol. III,

pt. 1) looks similar for both unscalloped and

scalloped mixers. On the other hand, the

velocity profile for the scalloped mixer at

X/D = 0.2 (see Figure 4.13) appears fairly

uniform with only a small local maximum at

the center, corresponding to the hot-spot

from the core stream. The corresponding

velocity profile for 20UH is not available,

but it is expected that it will be less uniform

at the periphery due to the larger hot-spots.

This would imply an increased shear in this

region for the unscalloped mixer, making it

an additional source of noise. The centerline velocity traverse out to one diameter

(see Figure 4.18) is similar for scalloped and

unscalloped mixers. Plume data is not

available for X/D greater than 10 for 20DH,

a region which is expected to govern the low

frequency noise behavior. However, from

the polar SPL data (Figure 5.3(a)) which

shows similar low frequency behavior at all

angles for both unscalloped and scalloped

mixers (20UH and 20DH) we can inversely

deduce that the far-field plume decay must

be similar for both mixers. Beyond this we

cannot really pin-point the process that leads

to noise reduction due to scalloping with the

available data.

Packman & Eiler (1_describes the differences

in flow-fields and noise of unscalloped and

scalloped mixers for a very low bypass ratio

nozzle for Pratt & Whitney's JTSD engine.

Although some of their observations and

conclusions are applicable here, some are

not. For example, as in our case, they also

found scalloping to decrease noise in the

mid-to-high frequency range. However,

their scalloped mixer was concluded to have

less nozzle exit uniformity and there was no

central hump. Although the details may

differ with scale and operating conditions

the final outcome relative to noise was

similar: scalloping reduces noise. Another

recent paper by Yu et al (19) attempts to

describe the fluid-dynamic effects of

scalloping but no connection to noise

production is made. All the processes

described in Yu et al (19) may not be

applicable to the mixers used in our study

due to different lobe/nozzle geometry and

NASA/C_2002-210823/VOL 1 131
operating conditions.

In summary, scalloping appears to produce

less intense hot spots by encouraging

azimuthal and radial mixing between axial

vortices which reduces the mid-to-high

frequency noise without changing much of

the far-field plume decay produced by an

unscalloped design. A more detailed plume

flow study is needed to uncover more of the

physics connecting the scalloped mixer flow

to the noise that it generates.

All the arguments that apply to the noise

benefit of scallops can be extended to

cutouts as used on the 12CL mixer. Since it

was shown to be quieter than 12UH it is

likely that the "leading-edge dipole source"

from the lobe-wall strip downstream of the

cutouts on 12CL, mentioned earlier in

Chapter 2, is a less important noise source

than the mixing of the flows downstream.

However, 12CL also had a lower lobe

penetration than 12UH. Penetration plays

some role in the radial migration of the axial

vortices generated by the mixers (as also

seen in the CFD simulations, Figures

2.22(b), 2.25(b)). This change in radial

migration will determine whether the

vortices interact with the outer nozzle-duct

wall and, hence modify the ambient/jet

shear-layer, thus generating more or less

noise. We unfortunately cannot distinguish

the lobe-penetration effect on noise from

this data directly other than noting that for

high-penetration the high-speed core flow,

which is radially thrown outwards,

immediately interacts with the shear-layer

between the jet and the ambient flow close

to the nozzle-exit plane, possibly increasing

the shear, and hence, increasing the mid-tohigh frequency sound. A low-penetration

mixer, such as 12CL, keeps the axial

vortices closer to the jet-axis. This prevents

the high-speed core flow from immediately

interacting with the ambient shear layer,

reducing shear there and, hence, also

reducing the mid-to-high frequency noise

from that region. However, these same lowpenetration axial vortices will continue to

modify the flow further downstream, what

Fisher et al (17) call the "interaction" region

for noise production, and modify the noise

characteristics.

Lobe Number Effect

Now, consider the effect of increasing the

number of lobes (20UH versus 12UH).

This obviously increases the interface area

between the fan and the core flows and

decreases the length-scale (diameter) of the

axial vortices which are now more

numerous. This should enhance mixing

between the two flows. A decrease in the

length-scale of the axial vortices would

seem to imply an increase in the dominant

frequency, but this is true only if their

strengths remained the same. As the number

of lobes increases, the number of vortices

occupying the space within the nozzle must

also increase. This promotes earlier

azimuthal interaction between the vortices.

This can, indeed, reduce their strengths. It

will be shown later that the mid-to-high

frequency sound of mixers 12UH and 20UH

is associated with these axial vortices and

their interaction with the jet/ambient shear

layer. In Figure 5.2(a), the mid-frequency

content of the 20 lobe mixer, 20UH, is less

than that of the comparable 12 lobe mixer,

12UH. This seems to imply that an increase

in the lobe count produces a reduction in the

strength of the axial vortices and the

jet/ambient shear layer due to better

azimuthal mixing of the axial vortices. On

the other hand, the 20 lobe mixer, 20UH, is

less effective in reducing the low frequency

portion of the spectrum which is typically

associated with the far downstream plume

characteristics. It is inferred that the larger

diameter and stronger axial vortices

produced by the 12 lobe mixer, 12UH,

NASA/CR 2002-210823/VOL1 132
probablypromotemorerapiddecayof the

plumethroughviscousdiffusion.

In summary,increasingthenumberof lobes

produces a decrease in mid-to-high

frequency sound. Greater lobe count

increasestheinterfaceareabetweenthe fan

andcorestreamsandpromotesalessintense

initial mixing region aided by more

numeroussmallerand,perhaps,lessintense

axialvortices. However,this typeof initial

mixingdoesnotappearto promoterapidfarfield plumedecay. This resultsin a slight

increasein low-frequencysound over a

mixer with fewer lobes. More definitive

workisneededto understandthereasonsfor

thelatterhalfof thisstory.

5.2.2Flight-Effect: Free-Jet Mach No. 0.2

This section examines the comparative

acoustics results for the various mixers with

the reference 100% nozzle-length, at

baseline operating condition TO #1 when

the free-jet Mach number is set to 0.2. This

is representative of aircraft speeds at takeoff for today's regional and business jet

applications. As before, the data

presentation will be in terms of different

groups of mixers in order to isolate specific

effects. Instead of analyzing all the details,

as in the previous section, the focus in this

section will be on the differences in noise

signatures between mixers caused by the

non-zero free-jet Mach number.

5.2.2.1 12-Lobed Mixers

Figures 5.4(a), (b), (c) show the polar

OASPL directivity, power spectrum and

flyover PNL-directivity with the

corresponding SPL-spectra or SPLdirectivity for 12CL, 12UH, 12TH and

CONF mixers for a free jet Mach No. of 0.2

at TO # 1. Compare these figures with

Figures 5.1(a), (b), (c) respectively where

the free-jet Mach No. was zero.

(i) The major difference observed in the

OASPL with a non zero free-jet speed

is that 12UH and 12TH appear much

louder that either the baseline CONF

or 12CL configurations for angles less

than or equal to 130 ° . For example, at

90 ° the difference in OASPL of 12UH

and CONF is almost 3.5 dB as

opposed to 1.5 dB at static conditions.

(ii) Over this angular range, the polar SPL

spectra of the lobed mixers in this

group are dominated by the mid-tohigh frequency range, as shown clearly

in Figure 5.4(a). Similar trends are

observed when comparing the PWL

spectra of Figure 5.4(b) and 5.1(b).

Using 2000 Hz as an example, the

difference in sound power between

configurations 12UH and CONF is 8.5

dB with a free jet Mach No. of 0.2 as

opposed to 4.5 dB at static conditions.

It is important to note that the presence

of the free jet flow produces a decrease

in the PWL spectra for both the lobed

mixers and the confluent configuration

but the confluent is more affected.

Similar trends hold for the tongue

mixer. Similar local trends are

observed in the SPL spectra for the

various directivity angles.

(iii) With the free jet operating, the

difference in sound power spectra

between the confluent configuration

and the forced mixers 12UH and 12TH

is larger in the mid-to-high frequency

bands than the difference in the low

frequency bands (less than 500 Hz).

Comparing the PWL of 12UH and

CONF at Mfj = 0.2, the difference is

8.0 dB at 2000 Hz versus 6.0 dB at

100 Hz; whereas, in the static case the

corresponding differences are 4.5 dB

NASA/CR 2002-210823/VOL1 133
OASPL Directivity Comparison

96 -- Mp_/_ __.4'.,_., MpD{_C} _- 4 _39..I"T_I_. -- 2.__, .M_,/._,}-- n_

_t _" I-'-_ k

86,t_ I.....12_,1 1_-12" I

55 65 75 85 95 105 t15 125 135 145 155 165

Angle from Inlet Axis (deg)

85,

83,

81,

79,

77,

73

71

69

67

65

10

85,

8O

,J

P

70

65

10

150' SPL Spectra Comparison at 60 degs

...... _ 12o. _ !!.....\

....

, , - ..... : ........ ; .......

100 1000 10000

freq (hz)

150' SPL Spectra Comparison at 125 degs

CONF

-m-- 12CL

..... 12UH

_X_ 12TH

i , " • "=,,: ........ : .......

100 freq (hz) 1000

84 q,

82 4,

80 4,

78 4"

74 4"

72 4"

7o

68I

10

85,

8O

_5

7O

65

6O

t50' SPL Spectra Comparison at 90 degs

100 1000

freq (hz)

150" SPL Spectra Comparison at 150 degs

CONF

--II-- 12CL

..... ,:,- 12UH

12TH

........ I ........ = ......

10000

10000 10

L

1O0 1000 10000

freq (hz)

Figure 5.4(a) OASPL directivity and SPL spectra of CONF, 12CL, 12UH, 12TH mixers

at T.O. # 1, Mt_ = 0.2

NASA/C_2002-210823/VOL 1 134
126

124

122

120

.J

118

11E

11,t

11;

Power Spectrum Comparison

NPR(f) = 1,44, NPR(c) = 1.39, TTR = 2.34, M(fj) =

0.2

--B-- 12CL ,

,.;:, 12UH i

/

12TH

10

, , , ..... : , = , ..... : ......

100 1000 10000

freq (hz}

86-

Directivity Comparison at 1600 hz

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

86-

84

82

80

76

74.

Directivity Comparison at 250 hz

--_-- CONF

-4k- 12CL / _ I

....12o. /

1

721

70

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

85.

Directivity Comparison at 5000 hz

--x-, 12[H

60 ' l ' I ' ]' "' I ' [ ' '"1 ' I " I ' I ' I '

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Icier Axis (deg)

Figure 5.4(b) Power spectrum and SPL directivity of CONF, 12CL, 12UH, 12TH mixers

at T.O. # 1, M_ = 0.2

NASA/C_2002-210823/VOL 1 135
1500' PNL Directivity Comparison

NPR(f) = 1.44, NPR(c) = 1 •39, TTR = 2.34, M(fj) = 0.2

82

7478 - ...... .. :,-,_" ._ _ ' _:_" ,,,

z ...."_

66

62

58 I ' ' I _.--

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

60,

55,

50

,.J

a.

45

40

35

10

63

61,

59,

57,

53

51

49

47

45

10

1500' SPL Spectra at 60 deg

lOOO 10000

freq (hz)

1500' SPL Spectra at 120 deg

CONF

--I-- 12CL I

12UH I

L-*- 12TH

I " " ='''=_ "=="

100 10000

= . • . ==•=1

1000

freq (hz)

3"

D.

t/J

6O

58

56

54

52

5O

48

46

44

42

40

10

1500' SPL Spectra at 90 deg

__

........ :

100 1000 10000

freq (hz)

60--

5O

4o

35

30

25

10

1500' SPL Spectra at 150 deg

:+

100 1000 10000

freq (hz)

Figure 5.4(c) Flyover PNL directivity and SPL spectra ofCONF, 12CL, 12UH, 12TH mixers

at T.O. # 1,Mfj =0.2.

NASA/C_2002-210823/VOL 1 136
(iv)

(v)

(vi)

(vii)

at 2000 Hz versus 7.5 dB at 100 Hz.

From the similarity of dB-difference at

low frequencies, whether the free jet

flow is on or off (compare SPL's in

Figure 5.4(a) and 5.1(a)), it is

concluded that the low frequency

spectra of both the forced mixers and

the confluent mixer are generated in

similar locations within the jet plume

and respond similarly to an external

flow. This is not true for the higher

frequency portion of the spectrum.

Under static conditions, the 150 fl

polar SPL spectra for configuration

12UH and 12TH show very similar

trends. This similarity is not altered

when a free jet flow is present. This

implies a similarity in source

distribution between the two

configuration which will be examined

in more detail in a subsequent section.

At 125 °, where lobed mixers have a

local OASPL peak, the relative

magnitudes of the peak SPL's between

12UH (or 12TH) and CONF (or 12CL)

change drastically with external flows.

Under static conditions, CONF has a

larger peak SPL than 12UH peak SPL

at this angle. With a free jet flow

present, the magnitudes are reversed.

The reduction in peak SPL of CONF

or 12CL due to a free jet flow is

typically larger than that for 12UH or

12TH at all angles upstream of 125 °.

That is, the free-jet appears to benefit

CONF and 12CL more than it does

12UH or 12TH.

The noise benefits of the free-jet Mach

No. for configurations CONF or 12CL

over 12UH or 12TH are most clearly

captured in the flyover PNL (Figure

5.(c). Note that both 12CL and CONF

are now quieter than 12UH or 12TH at

all angles. In particular, the decrease

in noise at shallow angles for CONF or

12CL is significantly greater than that

for 12UH or 12TH. The reason for

this will be discussed later.

(viii) As a result of the spectral changes

produced by the flee jet flow, the

confluent mixer, CONF, produces a

lower PNL than the 12 lobe

configuration with cutouts, 12CL, for

angles less than or equal to the peak

PNL angle of 120 ° or so. A very

different result was observed under

static conditions. In that situation,

configuration 12CL produced a lower

PNL for all angles than CONF.

In summary, forward flight-effect or nonzero free-jet speed benefits the confluent

mixer and 12CL much more than it does

either 12UH or 12TH. As a result, CONF

and 12CL are now much quieter at all angles

than 12UH or 12TH. Mixer CONF now

even competes with 12CL for overall

quietness but the EPNL discussion will be

postponed to a later section. This appears

mainly due to larger relative reductions in

the mid-to-high range frequency sound for

CONF and 12CL than those for 12UH and

12TH; the low frequencies are affected

similarly for all the mixers.

5.2.2.2 Lobe Number Effect

Figures 5.5(a), (b), (c) show the polar

OASPL, PWL, flyover PNL and

corresponding SPL's for 12UH, 16UH and

20UH for a free jet Mach number of 0.2 at

TO # 1. Primary attention will be paid to

the difference between the results for 12UH

and 20UH, recalling that 16UH has a much

higher fan to core area ratio. The

comparable plots for free jet flow are in

Figures 5.2 (a), (b) and (c).

NASA/CR 2002-210823/VOL1 137
OASPL Directivity Comparison

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj) = 0.2

96,

94,

g 92,

90, ...._'_--x---x

0 881

-..e--- CONF

84 ............. --×--20UHJ. I " " "

55 65 75 85 95 105 t15 125 135 145 t55 165

Angle from Inlet Axis (deg)

83i

8t.

79.

77°

75°

¢/1

73,

71,,

69, ,

67,,

65

10

86:

04.

82,

80,

_78,

a.

¢/J

76,

74,

150' SPL Spectra Comparison at 60 degs

CONF

-.B-- 12UH

16UH

-'-X-- 20UH

72,

70 . . .m .... :

10 100

85

83,1

81

79!

I

77.

_'75 •

¢/)

73.

71.

69.

67°

i

150' SPL Spectra Comparison at 90 degs

CONF

12UH

' 16UH

20UH

65 ........ : ........ ; .......

10 100 1000 10000

freq (hz)

100 1000 10000

freq (hz)

150' $PL Spectra Comparison at 125 degs 150" SPL Spectra Comparison at 150 degs

85.

80.

--4-- CONF l _'_

--.B-- 12UH

.....'._.. 16UH,

20UH

/

2,

1000 10000

freq (h=)

•---'e.-- CONF

"-tl-- 12UH

•.- _:., 16UH

'-_<-'-" 20UH

_75.

..J

u)

70,

65,

60 ........ : ........ : ......

10 100 1000 10000

freq (hz)

Figure 5.5(a) Effect of lobe number on OASPL directivity and SPL spectra at T.O. # l, Mfj = 0.2

NASA/C_2002-210823/VOL 1 138
126,

124,

122,

_, 12o,

...J

_110,

116. ,

114..

112

10

Power Spectrum Comparison

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(t]) = 0.2

CONF

--B-- 12UH

......:'_-- 16UH

20UH

, , , , .... _ = ....... : ......

100 1000 10000

freq (hz)

86

DirectivityComparisonat 1300hz

;i ......... ,,, c::: ...... ,,\

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

86

"I 82

m 80

;'°l

76 t

DirectivityComparisonat250hz

c°2

v2

55 66 75 85 95 105 115 125 135 145 t55 165

Angle from Inlet AxIs (deg)

8O

78

76

74

72

7o

in

68

60

64

62

6O

DirectivityComparisonat5000hz

I

I 1 J I --- --" 'If

55 65 75 85 95 105 t15 125 135 145 155 165

Angle from Inlet Axis [deg)

Figure 5.5(b) Effect of lobe number on power spectra and SPL directivity

at T.O. # 1, MO = 0.2

NASA/C_2002-210823/VOL l 139
1500' PNL Directivity Comparison

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(fj) = 0.2

80

75

70

65

60

55! , I , t.... _ , I ' ; ' ',...... ', ............. ', ' ; ' _ •i_=

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

60,

55

_- 5O

..J

D.

¢n45

4o

1500' SPL Spectra at 60 deg

35 i • ="

10 10000

| i -= • ....

100 1000

freq (hz)

1500' SPL Spectra at 120 deg

60

55

_50 ÷ ;_:

45 --II-- 12U

,of ........ I

35

10 100 1000 10000

freq (hz)

60.

58,

56,

54,

_- 52,

"0

_" 50,

48,

46,

44,

42

40

10

65

60

55

5O

e4o

35

30

25

20

10

1500' SPL Spectra at 90 deg

CONI_

--II-- 12U

•. ,1o . I

100 1000 10000

freq (hz)

1500' SPL Spectra at 150 deg

• ._e__ c_

..,,o.

-_ 20UH 1

100 1000 10000

freq (hz)

Figure 5.5(b) Effect of lobe number on flyover PNL directivity and SPL spectra at T.O. # 1, Mt_ = 0.2

NASA/C_2002-210823/VOL 1 140
(i)

(ii)

(iii)

For a free-jet Mach No. of 0.2, the

OASPL of 20UH is lower than that of

12UH at all angles rather than just

upstream of the local peak OASPL

angle of 125 ° as was tree for the static

case. The spectral results of Figures

5.2(a) and 5.5(a) provide the

explanation for this change in

behavior. For the static case in Figure

5.2(a), the 12 lobe configuration,

12UH, was more effective in

suppressing low frequencies for angles

near the jet axis than was the 20 lobe

design, 20UH. In the presence of an

external flow, both designs produce

essentially equal low frequency sound

pressure levels; at higher frequencies,

the free jet flow affects both

configurations equally for all angles,

with the overall result that the 20 lobe

configuration remains quieter than the

12 lobe configuration.

The improved low frequency

suppression of mixer 20UH in the

presence of an external flow is also

observed in the PWL spectra of Figure

5.5(b). It is now seen that the PWL of

20UH and 12UH are very similar at

low frequencies as opposed to the

static case (Figure 5.2(b)) where 20UH

produced more power output.

