experimental clean combustor program …...'for sale by the national technical lnfamat~on...

, '

(WASA-CR-15945E) EXPERIMENTAL CLERN N79-17656 COMBUSTOR FRCGFAM, PHASE 3 : Y O I S E MEASOREPIENT ACDEND'clM Final Report (General Elec t r ic Co.) 139 p HC A07/HF A01 CSCL 20A Unclas

63 /71 1 3 9 0 9

EXPERIMENTAL CLEAN COMBUSTOR PROGRAM - PHASE Ill NOISE MEASUREMENT ADDENDUM

FINAL REPORT

Prepared For

Natiomal Aeromartits and Spato Admlmistratiem

NASA Lewis Research Center CONTRACT NAS3-19736

https://ntrs.nasa.gov/search.jsp?R=19790009485 2020-03-28T09:23:32+00:00Z

'For sale by the National Technical lnfamat~on Service, Springfield, Virginia 22151

1

3. Rlcipknt's Catalog No.

6. Rlpon hn December 1978

6. Putorrnhg Ofg~izrtia~ Codr

8. Pr(onnhg Orpnintion Rlpon No.

R78AEC319 10. Work W No.

11. conm* or Gmt No.

NAS3-19736

13. TVP of RIpM nd Wiod CDwd

Final Re~0rt 14. Sponroring Agmcv Coda

I 1. Rrpart No. 2. Goummmt h i o n No.

NASA CR-159458 4. Title and Subtitlr

Experimental Clean Combustor Program - Phase I11 Koise Measurement Addendum Final Report

7. Author(,)

V.L. Doyle ,

9. Rrfaming Orprnlmtion Nnw and Addnrr

General Electric Company Aircraft Engine Group Cincinnati. Ohio 45215

12. Suomoring Nwn and M d m s

NASA Levis Research Center 21000 Brookpark Road Clev&nd. Ohio 44135

15 Suoprmntuv Nota

NASA Project Manager: Ronald G. Huff Lewis Research Center

16. AbstTrct

The experimental work sponsored under the Noise Measurement Addendum to the Experimental Clean Combustor Program was conducted to investigate the acouscic characteristics of the Double Annular combustor in a CF6-50 high hypass turbofan engine. Internal fluctuating pressure measurements were made in the combustor region and in the core exhaust. The transmission loss across the turbine and nozzle was determined from the measurements and compared to previous component results and present theory. The primary noise source location in the combustor was investigated. Spectral comparisons of test rig results from ECCP Phases I and I1 were made with the engine results. Com- parisons of measured overall power level were made with component and engine correlating parameters,

17. KW ad, ( ~ ~ t r d ~ J Y A U ~ ~ ~ ( S I J

Internal combustor noise measurements (26-50 engine core noise Double Annular cmbuetor noise

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21. NO. of P-

135

TABLE OF CONTENTS

Sect ion Page

1.0 SUMMARY

2.0 INTRODUCTION

3.0 DESCRIPTION OF TEST

3.1 CF6-50 Engine T e s t Vehicle 3.2 Double Annular Combustor, Duct Rig and Engine 3.3 Tes t Object ives 3.4 Engine Setup f o r Acoustic Measurements 3.5 T e s t Matrix 3.6 Data Acquis i t ion Systems 3.7 Data Reduction and Process ing

4.0 ANALYSIS AND DISCUSSION OF RESULTS

4.1 Spec t ra Comparisons 4.2 Turbine Trans fe r Function Comparison 4.3 Primary Noise Source Locat ion 4.4 Overa l l Power Level Comparison

5.0 CONCLUSIONS

5.1 Double Annular Combustor Spec t ra Comparisons 5.2 Noise Source Locat icns 5.3 Engine t o Duct Rig Comparisons

APPENDIX A - 2 Hz NARROWBAND SPECTRA RESULTS

APPENDIX B - 1 / 3 OCTAVE BAND SPECTRA RESULTS

APPENDIX C - TURBINE TRANSFER FUNCTION RESULTS

APPENDIX D - EVALUATION OF PROBE RESPONSE AT ELEVATED TEMPERATURES

APPENDIX E - CROSS-CORRELATION RESULTS

APPENDIX F - NOMENCLATURE

REFERENCES

iii

LIST OF ILLUSTRATIONS

Page

7

Figure No.

3-1 CF6-50 Engine Demonstrator Double Annular, Low Emission Combustor.

Internal Sensor Locations for Engine Combustor Noise Measurements.

Sound Separation Probe Location for Combustor Noise Measurements on a CF6-50 Engine.

E~lite Sensor Installations for ECCP Phase I11 Acoustic Test (7708283).

Combustor Borescope Kulite (282') Sensor Installation Sor ECCP Phase I11 Acoustic Test (7708284).

Setup of Core Exhaust and External Probes for ECCP Phase 111 Acoustic Test (7708286).

Schematic of Acoustic Data Acquisition System Setup.

Calibration Setup for Waveguide Probes.

Frequency Response of Waveguide Probe Systems.

Ambient Phase Calibrations of Waveguide Probe Systems.

Noise floor of Coherent Spectrum.

Typical Cross-Correlation Function for Flow with Sound Dominant . Comparison of Core Probe Power Levels with Farfield Measurements.

Core Exhaust Power Level Spectra Comparison.

Comparison of Common Measurements for Repeat Readings at 45.6 Percent Thrust.

Comparison of Common Measurements for Repeat Readings at 64.6 Percent Thrust,

Comparison of Common Measurements for Repeat Readings at 82.0 Percent Thrust.

LIST OF ILLUSTRATIONS (Continued)

Figure No.

4- 1

Page

4 1 Plane 3.0 (MADAR Probe) 113 Octave Band Spectra Variation with Engine Speed.

Plane 3.0 (MADAR Probe) Spectra Variation with Engine Speed.

Plane 3.5 (282O Borescope Probe) 1/3 Octave Band Spectra Variation with Engine Speed.

Plane 3.5 (282' Borescope Probe) Spectra Variation with Engine Speed.

Plane 4.0 (HPTN Probe) 113 Octave Band Spectra Variation with Engine Speed.

Plane 4.0 (HPTN Probe) Spectra Variation with Engine Speed.

Plane 8.0 (Exhaust Probe B) 1/3 Octave Band Spectra Variation with Engine Speed.

Plane 8.0 (Exhaust Probe B) Spectra Variation with Engine Speee.

Acoustic Probe Locations for Duct Rig Test.

Engine-to-Duct Rig Pressure Spectra Comparisoq for Double Annular Combustor at Idle,

Engine-to-Duct Rig Pressure Spectra Comparison for Double Annular Combustor at Approach,

Engine-to-Duct Rig Pressure Spectra Comparison for Double Annular Combustor at Takeoff.

Engine-to-Duct Rig Power Spectra Comparison for Double Annular Combustor at Idle,

Engine-to-Duct Rig Power Spectra Comparison for Double Annular Combustor at Approach.

Engine-to-Duct Rig Power Spectra Comparison for Double Annular Combustor at Takeoff.

Measured Attenuation Spectrum and Predicted Results for CF6-50 Six-Stage Turbine.


Figure No.

4-17 Attenuation Spectrum from 3-Stage Low Presaure Turbine.

Cross-Correlation of Plane 3.0 to Plane 3.5 (102" Probe).

Cross-Correlation of Plane 3.0 to Plane 3.5 (282" Probe).

Cross-Correlation of Plane 3.0 to Plane 4 .O.

Cross-Correlation of Plane 3.5 (102' Probe) to Plane 4.0.

Cross-Correlation of Plane 3.5 (282" Probe) t o Plane 4.0.

Cross-Correlation a t Plane 3.5 Between 102" to 282' Sensors.

Cross-Correlation of Plane 3.0 Probe to Plane 8.0 Probe, Element A.

Cross-Correlation of Plane 3.0 Probe to Plane 8.0 Pro.te, Element B.

Cross-Correlation of Core Probe Element B with .Element A,

Coherent Spectrum a t Core Nozzle Discharge, Plane 8.0.

Comparison of Core Probe Spectra a t Each Immersion.

Core Probe Immersion 4 a t 30 Percent Thrust.

Core Probe Immersion 4 a t 19.7 Percent Thrust.

Core Probe Immers&on 4 a t 29.7 Percent Thrust.

Core Probe Immersion 4 a t 64.6 Percent Thrust.

Comparison of Raw and Coherent OASPL with Core Speed.


Figure No.

4-34 Comparison of FPWheas t o PWLGE f o r In te rna l Sensors.

Comparison of PWL,eas " ~ E E F M '

Comparison of P b e a s t o PWLGEFA~ using Adjusted Fuel-Air Ratio f o r 100% P i l o t Points.

A s Measured Narrowband Spectra f o r CF6-50 Engine Condition a t 3.8 Percent Thruet, Reading 6 ( Id le ) .

A s Measured Narrowband Spectra f o r CF6-50 Engine Condition a t 19.7 Percent. Thrust, Reading 16.

A s Measured Narrowband Spectra f o r CF6-50 Engine Condition a t 30 Percent Thrust, Reading 23. (Approach Power with 100 Percent P i l o t Fuel).

A s Measured Narrowband Spectra f o r CF6-50 Engine Condition a t 29.7 Percent Thrust, Reading 39 (Approach Power with 50150 Fuel Sp l i t ) .

A s Measured Narrowband Spectra f o r CF6-50 Engine Condition a t 45.4 Percent Thrust, Reading 34.



A s Measured Narrowband Spectra f o r CF6-50 Engine Condition a t 64.6 Percent Thrust, Reading 13. ,

A s Measured Narrowband Spectra f o r (36-50 Engine Condition a t 82 Percent Thrust, Reading 18.

A s Measured Narrowband Spectra f o r CF6-50 Engine Condition a t 82 Percent Thrust, Reading 43.


Page

82

v i i

LIST OF ILLUSTRATIONS (Concluded)

Figure No.

B-1

B- 2

B- 3

B-4

B- 5

B-6

B- 7

B- 8

One-Third Octave Band Spec t ra f o r CF6-50 Engine Condit ion a t 3.8 Percent Thrust .

One-Third Octave Band Spectra f o r (26-50 Engine Condit ion a t 19.7 Percent Thrust.

One-Third Octave Band Spectra f o r CF6-50 Engine Condit ion a t 30 Percent Thrust .

One-Third Octave Band Spectra f o r CF6-50 Engine Condit ion a t 29.7 Percent Thrust .

One-Third Octave Band Spectra f o r CF6-50 Engine Condit ion a t 45.4 Percent Thrust .

One-Third Octave Band Spec t re f o r (26-50 Engine Condit ion a t 64.6 Percent Thrust .

One-Third Octave Band Spec t ra f o r CF6-50 Engine Condit ion a t 82.0 Percent Thrust .

One-Third Octave Baud Spec t ra f o r CF6-50 Engine Condit ion a t 94.6 Percent Thrust .

Coherent Spec t ra Comparisons f o r Turbine At tenua t ion from ECCP Phase I11 Engine Test .

v i i i

LIST OF TABLES

Number

CF6-50C Engine Specifications.

Double Annular Combustor Design Parameters.

Summary of ECCP Phase 111 Acoustic Test Conditions.

Ambient Frequency Response Corrections for ECCP Phase I11 Waveguide Sensor?.

ECCP Phase I11 Program Aerodynamic Performance Sunnnary.

Summary of Local Static Conditions for Combustor Test.

Test Condtions for Comparison of ECCP Phase I, I1 and I11 Results.

CF6-50 Turbine Attenuation Summary from ECCP Phase I11 Test Results.

One-Third Octave Band Spectra from ECCP Phase I11 CF6-50 Engine Test.

Time Delays and Amplitudes of Cross-Correlation Peaks for ECCP Phase I11 Test Data.

Page

4

5

14

2 2

1.0 SUMMARY

The purpose of the Noise Measurement Addendum t o the Fxperimental Clean Combustor Program sponsored by the NASA Lewis Research Center under Contract NAS3-19736 was t o inves t iga te by experiment the noise c h a r a c t e r i s t i c s of the Double Annular combustor i n the CF6-50 high bypass turbofan engine. The pro- grarc object ives were t o measure and compare the engine combustor i n t e r n a l pressure spec t ra with t e s t r i g r e s u l t s from ECCP Phases I and 11, t o determine the acoust ic t ransfer function, comparing the r e s u l t s with component t e s t s and theo re t i ca l predict ions, and t o inves t iga te the primary noise source loca t ion i n the combustor.

The s a l i e n t r e s u l t s and conclueions from t h i s inves t iga t ion a r e a s follows:

Engine t e s t r e s u l t s show spec t ra from i n t e r n a l measurments t o exh ib i t the t yp ica l core noise spec t ra shape a t approach power s e t t i n g s and below with peaks i n the 400 t o 600 Hz frequency bands.

Engine t o duct r i g r e s u l t s show l a rge d i f fe rences (10 t o 25 dB) on a pressure bas i s but exhib i t much c loser agreement (3 t o 8 dB) on a power l eve l bas i s . The spec t r a l shapes f o r both the i n l e t and discharge planes of the engine and duct r i g were s imi la r . Differ- ences i n t e s t conditions, f u e l s p l i t s , measurement loca t ions and the presence of t h e choked HPT nozzle diaphragm i n the engine were c i t ed a s possible causes fo r the l eve l differences.

The turbine t ransfer funct ion i n terms of a t tenuat ion agreed well with component r e s u l t s and theo re t i ca l predict ions.

%'he exact loca t ion of the primary noise source within the combustor was undefined from present measurements. The bes t ind ica t ion f o r the primary source locat ion is i n the region between the combustor e x i t and the region forward of the core nozzle e x i t .

Operation of the Double Annular combustor with p i l o t f u e l only r e s u l t s i n higher OAPWL than when f u e l s p l i t s between inner and outer burner s tages a r e used. This is pa r t i cu l a r ly c r i t i c a l at approach power s ince the difference can be a s much a s 9 dB.

A b e t t e r understanding of combut:tor noise during engine operation was gained from t h i s program. These r e s u l t s w i l l be usefu l i n evaluating the core noise transmitted from the engine t o the f a r f i e l d i n t e s t s t o be conducted under another NASA Lewis program.

2.0 INTRODUCTION

The NASA/General E l e c t r i c Experimental Clean Combustor Program has been s p e c i f i c a l l y d i rec ted toward developing an advanced low emission, low noise combustor f o r use i n the General E l e c t t f - 'F6-50 engine family, The CF6-50 engine family is the higher power so!.. L . \ h e two CF6 high bypass turbofan engine fami l ies developed by Gencre : G1ecti-i: .

Noise evaluat ion a s wel l a s emissions t e s t t n g have played important r o l e s i n the development work conducted during t h i s program. Acoustic measurements have been made on seve ra l f u l l s i ze , annular combustor configurat ions t e s t ed i n Phases I and I1 of the Experimental Clean Combustor Program (ECCP) during duct r i g t e s t i ng . These measurements ind ica ted t h a t t h e peak sound pressure l e v e l (SPL) occurred near a frequency of 1000 Hz. Generally, fa r - f i e l d da ta on engines have been found t o peak a t 400-500 Hz. This d i f fe rence between an i so l a t ed combustor running i n a component test s tand and a combus- t o r on an engine is the b a s i s f o r the study reported here. In order t o i nves t i ga t e the e f f e c t of the engine i n s t a l l a t i o n on the combustor acous t ics i t is necessary t o take acous t ic measurements on a given combustor both i n t h e combustor development r i g and i n t he engine. During Phase I and I1 of the ECCP, acous t ic measurements were made i n a combustor development r i g on a combustor s imi la r t o the one used i n t he engine demonstration tests under ECCP Phase 111.

The work sponsored under t h i s acous t ic addendum t o NASA Lewis Contract NAS3-19736 provided the measurements i n the engine necessary t o evaluate the combustor engine acous t ic s igna ture and the acous t ic transmission l o s s of the low frequency combustor noise through turbine.

3.0 DESCRIPTION OF TEST

CF6-50 ENGINE TEST VEHICLE

The t e s t s conducted under Phase 111 were performed on a CF6-50 high bypass turbofan engine. The basic engine descr ip t ion and d e t a i l s of the spe- c i f i c conf igura t ions tested fo r the emissions inves t iga t ion a r e found i n Reference 1. The CF6-50 engine model operating parameters were selected f o r use a s the combustor design and t e s t conditions of t h i s program. Key overa l l spec i f ica t ions of t h i s engitrd a r e presented i n Table 1.

3.2 DOUBLE ANNULAR COMBUSTOR, DUCT R I G AND ENGINE

3.2.1 Design Concept

In the Phase I and 11 Programs, four advanced combustor design concepts were evaluated i n CF6-50 engine-size f u l l annular combustor duct r i g tests (References 2-5). The bes t ove ra l l r e s u l t s fo r engine appl icat ion were obtained with the Double Annular combustor configuration D-12, which was the prototype f o r t he fl ightworthy demonstrator Double Annular combustor designed i n the Phase 11 Program f o r use i n these Phase I11 Program combustor r i g and CF6-50 engine t e s t s . Acoustic component t e s t s were conducted i n Phase I1 (Reference 5. on a Double Annular combustor configuration D-13 which was very s imi la r t o t h D-12 design. The D-13 difcnred from the D-12 by the percent- ages of a i r f low d i s t r i bu t ion through i l o t and main s tages as noted i n Table 2.

The Double Annular combustor comprises two annular primary burning zones, i n p a r a l l e l , separated by a shor t centerbody. Thir ty f u e l nozzles a r e used i n each annulus. The outer annulus is the p i l o t s t age and is always fueled. The inner annulus is the main s t age and is fueled only a t higher engine-power operating conditions. The a i r f low d i s t r i b u t i o n is highly biased t o the main s tage in order t o reduce both i d l e and high-power emissions. The p i l o t s t age a i r f low is spec i f i ca l ly sized t o provide nearly s toichiometr ic fuel-air r a t i o s and long residence times a t i d l e power s e t t i n g s , thereby minimizing C3 and HC emissions levels . A t high-power operating conditions, most of t he fue l is supplied t o the main stage. In t h i s s tage, the residence times a r e very shor t . Also, a t high-power operat ing conditions, lean fuel-air r a t i o s a r e maintained i n both s tages t o minimize NOx and smoke emissions levels .

3.2.2 Engine Demonstrator Combustor Design

With t h i s Double Annular combustor concept, program goals f o r CO and HC emissions a t i d l e operating conditions were achieved ea r ly i n the Phase XI Program, and maintained throughout the f i n a l combustor refinement test se r i e s . The emphasis i n t h i s f i n a l s e r i e s of refinement tests was therefore t o more nearly approach the N4, emission goal, and t o meet t h e engine i n s t a l l a t i o n

Table 1. CF6-50C Engine Specif icat ions.

Takeoff Rating (SLS) .Yilrust 224.2 kN (50,400 I.bf) Speci f ic Fuel Consumption 10.7 m g l ~ s (0.377 lbmllbf-hr)

Maximum Cruise (Mach 0.85110.7 km) Thrust 48 kN (10,800 l b f ) Spec i f ic Fuel Consumption 18.6 mg/Ne (0.656 lbmllbf-hr)

Weight; 3780 kg (8330 lb )

Length 482 c m (190 in)

Maximum Diameter 272 cm (107 in)

Pressure Rat io Takeoff Maximum Cruise

Bypass Rat io (Takeoff) 4.4

Tota l Airflow (Takeoff) 659 kg/s (1452 lbm/s)

U \ D U Q\Od 0 . . . . . . N U U N b U 4 0 40 A mrl el '6

ra c'd M O O 0) d V

and performance requirements. I n t h i s s e r i e s , very l i t t l e change i n NOx emission l e v e l s was obtained and engine i n s t a l l a t i o n and performance r e q ~ i r e - ments were most near ly met with conf igura t ion D-12 which was t h e r e f o r e se- l ec ted as the prototype f o r a more s o p h i s t i c a t e d second generat ion vers ion of t h i s ,~dv;~i lccd low e1111ssion c.omhustor design f o r use i n t h e Phase 111 demon- s t r a t i o n engine t e s t s . This second generat ion combustor conf igura t ion was needed because the prototype coaf igura t ion used i n Phases I and I1 was de- s igned t o accommodate d i f f e r e n t i a l thermal growths, p ressure loads , v i b r a t i o n loads , and mechanical assembly were n c t adequate t o permit t h e use of t h i s combustor i n engine t e s t s .

The r e s u l t i n g demonstrator engine combustor des ign is shown i n Figure 3-1. The aerothermal design f e a t u r e s of t h i s demonstrator engine combustor was pat terned a f t e r those of the prototype combustor. I n a d d i t i o n , advanced aeromechanical design f e a t u r e s der ived from o t h e r General E l e c t r i c programs were incorporated i n t o i ts design.

Key aerothermal design parameters of t h e Double Annular combustors a r e compared i n Table 2. Airflow d i s t r i b u t i o n s a r e very similar except t h a t t h e demonstrator combustor dome cooling a i r f l o w s a r e s l i g h t l y higher which is accomplished p r imar i ly by reducing p r o f i l e t r i m a i r f low. Key v e l o c i t i e s a r e a l s o very s i m i i a r except t h a t inner and o u t e r passage v e l o c i t i e s of t h e demonstrator combustor a r e more near ly equal ized t o reduce p a r a s i t i c p ressure losses . Dome he igh t s of the demonstrator combustor were increased about 20 percent t o provide the a d d i t i o n a l room wi th in t h e cowl t o accommodate t h e needed r a d i a l movements of t h e s w i r l cup s l i p j o i n t s .

3.3 - TEST OBJECTIVES

The a c o u s t i c t e s t of the Double Annular combustor i n t h e CF6-50 engine provided the f i r s t i n d i c a t i o n of i n t e r n a l f l u c t u a t i n g p ressure measurements f o r t h i s type of combustor i n an engine environment. Previous duct r i g measurements gave l e v e l s of noise based OIL t e s t s conducted a t sca led engine cond i t ions a t l o c a l s t a t i c p ressures of up t o 10 a t m . compared t o engine pressure l e v e l s of 26 atm a t takeoff . The engine test provided a base f o r comparison with the duct r i g r e s u l t s from Phase I and 11.

Other t e s t ob jec t ives included determining t h e a c o u s t i c t r a n s f e r func- t i a n across the tu rb ine by eva lua t ing t h e a t t e n u a t i o n of t h e a c o u s t i c s i g n a l upstream and downstream of t h e tu rb ine , and i d e n t i f y i n g t h e l o c a t i o n of t h e primary noise source wi th in t h e combustor from cross -cor re la t ion a n a l y s i s r e s u l t s .

