NASA CR-134969

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EXPERIMENTAL CLEAN COMBUSTOR PROGRAM i'

PHASE It

FINAL REPORT

by i

R. Roberts, A. Peduzzi,G. E. Vitti

UNITED TECHNOLOGIES CORPORATION

PRATT & WHITNEY AIRCRAFT GROU2----

COMMERCIAL PRODUCTS DIVISION

November 1976

,,

9

Preparedfor

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

NASA LewisResearchCenter

Contract NAS3-18544 ,_,_.-_-__

(hA_A-C_-IJq_ _9) _Xl- _,i._I i. NI A_ CL.e.A, _ N77" lJub9

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('R- 134969 I

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Novelnher 1976

6 Folfo_n.n_ Ot_nlzation Cod0

I-XI_I_RIMI'NTAL CI.EAN COMIIUSq'OR PROGII, AM, PIIASh II

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R, Roberls, A. Peduzzi, (;. l!. Villi -_ wc.k u_,an_

9. Plh'formingOt_tmlat,on Name irld Addresl Lhlited Tecbuok)gies Corpoiatiol-_---

Pratt & Whitney Aircraft Group "__ ColRr=ctOt GtasltNo.

C(munercial Products Division NAS3-18."_44

= _ fast Ilartford, Collnectictl[ 06108 13 Type el Report 8rid Petted COveCecl

Sl_onso,'incJ A_rtcy Nlm4_ and Ad(ffe_,s Colltrael l_.eport

National Acromlufics aud _pace Administration

Lewis I_eseart:b Celtter 14.SpOntormA9gencCy ot_

21000 Brookpark Road, ClevelamI, Obh> 44135

; 16. _ppilmenttry Nc,te_

16. Abttrlct

Priluary objectives of Ihe NASA/Pratt & Whitney Aircraft I_xperimetxtal ('le;m Combustor Progralu are

tile geuelatioll of CelllbuMor p,allution redllctiol_ Icchuok_gy for CGllllllercial ('1"Oi. ell#lieS aBd delnonstration of this technob)gy in a full-scale JTgD ell#lie ill 1076. Phase [l ¢OllSiMed of conll)OtlCllt riG refillelllelll tesling of the two best PIla_e I combustor concepts. These COllCepts are the Vorbix conlbustor,

¢oulinlied |roB) I)b;l_e I. add ;J IJybrid ¢oBlbuslor which ¢oJtlbiBe_, the pilot ZOlle of Ihe Staged Premix

etH11buslor alld tile nlain zolle of the Swirb('an ¢'.ultbllMor. Itolh Phase II coneelHs significantly reduced

all pollutant etllJssions relative to the JTqD.7 ell#u¢ coil)busier. IIowever. JleJlheJ concept met all program goals. 1"he Ilybrid ¢oulbuslc, r tiler polhltiou goals for unburlled hydrocarbons and carbon lllol'=-

oxide bu* _]i¢luot achieve tile oxides of nitrogen goal 11tis eombustor ]lad sigBJfieanl pearl)finance deftcieneies. The Vorbix combustor met goals for unburned hydrocarbons and oxides of nitrogen but did

not achieve tile carbon nlonoxide goal. PerformaJlce of file VorbJx ¢ombuslor approael)ed tile erJgJne

requirements. On rite basis of these results, the Vorbix combustor was selecled for tltc Pbam 111englue dell)oDslralJol) ])rograul.

A eol]trol sllldy was ¢oBdueled as part of lbe J)base II effort to establish fuel eolltrol requireu)ents imposed by the low-emission combustor co|loci)Is and to Identify coueeptl;al control system desigus. Coll-

. current efforts were also completed oil two addendums: an Alternate Fuels Adden(bmt aml _ Combustion Noise Aildeudmu.

It. Kly Wohlt (Su_AsI_ by AulhO/(lll Ifl. [tttrihutlon Slallmlnl

Main Burner I-inissh)ns

J'l't)l) (.'OlllbUSIor

Fmissioll Standards

CTOL I'misshms

Unclassified (/ttclassifled ] 1_7 t

.... ' F0t sale by theNal_onalTechmcalJnlomahonSemce SOtmeheldW. tem_a 72151
FOREWORD

+ This document describes tile work conducted and completed by tile Commercial Products

Division, Pratt & Whitney Aircraft Group of United Technologies Corporation during Phase

II of the Experimental Clean Combustor Program. This final report was prepared for the

National Aeronautics and Space Administration Lewis Research Center in compliance with

the requiremetlts of Contract NAS3-18544.

The authors of this report wish to acknowledge the guidance and assistance of Mr. Richard

; Niedzwi'ecki, NASA Project Manager of the Experimental Clean Combustor Program and ,,.,,

Mr. J. R. Baker of the Ad,,anced Controls Group of Pratt & Whitney Aircraft.

C

iii
TABLE OF CONTENTS

Section Title Page

1 F,XI)ERIMF.NTAI, CI,I:,AN COMI]USTOR PROGRAM DESCRIPTION 4

A. Gencral I)cscription of Overall Program 4

B. Program Plan 4

1. Plmse 1Program 4

2. Phase 11Program 5

3. Phase II1 Program 5

C. Program Schedule 6

D. Program Goals 6

I. Pollution Got:Is 6

2. l'erlk)rmance Goaks

11 EQUIPMENT AND EXI)EP, IMENTAL PROCEDURES 9

A. Reference Engine _lndCombustor 9

1. Reference Engine Description o

2. Reference Combustor Description I0

3. Reference Engine Combustor Pollution Levels 12

B. Tcst Combustor Concepts and Configurations 13

1. Ilybrid Combustor 13

2. Vorbix Combustor 19

C. Test Facilities and Equipment 28

I. Test Facilities 28

2. Test P,igs 31

D, Test Conditions 33

1. lligh-Prcssure Test Facility 33

2. Altitude Test Facility 35

E. I)ata Analysis Procedures 36

I. l'crformal_cc Data 36

2. Emission Data 37

I11 PIIASF 11PROGRAM RESULTS 41

A. Introduction 41

B. l lybrid Combustor 4 I

1. Emission Results 41

2. Perform:race Results 49

C. Vorbix ('oml_ttstor 55

1. Emission Results 55

2. Pcrlk)t-mancc Resulls 67

D. ConlbLisl¢)rStatus Assessment 74

I.:. Phase I11 Eulgine I)csign 77

IV FUF,L CONTP, OI, I)FS1C;N STUI)Y 79

A. Introduction 79

B. Fucl Control Design Rcquircments 79

¥

1l_.l,t,,l,,,INt3 PAGI._ BLANIC NO_? FILMI_)
: TABLE OF CONTENTS (Cont'd)

Section Title Page

C. Fuel Control Breadboard Design for Phase [1I 81

Engine Testing

V CONCLUDING REMARKS 83

Appendix A Combustor Configurations 85 .,,

Appendix B Experimental Test Data 109

Appendix C Fuel Control Design Study 133

Appendix D Nomenclature 173

Appendix E References 175

Distribution List 177

Z

vi
LIST OF ILLUSTRATIONS

Figu re Title Page

1 NASA/Pratt & Whitney Aircraft Experinlental Clean Combustor

Program Schedule 6

2 Cross-Sectional Scllem:ltic of tile JTgD-7 Reference Engine 9

3 Cross-Sectional Schematic of tile JT9D-7 Reference Combustor 10

4 Reference Combustor Exit Average Radial Temperature Profile 11

5 Cross-Sectional Side View of Hybrid Combustor 13

6 Exterior View of Hybrid Combustor 14

7 Front-End View of llybrid Corn buster 14

8 Pressure Loss Profiles Showing hnprovement at Main Burning Zone

Entrance Plane Obt'dned for Phase 11Hybrid Combustor 15

9 Swirl Vorbix Combustor Dcsign 19

10 Extcrior View of Vorbix Combustor Configuration S11 20

11 Front-End View of Vorbix Combustor Configuration S11 20

12 Designs of Baseline Swirl Vorbix Combustor anti Combustol with

Increased Pilot Zone Volume 24

13 Swirl Vorbix Combustor Pilot Burner Front-End Designs with

Pressure Atomizing Fuel Injector and Low Pressure Drop Aerating

Fuel Injector 25

14 Swirl Vorbix Combustor Main Zone Swirler Orientations Evaluated

in Phase I1 26

15 Swirl Vorbix Combustor Design with Premixing Tubes 27

16 Detailed View of Combustor Exit Annular Traversing Rake 29

17 Combustor Exit Ammlar Traversing Rake Ilead Showing Disassembly

of Thermocoul_lc llcad Froln (;as Sampling llead 30

18 Schematic I)hgram of Test Rig Installed in the Test Facility 31

vii
i

LIST OF ILLUSTRATIONS (Cont'd)

Figure Title Page

19 Comlmstor Rig Outer Case Showing Sectioning to Improve

Accessibility 32

20 V.orbix Combustor Design Showing Rear Outer Seal Arrangement 33

21 JT9D Relight Envelope 35

22 Combustor Carbon Monoxide Emission Levels as a Function of

Fuel-Air Ratio at Idle Conditions with Bleed 43 .............