From the flyover noise point of view,

Figure 5.5(c) shows that now 20UH is

quieter than 12UH at all angles rather

than just upstream of the peak PNL

angle as occurred in the static case

(Figure 5.2(c)). The Doppler-shifted

SPL in Figure 5.5(c) also show how, at

even shallow angles, 20UH now

matches 12UIt at low frequencies and

is, of course, always quieter at mid-tohigh frequencies as in the static case.

(iv) However, with a free jet flow, both

12UH and 20UH produce higher PNL

than the confluent baseline for angles

less than or equal to 135 °. As a result,

the confluent configuration will be

quieter than any of the unscalloped

mixers in flight as will be seen in the

EPNL values later.

In summary, when flight effects are

included, increasing the lobe count makes

the mixer-nozzle quieter at all angles,

especially due to larger relative decreases in

low frequencies as compared to smaller

relative decrease in mid-to-high frequency

spectra.

5.2.2.3 Effect of Scalloping

Figures 5.6 (a), (b), (c) show the polar

OASPL, PWL, flyover PNL with the

corresponding SPL-spectra or directivities

for 20UH, 20MH and 20DH mixers for a

free jet Mach number of 0.2 and at the

reference operating condition TO #1.

Comparable no free jet flow plots are shown

in Figures 5.3 (a), (b) and (c).

(i) The presence of a free jet flow

produces only a minor relative change

in the OASPL directivity for the

scalloped mixers compared to the

static results in Figure 5.3(a). Both the

polar SPL spectra (Figure 5.6(a)) and

sound power spectra (Figure 5.6 (b)) of

the scalloped mixers continue to show

that free jet flow primarily affects the

lowest frequencies. Little or no effect

on the differences between mixer

configurations is introduced by a free

jet flow at higher frequencies as

illustrated by comparison of Figure

5.3(a) and Figure 5,6(a) at an angle of

125 ° and a frequency of 2000 Hz for

mixers 20UH and 20DH.

(ii) From a flyover noise viewpoint (see

NASA/C_2002-210823/VOL 1 141
OASPL Directivity Comparison

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(t)) = 0.2

96

94

92

v

90

0

88

86

84 'I,_

55

--.-co,F7

20UH _ _Alr_ -

--B--20MH I

• ' ' : ' I ' : ' I ; ; : ; " : "

65 75 85 95 105 115 125 135 145 155

Angle from Inlet Axis (deg)

165

81 l

79.

77.

75.

73.

a.

o)

71.

69,

67,,

05

10

84 .p

82 4.

i.

80 4.

I

I.

_'78 "1"

LL

P

74 41,

!

70 4_

10

150' SPL Spectra Comparison at 60 degs

CONF

20UH

--B-- 20MH

......! 20DH

100 1000 10000

freq (hz)

150' SPL Spectra Comparison at 125 degs

---0-- CONF

20UH

--11--20MH

20DH

t

t00 1000 10000

150' SPL Spectra Comparison at 90 degs

01 -r

79 4.

77 4-

75 4-

_ 73 ,_,,

o.

I,o

71 4-

i

69 -Ik

CONF i

--_:-- 20UH

--B-- 20MH

......_. • 20DH

°7t

65 , ........ •

10 100 1000

freq (hz)

150' SPL Spectra Comparison at 150 degs

88 .p

10000

freq (hz)

10 100 1000 10000

freq (hz)

Figure 5.6(a) Effect of scalloping on OASPL directivity and SPL spectra at T.O. # 1, M_ = 0.2

NASA/C_2002-210823/VOL 1 142
126.

124,

122 -

, _'120 -

_118 •

116 •

114,

112

10

Power Spectrum Comparison

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, M(t]) = 0.2

"--_I--'CONF

--'_._--20UH

--I--20MH

+-_ -20DH

100 1000 10000

freq (hz)

84

Directivity Comparison at 1600 hz

82

80

+1 /

76 '' i ._, _:,+°: L.

74 /'

68 ........... +' ' ' ...............................................

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

86

"182

8O I

76i

74 i

72 I

Directivity Comparison at 250 hz

+CONF

+;

70_;,L+I+I_I,I,L_L,L'I'I'

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inl_ AxIs (dog)

78

Directivity Comparison at 5000 hz

72

.J

¢n 7O

76 :+:+ ......._:_'

74

+?' :+

66

66

64

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet AxIs (deg)

Figure 5.6(b) Effect of scalloping on power spectra and SPL directivity at T.O. # 1, M 0 = 0.2

NASA/C_2002-210823/VOL 1 143
1500' PNL Directivity Comparison

NPR(f) = 1.44, NPR(c) = 1.39, TTR = 2.34, i(fj) = 0.2

76

72

64

60

56

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

55

53

51

4g

_" 47

45

43

41

39

37

35

lO

61,

59,

57,

55

53

_" 51

49

47

¢/1

45

43

41

39

37

35

10

1500' SPL Spectra at 60 deg

20UH

1000

freq (hz)

1500' SPL Spectra at 120 deg

--e--- CONF

20

• . _..-__. ........... _....

100 1000 10000

freq (hz)

1500' SPL Spectra at 90 deg

55

53

51

--'4,.!: ---,-0o_-4_-1 -'tff

°,,l ]

37

35 _ml

10 100 1000 10000

freq (hz)

1500' SPL Spectra at 150 deg

61.

57.

53,

49.

45,

,,=1

41,

37,

33,

29.

25 ....

10 10000

20°. _,__

--II-- 20MH

' _,-_20DH I .T. ........

100 1000

freq (hz)

Figure 5.6(c) Effect of scalloping on flyover PNL directivity and SPL spectra at T.O. # 1, M 0 = 0.2

NASA/C_2002-210823/VOL l 144
PNL directivity Figure 5.6(c)) the

scalloped mixer remains quieter than

the unscalloped mixer at all angles in

the presence of a free jet flow. At

shallow angles near the jet exit axis,

the difference between scalloped and

unscalloped designs increases slightly

when a free-jet flow is present.

(iii) In terms of peak PNL, the confluent

mixer is superior to even the quietest

scalloped, lobed mixer (20DH) with a

free jet flow present for the low power

operating condition under

consideration. This superiority is also

observed for angles less than the angle

for peak PNL. This difference in PNL

continues to be driven by the higher

levels observed in the mid-to-high

frequency sound for the lobed mixers.

The penalty incurred in this frequency

range by lobed mixers is reduced with

scalloping but not removed whether a

free jet flow is present or not.

In summary, the beneficial effects of

scalloping are not influenced by the

presence of an external flow-field. The

scalloped designs tested in this program

were equally effective under static

conditions and with a free jet flow present.

5.2.2.4 Confluent Mixer Benefits

In the above section, it was found that the

confluent configuration produced lower

PNL values than any of the scalloped

designs for angles less than or equal to the

peak PNL angle when an external flow is

present. Comparison of the spectral

characteristics of the confluent mixer with

the lobed designs showed that the lobed

designs produced excess energy in the 1-4

KHz frequency band. This spectral shift for

the lobed mixers has been related to the

presence of the turbulence generated by the

strong axial vortices induced by these

devices. The confluent design produces no

axial vorticity, relying solely on the

formation of Kelvin-Helmholtz vortex sheet

instability and consequent turbulent

transport between the primary and the

secondary flow streams to accomplish

mixing of the two streams. This type of

mixing does not produce intense turbulence

in the length scales associated with acoustic

radiation in the mid-frequency ranges. This

type of viscous mixing is also much less

effective than that induced by the strong

axial vortices of forced lobe mixers and the

plume does not decay as fast as a fully

mixed jet. As a result, the confluent mixer

becomes a strong source of low frequency

noise under static conditions. The presence

of the free jet produces a reduction in shear

downstream of the primary potential core, as

will be analyzed later. Since this area is the

primary source of low frequency noise in a

jet, the reduced shear will result in reduced

low frequency noise. This is particularly

important for the confluent mixer, since its

spectra is dominated by low frequencies.

The external flow has little or no impact on

the mid-spectral bands, which dominate the

radiated noise from the lobed mixer

configurations.

Directivity differences between the lobed

and confluent mixers also affect integrated

metrics like EPNL. The confluent mixer

produces it maximum polar levels at shallow

angles near the jet axis. During flyover type

calculations, increases in slant propagation

distance and source convective deamplification associated with shallow angle

radiation further benefit the confluent

design. Lobed designs tend to produce

maximum polar levels closer to 90 ° and as a

result do not see the same propagation

distance benefits. Based on comparison at

the fixed operating condition TO #1, only

mixer configuration 12CL provides a PNL

NASA/CR 2002-210823/VOL1 145
benefit over a confluent mixer when flight

effects are included. Consideration of

detailed EPNL trends will be deferred to a

later section. However it is possible to draw

some conclusion at this point related to

flyover noise suppression. It has been

shown that the noise levels in the low

frequency portions of the jet mixing

spectrum are significantly reduced as the

flight speed increases. Lobed mixers

produce a dramatic decrease in the low

frequency spectrum at static conditions but

this is accompanied by an increase in higher

frequency bands. Forward flight does not

significantly influence this high frequency

hump. Removing portion of the mixer

sidewall through the use of cutouts or

scalloping, increasing the number of lobed

mixers and possibly reducing the lobe radial

penetration reduces the increase in the

higher frequency band sound produced by

lobed mixers, but not to the levels observed

on a confluent configuration. As a result,

with flight effect, lobe mixers produce only

a minimal reduction in jet noise when jet

velocities are low.

An important general lesson learned

regarding flyover noise suppression at low

jet speeds is the following: any mechanism

that produces less of the "annoying" mid-tohigh frequency noise is usually preferable

even if it does not significantly reduce the

less annoying low frequency noise; flight

effect helps reduce the low frequency noise

anyway.

In the next section, results at additional free

jet Mach numbers will be presented. This

data provides insight into the axial location

of the sources producing the mid-to-high

frequency range noise associated with lobed

mixers, prompted by the observations in this

section that free jet flow affects some

frequencies less than others.

5.2.3 Flight-Effect: Other Free Jet Mach

Numbers

Data was acquired over a range of free-jet

Mach numbers, varying from 0.0 to 0.3.

The effect of varying this parameter on the

PWL spectrum, polar OASPL directivity,

and flyover PNL directivity is presented and

discussed for each mixer in this section.

Figures 5.7 to 5.14 show these acoustic

metrics for, respectively, CONF, 12CL,

12UH, 12TH, 16TH, 20UH, 20MH, and

20DH mixers. Note that the vertical-axis

scale for a given metric may vary between

mixers. This should be borne in mind when

comparing changes in a given metric from

one mixer to the other.

Increasing the free-jet Mach number

produces the following effects for ever2

mixer:

(i) All acoustic metrics (PWL, OASPL

and PNL) decrease.

(ii) The reduction in low frequency PWL

is much greater than that at mid or

high frequency.

(iii) The decrease in aft quadrant OASPL is

much greater than that in the front

quadrant.

(iv) Similarly the decrease in the aft

quadrant flyover PNL is much larger

than that in the front quadrant.

(v) The angle at which the peak PNL

occurs moves upstream with

increasing free-jet speed.

All these effects were noted in the previous

section when studying Mfj = 0.2 and are now

seen to be extended also to MO = 0.3. But

the additional decreases due to a change of

M_ from 0.2 to 0.3 are typically not as

drastic as those from M 0 = 0.0 to Mfj = 0.2.

The fact that noise, in general, decreases

with increases in free-jet Mach numbers is

NASA/CR 2002-210823/VOL1 146
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NASA/C_2002-210823/VOL 1 154
well known experimentally for single-stream

nozzles and the same reason must apply to at

least fully-mixed nozzles with uniform exit

flow. A new and simple mathematical

explanation is given below using

Abramovich' s theory of turbulent jets (2°_ .

An increase in flee-jet speed reduces the

shear-layer thickness, 6, and increases the

potential core length. The highest

turbulence intensity and the dominant noise

source is known to be just downstream of

this potential core where the shear-layers

surrounding it interact most vigorously. The

magnitude of the radial gradient of axial

velocity, [bU/Orl, governs the dominant

noise source intensity and the jet diameter

there governs the source volume and, hence,

its net contribution to noise. A first-order

estimate of these quantities can be obtained

by using

O.U_U= U free-jet-U jet

dr

The shear layer thickness at the end of the

potential core, _5c,can be estimated by using,

for simplicity, Abramovich's expression for

incompressible, axisyn-nr_etric cold turbulent

jets (which assumes self-similarity in the

velocity in the initial portion of the

jet/ambient mixing layer), namely,

6 c = R/_a+brn

where R is nozzle-exit radius,

m = Ufree_je t /Uje t and a, b are universal

constants with values of 0.214 and 0.144

respectively.

The above two formulae for _U/Or and 6c

immediately show that the ratio of 10U/0rl

at the end of the potential core with a freejet present to that with no free-jet is

r_ = I1- m[_/_ +bm'a

This ratio can be easily shown to decrease

with increase in m for 0 < m < 1. For

observed self-similar velocity profiles (2°_ it

can be shown that, r/, as given above, is also

the ratio of the maximum axial velocity

gradients in the flow at this location. Thus,

when the free-jet speed increases, the

potential core length increases and the shear

at its end decreases, leading to a decrease in

turbulence intensity. In addition the volume

of the noise sources at this location,

characterized by the radius _c, is also

reduced. This decrease in turbulence

intensity, as well as, the volume of turbulent

source, leads to a decrease in the peak farfield noise. For hot jets, with jet-to-ambient

density ratio other than one, the above

argument can be extended by using more

complicated expressions for the shear-layer

growth given in Abramovich (z°_. With some

modifications this argument can also be

extended to coaxial jets where the

ambient/jet shear layer increases the annular

fan potential core length with an increase in

free-jet speed but not the central primary

potential core length.

The above explanation is applicable only for

the "peak" noise producing regions outside

the nozzle for fully-mixed or single-stream

nozzles. It cannot explain, for example, the

observed differences in sensitivity of high

and low frequency noise to free-jet speed

changes, since these originate in different

portions of the jet. Abramovich's theory (2°)

can be applied similarly to the upstream end

of the jet (jet/ambient shear layer or the

"initial' region of Fisher et al (17)) to study

the high-frequency noise, as well as, to the

far downstream self-similar region of the jet

("mixing" region of Fisher et al (17)) to study

low-frequencies. However, we do not

embark on that analysis here.

Returning back to Figures 5.7 to 5.14 we

note that, as discussed before, the smaller

NASA/CR 2002-210823/VOL1 155
changes in higher frequencies in the PWL

spectra with increasing M 0 are related to

smaller changes in OASPL and PNL in the

upstream angles where these frequencies

dominate. Similarly, the larger changes in

the low frequency PWL spectra with

increasing M 0 are related to larger changes

in the downstream angles for OASPL, PNL

where low frequencies dominate.

The flyover PNL directivity is influenced by

several additional effects due to aircraft

motion, which are not present in the polar

OASPL direcfivity and which are worth

emphasizing again. First is the source

convective amplification effect which

amplifies sound in the upstream quadrant

and decreases it in the downstream quadrant

solely due to the motion of the source with

respect to the observer. This tends to rotate

the PNL-0 curves clockwise around 90 °.

Secondly, in a horizontal flyover above the

observer, the slant distances from the source

to the observer and the Doppler shift in

frequency change with angle. The

directionality of the Doppler frequency

effect produces an increase in the observed

frequency spectrum, as the source

approaches the observer and a

corresponding decrease in the observed

spectrum as the source moves away from the

observer. As a result, the atmospheric

absorption applied to a particular frequency

band in the front and rear quadrants will

relate to different portions of the original

source spectrum. This creates further

asymmetry in PNL directivity around 90 °

than the true source asymmetry captured in

the polar OASPL directivity in the reference

frame of the nozzle. These effects combined

with the overall decrease in peak noise

source amplitude, discussed earlier, leads to

the observed downward shift in PNLdirectivity with clockwise rotation around

90 ° when the free-jet speed is increased.

As noted before, the effects discussed above

produce a steep decrease in flyover PNL at

shallow angles. Since EPNL is a temporal

integration of PNL (corrected for tones), but

only extends over the interval when the PNL

is within 10 dB of the peak value, reductions

in noise at shallow angles may have little

effect on EPNL since the levels at these

angles are near or below the 10 dB down

points. For example, note that at Mfj = 0.2

this happens at 160 ° for CONF, 155 ° for

12CL, and about 145 ° for the rest of the

mixers. The spectra associated with these

angles are dominated by low frequencies

whose strengths are in turn governed by the

rate of plume decay. Thus, devices which

promote only rapid plume decay near the jet

axis for the purpose of reducing noise levels

will not significantly reduce EPNL unless

they also reduce the mid-to-high frequency

noise at other angles.

A far more effective strategy is to focus on

reducing the PNL for angles less than or

equal to the peak. There are several reasons

for this. First, the rate of decrease of PNL is

lower for angles upstream of the peak than

downstream. As a result, the majority of the

"duration" correction for EPNL is associated

with angles upstream of the peak PNL angle

for the forced mixers. In addition the spectra

at the upstream angles peak at frequencies

with a higher annoyance weighting.

Perhaps the most striking feature of Figures

5.7 to 5.14 is the difference in sensitivity of

the PWL, OASPL, or PNL to free-stream

Mach number changes when we compare

one mixer against the other. For example,

all the lobed mixers show smaller changes in

high frequency PWL or upstream OASPL or

PNL than the confluent mixer when free

stream Mach number is changed. However,

mixer 12CL shows remarkably similar

changes to CONF as free stream Mach

number changes, except for the OASPL

NASA/CR 2002-210823/VOL1 156
signature at shallow angles. The noise

signatures of all the "unscalloped" mixers

(12UH, 16UH and 20UH) are also

remarkably similar to each other and

markedly different than CONF. The deeply

scalloped mixer, 20DH, has distinctly

different noise signatures than the rest of the

lot. (Note that the high-frequencies

recorded by the 55 ° microphone for the

20DH mixer test were in error and are not

shown in Figure 5.14.) Its PWL appears

closest to that of 12CL but not its OASPL or

PNL. The tongue mixer, 12TH, also has its

own distinct noise signatures but is, perhaps,

most similar to 12UH, as noted before.

Before leaving this section, we make two

final notes.

(i) If mid-to-high frequency sound is

affected less by free-jet Mach number

changes then the sources that produce

them may be hidden or shielded away

from the effects of the free-jet/nozzlejet shear layer.

(ii) Since changes in free-jet speed

produce changes in the direction of the

ray coming out from the jet (because

acoustic group velocity (ray) in the

free-jet is the vector sum of phase

velocity and free-jet velocity), in order

to study the change in the noise source

due to free-jet effects one needs to

study polar SPL's not at the same

angles in the far-field inside the freejet, but at angles obtained after

removing the additional refraction

changes in the transmitted rays.

The effect of a free jet flow on the

corresponding polar SPL's will be examined

in depth in a later section.

5.3 Mixers with Reference NozzleLength at Higher Nozzle Pressure Ratio

Conditions

It is expedient to examine how the noise

signatures of different mixers compare at the

two higher growth TO conditions with

higher thrusts or jet speeds.

We first examine the acoustic results at M 0

= 0.2 for the two higher pressure ratio TO

conditions for the various groups of mixers.

Although corresponding static results are

available, it is the case with a free jet flow

present (simulating aircraft motion), that is

the more realistic one from a flyover noise

point of view. Secondly, since we found in

the last section how important noise source

information can be obtained by studying the

effect of various free-jet Maeh numbers, we

will examine that effect for the highest

pressure ratio condition (TO #3) in detail.