3.4 - ENGINE SETUP FOR ACOUSTIC MEASUREMENTS

The a c o u s t i c t e s t s of t h e ECCP Phase 111 Double Annular combustor in- s t a l l e d i n CF6-50 engine number 445-10517 were conducted a f t e r completion of t h e combustor perfrnnance and emissions tests. The engine t e s t c e l l i n s t a l - l a t i o n and se tup f o r these t e s t s is descr ibed i n t h e ECCP Phase I11 f i n a l

report on the emissions t e s t , Reference 1. The acous t ic t e s t setup u t i l i z e 4 the same engine configurat ion except t ha t the gas sampling rakes i n t he tort: exhaust and those i n the turbine r e a r frame were removed. The basic aerodynamic performance and engine s a f e ty instrumentation remained e s s e n t i a l l y the same a s described i n Reference 1.

Dynamic pressure instrumentation was i n s t a l l e d i n ava i lab le access po ct ;

a t three a x i a l planes i n the combustor region and a t the core nozzle. An ex te rna l probe mounted i n the t e s t c e l l was used t o i nd i ca t e c e l l noisc l eve l s and t o attempt t o determine the amount of combustor noise t ranslerr .11 ou ts ide the engine. Acoustic probes were posit ioned i n the W A R port a t l h l e

compressor discharge, Plane 3.0; i n the borescope po r t s a t combustor r:nt..-ance, Plane 3.5; i n the borescope por t of the f i r s t s t a g e turbine nozzle a t the combustor e x i t , Plane 4.0; and a t the core nozzle discharge, Plane 8.0, as i l l u s t r a t e d i n Figures 3-2 and 3-3 respect ively, and shown i n the photographs of Figures 3-4, 3-5 and 3-6.

3.5 TEST MATRIX

211e t e s t matrix for acoust ic da ta pa ra l l e l ed the conditions s e t f o r the combustor performance and emissions tests. This was done t o match the poise s igna ture of the combustor with comparable emissions r e s u l t s i n order t o a s se s s the ove ra l l e f fec t iveness of the Double Annular combustar. The t e s t po in ts , based on percent net t h rus t , covered the s ea l e v e l s t a t i c operating l i n e of the CF6-50 engine and included i d l e , approach and takeoff power s e t t i ngs . A t o t a l of e igh t (8) test condi t ions were set. Repeat readink3 were taken t o check data r epea t ab i l i t y and t o supplement sensor measureme~lts of severa l t ha t were l o s t during the run due t o equipment malfunctions. Table 3 sunrmarizes the test mat .x f o r acous t ic da ta and shows t h a t a l l required condi t ions were obtained. Matching repeat po in ts and redundant probes a t two of the four a x i a l planes provided complete sets of measurements a t every plane f o r each test point. The e igh t test condi t ions included two approach poin ts a t 30% n e t th rus t . The f i r s t approach s e t t i n g consis ted of the 100% p i l o t f u e l only, while the second condition approximated the 50% f u e l s p l i t between inner and outer sets of f u e l nozzles.

Aerodynamic performance da ta was provided a t each t e s t condition through a combination of the on-line DMS (Data Management System) readings supplied f o r each engine point s e t t i n g and some performance da ta supplied by readings taken a t s imi l a r s e t points during the running of the emissions tests. This was necessary as a r e s u l t of d i g i t a l system problems encountered during the acous t ic run. The paired aero and acous t ic da ta a r e compatible f o r a l l t e s t conditions and represent t yp i ca l s e t t i n g s with t h i s engine.

3.6 DATA ACQUISITION SYSTEMS

The da ta acquis i t ion systems f o r t h i s test included both analog and d i g i t a l systems. The acous t ic da ta was acquired from combustor i n t e rnc r f luc tua t ing pressure measurements made with four single-element waveguide

Core Nozzle

Discharge

Plane

Probe Externally I

Figure 3 - 3 . Sound Separation Probe Location for Combustor Noise Measurements on CF6-50 Engine

Figure 3-6. Srwtup of Care Exhaust and External Probes For ECC1' Phase 111 Acoustic T e s t 17708286)

. -

'. . - . 6 a- .-w .- -- . -. .- I., ,- ., . , ..r.r.-*. :'.. .

probe systems loca ted a s i l l u s t r a t e d i n Figure 3-2. The i n t s r n a l measurements a t t h e core nozzle e x i t were acqu i red wi th a dual-element snund sepa- r a t i o n probe wi th 12.7 cm a x i a l spacing between sensors . Figure 3-3 shows t h i s probe,

The waveguide sensors cons i s t ed of K u l i t e p r e s s u r e t r ansducers (XGE-1S- 375-200D) AC-coupled and mounted i n a i r -cooled tee-blocks. A capped 12.192 meter long semi - in f in i t e c o i l was f i t t e d t o one e r d of the t e e , whi le t h e o t h e r end of the t e e was connected t o a s t andof f tube approximately 30.48 cm long which was mounted t o t h e engine t o complete t h e system. The K u l i t e system e x c i t a t i o n v o l t a g e was provided by a cons tan t 10 VDC source. The output s i g n a l was ampl i f ied p r i o r t o zecording on magnetic tape. I n i t i a l ampl i f ica- t i o n was provided by the tape recorder p r e a m p l i f i e r s which were s e t on cons tan t ga in s e t t i n g s . A second s e t of a m p l i f i e r s i n s e r i e s wi th t h e p r e a m p l i f i e r s had a d j u s t a b l e ga in s e t t i n g s which provided t h e f i n a l s i g n a l boost ensur ing the s i g n a l t r ansmi t t ed from the K u l i t e was w e l l above t h e system no i se f l o o r .

The s i g n a l s were recorded on a 28 channel Sabre I V magnetic t ape recorder a t 76.2 cm/s up t o 10000 Hz. Approximately 1 .5 t o 2 min. of d a t a were recorded a t each t e s t condi t ion. S i g n a l po1arii:y was checked f o r a l l sensors p r i o r t o recording data . A spectrum ana lyze r connected i n p a r a l l e l w i t h t h e t ape recorder and coupled t o an X-Y p l o t t e r enabled on-l ine 10 Hz narrowband s p e c t r a t o be acquired during t h e t e s t run. Th i s provided a check of the measurements and an i n d i c a t o r of sensor s i g n a l v a l i d i t y . The spectrum analyzer d i d no t , however, g ive a check on p o l a r i t y . Figure 3-7 is a schematic i l l u s ~ r a t i o n of t h e d a t a a c q u i s i t i o n system setup.

3.6.1 K u l i t e Waveguide Probe P r e t e s t C a l i b r a t i o n

P r i o r t o recording da ta , t h e K u l i t e sensor c a l i b r a t i o n s were e s t a b l i s h e d f o r each recorder channel. Bench s e n s i t i v i t i e s determincd f o r each K u l i t e sensor i n the l abora to ry were used t o se tup t h e tape recorder c a l i b r a t i o n . The bench s e n s i t i v i t i e s and e x c i t a t i o n v o l t a g e combinations were set w i t h a dial-a-source vo l t age f o r a h igh p ressure (0.68 atmj system used wi th t h e waveguide probes, and a low pressure (0.068 atm) system f o r the exhaust and c e l l probes. Gain s e t t i n g adjustments were made on both s e t s of a m p l i f i e r s t o match a 4 v o l t peak-to-peak t a p e recorder i n p u t s i g n a l requirement. Once t h i s was e s t a b l i s h e d an AC c a l i b r a t i o n from an e l e c t r i c a l source was recorded on a11 channels which referenced each K u l i t e sensor l e v e l .

3.6.2 Frequency Response and Phase C a l i b r a t i o n

Frequency response and phase c a l i b r a t i o n s a t ambient cond i t ions (room temperature) were performed on the K u l i t e sensors loca ted i n t h e combustac region (MADAR, 102', 282' and HPTN borescope p o r t s ) . This was done t o account f o r probe l o s s e s r e s u l t i n g from t h e s tandoff tube l eng th r e q u i r e d on each sensor . The a c o u s t i c probe loca ted i n t h e core exhaust d i d no t r e q u i r e c a l i b r a t i o n s i n c e the K u l i t e s were f l u s h mounted on t h e probe and previous c a l i b r a t i o n s showed a f l a t frequency response over t h e range of i n t e r e s t and up t o 10 KHz.

With Double Annular Combustor

Engine KuLitv Pcr~er Systems

Amplifier 0 Cascade Amplifier

Scope Monitor

I Tape Recorder

Spec truu Analyzer

Recorder n Pi g u n 3 - 7 Schematic of Acowt ic Data Acquisition Syatem Setup

C a l i b r a t i o n System

The apparatus se tup f o r t h e frequency response and phase c a l i b r a t i o n s is a s shown I n t h e schematic of Figure 3-8. The c a l i b r a t i o n .system cons i s t ed of a p lane wave tube wi th s u f f i c i e n t tube l eng th t o dampen r e f l e c t i o n s i n the frequency range of 100 t o 5000 Hz.

A t t h e t e s t s e c t i o n , a f l u s h mounted K u l i t e p ressure t r ansducer , type CQL-125-25D, was used a s t h e re fe rence . The t e s t probe i n each ca.e was posi t ioned a t t h e samb* a x i a l p lane and d i sp laced 90' from t h e re fe rence . The K u l i t e e x c i t a t i o n v o l t a g e was supp l i ed by 6 v o l t b a t t e r i e s . The output of each system went through a low n o i s e a m p l i f i e r and then i n t o t h e s i g n a l genera t ing u n i t .

The monitor sign; a pure tone swept through t h e frequency range. The s i g n a l was suppi le : . 3 sweep o s c i L l a t o r which was c o n t r o l l e d by a auto- matic l e v e l r e g u l a t o r & 2 l i f i e d by p r e and power a m p l i f i e r s be fo re being fed i n t o t h e speaker mounted a t one end of t h e p lane wave tube. The output of the sweep o s c i l l a t o r was connected t o t h e X-axis (frequency s c a l e ) of an X-Y recorder .

The frequency response of t h e t e s t probe and r e f e r e n c e K u l i t e were measured consecu t ive ly wi th a dynamic analyzer . The output of each K u l i t e system was p l o t t e d on t h e Y-axis (dB s c a l e ) of t h e X-Y recorder . The sens i - t i v i t y of each K u l i t e system was ad jus ted t o measure the same l e v e l of sound p ressure i n t h e tube a t the t e s t s e c t i o n . This assured t h a t both K u l i t e s sensed t h e same s i g n a l i n p lane wave form.

Phase c a l i b r a t i o n s were performed wi th the use of a phase meter. The output s i g n a l s of t h e r e f e r e n r e and t e s t probe K u l i t e were l ead i n t o t h e phase meter inpu t channels , A E B, r e s p e c t i v e l y . Adjustments t o account f o r any e l e c t r o n i c d i f f e r e n c e s between t h e K u l i t e systems ( s e n s i t i v i t i e s , e r c , ) were performed t o ensure t h a t t h e output s i g n a l p l o t t e d on t h e Y-axis (phase ang le , deg.) represented t h e complete phase d i f f e r e n c e between t h e f l u s h mounted re fe rence and t h e t e s t K u l i t e ,

L imi ta t ions t o t h e c a l i b r a t i o n system were as follows:

Frequencies below 100 Hz were not a t t a i n a b l e beczdse of t h e danger of exceeding t h e speaker maximum v o l t a g e l i m i t x i o n .

Ambient c a l i b r a t i o n s (frequency response and phase) do not repre- s e n t a c t u a l t e s t c o n d i t i o n s u n l e s s t h e s t a t e p r o p e r t i e s a r e mea- su red a t t h e K u l i t e l o c a t i o n . Care must t h e r e f o r e be used when applying ambient c o r r e c t i o n s ( e s p e c i a l l y phase) t o t h e hot t e s t d a t a .

Waveguide Probe C a l i b r a t i o n Resu l t s

Two types of probe systems were c a l i b r a t e d . One system (System A) con- s i s t e d of 0.635 cm 01). 0.00711 cm w a l l tubing and had a 33.02 cm long standoff tube. Th i s system was used i n t h e W A R por t and combustor borescope por t s a t 102" and 282". The o t h e r system (System 8) used i n t h e HPT nozz le borescope por t cons i s t ed of 0.635 cm OD, 0.01245 cm w a l l tubing wi th a 19.05 cm long s tandoff tube. Both types of systems used a s t andard t e e K u l i t e mount f i t t e d wi th an air-couled j a c k e t . A l l systems used a type XGE-lS-375- 200D Kul i t e p ressure t ransducer wi th t h e re fe rence p r e s s u r e tube pneumat- I - a l l y coupled t o the semi - in f in i t e c o i l mounted a t t h e end of t h e probe.

The frequency response of t h e systems a r e shown i n F igures 3-9a-c. The response of System A is shown i n Figure 3-9a wi th a s t r a i g h t s t andof f tube and i n Figure 3-9b wi th t h e curved tube a s run on t h e engine. Both response curves a r e smooth t o 1000 Hz wi th a l o s s of about 1.5 ds . A t 2000 Hz they a r e down 2 dB whi le a t 5000 Hz t h e l o s s is 3.5 and 4.5 dB, r e s p e c t i v e l y . Agreement w a s e x c e l l e n t , showing no s i g n i f i c a n t l o s s due t o t h e curved tube, and the response was w e l l w i t h i n t h e s p e c i f i e d requirements. The system response curves apply t o t h e borescope K u l i t e s a t 102' and 282" a l s o , s i n c e they were of t h e same conf igura t ion .

The KPT nozz le Kul i t e , System B, had a response a s shown i n F igure 3-9c. It is smoath t o 1000 Hz wi th a drop of 2.5 dB a t 2000 Hz. However, above 2000 Hz the dropoff r a t e increased t o 7 dB a t 5000 Hz which was a t t r i b u t e d t o the smal ler diameter tubing requ i red t o f i t i n t ~ t h e p h y s i c a l c o n s t r a i n t of the HPT nozzle borescope plug. Since t h e majol p o r t i o n of t h e frequency range below 3500 Hz had s i g n i f i c a n t l y l e s s than a 5 dB l o s s r h e r e was no s e r i o u s t e c h n i c a l l i m i t a t i o n on t h e d a t a acquired wi th t h i s sensor .

The response i n t h e low f requenc ies (30-100 Hz) is f l a t from previous experience. Therefore, a smooth e x t r a p o l a t i o n of t h e response curve from 100 Hz t o 30 Hz was performed t o complete the curve.

The phase c a l i b r a t i o n of System A is shown i n Figure 3-10a f o r t h e curved s tandoff tube conf igura t ion on t h e MADAR sensor . Th i s is r e p r e s e n t a t i v e of t h e o t h e r System A Kul i t e s (102" and 282" combustor borescope).

System B phase c a l i b r a t i o n i n Figure 3-lob shows t h e phase t o be smooth up t o approximately 2500 Hz. Above t h a t frequency, t h e r e s u l t s r e f l e c t those of the frequency response. However, t h e System B phase ang le h a s a d i f f e r e n t l i n e a r r e l a t i o n s h i p than t h a t of System A causing t h e phase informat ion from System B t o Le quest ionable .

The K u l i t e probe c a l i b r a t i o n s f o r both types of systems g i v e a good r e p r e s e n t a t i o n of t h e frequency response and phase r e l a t i o n s h i p s f o r t h e ambient s t a t e . A t a b u l a t i o n of t h e frequency response c o r r e c t i o n s f o r each type of waveguide sensor a r e presented i n Table 4.

.

I 1 1 0.386 cm (ID) x 19.05 cm LO^^ I 1

. Straight Standoff Tube

w-* -

a ) S y s t e m A I're-Test r Responsc? ' I

I

6.493 cm (ID) x 33.0; cm ~ o h g Curved Standoff Tube

- i -

I - I' 4 5dE

I I Standoff rube I I I I

1 m v m 1 v

b ) System A Post-Test Response

&

Response

v I 1 1 v w 1

". -f-

1

50 100 200 !500 lo00 2000 5000 Frequency, Hz

F i g u r ~ 3-90 Frequency Response o f Waveguide Probe Systems

w w I I

a. System A Phase Calibration

a 0.493 cm (ID) x 33.02 cm ~ o h g Curved Standoff Tube

0.386 cm (ID) x 19.05 cm Long Standoff Tube

loo0 2000 3ocQ 4000 5000 Frequency , Hz

b. System B Phase Calibration

Figure 3-10. Ambient ?hase Calibrations of Waveguide Probe Systems

Table 4. Ambient Frequency Response Corrections for ECCP Phase 111 Waveguide Sensors.

Frequency Planes 3.0 and 3 5 Plane 4.0 Hz (0 .492 cm I!) Svstem' (F.386 cm I D ~ ) r s t e m )

Applied to Measured 1 / 3 Octave Band 'Jalues

3.7 D.QTA REDUCTION AND PROCESSING

3.7.1 Description and Implementation of Techniques

Several d i f f e r en t technique9 were employed i n the reduction of the acoust ic data from the engine t e s t IF x?& thess were narrowband and 1 / 3 octave band s p e c t r a l analysis , and d i g i t a l f a s t r'ourier t ransf o w techniques of coherence ana lys i s and cross-correlat ion analysis .

Narrowband Spec t ra l A n a l y e

Narrowbana s p e c t r a l ana lys i s ,~cs:.. IS of inves t iga t ing a small segment of the acoust ic energy using a constant bandwidth f i l t e r . On-line 10 Hz Sandvidth spec t ra was obtained from a conventional spectrum analyzer f o r f requencies frcw fl t o 2500 Hz w i t - ,&ple times of 0.1 s ec and 32 averages t o y i e ld a t o t a l ana lys i s record length of 3.2 seconds. Considerably f i n e r resolut ion was d t a i n e d using a d i g i t a l fou r i e r transform analyzer. The spec t r a obtained from t h i s system at frequencies from 0 t o 2000 Hz was of 2 Hz bandwidth a t sample times of 0.1 s ec and employed 100 t o 150 averages t o y i e l d a t o t a l a,*alysis record length of 10 t~ 15 seconds. The high resolu- t i o n narrowband spec t ra obtained f o r each Kul i te sensor were used t o assist i n evaluating trends from the i n t e r n a l pressure measurements. Fluctuat ing pressure l e v e l (FPL) was used ins tead of sound pressure l e v e l (SPL) s ince the pressure s igna l contains turbulence i n addi t ion t o Sound. P lo t s of t he narrowband spec t ra f o r the working sensors from a l l e ight t e s t conditions a r e found i n Appendix A.

One-Third Octave Band Spectral Analysis

The 113 octave band spec t ra were processed from the recorded da ta by s tandard techniques f o r frequencies of 50 through 2000 Hz. Overall pressure l eve l s were computed from the resu l t ing spec t ra which were corrected f o r probe response losses. The 1 / 3 octave band FPL spec t ra (113 OBFPL) determined f o r the e igh t test conditions were used f o r evaluat ing the combustor i n t e r n a l measurement trends and f o r comparisons with the ECCP Phase I and I1 component spec t ra r e su l t s . Tabulations and p l o t s of the 1 /3 OEFPL r e s u l t s a r e presented i n Appendix B. The ove ra l l f luc tua t ing pressure l e v e l (OAFPL) obtained from these 113 OBPPL spec t ra were used t o determine the measured power l eve l (F'E'WLmeas) assuming the t o t a l pressure s i g n a l was acoustic.

Coherence Analysis

Coherence ana lys is is a measure of the amount of s i m i l a r i t y o r coherence between two s igna ls i n the frequency domain. Pressure s igna l s generated upstream i n the combustor and received downstream of the turbine contain frequencies which have the same or s imi l a r c h a r a c t e r i s t i c s as the upstream s igna l . The coherent port ion of the downstream spectrum -s considered t o be mostly sound i f the s igna l f a l l s above the noise f l o o r es tab l i shed f o r an uncorrelated random s igna l input. Coherent 113 OASPL spec t ra were used t o

determine the turbine t r ans fe r function between upstream and downstream Kuli tes i n the CF6-50 engine. The coherent spec t r a a t Plane 4 ( tu rb ine i n l e t ) were obtained r e l a t i v e t o Plane 3.5 (combustor i n l e t ) . S imi la r ly , the coherent spec t ra a t Plane 8.0 (core nozzle e x i t ) were obtained r e l a t i v e t o Plane 3.5. Thc turbine a t tenua t ion was computed from the d i f fe rence between the Plane 4.0 and Plane 8.0 coherent spectra . Processing of the coherent spec t ra was done with a constant number of averages (50) f o r a l l sensor pa i r s . The noise f l o o r between the raw and coherent spec t r a was an average of 18 dB below the raw s igna l f o r the 50 t o 2000 Hz frequency range.

The noise f l o o r of the coherent s p e c t r a l ana lys i s procedure was de te r - mined using an independent random noise source d i r e c t l y i n t o the analyzer from a white noise s i g n a l generator a s t he input sensor s igna l . The coherent s p e c t r a l ana lys i s was performed using the f l uc tua t ing pressure s igna l a s the output sensor.

The coherent spec t r a displayed represen ts t h a t port ion of t he input s i gna l t h a t i s apparently coherent with t he output s igna l . For a completely independent input the s i g n a l s a r e completely incoherent which would r e s u l t i n a zero dB coherent spectrum using an unlimited number of averages. However, t he da ta ana lys i s procedure used f o r t h i s program was l imited t o t he number of averages d i c t a t ed by the length of the sample. The sample length w a s designated by the 1.5 t o 2 min recording time a t each da t a point . The number of averages were set a t 25 f o r t he cross-correlat ion ana lys i s and 50 f o r the coherence analysis . The coherent noise f l o o r due t o the l imi ted number of averages f o r each sensor was approximately 18 dB below the raw s igna l . This can be seen i n Figure 3-11 which shows the t yp i ca l r e s u l t s of t he coherence d n a l y s i s a t the output sensor loca t ion using the 50 averages and an uncor- re la ted input s igna l .

The e l ec t ron i c no ise f l o o r of the da ta acqu i s i t i on system (Kulite senso r s , ampl i f ie rs and tape recorder) was checked p r i o r t o each test by recording an "ambient" reading with the engine o f f . Comparisons of these ambient readings with the da ta runs ind ica ted t he spec t r a of t he noise f l oo r t o be 25 t o 30 dB below the l eve l of the da t a a t the i d l e power s e t t i n g and 40 t o 50 dB below the da ta a t t h e higher power s e t t i n g s . The e l ec t ron i c no ise f l o o r did not present a problem i n the ana lys i s of the data.