23 Circumferential Exit Temperature Patterns for Hybrid Combustor

Configuration H 1 at Idle Conditions 50

24-_Circumferential Exit Temperature Patterns lor Iiybrid Combustor

Configuration H2 at Sea0Level Take-Off Conditions with Fue -l._kir__

Ratio of 0.616 51

25 Radial Exit Temperature for Hybrid Combustor Configuration H2

at Idle Conditions 53

26 Radial Exit Temperature Pattern for ttyhrid Combustor ConfigUration

112 at Sea-Level Take-Off Conditions with Overall Fuel-Air Ratio of

0.0160 and Pilot Zone Fuel-Air Ratio of 0.0052

27 Minimum Pressur_ Blowout Test Results for Itybrid Comhustor

Shown on JT9D-7 Altitude Relight Envelope 55

28 Relationship Between Emissions of Oxides of Nitrogen and Efficiency

for Vorbix Combustor Configurations $23 and $25 60

29 Efficiency Characteristics for Vorbix Combustor Configuration $25

as a Function of Pilot Zone Fuel-Air Ratio 66

30 Circumferential Exit Temperature Patterns for Vorbix Combustor

Configuration $20 at Idle Conditions 68

31 Circunfferenlial Exit Temperature Patterns for Vorbix Combustor

Configuration $20 at Sea-Level Take-Off Conditions wifl_ Fuel-Air

Ratio of 0.016 69

32 Radial Exit Temperature Pattern for Vorbix Combustor Confignration

SI6 at Idle Conditions 70

viii
LIS1 OF ILLUSTRATIONS (Cont'd)

: Figu re Title P.age

.. 33 Radial Exit Temperature Patterns for Selected Vorbix Combustor

Configurations at Sea-Level Take-Off Conditions 71

; 34 Altitude Stability and Relight Characteristics of Vorbix Combustor

Configurations S19 and $20 Shown on JT9D-7 Altitude Relight

Envelope 72 ...'.

35 Altitude Stability and Religbt Characteristics of Vorbix Combustor

Configurations $24 and $25 with Modified Pilot Swirler 73

36 Altitude Stability and Re!ight Characteristics of Vorbix Combustor

Configuration $24 with hlcreased Pressure Drop Pilot Fuel Nozzles 73

37 Emission Characteristics for Oxides of Nitrogeu Versus Percent

Engine Thrust 74

38 Emission Characteristics for Carbon Monoxide Versus Percent

Engine Thrust 75

39 Emission Characteristics tor Total Unburned Hydrocarbons Versus

Percent Engine Thrust 75

40 Preliminary Design of Phase III Vorbix Combustor for Demonstration

Engine Testing 78

41 JT9D Engine Combustor Pressure and Fuel Flow Envelopes 80

I - 42 Fuel Scheduling Requirements for Two-Stage Vorbix Combustor

" _ Based ou l'hase II Experimental Clean Combustor Results 80

43 l'reliminary Design of Breadboard Fuel Control for Phase Ili Engine

Testing 81

i .?. ix
J

LIST OF TABLES

'l':dfle Title Page

1 I_ollntion Goals and Current JT9D-7 Levels 7

11 ECCP Pertbrmance Goals 8

111 Key Specifications of tile JT9D-7 Engine 10

IV Key Operating Parameters of the JT9D-7 Reference Combustor 11

V -- Representative JT9D-7 Production Engine Emission Levels and

Experimental Clean Combustor Program Goals 12

VI Hybrid Combustor Configuration Summary 16

VII Airflow Distributions for Hybrid Combustor Configurations 17

VIII Vorbix Combustor Configuration Summary 22

IX Airflow Distribntions for Vorbix Combustor Configurations -- 23

X JT9D-7 Reference Engine Operating Conditions and Experimental

Clean Combustor Sector Rig Operating Conditions 34

XI Summary of Reported Combustor l'erformance Parameters 36

XII Definition of EPAP Conditions and EPAP Coefficients for Jl'9D-7

Engine 40

XIII Ilybrid Combustor Emissions At hlle Conditions 42

XIV llybrid Combustor Emissions At Sea-Level "l'ake-Off Col_ditious 44

XV Hybrid Combustor Emissions At Approach Conditions 45

XVI Ilybrid Combustor 13missions At Climb Conditions 46

XVll ltybrid (knnbustor Emissions At Cruise Conditions 47

XVIII llybrid Combustor EPA I'-'lrameter Emission._ Results 48

XIX Landing-Takeoff Cycle Point I';mission Indices tbr llybrkl Combustor 48

XX Sumlm_ry of I'ressure Loss Test Results lbr Itybrid Combustor 49

• I
I I I ,

LIST OF TABLES (Cont'd)

Table Title Page

XXI Sumlnary of Ilybrid Combustor Exit Tenlpcrattu'e Pattern Factor

P,estdt s ,S2

XXII Sunuuary of llybrid Coml)ustor Idle Lean Blowout Test P,esults 54

- XXlll Vorbix Colllbttstol' I'missions At Idle Conditions 56

XXIV Vorbix Combusto_ Emissions At Sea-Level 'l'ake_Off Conditions 58

XXV Vorbix Combustor F_missions At Approach Conditions 61

XXVI Vorbix Com_tstor F.missions At Climb Conditions 62

XXVII Vorbix Combustor Emissio!',s At Cruise Conditions 63

XXVIII Vorbix Combustor EPA Parameter Emissions Results 64

XXIX Landing-Takeoff Cycle Point Emission Indices for Vorbix Combustor 65

XXX Summary of Pressure Loss Test Results for Vorbix Combustor 67

- XXXI Summary of Vorbix Combustor Exit Temperature Patterns Factor

Results 70

XXXll Summary of Vorbix Combustor Idle Lean Blowout Test Results 71

• XXIII Hybrid Combustor Performance Results 76

,' XXXIV Vorbix Combustor Performance Results 76

• t d i I i

XXX.V___Steady_St,l!_. I erl_rrnanet/Enlission Test Points 84

?'2

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L 1
I

SUMMARY

" Tile Experimental Clean (_ombnstor I_l'ogran_ is directed toward the dcvcloplnent and (Icn'lt.l'Lst,'ation of technology for rcducillg polJulant emissions for application to both current and

future gas turbine engine ¢onlbustors. The prograln is being comlucte(I in three phases. Phase

Linvolved experimental rig screening of Colnbustor concepts to identify the best approach,'s

for reducing _:missio,, levels. Ph-'lse II, which is the subject of this report, consisted of evalu'ltion and refinement of tile best two combu_tor concepts identified ill Phase I. F,inph'-lsis was

placed on docunlentation of emission characteristics over the full rat_ge of operating conditions

mad development of satisfactory eombustor performance. A fuel control design study was

eondttcted to esta lish fuel nlanagenlent reqtdrcments for two-stage combustors. Phase Ill ,. ....

will consist of fnll-sc',de engine demonstration tests of the best combnstor selected from

Ehase 11.

Program gaseous pollution golds are expressed in two ways. as emission indices at the engine

operating mode where peak levels of each pollutant are generated, and as integrated EPA

parameter values. Compared with the current production JT9D-7 engine, the attainment of

these goals involves significant pollution reductions by factors of 2Yzto 5% on an emission

index basis. Program performance goals consist of maintaining or in_lproving ":tlrrent production combustor operating parameters.

Combustor testing in Phase 11was conducted in the same 90-degree sector rig used for Phase

I. The sector rig sinmlates tile JT9D engine combustor envelope and internal flowpatb. Tests

were conducted at conditions that simulated JqgD-7 engine operating conditions at idle, ap-

_ preach, climb, sca-level take-off, and cruise. The test rig matched tile engine conditions at

idle, and duplicated all parameters except inlet pressure and airflow at the higher powe_ levels.

The emissions data were corrected to account for differences between the fig operating conditions ,'rodactual engine conditions. Stability and relight capability at simnlated altitude

: conditions were also documented.

The Hybrid and Vorbix colnbustor concepts were selected for refinement in Phase I1. The

• ltybrid combustor was created by merging tile pilot zone of tbe Phase 1Staged Premix cornbuster with a main zone derived from the Swirl-Can comb_tstor. The Vorbix concept was

continued from l'hase I with evolutionary changes to the pilot and main burning zone dimensions.

Althougb none of tile llybrid combustors met all of tile EI'A parameter goals, substani.ial reductions of aU pollutants were achieved. Oxides of nitrogen levels were significantly belov,

tile JT9D-7 production engine levels, and surpassed tile program goal for two configurations.

Carbon monoxide am; total tmburned hydrocarbon levels were generally close to or surpassed

goal levels with only the pilot zone fueled at the .q_proacb condition. Smoke levels were well

below goal levels. Best rcsults were achieved with ccnfiguratiol_ 117 where EPAP wdues of

- 3.4, 3.7, and 0.7 tbr oxides of nitrogen, carben Inonc_xitlc, and total unburned hydrocarbons,

i_espeetively, were obtained.

t: _i *
The most significant pollutant problem uncovered with tile Ilyhrid combnstor occurred at

the approach condition. At this operating condition, the Ilybrid wasunable to operate with

high comhustion efliciency (and accompanying low I_ollutant levels) with the pilot and main

zones f'dly I'neled. With hoth zones fully fncled, combustion efficiencies wcrc below 91 percent. Solnc conligurations cxhibitcd combustion efficiency near I00 percent with all approach

fucl supplied to either the pilot or to the main zone. Selected configurations were tested with

fuel supplied to only a portion of tile main injectors, and, in all cases, efficiency improved as

tile number of mvJa nozzles fueled was reduced, hnplementation of any of the partial fuel

staging techniques is undesirable since it imposes additional demands on the engine fuel con ....

trol, involves ignition during a flight mode. and creates additional combustor exit temperatt,re

distribution problems.

None of the Vorbix combustor configurations met the EPA parameter goals for all pollutants.

llowever, the Vorbi.'- combustors exhibited emission levels for oxides of nitrogen that were

significantly bclow tile current JT9D production combustor level, in some cases surpassing

the program goals. Unbnnlcd hydrocarbon emission levels and smoke were also low, generally

exceeding the goal levels. None of the Vorbix combustors met the carbon monoxide emissions

goal, although some configtlrations exhibited 25 percent reductions from the production combustor level. Best results were achieved with configuration $25 where EPAP values of 2.2,

6.5, and 0.3 for oxides of nitrogen, carbon monoxide, and total unburned hydrocarbons,

respectively, were obtained. The high carbon monoxide EPAP levels recorded with the Vorbix

combnstor were generally due to the contribution at the idle test condition. Attempts to reduce carbon monoxide through pilot zone modifications including airflow redistributions,

fuel spray modification, and volume modifications were ineffective.