5.3.1 TO #2 and #3 with Free-Jet Mach

Number 0.2

The overall spectral information is captured

in the polar PWL-spectrum, whereas, the

flyover noise is best captured in the flyover

PNL directivity. For the sake of brevity, we

will present the PWL-spectra and flyover

PNL-directivity including the polar SPLdirectivity of the most energetic component

and flyover SPL-spectra at only the peak

PNL angle. These four plots, which

summarize the most important noise

signatures at any give free-jet Mach number,

will be presented in this section for each

group of mixers.

5.3.1.1 12-Lobed Mixers

Figures 5.15 and 5.16 show PWL, PNL and

SPL for the 12-lobed mixers and the

confluent baseline (Group 1) at TO #2 and

TO #3 respectively. Comparison of PWL

NASA/CR 2002-210823/VOL1 157
qD

,J

150' Power Spectrum Comparison

13

13

150' SPL Directivity at 2000 hz

92

.J

(/}

12

12

12

11

11

"-B-- 12CL

12UH - "_

12TH

a ....... : ........ : ........

10 100 1000 1O0OO

freq (hz)

1500' Flyover PNL Comparison

9o

85

8o

65 ' [ ' I ' I ' I

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

88

84

8C

7E

--_--12TH I , . ,

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

1500' Flyover SPL Spectrum at 115 degs

70,

u_ 45

40

10 100 1000 1000C

freq {hz)

Figure 5.15 Power Spectrum and Flyover PNL Comparison of Mixers at T.O.#2 and M(fj) = 0.2

(12 Lobed Mixers)

NASA/C_2002-210823/VOL 1 158
150' Power Spectrum Comparison

140

136

132

12.1

120

116

10

--4P--CONF

--B--12CL

100 1000 10000

freq (hz)

1500' Flyover PNL Comparison

9!

9,

lit

8:

150' SPL Directivity at 2000 hz

_12UH

--)(-- 12TH

I ' I _ I ' I ' I ' I ' I ' ] _ I ' J '

65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

CONF

....... :iiii.......... ..

100 1000 10000

95

9o

=a95

70 ' I ' I ' I ' I " ;

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

7,

55

1500' Flyover SPL Spectrum at 115 degs

75_

freq (hz)

70,

65.

60,

55,

_so,

,15,

40,

35

30

25

10

Figure 5.16 Power Spectrum and Flyover PNL Comparison of Mixers at T.O.#3 and M(fj) = 0.2

(12 Lobed Mixers)

NASA/C_2002-210823/VOL 1 159
shows how the CONF low frequency peak

rises above the high frequency peaks of

12UH and 12TH as nozzle pressure

increases. The lack of mixing in CONI c

starts hurting it at these high pressure or

thrust conditions. Although the relative

ranking of the mixers remains the same for a

given frequency, a secondary peak near

2000 Hz emerges for 12CL at TO #3 which

was not present in TO #1 or TO #2. The

decibel increase in this secondary peak

between TO #2 and TO #3 is considerably

larger than the increase in the low frequency

peak. The comparison of SPL directivity at

2000 Hz at TO #3 between different mixers

in Figures 5.16 clearly shows how 12CL

approaches the other two mixers near peak

PNL angle of 115 °.

This implies that at very high nozzle

pressure ratios the axial vortices and their

interaction with the ambient, which was

seen to produce the mid-to-high frequency

peaks, become even more important. This is

especially true for 12CL where its low

frequency peak is already higher than either

the 12UH or 12TH mixers. The inability of

the CONF mixer to reduce low frequencies

due to inferior mixing is accentuated at

higher pressure, as seen at the shallow

angles of the PNL-directivity (compare

Figures 5.15 and 5.16 with Figure 5.4(c)).

At TO #3, the peak PNL is shifted

downstream for CONF to 140 ° from 125 ° at

TO #1. In addition, the PNL associated with

angles downstream of the peak are still

within 10 PNdB of the peak and, hence, still

important as opposed to the TO #1 case.

The mid-frequency spectral peak of 12CL

produces a prominent PNL peak at 115 °

directivity which was not present at TO #1.

This is also, of course, captured in the

flyover SPL spectra at 115 °.

In summary, 12UH and 12TH are still

noisier than 12CL or CONF at higher

NPR's, especially at angles upstream of

130 ° . The inferior mixing of the confluent

configuration starts hurting it significantly at

shallower angles although it is still quieter at

angles upstream of 125°., A previously

unseen strong mid-to-high frequency peak

emerges in the spectrum of 12CL at an angle

of 115 ° . This new peak is the result of an

increase in the strength of the axial vortices

produced by this mixer due to the higher

pressure ratio.

5.3.1.2 Lobe Number Effect

Figures 5.17 and 5.18 show PWL, PNL and

SPL for 12UH, 16UH and 20UH mixers at

TO #2 and TO #3 respectively. The overall

relative behavior in terms of PWL spectrum

appears to have remained similar for all the

three mixers with changes in operating

condition, with perhaps 20UH producing a

slight increase in the high frequencies,

similar to those of 12UH at TO #3. The

peak PWL of 20UH appears to have shifted

by a couple of frequency-bands at TO #3.

The relative PNL-directivities of these

mixers also have not changed much.

In summary, increasing the lobe count has

the same effect at higher NPR's as at lower

NPR's with a small relative increase in the

high frequency power output.

5.3.1.3 Effect of Scalloping

Figures 5.19 and 5.20 show PWL, PNL and

SPL for 20UH, 20MH and 20DH mixers at

TO #2 and TO #3 respectively. Again deep

scalloping (20DH) appears to suppress noise

in terms of PWL as efficiently as at low

NPR's (Figure 5.6(b)) at all frequencies.

Even the mid-to-high frequencies

attributable to turbulence in the axial

vortices, shows only a minor peak PWL at

3200 Hz for TO #3 for mixer 20DH.

Actually the competition between the low

NASA/CR 2002-210823/VOL1 160
150' Power Spectrum Comparison

136

132

94

150' SPL Directivity at 1600 hz

128

124

12C

11E

11_

---4k-- CON F

_12UH

--B-- 16UH

--)('-- 20UH

10 100 1000 10000

freq (hz)

1500' Flyover PNL Comparison

65

6O

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

9O

86

U)

78

74 --B--16UH

--X--20UH

70 , I' I' ; ' I, I, I ' [ , J' I' I '

55 65 75 85 95 105 115 125 135 145 155 165

Angle _om InletAxis(deg)

1500' Flyover SPL Spectrum at 115 degs

-.-,2°.

....... _. ........ : ...... q

100 1000 10000

freq (hz)

55

45

40

35

10

Figure 5.17 Power Speetnma and Flyover PNL Comparison of Mixers at T.O.#2 and M(I]) = 0.2

Effect of Number of Lobes

NASA/C_2002-210823/VOL 1 161
140_

1364

132 4

.j 128 4

1244

120 4

116 '

10

95

9O

85

_=

80]

75

70

55

150' Power Spectrum Comparison

--b--12UH

150' SPL Directivity at 1600 hz

91

9,

8:

....... : ........ _= .......

100 1000 10000

freq (hz)

1500' Flyover PNL Comparison

65 75 85 95 105 115 125 135 145 155 165

Angle from inlet Axis (deg)

, t , I

55 65 75

i --e-- CONF

---_-- 12UH

--B-- 16UH

---X-- 20UH

' [ ' I ' ' I , r ,

85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

1500' Flyover SPL Spectrum at 115 degs

75-

70.

65 •

60.

_. 55.

"o

5_so.

u) 45-

40.

35,

30,

25,

10 100 1000 10000

freq (hz)

Figure 5.18 Power Spectrum and Flyover PNL Comparison of Mixers at T.O.#3 and M(fJ) = 0.2.

Effect of number of lobes.

NASA/C_2002-210823/VOL 1 162
150' Power Spectrum Comparison

136

132

128

124

120

116

112

85

83

81

79

77

75

73

71

69 4

-'qK---20DH

= , , ..... : ........ : ......

10 100 1000 10000

freq (hz)

1500' Flyover PNL Comparison

20MH :1_

-")(-- 20DH

671

Zl

65 _ ' ', ' _ ' _ = I ' I ' _ ' I r I ' '1 ' I

55 65 75 65 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

81

8,

7,

150' SPL Directivity at 2000 hz

--m--20UH

--_-'20MH

I _1 ]1 I I ] ---_K---20DH ':':' :.

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

1500' Flyover SPL Spectrum at 115 degs

7o

65

60

55

_so

_J

a,

cn 45

4O

35

30

25

10

_20MH

-'K'--20DH

100 1000

freq (hz)

100(30

Figure 5.19 Power Spectrum, Flyover PNL & SPL Comparison of Mixers at T.O.#2 and M(t_) = 0.2.

Effect of scalloping.

NASA/C_2002-210823/VOL 1 163
14l

13( 9

9

8

7

150' SPL Directivity at 3200 hz

13'

121

12,

121

150' Power Spectrum Comparison

111 ........ :

10 100 1000

------20UH

_20MH

_'__ _20DH l

freq (hz)

10000

1500' Flyover PNL Comparison

95-

9O

85-

i i

Q801

q

75  ̧ _ ....

70 I I I I J I I I , .,

55 65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

55

, I , I , I ' I ' I ' [ ' I ' : ; ' ; '

65 75 85 95 105 115 125 135 145 155 165

Angle from Inlet Axis (deg)

1500' Flyover SPL Spectrum at 115 degs

7;I 7O

/

4O

10 100 1000 10000

freq (hz)

Figure 5.20 Power Spectrum, Flyover PNL & SPL Comparison of Mixers at T.O.#3 and M(t]) = 0.2.

Effect of scalloping.

NASA/C_2002-210823/VOL 1 164
and the high frequency mechanisms of

20DH is made quite apparent in the double

humped nature of the PWL at TO #3 and the

plateau at TO #2 versus the single hump at

low frequencies for TO #1 (Figure 5.6(b)).

This also reflects in the flyover PNL of

20DH as a peak at 115 ° for TO #3, a plateau

from 90 ° to 115 ° for TO #2 and a peak at

100 ° for TO #1 (Figure 5.6(c)). From the

PNL comparison between 20DH and CONF

at TO #3, it appears that the PNL peak of

20DH is smaller than that of CONF, and due

to the steep reduction in noise for angles

higher than the peak angle the 20DH will be

much quieter than CONF on an EPNL basis

at these higher NPR-conditions. The same

can be said at TO #2. However, we will

defer the actual EPNL calculation for now.

On the other hand, the SPL-directivity at

3200 Hz for the moderately scalloped mixer

20MH now shows a larger reduction relative

to the unscalloped 20 lobe design than at

low NPR's. That is, unlike at low NPR's,

even moderate scalloping is now beneficial.

In addition the frequencies associated with

peak PWL increase with increasing NPR.

In summary, the effect of scalloping become

more pronounced as nozzle pressure ratio

increases. Even moderate scalloping

produces a noise benefit at the highest

pressure tested, a result not observed at the

lowest NPR. Deep scalloping produces a

considerable noise reduction relative to the

confluent baseline at the higher NPR's

tested. This is the result of producing

significant low frequency reduction with

only a moderate increase in the mid spectral

bands. A cross-comparison between all eight

mixers will be done later on EPNL basis.

However, for now we will examine the

effect of different free-jet Mach numbers at

the highest NPR condition TO #3.

5.3.2 Effect of Free-Jet Mach Number at

TO #3

The purpose of making a comparison at

different free-jet Mach numbers (0.0, 0.2

and 0.3) is to see the relative spectral

changes at high and low frequencies, and

also at upstream and downstream angles as

in Section 5.2.3, but for the highest NPR

tested for a given mixer so that we carl

understand the source location

characteristics. Hence, we focus on only the

changes in the power spectrum (PWL) and

OASPL-directivity, in the reference frame of

the nozzle, as opposed to changes in flyover

PNL, due to M 0. For convenience of

comparison across mixers we will present

each group on one sheet for each group of

mixers. Figures 5.21, 5.22, 5.23 and 5.24

show the effect on PWL-spectrum and polar

OASPL-directivity for, respectively, CONF,

Group 1 (12CL, 12TH, 12UH), Group 2

(12UH, 16UH 20UH) and Group 3 (20UH,

20MH, 20DH) mixers. Compare this with

similar plots at TO #1 (Figures 5.7 to 5.14).

The confluent mixer (Figure 5.21) shows

similar behavior at all free jet conditions

with a more uniform decrease in PWL over

the whole frequency spectra and in OASPL

for all angles than observed for all the lobed

mixers. All lobed mixers at this high NPR

condition show a marked insensitivity in

certain regions to free jet conditions.

(i) PWL for the mid-to-high frequencies

changes very little with changes in

free-jet speed. However significant

changes in PWL are observed at the

lower frequencies with changes in the

free jet flow. This is more evident at

TO #3 than at TO #1.

(ii) Similarly, the OASPL for upstream

angles does not change much with a

change in free-jet speed, especially

NASA/C_2002-210823/VOL 1 165
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NASA/C_2002-210823/VOL 1 169
compared to the larger changes at

downstream angles. This, too, is more

evident at TO #3 than at TO #1.

It is remarkable to note that local secondary

peaks in the mid-to-high frequency range for

mixers 12CL and 20DH start to become the

predominant features in the PWL-spectra as

the free-jet Mach number increases at this

high NPR condition. This was not evident

at the low NPR conditions of TO #1. So

whatever mechanism is causing that high

frequency peak now controls the noise

suppression ability of these two mixers. For

all other lobed mixers this mechanism was

dominant even at low NPR conditions. We

have remarked repeatedly in the past that

this peak at higher frequency must represent

the effect of turbulence intensity due to axial

vorticity and mixing of fan and core flows

both inside and outside the nozzle. We

examine this in more depth in the next

section by analyzing the SPL-spectra at

various angles.

5.4 Diagnosing High-Frequency

Source Locations with Changes in Free-Jet

This analysis was prompted by the

observations in the previous sections that the

sound power spectrum, PWL, does not

change much in the mid-to-high frequencies

for lobed mixers when the free-jet speed is

changed. This is also reflected in the

relative insensitivity to free-jet speed of the

polar OASPL and the flyover PNL

directivities at angles upstream of peak

PNL. It is important to understand the

reason behind this trend because the PNL

associated with angles near 90 ° is within 2-3

PNdB of the peak PNL value, which would

imply a significant contribution to EPNL.

This observation was not found to hold good

for the CONF mixer, which re-enforces the

suspicion that the turbulence in the axial

vortices of the forced lobed mixers may be

the cause of the high frequency noise. We

proceed to explore the polar SPL-spectra at

different free-jet Mach numbers in detail in

this section.

5.4.1 Ray Theory Application

Since our interest principally lies in

relatively "high" frequencies, we will use

ray theory or geometric acoustics. Consider

an acoustic ray radiating from a source

inside the jet at some angle to the jet axis

and coming out through the shear-layer

between the model-jet and the free-jet with

the observer inside the free-jet. When the

free-jet speed is changed, the outgoing ray in

the free-jet will change its direction due to a

change in the refraction of the wave-number

vector and also due to a change in the

convection speed of the wave-fronts at the

altered free-jet speed. Hence, if we want to

study the change in the source

characteristics due to a change in the free-jet

speed, then we should compare the rays

coming from the jet for different free-jet

speeds, not at the same angle in the free-jet,

but at different refracted ray angles

corresponding to the same incident ray from

the source. Also, in such a comparison we

are ultimately interested in the direction of

the rays, and not of the wave-number

vectors, because acoustic energy transmitted

to an observer inside the free-jet is carried at

the group velocity along the ray directions

_22._3).In a moving media, ray direction is not

the same as the wave-number vector

direction or nomaal to the wave-fronts.

Further, in the geometric acoustics limit

there is no reflection of the incident ray from

the shear layer and the incident ray is fully

transmitted. However, it is known from

Candel's planar wave analysis 21 that the

amplitude of the wave that is transmitted

through a shear-layer discontinuity can be

NASA/CR 2002-210823/VOL1 170
differentthanthe amplitudeof theincident

wave(theirratio beingcharacterizedby the

"transmissioncoefficient")and,importantly,

it varies with the incident wave-normal

angle.Hence,we alsoneedto estimatethe

changesin thetransmissioncoefficientwith

changesin free-jetspeedfor agivenincident

waveamplitudesothat we canaccountfor

thedifferencesif theyarelargeenough.

Theresultsfor theangularrelationshipsdue

to refraction and convection, and the

amplitude changes due to transmission

coefficient changes are given in this section

from two-dimensional considerations

assuming planar waves. This is an extension

of Candel's wave analysis 2_ to moving flows

on both sides of a shear-layer. Consider the

geometry of the angles on either side of the

shear-layer between the free-jet and the

model-jet. This shear layer is modeled, for

convenience, as a thin planar vortex-sheet

on the nozzle lip-line, as shown in Figure

5.25. The model-jet flow, assumed to be

uniform, on one side of the lip-line is

labeled as region 1. The free jet flow on the

other side, also assumed to be a uniform

flow, is labeled region 2. The source S,

located inside the model jet and assumed

stationary, radiates a wave at local speed of

sound c_ with wave-normal angle 01 with the

jet inlet axis. This wave-front convects

downstream with the local flow speed Us =

M1Cl to give the direction of the incident ray

from the source, as shown using the velocity

triangle. Such angular relationships between

the wave-front normals and the ray

directions in moving media through velocity

triangles are well-known and can be found,

for example, in Lighthil122. This incident

wave-front is refracted through the thin

vortex-sheet. When the free-jet speed is

zero, the transmitted wave-front normal

subtends an angle of 020 with the inlet axis

and the wave travels in region 2 at the local

speed of sound, c2, as shown in Figure 5.25.

This is also the direction of the transmitted

ray when the free-jet speed is zero.

However, when the free-jet speed is nonzero then the change in the propagation

angle of the transmitted ray can be

conveniently broken into two parts:

(i) refraction of the corresponding wavefront normal, denoted by angular change of

A¢r,

(ii) convection of the wave-fronts using

velocity triangle, denoted by angular change

of A0_.

This leads to the final transmitted ray angle

of

q_2 = q}20 + Aq_r + A_c"

(Note: We will use lower-case Greek

symbols for wave-normal angles and uppercase Greek symbols for corresponding ray

angles.) From the detailed analysis given in

Appendix C we obtain:

C2/cclcos _l

COS _2 =

1 - cos CI(M 1 - C_/cl M2)

tanA¢ c= M 2sin_2

1 - M 2 cos ¢2

where

02 = 020 + A0r"

is the angle of the wave-normal in the free

jet. Hence, 020 , which is 02 at M2 = 0, is

given by

c2//cc1 COS 01

COS _20 =

1 - M 1 cos ¢1

By using the above formulae, we can

directly relate ¢2 to 020 without referring to

the model jet (or flow 1) properties:

NASA/CR 2002-210823/VOL1 171
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NASA/C_2002-210823/VOL 1 172
cos 02 = cos 020

1 + M 2 cos q_20

This relationship only requires knowledge of

the free-jet Mach number Mz , without

involving the model-jet properties, M1 and

Ca. Thus given _20 and M2 we can find (I) 2 _-

d_z+A_c from the above equations without

knowing M1 and cl.

Since rays radiate from different portions of

the jet, it would be difficult to utilize these

results in the geometric near-field. However,

as the distance R from the jet to the observer

becomes sufficiently large (geometric farfield), some simplifications occur. For these

simplifications to be valid, R must meet two

conditions: firstly, R must be large

compared to the distance from the nozzle-lip

to the point where the incident ray crosses

the shear layer, denoted by l in Figure 5.25;

secondly, R must be large compared to the

nozzle exit radius, rexit. It is clear from

Figure 5.26, which depicts this situation,

that in the geometric far-field inside the free

jet flow when l/R << 1 and rexit/R << 1:

where 020 is measured from the center of the

nozzle-exit plane, as usual, and A_ is change

due to refraction or convection at the lipline. This can also be proven mathematically

very easily. Thus, instead of using _2o and

_20 + A_ one can use 020 and 020 + A0 at a

sufficiently long distance R inside the freejet. In other words, in the geometric far field

we can refer to observer distances and

angles from the center of the nozzle exit

rather than from the apparent source

locations.