The majori ty of the coherent spec t r a frequencies below 1000 Hz were above the noise f l o o r f o r the low power s e t t i n g s below 45% Fn. The coherent spec t ra above 800 Hz were a f fec ted by th s noise f l o o r a t almost a l l test conditions. A 3 dB reduction i n the noise f l o o r l e v e l requi res doubling the number of averages from 50 t o 100. The coherent spec t r a a t t he high f re - quencies could be 6 dB o r more below the noise f l o o r which would requi re a grea te r number of averages than the 90 sec t o 130 sec sample t i m e avai lable . The a t tenua t ion r e s u l t s from t h i s present study a r e concentrated pr imari ly between frequencies of 100 t o 800 Hz. Previous a t tenua t ion da t a were obta ined (Reference 6) between 100 t o 1200 Hz.

. - -

Raw Spectrum From Probe Signal -

9

- Average &18 db -

Coherent spectrun - , Fnrm Uncorrelated

Random Noime 9

- . t . . . . r . . . - . . - 90 - . - - . . . . . . . . . . .

125 250 500 lo00 2000

Frequency, Hz

Figure 3-11. Noise Floor of Coherent Spectrum

Negative time delays i n t h e c ross -cor re la t ion of K ~ i l i t e p a i r s were no t d i r e c t l y incorporated i n t o the coherent s p e c t r a r e s u l t s , s i n c e t h e a n a l y s i s only handles p o s i t i v e time delays. The b i a s e r r o r s due t o these negat ive time delays were computer from t h e method d e t a i l e d i n Reference 7. The b i a s e r r o r c a l c u l a t i n g procedure p e r t i n e n t t o t h e a n a l y s i s method used dur ing t h i s program i s found i n t h e appzndix of t h a t paper. Typical negat ive delay times ranged from -0.0015 t o -0.007 seconds which corresponded t o e r r o r s i n co- he ren t s p e c t r a l e v e l of 0 .1 t o 0.6 dB g r e a t e r l eve l . These d i f f e r e n c e s were sub t rac ted from t h e o v e r a l l coherent s p e c t r a l e v e l s . Appendix C p r e s e n t s t h e coherent s p e c t r a p l o t s used t o determine t h e t u r b i n e t r a n s f e r func t ion on t h e CF6-50 engine.

Cross-Correlation Analysis

While s p e c t r a l a n a l y s i s c o n s i s t s of examination of a s i g n a l content i n t h e frequency domain, c o r r e l a t i o n a n a l y s i s may be thought of a s a n analogous inspec t ion i n t h e time domain. Cross-corre la t ion a n a l y s i s is a measure of t h e propagation-time-delay c h a r a c t e r i s t i c s of s i g n a l t ransmiss ion. I n cross- c o r r e l a t i o n , one s i g n a l i s compared wi th a time-delayed second s i g n a l t o determine the amount of s i m i l a r i t y between t h e two s i g n a l s . By t h i s tech- n ique, time de lays assoc ia ted with a c o u s t i c and tu rbu len t s i g n a l s over s h o r t d i s t a n c e s can be i d e n t i f i e d and t h e i r r e l a t i v e s t r e n g t h s determined from t h e anlplitudes of t h e normalized c o r r e l a t i o n c o e f f i c i e n t , Rv. For example, Figure 3-12 d i sp lays the c ross -cor re la t ion between two a x i a l l y spaced sensors loca ted i n a flow. The c ross -cor re la t ion funct ion d i s p l a y s s e v e r a l peaks. The l a r g e s t peak (Rxy = 0.75) is r e p r e s e n t a t i v e of a s i g n a l propagating a t t h i s speed of sound r e l a t i v e t c t h e f low and corresponds t o t h e a c o u s t i c pe r tu rba t ions . Two s i m i l a r peaks, smal le r i n magnitude (Rv = 0.18), but wi th opposi te time delays a r e noted on t h e f i g u r e s . The p o s i t i v e time delay is a s i g n a l convected by t h e flow and corresponds t o t h e turbulence. The negat ive time delay represen t s a s i g n a l t r a v e l l i n g upstream a t a v e l o c i t y corresponding t o t h e d i f f e r e n c e between t h e flow and a c o u s t i c v e l o c i t i e s .

With l a r g e spacings between sensors , (d i s t ances much g r e a t e r than t h e q u a r t e r wavelengths of t h e f requencies of i n t e r e s t ) only t h e a c o u s t i c s i g n a l i s w e l l c o r r e l a t e d . The turbulence is l e s s c o r r e l a t e d . I n t h i s case , the peaks observed w i l l be p r imar i ly acous t i c . This is t y p i c a l of cross-corre la- t i o n s between combustor sensors and t h e downstream probes.

3.7.2 Data Process ing

Frequency Response Correct ions

The a c o u s t i c d a t a acquired during t h e combustor test w a s cor rec ted on a 1 / 3 OBFPL b a s i s by t h e ambient frequency response determined from room tem- p e r a t u r e c a l i b r a t i o n s . The ambient response c o r r e c t i o n s app l ied t o t h e measured d a t a (between 50 t o 2000 Hz) acquired from t h e waveguide s e n s o r s a r e presented i n Table 4.

ECCP Phase I11

-.03 -.02 -.oi o .01 .02 -03 Time Delay 7 , mconds

Figure 3-12. Typical Cross-Cormlation Function for Flow with Sound Dominant

The e f f e c t of e leva ted temperature and p ressure on t h e response correc- t i o n was evaluated. Appendix D g i v e s t h e d e t a i l s of t h e eva lua t ion which is based on t h e t ransmiss ion l o s s due t o viscous damping wi th in a tube of c i rcu- l a r cross-sect ion assuming no i n t e r n a l r e f l e ~ t i o n s and no non-linear e f f e c t s a ssoc ia ted wi th sound pressure l e v e l s aLove 120 dB. The waveguide s e n s o r s i n the combustor region were loca ted under t h e engine cowling where t h e volume of a i r may be considered t o reach a s i n k temperature of about 477.813 forward of t h e f i r e w a l l and 533.313 i n t h e a f t regions . Using t h i s temperature (477.8K) and t h e s t a t i c p r e s s u r e a t Plane 3.0 and 3.5 of 26 atm f o r t h e takeoff con- d i t i o n t h e t ransmiss ion l o s s of t h e 2000 Hz frequency was determined and compared t o the t ransmiss ion l o s s ca lcu la ted from t h e ambient c a l i b r a t i o n condi t ions . The r e s u l t showed t h e l o s s a t e leva ted cond i t ions t o a c t u a l l y be smal le r than t h e ambient c a l i b r a t i o n l o s s by approximately a r a t i o of 4 t o 1. Comparing these c a l c u l a t e d r e s u l t s t o the a c t u a l ambient response l c s s of 2 dB a t 2000 Hz sugges t s t h a t t h e a c t u a l l o s s was approximately 0.5 dB. Since 2000 Hz was t h e upper l i m i t of our s p e c t r a l comparisons and t h e lower f r e - quencies had much lower l o s s e s (1 t o 1.5 dB) any c o r r e c t i o n f o r e leva ted cond i t ions would be i n f r a c t i o n s of a dB and i n t h e o rder of measurement accuracy. Therefore, the ambient response c o r r e c t i o n s were app l ied t o t h e data .

The phase c a l i b r a t i o n s determined f o r t h e waveguide sensors f o r t h e ambient c a l i b r a t i o n s were not app l ied t o t h e d a t a a t e leva ted cond i t ions s i n c e the re was no way of determining the temperature g rad ien t i n t h e sensor tube without e x t e n ~ i v ? measurement. Ins tead , t h e t i m e delay between t h e p a i r s of sensors used i n c ross -cor re la t ion a n a l y s i s was removed which e f fec - t i v e l y accounted f o r the phase d i f f e r e n c e s between t h e p a i r s of sensors . S i m i l a r l y , t h e same cons idera t ion was appl ied during t h e coherence a n a l y s i s between p a i r s of sensors i n determining t h e t r a n s f e r func t ion a c r o s s t h e turbine .

Exhaust Probe Level Correct ions

The exhaust probe K u l i t e A and B s p e c t r a were checked f o r SPL l e v e l and s p e c t r a l shape. Both Probe A and B had similar s p e c t r a l shapes a s would be expected s i n c e t h e K u l i t e s were on t h e same prctbe separa ted by a 12.7 cm a x i a l spacing. The Probe B l e v e l was c o n s i s t e n t f o r t h e major i ty of t h e p o i n t s on both t e s t runs which included repea t reqdings t h a t gave t h e same r e s u l t . The s i g n a l l e v e l s of Probe A matched those of Probe B f o r one run b u t were h igher f o r t h e second run, making t h e Probe A d a t a suspect . An adjustment of -6 dB corresponding t o an a m p l i f i e r a t t e n u a t i o n change was app l ied t o Probe A t o r e a l i g n the d a t a wi th Probe B. The Probe A narrowband s p e c t r a r e f l e c t t h i s cor rec ted l e v e l .

Power Level Ca lcu la t ion

The measured power l e v e l , FPWLmeas was ca lcu la ted a t each measurement plane assuming t h e e n t i r e f l u c t u a t i n g p ressure measurement was a c o u s t i c s i g n a l propagating i n a plane wave a x i a l l y through t h e engine. The power

level was calculated using Blokhintsev's results (as noted in Reference 6), for the acoustic intensity f l u vector which can be written:

+ P, p and c are used in the conventional sense, where V is the absolute flow velocity and Gp the unit vector normal to the acoustic wave front.

We are interested primarily in the axial component, hence

where 8 and $ are the angles made by the acoustic wave front and the flow with the axial direction. M is the flow Mach number. The flow at the measuring planes is near axial and if a plane wave assumption is used here,

The plane wave assumption also permits the acoustic power to be computed from a measurement at any point of the cross-section. Using a consistent set of reference pressure (Po) and specific impedance (poco), the acoustic power level (PWL referenced to 10-l3 Watts) is given by:

P W L = S P L + 2 0 log ( 1 t M ) + 10 log($#)+ 10 logA+9.9 ( 5 )

were SPL = sound pressure level re 2 x 10-5 ~ / m 2

Ps, TS = static pressure and temperature at the measuring etation

Po, To = ambient (standard day) pressure and temperature

A = cross-sectional area in m2.

3.7.3 Engine Test I n t e r n a l Measurements V e r i f i c a t i o n

Checks on Measurement Level - Several c ross checks of t h e d a t a obta ined dur ing t h e CF6-50 engine t e s t

were performed t o ensure t h e v a l i d i t y of t h e f l u c t u a t i n g p r e s s u r e l e v e l s determined. These included on-line 1 0 IIz narrowband s p e c t r a compared wi th t h e same d a t a processed p o s t t e s t through t h e - irrowband s p e c t r a l analyzer ; com- par i son of Ps3 l e v e l s obta ined dur ing emissions t e s t i n g wi th a c o u s t i c t e s t r e s u l t s ; and checks of peak-to-peak p r e s s u r e l e v e l s obta ined on similar engine t e s t s . The l e v e l s agreed i n each case w i t h i n '1.5 dB.

The probe o v e r a l l FPWL ( r e w a t t s ) f o r t h e engine t e s t was compared t o f a r f i e l d power l e v e l s determined from unpublished d a t a from previous t e s t s on CF6-50 engines t o e v a l u a t e t h e reg ion of core n o i s e d o m i ~ ~ x e . The PWL was ca lcu la ted f o r t h e low f requenc ies (50 t o 2000 Hz) a t which core n o i s e would be manifested at each test condi t ion encompassing t h e opera t ing range of the engine. The PWL were p l o t t e d a g a i n s t an e f f e c t i v e jet v e l o c i t y of t h e f a n and core exhaust streams. This v e l o c i t y , Ve was determined from t h e f a n bypass r a t i o BPR and j e t v e l o c i t i e s of t h e f a n and c o r e streams using:

(BPR) Vf + Vc - - ve - BPR + 1

The r e s u l t s of t h e comparison a r e shown i n Figure 3-1.3 and i n d i c a t e t h e region of core n o i s e dominance a t Ve below 230 m / s . Above t h i s v e l o c i t y t h e j e t no i se over takes and dominates t h e o v e r a l l power l e v e l . This t r end is a s expected and p a r a l l e l s c l o s e l y the r e s u l t s of Reshotko (Reference 8).

I n t e r n a l p ressure measurements taken i n t h e core exhaust dur ing t h e p resen t t e s t s wi th t h e sound s e p a r a t i o n probe were obta ined a t f i v e i m e r - s i o n s loca ted on c e n t e r s of equal a r e a s a c r o s s t h e annulus of t - i ~ c o r e ex- haus t nozzle. Comparisons of PWL c a l c u l a t e d from t h e average SPL s p e c t r a frcm t h e f i v e inrmersions agreed very w e l l w i th t h e PWL c a l c u l a t e d us ing t h e p i t c h l i n e or cen te r immersion and t h e t o t a l annulus a r e a i l l u s t r a t e d by t h e power l e v e l s p e c t r a comparison i n Figure 3-14. The comparison i s f o r the approach (30% Fn) condi t ion, b u t is a l s o t y p i c a l of t h e o t h e r test condi t ions . The da ta presented f o r t h e exhaust probe i n t h e a n a l y s i s s e c t i o n of t h i s r e p o r t a r e based on p i t c h l i n e immersion measurements.

Aerodynamic Acoustic Readings

The ae ro performance parameters requ i red by t h e ECCP Phase I11 c o n t r a c t a r e l i s t e d i n Table 5 f o r t h e e i g h t t e s t condi t ions . These parameters were obta,ined from t h e engine performance program read ings obta ined f rorn t h i s ECCP Phase 111 t e s t s . The d e t a i l e d c a l c u l a t i o n procedure f o r t h e major i ty of t h e s e parameters are found i n Reference 1. Other aerodynamic parameters assoc ia ted wi th t h e o v e r a l l power l e v e l c a l c u l a t i o n s a r e presented i n Table 6. Included i n t h e t a b l e a r e t h e s t a t i c p ressures and temperatures determined

CFb-50 Engine Results

a OAPWL From $0 to 2000 Hz

I 0 -Core Probe Result. on ECCP Phase I11

1 0 -~revious Farfield Results, 1972

I L ~ e t Noise

/ Prediction

140 I I I I I 1 1 I

0 50 lo0 150 200 250 300 350 400

Effective Velocity Ve, m/s

Figure 3-13. Comparison of Core Probe Power Levels with Farfield Measurements

a CYb-50 ECCP Phase XI1 Test

Approach Power, 30 qb Fn

OIPWL Probe B Measurements

156.9 dB 0 Pitchline Immersion

156.2 dl3 0 Average cf F i v e Imi~rsionm

130 A ' 1 L I I 1 1 I I 50 80 125 200 315 goc 8001250 2000

Frequency, Hz

Figure 3-14. Core Exhaust Pasor Level Spectra Comparieon

Table 5.

ECCP Phase

I11 Program Aerodynamic Performance Sumnary.

Acoustic Test, 1977

PCT

Acoustic

Test

Net

hero

Tape

Pt.

Thrust Rdg

Rdg

1 3.8

147

6

2 19.7

149

16

3 30.0

151

. 23

4 29.7

119

39

5 45.4

154

28/34

6 64.6

163

13/40

7 82.0

156

18/43

8

94.6

107

33

PCT

Acoustic

Test

Net

Aero

Tape

Pt.

Thruat Rdg

Rdg

1 3.8

147

6

2 3.7

149

16

3 30.0

151

23

4 29.7

119

39

5 45.4

154

28/34

6 64.6

163

13/40

7

82.0

156

18/43

8

94.6

107

3 3

:

-.

NC

Nf

Po

TO

PT3

Ps~ Ti3

W3

W36

kg dK

vr -

rprn

rpm

atm

K atm

atm

kg/sec kgleec sec-atm-m2 m/s

kg!hr

6376

877

0.981

296.7

2.95

2.86

442.8

14.9

12.5

842.7

16.9

699.9

8144

1952 0.978

297.2

8.23

7.96

586.7

39.2

32.9

969.4

21.2

1843.4

8645

2384 0.976

298.3

10.94

10.61

641.1

43.8

36.8

852.7

19.4

2728.4

8661

2393 0.976

298.3

11.48

11.19

642.8

53.6

45.0

995.0

22.2

2698.9

9125

2847 0.974

300.6

15.15

14.69

709.4

67.0

56.2

989.2

23.8

3938.6

9522

3223 0.976

293.9

19.66

19.07

749.4

83.9

70.5

982.8

24.2

5617.4

9878

3506 0.970

301.1

23.38

22.67

808.3

95.5

80.3

977.6

25.0

7315.2

10191

3726 0.971

298.3

26.95

26.27

830.6

105.8

88.5

951.8

25.3

8814.4

APT -

PCT

PCT

PT~ TT4

bP~34 PT3

AT43

HRR

PTg

TT8

Out

In

kgE/kga atm

K atm

PCT

K kw

atm

K 'l b

100

0 0.01555

2.85

1042.8

0.102

3.5

600

27.15

0.987

672.2

58.55

100

0 0.01555

7.89

11'0.0

0.340

4.13

583.3

71.52

1.048

693.3

99.57

100

0 0.02060

10.51

1374.4

0.429

3.91

733.3

105.85

1.096

711.7

99.60

46.9

53.1

0.01664

10.98

1256.7

0.504

4.39

613.9

104.71

1.102

712.8

97.24

17.0 83.0 0.01944

14.46

1403.9

0.687

4.55

694.4

152.80

1.191

763.9

99.65

17.8 87.2

0,02213 18.76

1521.7

0.885

4.48

772.2

217.93

1.313

802.2

99.92

17.5

82.5

0.02530

22.35

1647.2 1.034

4.43

838.9

283.80

1.429

873.9

99.96

12.6

87.4

0.02754

25.76

1733.3

1.184

4.39

902.8

341.96

1.565

908.9

99.96

i b F ( r C ) * U L ? l n l n V I m m m m m s n m v l . . . . . . . . 2V O O O O O O P P

a \ D \ D \ D a 4 d d d 4 . . . . 12 o o o o d o o o

, ? g Z d I $ S ! i j 6 1 W U . . . . .

PIS rn r. s z z z n i l

from i s en t rop i c re la t ionsh ips along with an est imate of the l o c a l Mach number 3 c each measurement plane.

The eight conditions l i s t e d i n the t ab l e s were obtained from two engine runs on consecutive days. Repeat readings taken during the engine test served a dual purpose of es tab l i sh ing a check f o r da ta r epea t ab i l i t y and providing a replacement source f o r sensor measurements l o s t due t o equipment problems a t the same t e s t condition.

The low nominal power s e t t i n g s ( 3.4, 19.8, and 30% Fn) have a f u l l com- pliment of measurements a t a l l planes. Repeat readings were used i n combina- t i on with the i n i t i a l readings taken a t 45, 65, and 85% Fn t o obtain a f u l l s e t of measurements a t each plane f o r these conditions. A t takeoff (100% Fn) and approach power with the 50150 s p l i t f u e l flow (29.7 X Fn), measurements were obtained a t each plane; except f o r the redundant sensor 102' a t Plane 3.5.

C e l l probe measurements were acquired a t a l l test conditions. However, only the data a t 65% Fn and above a r e presented s ince t he l eve l s b e l w t h i s condition were i n the system noise f loor . Cell probe narrwband spec t r a a r e included i n the spec t ra p l o t s of the 65, 85 and 100% Fn poin ts found i n Appendix A.

Comparison of the measurements acquired during the i n i t i a l and repeat readings was exce l len t a s can be a t t e s t e d by the following 2 Hz narrowband FPL spec t r a comparisons f o r the nominal power s e t t i n g s of 45, 65 and 85 percent th rus t .

Figure 3-15a-c shows the MADAR, 282' borescope and exhaust Probe B spec t ra comparisons f o r separa te readings a t 45.6% Fn taken on the d i f f e r e n t days. Agreement is good a t a l l planes compared.

Figure 3-16a and b i l l u s t r a t e s the exce l len t agreement obtained from separate readings a t 64.6% Fn f o r the MADAR Probe and core exhaust Probe B.

A t 82% Fn, t he WAR comparison i n Figure 3-17a and the 282' borescope probe comparison i n Figure 3-17b a l s o show exceptually good agreement a t t he d i f f e r e n t readings. These examples serve t o subs t an t i a t e the r a t i o n a l e f o r combining the i n i t i a l and repeat readings i n order: t o form the 113 OBSPL spec t ra tabula t ions a t the e igh t test conditions.

ECCP LlUSE I11 TEST, 1977

2 Hz Narrowband A s Measured Spectra . - -- -

RDG 28 RDG 34

fs e

130 u\ I

2 I10

Q, Ll a. Plane 3.0, Madar Probe

b. Plane 3.5, 282' Borescope Probe

Frequency, Hz

c. Plane 8.0, Exhaust Probe B

Figure 3-15. Comparison of Caamon Measurements For Repeat Readings at 45.4 Percent Thrust

-- -- -- -.------ ... -- 1_-. -

. . > & 2- - ..* ' - . I... L y r . , , - . 4 I

ECCP PHASE I11 TEST,1977

2 Hz Narrowband A s Measured Spectra

a. Plane 3.0, Madar Probe

RDG 40 .

r ! 1 1 1 n D * s i ~ 1 1 1 ~ ~ r v r 4 loo0 2030

b. Plane 8.0, Exhaust Probe B

Figure 3-16. Comparison of Common Measurements For Repeat Readings a t 64.6 Percent Thrust

ECCP PHASE I11 TEST, 1977

2 Hz Narrowband As Measured S~eatrs

In I

: 110 0,

i a. Plane 3.0, Madar Probe

2

RDG 43 -

r q - - - ~ m v - ~ ~ - r m ~ H ~ 0 ?m 0 lo00 2000

Frequency, H z

b. Plane 3.5, 2820 Borescope Probe

Figure 3-17. Comparison o f Caarmon Measurements For Repeat Readings a t 82.0 Percent Thrua t

4.0 ANALYSIS AND DISCUSSION OF RESULTS

4.1 SPECTRA COMPARISONS

Comparisons of the f l uc tua t ing pressure measurements a t each measurement plane a r e discussed fo r the engine t e s t . Results a r e based on the measured pressure da ta analyzed a s 113 octave band and narrowband spectra . Frequency response correct ions were applied t o the 113 OBFPL, but not t o the narrowband FPL .