The Vorbix combustor appeared to offer the greatest potential for meeting the performance

goals without excessive comprumise of the emission levels. The ltybrid combustor was deficient in both idle stability and altitude relight capability. On the basis of the pollution and

performance results, the Vorbix combustor was selected for engine demonstration testing in

Phase III.

Work was also conducted on two addendums Gothe Phase II program, an Alternate Fuels Addendum and a Combustor Noise Addendum. Details and results of the fuel program are included in NASA Report CR-134970. Results of the combnstor noise program are described

in NASA Report CR-135106.

2
INTRODUCTION

This report describes the low pollution combustor concepts invesdg_ted aml the combnstor

component test results obtained in Phase II of tile NASA/Pratt & Whitney Aircr;fft F,xperimental Clean Comlmstor Program (ECCP). Also described are thc program objectives, plan,

schedule, pollution and performance goals, and current and fL_tureprogram efforts emphasizing

demoflstration tests of a low pollution combustor installed in a full-scale JT9D engine.

The concern with air quality in the vicinity of airports has led to the issuance of emission

standards by the U. S. Environmental Protection Agency for aircraft engines manufactured

after January 1979 (Reference 1). These standards limit the emission of carbon monoxide

(CO), total unburned hydrocarbons (TItC), oxides of nitrogen (NOx), and smoke at altitudes

under 914 m. Recently introduced gas turbine engines, such as the JT9D fancily, already

meet the requirement for producing no visible smoke. Ilowever, compliance with the standards

for the gaseous pollutants will require substantial improvements relative to current engine

emission levels.

The rudiments of pollution control are understood, llowevcr, when incorporating pollution

reduction features, aircraft combustors must also accommodate a diversified range of factors

that greatly add to the development complexity of a practical low-emission eombustor system.

Physical constraints on fuel vaporization, turbulent =nixing rate, di!ution air addition, and

residence time impose absolute limits on the combust on process. Performance requirements

for uniform exit temperature distribution, combustion stability, relight capability, durability,

and operational safety must also be considered. Furthermore, it is desirable to maintain com.

ponent weight, costs, and mechanical complexity at a minimum.

Specific combustor-engine designs have not yet demonstrated the required pollutant reductions without compromising other performance parameters, indicating the need for a,Jditiona)

technology. In response to this need, the National Aeronautics and Space Administration

initiated the l';xperimental Clean Combustor Program in December 1972.

The first two phases, comprising the rig-test portion of the progr,lnl, have been completed.

The results of Phase I were reporte-t in NASA CR-134736. The results of the Phase 11refinenlent and optimization testing are reported in this document.

A summary of the program plan and goals of the Experimcnt:tl Clean Combustor Program is

provided in Chapter 1. Chapter 11contains a description of the reference engine (.1T91)-7)

'4nd combustor used as a basis for the program work, a description of the combuslor design

concepts tested in l'hase 11,and a description of the test and analysis procedures. Chapter

lll In'esents the detailed emissions and performance rest=ItsIbr tile Phase II eombustors.

Chapter IV discusses the It=elcontrol design study in which the control system implications

of two-stage combustor concepts were explored. The relation of tl)e completed work to the

work remaining in I'hase III is discussed as a part of the concluding remarks cent:tined in

Chapter V. l)etailed descriptions of tile Phase 11combustor co'd]gurations, detailed data

tables, and control concept schematics are contained in Al_pendices A through (', respectively.

A table of nomenclature is presented in Appendix D, and the references in Appendix F.

3
CHAPTER I --

EXPERIMENTAL CLEAN COMBUSTOR PROGRAM DESCRIPTION

A. GENERAL DESCRIPTION OF OVERALL PROGRAM

The F+xperimental Clean Combustor Program is a multiyear effort tbat was initiated in December 1972 and is scheduled lot completion in late 1976, 1lie program is directed towards

two primary objectives: ,...

I. The generation of the tecbnology required to develop advanced commercial CTOL

aircraft engines with lower exhaust pollutant emissions than those of current technology engines, and

2, Tbe demonstration of the emission reductions and acceptable performance in a

full-scale engine in 1976.

The program specifically addresses the development of a eombustor with low emission characteristics for the Pratt & Whitney Aircraft JT9D-7 engine, ltowever, the technology developed

during the program will be tr_mslatable to other combustors, either for commercial or military

applications. Tire technology will also provide the foundation for developing further refinemcnts and for klentifying other avenues for continued exploration and experimental research.

B. PROGRAM PLAN

The program is divided into three phases which provide a step-by-step approach for developing the technology required for reducing emissions. These phases are described below:

I. PHASE I PROGRAM

l'base 1 was directed toward identifying promisin_ concepts and _stablishing the design trends

in sufficient detail to provide a Iirm basis for refinement of the more promising concepts in

Phase I1. Three concepts were tested in a 90-degree sector component rig at simulated engine

idle and sea-level take-off operating conditions. "lllese were a Swirl-Crn combustor concept,

a Staged Premix cu::xbustor concept, and a Swirl Vorbix eombustor concept. Thirty-two

conl_gtlratious were evaluated.

The lowest emissions at idle engit_e operating conditions were obtained with the Staged Premix combnstor, The carbon monoxide emission index level was 55 percent below the goal

and the total hydrocarbon elnission index level was 75 percent below the goal. The Swirl

Vorbix combustor approacbed but did not meet the goals, while tbe Swirl-Can combustor

exhibited significantly higher emissions that were close to fire levels produced by current production JT9D-7 combustot._.

At sea-level take-off engine conditions, none.of the combustors met tbe goal for oxides of

nitrogen, .'dtbough some combnstor configurations provided significant imptovcmCldn .:lativc

to tbc cnrrent production J'1"91)-7combustor, The best results were obtained with the Swirl

4
Vorbix aml tile Swirl-t'an combustors, both or whic[_ provided approximately 60-percell|

lower clnissions or nitrogen oxides than the current production JT91)-7 cofnbnstor. All three

¢onlbtlslm" COllCepIs Inel tile snloke goal.

It) addition to the basic l'hase 1 lrrogram, s{.,verlteen combustor eolltlgul'at[ons were evaluated

titsitnulated supe,'sonic cruise conditions, and ,Ilow-emissions conceptual design was evolved.

A combustion noise study was also conducted to document the noise characteristics of the

three low emission combustor concepts and to provide a data base for correlating combllstor

noise levels with performance parameters ....

l)etailed results 6f the basic program including tile studies for advanced supersonic technology

applications are presented in NASA Report CR-134736. Tile results of the noise studies are

l)rcscnted in NASA I(eport CR-134820.

2. PHASE II PROGRAM

The Phase 11prograln involved refirJcnlent and optimization of tile most promisitlg concepts

identified in Phase I. The concepts selected for Phase 11were tile Vorbix combustor and a

l lybrid combustor created by merging the pilot zone of the Staged Premix combustor with a

main burnil_g zone derived from the Swirl-Can combustor. More coml_rehensive sector rig

testing simulating the full range of engine opcrating conditions was conducted to fully document pollutant emission characteristics, to identify previously undetected problem areas, and

to assess combustor pertbrm;mce. After initial testing, tile program was reduced to the Vorbix

combustor concept and the renlairdng test eflbrt was devoted to development of performance

characteristics in preparation for the Phase 111engine demonstration tests. Details of the

l_hase I1 work are cont_tined ill the lbllowing chapters of this report.

Work was also conducted oil two adden,lums to the Phase 11program, im Alternate Fuels Addendum and a Combustion Noise Addendum. I",esults of the Alternate l:uels Addendum arc

contained ill NASA Report CP,-134970. Thc results of file Combustion Noise Addeudtnu

are contained ill NASA Report CR-1351{)6.

3, PI4ASF III PROGRAM

The objectiw' of tile I'hasc Ill i_rogram, currently ill progress, is to substantiate tile pollution

reduction technology develolretl ill I)h;lses I and II ill aP, actual engine environlnent. Ph:lse 111

provides ewduation of engine-quality, :mnular combustor hardware at full cugine pressure

levels ;Is well as assessment or eombustor and I'llei control system l_crformance um!er transient

ol)eratiag collditions or acceleration a!ld decelcratiotl which cannot be simulated ill a conlponent rig. The work will cuhnimde with the testing o1"the Vorbix combustor concept in a

JTgl) engillc d_ring 197().
C. PROGRAM SCHEDULE

The program schedule for tile Expcdlnental Clean Combustor Program is shown in Figure 1.

Phase 1was m_eighteen-month effort wlfich has been completed. Phasc II was a nineteenmoRth effort which has also been completed. Phase 111,currently in p_r_og__sswill , be a sixteen month effort and is scheduled for completion during 1976.

I

PHASE r- COMBUSTOR I_BmBi COMPLETED

SCREENING TESTS _ IN PROCESS

UASIC PROGRAM .... I I I

AST ADDENDUM I

NORSE ADDENDUM 1

PHASE 17 - COMBUSTOR

REFINEMENT AND

OPTIMIZATIONTESTS

BASIC PROGRAM __ 1 I

FUELS ADDENDUM ........ 1 1

NOISE ADDENDUM ......... 1 1

PHASE I]] - ENGINE

DEMONSTRATION TESTS ........... I

" I I I I I I 1 I I I I I

lS73 f974 1B75 1976

Figure I NASA/Pratt & BqlitneyAircraftExperimental CleanCombustorProgramSchedule

D. PROGRAM GOALS

• l'rogram goals were defined for both pollutant emissions and combustor aero-tllennodyn_umc

perfomlance. Thc goals for gaseous pollutaDts rcprcsent the primary program focus. The performance aRd smoke goals were set to rcquire maintenance of current JT91)-7 combustor

_- pertbrmance levels, m_d are imposed to ensure that the reductions in pollutant emissions are

not achieved at the expense of perforlnance or smoke levels. All goals are predieatdd on the

use of commercial 3fade Jct-A aviation turbine fnel.