At full-scale under study, rexit -" 1.208 ft and

R = 150 ft, so that rexit/R = 0.008. For midto-high frequencies of interest in such

comparisons, the responsible sources are not

too far from the nozzle-exit plane making l

of the order of rexit. Even taking the worst

case of l = 10rexit gives l/R = 0.08 which

creates hardly any error by using 02o and A0

instead of _2o and A#, respectively. Recall

from Chapter 3 that the polar SPL data at

150 ft with a free jet flow has been obtained

from a free jet of finite diameter after first

correcting for the refraction effect from the

shear layer between the free jet and the

ambient using Ahuja et al (_5)type of method

(see Appendix C). This polar 150 ft SPL

data corresponds to the angle of the wavenormals, 02 (or the "emission" angle (15))

inside the free-jet and in the reference frame

of the nozzle.

Table 5.1 lists the final refracted ray angle

1_2 (=_2) and the refracted wave-normal

angle, 02 (= q2) for a range of input "static"

free-jet angles 020 (= _20) at two different

free-jet Mach numbers of interest, M 0 = 0.2

and 0.3. The intermediate angles which go

into this calculation are also given, namely,

the pure refraction angular change Arefr (=

ACt) and the pure convection angular change

Aconv (= A_c). The blank spaces for the

shallow angular range (larger values of 020)

give complex values of O

2 and correspond to the "zone of silence,"

which begins at slightly different values of

020 for the two free-jet Mach numbers.

Figure 5.27 shows the two angular changes,

A_r and A_)c as functions of ¢20. Thus note

that at 90 ° there is no refraction effect on the

wave-normals and the bending of the rays is

purely due to convection of the wave-fronts.

This is also true approximately for angles

near 90 ° . For very small angles and angles

just before the zone of silence begins the

refraction effect is much larger than the

convection effect.

NASA/CR 2002-210823/VOL1 173
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r

NASA/C_2002-210823/VOL 1 174
Table 5.1 Refraction & Convection Effects with Free-Jet On

M0 = 0 MO = 0.]2 MO = 0,3

020 02 02 02 02

0

5

10

15

20

25

3O

35

40

45

50

55

60

5

70

75

80

85

90

95

100

105

110

115

I20

125

130

135

140

145

150

155

160

165

170

175

180

33.56

33.83

34.64

35.95

37.72

39.89

42.42

45.26

48.37

51.72

55.28

59.03

62.96

67.07

71.33

75.75

80.34

85.09

90.00

95.09

100.36

105.84

111.54

117.49

123.75

130.38

137.53

145.44

154.78

168.42

Arefr Aconv

33.56 7.56

28.83 7.61

24.64 7.75

20.95 7.98

17.72 8.27

14.89 8.62

12.42 8.99

10.26 9.39

8.37 9.78

6.72 10.16

5.28 10.51

4.03 10.82

2.96 11.09

2.07 11.30

1.33 11.44

0.75 11.52

0.34 11.53

0.09 11.46

0.00 11.31

0.09 11.08

0.36 10.75

0.84 10.34

1.54 9.83

2.49 9.23

3.75 8.51

5.38 7.68

7.53 6.71

10.44 5.56

14.78 4.13

23.42 1.92

41.11

41.44

42.39

43.92

45.99

48.51

51.42

54.65

58.16

61.88

65.79

69.86

74.05

78.36

82.77

87.28

91.87

96.55

101.31

106.16

111.12

116.18

121.37

126.72

132.26

138.06

144.24

151.01

158.90

170.34

39.72

39.92

40.52

41.50

42.86

44.56

46.57

48.89

51.47

54.31

57.39

60.70

64.23

67.97

71.93

76.10

80.50

85.13

90.00

95.13

100.56

106.30

112.40

118.95

126.03

133.85

142.78

153.83

174.06

Arefr Aconv

39.72 13.99

34.92 14.04

30.52 14.17

26.50 14.38

22'86 14.66

19.56 14.99

16.57 15.35

13.89 15.73

11.47 16.10

9.31 16.45

7.39 16.78

5.70 17.05

4.23 17.26

2.97 17.40

1.93 17.46

1.10 17.42

0.50 17.29

0.13 17.05

0.00 16.70

0.13 16.22

0.56 15.62

1.30 14.87

2.40 13.98

3.95 12.91

6.03 11.65

8.85 10.15

12.78 8.33

18.83 5.95

34.06 1.37

53.71

53.95

54.69

55.89

57.52

59.54

61.92

64.61

67.57

70.77

74.17

77.75

81.49

85.37

89.39

93.53

97.79

102.18

106.70

111.36

116.17

121.17

126.38

131.86

137.68

144.01

151.12

159.78

175.43

$

020 = Angle of wave-front normal with inlet axis in the static free-jet (M 0= 0)

02 = Angle of wave-front normal with inlet axis in the free-jet purely due to refraction

Arofr= 02 - 020, Pure refraction effect without convection

Aconv= Purely convection effect after accounting for refraction effect

02 = 020 + Arotr+ Aconv,Final angle of the ray in the free-jet

* = Gives complex values of 020corresponding to "zone of silence"

NASA/C_2002-210823/VOL 1 175
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NASA/C_2002-210823/VOL 1 176
Therelationshipsbetweenthevariousangles

given above are not restricted by the

assumptionof plugflow in themodeljet or

the thin vortex-sheetmodel of the freejet/model-jet shear layer but are also

applicablewhen there is variationin the

flow propertiesin the verticaldirectionand

the shear-layeris thicker. This is true so

long as the static pressurein the vertical

directionremainsfixed. In that casethe

model-jet flow properties in the above

relations,namely, M1 and cl are to be

replaced by the local model-jet flow

propertiesat the source location. The final

relation between _2 and #20 or _2 and q_20is,

however, independent of model-jet flow

properties, as mentioned earlier.

Further, these relationships between angles

are applicable not only to planar twodimensional flows but also to the cylindrical

case if the two flows are assumed uniform,

the thin vortex-sheet is a cylindrical surface

and the point source is on its axis. In this

case, any incident ray from the source is

always in the meridional plane, rx

(containing the radial direction r and the

axial streamwise direction x), whence, its

local interaction at the cylindrical vortexsheet can be considered as if both the

vortex-sheet and the incident wave-fronts

are locally planar. This results in the

transmitted ray being in the same meridional

plane as the incident ray. That is, Figure

5.25 can be interpreted as if it is drawn in

the meridional plane of the incident ray and

all the angular relationships remain the same

as derived before for the planar case.

Next we tackle the issue of "transmission

coefficient" changes at different free-jet

Mach numbers and the corresponding

observation angles. The detailed analysis

and results are again given in Appendix C.

These results show that for a range of

"static" free-jet angles 020 not too far from

90 ° the changes in the SPL of the

transmitted ray-tube due to changes in the

free-jet Mach number up to 0.3 are at most

of the order of + 0.4 dB. The errors are less

in the aft quadrant than in the front quadrant

and for hot jets the errors are even smaller.

This is small enough for our purposes of

comparison and, hence, can be neglected if

we restrict ourselves to that angular range.

Recall that, in any case, in the geometric

acoustics limit the rays are not reflected and

our analysis will be consistent if we indeed

neglect the transmission coefficient changes.

5.4.2 The New Diagnostic Method for

"Excess" Sources

Let's say that we know the 150 ft polar SPL

data inside the free jet at different

"emission" angles, 02 (= 02, see Figure 5.25)

for different M_, as in our data base. Then

in order to study the effect of free-jet Mach

number on the high-frequency source

characteristics we should compare the farfield polar SPL of Mfi = 0 at angle 020 with

that of Mfj > 0 at angle 0z = 020+A0r as given

in Table 5.1. Thus, we will compare the farfield polar SPL's of different M0's at

corresponding "shifted" angles, denoted by

the following pairs:

(Mfj = 0, 020) _ (Mrj>0, 02).

If the 150 ft polar SPL data inside the free

jet is referred to the ray angle, 02, or

"reception" angle (15) then comparison of

SPL's at different Mtj's should be made at

the corresponding shifted ray angles, also to

be found in Table 5.1, and this relationship

is denoted by:

(Mfj = O, 02o) +-->(Mfj>O, 02).

It is possible that the polar SPL data is

available at only certain angles, as in our

case, from 02 = 55 ° to 165 ° with 5 ° interval.

NASA/CR 2002-210823/VOL1 177
In that case,data may not be available

exactlyat the angle0z demanded by Table

5.1 for different M_. As an approximation,

SPL data at the closet available angle will be

used. With data available at 5 ° intervals the

maximum error in angle will be 2.5 °. When

we actually compare data we will select

angles which will minimize this error for the

available Mfj of 0.2 and 0.3.

Since our whole analysis here was driven by

the observation that the sound power in the

mid-to-high frequencies does not change

much with free-jet speed, let us see from the

viewpoint of source location, what this may

mean. Figure 5.28 shows several sourcelocations for a jet with internal mixing and a

table listing if those sources should change

in intensity due to free-jet speed changes.

Generally speaking a change in free-jet

speed changes the mixing characteristics

between the nozzle-jet and the free-jet at the

lip shear-layer and further downstream

where the two lip-shear layers from

diametrically opposite ends of the nozzle

interact as well as in the far downstream

plume flow-field. However, if there are any

other sources of noise away from the free-jet

related shear-layers then their turbulence

intensity will not be modified much due to

changes in the free-jet speed, if we neglect

"back-reaction" on those sources due to an

impedance change at the lip shear-layer.

For example, source 1 in Figure 5.28 is

inside the nozzle and may arise due to

internal fan-core flow mixing and hence, not

change in strength due to changes in free-jet

speed with a fixed ambient pressure at the

lip. Similarly source 2 in that Figure, shown

away from the lip shear-layer, can arise due

to fan and core flows not being completely

mixed at the nozzle-exit plane and will not

change in strength due to changes in free-jet

speed. Back-reaction on both these sources

can be neglected. Both these sources, 1 and

2, are in some sense "excess" noise sources.

The first is an "internal" excess noise source

and the second an "external" excess noise

source. These are "excess" noise sources in

the sense that they will not occur in a single

stream nozzle with "uniform" exit jet

velocity producing the same thrust. In the

case of forced lobed mixers these are driven

by axial vorticity, have small length-scales,

and hence, are candidates for mid-to-high

frequency noise sources.

On the other hand, sources in locations 3, 4

and 5 in Figure 5.28 are in direct interaction

regions with the free jet flow and will

change their strengths as M 0 changes. For

example, earlier in Section 5.2.3 we had

shown how the radial gradient of velocity at

the end of the potential core near region 4

decreases with an increase in M 0 thus

reducing noise in fully mixed jets.

However, this should not be taken to imply

that effects of axial vorticity are only

confined to source locations 1 and 2. For

example, in the case of high lobe penetration

the axial vortices may not completely

diffuse internally, and will then interact with

the lip-shear layer at location 3. For lowlobe penetration these vortices can continue

to convect all the way up to location 4.

From the table in Figure 5.28, we can

conclude that if we find that the far-field

high-frequency SPL's for different free-jet

Mach numbers at the appropriately "shifted"

angles do not change, then they probably are

coming from "excess" sources either from

source location 1, inside the nozzle or from

source location 2, close to the nozzle exit

plane and away from the lip-shear layer.

Further, if we have flow data at the nozzle

exit plane which shows uniform flow

(velocity and temperature) then it would

mean that there are no "external" excess

sources at 2 and the burden of invariant

SPL-data will completely fall upon

"internal" excess noise sources. In the case

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NASA/C_2002-210823/VOL1 179
of forced lobe mixers, these are nothing but

the turbulence generated by the dominant

axial vortex structures. On the other hand, if

the SPL's at shifted angles do change then

this method fails to say anything definite

about the excess noise sources.

It should be noted that internal sources of

noise can also radiate to the far-field

upstream quadrant. This is illustrated in

Figure 5.29 for an internal source in a duct

with uniform flow and zero free-jet speed,

for simplicity. Consider a noise source

inside the nozzle with its wave-number

vector k having an upstream facing

component, kx < 0, with phase velocity

vector of magnitude c (local speed of sound

in the jet) and group-velocity vector or

acoustic ray velocity facing in the

downstream direction obtained by

vectorially adding c (= ck/lk[) and the local

jet velocity vector V. We assume here that

kx is such that it does not correspond to total

internal reflection. This downstream going

ray in the jet finally interacts with the lip

shear-layer, represented as a vortex-sheet,

either directly or after reflecting from the

duct-walls. At the vortex-sheet between the

ambient and the nozzle-jet, the equality of

static pressure implies that the xcomponents of the wave-number vectors in

the ambient and the jet must be equal, i.e.,

kxl = kx where subscript 1 stands for flow

outside the jet. Hence if kx has an upstream

facing component at the source kx < 0, then

kxl < 0 too, and the far-field ray in the

quiescent ambient will propagate in the

upstream quadrant. With a free jet flow

present and V_ > 0 this far-field ray will be

further deflected due to convection, as

already discussed using Figure 5.25. While

a vortex-sheet model was used here for

convenience, this whole argument is also

true for a thin lip shear-layer with constant

static pressure within its thickness because

the x-component of the wave-number

vector, kx, remains constant throughout the

shear-layer thickness.

In summary, only downstream going rays

from internal sources within the jet, which

do not correspond to total internal reflection

angles, can radiate outside the nozzle; all

upstream going rays from internal sources

within the jet are trapped inside the nozzle.

As a special and important subset of these

rays, only the downstream going rays from

the internal source within the jet which

correspond to upstream facing wave-number

vectors at the source will radiate into the

upstream far-field quadrant when no free jet

flow is present.

With this new diagnostic method in hand, let

us examine the high NPR condition TO #3

because it showed better invariance of PWLspectrum than TO #1 which corresponds to

low NPR.

5.4.3 Application to Operating Condition

TO# 3

Figure 5.30 shows the comparison of polar

SPL at a 150 ft radius inside the free-jet for

mixer 12UH at free-jet Mach numbers of

0.0, 0.2, and 0.3 at the "shifted" emission

angles obtained from the equation of the

previous section or Table 5.1. Here, the

reference case is taken to be at 020 = 90 ° for

M 0 = 0.0 (the "static" angle). Since we are

using emission angles we obtain, from Table

5.1, 02 = 90 ° also for both Mtj = 0.2 and 0.3

as the "shifted" angle. In this case data is

indeed available exactly at 90 ° for Mcj = 0.2

and 0.3 and, hence, the error in the angle for

which data is plotted is zero. This error in

angle is shown in brackets after the

corresponding shifted angle in the legend of

Figure 5.30. We follow this nomenclature in

this section. We see that frequencies beyond

1000 Hz do not change significantly (they

are all within the margin of error for SPL)

NASA/CR 2002-210823/VOL1 180
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NASA/C_2002-210823/VOL 1 182
with changing Mfj. This implies that the

invariant high-frequency noise at angles at

90 ° is most likely associated with "excess"

noise sources lying either inside the nozzle

or close to the exit plane and is due to the

turbulence caused by the axial vortices

generated between the fan and core streams.

We unfortunately do not have the plume

data at the exit plane for this condition to

nail down the excess sources any further as

"internal" or "external."

Let's examine the "shifted" SPL's at

different angles.. In Figure 5.31(a) the data

at "static" angles of 80 ° and 100 ° are

compared to appropriately shifted angles for

the various free-jet Mach numbers tested.

The shifted angles, 02r, for both free-jet

Mach numbers are also 80 ° and 100 °

respectively from Table 5.1. It appears that

at 80 ° the SPL's for frequencies around

2000 Hz are quite invariant for the two freejet Mach numbers, implying that internal

and/or external excess sources are

responsible for these frequencies at this

angle. The downstream angle associated

with 100 ° does not show as much invariance

at higher frequencies. If we further compare

in the upstream quadrant at angles 85 °, 75 °

and 55 °, as shown in Figure 5.31(b), then

the previous invariance of SPL appears to

hold good at intermediate frequencies near

2000 Hz. For angles less than 70 ° or more

than 95 ° this SPL invariance appears to

falter. Hence, it appears that at least for a

certain angular range near 90 ° the internal

and/or external excess sources are, indeed,

responsible for the noise at frequencies near

2000 Hz and higher. For the corresponding

range of angles near 90 ° for M 0 = 0.2, the

resulting PNL is within 3 to 4 PNdB of the

peak PNL value for mixer 12UH (see Figure

5.18) and, hence, the excess noise source

should be a significant contributor to EPNL.

We have analyzed all the mixers at

operating condition TO #3 using the

diagnostic technique of comparing SPL's at

"shifted" angles for different Mfj's. Some of

these results are attached in Appendix D and

show comparisons at "static" angles of 75 ° ,

90 ° and 100 ° and corresponding "shifted"

angles at higher Mfj's. (Note from table 5.1

that the corresponding shifted angles are

also approximately 75 ° , 90 ° and 100 °

respectively.) Examination of all these

figures given in Appendix D shows the

following:

° All lobed mixers show strong

invariance of SPL at mid-to-high

frequencies for "static" 75 ° when Mfj

is varied from 0.0 to 0.2 to 0.3 similar

to the 12UH mixer discussed earlier.

Actually this was seen to hold good for

a small range of angles in the upstream

quadrant near 90 ° .

, In the downstream quadrant away from

90 ° the SPL invariance breaks down.

. In the scalloped lobe mixers 12CL and

20DH although the SPL variation at

higher frequencies was less that at low

frequencies, clean SPL-invarianee was

not observed even at 90 ° . That is,

compared to the SPL invariance in

unscalloped mixer 20UH, the SPL's

for 12CL and 20DH mixers at shifted

angles varied more than 20UH for high

frequencies when M 0 was varied.

. The sound pressure levels of the

confluent mixer (CONF) do not show

invariance at "shifted" angles for

different free jet Mach numbers at any

frequency. This means that the

dominant noise sources are either not

the internal or the external excess

noise sources, or that if any excess

noise sources are present then they

have merged with the free jet outside

NASA/CR 2002-210823/VOL1 183
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NASA/C_2002-210823/VOL 1 185
the nozzle. This configuration does not

produce axial vorticity structures and

little mixing occurs inside the nozzle.

The fan/core flow mixing outside the

nozzle near the exit-plane, which is

away from the free-jet, is a region

which is not a dominant noise

producing region as explained in the

model of a coplanar coaxial jet for the

"initial" region contained in Fisher et

al(17).

In summary, application of the new

diagnostic technique at operating condition

TO #3 shows that all unscalloped mixers

(12UH, 16UH, 20UH) appears to have

"excess" noise sources which emit noise at

upstream angles close to 90 ° . These excess

sources can be either inside the nozzle duct

or close to the nozzle exit plane near the jet

axis representing partial fan/core stream

mixing. The turbulence generated by the

strong axial vortices found in these forced

lobed mixer configurations is the probable

candidate for the excess source.

Perturbation of streamwise vortices which

have convoluted vortex surfaces with

normals in the transverse plane will emit the

strongest sound in the radial direction or at 0

= 90 ° which is what has been observed.