The Double Annular combustor r e s u l t s from duct r i g tests conducted during ECCP Phase I and I1 were compared t o se lec ted po in ts from the ECCP Phase I11 engine t e s t s .

4.1.1 Measurement Plane Comparisons of 113 Octave Band and Narrowband Spectra

The i n t e r n a l f l uc tua t ing pressure da ta acquired during the t e s t of t h e Double Annular combustor i n the CF6-50 engine were compared on both a 113 octave band a 2 Hz narrowband spec t r a bas i s . Comparisons were made t o iden- t i f y and inves t i ga t e t rends a t each measurement plane fo r severa l test con- d i t i o n s covering the f u l l operat ing range of t h e engine. The spec t ra en- compass frequency ranges f o r core noise ana lys i s from 50 t o 2000 Hz i n 113 octave band and from 0 t o 2000 Hz i n 2 Hz narrowband. The complete set of narrowband spec t r a f o r a l l t e s t condi t ions is found i n Appendix A. Similar ly , t h e 113 octave band da ta is presented i n both tabular and graphical form i n Appendix B.

Five of t h e e igh t test conditions were se lec ted fo r t h i s ana lys i s as represen ta t ive of combus t o r operat ion over t he engine operat ing l i n e . These f i v e conditions include the percent ne t t h r u s t s e t t i n g s of 3.8 ( i d l e ) , 19.7, 30 (approach), 64.6 and 94.7 ( takeoff) percent. The f i r s t t h r ee conditions; 3.8, 19.7 and 30 percent ne t t h r u s t a r e with p i l o t f u e l only, while the 64.6 and 94.6 percent ne t t h r u s t po in ts a r e with f u e l s p l i t s between p i l o t and main burner s e c t ions.

A common s e t of four sensors were used i n the comparisons which were se lec ted on the bas i s of measurement r e l i a b i l i t y and a v a i l a b i l i t y . The set included one probe a t each measurement plane and consis ted of the MADAR probe a t Plane 3.0, t he 282' borescope probe a t Plane 3.5, t he HPTN borescope probe a t Plane 4.0 and the core exhaust probe, element B, on the sound separat ion probe located a t Plane 8.0. The comparisons include the p i l o t f u e l s p l i t s a s a percentage of the p i l o t f u e l t o t o t a l f u e l flows (PCT OUT).

Plane 3.0 - Compressor Discharge

The 113 octave band s p e c t r a comparison f o r t h e MADAR Probe i n Figure 4-1 show a peaked shape a t i d l e along wi th a predominate tone a t 400 Hz. The f l u c t u a t i n g pressure l e v e l (FPL) i n c r e a s e s uniformily by approximately 10 dB over t h e i d l e s e t t i n g f o r 19.7 and 30 percent t h r u s t points . The tone a t 400 Hz i s no longer a c h a r a c t e r i s t i c of t h e s p e c t r a l shape a t these condi t ions . A t t h e h igher power s e t t i n g s of 65 and 94.6 percent t h r u s t t h e pressure l e v e l s inc rease by 4 t o 5 dB over those a t 30 percent t h r u s t i n t h e low frequencies , but show a s i g n i f i c a n t inc rease of 10 dB aL f requencies above 800 Hz. The FPL inc rease i n t h e h igher f requencies tends t o f l a t t e n t h e s p e c t r a shape. The peaked s p e c t r a , c h a r a c t e r i s t i c of co re no i se dominated s p e c t r a , seem t o be l imi ted t o power s e t t i n g s a t approach and below.

The r e s u l t s of t h e 2 Hz narrowband s p e c t r a a n a l y s i s of these cond i t ions a t Flane 3.0 a r e shown i n Figure 4-2. The predominate tone a t 400 Hz f o r t h e i d l e condi t ion appears t o be a low l e v e l resonance o r combustor growl phenom- enon which i s no t apparent a t o t h e r cond i t ions but i s p resen t a t o t h e r p lanes f o r t h e i d l e condi t ion. There i s a s l i g h t reduct ion i n combustor e f f i c i e n c y (see Table 5) of about one point a t i d l e compared t o t h e 19.7 and 30 percent t h r u s t cond i t ions wi th p i l o t f u e l only. However, i t does not appear t h a t t h i s amount of change would be s u f f i c i e n t t o cause the l a r g e 400 Hz tone observed i n t h e da ta . There is some e l e c t r o n i c noise a t 60 and 180 Hz. A "hay- stacked" torre is evident a t 1 9 . 7 and 30 percent t h r u s t occurr ing a t approximately 700 Hz which corresponds roughly t o the 5 th per rev of the engine spool. A ;road 5 t o 7 dB p l a t e a u i s a l s o apparent between 700 and 900 Hz f o r these condi t ions . A t t h e h igher power s e t t i n g s the narrowbands appear q u i t e s i m i l a r and show an inc rease i n the broadband above 1000 Hz which re- f l e c t s t h e 1 / 3 octave band s p e c t r a l r e s u l t s .

P lane 3.5 - Comb~~stor In leL

A t t he combustor i n l e t , Plane 3.5, t h e combustor borescope probe a t t h e 282' p o s i t i o n e x h i b i t s s i m i l a r t r ends i n t h e 113 octave band s p e c t r a (Figure 4-3) a s were shown f o r the MADAR Probe a t Plane 3.0. Peaked s p e c t r a (400 t o 800 Hz) a r e evident f o r power s e t t i n g s of 30 percent t h r u s t and below. The high power p o i n t s show an inc rease i n FPL of 5 t o 10 dB above 1000 Hz and t h e 400 Hz tone a t t h e i d l e condi t ion is s t i l l apparent .

The narrowband a n a l y s i s show s p e c t r a i n Figure 4-4, f o r 3.8, 19.7 and 30 percent t h r u s t which compare w e l l wi th s i m i l a r s p e c t r a f o r t h e Plane 3.0 probe i n Figure 4-2. Large p ressure f l u c t u a t i o n s of t 10 t o t 18 dB a r e seen below 500 Hz. There is evidence of e l e c t r o n i c no i se contamination a t 60 and 180 Hz which does not compromise t h e q u a l i t y of t h e d a t a but might c l a r i f y a por t ion of these l a r g e f l u c t u a t i o n s below 200 Hz. A p o s s i b l e explanat ion f o r t h e remainder might be due t o t h e l o c a t i o n of t h e probes. They a r e loca ted i n t h e pre-mix region of t h e combustor j u s t a f t of t h e f u e l nozzle d ischarge where f u e l i n j e c t i o n and i g n i t i o n take place which r e s u l t s i n a high r a t e of energy r e l e a s e . Note t h a t t h e s p e c t r a f o r 64.6 and 94.6 percent t h r u s t p o i n t s a r e f l a t i n the high f requencies a s opposed t o t h e s p e c t r a l drop o f f a t lower power s e t t i n g s , This is c o n s i s t e n t wi th the 113 octave band r e s u l t s .

170- ECCP P W E IS5 COnBWITOR TEST +

PLANE 9.0, NADCIR PROBE

Pilot I Fuel Sp l i t 1

PCT Fn PCT Out I

Figure 4-1. Plane 3 . 0 (MADAR Probe) 113 Octave Band Spectra Variation with Engine Speed

_. -_ . -. - -.--- - --- - - - -

94.6 % Fn 150 N = 10191 rpm

C

PCT OUT = 12-6

64.6 '% Fn

N~ = 9522 rpm

PCT OUT = 17-8

3C % Fn

N = 8645 rpm 0 C Li

130

C

JQ.7 1 F,,

N~ = 8144 rpo

K T OUT = 100 1

N 150

Z CU

130

3.8 % Fn N = 6376 rpm

C

pcT OUT = 100

I 1

500 lo00 1500 ZQOo 0

Frequency, Hz

Figure 4-2 Plane 3 - 0 (MALUR Probe) Spectra Variation with Engine Speed

- ECCP P H ~ E 111 ConstlsroR TEST

PLANE 3.6. 262 DEOUEE BORESCOPE PROBE

Pi lo t Fuel Spl i t

PCT Fn PCT Out

Figure 4-3. Plane 3 .5 (282O Borescope Probe) 113 Octave Band Spectra Variation with Engine Speed

Nc = 10191 rpm

X'I OUT = 12.6

130

0 64.6 % F n i

N = 9522 rpm I C

OUT = 17.8 ,

I 1

jo :% F n

I i 8 6 4 5 r p m

PC1 OUT = 100 i I I

'CI j 2 13Q - I P $ 1 5 0 1

I

I

19.7 % F i

n Nc = 8144 rpm

N PCT OUT = 100

13'3

140 ! I

3.8 % Fn ? i I

120

110 - 1- ;-

Frequency, Hz

Figure 4-4 Plane 3.5 (282O Borescope Probe) Spectra Variation with Engine Speed

4 4

Plane 4.0 - Combustor Discharge

The HPTN borescope probe a t Plane 4.0 e x h i b i t s inc reas ing s p e c t r a f o r a l l cond i t ions a s shown i n Figure 4-5. This sensor does no t i n d i c a t e a core no i se spectrum shape a t any of t h e t e s t cond i t ions and appears t o be s a t u r a t e d wi th high frequency noise . This may be a consequenp2 of t h e probe l o c a t i o n between two vanes i n t h e high p ressure t u r b i n e nozzle diaphram.

The 2 Hz narrowband r e s u l t s presc:~ted i n Figure 4-6 show f l a t s p e c t r a t o 2000 Hz and high p ressure f l u c t u a t i o ~ ; of + 1 0 t o + 12 dB i n t h e same region a s the Plane 3.5 sensor. The h ighes t l e v e l s a r e recorded a t t h i s probe loca t ion . This agrees wi th t h e physics of t h e combustion process which is culminated a t t h e combustor e x i t a f t e r complete mixing and burning of t h e f u e l - a i r mixture have taken place. It is at t h i s plane t h a t t h e maximum energy is a v a i l a b l e t o do work on :he turbine .

Plaae 8.0 - Core Nozzle Discharge

The exhaust Probe B 113 octave band l y e c t r a shown i n Figure 4-7 e x h i t i t t y p i c a l core n o i l e shapes between 200 and 1250 Hz f o r t h e 3.8, 19.7 and 30 percent t h r u s t points . A t t h e higher power s e t t i n g s t h e high frequency n o i s e (above 800 Hz) increased t o produce an e s s e n t i a l l y f l a t spectrum a t t h e take- o f f condit ion. Below 200 Hz t h e FPL f o r a l l cond i t ions , except i d l e , is genera l ly about t h l same level., wi th in + 5 dB. This may be t h e r e s u l t of low frequency tdrbulence generated off t h e c e n t e ~ h o d y of t h e f a c t o r y plug nozzle.

The narrowband s p e c t r a f o r Probe B a t t h e exhaust p lane is shown i n Figure 4-8. The 400 Hz tone a t t h e i d l e cond i t ion is again c l e a r l y v i s i b l e . The s p e c t r a below 30 percent t h r u s t a r e dominated by t h e low f r e q u e n c i e s . (el000 Hz). High f r e q ~ e n c y n o i s e above 1000 Hz is increased a t t h e power s e t t i n g s above 30 percent t h r u s t a s i l l u s t r a t e d i n Figure 4-8.

4.1.2 Double Annu* Conbustor Spcztra Comparisons

Comparisons were made between t h e Double Annular combustor s p e c t r a measured on t h e engine t e s t and t h e combustor s p e c t r a obtained from f u i l annular duct r i g tests conducted dur ing Phase I and I1 of t h i s program. The engine condi t ions a t i d l e (3.8% Fn), approach (36% Fn wi th 100% p i l o t f u e l and 29.7% In with 50% f u e l s p l i t ) and takeoff ( 9 4 . 6 2 Fn) were s e l e c t e d f o r comparison with t h e duct r i g data .

A number of cond i t ions approximating t h e above er . : ? power s e t t i n g s were a v a i l a b l e f o r cornpatison from t h e duc t r i g da ta . : e se lec t io i l of t h e p a r t i c u l a r duct r i g t e s t p o i n t s f o r comparison wi th t h e engine d a t e was based on matching t h e i n l e t and discharge temperatures i n c ~ n j u r c t i o n wi th keeping w ~ / P ~ constant f o r t h e approach and takeoff points . The i d l e set po in t s e l e c t i o n was based on t h e i n l e t p ressure and temperature match. Considera- t i o n was a l s o given t o matching t h e f u e l - a i r r a t i o s between duc t r i g and

- ECCP PWIIE I I r COHBU~TOR TEST

- PLRHE 4-01 HPTN PROBE

P i l o t Fuel S p l i t

PCT F~ PCT nut

a 3.8 100 0 19.7 loo A 30.0 LOO + 64.6 17.8 a 94.6 12.6

l L B . . . l . . I 1 1 1 1 . 1 1 . 1 . + - - - : . . I * 1

50 80 ie6 POO 31s 600 800 teso 0000 FREQUENCY. HZ

Plane 4 .0 (HPTK Probe) 113 Octave Band Sp. ,ctra Variatl.on with Engine Speed

lbo

Nc = 10191 rpm

140 .

a . 6 76 Fn

Nc = 9522 rpm I

m' OUT = 17.8

L? I

2 lbO 1 30 % F~ N = 8645 rpm

C

PC'I' OUT = 100

19.7 % Fn

N = 8144 rpm C

150 PC:' OUT = 100

140

3.8 % F n Nc = 6376 rpm

I I

Frequency, Hz

Figure 4-6 Plane 4.0 (HPTN Probe) Spectra Variation with Engine Speed

ECCP PW8E 5 5 5 CUlWSTdR TEST PLCIIIE 8-0. EXHW8T PROBE 0

P i l o t Fuel S p l i t

PCT Fn PCT O u t

o 3.8 IOO O 19.7 100 a 30.0 KIO + s . 6 17.8 0 94.6 12.6

Figure 4-7. Plane 8 .0 (Exhaust Probe B) 113 Octave Band Spectra Variation with Engine Speed

4 9 4 . 6 % F n

K T OUT = 12.6

I I

PCT OUT =

3.8 % Fn

Nc = 6376 rpm

PC'T OUT = 100

-l-----l- -

0 500 lo00 1500 3000

Frequency, Hz

Figure 4 - 0 . Plane 8.0 (Exhaust Probe B) Spectra Variation with Engine Speed

engine within t h e l i m i t s of da t a a v a i l a b i l i t y . Table 7 lists the compared t e s t points f o r these conditions.

The s e t up f o r t he duct r i g tests is de t a i l ed i n t he ECCP Phase I and 11 f i n a l repor t s (References 3 and 5). The acous t ic probe loca t ions f o r these t e s t s were a s i l l u s t r a t e d i n Figure 4-9. The engine and duct r i g acous t ic measurement plane loca t ions were not i den t i ca l a s noted i n reviewing the engine plane loca t ions i n Figure 3-2.

The i d l e point 113 OBFPL s p e c t r a l comparison is shown i n Figure 4-10 f o r t he combustor discharge sensors. No upstream probe measurements were ava i l - ab l e f o r the duct r i g a t t h i s condition. The engine da ta a r e 17 t o 20 dB grea te r than the duct r i g d a t a f o r the frequency range invest igated.

Figures 4-lla and 4-llb present t he combustor i n l e t and discharge f luc- tua t ing pressure l eve l comparisons respec t ive ly f o r the approach condition. A t the combustor i n l e t (Figure 4- l la ) , t he Plane 3.0 engine probe shows r e s u l t s t h a t a r e from 5 t o 12 dB g rea t e r than the duct r i g i n l e t probe f o r both the 100% p i l o t and f u e l s p l i t approach condition. Larger d i f fe rences a r e noted f o r t h e combustor discharge comparison i n Figure 4-llb with t he p i l o t only f u e l than with t he f u e l s p l i t point . The engine Plane 3.5 sensor shows a c h a r a c t e r i s t i c core no ise s p e c t r a l shape peaking between 500 and 1000 Hz a t both approach conditions.

A t takeoff the Plane 3.0 engine and dac t r i g pressure l e v e l measurements i n Figure 4-12a a r e within 1 t o 4 dB a t frequencies above 315 Hz. Below 400 Hz the engine measurements a r e 4 t o 15 dB g r e a t e r than those of t he duct r ig . The discharge Plane 4.0 measurements i n Figure 4-12b show the same t rends observed f o r t he i d l e and approach point comparisons. The Plane 3.5 engine borescope spectrum shown i n t he f i g u r e is 10 t o 15 dB g rea t e r than t h e duct r i g data .

These engine t o duct r i g 113 OBFPL s p e c t r a l comparisons i nd i ca t e l a rge l e v e l d i f fe rences a t both i n l e t arid discharge measurement planes. The inlet plane comparisons a r e much c lo se r (5 t o 10 dB) than those a t t he discharge (15 t o 20 dB). Some of t he d i f f e r ences can be explained due t o test condi- t i o n d i f fe rences a s noted i n Table 7. The l o c a l s t a t i c pressures i n the engine (Plane 4.0) a r e approximately 2.5 t i m e s g rea te r than t h e duct r i g s t a t i c pressures f o r t h e condi t ions above i d l e .

In order t o minimize these d i f fe rences t h e duct-rig and engine 113 OBFPL spec t ra were converted t o power spec t r a assuming a l l t h e neasu c?d s igna l was acous t ic and plane wave propagation occurred a t each measurement plane. The F P b e p s w a s ca lcu la ted using Equation 5. Comparisons of t he spec t r a from the f l uc tua t ing power l eve l ( F w e a s ) showed much c loser agreement between engine and duct r i g data . Figure 4-13 shows t h e meas spec t r a comparison a t Plane 4.0 f o r t h e i d l e condition. The engine spec t ra is approximately 8 dB above t h e duct r i g r e s u l t s ( so l id symbols) f o r f requencies above 160 Hz where t he spec t r a l shapes a r e s imilar . The Plane 3.5 engine r e s u l t s (from the 282' borescope sensor) show good agreement from 160 t o 500 Hz with t he duct r i g data . However, above 500 Hz t h e duct r i g da t a is 6 t o 10 dB higher than t h e engine data .

W \Dm rnr-lco ,-I9 n U B N Q O TyCJ! O h W etu . . . .

& a : mcu m o o mtn u 4 d N

- m m C i PI 0 O""0 0 9 I

w r n N A G o I H , N l j a u u 4 m a0 9 - J 9\09 Q O Q )

U t

m ~ l m e a ~ m t n urn m m u VlQI . . . . . . . m h l m o d rn\O

A 4 N

h U h 0 U l-l 0 .I4 4 P4 4

i Ch ! bP

0 0- 1 0 0 i FI ~l I Q W

Table 8.

CF6-50 Turbine Attenuation Su

mmar

y from ECCP Phase I11 Test Results.

Results Based on Difference Between Coherent Spectra at Planes 4.0

and

8.0

Relative to Plane 3.5

-

94.6

107/33

&SF

%

bPk'

L

---

---

---

-- -

>32

> 9.5

---

-- --

- --

- 30

7.5

30

7.5

~29 <

G

.5

-- --

31

8.5

-- -

---

-- ---

-

---

-- 7.9

Note:

< Actual Attenuation is Less than Value Shown

> Actual Attenuation 18 Greater than Value Show

82.0

156118

ASPL

APWL

>26

. 4.4

~16.5

5.1

25

3

.4

>30

> 8.4

-- ---

>36

>14.4

<27

C 5.4

<28

< 6.4

---

-

33

11.4

-- --

-- --

-- --

-

---

.- - 7.4

64.6

163140

ASP

L

APWL

~19 <-1.3

---

---

34

13.7

>35

>14.7

--- -

---

---

---

-

37

16.7

35

14.7

34

13.7

---

-- .--

---

--

-- --

-- 14.7

45.4

154/28

ASPL

APWL

---

---

>21

>2.4

4

<-4.6

---

-

---

-- 27

8.4

---

-- 35

16.4

30

11.4

30

11.4

-- ---

---

-

---

---

---

---

10.0

29.7

119/39

ASPL

APWL

52

6

> 9.4

.23.5

>6.9

28

11.4

~24 <7.4

---

---

23

6.4

29

12.4

---

---

39

22.4

~35 ~18.4

---

---

---

---

---

---

---

---

11.8

30.0

151/23

ASPL

APWL

18

1.4

23

6.4

20

3.4

e27

<10.4

23

6.4

19

2.4

<18.5

< 1.9

25

8.4

27

10.4

31

16.4

<34

~17.4

---

---

--- -

---

---

7.5

19.7

149116

ASPL

APWL

<I9

< 4.2

22.5

7.7

26

11.2

30

15.2

19

4.2

13.5

-1.3

~16

< 1.2

<22

< 7.2

31

16.2

30

15.2'

-- ---

---

--

---

---

---

-- 9.1

Pct Thrust

Aero /Tape

Reading

113 OB O

ASP

L

Frequency

(Hz)

100

125

163

200

2 50

315

400

500

630

8 00

1000

1250

1600

2000

Average

A~

BP

~JL

3.8

147/6

ASPL

APWL

16

6.9

19

9.9

19

9.9

<20

~10.9

25

15.9

<13

< 3.9

14

4.9

24

14.9

~22.5

~13.4

~27.5

<18.4

33

e23.9

<35

q25.9

<31

e21.9

-- -- 13.9

170- ECCP CUWUBTOR TEBf CONPCIRIBON8 A rrmY 11-71a1 IDLE M U S T CONDITION

4.00

Discharge Plane

'Id bb ' 00 L . L I . L . L . I . b . . l . . . . . . i26' $00' $16' doo ' 1d60 ' ld00 '

FREQUENCY l HZ

Figure 4-10. Engine-to-Duct Rig Pressure Spectra Comparison for Double Annular Combustor at Idle.

ECCP COtlBUBTOR fE8T COHPRR18ONB RPPROACH THRUBT CONDITION

a) Inlet Plane

1 2 S , 1 1 1 1 r t ' I ~ ' t I 1 1 " ' I

"1 **\ b) Discharge Plane

Figure 4-11 Engine-to-Duct Rig Pressure Spectra Comparison for Double Annular Combustor at Approach

ECCP COHSUBTOR TEGT COflPflRISON8 TflKEUFP THRUST CONDITION

a) Inlet Plane

b) Discharge Plane

Figure 4-12 Engine-to-Duct Rig Pressure Spectra Comparison for Double Annular Combustor at Takeoff

ECCP COHDUSrOR TEST COHPARISON~- A PHASE 11-713 IDLE THRU6T CONDITION 4.0, 0 PHASE 111-6

4101 HPlW I

5

Discharge Plane

lf0l ! i0 ! 6 do ! As! Qbo! dbo! ! FREQUENCY. HZ

Figure 4-13 Engine-to-Duct Rig Power Spectra Conpariaon for Double Annular Combustor at Idle.