1, POLLUTION GOALS

The pollutant emissiongoals;ire sunml.',rizedin T;LbleI. The gaseouspollutant emissions

goalsare expressediR tWOways: asemissionindices at the engineoperating mode where the

peak levels of each polltitant arc generated, and as integrated EI)A parameter values. EmissioD

index is defined as the ratio of grams of pollutant formed per kilogram of fuel consumed.

The rl'A parameter, Reference I, is a thrust-normalized mcasnrc of the total mass of poilutart emitted il_a prescribed landing and takc-off cycle. As shown by a comparison of the

goals with the current production JT91)-7 engine, tbe attainRrent of these goals involves

signilicant pollnlant reductions I)y factors of 2Yzto 51⁄2 on anemission index basis.

6

Two ground idle operating points are defined in Table I, corresponding to engine operation

with and withot, t lifteenth-stagc compressor bleed air extraction to meet airframe requiremcnts. Currcn! I'ratt & Whitney Aircraft experimental engine and production acceptance

tests are conducted without extraction of bleed air. IIowever, engines ill the field operate

with varying amounts of bleed extraction. Owing to the particular design of tbe ]T9D fuel

control, compressor bleed extraction results ill a drop in high-pressure compressor rotor speed,

and corresponding changes ill combustor inlet pressure, temperature, and fuel-air ratio. Since

the basis for I:.PAF.mission Standard compliance testing is presently open to interpretation,

both conditions were included in the test program.

2. PERFORMANCE GOALS

The Experimental Clean Combustor Program performance goals arc summarized in Table II.

Tbcsc goals do not represent an appreciablc departure from current JT9D-7 operating levels

except for the pattern factor and the combustion efficiency at idle. The combustor exit

temperature pattern faetor represents a level that is difficult to achievc on a production basis.

Also required for satisfactory performance is the actfievement of a radial average temperature

prolile at the combustor.exit that is substantially equivalent to that produced by the current

production .IT9D-7 combustor.

TABLE I1

ECCI' PERFORMANCE GOALS

Total Pressure Loss (%) 6

Exit Temperature Pattern Factor 0.25 at take-off

Combustor Efficiency (%) 99 or better at all operating conditions

Lean Blowout Fuel-Air Ratio 0.004 +-0.001

Altitude Relight Capability

Altitude (m) 9144

Flight Mach Number 0.5 - 0.8

An additional performance goal is the requirement that the combustor mechanical durability

be consistent with long-term engine operation, equivalent to the current JTgD-7 eombustor.

This goal encompasses structural integrity, liner coolant air level, liner pressure drop, fuel

system metal temperature, etc. In addition, the rear eombustor liner pressure drop must not

be less than 1.8 percent to ensure that the coolant flow requirements for the first turbine

inlet guide vane will be met.
CHAPTER II

EQUIPMENT AND EXPERIMENTAL PROCEDURES

A, REFERENCE ENGINE AND COMBUSTOR

" " 1. REFERENCE ENGINE DESCRIPTION

Tile JT9D-7 engine was selected as a reference for tile NASA/Pratt & Whitney Aircraft Ex- --..

perimental Clean Combustor Program. This engine is tile current production version of the

basic JT9D engine model, which was designed and developed by Pratt & Whitney Aircraft.

Since its introduction into commercial service, this engine has acquired widespread use as the

: powerp3ant f6r both the Boeing 747 and the Douglas DC-I 0-40 aircraft.

The JT9D engine is an advanced, dual-spool, axial-flow turbofan engine designed with a high

overall compression ratio and a high bypass ratio. The medlanical configuration is shown in

Figure 2. The engine consists of five major modules: a fan and low-pressure compressor

module, a high-pressure compressor module, a combustor module, a high-pressure turbine

module, and a low-pressure turbine module. Tile low-pressure spool consists of a single-stage

fan and a three-stage low-pressure compressor driven by a four-stage low-pressure turbine.

The high-pressure spool consists of an eleven-stage high-pressure compressor driven by a twostage high-pressure turbine. Tile accessory gearbox is driven through a towershaft located

between the low- and high-pressure compressors. Selected key specifications for the JT9D-7

- engine are listed in Table I11.

FAN

Figure 2 Cross.SectionalSchematic of the JTgD-7 ReferenceEngine

×, 9
TABLF. 111

KEY SPECIFICATIONS OF THE JT9D-7 ENGINE

Weight (kg) 3982.5

Length (Ill) 3.912

Maximum Diameter, cold (m) 3.427

Pressure Ratio 21.7

Airflow Rate (kg/s) 691

Maximum Sea-Level Static Thrust (kN) 197

Cruise Performance ""

Mach Number 0.85

Altitude (m) 10668

Thrust (kN) 44.6

Specific Fuel Consumption (kg/Ns) 1.979 x 10-5

2. REFERENCE COMBUSTOR DESCRIPTION

The mechanical design of the JT9D reference combustor is shown in Figure 3. The combustor is annular in design with an overall length between the trailing edge of the compressor

exit guide vane and the leading edge of the turbine inlet guide vane of 0.6m. The burning

length between the fuel nozzle face and the turbine inlet guide vane leading edge is 0.45m.

Key performance parameters of the JT9D-7 reference combustor are summarized in Table

IV. Tile reference combustor exit average ,'adial temperature profile is shown in Figure 4.

ENGINE CENTER LINE

FUEL NOZZLE

ASSEMBLY (20) DIFFUSER STRUT

TRAILING EDGE

TONGUE-IN-GROOVE SLIP JOINT

LEADING EDGE

INNER DIFFUSER RAMP

ONE

COMPRESSOR

EXIT

GUIDE VANE

OUTER DIFFUSER

TRIP

TURBINE

INLET

GUIDE

AFT OUTER VANE

SLIP JOINT

Iqgure 3 Cross-SectiottalSchematicof the J79D.7 Reference Combustor

I0
' I I I

"FABLI'?,IV !

KEY OPERATING PARAMETERS OF THE

JT9D-7 REFERENCF, COMBUSTOR

Compressor Exit Axial Mach Number 0.26 i

Compressor Discharge Temperature (K) 767

Combustor Temperature Rise (K) 719

Combustor Section Pressure Loss (%) 5.2 _ ....

Combustor Exit Temperature Pattern Factor 0.42

Average Combustor Exit Temperature (K) 1486

Note: All data for standard sea-level static take-off conditions.

150

0p- _ -5O

--I50

20 40 60 80 100

SPAN_"PERCENTTURBINEHEIGHT

Figure4 Reference Combustor Exit AverageRadialTemperatureProfile

The JT9D-7 combustor incorporates a number of advanced features. The primary diffuser

incoqlorates an inner ramp and outer trip followed by a dump section, and a burner hood is

used to provide a positive pressure feed to the combustor front end. The hood is indented

locally in ten places downstream of each diffnser case strut, A film-cooled louver construction is used for the combustor liners. The liner assembly features inner and outer slipjoints

to facilitate assembly as well as to allow Ibr liner thermal expansion. The fitel system features

direct liquid fuel injection by th,: use of twenty duplex-pressure atomizing fuel nozzles. The

nozzle portion of the fuel injector is enclosed in twenty short-cone swirler modules, which

provide primary zone flame stabilization. Optional take-uff thrust augmentation is provided

by water injection through the fuel nozzle heatshi61ds.

11
r I 1 I 1

3. REFERENCE ENGINE COMBUSTOR POLLUTION LEVELS

Since tile JTgD engi)lc and comhustion system were designed prior to eurPent concerns re-

• garding gaseous poUutants, tile combustor was not specifically intended to provide low

emissions. It does [zreorporate snloke reduction features alld.produces no visible smoke at

any operating condition.

Exhaust emissions are periodically monitored durizzg JT9D production acceptance tests, and

typical results for the idle, 30-percent_ 85-percent, and 100-percent sea-level static thrust on- =lira.

gine power settings are shown in Table V. These power settings correspond to the EPA-speci-

- fled simulatexl ground idle, approach, climb-out, and take-off conditions which are used in the

establishment of aircraft engine emission standards. The data presented in Table V represent average emission levels for a JT9D-7 production engine incorporating combustor configuration EC 289386. This eombustor configuration has been installed in production engines shipped since November 1975. The data have been corrected to standard day temperature and pressure and to an ambient humidity level of 6.3 g H20/kg dry air. Jet-A fuel was

used for the tests. The corresponding values of EPA Parameter (EPAP) are also presented in

Table V. This parameter combines emission rates at the engine idle, approach, climb, and

[ take-off operating modes, integrated over a speci fled landing-take-off cycle (Reference 1).

TABLE V

t • REPIZ.I'.'SENTATIVE JTgl)-7 PROI)U(_I'ION ENGINI! EbIISSION LEVELS

AND EXPI_RIMENTAL CLEAN COMBUSTOR PROGRAM GOAI.S

_ A. Emilio. Indices

Total

Carbon Monoxide Unburned llydrocarbons Oxides of Nitrogen Smoke

; Emissions fsJkg fuel) Emissions (g]xg fuel) }_missions (g/kg fuel) (SAE Smoke Ntmzber)

L Operating ECCP ECCP _'_C(_ T

Conditloll JTgD-7 Goal JT9D-7 Goal JTgD-7 Goa_.L .I]_9D-7 Goal

Ground Idle 46.7 20 21.6 4 3.9 ......