However, the deeply scalloped mixer

(20DH) and the cutout mixer (12CL) do not

show this behavior clearly. Hence, their

excess sources, if any, are either weak or

they interact with the free-jet/model-jet

shear layer to change the far-field SPL's at

"shifted" angles, as observed. 20DH mixer

also typically shows fairly uniform velocity

profiles at the nozzle exit plane, except for a

small hump at the center (see Figure 4.13).

Hence, its internal noise sources must be

weak radiators. Recall that 20DH and 12CL

are also the quietest mixers, especially in the

mid-to-high frequency range (see Figures

5.19, 5.16). Thus, this "shifted" angle SPL

comparison in a way complements the

previous acoustic findings.

We have applied this technique also at the

lower NPR condition, TO #1, to all mixers,

and the results do not show much SPL

invariance with free jet Mach number,

implying that these "excess" sources

become less important at the lower jet

speeds (830 ft/s) associated with this

operating condition.

5.5 Effect of Nozzle-Length

In this section acoustic data is presented and

analyzed for mixers with different nozzlelengths. The nozzles tested varied the

mixing-length or the distance between the

mixer exit plane to the nozzle exit plane, L,

between 50%, 75% and 100% of the

baseline mixing length. These correspond,

respectively, to L/Dmp = 0.55, 0.83 and 1.10

where Dmp is the mixing-plane diameter.

The 75%L configuration was tested in

conjunction with only three of the available

lobed mixers (12TH, 20UH, 20DH). In each

case the nozzle exit plane diameter was kept

the same, as shown in Figure A.5.

Since the recording microphones were at

fixed locations in the APL dome, the

distance between them and the nozzle exit

plane changed slightly as the nozzle-length

was varied. The changes in the nozzlelength, AL, are at most 5.5" (distance

between the exit planes of 100%L and

50%L) and the microphones are at R = 48 ft.

As a result, AR/L = 0.0095 and the changes

in the microphone angles and distances

which are measured from the center of the

baseline (100%L) nozzle are negligible.

Initially we examine the effect of nozzlelength on the overall acoustic metrics at all

performance conditions. In the second

section we examine the SPL-spectra at

various angles. Finally we examine the

NASA/CR 2002-210823/VOL1 186
probable mechanismthat produces the

observedbehavior.

5.5.1 Overall Acoustic Metrics

Earlier studies with the confluent nozzle

(CONF) showed that the nozzle-length or

mixing-length, L, did not make any

difference in far-field noise. (Some results

from 1995 tests for CONF mixer are

reported in Mengle et al(16).) This is an

expected result because CONF does not

produce any appreciable internal mixing

and, therefore, changing mixing lengths

does not change the fan/core or jet/ambient

mixing behavior. Hence, we will not deal

with the length effects for CONF mixer.

The focus here will be on mixing-length

effects for the lobed mixers and the tongue

mixer.

Figures 5.32 to 5.38 show the power

spectrum, polar OASPL directivity and

flyover PNL-directivity for, respectively,

12CL, 12UH, 12TH, 16UH, 20UH, 20MH

and 20DH mixers at conditions TO #1, #2

and #3 with a free jet Mach number of Mfj =

0.2 for all nozzle-lengths tested. These are

very comprehensive plots. Examination of

these plots reveals the following interesting

facts regarding mixing-length effects:

. Reducing mixing length does not

necessarily produce an increase in

noise, as commonly believed. For

example, for mixers 12TH, 20UH and

20DH at TO # 3 (Figures 5.34, 5.36,

5.38) both PNL and OASPL decreases

at most angles as the mixing length is

decreased from 100% to 75%L and

then increases with further reduction to

50%L, showing a local minimum for

noise. This is also true for mixer 12CL

(Figure 5.32) where a reduction from

100%L to 50%L produces a decrease

in noise for, all the operating

.

.

°

conditions. This noise decrease due to

the reduction in L looks substantial

and will be quantified in terms of

EPNL later.

The reduction in noise with a decrease

in L is not always observed for the

lower NPR conditions for these

mixers. For example, a noise increase

is observed for mixer 12TH as the

nozzle-length decreases for conditions

TO #1 and #2. The PNL directivity of

the 20 lobe mixer with deep scallops,

20DH, is essentially unchanged when

the mixing-length is reduced from

100%L to 75%L at the lowest pressure

ratio, TO #1. However as the NPR is

increased, a PNL reduction is noted as

the mixing length is reduced from

100%L to 75%L for configuration

20DH at angles less than or equal to

125 °. However, 20UH shows a local

minimum in noise at 75%L for all

take-off conditions TO #1, #2 and #3.

The three mixers 12UH, 16UH and

20MH (Figures 5.33, 5.35, 5.37), for

which no 75%L tests were done,

always show an increase in noise when

the mixing-length is reduced by half.

Since all the other noise characteristics

of these three high-penetration mixers

are very much like those of 20UH, it is

very likely that they too will show a

local noise minimum if data were

obtained at 75%L.

In terms of the PWL-spectra, changes

in mixing-length produces substantial

changes in the mid-to-high frequency

range, but the low frequency is

unchanged at all operating conditions

for all mixers. Since near the peak

PNL angle and upstream of it mid-tohigh frequencies dominate, this

spectral behavior is reflected as larger

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changes in PNL in the corresponding

angular regions and smaller changes at

the shallow angles. Similar changes

were noted in the OASPL directivity.

Thus, using 20UH as an example

(Figure 5.36), large changes in peak

PWL at 2000-3000 Hz are reflected as

large changes in peak PNL at 100 °-

115 °. Consequently if there is a

reduction in the PWL-spectra at 2000

Hz going from 100% to 75%L then it

also appears as a reduction in the PNL

at the corresponding angles near the

peak PNL angle.

. The striking independence of low

frequency PWL on mixing-length for

all mixers at all conditions implies that

the far downstream plume decay must

be independent of the mixing length.

Any reduction in noise with decrease in

nozzle length is always welcome from the

point of view of nozzle weight but its effect

on thrust should also be considered. We

will investigate the reasons for this

interesting acoustic behavior by scrutinizing

the polar SPL-speetra in the next section.

5.5.2 Polar SPL-Spectra

Figures 5.39 to 5.45 show the 150 ft polar

SPL-spectra for all nozzle lengths at 60 °,

90 ° and 115 ° for the lobed mixers 12CL,

12UH, 12TH, 16UH, 20UH, 20MH and

20DH respectively for all operating

conditions and at M 0 = 0.2. These Figures

correspond to those in the previous section

for each mixer and should be used for

reference.

Examination of these SPL-plots reveals the

following:

. Changes in OASPL due to mixinglength changes at a given angle arise

.

.

mainly due to changes in SPL in the

mid-to-high frequency spectra. The

low frequency portion of the SPL is

almost invariant to nozzle-length

changes at all angles. The very high

frequency spectra also appears

insensitive to nozzle-length,

particularly at the low nozzle pressure

ratio of TO #1. This spectral behavior

is reflected in the PWL spectra

discussed in the last section.

Using mixer 20UH at operating

condition TO #2 as an example

(Figures 5.43 and 5.36), the changes in

OASPL or PNL at 115 ° due to mixinglength changes must be due to

frequencies larger than 800 Hz since

frequencies lower than 800 Hz do not

change. Hence changes in the flow

structures that cause the mid-to-high

frequency sources must also be the

dominant reason for changes in PNL

or OASPL due to mixing-length

changes. Earlier (Section 5.4.3 and

Figure D.5) we have found that

internal and external "excess" noise

sources contribute largely to this midto-high frequency range at angles near

90 ° at Mfj = 0.2, and that the

turbulence in the axial vortices are the

likely noise generators. Interaction of

these vortex structures with the free-jet

in the nozzle lip-shear layer or further

downstream near the end of the jet

"core" will also produce mid-to-high

frequency noise at other angles, e.g.,

where PNL peaks.

In order to understand why a local

minimum for sound exists when

mixing-length is decreased, we focus

on a fixed frequency, say, at or near

the spectral peak and see how its SPL

changes at all angles. Recall that the

polar SPL being examined is in the

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.

reference frame of the nozzle so that

we do not need to Doppler-shift the

frequency at different angles for

comparison. Again using 20UH as an

example (Figure 5.43), we see that the

2000 Hz component of SPL decreases

by similar amounts at all angles when

the mixing-length is reduced from

100% to 75%. This is also true for all

the neighboring frequencies in the

mid-to-high frequency bands. Since

the far-field radiated sound due to a

source can be thought to be made up of

the product of an amplitude factor

(intensity) and a directivity factor, this

observation at all angles implies that

the source producing these mid-to-high

frequencies is decreasing in intensity

as the length is decreased from 100%

to 75%. Since the turbulence in the

axial vortex structures is the cause for

this frequency range, this turbulence

intensity which is related to the

strength of the axial vortices must be

getting weaker. It was also noted

previously that the noise starts

increasing when the mixing-length is

reduced further to 50%L. This

suggests the presence of "competing"

mechanisms - one which decreases

noise and the other which increases it

when the mixing length is reduced.

We explore the competing

mechanisms in the next section.

From Figure 5.45A it is also observed

that as the mixing length is decreased

from 100%L to 75%L the peak SPL

frequency (in terms of 1/3 ra octaveband center frequency) generally

decreases, and when the mixing length

is further decreased to 50%L the peak

SPL frequency always increases, thus

exhibiting a local minimum in the peak

SPL frequency at a given angle. This

curious fact is worthy of attention

.

because one would otherwise expect

that as the nozzle length is decreased

the peak SPL frequency ought to

monotonically increase due to more

and more exposure of the smaller

eddies produced by the mixing of fan

and core streams. Also the decrease in

peak SPL frequency for 75% L takes it

away from the most annoying spectral

region of 3000-4000 Hz. This

combined with its reduced intensity,

observed earlier, no doubt will reduce

the perceived noise levels for 75% L.

Such perceived noise levels will be

studied in Section 5.6.

With a free-jet Mach number of Mvj =

0.2 in the above data, changing the

nozzle length by a few inches also

changes the boundary-layer thickness

on the outside of the nozzle wall at the

exit-plane by a small amount. It can

be argued that this change in boundary

layer thickness can affect the evolution

of the jet plume downstream of the

nozzle exit and, hence can affect this

peculiar phenomenon of "local

minimum of noise." In order to check

if this unique phenomenon exists

without any external boundary layer,

we examine the static case, M_ = 0,

where there is no external boundary

layer. Figure 5.46 shows the static

polar SPL's for 20UH at TO #3 with

different nozzle-lengths. We see that

75% nozzle-length again shows a local

minimum in SPL at mid-to-high

frequencies, as well as a local

minimum for the frequency of peak

SPL as before. This implies that the

effect of the external boundary-layer is

not an important part of the

mechanism which produces the local

noise minimum.

NASA/C_2002-210823/VOL 1 203
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TO #3, M(fj) = 0.2.

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NASA/C_2002-210823/VOL 1 205
5.5.3Competing Mechanisms for Local

Minimum of Noise

We have observed from Figures 5.41, 5.43

and 5.45 that all the lobed mixers tested with

all three mixing lengths showed a local

minimum for noise at some of the operating

conditions. A local minimum exists when

there are at least two competing mechanisms

which are functions of the parameter being

varied -- one which increases noise and the

other which decreases it. We discuss the

probable reasons for this behavior.

It is not difficult to see why noise for

configurations with lobed mixers should

increase when the mixing length is

decreased. As we have shown earlier, these

flows are dominated by axial vorticity and

the consequent turbulent mixing between

fan and core flows inside the nozzle duct.

These turbulent vortices comprise the

internal noise generator. When the duct

length is decreased the sources in the aft

portion are acoustically exposed or

"acoustically unshielded" as shown in

Figure 5.47. For a "shielded" source inside

the duct, like A in Figure 5.47, all upstream

going rays from it are trapped and only

downstream going rays can be radiated

outside the nozzle, if their wave-number

vector angles do not correspond to total

internal reflection angles, as discussed

before. For an unshielded source, like B in

Figure 5.47, rays over a much larger

propagation angle range can reach the farfield. For the example of source B in Figure

5.47, the angle between the limiting rays is

greater than 180 ° and is limited by only total

internal reflection angles. Thus, a higher

percentage of the total acoustic power of the

source is radiated in the far field. Hence, if

this source is dominated by mid-to-high

range frequencies then there will be an

increase in noise in that spectral range in the

far-field. With many such sources being

"unshielded" a higher percentage of their

total acoustic power will be radiated to the

far-field. However, the total far-field noise

will go up only if the strength of these

sources remains the same after reducing the

mixing-length. If their strengths can

decrease then a competing mechanism

which can lead to a local minimum in noise

has been identified. Let's explore that

possibility.

When the nozzle length is reduced, the

mixing length, L, is reduced but the nozzle

diameter D is kept the same. This leads to

"squeezing" of the flow in a shorter distance

which would increase the acceleration in a

subsonic flow more than before. But, more

importantly, the "pressure release" condition

to the ambient pressure applied at the nozzle

lip on this squeezed flow now occurs earlier.

This "squeeze and early release" mechanism

now exposes axial vortex flow structures to

the ambient pressure which would have been

otherwise in a bounded duct-flow

environment, as also explained earlier in

Section 4.1.4. What the previous SPL data

(§ 5.5.2) suggests is that when this happens

the noise source intensity and, hence, the

turbulent intensity of these "exposed" axial

vortex flow structures can decrease. These

same "exposed" weaker axial-vortex flow

structures convect downstream, so that now,

with reduced mixing length, there are

weaker noise sources further downstream at

the same axial distance from the mixer-exit

plane than when the nozzle length is longer.

This we will call as the "aerodynamic

unshielding" effect. In order for the total

noise to decrease, as observed for the 75%L

cases, the weakening of these noise

generators must exceed the increase in noise

due to the "acoustic unshielding"

mechanism discussed earlier by referring to

Figure 5.45. We emphasize that this

weakening of axial-vortex flow structures is

only implied by the SPL data; we do not

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have definite plume turbulence intensity

data or spectral data to prove this

hypothesis.

Recall that for 12CL mixer the jet "mean"

flow data in the transverse plane at X/D =

0.2 (ref. 100%L nozzle) and center-line

velocity beyond X/D = 4 (see Figures 4.16,

4.17) did not show much change between

100%L and 50%L cases when compared at

the same distance from the mixer exit plane.

This means that the changes in noise source

strengths speculated here must occur

upstream of X/D = 0.2 (referred to nozzle

exit plane of 100%L), that is, upstream of

X/D = 1 when referred to the nozzle exit

plane of 50%L. This calls for more scrutiny

of the full plume data in future experiments

near the nozzle exit plane and also inside the

nozzle.

The "optimal" length of the nozzle where

the local noise minimum occurs depends on

the turbulence intensity governed by the

strength of these axial vortex structures

which in turn depends on the geometry of

the lobed mixer. Hence, for example, the

optimal mixing-length may vary for the

20UH and 12CL mixers, which produce

streamwise vorticity of different strengths.

The optimal nozzle length can also vary

according to the operating condition, since

this also affects the strength of the axial

vortices, as seen in the noise of 20DH with

differing nozzle-lengths (Figure 5.38). For

lengths less than the optimal length, the

initial vortices which are the strongest (as

seen from the CFD simulations (see Figures

2.22, 2.25)) are no longer acoustically

shielded, allowing the first mechanism to

dominate the second quieting mechanism,

thus producing an increase in overall noise.

How this can produce the observed

minimum in peak SPL frequency with

changes in nozzle length needs to be

scrutinized in the future.

In summary, to explain the existence of a

local noise minimum with decreasing

nozzle-length in lobed mixers two

competing mechanisms are proposed: (i)

"acoustic unshielding" of axial-vortex flow

structures which increases the far-field

radiated noise, and (ii) "aerodynamic

unshielding" which decreases the

turbulence intensity of the "exposed" axialvortex flow structures due to earlier pressure

release to the ambient conditions which

decreases the far-field noise. The second

mechanism especially needs to be verified.

5.6 Flyover EPNL

Effective perceived noise level or EPNL is

an integral measure of noise observed at a

stationary location over a finite time-interval

as an aircraft flies by. We obtain EPNL in

this section for a constant altitude flyover at

1500 ft flight-profile assuming a nonreflecting ground. EPNL is found from the

PNL-data presented in the previous section

after correcting it for "tones" and "duration"

as is implied in the definition of EPNL _z4).

With a single acoustic metric in hand, it

becomes easier to compare the acoustic

performance of one mixer to another over a

range of operating conditions.

We first examine the effect of free-jet Mach

number (or flight effect) on EPNL for a

given mixer and see if it collapses with any

particular choice of the horizontal

coordinate in the EPNL plots. Having

observed such a collapse with a particular

coordinate, we then compare EPNL of one

mixer against the other at different operating

conditions. This comparison brings out the

"relative noise merit" of each mixer and tells

us the acoustic "winners" and "losers"

amongst the mixers tested at different thrust

levels or jet speeds.

NASA/CR 2002-210823/VOL1 208
5.6.1 Effect of Free-Jet Mach Number

EPNL can be plotted against a number of

variables representing the operating

condition. For example, the operating

condition can be represented by the gross

thrust, net thrust (which removes the ram

drag), effective mixed velocity, effective

mixed Mach number, or some other

variable. (Here, effective mixed velocity is

defined as the ratio of gross thrust to total

mass-flow rate.) There is no agreement in

the literature on this and different variables

can be suitable for different purposes. For

example, it can be argued that the noise

from different mixers used in nozzles of the

same diameter should be compared at the

same thrust level for application to engines.

In single stream nozzles of different sizes

the jet velocity, which is specific thrust or

thrust per unit mass-flow rate, is often used

for noise comparison leading to Lighthill's

well-known eighth-power law of noise in

subsonic jets, for example. Effective mixed

jet velocity, on the other hand, may not be

the right choice for the horizontal coordinate

for two-stream nozzles since at a fixed

mixed velocity, different mixers can have

different total mass-flow rates which would

imply that their thrust would be different.

Comparison of noise of two different mixernozzle configurations of same size but

producing different thrusts does not seem

appropriate to us.

5.6.1.1 Effect of Observable Angular

Limits on EPNL

In this section, we try several horizontal

coordinates and look for a collapse of the

EPNL-data when the free-jet Mach number

is varied. Before presenting such plots,

however, it is important to examine the

errors incurred in calculating EPNL from the

current data set. As an example of a

potential source of error, the limit on

observation angles may not allow PNL to

reach a full 10 dB below the peak for some

of the mixer configurations as required in

the definition of EPNL (24_.

Figure 5.48 shows a typical plot of PNL

versus Receiver Time for all the mixers at

operating condition TO #3 and MO = 0.2

based on the available emission angles of

55 ° to 165 ° . The stark contrast in the noise

signatures between confluent mixer (CONF)

and forced lobe mixers is brought out in this

Figure. The confluent mixer (CONF) has a

long duration correction because its noise

dies out very gradually and it has a relative

low peak PNLT. All the lobed mixers show

a sharp rise and a sharp fall with, typically, a

higher peak PNLT (with the exception of

20DH). EPNL calculations need PNLT data

10 PNdB below the peak PNLT. We see

that the data for the CONF mixer, with a

peak PNLT of 86.8 PNdB does not reach the

10 dB down points at either end of the

"event" (which translates back to upstreamand downstream-angles). The PNLT time

history for the confluent mixer is much

flatter (lower rate of change) at the later

times after the aircraft has passed overhead

than in the initial interval before the aircraft

arrives at the observer location. As a result,

failure to reach the 10 dB down point during

the receding portion of the flyover will

produce a larger EPNL error than a similar

occurrence during the initial interval. On

the other hand, since the noise of lobed

mixers falls sharply on the downstream side

of the peak, sufficient data exists to reach 10

PNdB below peak PNLT. However, we

lack some data in the beginning or front

quadrant for these lobed mixers; but, since

the PNLT rises sharply the error incurred by

neglecting data below 55 ° will be small

compared to that for CONF.