The approach power s e t t i n g FFWLea, comparisons a t t h e i n l e t plane a r e shown i n Figure 4-14a. Agreement wi th in 3 t o 5 dB is apparent i n frequency regions between 160 t o 250 Hz and above 630 Hz f o r t h e duct r i g r e s u l t s wi th both the 30% Fn (100% p i l o t f u e l ) and 29.72 Fn (50/50 f u e l s p l i t ) po in t s .

The e x i t plane comparison Eor t h i s a p p r ~ ~ a c h condition shown i n Figure 4-14b i n d i c a t e s good agreement i n both s p e c t r a l shape and l e v e l between t h e duct r i g r ? s u l t s and t h e engine Plane 3.5 da ta . The Plane 4.0 r e s u l t s from t h e engine a r e g r e a t e r than t h e duct r i g measurements by approximately 5 t o 15 dB f o r t h e 30% Fn po in t and zero t o 8 dB f o r t h i s 29.7% Fn condi t ion.

A t t akeof f , t h e i n l z t plane FPWL,eas comparison i n Figure 4-15a shows t h e duct r i g d a t a t o be 1 0 t o 12 dB higher than t h e engine Plane 3.0 power l e v e l above 400 Hz. The shape of t h e s p e c t r a appear q u i t e s i m i l a r over t h i s f u l l frequency renge.

Figure 4-15b s i ~ ~ x z t h e e x i t p lane comparison a t t h e takeoff condi t ion. The duct r i g F P G e a s s n e c t r a matches t h e engine Plane 3.5 r e s u l t s below frequencies of 315 Hz, ~ h i l e a t t h e higher f requenc ies t h e Plane 4.0 d a t a appears t o g i v e a b e t t e r natch.

The comparisons of t h e engine and duc t r i g r e s u l t s on a f l u c t u a t i n g power l e v e l s p e c t r a b a s i s show genera l ly c l o s e r agreement than on a 113 OBFPL s p e c t r a l b a s i s as previously noted. The shapes of t h e engine s p e c t r a a r e very s i m i l a r t o t h e duct r i g s p e c t r a a t a l l p lanes inves t iga ted . The Plane 4.0 HPTN probe located i n t h e engine t u r b i n e nozz le g i v e s c o n s i s t e n t l y h igher f l u c t u a t i n g p ressure l e v e l s than o t h e r engine probes. This may be due, i n p a r t , t o t h e presence of t h e choked t ~ r b i n e nozzle diaphragm which s u p p l i e s an e n t i r e l y d i f f e r e n t end e f f e c t than t h e open plenum of t h e duct r i g f a c i l i t y .

The d i f f e r e n c e s i n measurement p lane l o c a t i o n , t h e t u r b i n e nozzle diaphragm and t h e v a r i a t i o n s i n t e s t cond i t ions a l l c o n t r i b u t e t o t h e d i f f e r e n c e s i n l e v e l s apparent between engine and duct r i g . However, t h e o v e r a l l engine- to-duct r i g co;nparisons show good r e s u l t s .

A more ~ t ~ n t r o l l e d t e s t , such a s a one-on-one t e s t between t h e engine and duct r i g us-w, t h e same combustor, dynamic ins t rumentat ion, ins t rumentat ion l o c a t i o n s , and a t u r b i n e nozz le diaphragm con£ ig i l ra t ion i n t h e duct r i g t e s t would provide a b e t t e r comparison.

4.2 TURBINE TRANSFER FUNCTION COMPARISON

The t u r b i n e a c o u s t i c t r a n s f e r f u n c t i o n s determined f o r t h e ECCP Phase I11 Double Annular combustor were c a l c u l a t e d i n t h e form of b lade row at tenu- a t i o n s a c r o s s t h e t u r b i n e us ing coherence a n a l y s i s techniques. I n i t i a l l y , a t t e n u a t i o n s were determined from t h e d i f f e r e n c e i n t h e Plane 4.0 HPTN probe raw s p e c t r a wi th t h e coherent s p e c t r a of t h e Plane 8.0 core probe a t t h e core nozzle discharge. This procedure proved unsuccessful due t o t h e excess ive amount of h igh frequency n o i s e i n t h e raw s i g n a l of t h e Plane 4.0 sensor which was incoherent wi th t h e downstream c o r e probe s i g n a l . When t h e coherent

ECCP COHBUSTOH TEST COHPRR18CN8 RPPRORCH THRUGT CONDITION

a ) I n l e t Plane

~ ~ ~ + ' I I I I I J I I I J 17--..

4 P l l M 1-468 4.0.

O WR3E 111-29 $.I;. 80R8CPE

0 PURGE I11 -2s 4.0. nprn

t> PHAsE Ilk39 160.- 4.0. HPTN

0 PHASE 111-39 3.5. bORSCP

160.-

b) Discharge ~ i a n e I

Figure 4-14 Engine-to-Duct Rig Power Spectra Comparison f o r Double Annular Combustor a t Approach

------ -- 'I, ,- ---...

- - . ,,r ' W' " I

ECCP ~ 0 t i 9 ~ 8 1 0 d TEST COlrPRR180N8 TRKEOFF THRbST CONOXTION

b)Discharge Plane

S

Figure 4-15 Sngine-to-Duct Rig Power Spectra Comparison for Double Annular Combustor at T~keoff

spec t ra of t h e d m s t r e a m probe was compared t o t h e Plane 4.0 raw s p e c t r a t h e r e s u l t i n g a t t e n u a t i o n amplitudes were excess ive ( i n t h e o r d e r of '5 t o 50 dB). The a n a l y s i s technique employed d i d not lend i t s e l f t o a d i r e c t comparison of t h e u p s t rcam raw s i g n a l wi th t h e downstream coherent s p e c t r a .

'ro m ~ n i r n i z c this t.f f c r t , t h e a t t e n ~ a t i o n was obta ined from the d i f f e r - ence i n coherent 113 OBSPL s p e c t r a a t Planes 4.0 and 8.0 r e l a t i v e t o t h e upstream combustor s i g n a l a t Plane 3.5. The coherent s p e c t r a between p a i r s of sensors incorporated p o s i t i v e time de lays determined from c ross -cor re la t ions t o account f o r phase d i f f e r e n c e s between sensors . I n c a s e s where nega t ive time de lays were observed (which ind ica ted t h a t t h e s i g n a l s were t r a v e l i n g opposi te t o t h e assumed d i r e c t i o n ) , a ze ro t ime de lay was inpu t . The e r r o r introduced by t h i s procedure r e s u l t e d i n coherent s p e c t r a l e v e l s which were 0 .1 t o 0.6 dB h igher than t h e t r u e l e v e l . Adjustments were made t o t h e r e s u l t i n g spectrum t o account f o r t h e s e d i f f e r e n c e s . The coherent s p e c t r a determined i n t h i s manner had some reg ions t h a t were a f f e c t e d by t h e n o i s e f l o o r a s soc ia ted with t h e s n a l y s i s procedure f o r determining t h e coherent spec t ra . This f l o o r , based on sample length l i m i t a t i o n s which r e s t r i c t e d t h e number of averages used i n t h e a n a l y s i s , w a s 18 dB down from t h e raw s i g n a l l e v e l . These regions were loca ted i n t h e h igher f r equenc ies above 800 Hz and i n a low frequency region around t h e 250 Hz band a t t h e h igher t e s t condi- t i o n s above 30 percent t h r u s t . Previous a t t e n u a t i o n r e s u l t s (Reference 6) covered a range of f requencies from 100 t o 1209 Hz.

4.2.1 Turbine At tenuat ion from Coherent Spec t ra -- The a t t e n u a t i o n s determined f o r t h e ECCP Phase I11 d a t a from t h e 113

OBSPL coherent s p e c t r a a t Planes 5.0 and 8.0 r e l a t i v e t o t h e borescope Plane 3.5 were converted t o PWL a t t e n u a t i o n s t o account f o r impedance changes, Mach number r f f e c t s and a r e a d i f f e r e n c e s a t t h e measurement p lanes .

For cond i t ions above 30 percent t h r u s t , t h e increased probe turbulence i n the lower f requencies and incoherent h igh frequency n o i s e produced coherent s p e c t r a a t Plane 8.0 t h a t were i n t h e n o i s e f l o o r f o r s e v e r a l of t h e 113 octave bands. Consequently, t u r b i n e a t t e n u a t i o n s a t t h e s e cond i t ions were obtained a t about 60 percent of t h ? frequency bands up t o 800 Hz. The a t t r n u a t i o n s determined froin t h e d i f f e r e n c e s i n coherent s p e c t r a a r e s m a - r i zed i n Table 8 f o r each test condi t ion. Bath t h e coherent SPL and PWU a t t e n u a t i o n s a r e l i s t e d a t t h e 113 oc tave band c e n t e r f r equenc ies from 100 t o 2000 Hz. The missing a t t e n u a t i o n s i n d i c a t e both coherent s i g n a l s were i n t h e no i se f l o o r . Where e i t h e r one o r t h e o t h e r spectrum was inf luenced by t h e r.oise f l o o r , t h e a t t e n u a t i o n l i s t e d i s l e s s than o r g r e a t e r than t h e va lue shown, a s noted on the t a b l e . Appendix C p r e s e n t s t h e coherent s p e c t r a at Planes 4.0 and 8.0 r e l a t i v e t o Plane 3.5 f o r each t e s t cond i t ion used t o determine t h e a t t e n u a t i o n ac ross t h e tu rb ine .

Typ ica l ly , . h e a t t e n u a t i o n s p e c t r a show an S-shaped d i s t r i b u t i o n a s indicated i n Tigure 4-16 f o r 30 percent t h r u s t . This is similar t o component tu rb ine a t t e n u a t i o n r e s u l t s b u t i r x l u d e s any change i n a t t e n u a t i o n i n prop- a g a t i ~ ~ g from t h e t u r b i n e e x i t t o nozz le d ischarge . Average a t t e n u a t i o n s

computed from the arithmetic average of the A d B p w ~ l ~ at each frequency for all conditions ranged from 7.4 to 14.7 dB as listed in Table 8.

4.2.2 Comparison of Engine Results with Component Tests --

Component tests were conducted on a single stage high pressure turbine and on a low pressure turbine tested in both a 1- and 3-stage build. A siren was used to similate combustor noise. The turbine attenuations obtained from these component tests conducted under another NASA Lewis sponsored program and reported on in Reference 6 were based on AdBpw~ determined from raw spectra upstream of the turbine to coherent downstream spectra. Average attenuations were obtained for each configuration in the same manner as the engine test. The freque~cy range for the average AdBpw~ was between 100 to 1200 Hz.

The average attenuation (AdBpn) for the 3-stage low pressure turbine ranged from 5.9 to 15.5 dB over the operating range tested as noted in the reference. These attenuations are of the same order of magnitude as determined from the engine test. The 3-stage low pressure turbine results used in the comparison were based on attenuations of siren tones at warm ambient inlet conditions to subambient coaditions at discharge. The Phase 111 results are from an engine run at SLS operating line conditions.

The shape oL the attenuation spectra for all component tests was similar and exhibited the saine characteristic S-shape trend seen for the engine spectra. Figure 4-17 (from Reference 6) shows a typical example of the attenuations obtained from the component turbine test using siren tones. A comparison of this figure with Figure 4-16 shows close agreement between engine and component turbine attenuations, both in level and spectral shape. Variations of 3 to 5 dB in the component turbine attentuations based on tones at the same frequency were com:n in the previous test results. The engine attenuations, based on broadband results, match those of the component test with a similar data spread over the frequency range.

4 .2 .3 Comparison of Engine Results with Theory

The theoretical prediction of turbine transmission loss (attenuation) for supersonic blade rows as p~esentcd in Reference 9 was used to determine the attenuation for the 6-stage turbiile on the CF6-50 engine. The takeoff design point condition was used in the analysis because of the availability of cycle conditions, but is considered reprzsentative of the other conditions since the turbine operates near choke for conditions above idle. The prediction program was set up following the recommendations noted in Reference 9 for the CF6-50 turbine. The attenuation results of the theoretical predic- tions are shown in comparison with the engine results on Figure 4-16. A maximum of 8.54 dB attenuation is recorded at 630 Hz after reaching the first cut-on frequency of 570 Hz. The attenuation drops off to 7.12 and 6.72 dB at 800 and 1000 Hz respectively. Prior to cut-~n, no significant attenuation (1.43 dB) is computed.

ECCP PHASE 111

I - T - - l - l l - r

Less Than Shown

ased on Coherent Spectra

pproach ~ ~ n d i t i o n

- 10 50 100 200 500 lo00 2000

Frequency, Hz

Figure 4-16 Meaeured Attenuation Spectrum and Predicted Resu l t s For CF6-50 S i x Stage Turbine

50 100 500 1000

Frequency, Ilz

Figure 4-17 ~t tenur t t ion Spectrum from +)-Stage Low Pressure Turbine

These r e s u l t s :,re s i n d l a r t o those obtained on the engine i n t ha t a t tenuat ions above :iOO Hz a r e general ly higher than those at t he lower frequencies. The magnitude of t he a t tenua t ions determined from the theore t i - ca l model a r e c lose t o those obtained from the engine t e s t . The engine re- s u l t s a r e 5 t o 8 dB higher than the predicted a t tenua t ion i n t h e region around 200 Hz. However, t he e f f e c t of t he cut-on frequsncy around 570 Hz is c l ea r ly seen i n t he engine da t a t o follow t h e predicted r e s u l t s within 1.5 t o 2 dB.

4.3 PRIMARY NOISE SOURCE LOCATION

Cross-correlation ana lys i s was used i n an attempt t o i den t i fy primary noise source locat ions within the combustor. Pa i r s of sensors combined f o r t h i s ana lys i s included those a t :

Compressor e x i t (Plane 3.0, MADAR Probe) t o combustor i n l e t (Plane 3.5, 102" o r 282' borescope probe)

Compressor e x i t (Plane 3.0) t o combustor exit (Plane 4.0, HPTN probe)

Combustor i n l e t (Plane 3.5) t o combustor e x i t (Plane 4.0)

Compressor i n l e t (Plane 3.0) t o core exhaust (Plane 8.0, Exhaust Probe A o r B)

Combustor e x i t (Plane 4.0) t o core exhaust (plane 8.0)

Comparisons were a l s o made between the 102" and 282' borescope probes located a t Plane 3.5 f o r a f e w of t h e low t h r u s t s e t t i ngs .

The d a t a f o r ihe cross-correlat ions used 25 averages i n t he ana lys i s and were high pass f i l t e r e d above 80 Hz t o remove a test c e l l resonance and el iminate some 60 Hz e l ec t ron i c no ise i n the data .

In gener; qood cross-correlat ions were obtained f o r a l l condi t ions between pa i r s . a t e rna l combustor sensors. A t approach power and below, i n t e r n a l combus r s igna l s cor re la ted w e l l with the core exhaust probe sensor. L i t t l e o r no co r r e l a t i on was observed between i n t e r n a l combustor sensors and the core exhaust probe above 30X Fn, however. This was due t o increased turbulence on the probe and uncorrelated noise i n t he high frequencies. The s igna l between Plane 3.5 t o 4.0 was observed t o be uncorrelated a t conditions of 82 and 94.6% Fn.

A review of the de t a i l ed r e s u l t s f o r the approach condition (30 percent t h rus t ) i s presented f o r i l l u s t r a t i o n . Similar r e s u l t s were obtained a t the lower power s e t t i n g s a t a l l measurement planes. The higher power s e t t i n g r e s u l t s were a l s o s imi l a r f o r the i n t e r n a l combustor sensors , but t he low s igna l co r r e l a t i on between the combustor sensors and the core probe l imited these r e s u l t s .

Coherent spec t ra derived from the cross-correlat ions were used t o com- pliment the cross-correlat ion r e s u l t s and a r e a l s o presented where appro- p r i a t e .

4.3.1 Combustor In t e rna l Sensor Comparison

The in t e rna l sensors were paired f o r cross-correlat ion ana lys i s using the upstream sensor a s the sender and the downstrean sensor a s the rece iver . The t i m e delays of the maximum peaks associated with the cross-correlat ions between Planes 3.0 t o 3.5 a s seen i n Figures 4-18 and 4-19 a r e negative (-1.5 msec) and correspond t o a ve loc i ty of 490 m / s which was computed using the point-to-point d i s tance i n Table E-1, Appendix E. The average acous t ic ve loc i ty between these planes is around 482.5 m / s a t the approach condi t ion which ind i ca t e s t he peak represen ts an acous t ic wave t rave l ing upstream against the flow. Similar r e s u l t s a r e observed i n Figure 4-20 f o r the cross-correlat ion between Planes 3.0 and 4.0. I n t h i s instance, the negat ive time delay corresponds t o a ve loc i ty of 581 m / s which is very c lo se t o the average acous t ic ve loc i ty of 566 m / s between these planes a t t h i s condition.

The cross-correlat ior between the 102' borescope sensor a t Plane 3.5 with the Plane 4.0 probe (Figure 4-21) shows a zero time delay f o r the pzak, ind ica t ing the acous t ic s igna l i s reaching both sensors simultaneously. The cross-correlat ions between the 282' borescope sensar a t Plane 3.5 with Plane 4.0 shows a negative time delay f o r the peak i n Figure 4-22 which i s s imi l a r t o those observed i n Figures 4-18, -19, and -20.

The i n t e r n a l combustor sensor comparison is completed with a cross- co r r e l a t i on a t Plane 3.5 between the 102" and 282' borescope probes. Two peaks a r e observed frcm the r e su l t i ng cross-correlat ion i n Figure 4-23, having equal time delays of opposite s ign ('2.0 msec) which correspond t o v e l o c i t i e s of r634 m / s . The speed of somd a t t h i s plane is 482.5 m / s . Acoustic s igna l s a t t h i s plane appear t o be t rave l ing i n a s p i r a l pa t t e rn across the sensor and moving i n a general ly forward d i r ec t i on aga ins t the flow a s indicated by the time delays obtained from cross-correlat ions between the Plane 3.5 sensors with the sensor a t Plane 4.0 (see Table E-1 i n Appen- d i x E) .

4.3.2 Combustor In t e rna l t o Core Nozzle Sensor Comparisons

A s imi la r anklys i s conducted between the i n t e r n a l combustor sensors paired with the downstream core probe showed pos i t i ve time delays f o r the major peaks of the cross-correlat ions a s noted i n Figures 4-24 and 4-25. The Plane 3.0 t o Plane 8.0 (element A) cross-correlat ion i n Figure 4-23 shows a pos i t i ve time delay of +5 msec f o r the l a r g e s t pos i t i ve co r r e l a t i on peak. This time delay corresponds t o a ve loc i ty magnitude of 553 m / s which i s comprised of both acous t ic and flow v z l o c i t i e s . .i negat ive co r r e l a t i on peak is a l s o observed i n the f i gu re which i s s imi l a r t o the r e s u l t obtained by Karchmer and Reshotkc (Reference 10). The Plane 3.0 t o Plane 8.0 (element B) cross-correlat ion i r . Figure 4-25 shows s imi l a r r e s u l t s .

- A h - .--AI_._ - _. _ . . . .. ..- . . .,.-r--Um+. ,. ,

-.G2 0 02 Time I)e'lay 7 , seconds

I I 1 . b

ECCP Phase 111 -- - * 30 % Fn, RDC 23

. . i

k Figure 4-18. Cross-Correlation of Plane 3.0 to Plane 3.5 ( 102O Probe ) 0

-.04 -,O2 0 02

Time Delay 7 , seconds

Figure 4-19. Cross-Correlation of Plane 3.0 to Plane 3.5 (282' Probe )

ECCP Phase I11

-902 0 . 02 Time Delay 7, seconds

$ Figure 4-20. Cross-Correlation of Plane 3.0 to Plane 4.0

-4

Figure 4-21 0 Z

-.02 0 002 Time Delay 7 , seconds

Cross-Correlation of Plane 3.5 (102O Probe ) to Plane 4.0

-.O2 O .O2 l 04 Time Delay T , seconds

Figure 4-22. Cross-Correlation af Plane 3.5 ( 2820 Probe ) t Plane 4.0

BCCP Phase I11 30 % Fn, RDG 23

- -02 0 -02 .04

Time Delay T, seconds

Figure 4 - 2 3 . Cross-Correlation at Plane 3.5 Between 102O to 2820 Sensors

ECCP Phase I11 3 0 Yb Fn, RDG 23

t I

-06 -.04 -.O2 O .O2

Time Delay T, seconds

Figure 4-24. Cross-Correlation of Plane 3.0 Probe to Plane g.0 Probe Element A

-.O2 0 • Q2 .04


Figure 4-25. Cross-Correlation of Plane 3.0 Probe t o Plane 8.0 Probe Element b

The amplitudes of t!ie c r o s s - c o r r e l a t i o n peaks g ive an i n d i c a t i o n of t h e amount of c o r r e l a t i o n e x i s t i n g between s i g n a l s . I n t h e previous r e s u l t s t h e amplitudes of t h e n o ~ m a l i z e d c o r r e i s t i o n c o e f f i c i e n t , Rxy ranged from +0.25 t o +0.075. CompleLe s i g n a l c o r r e l a t i o n s a r e obta ined when Rxy is equal t o 1.0. Therefore. t h e degree of s i g n a l c o r r e l a t i o n f o r t h e sensor p a i r s pre- v ious ly d iscussed is low.