, " Approach (30% Power) 3,4 0.6 7.8 ..........

tlimbout 185% Power) 0.5 -- - 0.4 - • 28.8 .......

_¢a Love) "rake-off 0.4 0.4 _9.9 [0 4 15

( 100% Power)

B. I'EPar=nlete A r (Ibm. poltutanI/1000 Ibf. thrust - hr.]landhlg-lake-off cycle)

Total

Carbon blonoxi(I.,, Unburned Ilydrocarbons Oxides of Nitrogen

... E(_ I!C('P fS(_'P

J'l'9 I)-7 Goal J'I'gD-7 (k_al J1"91)-7 (;oal

8.5 4,3 3.9 0.8 5.9 3.0

Notes; Oxides of nit rogen data presellled as ililrol_.,n dioxide gqtliValeltL

. . Data [epr=.,_ltI average emIssioB levels for a J 191)-7 ptoductiotl engine iltcorptlralillg COllll)uslor c()n tqgtlration I{("28937_6.

GrourRI idle dala is without ¢olnpres,a)r air bleed.

12
B. TEST COMBUSTOR CONCEPTSAND CONFIGURATIONS

1. HYBRID COMBUSTOR

a. GeneralDescription

The initial ltybrid combustor Col:figuration HI is shown in Figures 5, 6, and 7. This cornbuster concept incorporates two burning zones that are separated in both tile axial anti radial

dh'ections. The design basis for each zone evolved from two combustor concepts that were

developed during the Phase 1 program. Tile pilot zone duplicates the physical geometry of ,,.

tile Staged Premix eombustor Configuration P3 pilot which exceeded the low power emission

goals in Phase I. The mail1 zone was based on the Swirl-Can eombustor Configuration N9,

which demoustrated significant reductions in high power emissions of oxides of nitrogen.

• ii i --

¢..

PLLOT PRE-MIXING pASSAGE

PRES$UR6 ATOMIZf_G_ / Y

FUEL NOZZLE _ALL ®

d J

4

_ _ " MAIN OUTER - *

?" . , ,. _.T;.__.._.--FUELSPLASI-IPLATE '_"_

{ I " "' _ 5,_W'--

L. _ -,_ . / _ _""_'MA,_,NNERSWfNg.(IIt

M( N LOW PRESSURE _ _s i

DROP FUEL INJECTOR

Fisure5 Cross-SectionalSide View of Hybrid Combustor

The pilot zone was designed lbr an equivalence ratio of 0,8 at idle conditions to p,ovide low

levels of carbon monoxide and unburned hydrocarbon emissions as well as gee," stability during lighting. Fuel is introduced into the premix passage through forty pressure-atomizing

fncl injectors (on a full annular basis), where it mixes with air and is partially vaporized. The

combustible mixture then passes through boles in the flameholder and into the combustion

zone. At high power conditions, the pilot zone equivalence ratio is reduced to reduce lbrmation of oxides of nitrogen.

The main zone was designed to provide lean conlbnstion followed by quick quctlching with

dilution air to reduce emissions of oxides or nitrogen at high power conditious. As shown

in Figure 7, the zone consists of an arrangement of concentric conntcr-rotati_)g swirlers located downstream and displaced radially fi'om the pilot zone. Fuel is introduced through low

13
pressure drop injectors located at tile centcr of tile ,,;,wirlers. The fuel impinges on a splash

plate located in the exit plane of the swirlcrs, where il is broken into small droplels and then

entrained in the swirling flow.

i,

I

Figure 6 Exterior View of llybrid Cot_bustor (XPN-52117)

%

l,'igacre7 Front-I:)td View of llybrld Combustof (XI'N-33775)

__-_ 14
i i

During Phase 1, aspiration ploblen s were encountered with lhc .',;wirl-Canmoduhu" comlms-

- tor as a result of poor inlet air llow distribution. Consequcndy, a transparent plastic flow

model of the Hybrid combustor was fabricated and tested at cold flow conditions prior to

finalizing the Phase II design. The tests revealed nommiform airflow distribution resultlug from aerodynamic blockages caused by radial flows issuing from the strut cutouts in the

burner hood. To correct this prcd_lern,the Phase 11hybrid design incoq_oratcd the following changes:

i l 1. The burner hood strut cutouts were extended and contoured to provide a smooth

transition to the inner and outer combustor liners, and -'"

2. The outer diffuser case wall was recontoured to reduce the diffuser area ratio

approaching the main zone air inlet.

The effectiveness of the final design in improving airflow distribution is shown in Figure 8.

fI .... o.soo __--2.ON: ,RuN } *_ER. OF ._E_..E_C,.CO.E_R_N_,A_ TRAVER.ES

[.

RECONTOUREO

_" 0.985 /¢_ STRUT RELIEF

/ k /

o.o.o, / X / X /%

' °.o,o, \

SIDE FUEL f:UEI DIFFUSER FUEL FUEL SIDE

WALL INJECTOR INJECTOR STRU1 INJECTOR INJECTOI_ WALL

o.s.oII-1 El I-I • I-I I

ClR,"UMFERENTIALLOCATION

lqgure 8 PressureLoss ProfilesShowing Improvement at Main IhtmhtgZmw I:)_tram'e/'laneObtained

- [or PhaseHHybrid Combustor

15
I

 ̧l

b. TestConfiguratinns

Optimun_ airflow fractious for the Prenlix pilot zone and Swirl-Can main zone were detorlnincd during tile Phase I lost program. Ilowcver, tile quantity of air available after liner

cooling requirements had been satisfied was insufl'icient to permit both pilot and main

zones to be operated at their respective optilnnm equivalence ratios. Tllerefore, the primary

focus of the configuration modifications was to shift the pilot and main zone equivalence

ratio to lninilllizc pollutant formation. Other variables !nvestigatcd were the main zone

swirler arrangement and tile pilot zone fuel injector type. ,...

The configurations tested are summarized in Tables Vl and Vii. Specific types of configuration change are discussed below. Detailed configuration descriptions am presented in

Apep._dix A.

Pilot Zone EquivalenoeRatio

The effect of two different pilot zone equivalence ratios on idle emissions was evaluated.

The initial configurations, H1 through I14, wer,', tested with a pilot zone equivaleuce ratio

of O.8, In Configuratien 1t5, the amount of premix passage airflow was reduced to increase

the equiwdence ratio to 1.2. In the final two configurations, tl6 and H7, the pilot zone

equivalence ratio was retnrned to 0.8.

rABLE VI

EIYBRII) C(._MI|US I'OR ('ONIrlC;IJRA ['[ON SUMMARY

Pihfl Zone Main Ztme

I!quivalett¢¢ Fu¢l l:uel I)ihllion Air

Ratio at hll¢ =Ntutll_.*etl htj¢,:tul" _v,[tlet brumb_r tff Inj¢¢lor Iniler OUlt'r

Configuration '¢¢ilh diced Fu,:l h_cctot_ q_pe* Arrangement Fuel Injvclors "lype ' liner Liner

In 0,8 I0 . PA-II( q_vo (_nceat tic I1 I.PI) Yes I1o

(_!lltl¢ r-I_l!lali.lg

Swhle_

112 0,8 It) PA-II(" 'l_{_totatiag II I.PD Yes N_

Sw[rle_s

113 0.8 IO PA-II(" Out cr Swider BIc_cked II IPD *l'e_ Yes

,'lnd Inner Area Rt'dw:¢d

2if;

114 0,8 IO I'A-S(" _)l_tvr S'_irlcl Illockcd I I I PD Yes Yt's

and Inuvr '_a Reduced

2(;':

1t5 1.2 I t) PA-II(; Satnc .L_I11 Iht) Swirlcr I I I PI) "¢cs No

Alva [llvlvav:tL .IG':

116 (}.8 It) PA-II(" Sa,w a_ Ill lhlt Swirlct I I I.PD No N_

Alt_.l Increased 4(¥:

117 t).8 I(} PA-S(" SalUV a_ III But 5wilier I I LPI} No Nt_

Aica hLerca_d ,It)':

*Fuel I@:¢tor I ypc_. PA IIrt'_surc Att3mi/iag

II(" Iht)h,w t'c_ne

S(" _llid { _lne

IPD I iiw Pftr_ure DrLip

16 ORIGINAL PAGIO IS

()F I'OI)R QUALITY
000 000 0

•g#l ,_8 8r.,8 8 '_

I_,_I .#,#_ _,-4,,: ,# ...o

•_ ___ _ _ i

.[

]7
Main Zone EquivalenceRatio

Main zone modifications were made to evahBl[e tile effects of both equivalence ratk_ and

airflow distribution on cnlission levels, Ill the initM design, tile inner and outer main zone

swirlcrs operated witll approxinlately equal flow rates that provided a bulk equivalencc ratio

_f 0.6, with dilu tioll air illtroduccd only through tile inller liner. Ill Configurations 1t3 and

I|4, tile outer swirler was blocked, and tile flow was diverted to tilt otlter liner. In addition,

the flow through the inner liner was inclx:ased. This inere&sed tile bulk main zone cquivalence ratio to 1.6. In Coltfiguration 115, tile dilutiol_ tlow through tile outer liner was

blocked and the flow was returned to the outer swirler, and the dilutiol3 flow through the '-'"

inner liner was retnnled to approxhnatcly its l'omler level. Tiffs configuration included a

richer pilot zone, and the airflow removed from the pilot zone was divertc4 to the main

burner outer swirlers.