An estimate of the "relative" error in EPNL

between two mixers due to the neglection of

NASA/CR 2002-210823/VOL1 209
rn

Z

D.

93

91

89

87

85

83

81

79

77

75

Figure 5.48 1500' Flyover PNLT vs Receiver Time

T.O. # 3 and M(fj) = 0.2

CONF (89.2 EPNdB)

12CL (88.1 EPNdB)

12UH (90.1 EPNdB)

12TH (90.7 EPNdB)

16UH (88.1 EPNdB)

--o--20UH (89.1 EPNdB)

20MH (88.3 EPNdB)

--20DH (86.0 EPNdB)

0 10 2O 3O

Receiver Time (sec)

4O 50

NASA/CR 2002-210823/VOLl 210
PNLT-data beyond the observed angles can

be made as follows. Using the definition of

EPNL (24) , we can show that the relative

error between two mixers (denoted by

subscripts 1 and 2 below) is:

exa2/A2

where A = area under 10 PNLT/10 versus

receiver time curve for available data, and

AA = additional area needed to reach 10

PNdB below peak PNLT.

Note that the y-coordinate of the curve

under which the area is being found is

10 PNm'/l° and not PNLT itself. Due to the

appearance of PNLT in the exponent, the

peak PNLT value contributes the most to A

and much lower values of PNLT than the

peak will have a much smaller effect on A.

In any case, if AA1/A1 --- ZXA2/A2 then there

is no relative EPNL error and AEPNL = 0.

However, as an example, if zXAI/A1 = 5.0%

and AA2/A2 = 2.5%, then ZkEPNL = 0.1 dB.

Due to the flatness of the PNLT curve for

CONF in the final stages of the flyover

event we expect its EPNL error to be higher

than for the lobed mixers. This error

analysis shows, however, that the "relative"

error in EPNL between two mixers due to

lack of acoustic data at extreme angles will

be very small indeed.

Next let us consider the error made in one of

the possible candidates for the horizontal

coordinate, namely thrust, by using unmixed

thrust or separate flow thrust versus the

actual thrust. All the acoustic data was

collected in the NASA APL's NATR but

simultaneous thrust measurements were not

done because the thrust balance was not

working properly. However, accurate

individual mass-flow rates for the fan and

core flow allow unmixed thrust to be

calculated. The gross thrust coefficient Ct_,

defined as the ratio of actual thrust to the

unmixed thrust, gives a good measure of this

error. As we saw in the previous chapter

(see Table 4.1), for the lobed mixers tested

at the static thrust-stand in FluiDyne, the

deviation of actual thrust from the unmixed

thrust captured in ACtr loss = (1-Ctr) is

generally less than 1% at TO conditions. For

higher nozzle pressure ratios and hot total

temperature ratios the Cu loss decreases to

less than 0.5%. Moreover the "relative"

error in thrust between two mixers due to the

use of unmixed thrust instead of actual

thrust is even smaller than this because Cu's

for all mixers are closer to each other than

they are to 1.0. This means that using

unmixed thrust for EPNL plots instead of

actual thrust values will not create

significant errors when examining a range of

thrusts.

In summary, our error analysis shows that

the "relative" error in EPNL due to limits on

observable angles is small and the "relative"

error in thrusts made by using unmixed

thrust values instead of actual values in the

EPNL plots is also very small.

5.6.1.2 Collapse of EPNL Plots

We examine the change in EPNL, across a

whole range of operating conditions, due to

changes in free-jet Mach number by plotting

it against four variables:

(i) Normalized Separate Flow Gross

Thrust - This is calculated from

_Yr_ + mVcia___r¢/P_J

where me, mc are the observed massflow rates for the fan and core

passages at test-day ambient pressure,

Pamb; velocities Vf ideal, Vc ideal are

NASA/CR 2002-210823/VOL1 211
(ii)

(iii)

calculatedfor fan and core streams

when each expands isentropically/

ideallyandseparately(unmixed)to the

ambientpressurefromtheknowntotal

pressureand total temperature.The

last factoris usedto "normalize"the

thrustto removeday to day variation

in Pamb(this formula can be easily

provedfrom thedefinitionof thethrust

coefficient,Ctr,which itself doesnot

varywithPamb).

NormalizedSeparateFlowNet Thrust

- Thisisdefinedas:

Normalized Separate Flow Gross

Thrust - ( my + me) V£ [Pre/Pamb]

where V 0 = free-jet speed. That is, for

net thrust calculations we remove the

"ram drag" from the gross thrust.

Effective Mixed Velocity, Vmix,

defined as the ratio of separate flow

gross thrust to the observed total mass

flow rate. This is the same as the

mass-averaged separate flow velocity:

m£

Vmix- ms V_eeal + Vcideat

my + rn c ms + m c

(iv) Vmix/Camb, a non-dimensional variable

to remove day-to-day ambient

temperature effects.

Figures 5.49, 5.50 and 5.51 display EPNL as

a function of the above four variables across

the whole range of operating conditions

(including approach and all take-offs) with

Mfj = 0.2 and 0.3 for, respectively, mixers

12CL, 12UH and 12TH with the baseline

nozzle length. Only when EPNL is

displayed as a function of net thrust, the

variation with free jet Mach number

collapses into a single curve. This behavior

is not observed when any of the other

parameters are selected as the independent

variable. This is a significant new

discovery. It is expedient to examine this

behavior for all other mixers.

Figures 5.52(a) and (b) shows the EPNL

versus net thrust plots for all the mixers

studied with the baseline nozzle and Mr] =

0.2 and 0.3. Excellent collapse is observed

for 12UH, 12TH, 16UH, 20UH and 20MH.

However, the collapse does not appear as

good for 20DH and at the low thrust

conditions for 12CL. In addition, the EPNL

curves for the CONF mixer are close to each

other for the two free-jet Mach numbers but

the collapse is not nearly as good as for the

lobed mixers.

A couple of remarks are in order here:

(i) In constructing the previous plots,

EPNL is assumed to vary smoothly

and uniquely with net thrust. This

implies that interpolation between

calculated points is acceptable.

However, the operating condition

which will produce the required thrust

at some intermediate interpolated point

is not well defined. That is to say,

although the thrust is known, the

operating condition which will

produce this thrust is not unique. One

way to minimize this problem is to

obtain noise data at closely spaced

points on the operating line of the

engine/nozzle.

(ii) The normalized separate flow net

thrust can be further "nondimensionalized" by (mf+mc)Camb

which essentially gives the "relative"

mixed Mach number defined as

Mmix relative -- (Vrnix- Vfj)/Camb

NASA/CR 2002-210823/VOL1 212
NASA/C_2002-210823/VOL 1 213
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NASA/C_2002-210823/VOL 1 215
9O

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Confluent Mixer

(CONF)

o

--e-- M(rj)=0.2 /

_M(_) = 0.3

CoO , ' ' I ' ' ' I , ' , I ' ' ' I ' ' '

0 2000 4000 6000 8000 10000

Normalized Ideal Net Thrust (Ibf)

go

85

8O

75

70

65

12 Lobe Unscalloped Mixer

(12UH)

=o12I /

60 ' ' ' I ' ' ' I ' ' ' I ' ' ' I' ' ' '

0 2000 4000 6000 8000 10000

Normalized Ideal Net Thrust (Ibl)

12 Lobe Mixer with Cut-Outs

9o (12CL)

--b--M(f]) = 0.2 I /

85- = •

80-

70-

65.

60 , , , J , ' , I ' , ' I ' ' ' I ' ' '

0 2000 40O0 600O 8O00 10000

Normalized Ideal Net Thrust (Ibf)

9O

85

8O

7O

65

6O

12 Pair Tongue Mixer

(12TH)

' ' ' I T , , I ' ' ' I ' ' ' I ' ' '

2000 4000 6000 8000 10000

Normalized Ideal Net Thrust (Ibl}

Figure 5.52(a) Collapse ofEPNL vs Normalized Net Ideal Thrust data with free-jet Mach numbers.

NASA/C_2002-210823/VOL 1 216
16 Lobe Unscalloped Mixer

(16UH)

I /'0

M(t'j)= 0.3

, _ r I ' ' ' I ' _ ' [ r , , [ ' ' '

2000 4000 6000 8000 10000

Normalized Ideal Net Thrust (Ibf)

20 Lobe Moderately Scalloped

Mixer (20MH)

--41,-- M(_) = 0.2 ,_

I t E I

0 2000 4000 6000 8000 10000

Normalized Ideal Net Thrust (Ibf)

20 Lobe Unscalloped Mixer

(20UH)

-'4k'- M(_) = (_-2 1 /

M(rj) = 0.3

' ' = J = ' ' I ' ' ' i ' ' ' : ' ' '

2000 4000 6000 8000 10000

Normalized Ideal Net Thrust (Ibf)

20 Lobe Deeply Scalloped Mixer

(20DH)

_0

--4k--- M(fj) = 0.2 I

[]

6( , , , I , , , I , ' ' I , , , I , , ,

0 2000 4000 8000 8000 10000

Normalized Ideal Net Thrust (Ibf)

Figure 5.52(b) Collapse ofEPNL vs Normalized Net Ideal Thrust data with free-jet Mach numbers.

NASA/C_2002-210823/VOL 1 217
The EPNL variation with free jet Mach

numbers can also be expected to

collapse with this Mmix retative similar to

the collapse of EPNL versus net thrust

plots. However, for nozzles of same

size the comparison between mixers

for their noise merit should still be

done on equal net thrust basis rather

than the relative mixed Mach number.

The collapse of EPNL with net thrust for

varying free-jet speeds can be used to

advantage. Let us assume that EPNL has

been obtained at various operating

conditions with known total mass-flow rates,

and one free-jet speed, V 0. Then we can

find the EPNL at a slightly different free-jet

speed VO + AVfj, at the same operating

conditions as before, as shown in Figure

5.53 due to the invariance of the EPNL with

net thrust. The scaling of EPNL with net

thrust implies we only need to know the net

thrust at the new free-jet speed and the same

operating condition. Since the free-jet speed

does not significantly alter the total massflow rate at a given operating condition (see

Figure 4.5 for BPR, for example), the

difference in net thrust at the reference Vq

condition and at the new free-jet speed, Vnew

=Vt] + AVfj , is approximately [ mtotalAV0 ].

Since we are using pressure normalized

quantities on the net thrust axis, we assume

that a factor of (Pref/Pamb) is absorbed while

calculating mtotal above from the known total

mass-flow rate at V 0 and Pamb. Once the net

thrust has been calculated at the new free-jet

speed then the corresponding EPNL can be

quickly read off the EPNL-axis as shown in

Figure 5.53. If AVq = Vnew - Vref is positive

then the new net thrust is less than the

reference free-jet case and we move to the

left of the reference thrust; if AVfj > 0 then

we move to the right. This shows that, at

least, for those mixers for which collapse

has been observed we do not need acoustic

data at multiple free-jet speeds; only one

free-jet speed will suffice; noise at the other

free-jet speeds can be deduced. This will

reduce the test-matrix considerably for these

mixers, as well as for other similar mixers

which are expected to have same behavior.

As an example of application of this method

consider the case of 20UH mixer with

known results of 1500 ft flyover EPNL

versus net thrust for MO = 0.2 (see Figure

5.52(b) for 20UH). At the four operating

conditions (see Table 3.1) for which noise

data has been collected the total mass-flow

rates at M0 = 0.2 are also known. Now let us

say we want to find the 1500 ft flyover

EPNL at a different free-jet Mach number at

TO #3 only by using this curve for Mq = 0.2.

First to verify this method let us find the

EPNL at Mfj = 0.3 for which the actual

EPNL is known and then we find it for M0 =

0.25 for which there is no data.

At TO #3, MO = 0.2 we have Tamb "" 495.2

°R, Pamb = 14.48 psia and mtotal = 21.69 lb/s

from the scaled nozzle data. This

corresponds to VO = 218.1 ft/s and

normalized unmixed net thrust of 9192 lb.

Hence, at TO #3 but Mr? = 0.3, we get Vq =

327.2 ft/s at same Tamb -- 495.2 °R and the

normalized net thrust difference from that at

M 0 = 0.2 at full scale (scale factor of 4) as

mtotal* AV_(Pref/Pamb)=(4 * 21.69/32.2)*(327.2

- 218.1)*(14.7/14.48) = 1193.7 lb. This

means a normalized net thrust of (9192 -

1194) = 7998 lb at Mr5 = 0.3 and TO #3.

Reading the EPNL for this net thrust from

the EPNL plot in Figure 5.52(b), we get 86

EPNdB approximately. The actual data for

TO #3 at Mq = 0.3 gives EPNL of 86.02

EPNdB. Thus the method gives fairly

accurate results.

Now let us find EPNL at Mr1 = 0.25 for the

same TO #3 condition. Going through the

same process as above, we get for M0= 0.25:

normalized net thrust difference = 597 lb for

NASA/CR 2002-210823/VOL1 218
EPNL

EPNL

@ Vo + AVo

"Static" EPNL .2

llltotal*Vf j

dltotal*AV q

>

Normalized Ideal Net Thrust

Figure 5.53 Use of collapsed EPNL plots.

which net thrust = 8595 lb and EPNL is read

off as 87.5 EPNdB. This process of reading

off from the EPNL versus net thrust curve

for a given net thrust can be automated by

using a curve-fit for the EPNL versus net

thrust plot at the reference free-jet Mach

number rather than the graphical method

used here for demonstration purposes.

Another concept that we introduce here due

to this scaling of EPNL with net thrust is the

idea of an equivalent "static" EPNL, as also

shown on Figure 5.53. This is simply the

value obtained from the EPNL versus net

thrust curve using gross thrust value at that

operating condition as opposed to net thrust.

The corresponding "static" EPNL is then

read off from the y-axis on this curve. Note

that calculation of EPNL always needs a

non-zero aircraft speed or non-zero free-jet

speed - that is, the aircraft needs to move

relative to the observer for PNLT to fall by

10 PNdB, leading to a finite EPNL number.

If the aircraft did not move the usual

definition of EPNL will, of course, lead to

an infinite value of EPNL. So this new

concept of "static EPNL" is, at best, an

artifice and is to be considered as some sort

of datum for a given operating condition for

examining the flight effect and no more. It

would be difficult to give any physical

interpretation for EPNL values

corresponding to net thrust greater than the

gross thrust. This also implies that such

EPNL inferences by using this extrapolation

method must be limited to some small finite

interval of free-jet Mach numbers around the

reference Mt_.

5.6.2 Effect of Nozzle-Length

Figures 5.54 to 5.58A show the effect of

nozzle length on EPNL for free-jet Mach

NASA/C_2002-210823/VOL 1 219
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Figure 5.58A. Flyover EPNL vs nozzle length at various net thrust levels

(shown in legend) for 20UH, 20DH and 12Th mixers at M(fj) = 0.2.

NASA/CRy2002-210823/VOL 1 225
numbers of Mrj = 0.2 and 0.3 for mixer

configurations 12CL, 12UH, 12TH, 20UH

and 20DH for the three TO conditions.

,, For configuration 12CL a noise benefit

of more than 1 EPNdB is observed at the

highest thrust conditions and Mt] = 0.2

by reducing the mixing-length by half

(Figure 5.54); at Mtj = 0.3 a benefit of

more than 1.5 EPNdB is obtained at

mid-thrust conditions by reducing the

mixing-length in hall

• 12UH (Figure 5.55), on the other hand,

becomes considerably noisier when its

mixing-length is reduced by half for both

M 0 = 0.2 and 0.3.

• For the tongue mixer, 12TH, reducing

the mixing length from 100% to 75%

produces little less than 1 EPNdB noise

reduction (see Figures 5.56 and 5.58A)

at the highest thrust conditions for freejet Mach numbers of My = 0.2 and 0.3.

At the lower thrust conditions reducing

the nozzle length from the baseline

(100%) length produced a noise increase

for all free-jet conditions. Since the

observed differences in EPNL are

significantly larger than the error

associated with the EPNL calculation

method, the trends are deemed to be real.

• The data of Figures 5.57 and 5.58A for

configuration 20UH clearly captures a

local minimum in noise with a change in

the mixing length at all thrust conditions

and both Mfj = 0.2 and 0.3. This was, of

course, suspected from the PNLdirectivities and SPL-spectra discussed

earlier where the competing mechanisms

which create this phenomenon were

provided. Note, for example, that 20Utt

benefits anywhere from 1 to 2 EPNdB

by reducing its mixing-length by 25%

from the baseline length at Mg = 0.2. At

the higher Mfj = 0.3, the benefit is even

higher - about 2.2 to 3 EPNdB. This is a

significant reduction in noise and it is

obtained by a reduction in nozzle-length.

Due to the excellent static cruise thrust

performance of the unscalloped, highpenetration mixer 20UH (see Figure

4.2(b)) this is a phenomenon worthy of

attention and future scrutiny.

Figures 5.58 and 5.58A for mixer

configuration 20DH again clearly

illustrate the existence of a local noise

minimum with decreasing mixinglength, especially at the higher thrust

end. At M_ = 0.2 the decrease in noise

is as much as 1 EPNdB and at Mfi = 0.3

it is as much as 1.2 EPNdB when the

mixing-length is decreased by 25% from

the baseline length.

5.6.3 Comparison of EPNL Between

Mixers

To sort out the effect of individual lobe

mixer geometric parameters, the EPNL net

thrust variation is displayed for each group

of mixers, with the confluent configuration,

CONF, as the baseline. It is important to

recall that although the comparison is made

on an equal net thrust basis, some of the

mixers have very different area ratios (see

Tables 2.1, 2.2). As a result, there is a

significant bypass ratio variation between

the various mixers, even at constant net

thrusts as mentioned earlier (see Figures 4.6

and 4.7).

Figure 5.59 shows the EPNL plot for the 12-

lobed mixers at M_ = 0.2 and two different

nozzle-lengths, 100% and 50%. As

expected from the PNL-directivity curves

discussed earlier, the 12CL mixer is the

quietest, with I2UH second and 12TH being

the noisiest. Only slight relief is obtained

for the tongue mixer with the 50% L nozzle

at the highest thrust condition as compared

to 12UH. However, for both nozzle-lengths

12CL is the only configuration quieter than

the baseline CONF mixer; but even that

NASA/CR 2002-210823/VOL1 226
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NASA/C_2002-210823/VOL 1 227
effect appears to die out at high thrusts for

100% L. The maximum benefit of 12CL

compared to CONF is about 1.8 EPNdB at

TO #2 (at about 7000 lb net thrust) with the

short (50%) nozzle. It is important to

compare 12UH and 12CL on an equal net

thrust basis because of certain advantages of

no-scalloping in 12UH: With the 100%L

nozzle-length 12CL is nearly 3.75 EPNdB

quieter than 12UH near the reference TO #1

condition which corresponds to 5000 lb net

thrust. This difference jumps to 5.4 EPNdB

when the comparison is made with the 50%

nozzle-length at the same net thrust. The

noise benefit of 12CL compared to 12UH

decreases slightly as net thrust increases for

the baseline 100% nozzle-length. With a

higher cruise thrust coefficient for a

"cleaner" version of 12CL than 12UH (see

table 4.1) and such excellent noise

suppression, the 12-lobed low penetration

mixer with cutouts (12CL) is clearly

superior to the high-penetration unscalloped

mixer (12UH).