A c r o s s - c o r r e l a t i o n determined f o r the forward t o a f t sensors on t h e downstream c o r e probe i n F igure 4-26a shows a h igh degree of s i g n a l c o r r e l a - t i o n (4-0.74) wi th ze ro t ime delay. This r e s u l t is presented f o r t h e approach cond i t ion bu t i s r e p r e s e n t a t i v e of the o t h e r lower power s e t t i n g s . The high degree of c o r r e l a t i o n and ze ro time de lay i n d i c a t e s t h a t t h e major i ty of t h e s i g n a l is a c o u s t i c . A c l o s e r look a t t h e c r o s s - c o r r e l a t i o n between t h e c o r e exhaust probe A and B K u l i t e s wi th an rxpanded t ime s c a l e of 2 msec p e r d i v i - s i o n and increased sample r a t e is shown i n Figure 4-26b. Delay t imes corresponding t o t h e c a l c u l a t e d a c o u s t i c and flow v e l o c i t i e s f o r t h i s cond i t ion a r e 0.25 msec and 0.6 msec r e s p e c t i v e l y . The c o r r e l a t i o n c o e f f i c i e n t , Rxy, c f approximately 70% is seen f o r a zero time l a g peak, which corresponds approximately t o t h e a c o u s t i c s i g n a l a t the probe l o c a t i o n . No peak due t o turbulence i s observed, even wi th t h e t ime r e s o l u t i o n increased 25 t imes t h e s t r e n g t h of t h e a r 9 u s t i c s i g n a l appears t o o v e r r i d e any s i g n a l due t o t u r - bulence a t t h e l o c ~ t i o n . The r e s u l t i n g coherent s p e c t r a f o r t h i s cond i t ion a r e , t h e r e f o r e , considered t o be p r imar j ly acous t i c . The coherent spectrum r e s u l t i n g from t h i s c ross -cor re la t ion is shown i n Figure 4-27 t o confirm t h i s s i n c e t h e r e i s almost ccmplete coherence over t h e frequency range of 250 Hz t o 800 Hz, which is t h e core n o i s e region of i n t e r e s t .

4.3.3 Primary Noise Source Summary

The c r o s s c o r r e l a t i o n of t h e i n t e r n a l p r e s s u r e measurements obta ined on t h e ECCP combustor i n t h e CF6-50 engine were reviewed t o determine t h e maximum time delay a t t r i b u t e d t o a c o u s t i c propagation. Vectoring of t h e t ime da lays from each of t h e c r o s s - c o r r e l a t i o n s gave an i n d i c a t i o n ~f t h e d i r e c t i o n of s c o u s t i c s i g n a l t r a v e l between sensors . Amplitudes of t h e peaks were compared t o those of t h e t i m e de lays assoc ia ted wi th t h e f low v e l o c i t y i n o r d e r t o eva lua te t h e s t r e n g t h of t h e a c o u s t i c s i g n a l . A t a b u l a t i o n of t h e ampl i tudes and time de lays f o r these c r o s s - c o r r e l a t i o n s is faund i n Apperidix D. This procedure, performed a t each t e s t cond i t ion , i n d i c a t e d a c o u s t i c s i g n a l s were t r a v e l i n g upstream between p a i r s of sensors i n t h e combustor r eg ion (Planes 3.0, 3.5 and 4.0). A downstream t r a v e l i n g s i g n a l was apparent f o r a l l vectoring done between t h e combustor sensors and t h e c o r e exhaust probe.

The c ross -cor re la t ion between the 102' and 282" s e n s o r s a t Plane 3.5 shows p o s i t i v e and nega t ive t ime delays corresponding approximately t o a c o u s t i c v e l o c i t y and of n e a r l y equal amplitude. This sugges t s t h a t c i r c u m f e r e n t i a l t r a v e l i n g waves a r e a l s o p resen t i n the c?mbusior.

The forward t r a v e l i n g s i g n a l s occur between P lanes 3.0 t o 3.5, 3.0 t o 4.0, and 3.5 t o 4.0. Th i s r e s u l t sugges t s t h s t t h e primary n o i s e source is loca ted nea r t h e combustor d i scharge s i n c e t h c h ighes t energy l e v e l is i n t h i s

I I I J 4

ECCP Phase 111 30 76 Fn, RDG 23

1 L

- - It

-002 0 . 02 Time Delay T , seconds

Figure 4-26.a. Cross-Correlation of Core Probe Element B With Element A

1 1 I

o ECCP Phase I11 o 30% Fn, RDG 23

,--- -------- - 1 - --

I \ I

v I *

Time Delay T, Seconi

Figure 4-Z6b. Cross-Correlation of Core Probe Element B with Element A, On Expanded Scale

ECCP Phase I11 30 ?1. Fn, RDG 23

J Raw Spectrrm - of Element B

I - r- -;=--bcl L

2 ;a . - .I

of Element B - - - I

I

r ~ r v r r r r -

l /3 Octave Band Center Frequency, Hz

Figure 4-27 . Coherent Spectrum a t Core Nozzle Discharge, Plane 8.0

area . The c ross -cor re la t ions f o r cond i t ions below 45% Fn between t h e combustor sensors and t h e exhaust probe i n d i c a t e the sound i s t r a v e l i n g a f t .

A review of t h e c ross -cor re la t ion a n a l y s i s r e s u l t s i n d i c a t e s t h e combus- t o r environment is a highly t c r b u l e n t flow f i e l d wi th numerous n o i s e sources . The e l l t i r e combustor region might be c o n s i d e r ~ d a s a source s i n c e t h e wave l eng ths of predominate concern f o r core no i se a r e from 0.914 t o 1.524 meters long a t t h e engiae opera t ing cond i t ions and t h e d i s t a n c e s between probes average 0.304 t o 1.219 meters.

4.4 OVERALL POWER LEVEL COMPARISONS

Comparisons were made a t t h e core nozzle discharge plane wi th t h e o v e r a l l sound power l e v e l computed from probe raw s p e c t r a and wi th coherent s p e c t r a considered t o be mostly sound.

Overa l l power l ev01 comparisons were made between t h e measured r e s u l t s ( F P b e a s ) and the G e n ~ t a l E l e c t r i c component (PWLGE) and engine (PWLGEFAA) c o r r e l a t i n g parameters used f o r OAPWL pred ic t ions . The measured r e s u l t s assume t h a t the measured p r e s s u l e is e n t i r e l y made up of a c o u s t i c p lane waves both i n the combustor and core nozzle.

4 .4 .1 Acoustic Power Levsl i n Exhaust Duct

The power l e v e l i n t h e engine c o r e exhaust duc t was c a l c u l a t e d from pressure measurements made wi th t h e Plane 8.0 sound s e p a r a t i o n probe. The probe was t raversed t o f i v e r a d i a l immersions loca ted on c e n t e r s of equal annulus a r e a s . The FPL 113 octave band s p e c t r a from 50 co 2000 Hz obtained a t each immersion were compared f a r s i m i l a r i t y . Figure 4-28 shows t.hiw comparison f o r both K u l i t e s on t h e probe &t t h e 30% Fn approach power s e t t i n g . It il l -

d i c a t e s c l o s e agreement i n s p e c t r a l shape ?nd content f o r a l l immersions except t h e i n n e r immersion ( 5 j which is near t h e tu rbu len t region of t h e plug nozzle centerbody. This r e s u l t i n zenera l , t y p i f i e s colnparisons of t h e s e measurements a t o t h e r condi t ions . Because of t h e c l o s e agreement a t a l l immersions t h e p i t c h l i n e immersion (3) was t h e primary i m e r s i o n used i n a l l measurement compsrisons wi th upstream sensors . Cross-corre la t ion a n a l y s i s was performed between t h e "A" and "B" K u l i t e s on t h e c o r e probe a t each immersion f o r a l l test condi t ions . A high degree of c o r r e l a t i o n between K u l i t e s i g n a l s was apparent a t a l l immersions f o r power s e t t i n g s of 30% Fn (100% p i l o t f u e l ) and below.

Coherence a n a l y s i s performed on these sensors a t t h e s e cond i t ions showed good coherence between t h e forward and a f t s e n s o r s on t h e probe i n t h e f r e - quency range between 315 t o 630 Hz which is t h e primary region f o r c o r e no i se . Figure 4-29a and -29b i l l ~ u t r a t e t h e c ross -cor re la t ion and coherence a n a l y s i s r e s u l t s f o r t h e 30% Fn p o i n t (wi th 100% p i l o t f u e l ) a t immersion four . The c o r r e l a t i o n c o e f f i c i e n t Rxy, i s 0.75, a s seen i n Figure 4-29a. Note t h a t t h e pro5e 1 / 3 octave band coherent spectrum (frequency range of 100

ECCP PHASE I11

- 0 6 3 -.02 - .01 0 -01 . 0 2 03 Time Delay 7 , seconds

a. Cross -corre la t ion o f A (Delayed) with B 170 - 1 i : LOO yb P i l o t Fuel

150 Haw Spectrum

s .. a

130 rn 0

(?

2 110

K u l i t e B

90 125 250 500 lo00 2000

Frequency. Hz

b. Coherent Spectrum of Element B

Figure 4-29. c o r e Probe Immersion 4 a t 3 0 Percent Thrust

through 2000 Hz) i n Figure 4-29b, wi th i ts peak a t 400 Hz, f i t s almost e x a c t l y , wi th the GE co re n o i s e p red ic ted spectrum shape (Reference l l ) , t h e l e v e l of which was s h i f t e d :o agree wi th t h e coherent s p e c t r a l l e v e l a t 400 Hz. Simi- l a r r e s u l t s f o r t h e 19.72 Fn p o i n t w i t h 100% p i l o t f u e l a r e presented i n Figures 4-30a and -30b. A l e v e l of 0.55 i s observed f o r t h e c o r r e l a t i o n c o e f f i c i e n t . The p red ic ted c o r e n o i s e spectrum i s aga in , i n veyy good agreement wi th t h e coherent spectrum i n t h e reg ion of h igh coherence, i n d i c a t i n g t h e sound observed is predominantly c o r e no i se .

The approach power s e t t i n g a t 29.7% Fn wi th a nea r 50/50 f u e l flow s p l i t between p i l o t and main burner s t a g e s shows a l e s s e r degree of c o r r e l a t i o n (below 0.2) than f o r t h e 100% p i l o t f u e l cocd i t ion . The h i g h e r power s e t t i n g s which opera te on a s p l i t f u e l f low schedule a l s o showed a much lower c o r r e l a - t i o n between s i g n a l s a t a l l immersions. The coherence a n a l y s i s of t h e s e cond i t ions showed a reduc t ion of t h e coherent p o r t i o n of t h e spectrum t o between 315 t o 500 Hz. The amount of coherence between t h e c o r e probe s i g n a l s f o r these mixed f u e l f low p o i n t s was l e s s than f o r t h e 100% p i l o t f u e l p o i n t s . F igures 4-31a and b i l l u s t r a t e t h e r e s u l t s obta ined from t h e c r o s s - c o r r e l a t i o n and coherence a n a l y s i s of t h e approach po in t a t 29.7% Fn (wi th 46.9% p i l o t f u e l ) . The coherent spectrum a t t h i s cond i t ion s t i l l r e t a i n s t h e c o r e n o i s e shape but t h e r e i s l e s s agreement wi th t h e core n o i s e p r e d i c t e d spectrum. This i s c h a r a c t e r i s t i c of the r e s u l t s obta ined a t t h e high t h r u s t , s p l i t f u e l flow cond i t ions , a s i l l u s t r a t e d by Figures 4-32a and b which show r e s u l t s f o r 64.6% Fn wi th 17.8% p i l o t f u e l . The amount of c o r r e l a t i o n is aga in low ( t h e l e v e l of the normalized c o r r e l a t i o n c o e f f i c i e n t is 0.2:. and t h e amount of coherence betwten t h e r z w and coherent s p e c t r a is l e s s than a t t h e lower power s e t t i n g s v i t h 100% p i l o t f u e l .

Average o v e r a l l p r e s s u r e l e v e l s were determined f o r a l l cond i t ions from s e l e c t e d immersions. The immersions s e l e c t e d were based on those w i t h t h e h igher degree of c o r r e l a t i o n and t h e l e a s t no i se f l o o r of t h e coherent s p e c t r a .

The o v e r a l l p r e s s u r e level - ; determined from t h e c o r e probe immersion averaged raw s i g n a l l e v e l s and t h e coherent s p e c t r a (which are considered t o be mostly sound) were compared on a c o r r e c t e d c o r e speed b a s i s . The r e s u l t s of t h i s comparison i n Figure 4-33 shows t h e e f f e c t of t h e o p e r a t i o n w i t h p i l o t f u e l only , and wi th a f u e l s p l i t t o t h e p i l o t and main burner systems on t h e o v e r a l l sound p ressure l e v e l i n t h e c o r e exhaust . A r educ t ion i n OASPL is r e a l i z e d when opera t ing i n t h e s p l i t f u e l f low mode. Th i s r educ t ion a t approach i s seen t o be a s much a s 7.5 dB wi th t h e raw s p e c t r a t o 9 dB wi th t h e coherent s p e c t r a .

The a c o u s t i c power l e v e l i n tb.- c o r e nozzle is c l e a r l y dominated by core n o i s e a t approach cond i t ions and below. These c o n d i t i o n s which o p e r a t e wi th t h e p i l o t burner only a l s o show a high OASPL. A t h igher power s e t t i n g s which opera te wi th f u e l f low s p l i t s t o t h e p i l o t and main s t a g e burners t h e OASPL is l e s s than the low power s e t t i n g s . The e f f e c t is most dramat ic a t t h e approach c o n d i t i o n wi th and wi thout t h e f u e l f low s p l i t .

ECCP PHASE I11 % J

Core Probe 100 y6 P i l o t Fuel -- -

r

-.I

I I

- .6 -.03 -.02 -.01 0 -01 -02 0 3

Time Celay T, seconds

170 - a. Cross-Correlation of A (Delayed) with P

1 100 96 P i l o t Fuel 1 f Raw Spectrum

Kulite B

- - - - - 1

I Core Noiss Spectrum - Spec trun Kulite B

I l l l # r l I l l l l l ,

125 250 500 lo00 2000 Frequency, Hz

b. Coherent Spectrun of Element B

Figure 4-30* Core Pr,,be Innnersior~ 4 a t 19.7 Percent Thrust

ECCP PHASE 111

c 04

Sp (d

4 02 8 - h 4J & c 5 -2 0

0 a .rl 8 : '4 Q - .2 4 0 e * 0 z - .4

-.03 -.02 -.oi o .01 .oz .03 Time Delay 7, seconds

a. cross-Correlation of A (Delayed) with B

Fuei S p l i t 66.9 % P i l o t

- - ..

- -

Spec trun - Kulite B Core Noise

Spectrun w-7 r r i r

Raw Spectruu

125 250 500 loo0 2000 Frequency, Hz

b. Coherent Spectrun of Element B

Figure 4. -1. Core Probe Immersion 4 a t 29.7 Percent Thrust

ECCP PHASE 111


a. Cross-Correlation of A (Delayed)

03

with B

90 . . . . - - - - - -

125 250 500 lo00 2000 Frequency, Hz

b. Coherent Spectrm of Element R

- - Fuel S p l i t -

17.8 % P i l o t - .. Raw Spectrm -

- - - - - a

Spec trun - Core Noise Kulite B

- Spec trm

W V I V D I D D ~ ~ K ~ X .

Figure $-32. Core Probe Immersion 4 at 64.6 Percent Thrust

ECCP Phase 111 Engine Test Double Annular Combustor

Percent Net Thrust 30

3.8 19.7 29.7 45.4 64.6 82 94.6 I 1 I I I

e- 100 % P i l o t Fuel Only

00- Fuel S p l i t Between P i l o t and Main Stage

Corrected Core Speed, Nc rpm

Figure 4 - 3 3 . Comparison o f Raw and Coherent OASPL with Core Speed

4.4.2 Comparison of Measured Power Level with Component and Engine Pred ic t ions

The component and engine ove ra l i power l e v e l p red ic t ions were based on cor rec t ions derived from t e s t s conducted under the GE Core Engine Noise Inves t iga t ion Program, Reference 11.

Component (PWLGE) Comparison

The component predict ion is based on the co r r e l a t i on derived from component t e s t s conducted i n Reference 11. The ove ra l l powe; l e v e l PWLGE is calculated from:

PWLGE = 83.6 + 10 log [ ( P ~ ~ / T ~ ~ ) ~ * ~ Vr 3'5 (TT4-TT3) 1 , r e 10-l~ Warts (7)

where = Combustor i n l e t t o t a l pressure, k ~ / m 2

'T3

TT3 = Combustor i n l e t t o t a l temperature, K

T T ~ = Combustor e x i t t o t a l temperature, K

V r = Reference ve loc i ty defined a s

W36 = Tota l combustor a i r flow, kg/sec

2 = Combustor reference a rea , m2 ( fo r CY6-50 Are, = 0.37287 m )

The aero parameters used i n ca lcu la t ing the PWLGE values a r e Zound i n Table 5 (Section 3.7).

The measured r e s u l t s were based on an ove ra l l power l e v e l determined from the i n t e r n a l r ' luctuating pressure measurements reduced t o 113 octaveband spec t ra and corrected fo r frequency response loss . The o v e r a l l l e v e l s determined from these pressure measurements (considered to be acous t ic plane wave s igna ls ) were converted t o power l e v e l accounting f o r impedance changes due t o d i f fe rences i n t e s t conditions, Mach number e f f e c t s , and a r ea d i f fe rences a t each measurement plane a s described i n Section 3.7. The r e s u l t i n g power l e v e l was designated F P h e a S with u n i t s i n dB r e 10-13 wat t s . The parameters used i n the conversion to FPWLmeas a r e found i n Table 6.

Figure 4-34 shows the comparison of FPWLmeas t o PWLC,E f o r the i n t e r n a l combustor sensors a t Planes 3.0, 3.5 and 4.0. The Plane 3.0 FYWLmeas is 5 t o 15 dB lower than the FPWLmeas = PWLCE l i n e a t condi t ions above i d l e giving a f l a t t e r trend t o t he data. The FPL measured a t Plane 3.0 is i n t he compressor discharge region p r i o r t o the s t a r t of combustion.

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The FPWheas a t Plane 3.5 agrec e l l wi th t h e p red ic ted PWL a t a l l t e s t colldit ions except a t t h e extre. .s. Plane 3.5 is i n t h e pre-%x region of t h e combustor a f t of t h e f u e l noz7les. I g n i t i o n t akes p lace a t t h i s p lane t o s t a r t the combustion process.

A t Plane 4.0, t h e F P ' h e a s is higher than p red ic ted a t t h e lower h a l f of t h a t e s t cond i t ions by 15 t o 5 dB and g i v e s a s i m i l a r d a t a t r end as seen f o r Plane 3.0. The FPL a t t h i s p lane is determined from measurements i n t h e borescope p o r t located between two vanes i n t h e high p ressure t u r b i n e nozz le diaphragm. As previously discussed, t h i s region, which opera tes a t a near ci.,;stant Mach number of 0.55, con ta ins a l a r g e amount of high frequency n o i s e a t a l l t e s t condi t ions .

The b e s t co~uparison wi th YWLGE of t h e engine F P b e a s from FPL measurements appears t o be a t Plane 3.5 f o r v i r t u a l l y a l l test cond i t ions .

engine^^) Comparison

The PWkEFAA engine p r e d i c t i o n i s based on a c o r r e l a t i o n der ived from engine f a r f i e l d d a t a and c o r e probe r e s u l t s obta ined dur ing t e s t s conducted under the Core Engine Program, keference 11. The o v e r a l l power l e v e l from t h e engine c o r r e l a t i o n is c a l c u l a t e d from

2 2 P W L ~ ~ ~ ~ ~ = 169.3 + 1 0 l o g W36(TT4-TT3) (P3/Po) - 40 l o g ( T T ~ - T T ~ ) ~ ~ ~ ~ ~ ~

r e 10-l~ w a t t s (8)

where the nomenclature i s descr ibed above and (Ty4 - T T ~ ) ~ ~ ~ ~ ~ ~ is t h e t o t a l temperature drop a c r o s s t h e high and low pressure t u r b i n e s a t t h e c y c l e design po ia t .

The comparison wi th t h e measured power l e v e l i s based on t h e p ressure measurements taken w i t h t h e sound s e p a r a t i o n probe i n t h e core nozzle. The power l e v e l determined from +t.e a s measured f l u c t u a t i n g p r e s s u r e measurements i n the core exhaust was h igher than t h e predic ted p a r e r l e v e l from t h e engine c o r r e l a t i o n . To more a c c u r a t e l y compare t h e ECCP Phase 111 d a t a wi th t h e engine c o r r e l a t i n g parameter, t h e coherent p a r t of t h e probe K u l i t e B s p e c t r a wi th t h e probe K u l i t e A s i g n a l was used t o compute t h e P h e a S . An average OASPL obta ined from t h e i n d i v i d u a l immersions was used i n conjunct ion wi th t h e PWL calcu1a;ion procedxre noted i n Sect ion 3.7.

F igure 4-35 shcws t h e comparison of PWLmeaS computed from coherent s p e c t r a a s d iscussed wi th t h e engine p r e d i c t i o n parameter PWLGEFAA. The s o l i d symbols a r e f o r t h e lower pcwer s e t t i n g s using 100% p i l o t f u e l , whi le t h e open symbols a r e f o r s p l i t f u e l f low condi t ions . The i d l e po in t was highly contaminated wi th a l a r g e 400 Hz s i n u s o i d a l tone, a s previously discussed. A tone correct ior , of 7.9 dB brought t h i s po in t i n agreement wi th t h e o t h e r 100% p i l o t po in t s (dark symbols) which then p a r a l l e l e d t h e equal PWL l i n e . The

ECCP Phase I11

Double Annular Combustor

Pameas Based on Coherent

Spectra frw, Core Probe (B)

Fuel t o Outer Fuel Nozzles Only

0 Fuel S p l i t Between Inner and Outer Fuel Nozzles

A

Tone - 4 A -1

1 Repeat Points - -

I , . \ a t 45.4% Fn

tone c o r r e c t i o n was app l ied t o t h e OASPL from t h e 113 OBSPL coherent s p e c t r a a t each immersion. An average "tone corr-cted' ' OASPL was then obtained and con- ve r ted t o PWLmeas. This f i g u r e c l e a r l y shows t h e e f f e c t of burning i n t h e o u t e r p i l o t s t a g e only a s r e s u l t i n g i- higher o v e r a l l power l e v e l s by about 15 dB. The s p l i t f low condi t ions , however, agree wi th t h e c o r r e l a t i o n as repea t readings a t 45.4% Fn f a l l above and below t h e equal PWL l i n e . The approach point a t 30% Fn (100% p i l o t f u e l ) is 9 dB above t h e same condi t ion a t 29.7% Fn with a f u e l s p l i t of about 50%.