Main Zone Swirler Arran_iement

Tile swirlers for tile main burning zone consisted of two concentric swirlers around each

fuel injector. In the initial configuration, these concentric swirlers produced counterrotating flows. Tests were also conducted with both tile inner and outer swirlers producing

co-rotating flow, and with the outer swirler blocked. The co-rotating swirl flow was tested

in Configuration 112, and the outer swirler was blocked ill Configurations It3 and H4. For

these two configuratiol_s, the inner swirler flow area was reduced approximately 20 percent.

For the remaining configurations, Configurations 1t5 through 117, the initM counter-rotating

swirler arrangement was used with the outer swirler area increased approximately/40 percent.

Pilot Zone Fuel injector Tyoe

Two pilot zone fuel nozzle types were tested to determine the effect of changes in fuel distribution. Both were pressure atomizing nozzles. One provided a hollow spray cone while

the other provided a solid spray cone.

Pilot Zone FfameholderAttachment and Cooling

Modifications were incorporated thronghout the program to improve flamellolder durability. Ill tile iuiti'll design, tile ffalnehokler was welded directly to tile inner eombustor

liner. This arrangement was found to restrict thermal growth oFthe flatnchokler, resulting

in defornlation of the flalnehokler and premix passage, and inadequate flameholdcr coolant

flow. The flanlcbohlcr.lttachment was modified il_('onfiguratk)n 112 to provide a rigid

attaclunent at (lilly one sklewall. Retaining labs were used to axially position the tlalnebolder along the outer liner, the FINWALL ® ring betwcen the pilot and nlain zones, and

the opposite sidewall, ht f'onfiguration 113, cooling holes were added in the hailing cdgc

of the flanlcholder, and in Configtmltion 116, cooling of the outer porlion of the trailing

edge of tile flamelmlder was increased.

18
Pilot-to-MainZone Fuel Split

qlle fuel flow split between tile pilot ,rod main zones at operatiqg conditions above idle was

investigated for ;dl configurations. By varying the fuel split during the tests, the relative

equiv:dence ratios between zones could be optimized for 'all emissions. Ill early COlffigurations, the results were also used to indicate where adjustment of the equivalence ratio produccd by eli:rages in the airflow distribution would be beneficial. Tests were conducted at

the approach condition with fuel supplied only to the pilot zone, only to the main zone, or

with various fuel-flow splits witll fuel supplied to both zones. _'

2. VORBIX COMBUSTOR

a. GeneralDescription

A cross-section schematic drawing of the initial Phase II Vorbix combustor design is shown

in Figure 9, and representative photographs are shown in Figure 10 and 11. The design was

evolved from the Phase 1Vorbix combustor and incorporates two burning zones separated

axi_dly by a high velocity throat section. The pilot.zone is a conventional swirl-stabilized,

direct-injection combustor employing tlfirty fuel injectors (on a full annular basis). It

was sized to provide tire required heat release rate for idle operation at high efficiency.

Enfissions of carbon monoxide and unburned hydrocarbons are minimized at idle operating

conditions primarily by maintaining a sufficiently high pilot zone equivalence ratio to 'allow

complete burning of the fuel.

PILOT ZONE

NOZZLE MAIN ZONE

! ,d _ THROAT

• ' ,NJEcro.s ,! I

Iqgure 9 Swirl Vorbkv Combustor Design

19
b_gure10 l:,_teriorView of Vorb£t"CombustorConfiguration$11 (XPN-51834)

• _ de"

Hgure 11 I,'ront-I:hd l'_ewof VorbLx"Comb tstor ConfigurationSI I (XPN-51835)

20
i i

At take-off conditions, the pilot exhaust equivalence ratio is reduced as low as 0.3 (including pilot dilution air) to minimize formation of oxides of nitrogen. The minimum equivalence ratio for the pilot zone is determined by the overall lean blowout limits, combustion

efficiency, and the need to maintain sufficient pilot zone temperature to vaporize and ignite

the main zone fuel.

Main zone fuel is introduced through fuel injectors located at the outer wall of the liner

downstream of the pilot zone discharge location. Either thirty or sixty fuel injectors were

used (on a full annular basis). Main zone combustion and dilution air is introduced through ,..,..

sixty swiders positioned on each side of the combustor (on a full annular basis).

'll_e Phase II Vorbix design differed from the Phase 1 design in that the length of the pilot zone

was increased by 3.8 cm and the main zone length was reduced by 3.8 era. The increase in

pilot zone length was made to reduce residual carbon monoxide and unburned hydrocarbons

produced at idle conditions. The corresponding decrease in main zone length was expected

to reduce the main zone residence time, thereby reducing the formation of oxides of nitro-

_11.

b. TestConfigurations

Initial testing of the Vorbix combustor was concentrated on demonstrating further reductions in emissions. Modifications were made to the pilot zone to reduce emissions of carbon

monoxide and total unburned hydrocarbons at idle conditions. Changes were made to the

main zoue to reduce emissions of oxides of nitrogen at high power conditions and also to

identify those design features that would affect the radial temperature profile and pattern

factor. Later testing was directed toward correcting performance deficiencies and maintaining or improving the emission characteristics. The effect of pilot-to-main zone fuel split

was investigated for all configurations.

"111econfigurations tested are summarized in Tables VIII and IX. Complete specifications

are presented in Appendix A. In the pilot zone, changes were made to burning volume,

swirler/airflow schedule, and fuel injector type. In the main zone, modifications were made

to the swirler orientation, fuel source density, fuel injection technique, and airflow schedule.

Details of these modifications are discussed in the following sections.

Pilot Zone Volume

The pilot zone volume was increased approximately 30 percent in Configuration SIS in an

attempt to improve idle combustion efficiency. The revised pilot zone geometry is shown

in Figure 12. The volume increase _,as specified based on results presented in Reference 2,

which indicate that combustion efficiency is proportional to available combustion vohnne,

In Configuration S19 through $21, the e:_larged pilot volume was evaluated without a hood

attachment. For the remaining configurations, a hood derived from the current JT9D-7

hood was installed to reduce diffnser section pressure loss and to improve airflow feed to

the pilot zone.

21
i i

22

Ol{l(_]N&h PAGI_

()i_' POOR qUALITY.

24
Pilot Zone Swirler/Airflow Schedule

Various pilot swirler arrangements and airflow distributions were evaluated to assess the

effect on combustor emission and performance characteristics. Tile pilot zone utilized for

the first five configurations was transferred directly from the final Phase I configuration,

with the exception of increased length. Tile swirlers admitted 13.3 percent of the total

combustor airflow and no liner dilution airflow was used. The swirler airflow was metered

by swirier open area in conjunction with an air cap in the fonn of a toroidal deflector

located downstream of the swifter vanes, The air cap, visible in Figure 13, directs the swirl- ..,.

ing airflow towards the swirler fuel nozzle centefline and augments file tangential velocity

component.

BLOCKAGE

RING

AIR CAP

(HATCHED PORTION)

PRESSURE ATOMIZING LOW PRESSURE DROP

FUEL INJECTOR AERATING FUEL INJECTOR

Figltre13 Swirl Vorblx CombustorPilot BurnerFront-End Designswith PressureAtomizing Fuel

b_]ectorandLow PressureDropAerating Fuel Injector

For Configurations S16 through SI9, blockage tings were added to the pilot swirlers to reduce airflow to the range of 2 to 5 percent of the total combustor flow, This produced a

pilot equivalence ratio of approximately 2. Configurations S18 and S19 included approximately 6 percent pilot zone dilution air.

Pilot zone equivalence ratio was reduced for Configuration $20 by increasing the swirl airflow to approximately 13 percent. This was accomplished by removing the air cap, increasing the swifter size, and increasing the dilution airflow by approximately 0,7 percent.

In Configuration $21, the swirler flow was increased to 15.7 percent by the addition of a

concentric swirler and a revised air cap, In Configuration $22, the dual concentric swirlers

were replaced by a single JT9D-70 type swirler with approximately the same flow rate as

the concentric swirlers. This swifter was utilized for Configurations $22 through $25, with

adjustment of swirler airflow accomplished by changing the metering area of the toroidal

deflector air cap,

25
,/

Pilot Zone_Fuelinjector Type

Two fuel injector types were investigated. Most testing was conducted with pressure atomizing fuel injectors which depend on the fuel pressure to atomize the fuel. Two tests were

conducted with low pressure drop aerating fuel injectors which utilize the pressure drop

within the combustor for fuel atomizing. The two fi=elnozzle installations are shown in

Figure 13.

Main Zone Swirler Orientation ....

Main zone swirlers are located between main fuel injectors to promote mixing of tile fuel and

air. In the Phase I program, testing revealed that the combustor exit radial temperature pattern

was influenced by the orientation of the swirlers, that is, whether adjacent swirlers provided

swirl in the same direction or in opposite directions. In Phase II, the effect of swirl direction

was also investigated to determine if the Phase I results were applicable to tile larger swirlers

used in Phase II and also to determine the effects of swirl direction on combustor performance.

The swirler arrangements evaluated in Phase II are shown in Figure 14.

MAIN ZONE INJECTOR IN LINE

WITH PILOT NOZZLES (13)

i _ CONFIGURATIONS $11,S16

- MAIN ZONE INJECTOR IN LINE

WITH PILOT NOZZLES

CONFIGURATIONS S12 TO SlS "_1 IIN CONFIGURATIONS

$12 13 INJECTORS / /s S12 AND SlS

SlS 7 INJECTOPS IIt_l _ THESE NOZZLES

S14 7 INJECTORS __1 t ALSO FLOW FUEL

..- Sl. 13,NJEcroas

Fi_tre 14 Swirl Vorblx Combustor MahtZone Swirler OHentatlons/';valuatedin Phase11

2

} 26
'I I

i t

Main Zone FuelSourceDensityand Fuel Injection Technique

Three main zone fuel system variables were investigated as part of the Phase II program. The

first of these was fuel source density. Tests were conducted with seven and with tldrtcen

main zone fuel injectors (corresponding to thirty or sixty fuel injectors on a full annular basis).