Figure 5.60 shows the effect of number of

lobes for the second group of mixers (12UH,

16UH, 20UH). Comparing 12UH versus

20UH, the mixer with the higher number of

lobes (20UH) is quieter than 12UH at 100%

nozzle-length but the noise benefit decreases

considerably as the thrust increases. It is

significant to note that 20UH has a slightly

higher cruise thrust coefficient than 12UH

and it is also quieter at the baseline TO #1

condition by 1.2 EPNdB. However, it is

equally significant to note that none of these

lobed mixers, all of them "unscalloped,"

were quieter than CONF mixer either at

100%L or 50%L. Even 16UH, with its very

high bypass ratio is noisier than CONF at all

thrusts. Their unscalloped lobes with high

penetration are the controlling negative

factors from the viewpoint of noise

suppression.

Figure 5.61 shows the effect of scalloping

on EPNL by comparing mixer

configurations 20UH, 20MH and 20DH. At

100% nozzle-length, deep scalloping

produces a benefit that ranges from about

1.7 EPNdB at low thrust (5000 lb) to 2,8

EPNdB at high thrust (9200 lb) compared to

the unscalloped mixer. However, the

considerable loss in cruise thrust coefficient

due to deep scalloping (20DH versus

20UH) of about 0.38% (see Table 4.1)

should also be kept in mind for applications

to engines.

Figure 5.61A shows that compared to CONF

the noise benefit of the deeply scalloped

mixer (20DH) with 100% nozzle-length

increases with thrust, ranging from

negligible improvement at TO #1 to about

1.8 EPNdB at 9000 lb net thrust. At 100% L

the benefit of "moderate" scalloping

(20MH) over unscalloped mixer is also

negligible at low thrusts and improves to

only about 0.8 EPNdB at the highest thrust.

From Figure 5.61A we observe that with

half the nozzle length (50%L) none of the

20-lobed mixers are quieter than CONF, but

moderate scalloping is quieter than both

deeply scalloped and unscalloped mixers at

all thrust levels. This shows that there is an

'.'optimal depth of scalloping" for shorter

nozzles. But for longer nozzles the noise

benefits of scalloping appear monotonic.

Also note from Figure 5.61A that the

unscalloped mixer, 20UH, is quieter than

CONF by approximately 1 EPNdB only

with 75% L at the highest thrust level (9000

lb); otherwise with either the shortest or the

longest nozzle 20UH is noisier than CONF

at all thrust levels. Since 20UH also has the

highest cruise thrust coefficient such a

minimization of its noise by fine tuning the

nozzle length can be used in engine

applications where cruise thrust efficiency

cannot be sacrificed.

NASA/CR 2002-210823/VOL1 228
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Figure 5.61A Effect of scalloping and nozzle length on flyover EPNL benefit

compared to CONF at different net thrust levels and M(fj) = 0.2.

NASA/C_2002-210823/VOL 1 231
Figure 5.62 summarizes the EPNL

comparison of all the mixers with baseline

nozzle-length at Mr5 = 0.2 on one plot. It

clearly shows that out of all the mixers

tested only 12CL and 20DH are quieter than

CONF, with 12CL being more beneficial at

low thrust conditions (TO #1). Both are

equally beneficial at mid-thrust condition

(TO #2) and 20DH is the best noise

suppressor at the highest thrust conditions

(TO # 3).

Figure 5.63 summarizes the EPNL for all

mixers at M 0 = 0.2 with the shortest nozzle

length (50%L), except CONF which has

100%L but whose behavior is not expected

to change with nozzle length, as discussed

earlier. This Figure clearly shows that at

50%L only the 12CL mixer remains quieter

than the CONF configuration throughout.

Figures 5.64 and 5.65 summarize the EPNL

for all mixers at an increased free-jet Mach

number of 0.3 for 100% and 50% nozzle

length respectively. At this high M 0 none of

the lobed mixers is quieter than the simple

confluent nozzle. 12CL produces only a

marginal benefit at 50%L and low thrust

(Figure 5.65) but the differences observed

are within the margin of error for calculating

EPNL. Thus, the high free-jet speed of M S =

0.3 produces so much noise reduction in the

confluent mixer that all the intricacies of

spectral changes and reductions in noise of

lobed mixers is of no use at this scale factor.

5.6.4 Quietest Mixer-Nozzle Combinations

Figures 5.66(a) and (b) show the quietest

mixer-nozzle combinations compared to the

baseline confluent nozzle in terms of EPNL

versus Net thrust and EPNL versus Vmix,

respectively, at a free-jet Mach number of

0.2. Table 5.2 summarizes the noise

benefits in terms of AEPNdB on the basis of

equal net thrust, as well as, equal mixed jet

velocity compared to CONF at three

different thrusts and jet velocities. Note that

the noise benefit at the highest thrust level of

9500 lb for 20DH with 75% L is obtained by

linear extrapolation of EPNL data in Figure

5.66(a). Comparing AEPNdB at same jet

velocity increases it by about 1 EPNdB

(anywhere from 0.8 to 1.1 EPNdl3) over

suppression obtained on the basis of same

net thrust. Although the values of AEPNdB

benefit depend on whether we compare the

mixer-nozzles on the basis of net thrust or

jet velocity, the following trends are

common to both:

0) Only scalloped or cutout mixers (12CL

or 20DH) prove to be the quietest at all

thrust levels tested.

(ii) Noise suppression increases with

thrust or jet velocity. At low thrusts

the 12-lobe mixer with cutouts (12CL)

produces a 1 EPNdB reduction with

the 50%L nozzle. At high thrust, the

deeply scalloped 20 lobe mixer

(20DH) produces approximately 3.0

EPNdB noise reduction with the 75%L

nozzle. When compared on the basis

of equal jet velocity, these translate

into 1,8 EPNdB benefit at the lower jet

speeds and about 4.0 EPNdB at the

higher jet speeds. Since the EPNL

curves for the confluent nozzle and the

20DH-75%L configuration diverge

from each other at higher thrust levels

the noise benefits will be even more at

higher thrusts or jet-speeds.

(iii) In all cases, the maximum noise

reduction with a forced mixer was

obtained with a nozzle shorter than the

baseline length of L/Drop -" 1.10.

It is worth noting that only these two

quietest mixer configurations, namely, 12CL

and 20DH were also successful at reducing

the low frequencies without too much

increase in the annoying mid-to-high

NASA/CR 2002-210823/VOL1 232
92

Figure 5.62 1500 ff Flyover EPNL vs Ideal Net Thrust Comparison

100% Nozzle Length, M(i]) = 0.2

m

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91

90

89

88

87

86

85

84

83

82

81

80

79

78

77

76

75

74

4500

/

/

/ J _ _ Confluent Mixer (OONF> t - +12 Lobe Mixer with Cutouts (120L)

/Z_ _ 12 LObe unscalloped Mixer (12UH)

12 Pair Tongue Mixer (12TH)

.__ ----lt_ 16 Lobe Unscalloped Mixer (16UH)

-_e---20 Lobe Mixer- Unscalloped (20UH)

20 Lobe Mixer- Mod. Scalloped (20MH)

....... r , , , 1 , , , T--

5500 6500 7500 8500 9500

Normalized Ideal Net Thrust (Ibf)

NASA/CR 2002-210823/VOLl 233
I.g

Figure 5.63 1500 ft Flyover EPNL vs Ideal Net Thrust Comparison

50% Nozzle Length, M(fj) = 0.2

93

92

91

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78 _ I _ 'Confluent Mixer (CONF)

__/,_ _ 12 Lobe Mixer with Cutouts (12CL) 77 fy _ 12 Lobe Unscalloped Mixer (12UH)

#

76 _ 12 Pair Tongue Mixer (12TH)

-_ ,/ _ 16 Lobe Unscalloped Mixer (16UH)

75 _// .--e----20 Lobe Mixer- Unscalloped (20UH)

74 //, _ 2020 Lobe Lobe Mixer- Mod. DeeplyScalloped Scalloped (20MH) (20DH)

73

4500 5500 6500 7500 8500 9500

Normalized Ideal Net Thrust (Ibf)

NASA/C_2002-210823/VOL l 234
88

Figure 5.64 1500 ff Flyover EPNL vs Ideal Net Thrust Comparison

100% Nozzle Length, M(t]) = 0.3

L .L J

i""'_ 'Confluent Mixer (CONF) _

m_- 12 Lobe Mixer with Cutouts (12CL)

12 Lobe Unscalloped Mixer (12UH)

-- _ 12 Pair Tongue Mixer (12TH) p

+ 16 Lobe Unscalloped Mixer (16UH) _.

-- "me--- 20 Lobe Mixer-Unscalloped (20UH) / __

20 Lobe Mixer- Mod. Scalloped (20MH) / /

-- 20 Lobe Mixer - Deeply Scalloped (20D/ ,//// //#_,

. ,'2////_

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'/// /

i t i

3000 4000 5000 6000 7000 8000

Normalized Ideal Net Thrust (Ibf)

NASA/CR 2002-210823/VOLl 235
89

Figure 5.65 1500 ft Flyover EPNL vs Net Ideal Thrust Comparison

50% Nozzle Length, M(fj) = 0.3

"o

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87

86

85

84

83

82

81

80

79

78

77

76

75

74

73

72

71

70

•---e, 'Confluent Mixer (CONF)

----=--- 12 Lobe Mixer with Cutouts (12CL)

12 Lobe Unscalloped Mixer (12UH)

12 Pair Tongue Mixer (12TH)

16 Lobe Unscalloped Mixer (16UH)

---e---20 Lobe Mixer- Unscalloped (20UH)

-_+---20 Lobe Mixer- Mod. Scalloped (20MH)

20 Lobe Mixer- Deeply Scalloped (20DH)

69

3500 4500 5500 6500 7500 8500

Normalized Ideal Net Thrust (Ibf)

NASA/C_2002-210823/VOL l 236
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NASA/C_2002-210823/VOL 1 237
Table 5.2 Quietest Mixer-Nozzle Combinations at Mfj = 0.2

Normalized Net Mixer Nozzle

Thrust (lb) Length (L)

5000 12CL 50%

7000 12CL 50%

9500# 20DH 75%

zkEPNdB

Benefit*

1,0

1.8

3.0

Wmix

fit/s)

830

950

lO6O

Mixer

0.76 12CL

0.87 12CL

0.97 20DH

Nozzle AEPNdB

Length (L) Benefit*

50% 1.8

50% 2.9

75% 4.0

* AEPNdB Benefit = EPNL(Confluent) - EPNL(Mixer)

#After linear extrapolation of EPNL plot for 20UH+75%L

frequency noise compared to CONF.

Hence, we conclude that at lower net thrusts

a forced mixer using scallops/cutouts

combined with low lobe penetration and

fewer numbers of lobes will produce a noise

reduction. At high net thrust, higher

numbers of lobes and lobe scalloping helps.

However, the noise benefits must be

balanced against thrust losses at cruise

conditions. In that regards, the fine tuning of

the nozzle length for the most thrust

efficient unscalloped lobe mixer 20UH to

reduce noise below CONF should also be

borne in mind although 20UH is noisier than

20DH or 12CL.

As a final note, it is appropriate to point out

that the results presented in the report are

somewhat scale sensitive. This is directly

related to the frequency-weighting noy

function applied to the SPL spectra in the

definition of PNL. As the scale factor is

increased or decreased from the value of 4

used throughout this report, key spectral

features will translate to new frequency

bands. Depending on the scale factor used,

the resulting weighting factor applied to

these features may alter their importance at

the new scale. As a result, generalizations

of the conclusions contained in this report to

applications of different size should be

undertaken with caution.

NASA/CR 2002-210823/VOL1 238
Chapter 6

Concluding Remarks

A comprehensive test database for the

acoustic, aerodynamic and plume flow

characteristics of model scale forced lobe

mixers has been created for nozzles with

bypass ratio of 5 to 6. The operating

conditions for acoustic tests ranged from

typical approach through take-off points

with effective jet speeds from 600 ft/s to

1080 ft/s. Forward flight effects were

simulated through the use of a separate free

jet flow, with data obtained at free jet Mach

numbers between 0.0 and 0.3. The static

thrust performance was measured at cruise,

as well as, at most of the take-off conditions.

The data was scaled to a nozzle exit

diameter of 29 inches (scale factor of 4),

representative of small-to-medium thrust

turbofan engines. The effect of several

mixer and nozzle parameters was analyzed:

scalloping/cutouts in lobe side-walls,

number of lobes, mixing length, and also

lobe penetration and fan-to-core area ratio.

Several new concepts, methods and

mechanisms were identified during the

conduct of this project which helped us

explain some of the observed results and

these ideas can be applied to the design of

jet noise suppressors.

We summarize below some of the new

concepts and important findings:

New Concepts/Methods

1. Gradual introduction of vorticity and

the "trailing-edge" rule for scalloping

leading to "boomerang" scallops to

reduce internal fan/core mixing noise.

2. "Tongue" mixer - a unique limiting

case of deep scalloping.

3. Comparison of noise for different freejet Mach numbers at "shifted" angles

to examine "excess" internal noise

from fan-core mixing.

4. Extrapolation of flyover EPNL for

lobed mixers from data known at one

free-jet speed to another speed.

New Generic Findings

.

,

Horizontal flyover EPNL for different

aircraft speeds was found to correlate

with "net" thrust and not gross thrust

or absolute jet speed. The collapse of

the EPNL versus net thrust curves for

different aircraft speeds was found to

be good for most lobed mixers but less

impressive for the confluent

configuration.

Using the "shifted" angle comparison

method for several free-jet speeds, it

was found that the annoying mid-tohigh frequencies upstream of the peak

PNL angle were the result of "excess"

noise sources which could be either

"internal" to the nozzle duct or

"external" to it but close to its exit

plane. This was particularly notable in

the unscalloped lobe mixers and the

tongue mixer, and was less important

in the two scalloped mixers which

were also the quietest. The annoying

excess noise spectra was deemed to be

produced by the turbulence in the

strong streamwise vortices from the

forced mixers.

NASA/CR 2002-210823/VOL1 239
Important Specific Findings

.

.

.

Forced lobe mixers reduce noise in

comparison to a coaxial jet by

significantly reducing the low

frequency contribution but, if not

designed properly, such mixers can

lead to an increase in the mid-to-high

frequency sound which is heavily

weighted for annoyance.

Sudden introduction of streamwise

vorticity inside the nozzle as with

unscalloped, high-penetration lobemixers is not a desirable feature from

the point of view of noise suppression

because of increases in the annoying

spectral component which nullifies any

benefit accrued by decreasing the low

frequency spectra. On the other hand,

gradual introduction of axial vorticity,

as with "boomerang" scallops, is

preferable on an overall basis even

though it may not benefit the lower

frequencies as much.

The most prominent geometric feature

of the lobed mixers for reducing the

annoying portion of the spectral band

is "scalloping" of the lobes, but lobepenetration and lobe count also have a

role. For the size of engine considered

(29 in. diameter), a scalloped lobe

mixer with fewer lobes and lower

penetration is the most effective mixer

at low net thrusts (5000 lb), producing

a maximum benefit of 1 EPNdB. At

higher net thrusts (9500 lb) a deeply

scalloped lobe mixer with higher

number of lobes and higher lobe

penetration is the most effective mixer,

producing a maximum noise benefit of

about 3 EPNdB. In all cases, noise

reduction is measured against a coaxial

or confluent configuration. These

noise benefits increase by about one

more EPNdB when compared on the

basis of equal mixed jet velocity rather

.

°

,

than equal net thrust.

For a given lobe mixer, an optimal

mixing length was found to exist from

the point of view of noise suppression.

That is, decreasing the mixing length

(or nozzle length) from the baseline

length of L/D,_p = 1.1 also decreased

the noise up to a local minimum value

and any further decrease in nozzle

length increased the noise. For

unscalloped mixers such a decrease in

mixing length gave a noise benefit of

as much as 3 EPNdB at high nozzle

pressure ratios. A working hypothesis

in terms of two competing mechanisms

is proposed, and a deeper look at this

new and interesting phenomenon is

required. It is obviously beneficial

from the viewpoint of weight but its

repercussions on the external boat-tail

angle and consequent cruise thrust

performance must also be considered.

At low to medium thrust levels,

unscalloped lobe mixers with high lobe

penetration do not reduce noise below

that of a coaxial jet regardless of the

number of lobes and nozzle-length.

However, at higher thrusts fine tuning

of nozzle length for a 20 lobe

unscalloped mixer produced a 1

EPNdB benefit. The cruise thrust

coefficient and the thrust-mixing

efficiency of such unscalloped mixers

are typically higher than scalloped lobe

mixers. Hence, in applications where

thrust efficiency is of more importance

than noise suppression minimizing

noise of unscalloped lobe mixers by

adjusting the nozzle length may be a

good alternative.

It should be noted, however, that the

12-lobe cutout mixer with low

penetration has a good noise benefit

throughout the jet-speeds tested, as

well as, one of the highest cruise thrust

performance. Scalloped lobe mixers, in

NASA/C_2002-210823/VOL 1 240
o

general, still have far better cruise

thrust coefficients and thrust mixing

efficiencies than a coaxial jet and are

the quietest mixers. For certain nozzle

lengths the moderate scalloping was

better than unscalloped or deeply

scalloped mixers. Thus the amount of

scalloping can be used as a trading

parameter between noise suppression

and cruise thrust loss.

The tongue mixer model, designed for

proof-of-concept in this task, provided

the maximum suppression of the low

frequency spectra amongst all the

mixers, but this was accompanied by

an unacceptable increase in the

annoying mid-to-high frequency noise

making it one of the noisiest mixers. It

also produced low thrust coefficients.

The primary reason for this behavior

is its smaller fan-to-core area ratio and,

hence, smaller bypass ratio. However,

its simplicity of manufacturing,

reduced weight and ease of reparability

demand a second look at this concept

after appropriately sizing it with a

larger fan-to-core area ratio.

This tongue mixer concept has also

been extended by Rolls-Royce Allison

for use in externally mixed or separate

flow nozzles under a separate NASA

Contract NAS3-27720 to General

Electric Company (2 5), and it has

shown noise benefits of about 2

EPNdB over coaxial jets when applied

to larger engines with higher jet speeds

(see figure 87, Ref. 25). In this regard,

it should be noted that in applications

to real engines the concepts that

produce large decrease in low

frequency noise but moderate increase

in mid-to-high frequency noise may

not be so bad if the latter frequency

range is dominated by other engine

noise sources, such as, fan tones. This

shows that the tongue mixer is, indeed,

a viable concept from noise

suppression point of view and should

be further explored for application to

full-scale engines.

NASA/CR 2002-210823/VOL1 241

Chapter 7

References

1. Packrnan, A. B. and Eiler, D. C. (1977):

Internal Mixer Investigation for JT8D

Engine Jet Noise Reduction, Vol. I -

Results. U. S. Dept. of Transportation,

Federal Aviation Administration Report

No. FAA RD-77-132.1, Dec. 1977.

2. Shumpert, P. K. (1980): An

Experimental Model Investigation of

Turbofan Engine Internal Exhaust Gas

Mixer Configurations. AIAA Paper 80-

0288, AIAA 18th Aerospace Sciences

Mtg., Jan. 14-16, 1980.

3. Kuchar, A. (1980): Scale Model

Performance Test Investigation of

Exhaust System Mixer for an Energy

Efficient Propulsion (E a) System. AIAA

Paper 80-0229, AIAA 18th Aerospace

Sciences Mtg., Jan. 14-16, 1980.

4. Larkin, M. J. and Blatt, J. R. (1984):

Energy Efficient Engine Exhaust Mixer

Model Technology Report Addendum,

Phase 111 Test Program. NASA CR

174799, April 1984.

5. Saiyed, N. H., Bridges, J. E. and Krejsa,

E. A. (1996): Core and Fan Stream's

Mixing Noise outside the Nozzle for

Subsonic Jet Engines with Internal

Mixers'. AIAA Paper 96-1667, AIAA

Aeroacoustics Conference, May 6-8,

1996.