Operating t h e combustor wi th the o u t e r r i n g of 30 f u e l nozzles only , a t t h e power s e t t i n g s of approach and below i s d e s i r a b l e from an emissions s t andpoin t a s d iscussed i n Reference 1. With f u e l t o t h e p i l o t s t a g e only t h e 24% of t h e t o t a l combustor a i r t h a t passes through t h e o u t e r burner .nixes wi th the t o t a l amount of f u e l added. The main burner a i r f low (49% of t h e t o t a l combustor a i r flow) may be considered t o a c t a s cool ing o r d i l u t i o n air ~ h i c h mixes wi th t h e o t h e r 29% of W36 used f o r l i n e r cool ing, e t c . Th i s has t h e e f f e c t of inc reas ing t h e l o c a l f u e l - a i r r a t i o and r e s u l t s i n ho t f low i n t h e o u t e r annulus wi th a higher v e l o c i t y , whi le t h e f l o v i n t h e i n n e r annulus remains r e l a t i v e l y coo l and has a lower v e l o c i t y . Mixing of t h e s e two d i s - s i m i l a r r eg ions may r e s u l t i n no i se genera t ion due t o s h e a r l a y e r i n t e r a c t i o n between t h e hot and coo l streams. There is , of course, no way of determining t h i s from t h e p resen t scope of work.

With both p i l o t and main s t a g e burning, t h e temperatures i n both streams a r e i n much c l o s e r agreement and consequently, t h e stream v e l o c i t i e s a r e more near ly equal , t h ~ s reducing any shear l a y e r i n t e r a c t i o n noise . P ressure measurements taken a t Plane 4.0 a t 30% F, wi th t h e 100% p i l o t f u e l and a t 29.7% Fn wi th both p i l o t and n a i n s t a g e s fue led show a d i f f e r e n c e of approximately 6 dB lower l e v e l wi th t h e 60 f u e l nozzles i n both s t a g e s burning.

Agreement wi th t h e engine c o r r e l a t i o n is obtained f o ~ t h e 100% p i l o t f u e l only p o i n t s i n Figure 4-36 when an e f f e c t i v e temperature, T4p, i s computed from t h e f u e l - a i r r a t i o determined f o r t h e percentage of t h e t o t a l combustor a i r f l o w t h a t passes only though t h e o u t e r p i l o c burner annulus. The fuel-ai-r r a t i o f o r these 100% p i l o t only p o i n t s is considerably h igher s i n c e on ly t h e p i l o t a i r e n t e r s i n t o t h e c a l c u l a t i o n . The main s t a g e air i n t h i s case , i s considered t o a c t only a s cool ing and d i l u t i o n a i r and dnes n o t e n t e r t h e combustion process.

The higher T4p obtained from t h e increased f u e l - a i r r a t i o r e s u l t s i n a g r e a t e r PWLGEFAA which s h i f t s t h e 100% p i l o t p o i n t s a t 30, 19.72 Fn and t h e tone cor rec ted i d l e cond i t ion wi th in t h e d a t a c o r r e l a t i o n i eg ion .

Resu l t s oE t h e ECCP Phase 111 t e s t s have shown a s i g n i f i c a n t i n c r e a s e i n o v e r a l l sound power l e v e l (9 dB) a t approach wi th 1.00% p i l o t f u e l , conpared t o an approximate 50/50 f u e l s p l i t between s t a g e s . Considerat ion must, t h e r e f o r e , a l s o be given t o t h e i n f l u e n c e of n o i s e on t h e opera t ion of t h e Double Anr4.dar combustor system dur ing approach power, i n a d d i t i o n t o t h e emission t radeof f s .

ECCP Phase 111

Double Annular Combustor

' PULmeaa Baaed on Coherent

Spectra from Core Probe ( 8 )

*Fuel t o Outer F u e l S o z z l e s only S h i f t e d f o r Increased Fuel-Air R a t i o

0 Fuel S p l i t Between Inner and Outer Fuel Nozz le

130 140 150 160

P W I ~ ~ ~ ~ ~ A d~ re w a t t s

Figure 4 - 3 6 . Comparison cif PkL t o P W L ~ ~ ~ ~ ~

Us ing Adjusted meas

Fuel -Air R a t i o f o r 100 % P i l o t P o i n t s

5.0 CONCLUSIONS

This was t h e f i r s t time i n t e r n a l p ressure me isurements w e r L acqufred on t h e CF6-50 engine wi th a Double Annular combustor. The s i g n i f i c a n c e of t h i s engine test was t o e s t a b l i s h a base f o r comparison wi th component d a t a and provide an understanding of t h e opera t ion of a Double Annular combustor system i n an engine environment,

5.1 DOUBLE ANNULAR COMBUSTOR SPECTRA COMPARISONS - - The r e s u l t i n g s p e c t r a from ~ 4 e i n t e r n a l comb.ustor measurements i n d i c a t e

t h a t cond i t ions a t t h e lgwer power s e t t i n g s (30%'Fn and below) a r e more c h a r a c t e r i s t i c of c o r e n o i s e spec t ra . The Pldne: 3.5 1 /3 octave band s p e c t r a a r e peaked i n t h i s frequency range between 250 and 1000 Hz f o r those lower power p o i n t s , which is t y p i c a l of c o r e noise. The measurements a t t h i s p lane a r e more r e p r e s e n t a t i v e of combustion n o i s e than e i t h e r t h e compressor d i s - charge, ? lane 3.0 o r combustor d i scharge , Plane 4.0. Values of F P h e a s computed from t h e Plane 3.5 measurements match t h e p red ic ted va lues of P ~ E from t h e component cc -:elation.

The c o r e probe measurements a t these cond i t ions match t h e empi r ica l combustor shape, peaking a t 400 t o 600 Hz. Levels of P h e a S computed from t h e core probe average OASPL from coherent s-rectra (which were mostly sound) obtained from as many as f i v e r a d i a l innnersidas showed good agreement wi th t h e engine PWLGEFAA c o r r e l a t i o n f o r t h e approach power s e t t i n g and hI qher po in t s operated wi th f u e l s p l i t s t c t h e inner and o u t e r r i n g s of f u e l nozzles . The power l c v e l s frorr t h e approach cond i t ion and lower p o i n t s operated wi th 100% f u e l t o t h e oglter p i l o t rIt: of f - e l nozzles were h igher than predic ted. Adjusting t h e T T ~ t o account f o ~ t h e increahed frq.el-air r a t i o i n t h e o u t e r annulus wi th p i l o t f u e l only increased t h e PWLGEFA~ and s h i f t e d t h e lower power p o i n t s wi th in t h e reg ion of d a t a c o r r e l a t i o n .

The r e s u l t s from t h i s Double Annular combustor t e s t show a s i g n i f i c a n t i n c r e a , ~ i n o v e r a l l power l e v e l (9dB) a t approach wi th 100% p i l o t f u e l , compared t o an approximate 50150 f u e l s p l i t br -ween p i l o t and main s tage . Considerations must t h e r e f o r e be given t o t h e no i se impact as *re11 a s t h e emissions t r a d e s i n s e l e c t i n g approach power opera t ing f u e l schedules.

The t ransmiss ion losn a c r o s s t h e CF6-50 t u r b i n e and nozzle computed from t h e d i f f e r e n c e '~ctwef ,I s p e c t r a which were predominately a c o u s t i c matched component t e s t r e s u l t s snd t h e o r e t i c a l p red ic t ions .

5.2 NOISE SOURCE LOCATIONS

The r r . su l t s rf t h e c ross -cor re la t ion a n a l y s i s showed nega t ive time delays f o r t h e i n t e r n a l combustor s e n s s r s p r i r e d upstream t o downstream,

i n d i c a t i n g for ward sound wave propagatior., oppos i t e t o t!~e d i r e c t i o : ~ ~i flow. The i a t c r n a l combustor t o core nozzle Kul i t e p a i r s gave p o s i t i v e t i u e delays , i n d i c a t i n g a c o u s t i c s i g n a l s t r a v e l i n g aft. Circumferent ia l waves were a l s o i d e n t i f i e d between t h e Planes 3.5 senso l s a t 102" and 282".

The predominate r%e de lays between a l l tht: K u l i t e se1,sar p a i r s a r e l i s t e d i n Table E-1 of Appendix E. A s seen i n the t a b l e , p o s i t i v e t i m e de lays a r e recorded f o r combustor sensors pa i red wi th t h e eensars on t h e probe i n t h e core exhaust , Plane 8.0. Negative time de lays a r e observed between any combination of i n t e r n a l combustor s e n s o r s pa i red upstream t o downstream. The time delay f o r c ross -cor re la t ions between Plane 3.5 (102") wi th Plane 4.0 is zero (a t 30% Fn) o r very near zero (-0.3 msec ac 19.72 Pi,) f o r the l imi ted dat.3 a v a l l a b l c , i n d i c a t i n g t h e source may be between Plane 3.5 (102") ~ n d Plane 4.0. Howeccr, t h e p o s i t i v e time d e l a y s a s s o c i a t e d wi th Plane 4 .0 t o Plane E.0 a d t h e o ther combustor sensor ,) lanes wi th t h e core exhaust suggest the pvcsexlce of another source downstream of t h e t u r b i n e thi.C J a s no t of a propagating a c o u s t i c na tu re u n t i l a f t e r pass ing through t h e tur- b ine .

5.3 ENGINE TU DUCT= COMPARISONS

The comparison of t h e engine-to-duct r i g measurements made oti the Dotible .4nnular combustor dur ing t h i s program showed l a r g e d i f f z r e n c e s ~f 1 0 t o 25 dB on a p ressure l e v e l s p e c t r a b a s i s between both i n l e t and discharge meauure- ment planes. The power l eve l romparisons showed considerably c l o s e r agreement wi th d i f f e r e n c e s i n l e v e l s of aboqlt 3 t o 8 dB over t h e frequency range. The s p e c t r a ? shapes of t h e engin, and duc t r i g measurements a r e very c ) ? . i l a r .

Some cf the d i f f e r e n c e s j.n l e v e l a r e concluded t o be t h e r e s u l t of t h e d i f f e r e n c e i n s t a t i c p ressure t e tweer~ engine and duc t r i g a t t h e high Fcwer po in t s , measurement p lane l o c L ~ i a ~ d i f f e r e n c e s , and t h e presence of t h e choked nozzle diaphragm i n t h e engine which i s ab!.ent f r a ? t h e duct r i g . . r : ~ t h e r f a c t o r i s thc d i f f e r e n c e due t o combustor , ,peration wi th p i l o t fc;: ur.iy f o r t h e low power s e t t i n g s on ~ 1 . i ~ engice as oppossed t o opera t ion wi th fue l ~ p l i t s t o t h e l r i l o t and maia .>tage burners i n t h e test r i g f o r t h e Phase I and I1 noise measarement?.

. - '-z

APPENDIX A - 2 HZ NARROWBAND SPECTRA RESULTS

The 2 Hz narrowband spec t ra r e s u l t s presented i n t h i s Appendix a r e based on f luc tua t icg pressure l e v e l s (FPL's) from the raw i n t e r n a l measurements on the CF6-50 engine with the ECCP Phase I11 Double Annular combustor. The FPL spectra p l o t s cover the frequencies up t o 2000 Hz and a r e presented fo r the ava i lab le sensors a t each of the e igh t test condit ions including repeat points.

ECCP Phase 111 Combustor Test, 1977

2 Hz Narrowband Spectra

Reading 6

MADAR Probe OASPL = 156.5dB. OASPL = 155.6dB

. Plane 4.0 Plane 5.5 HPTN Probe

150 ' 282' Borescope : OASPL = 160.3dB . 140. OASPL = 155.6dB .

110 1- ,-

al h

140 i bO e Plane 8.0 + u 130 a Exhaust Probe A . 1 OASPL = 145.7dB . u 120' 0 1 4

110

100

Exhaust Probe B . OASPL = 142.9dB ,

0 1000 2000

Frequency, Hz

Figure A-1. As Measured Narrowband Spectra for CF6-50 Engine Condition at 3.8 Percent Thrust, Reading 6 ( Idle )

a ECCP Phase I11 Combustor Test, 1977

a 2 Hz Narrowband Spectra

a Reading 16

150 1 a Plane 3.0 a MADAR Probe ,

a OASPL = 161.9dB

120

1

a Plane 3.5 - 150 . a 282' Borescope a OASPL = 166.8dB

130 ' a Plane 4.0 o HPTN Probe

. l l o ~ . . . - ~ ' . . ' ~ . . " ' " " 110. . . . . I ....I . . . - r e - . .

a Plane 8.0 a Plane 8.0

a Exhaust Probe 130. a Exhaust Probe B -

a OASPL = 147.3dg 120 '

a OASPL = 146.6dB -

Frequency, Hz

Figure A-2. AS Measured Narrowband Spectra for CF6-50 Engine Condition at 19.7 Percent Thrust, Reading 16

a ECCP Phase I11 Combustor Test, 1977


a Reading 23

OASPL = 164.6dB

120

160,

140 a Plane 4.0

150.

HPTN Probe OASPL = 176.4dB

110

a Plane 3.0 a MADAR Probe

Plane 8.0 Exhaust Probe A

120 OASPL = 149.7dB

110

100

130

120 OASPL = 171.ldB

a OASPL = 168.7dB . . . . , . ' . . , . . . . 1 . . . .

a Plane 8.0 ' a Exhaust probe B -

110 . 100

9 0 . . . . . , . . . . , . . . . , . . . . - 0 1000 2000

Frequency, Hz

Figure A-3. As Measured Narrowband Spectra for CF6-50 Engine Condition at 30 Percent Thrust Reading 23 (Approach Power with 100 Percent Pilot Fuel)

a ECCP Phase I11 Combustor Tes t , 1977


a Reading 39

130 Plane 3.0

120 MADAR Probe a OASPL = 166.ldB

. . ' I - . - ' I v ' g ' . I ' ' ' - . -. Z

- . I . 160

+ *

150 .

130 . rl Plane 4.0 Plane 3.5 a 5 1 2 0 . 120 <

a HPTN Probe a 282' Borescope c] a OASPL = 170.9dB OASFL a 170.3dB a l l 0 . , v v w r , . . , . , . . . , . . - . 1 1 0 . k 1 . ' . ' 1 " . ' 1 " "

a Plane 8.0 a Plane 8.') Exhaust Probe A a Exhaust Probe B -

a OASPL = 149.0dB OASPL = !48.ldB

100'

9 0 ? ~ ; 1 1 ~ - u ~ 1 ~ v i ' " 0 ' 90,. . . . . , . . . . , . . . . , . . . . 0 1000 2000 0 1000 2000

Frequency, Hz

Figure A-4.As Measured Narrowband Spectra f o r CF6-50 Engine Condit ion a t 29.7 Percent T h r u s t , Reading 39 ( Approach Power w i t h 50/50 Fuel S p l i t )



Reading 34

MADAR Probe

. . . ' , . . . . 1 . . . ' 1 " . . OASPL = 168.0dB

P) k 3 m

160 . Plane 8.0 Exhaust Probe B 150 -

e OASPL = 150.8dB

100. OASPL- 172.0dB

90. . . . . , . . . . , . . . . , . . . '. 0 1000

Frequency, Hz

Figure A - 5 . A s Measured Narrowband Spectra for CF6-50 Engine Condition at 4 5 . 4 Percent Thrust, Reading 34

ECCP Phase I11 Combustor Test, 1977


e Readfiig 28

Plane 3.0 e W A R Probe o OASPL = 167.8dB

? ~ ~ . ~ l l l ~ l [ l ~ ~ ~ l ~ t ~ . ~

1301' Plane 4.0 e HPTN Probe

120' OASPL = 175.5dB

130. Plane 3.5 282' Borescope

120. OASPL = 172.8dB

Plane 8.0 Plane 8.0 o Exhaust Probe A Exhaust Probe B

OASPL = 149.6dB

ioa

Frequency, HZ

Figure A-6.As Measured Narrowband Spectra for CF6-50 Engine Condition at 45.4 Percent Thrust, Reading 28

ECCP Phase I11 Combustor Test , 1977


e Reading 40

, 160 ..

r Plane 3.0 cs 120 . r, Pla:le 3.5 TI W A R Probe 1 2 0 . 102' Borescope

F7 110 ,.. , . . , , . . . , , , , , , OASPL = 169.0dB OASPL = 174.4dB

2 110 -, . . . , I . . . . , . . . . , . , , . ,

7

r, Plane 8.0 Plane 3.5 Exhaust Probe B 282' Borescope

a OASPL = 150.5dB

c) 5 100 . 120. ;1 I

9 0 t 7 . . , , , , , , . , , lo.. . , , . I . . , , , . , . . 0 560 10A0 1560 2000 0 5 00 1000 l sb0 2000

Frequency, Hz

Figure A-7.As Measured Narrowband Spectra f o r CF6-50 Engine Conditi0.n a t 64.6 Percent Thrus t , Reading 40



Reading 13

Plane 3.0 Plage 3.5 MADAR Probe 282 Borescope . OASPL I 168.9dB OASPL = 173.5dB .

In 1104.. , , . , . , , . , . . . . , . . , . - I 1 " " 1 " " 1 9 ~ ~ . 1

Plane 8.0 Exhaust Probe A' OASPL = 149.8dB

. . N

\

140. 9)

5 Plane 4.0 130. HPTN Probe . 110

9) LI OASPL = 178.7dB 3 120.. rn

. 100 rn PI a '

110" ' . ' . , ' . . . , . . . . , , . . . , P.l 9 0

ca d .I4 U

9 140 U

- 140 - C)

Plane 8.0

3 130 e Cell Probe 130 Exhaust Probe G F OASPL - 126.0dB OASPL - 149.7dB

120

110 110.

100

--r -1- -1 90 - . ~ 0 1000 2000 0 1000 2000

Frequency, Hz

Figure A-8.As Measured Narrowband Spectra for CF6-50 Engine Condition at 64.6 Percent Thrust, Reading 13

ECCP Phase I11 Combustor Tes t , 1977


Reading 1 8

160 . - - 160

150 - 150

130 - Plane 3.0 , Plane 3.5

120 o MADAR Probe 120 . 282' Borescope e OASPL = 170.6dB OASPL = 174.6dB

110 - . . . . , . , . . , . . . . , . . . . 110 . , . , , hl

E \

Plane 8.0 Exhaust Probe A OASPL = 156.ldB

Plane 4.0 o HPTN Probe = 110 .

OASPL = 178.5dB - 120 ' 4

100 . 4

1 1 0 - . . . - 1 , . . . . I . . . . , . . . . , 4

90 , 1 " " 1 ' ~ ' . ~ " . ~ ~

0)

k

C e l l Probe M c OASPL = 130.5dB .;1 120 1

Plane 8.0 Exhaust Probe B OASPL = 151.7dB

.--I. 0 1000 2000

Frequency, !IZ

Figure A-9.As Measured Narrowband Spectra f o r CF6-50 Engine Condit ion a t 82 Percent T h r u s t , Reading 18

ECCP Phase I11 Combustor Test, 1977


Reading 43

rl

X N

QI bd

cs 130 a Plane 3.5

120 1 102O Borescope 0) 5

OASPL = 175.5dB

: Plage 3.5 282 Borescope

o. OASPL = 175.8dB

Frequency, Hz

Figure A-10. A s Measured Narrowband Spectra for CF6-50 Engine Condition a t 82 Percent Thrust, Readinn 43

o ECCP Phase I11 Combustor Teet, 1977


Reading 33

160..

Plane 3.0 MADAR Probe

o OASPL = 172.0dB OASPL = 175.8dR 110. . . . . , . . . . , , , , . , . . . .

\ Z

160 4 I 0

X

CJ 140 - aJ k e Plane4.0 . . 110. 130 HPTN Probe a OASPL = 180.5dB Plane 8.0 2 120 Exhaust Probe A a~ 'a OASPL = 156.9dB

110 ;. . . . , . . . 9r - r , l l l * l - v ~ q * d

-t 140 - --- Cell Probe OASPL = 130.6dB

Plane 8.0 Exhaust Probe B OASPL = 151.9dB

0 1000 2000 0 1000 2000 Frequency, Hz

Figure A-11.A~ Measured Narrowband Spectra for CF6-50 Engine Condition at 94.6 Fercent Thrust, Reading 33

APPENDIX B - 113 9CTAVE BAND SPECTRA RESULTS

This Appendix contains the 113 octave band spectra resul tn from 50 t o 2000 Hz in both tabular and graphical form for the eight res t co~ . . ' ~ . i ons noted i n the acoustic t e s t matrix. The data presented is the measured internal spectrr; with corrections for ambfent frequency response l o s s .

Table B-1.

One-Third

Octave Band

Sp

ectr

a f

rom

EC

CP

Ph

ase

111

CF6-50

En

gin

e T

es

t.

I a)

3.8

XF

(Idle)

n I I

- -

----

----

-.--

--.- --

---

--

----

--..._--

_-

I___

-----.

. _j

Fr

eque

ncy

3.0

3.5

3.5

4.0

8.0

8.0

3.0

3.5

3.5

4.0

8.0

8.0

I

Tab

le

B-1

. O

ne-

Th

ird

O

ctav

e B

and

Sp

ectr

a f

rom

ECC

P P

has

e 111

CF

6-50

E

ngi

ne Test,

(co

nti

nu

ed).

-.

. --

---

-

c)

30%

F

(Ap

pro

ach

) n

d)

29

.7%

Fn

(Ap

pro

ach

)

t (1

00

% P

ilo

t F

uel

)

i (5

0/5

0 Fuel

Sp

lit)

i

I - -

---

- --.

- .-

-. -

.. ....

. .

- - - -

+ -

. - - -

--

. - -

.- - - .

- - -

- -- ---

--

- ---

.---

---.-.

i I

Fre

qu

ency

3

.0

3.5

3

.5

4.0

8

.0

8.0

3

.0

3.5

3

.5

4.0

8

.0

8.0

i (HZ

1 (1

02

~) (2

82')

(A

) (B

) I

(10

2~

) (282

O)

(A)

(B)

!

FPL

I

t-' 0

6,

~ Y W ' $ ? " ? 1 ~ ? 9 9 ~ m m n b-o - N O Q Q ~

C ) 0 P J e n O N N N C C C c r - V C C C P C

- . . . . . . . .4"";1 C1aa Q b b b b V m o o o n o o n n m o o - - - --- --- ---1

oJGcDvQJ~ID~~~~~~~~cD~~ v v o v v v nnn

. C.. e l ' .. . ' I .._- '

Table K

-1.

One-

Thir

d Octave Band

Spectra from ECCP Phase 111

CF6-

50 Engine Test,

(concluded)

- . -

. -

- . . .

. .

- . . -

. .

. . . . .

Frequency

1 3.0

3.5

3.5

4;0

8.0

8.0

Cell

-1

overall

1 166.2 172.0 172.0 176.3 147.9 145.1 127.6

FPL

1

- .