Tile tests were directed toward confirming exit temperature distribution trends identified in

Phase 1and determining the effect of reducing the number of injectors on emissions.

Tile second variation was fuel injector pressure drop. Tests were conducted with two different ,..

pressure drop levels, utlhzing both seven and thirteen injectors at each pressure drop level, to

assess tile effect on emissions and radial temperature profile.

The third variation was the use of premixing tubes for each of tile main zone fuel injectors, as

shown in Figure 15. Tile premixing tubes are designed to provide a rich, partially vaporized

fuel-air mixture with an equivalence ratio of three. This design was based on results obtained

from other Pratt & Whitney Aircraft programs and was expected to provide reduced emissions

of oxides of nitrogen at high power levels with additional benefits related to reduction in the

fuel nozzle coking tendency and simpler liner removal.

__ _ q.

I'fl" J ,c_ : " " i

';i _ -1'I' _\, _ _

1UIBES t3 LOCATIONS ii

/,igure15 5w'rl Vorbix Combustor Design It_'thPtemlx#lg 7kbes

27
) i

[VlainZone Airflow Schedule

Most of the changes in main zone airflow schedule were made to adjust the exit temperature

.... profile and pattern factor. Tile modifications inc(udcd changes in air distribution to the

swirlers and changes to the dilution flow. The dilution hole patten_ was keyed to symmetrical

"_ combustor features, such as stn_ts and fuel nozzle locations, since tile air distribution developed

- in l'hasc ll would be transferred to full annular hardware for£hase 111.

Initially, the flows through tile inner and outer liner swirlers were approximately equal, with

a combined flow of 54.4 percent of the total combustor airflow. Begimling with Configuration

$20, the total main zone swirl flow was reduced to approximately 45 percent and file flow

split between the inner and outer swirler rows was progressively biased to provide decreasing

anlounts of flow through the inner swirlers. In Configuration $25, tile inner swirler flow was

11.5 percent and the outer swirler flow was 25 percent. Configuration $21 employed carbureted fuel injection with 7.8 percent of the flow passing through the carburetor or premixing

tube passage. Most of this flow was subtracted from the swider flow, reducing the combined

swirler flow to 38.8 percent.

No dilution flow was initially provided in the main zone. Beghming with Configuration S16

: dilution flow was introduced through the main zone inner and outer liners with approximately

equal amounts through each liner. This flow split was progressively modified to tailor the

radial exit temperature pattern until, in Configuration $23, all of the main zone flow not

. passing through tile swirle,s was introduced through the outer liner, with a total dilution flow

of 11.8 percent. In Configurations $24 and $25, tile circumferential distribution of this flow

:: was varied to tailor tile circumferential temperature pattern.

C, TEST FACILITIES AND EQUIPMENT

.... 1. TEST FACILITIES

' "llle combustor tests were conducted in two test facilities. All of the emissions and performance evahtations except for tile altitude relight tests were conducted in a high-pressure test

facility, Stand X-903, located at Pratt & Whitney Aircraft's Middletown, Connecticut, plant.

The altitude relight testing was conducted in an altitude test facility, Stand X-306, located at

the Rentschler Airport Laboratory in East llartford, Connecticut. Tile capabilities of these

facilities and the procedures used have been fully described in the Phase I final meport, NASA

CR-134736.

Two significant improvements were made to the high-pressure test facility. An automatic

traversing rake system was installed to replace tile fixed rakes for gas sampling, and a chemilumemlscnce analyzer for oxides of nitrogen was added to tile exha_st ga_ iuatiumcutation.

The Ph:tse t traversing rake was modified to inchale provisions for automated sequencing, incorporation of .'lremovable thermocouple rake I|ead, and thermocouple air cooling. The

automated sequencing system was specifically designed for compatibility with the gas :malysis

instrtuncntation response time to ensure accnrate sanlpling. The rake llead, which is shown

in Iqgtlrcs 16 and 17, was split in(o two sections to permit rapid replacement of either head

28

Ssection. In addition, a platinunt-rhodium air supply manifold was incon_ora':ed to provide air

cooling of the thernlocouples when temperature data were not being takan. This system was

- effective in protecting the-thcnnocoul_les when local temperatures exceeded the lil;lits of

tile thermoeouples,

_: Tile chemilumensence analyzer became the primary measuring system for the Phase 11program. The Phase I nondispersive infrared and nondispersive ultraviolet systems were retained

: as backup systems and for measuring concentrations of NO and NO2, respectively.

ab,,

WITH COOLING AiR _

COOLING AIR

_.K._ x__ __i ,t _-_

HTO__ --_ ;: _TsEALML OLYEXTTU/P

,L?.'E.... T _.

" COOLING TEMPERATURE PROBE TIP r ------- STEAM

AIR) _IAMPLE_ _ -- -_1

S1-EAM COOLED SUPPORT GAS SAMPLE OR PRESSURE PROBE TiP . • .

PROTECTING

PLATINUM-RHODIUM THE COOLING AIR SUPPLy

/

STAINLESS STEEL

THERMOCOUPLE

I,Igttre16 Detailed Viewo[ Cot buster Exit Ai *Mar'l}'mers#g Rake

-" 29

lib, m ____ /
II 3O . • il
t I t

2. TEST RIGS

A schematic diagram of a test rig and adapting hardware installed in tile test facility is shown

in Figure 18. "l_herigs used for Iqlase 11were essentially tile same as_those used in Phase I.

•0 . .

t;, eh g co _ssled of a 90-degree sector simulating a full annular.s.:onlbustor.

• _" ', ll(iq_ ." 90o

I _ • ." _./DVAIIIEPIIOII[

S[(:|(Ofl A A SI:CI IO_t _.EI _1 CHe)I_C_: O TOI_AI.FR[SSU_E X STATIC Pill: SStlR E _'I

PllE$SURf pROBf S 1 [I.IPI lit. 1t3_tEPllOIIE$

INS1flIjhlEN i ATION pt/_NE S • TOTAL 1 EMP[IIA lull f

OI_XIST1_4($ . _ i la I, I, • :_

EXPhNSION SEClrION ! [ I_ IRAVI:RSE IIAKE

" ;': / t._LE. EXlS'ING PI'E SSUR£ COM_IttS_O" _ I'I' _" ._'_N "

.. : , _t-___._L.. _ -u ,..: :. ,^KE _,

: : i .. , _ . ,

. _ I I ._¢_EOL O_ COMBt/STOH

i KII_C_A.NISM " $tCI"ION O O

I : i ,

i "' ! "j=

". I , , !

I I;i :

Figure 18 SchematicDiagramo[ Test Rig Installedin the TestFacility

The combustor cases for tile test rigs duplicated the JT9D-7 engine diffuser and combustor

case design as well as the diffuser strut orientation. The combustor cases for the Vorbix

combustor were the same as those used in the Phase I program while tile combustor cases for

the liybrid combustor were provided by modifying the Phase I premix combustor eases.

The modifications consisted of changing the main fuel nozzle bosses to accommodate the

eleven main nozzles for the hybrid combustor and capping the ten premix main nozzle bosses.

The test rigs differed in two respects from those used in Phase I. First, tile rear outer portion

of each ,:ase was sectioned and then joined with a mechanical attachment. As shown in

Figure 19, this modification improved the accessibility to tile combustor for visual and borescope inspection, and also permitted rapid modifications to the rear outer liner hole patterns

without requiring removal of the rig from the test stand.

The second modification was the incorporation of positive sealing at the rear liner seal. In

Phase 1,sliding finger seals had been used at the rear of the liners, and leakage had occurred

as a result of nommiform thermal growth. For l'hase II, the rear seals were fabricated as an

integral part of the rear liners, and the assembly was then bolted to the rig backplate. Radial

and axial thermal growth was accommodated by deflection of the seal. A typical re:_rseal

arrangement is shown in Figure 20.

31

• I
The fi_cl _,ystem for Ih,.' Vorl_ix c(HIIhtl.Sh)r ill,.'orpt)rated up Io |Jltut_ sL'pat'ah' lu:mil'(_lds, ()lie

for the i-,ilol /.t_ll¢ [llld Iwo for the II|[lill ZI)II_2, The two i_lain ,<(_11¢lllalli]'oJdg permitted staging

of L'J[hL']SL'_,SCV¢II,0:":1]1Ihirlcen ['tlCJJll.leCtOl'S.

The fuel system for Ihe Ilyl_ritl COlll[_LlSt.orWas similar to tl_at for the Vorbix colnbLl_|Ol', with

011_.']dCI ll1_lllifoJd ['()I"|II¢ pilot 7,(.)11_I¢II_,] !.WO for th¢ 111ai11;_on¢. "Hie lll|lill 7.011¢nlallifo](Js

WCl'__.!ll'r_in_,.'d,_.)l'_¢l*nlit._t_|_,Jl1,oJ" _ ]'OLII'_$¢vcn, or _iIJL'Jg*,,ell|'ii,2] llOZZJ,2s.

Iq,_,ltr(,19 ('.mlmst,,r Rig ()ttter ('else ,_'howillgS(,ctimliltg to h pro "/(" "¢s.dhilil.v ._'1W-52120)

32
]

Figure20 VorbLxCombustor DesignShewb_gRear Outer SealArrangement

D. TEST coNr31TIONS

1, HIGI't-PRFSSURE TEST FACILITY

The combustor rig test conditions were chosen to simulate the desigzl table estimates of the

JT9D-7 engine. I)roduction engine performance data has verified that the design table estimates are representative of actual engine comlitions. The data were taken during sea-level

static production engine acceptance tests and included operation at idle, approach, climb,

_md take-off condilions for on tminstalled engine.