6. Booher, M. E., Kwon, O., Barta, A. B.,

Vittal, B. R. and Krishnan, M. R. (1993):

Development of an Advanced Exhaust

Mixer for a High Bypass Ratio Turbofan

Engine. AIAA Paper 93-2435, 29th

AIAAJASME/S AE/ASEE Joint

Propulsion Conference, June 28-30,

1993.

7. Barta, A. B., Vittal, B. R. and Dalton,

W. N. III (1997): Experimental

Verification of the Mixing Process in

Turbofan Engine Exhaust Systems.

AIAA Paper 97-2640,

AIAA/ASME/S AE/ASEE Joint

Propulsion Conference, July 6-9, 1997.

8. Meade, J. P. (1994): Acoustic Evaluation

of a Forced Mixer on a High Bypass

Ratio Engine. AIAA Paper AIAA-94-

2954, 30th AIAAJASME/SAE/ASEE

Joint Propulsion Conference, June 27-

29, 1994.

9. Elliott, J. K., Manning, T. A., Qiu, Y. J.,

Greitzer, E. M. and Tan, C. S. (1992):

Computational and Experimental

Studies of Flow in Multi-Lobed Forced

Mixers. AIAA Paper 92-3568, 28th

AIAA/ASME/SAE/ASEE Joint

Propulsion Conference, July 6-8, 1992.

10. McCormick, D. C. and Bennett, J. C.

(1993): Vortical and Turbulent Structure

of a Lobed Mixer Free-Shear Layer.

AIAA Paper 93-0219, 31st AIAA

Aerospace Sciences Mtg., Jan. 11-14,

1993.

l l. Belovieh, V. M., Samimy, M. and

Reeder, M. F. (1994): Dual Stream

Axisymmetric Mixing in the Presence of

Axial Vorticity. AIAA Paper 94-3084,

30th AIAA/ASME/SAE /ASEE Joint

Propulsion Conference, June 27-29,

1994.

12. The NPARC Alliance (1996): A User's

Guide to NPARC Version 3.0. Arnold

Engineering Development Center,

Tullahoma, Tennessee, Sept. 1996.

13. Pointwise, Inc.: GRIDGEN User's

Guide., Bedford, Texas 76095.

14. Castner, R. S. (1994): The Nozzle

Acoustic Test Rig - An Acoustic and

Aerodynamic Free-Jet Facility. NASA

NASA/CR 2002-210823/VOL1 243
TM 106495,1994.

15.Ahuja,K. K., Tester,B. J.andTanna,H.

K. (1978):The Free Jet as a Simulator

of Forward Velocity Effects on Jet

Noise. NASA-CR-3056, Oct. 1978, 324

p.

16. Mengle, V. G., Dalton, W. N., Bridges,

J. C. and Boyd, K. C. (1997): Noise

Reduction with Lobed Mixers: NozzleLength and Free-Jet Speed Effects.

AIAA Paper 97-1682, 3rd AIAA/CEAS

Aeroacoustics Conference, May 12-14,

1997.

17. Fisher, M. J., Preston, G. A. and Bryce,

W. D. (1993): A Modelling of the Noise

from Simple Co-axial Jets'. AIAA Paper

93-4413, 15th AIAA Aeroaeoustics

Conference, Oct. 25-27, 1993.

18. Balsa, T. F. and Gliebe, P. R. (1977):

Aerodynamics and Noise of Coaxial Jets'.

AIAA J., Vol. 15, No. 11, Nov. 1977,

pp. 1550-1558.

19. Yu, S. C. M., Yip, T. H. and Liu, C. Y.

(1997): Mixing Characteristics of

Forced Mixers with Scalloped Lobes. J.

Propulsion & Power, Vol. 13, No. 2,

March-April 1997, pp. 305-311.

20. Abramovich, G. N. (1963): The Theory

of Turbulent Jets. M.I.T. Press,

Cambridge, MA, Chapter 5.

21. Candel, S. M. (1972): Acoustic

Transmission and Reflection by a Shear

Discontinuity Separating Hot and Cold

Regions. J. Sound and Vibration, Vol.

24(1), pp. 87-91, 1972.

22. Lighthill, James (1978): Waves in

Fluids. Cambridge University Press,

Cambridge, Great Britain.

23. Thompson, R. J. (197l): Ray Theory for

an lnhomogeneous Moving Medium. J.

of the Acoustical Society of America,

Vol. 51, No. 5 (part 2), 1971, pp. 1675-

1682.

24. Society of Automotive Engineers

(1973): Definitions and procedures for

Computing the Effective Perceived Noise

Level for Flyover Aircraft Noise. ARP

1071.

25. Janardan, B. A., Hoff, G.E., Barter, J,

W., Martens, S., Gliebe, P. R., Mengle,

V.G. and Dalton, W.N. (1998): Separate

Flow Exhaust System Noise Reduction

Concept Evaluation. Final Report,

NAS3-27720, Area of Interest 14.3,

NASA CR (to be numbered), GEAEA

Report No. R98AEB 152, May 1998, 273

p.

NASA/CR 2002-210823/VOL1 244
Appendix A

Geometrical Details of Mixers and Nozzles

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Appendix B

Refraction Correction Procedure for Free-Jet/Ambient Shear-Layer

and Angular Interpolation Scheme

When we have a free-jet facility of finite

diameter surrounding the model jet to

simulate aircraft forward speed effects, a

shear-layer is created between the free-jet

and the ambient quiescent atmosphere in the

test cell. Sound from the model jet then

refracts through this shear-layer before

being recorded by the microphones which

are outside the free-jet in the test cell. This

shear layer is not present in the actual

flyover case. In order to retrieve the original

directivity of sound from the model jet we

then need to account or correct for this

refraction. This correction procedure is

summarized below, as given to Rolls-Royce

Allison by Dr. James Bridges of NASA

Lewis Research Center. NASA has

processed all the acoustic data using this

procedure which forms part of the overall

data processing scheme shown in Figure 3.5.

The free-jet shear-layer refraction correction

is taken from Ahuja et al is. It is similar to

the "Amiet" method with the main

difference being that NASA has generalized

the algorithm to work for narrow-band data

instead of third octave spectra. In this

method, one starts by considering the SPL of

each frequency band (narrow-band) and

each microphone individually. One assumes

a location for the source of this sound and

the location of the shear-layer based on the

geometry and flow-conditions of the free-jet

and the model. One also assumes that the

shear-layer is infinitely thin, and use is made

of acoustic ray theory, an assumption that

was verified in Ahuja et a115 for frequencies

generated with a relatively small speaker.

From this geometry the refraction that the

sound underwent to reach the microphone is

calculated and the original direction of the

sound during propagation between model and

free-jet/ambient shear-layer is found. This

will not, in general, correspond to a location

where a microphone exists. In fact, when a

source location model is used, each frequency

for a given microphone will have been

directed to a unique angle because it

originates from a different place. In addition,

the difference between the geometric distance

from the source to the microphone and the

actual path-length that the sound traveled due

to the refraction is calculated. This procedure

of finding the original angle and the refracted

path-length as a function of refracted angle is

done for each frequency. Now if the

attenuation due to spherical spreading over

the refracted path-length is factored out, then

one retrieves the original directivity of the jet

as it would appear inside the free-jet. This is

the same as the directivity in the reference

frame of the nozzle.

As mentioned above, since we can assume

that the effective source location is different

for each frequency band, the directivity is

known at different locations for each

frequency band. To take this back to a

common set of observer locations, the

directivity of each frequency band is

interpolated (in SPL) onto a convenient

regular array, corresponding to an arc with 5

degrees intervals for this test. There are

situations where sound at some frequencies

are refracted strongly enough that none of the

measured sound covers the interpolated array.

NASA/CR 2002-210823/VOL1 269
In this case, the data is set to "unknown" and

is not used in the processing. In addition,

for bands in which the data was not

sufficiently above the background level or

for which the microphone was found to be

suspect in post-test analysis, the data was set

to "unknown". This is generally a problem

for high free-jet Mach numbers, low

frequencies, and shallow angles, which have

little impact on overall evaluation of noise.

Interpolation gives values only for those

angles where sufficient information is

known; data is never extrapolated outside

the measured information. If insufficient

data is available over a region because of

background noise, then the whole region is

taken as "unknown" in the interpolation

scheme rather than filling in values based on

very limited information. Specifically the

algorithm calls for at least as many good

points as unknown points in a region before

the interpolating spline is evaluated for

output. For example, if at the end of the

array of N microphones a measured value

exists at angle N but the two measurements

at N-1 and N-2 were suspect while the

measurement at N-3 was good, SPL values

will be recorded as "unknown" for any

locations which lie at angles greater than

that of microphone N-3.

Interpolation of SPL between angles for a

given frequency band is a required step and

is not used for smoothing. The interpolation

is done by strict cubic spline with zero

derivative knots at the ends. Tests have

shown this choice of interpolating

polynomial to capture, but not overshoot,

peaks very well, as opposed to linear

interpolation which showed some

smoothing. Interpolation can be performed

on either SPL (in dB) or sound pressure (in

Pascals squared). The final result will be

affected by the choice, but again tests found

that for representative spectra, sound power

differences due to interpolation were less than

0.05 dB. Using the log scale makes it easier

for the interpolating function to fit without

spurious excursions, although it biases the

error to the high side.

Because the source distribution is actually

unknown and probably different for each jet,

NASA has adopted as standard procedure the

assumption that the sound sources for all

frequencies is at the center of the nozzle exit

plane. This is highly incorrect for low

frequencies but without better information we

prefer this simple model over an arbitrary

one. This was mostly driven by the need to

compare results with other facilities which

have other shear-layer correction schemes.

Data being compared came from G.E.'s Cell

41, Boeing's LSAF, NASA Ames' 40 x 80

wind-tunnel, NASA Langley's JNL Quiet

Flow Facility, and NASA Lewis' APL. This

source location assumption has been made for

all NASA/Industry programs. However, until

a better model comes up, this will remain an

open issue.

NASA/CR 2002-210823/VOL1 270
Appendix C

Refraction Angle_ Convection Angle and Transmission Coefficient Changes

Due to Free-Jet Speed Changes

We extend Candel's analysis 21 to moving

flows on both sides of a thin shear-layer

represented by a thin straight vortex sheet at

y = 0, x > 0 as shown in figure 5.25. In

region 1, y > 0, the speed of sound is cl and

the flow Mach number is M1. In region 2, y

< 0, the speed of sound is c2 and the Mach

number is M2. The pressure fluctuations are

piexp(-icot), Rprexp(-i03t) and Tptexp(-i03t)

for the incident, reflected and transmitted

waves, respectively, and v(x)exp(-i03t) is the

y-displacement of the vortex-sheet where 03

is frequency, t = time and i = _-1. The

incident wave has a normal direction 001;the

reflected wave propagates in the direction -

00t, and the transmitted wave in the direction

002(see figure 5.25):

=exp( -i klxc°s4_ Pi

1 - M 1cos 41

klx cos

Pr = exl: - i

1 - M 1cos01

kzxcos4 z

Pt = exp - i

1 - M 2 cos42

k 1 = _el,k2 =_c z

0<41 <z.

i klysin41

1 - M 1 cos 41

k1y sin 41 +i

1 -- m 1 cos 41

-i kzysin42

1- M 2 cos4z

All waves have the same wave-number

vector components in the x-direction:

k1 cos 41 k2 COS_ 2

1- M 1COSq_l 1 - M 2 cos 42

This leads to the first useful relation

between angles which denotes pure refraction

effects:

COS42 = (c2//Ccl)COS41 (C.1)

1 - cos41(M 1 - (c2//q)M2)

We obtain 1_20, which is ¢2 for M2 = 0, from

above as:

cosO2o- (%) cos41 (c.2)

1 - M t cos41

Hence, 02 can be written directly in terms of

0020without involving M_ or 00_as:

cos¢2 = cos420 (C.3)

1 + M 2 cos4zo

And, from eqn (C.2), 001 can be found in terms

of 0020as:

COS _20

cos 41 = c_22+ M1 cos 420

Cl

(C.4)

To find the effect of convection of the

refracted wave-fronts on the convected angle

A00c(see figure 5.25) we apply the sine-rule to

the velocity triangle in region 2 which

includes the transmitted ray and the refracted

wave-front normal direction:

U 2 c2

sin A4c sin(zr - (02 + A4c ))

NASA/C_2002-210823/VOL 1 271
This gives, after some manipulation, the

second desired angular relation:

tanA¢c = M 2 sine2 (C.5)

1 - M 2 cos q)2

The numerical results for these angles have

been given in Table 5.1, plotted in figure

5.27 and discussed in the main text. As a

further independent check on these

calculations, we spot-checked it with values

obtained from Thompson's ray equations

(eqn. (12), Ref. 23) where the "generalized

Snell's law of refraction" for moving media

is derived by a wholly different method

using ray-paths obtained from the

characteristic eikonal equations. The values

for the refracted ray angle, qb2, for given

"static" ray-angles, _0 , matched exactly

with the ones given in Table 5.1.

To obtain the transmission coefficient T and

reflection coefficient R, the conditions of

continuity of pressure and particle

displacement at the interface (vortex-sheet)

must be applied:

(Pi + gPr)y=O+ =(WPt)y=O_

-ion+ UI-_)0 2 v(x)=-_l-_(Pi1 0 + RPr)y=O+

- (-io) + U 2 v(x) = O (Tpt)y=0_

02 Oy

where P is density. Manipulation of these

three equations with the previous result from

the equality of x-component of the wavenumber vector leads finally to the following

relations between T and R:

I+R=T

1 - R = plcl2 sin 202 T

P2C22 sin 201

(C.6)

This yields:

2

T = (C.7)

1 -t plcl2 sin 2¢2

P2c22 sin 291

The static pressures on both sides of the

interface are equal and consequently the ratio

PlCl2 reduces, for perfect gases, to Y_/'/2,

P2c22

where y is the ratio of specific heats. The final

relation of the transmission coefficient T in

terms of q_2and _l is the same as in Cande121;

however, the difference is in the relation

between _)2 and M2, M1 as derived in eqn

(C.1).

In order to examine the effect of M2 on the

transmission coefficient T for a given incident

wave angle q_l we can use equations (C.7) and

(C.1) for some assumed values of flow in

region 1. However, since t_l and _20 have a

one-to-one relation from eqn (C.2) and _20 is

the observable angle when no free-jet flow

exists we will discuss the changes in T due to

free-jet Mach numbers as function of _?zo.

Thus for a given value of _20 we first find _1

and _ from eqns. (C.4) and (C.3) respectively

and substitute that in eqn (C.7) to find T. We

will also assume for simplicity that Y1/'/2 = 1.

Figure C.1 shows a sample of plots of T Vs

B20 (= _20 ) for two typical cases (M1 = 0.8,

c2/cl = 1.0) and (Ma = 0.8, c2/c1 = 0.8), for a

range of Mz values from 0.0 to 0.3. This is

shown for a range of 0z0 values from 0 to 180

degrees. A note on interpreting figure C.I:

For a given _z0 the value for T for some

parameter M2 means the transmission

coefficient for the corresponding refracted

wave in region 2 whose angle is ¢2 •

Figure C.1 shows that the variation in the

transmission coefficient T for a given 02o (or

incident wave) due to free-jet variations from

NASA/C_2002-210823/VOL 1 272
0.0to 0.3is fairly smallin theangularrange

above40° or so; below40° the difference

betweentheno free-jetflow case(M2= 0),

andtheM2> 0casecanbelarge.Thevalues

of 020aboveapproximately125° in Figure

C.l(a) correspondto the "zoneof silence."

Our acoustictestdatarangevariesfrom 55°

to 165°. Soit is of someinterestto examine

whatthesmalldifferencesin T from theno

free-jet flow casein that range meanin

terms of SPL differencesin dB which is

givenby20*log(T(M2)/T(M2=0)).

Figures C.2(a) and (b) show this SPL

differenceasa functionof M2 with 0z0(or

0020) heldfixedfor a widerangeof region1

flow parameters.From thesefigureswe can

makethefollowinggeneralobservations:

1. The SPL difference is no more than

about 0.4 dB for the angularrange

investigated(60° to 110°) andis smaller

for anglescloserto 90° .

2. At 90 ° there is no SPL difference at

different M2, and any Ma, cl, c2 value.

This can be proven mathematically from

the above formulae. That is, T = 1 for 0020

= 90 ° and the corresponding refracted

wave-normal angle at any M2 value,

regardless of region 1 values.

3. Hotter the region 1 flow (lower the c2/ca

value) the smaller is the SPL difference.

4. The SPL difference is almost linear with

M2 in this range for a given value of _0.

Thus, the smaller the M2 value the

smaller is the SPL difference from the

static case Mz = 0.

NASA/CR 2002-210823/VOLl 273
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Comparison of SPL's at Different Free-Jet Mach Numbers

and "Shifted" Angles for T.O. # 3

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NASA/C_2002-210823/VOL l 284
Form Approved REPORT DOCUMENTATION PAGE

OMB No. 0704-0188

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 Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson

Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

July 2002 Final Contractor Report

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Lobed Mixer Design for Noise Suppression

Acoustic and Aerodynamic Test Data Analysis

6. AUTHOR(S)

Vinod G. Mengle and William N. Dalton

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

Rolls Royce Allison

RO. Box 420

Indianapolis, Indiana 46206-0420

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

National Aeronautics and Space Administration

Washington, DC 20546-0001

DlJ-781-30-12-00

NAS3-27394

Task Order No. 6

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-12741-1

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA CR--2002-210823-VOL1

EDR No. 18580

11. SUPPLEMENTARY NOTES

Contents were reproduced from the best available copy as provided by the authors. Project Manager, James Bridges,

Structures and Acoustics Division, NASA Glenn Research Center, organization code 5940, 216-433-2693.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category: 07 Distribution: Nonstandard

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

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

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

A comprehensive database for the acoustic and aerodynamic characteristics of several model-scale lobe mixers of bypass

ratio 5 to 6 has been created for mixed jet speeds up to 1080 ft/s at typical take-off (TO) conditions of small-to-medium

turbofan engines. The flight effect was simulated for Mach numbers up to 0.3. The static thrust performance and plume

data were also obtained at typical TO and cruise conditions. The tests were done at NASA Lewis anechoic dome and

ASK's FluiDyne Laboratories. The effect of several lobe mixer and nozzle parameters, such as, lobe scalloping, lobe

count, lobe penetration and nozzle length was examined in terms of flyover noise at constant altitude. Sound in the nozzle

reference frame was analyzed to understand the source characteristics. Several new concepts, mechanisms and methods

are reported for such lobed mixers, such as, "boomerang" scallops, "tongue" mixer, detection of "excess" internal noise

sources, and extrapolation of flyover noise data from one flight speed to different flight speeds. Noise reduction of as

much as 3 EPNdB was found with a deeply scalloped mixer compared to annular nozzle at net thrust levels of 9500 lb for

a 29 in. diameter nozzle after optimizing the nozzle length.

14. SUBJECT TERMS

En_ne noise; Jet noise; Lobed mixers; Aeroacoustics; Internal noise; Excess noise; Nozzle length;

Mixing length; Lobe scalloping; Thrust coefficient; Turbofan mixing; Jet plumes; Flight effect;

Lobe count; Lobe penetration; Lobe scalloping

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION

OF REPORT OF THIS PAGE OF ABSTRACT

Unclassified Unclassified Unclassified

NSN 7540-01-280-5500

15. NUMBER OF PAGES

295

16. PRICE CODE

20. LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102