. . . - - . - - -

. . - . . - -

.

h)

94.6%

F (T

akeo

ff)

n -

- .

---

- -

- - .

.-. . - - - . -

. . - - - --

- - -

.

,

3.0

9.5

3.5

4.0

8.0

8.0

Cell

1

I 1 1 COHBUGTOR THRU8T. IDLE

1 1 0 4 : : : : : : : : : : : : : : : : : : : l 60 80 126 400 916 600 800 1260 ZOO0

FREQUENCY I HZ

Figure B-1. One-Third Octave Band S p e c t r a f o r CF6-50 Engine Condition a t 3 .8 Percent Thrus t

PHABE I11 COMBUSTOR TEST a a 7 PCT. THRUBT 0

I . L l l . . . I . . I . A . . . . . . . . 1 . 1 1 . 1 . 1 1 1 . A I L

SO 80 126 400 316 600 800 1460 2000 FREQUENCY. HZ

F i g u r e B-2. One-Third Octave Band S p e c t r a f o r CF6-50 Engine Condi t ion a t 19.7 P e r c e n t T h r u s t

180 ECCP PHB8E $11 COWBUGTOR TEST ElPl.mtE3.0 30 -0 PCT- THRUST* QPPROflCH )#OM PROW

8 P U N E 3.6 BORE8CWE

A PUWE 4.0 nrtw PROBE + P U W a.0 a n . PROW 0

130-

1 P m : ; : : : : : : : : : : : : : : : : 60 80 I05 200 316 600 800 1260 PO00

I FREQUENCY HZ

Figure B-3. One-Third Octave Band Spectra f o r CF6-50 Engine Condit ion a t 30 Percen t Thrust

ECCP P W 8 E I I I COWU8TOR TEST L9 KCWlE 3.0 ' 28.7 PCT. THRUST ~ r R O S aw

A M E 4.0 tttn r- + M E 8.0 . r m b

Figure B-4. One-Third Octave Band Spec t ra f o r CF6-50 Engine Condit ion a t 2 9 . 7 Percent Thrust

ECCP P H ~ E f I 46.4 PCT THRUST

A ? W E 4.0 HPTN tR= + ?W 8.0 LXH. PROM B

F i g u r e 0-5. One-Third Octave Band S p e c t r a f o r CF6-50 Engine Condi t ion a t 45.4 P e r c e n t Thrus t

- --

ECCP PHA8E 1 I I O H B U ~ T O R ~ € 8 1 CI LA)# 8-0 84.6 PCT. THRUST rmm rftm O W

r ruw 4.0 WTM m e + KCIWL 8.0 # H e m 0

t 2 0 4 : : : : : : : : : : : : : : : : : : : d 60 80 126 ZOO 316 600 800 1260 LOO0

FREOUENCY I HZ

Figure B-6. One-Third Octave Band Spectra for CF6-50 Engine Condition a t 64 .6 Percent Thrust

. , . , , . . , , . , I . , . . . . , 1 . 1 I . . . . . . . . 1 1 1 1 . 1 1

60 00 126 200 916 600 800 1260 2000 FREQUENCY I HZ

Figure B-7. One-Third Octave Band Spectra f o r CJ6-50 Engine Condit ion a t 82.0 Percent Thrus t

1 2 0 ' : : : : : : : : : : : : : ~ 60 80 125 200 315 600 600 1260 2000

FREQUENCY. HZ

Figure B-8. One-Third Octave Band Spectra for CF6-50 Engine Condition a t 94.6 Percent Thrust

APPENDIX C - TURBINE TRANSFER FUNCTION RESULTS

The t u r b i n e t r a n s f e r func t ion r e s u l t s presented i n t h i s Appendix a r e i n the form of coherent 113 o c t w e band SPL s p e c t r a comparisons a t each t e s t condi t ion. Turbine a t t e n u a t i o n s were obtained between t h e coherent s p e c t r a a t Plane 4.0 and t h e coherent s p e c t r a a t Plane 8.0, both of which a r e r e f - erenced t o Plane 3.5. Noted on t h e s p e c t r a a r e reg ions where both s p e c t r a a r e i n t h e n o i s e f l o o r 18 dB d ~ w n from t h e raw s i g n a l . Also noted on t h e s p e c t r a a r e regions of f requen .2~ bands which a r e 15 t o 18 dB down from t h e raw s i g n a l f o r e i t h e r Plane 4.0 qr Plane 8 . 0 , but no t both. The a c t u a l a t t e n u a t i o n s obtained i n these reg ions a r e e i t h e r less than o r g r e a t e r than t h e observed value read from t h e p i c t s .

lgO T N o l s e Floor 1

a) 3.8 ;d Fn, RM; 6

, *- Noise Floor -1

Plane 4.0 re 3.5

Plane 8.0 re 3.5

I i I

Frequency, Hz

b) 19.7 % Fn. RDG 16

Figure C-1. Coherent Spectra Comparisons Eur Turbine Attenuatioas From ECCP "haae 111 Engine T e s t

-I *- Noise Floor

i I y P l a n e 4.0

C) 30 % Fn, RDG 23 ( 100 % P i l o t Fuel )

*- Noise lo or i J

Plane 4.0 re 3.5 : r,/A --*L, -

- -

.* - - - - - - -

I I I I ~ ~ ~ I I I I I I I ~

Frequency, Hz

d ) 29.7 % Fn, RDG 39 ( 50 / 50 Fuel Spl i t )

Figure 8;-1. Coherent Spectra Comparisons For Turbine Attenuation From ECCP Phase 111 Engine Test ( Continued )

e ) 45.4 % Fn, RDG 28

f Plane 4.0 3.5

f Plane 8.0 re 3.5 I 1

Frequency, Hz

f) 64.6 % RDG 60

Figurec-1. Coherent Spectra Cocnparisone For Turbine Attenuation From ECCP Phase I11 Engine Test (Continued )

f Plane 4.0 re 3.5 1 190 7 1

9) 82 % Fn, RDG 18

-

-

2 1 9 Q

2 I - 0

1 *- Noise Floor

*- Noise Floor -

Plane 4.0 re 3.5

Frequency, Hz

h) 94.6 % Fn, RDG 33

Figure C-1 , Coherent Spectra Comparisons For Turbine Attentuat ion From ECCP Phase 111 Engine Test (concluded )

APPENDIX D. EVALUATION OF PRCBE RESPONSE AT ELEVATED TPIPERATURES -.

This Appendix contains the equations, basic rs~unptions and results of a study conducted on the SCCP Phase I11 lntersal KIite probe measureaents to evaluate the effect of elevated temperature an t b engine test results.

APPENDIX D - EVALUATION OF PROBE RESPONSE AT ELEVATED TEMPERATURES

The theo re t i ca l value of probe a t tenuat ion determined from the mathe-

matical equations of Arthur S. I b e r a l l (Reference 12) includes e f f e c t s of

compressible f l u id flow, f l u i d acce lera t ion , an.! h??t condvction. The pres-

sure a t any point i n the l i n e t o the pressure a t the probe entrance is re-

la ted in the form of an amplitude r a t i o &ich .s a function of frequency.

A s implif ied computation can be made when assuming no r e f l e c t i o n i n the

system such as i n the c2se of a probe terminated i n a constant cross-section

" inf in i te" l i n e (a l i n ~ i n which r e f l ec t ions from its end can be neglected).

The bas is of t h i s s implif ied calculat ion, a s given i n Olson (Reference 13) is

the l o s s i n tubes due t o t he v iscos i ty of a i r .

Transmission l o s s i n tubes of c i r cu l a r sec t ion can be expressed as

where A = amplitude (pressure) a t a d i s tance x from the amplitude Ao, which

e x i s t s a t the probe entrance. The other terms a r e given below:

, transmission coe f f i c i en t

R = radius of tube

x = length from entrance t o point of i n t e r e s t

c = veloc i ty of sound

w = 2 r f

f = frequency i n Hz

p = v i scos i ty coe f f i c i en t , f o r a i r

p = densi ty of medium

y = r a t i o of spec i f i c hea ts

The amplitude r a t i o becomes

A/Ao = e , a function of frequency.

The transmission l o s s (at tenuat ion) is

TL = 20 log

This a t tenuat ion is due so l e ly t o viscous damping within the probe tube and

assumes no i n t e r n a l r e f l ec t ions and no non-linear e f f e c t s associated with

sound pressure l e v e l s over 120 dB.

The waveguide sensor systems used i n t he ECCP Phase I11 program under-

went room temperature ambient pressure ca l ibra t ions . The transmission l o s s

i n the sensor tubes located i n the Double Annular combustor region as a

r e s u l t of the elevated operating temperatures was estimated assuming t h a t - 1. The temperature i n the waveguide tube reaches the s ink temperature

under the engine cowling of approximately 477 t o 505K.

2. This temperature is constant f o r a l l operating conditions.

3. The pressure i n the tube w i l l be t he s t a t i c pressure in the combustor a t t he measuring plane.

4. An a i r mediumexists.

For waveguide systems a t Planes 3.0 and 3.5

Conditions

A. Calibrat ion

B. Takeoff (Worst Case)

Transmission coe f f i c i en t at f = 2000 HZ

a, = 0.2789 l/m f c r ambient ca l ib ra t ion

a1 = 0.0643 l / m f o r takeoff condition

Transmission l o s s across tube length x = 0.3048 m f o r f = 2000 Hz

A TL = 20 log -

A,

TL, = - 0.74 dB f o r ambient ca l ib ra t ion

TL1 = - 0.17 dB fo r takeoff condition

Ratio of transmission lo s ses

The frequency response l o s s determined from ambient ca l ib ra t ions is

approximately four (4) times grea te r than the l o s s determined a t the takeoff

condition of the CF6-50.

This amounts t o a reduction of the response correct ion from +2 dB a t

2003 Hz t o M.5 dB a t takeoff operating condtion. This 1.5 dB change is

based on the l i n e a r assumption f o r SPL's less than 120 dB. Since the major-

i t y of the f luc tua t ing pressure measurements were well above 120 dB (140 to

160 dB) and the AdB change was approaching the measurement accuracy, t h i s

temperature e f f e c t w a s neglected and the ambient response r e s u l t s were used

t o correct the data.

APPENDIX E, CROSS-CORRELATION RESULTS

This Appendix contains a tabulated summary of time delays and amplitudes for the major peaks of the cross-correlations between pairs of Kulite sensors for a l l t e s t conditions on the ECCP Phase 111 CF6-50 engine test.

Tab

le E

-1.

Tim

e D

elay

s an

d A

mp

litu

des

of

C

ross

-Co

rrel

atio

n

Pea

ks

for

EC

CP

Ph

ase

I11 T

est

Dat

a.

-.--

-.--

---.

-..-

.-

-

L- P

ct

Th

rust

T

ape

Rea

din

g

- ___

____

___ _

Pla

ne

3.0

to 8

.0

TI

(mse

c)

R~

yl

'2 (m

sec)

%y2

C

(m

/s>

Pla

ne

3.0

to 4

.0

TI

(mse

c)

R~

l

'2 (m

sec)

!hy2

C

(m/s

)

Pla

ne

3.0

to 3

.5

'1

(mse

c)

R~

l

r2

(mse

c)

hy

2

C

(m/s

>

Pla

ne

3.5

to

4.0

(282

O)

'1

(mee

c)

R~

l

~2

(mse

c)

R~

2

(102

')

TI

(mse

c)

Ex

yl

C

(m/s

>

,

. ._

____ _

_. ..

__._

_______._

--___I-

-.--

---

-- - .--

-

3.8

19

.7

30

.0

29.7

45

.4

64

.6

82.

94.6

6

16

23

3 9

3

4/2

8

4011

3 4

3/1

8

33

-___

_ .

_ --

--

.. __

..

-_ .

- .

--

--.-

--

----

Po

int-

to-P

oin

t D

ista

nc

e,

d =

2.7

64

m,

dZ

= 2

.527

m

+2.

5 +

5.0

+5.

0 +

6.0

+4.

0 +

5.0

--

- +

2.0

.6

.1

.13

.0

7 .0

4

.03

--

- .0

4

+5.

0 +

2.5

+2.5

--

- --

- --

- --

- --

-5

6

.05

.0

8

---

---

---

---

---

46

4.5

49

8.0

510.

2 51

0.8

529.

7 5

39

.5

557.

8 56

3.6

-- - -

---

Po

int-

to-P

oin

t D

ista

nce

, d

= 0

.87

2

m,

d,

= 0.

634

m

-1.0

-1

.5

-1.5

-1

.5

-1.5

-1

.5

-1.2

-1

.5

.18

.1

3

.8

.14

.1

.08

.1

.06

+

4.0

+4

.0

+6.

0 +

7.5

+l.

0

+1

.0

+3.

5 +

2.5

. 23.

.03

.0

2 .0

5

.06

.04

.0

3

.03

5

15

.1

566.0

59

5.6

595.

9 6

18

.1

64

4.6

6

68

.4

68

2.1

Po

int-

to-P

oin

t D

ista

nce

, d

= 0

.734

m

, d

Z =

0.

357

m

-2.0

-1

.5

-1.5

-1

.5

-1.5

-1

.5

-1.5

-1

.5

.3

.3

.24

.2

1.5

.1

2

.07

.14

-6

.0

-3.0

+

6.0

+4.

5 +

4.0

+4

.0

---

+4.

0 .1

8

.16

.0

7 .0

6 .0

5

.03

--

.0

6

41

9.1

48

2.5

50

4.1

5

05

.1

530.

6 5

45

.3

56

6.3

57

3.9

---

Po

int-

to-P

oin

t D

ista

nce

, d

= 1

.27

1 m

, d

, =

0.2

80

m

-2.0

-2

.0

-1.8

+

1.5

-1.5

-1

.5

-1.5

-1

.5

15

.0

8 .0

7 .1

0

.04

.04

.03

.0

3

-7.0

+

1.5

+

1.5

-2.0

+

5.0

+6.

5 +

1.5

+6.

5 .1

6

.08

.0

2 .0

7 .0

4

.03

.0

3

.03

Po

int-

to-P

oin

t D

ista

nce

, d

= 0

.28

1

m,

dZ

= 0

.28

0

m

---

-.30

0

---

---

---

---

-- .-

--

.18

.1

0 --

- --

- --

- --

- --

51

5.i

5

66

.0

595.

6 59

5.9

61

8.1

6

44

.6

66

8.4

6

82

.1

Tab

le E

-1.

Tim

e D

elay

s an

d A

mp

litu

des

of

C

ross

-Co

rrel

atio

n

Pea

ks

for

ECCP

P

has

e I11 T

es

t D

ata.

(C

on

clu

ded

)

Pla

ne

4.0

to

8.0

I

Po

int-

to-P

oin

t D

ista

nce

, d

=

1.8

93

m

, d

, =

1

.89

3 m

I

PC t

Th

us

t T

ape

Rea

din

g

TI

(mse

c)

Rx

yl

r2

(mse

c)

3.8

1

9.7

30

.0

6 1

6

2 3

39

34

/28

40

113

43

/18

29

.7

45.4

6

4.6

8

2.0

Pla

ne

3.5

(1

02O

to

282

O) I

Po

int-

to-P

oin

t D

ista

nce

, d

= 1

.26

8

m,

d,

= 0

m

TI

(mse

c)

Rx

yi

~2

(mse

c)

,Rxy

2 C

(r

nls

Pla

ne

8.0

(A

to

B)

I P

oin

t-to

-Po

int

Dis

tan

ce,

d =

0.1

28

m, dZ =

0.1

28

m

I

Pla

ne

3.5

to

8.0

0.0

0

.0

0.0

0

.0

0.0

---

--

- .8

.7

6 .7

5

.17

.4

---

---

+3.

0 +.

25

+. 3

---

---

---

.2

.1

.06

--

- --

- 51

0.2

51

3.3

51

6.6

51

6.6

52

8.5

533.

7 5

49

.6

553.

2

Po

int-

to-P

oin

t D

ista

nce

, d

= 2

.030

m

, dZ

=

2.20

4 m

'1

(mse

c)

R~

l

T2

(mse

c)

3972

C

(m

/ s

+3.

5 -4

.5

+5.

5 +

6.0

+. 7

+

.3

---

-- .2

4

1.6

1.1

.0

4

.04

.03

--

- --

---

+2.

0 +

3.0

---

---

---

---

---

---

.06

-08

--

- --

- --

- --

- --

464.

6 49

7.9

510.

4 51

0.8

529.

6 5

39

.5

55

8.0

56

3.6

APPENNDIX F - NOMENCLATURE

Symbol Description

Annulus area, cross-sectional area at combustor measurement planes

A, A. Amplitued (pressure) of wave

Ar ef Combustor reference area

BPR Fan-to-core bypass ratio

c, C Acoustic velocity - ,. L Average acoustic velocity

d Point-to-Point distance between Kulites

dB Decibel

dz Axial distance between Kulite sensors

A~BPWL Turbine attenuation based on PWL

@P Unit vector nornal to acoustical wave front

ex Unit vector, axial component

f Frequency

f /a Fuel-air ratio

FAR3 6 Combustor fuel-air ratio based on W36

% Fn, PCT FN Percent net thrust

FPL -5 2 Fluctuating pressure level re 2 x 10 ~ / m

FPWL Power level based on fluctuating pressures re 10-13 watts

HPT High pressure turbine

HPTN High pressure turbine nozzle

HRR Heat release rate (total fuel flow x heating value of JP5 fuel)

Hz Hertz: cycles/second

Units - m

2

watts

In t ens i t y f l u x vector

Ins ide diameter

Low pressure turbine

Mach number

Average Mach number

Malfunction, detect ion, ana lys i s and recording subsys tem

N c

N f

OAFPL

Core speed, corrected

Fan speed, corrected

Overal l f l uc tua t ing pressure l e v e l

OAPWL Overal l sound power l e v e l

Overal l sound pressure l e v e l OASPL

One-third octave band f l uc tua t ing pressure l e v e l 113 OEFPL

113 OBSPL One-third octave band sound pressure l e v e l

Outside diameter

Acoustic pressure (rms)

Percent f u e l s p l i t t o main s t age burner

Percent f u e l s p l i t t o p i l o t s t a g e burner

PCT I N

PCT OUT

Ambient, standard day preseure atm

atm

8:'

4,

S t a t i c pressure

Tota l pressure

PCT combustor pressure drop r e l a t i v e t o upstream pressure

Combustor t o t a l pressure drop atm

dB Sound power l eve l , re 10-l3 wat t s PWL

Gas constant f o r a i r

Radius of tube

RDG, Rdg Reading number

R x ~ Cross-correlation function

SLS Sea level static

SPL

f

w3

W36

WFT

Sound pressure level, re 2 x ~h~

Ambient, standard day temperature

Total temperature at turbine inlet based on pilot fuel-air ratio

Static temperature

Total temperature

Combustor total temperature rise

Transmissicn loss

Mean flow velocity

Absolute flow velocity vector

Core jet velocity

Effective (mixed jet) velocity

Fan jet velocity

Compressor discharge airflow

Total combustor discharge airfl~w

Total fuel flow

Combustor discharge flow corrected to corn; ,:stor inlet conditions

Axial Cartesian coordinate

3istance along probe tube

Tangential Cartesian coordinate

Transmission coefficient

SUBSCRIPTS

a

C

f

0

1

1

?,

Difference, i n c r e a s e o r decrease

Ra t io of s p e c i f i c h e a t s

Time de lay

Angle between a c o u s t i c wave f r o n t arid a x i a l d i r e c t i o n

Vi.sccsity coef f i c i ency f o r a i r

Density

Angle between flow and a x i a l d i r e c t i o n

Angular f requsccy, 2rf

A i r

Core

Fuel, f a n

I n i t i a l o r c a l f b r a t i o n cond i t ion

T ~ S L -' ..-.-. . ~ n

Largest. peak i n crcibs-.correlation

Next l a r g e s t peak i n c ross -cor re la t ion

Compressor d ischarge plane

Combustor i n l e t p lane

Combustor d ischarge / t u r b i n e i n l e t p lane

Core nozzle diecharge pla,ae

sec

deg

Pa. s e c

kg i m 3

REFERENCES

Gleason, C. C., auld Bahr , D.W. , "Experimental Clean Combustor Program Phase 1x1 F i n a l Report," Report No. NASA CR-135384, t o b e published.

Bahr, D.W. and Gleason, C .C . , "Experimental Clean Combustor Program - Phase I F i n a l Report," Report No. NASA CR-134737, June 1975.

Emm~rling , J. J. , "Experimental Clean Combustor Program - Noise Measure- ment Addendum - Phase I F i n a l Report," Repr r t No. NASA CR-134853, J u l y 1975.

Gleason, C.C., Rogers, D.W., Bahr, D.W.; "Experimental Clean Combustor Program - Phase 11 F i n a l Report," Report No. NASA CR-134971, August 1976.

Ermerling, J. J. and Bekofske, K. L. , "Experimental Clean Combustor Program - Noise Measurement Addendun - Phase I1 F i n a l Report ," Report No. NASA CR-135045, January 1976.

Doyle, V.L. and Matta, R.K., "Attenuation of Upstream-Generated Low Frequency Noise by Gas Turbines, "Report No. NASA CR-135219, August 1.977.

Karchmer, A.M., Roskotko, M. , and Montegani, F .J . , "Measurement of I a r f i e l d Combustion Noise From a Turbofan Engine Using Coherence Func- t ions ," AIAA Paper 77-1277, October 1977.

Reshotko, M. e t a l . ; "Core Noise Measurements on a YF-102 Turbofan ~ n g i n e , " NASA TMX-73587 (AIAA Paper No. 77-21), January 1977.

Matta, R. and Mani, R., "Theory of Low Frequency Noise Transmission Through Turbines," NASA CR 159457 , t o be published.

Karchmer, A. and Reshotko, M., "Core Noise Source Diagnost ics on a Turbofan Engine Using Cor re la t ion and Coherence Techniques ," NASA RIX- 73535, November 1976.

Matta, R.K., Sandusky, G.T. and Doyle, V.L., "GE Core Engine Noise I n v e s t i g a t i o n - Low Emission Engines," Report No. FAA-RD-77-4, February 1977.

I b e r a l l , Arthur S. , "Attenuation of O s c i l l a t o r y Pressures i n Instrument Lines," Journa l of Research of t h e NBS, Vol. 45, p85-108 J u l y 1950.

Olson, Harry F. , "Acoustical Engineering," 1957 D. Van Nostrand Company, Inc .

experimental clean combustor program …...'for sale by the national technical lnfamat~on...

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