Rig test conditions arc shown in Table X together with the corresponding production engine

conditions. As shown, the engine idle conditions were exactly duplicated in tile rig. At higher

power levels, the rig duplicated all engine parameters except for tile inlet pressure and airflow.

'rhe inlet pressure was limited by the test facility capabilities. The airHow was reduced accordingly to provide the engine levels of combustor reference velocity.

33
I t I i

I

34 OItI(_INAI5PA(_ IS

Ol,'PO01_QUALITY
Tile test figs incorporated shroud bleeds to simulate bleed flow to meet turbine cooling requirements. 311e bleeds used in the rig duplicated those required in the engine, thereby

providing representative shroud airflow characteristics and combustor pressure loss.levels in

the rig.

In addition.to tile tests conducted at simulated engine conditions, tests were conducted with

parametric variations of overall fuel-air ratio at a!l simulated engine operating conditions.

Parametric variations of inlet temperature aud reference velocity were also made for selected

comimstor configurations. ,,.

The test program included main burner fuel staging tests for all configurations at simulated

sea-level take-off conditions. In addition, circumferential staging was evaluated for selected

configurations _t the approach and take-off test conditions.

2. ALTITUDE TEST FACILITY

Altitude stability and windmilling starting tests were conducted at simulated JT9D-7 engine

windmilling conditions. Actual engine combustor inlet and pressure conditions were simulated

while fuel flow and airflow levels were scaled for the one-quarter sector rig. The range of

conditions that were simulated are shown in Figure 21 together with a typical JT9D engine

relight envelope.

The windmilling starting tests conducted in tile combustor rig were made at stabilized conditions wlfile the engine is required to relight immediately after a blowout with the eombustor

trod fnel system still warm. As a result, the limits demonstrated in the rig are conservative

relative to the relight capability at actual eugine conditions following blowout.

Ti4 747K _K /lilK _K

9O

ii: /

FLIQH'_ MACIt NUMtlE _1

Iqgure21 JI_gDRelight I':nwlope ,

35
E. DATA ANALYSTS PROCEDURES

1, PERFORMANCE DATA

Measured and calculated combustor performance parameters are listed in Table X1 and defined below.

TABLE XI

re=l,,.

SUMMARY OF REPORTED COMBUSTOR PERFORMANCE PARAMETERS

Parameter Symbol Units Measured Calculated

Total Airflow Wa4 kg/s X. --

Total Combustor Airflow Wab kg/s X

Pilot Fuel Flow Wf pri kg/s X

Main Fuel Flow Wf see kg/s X. _

Total Fuel Flow Wf tot kg/s X

Fuel Temperature Tfuel K X

Inlet Total Temperature Tt4 K X

Inlet Total Pressure Pt4 arm X

Reference Velocity Vre f m/s X

Pattern Factor PF -- X

Inlet Air Humidity H gH20/kg air X

Fuel-Air Ratio f/a -- X

Total CombustorAirflow

"Nit total combustor airflow is calculated by subtracting the measured inner and outer turbine

cooling air bleed flows and the estimated combustor liner sidewall cooling airflow from the

total airflow.

ReferenceVelocity

The rclbrence velocity is defined as that flow velocity that would result if the total combustar airflow, at the compressor discharge temperature and static pressure, were _assed through

the combustor liner at the maximum cross-sectional area. This area is 0.098 m" for the ttybrid combustor sector rig and 0.056 m2 for the Vorbix combustor sector rig.

Pattern Factor

The pattern factor at the combustor exit is defined by the expression:

Tt5 max Tt5 avg.

I)attcrn Factor =

Tt5 avg.- Tt4

36
whore:

Tt5 max = Ilighest local temperature observed at the combustor exit plane

Tt5 avg = Average combustor exit_emperature (calculated from the

carbou-balance fuel-air ratio and the corresponding combustor

. inlet temperature and pressure) _

• Tt4 = Combustor inlet temperature ,,.,.

Fuel-Air Ratio

The fuel-air ratio is the ratio of fuel flow to total co:nbustor airflow. Fuel-air ratio was

calculated from measured values of total fuel flow and airflow. The independent fuel-air

ratios for the pilot and main zones were determined by dividing the total fuel-air ratio in

proportion to the measured fuel flow rates to each of the burners. Hence, the sum of the

" pilot and main zone fuel-air ratios equals the total fuel-air ratio.

" 2. EMISSION DATA

a. Fuel-Air Ratio Calculations

.-_ Fuel-air ratios are reported on tim basis of measured fuel and air flows (performance basis).

This differed from Phase I where the carbon balance technique was used as the basis for the

reported data. The change was made possible by the addition of the automated traverse rake

and by the use of data obtained between the diffuser strut projections where average values

can be obtained free from sidewall effects. In analyzing the data, emission indiczs were calculated using the local carbo,l-balance fuel-air ratios, and correlations were tllen made using

overall average fuel-air ratios calculated on the performance basis.

b. Combustion Efficiency

The combustion efficiency was calculated on a deficit basis using the measured concentrations

of carbon monoxide and total unburned hydrocarbons from the gas sample data. The calculation was based on the assumption that the total concentration of unburned hydrocarbons

could be assigned the beating value of methane (CH4). The equation was

-: .c=,00-,00L 18.4.oJ 6

" where:

x = Measured carbon monoxide concentration in g/kg fuel

y = Measured total unburned hydrocarbon concentration in g/kg fuel

37

I

L
c. Extrapolationof Rig Data to EngineConditions

Since tile conthustor sector ng was unable to simulate the combustor inlet pressure at conditions above idle, the emissions data for oxides of nitrogen, carbon monoxide, and total

unburned hydrocarbons obtained at the rig test conditions required correction to the engine conditions to permit comparison of the results with both the Experimental Clean Combustor Program goals and with the current JT9D-7 emission levels. The correlations used

are described in the following paragraphs.

Correlaticn for Oxides.of_Nitrogen

Oxides of nitrogen are reported as equivalent NO 2.

The correlation used to scale oxides of nitrogen values to engine pressure levels and to correct

the values for small differences between the actual rig conditions and the desired engine conditions is as follows:

0.5 )

/ _/1 t4 corr._ /Vref. meas._ (Tt 5 corr.

NOx corr. = tNOx mea_ltPt-'_"_'eme_) _e f, c-'_---)_.t"_tt5 m--_e '

where:

NOx = Emission index of oxides of nitrogen.

Pt4 = Inlet total pressure (atm)

I14 = Inlet total temperature (K)

Vref. = Reference velocity (m/s)

tt = Inlet specific humidity (g 1t20/kg air)

Tt5 = Combustor exit temperature

luld subscripts:

corr. = Relates to value at corrected condition

meas. = Relates to value at nleasured condition

38

i , I ii , i I " -
in this program, this correlation was used to extrapolate tile experimental data to tile conditions shown in Table X. It could also be used to extrapolate the experimental data to other

conditions of inlet pressure, temperature, fuel-air ratio, reference velocity, and humidity.

However, the correlation factors for inlet temperature and fuel-air ratio are sensitive to the

specific combustor design features, particularly to the equivalence ratio in the burning zone.

Consequently, use of this correlation was restricted to relatively small adjustments.

Correlationsfor.Carbon Monoxide andTotal Unburned.Hydrocarbons

Total unburued hydrocarbons arereported asequivalentCH4.

Emissionindices for carbon monoxide and total unburned hydrocarbons were scaledto engine

pressure levels by scahng inversely to inlet pressure. This correlation has been used successfully to reduce data scatter in production JT9D-7 engine carbon monoxide and unburned

hydrocarbon data, as discussed in Reference 2.

d. Calculation of EPAP Values(ClassT2)

Values for the Environmental Protection Agency Parameter (EPAP) were calculated on the

basis of the emission indices extrapolated to engine conditions. The parameter is defined as

follows:

4

X Eli Wfi TIM i

1

EPAP -

4

_. FN i TIMi

1

where:

El = Emissions index

Wf = Fuel flow

TIM = Time in mode

i = Mode index

FN = Net thrust

The mode indices and times in each mode are defined in "Fable XI1.

Since tlle fuel flow, time in mode, and net thrust for the JT9D-7 engine are all l:nown for

each operating condition, the calculations for this program were simplified by defining coefficients combining these terms for each operating condition. These coefficients were defined as:

Wfc TIMc

EI'AP Coefficient c 4

FNi TIM i

i

39
where c denotes tile operating condition for the particular cocfficicnt. The resulting values

for the coefficients are presented in Table XII. With these coefficients, EPAP values could

be calculated by multiplying the emission indices for each operating condition by the appro- . .....

priate coefficient and summing.

TABLE XlI

DEFINITION OF EPAP CONDITIONS

AND EPAP COEFFICIENTS FOR JT9D-7 ENGINE _"

Time EPAP

Index Operating in Mode Coefficient

- Number Condition (Minutes) (lb/hr lb)

1 Idle 26.0 0.1763 (Unbled)

0.1728 (Bled)

2 Approach 4.0 0.0682

3 Climb 2.2 0.1065

4 Take-off 0.7 0.0414

- 40
t .... I1 1

CHAPTER III

PHASE II PROGRAM RESULTS

A, INTRODUCTION

The following sections present emi_ions and performance data for both eombustor con- --

cepts. The data sections are followed by a discussion of the development status of each

concept and identification of tile concept selected for the Phase III demonstration program. "

The data are the results of tests conducted on seven Hybrid and 15 Vorbix combustor configurations. The program accumulated 338.5 hours of rig testing. The emissions data except for the smoke data have been scaled from the rig test ccnditions to typical engine operating conditions using file procedures described in Chapter II.

B. HYBRID COMBUSTOR

1. EMISSION RESULTS

Idle Results

The test results obtained for the Hybrid combustor configurations at idle conditions are

summarized in Table XIII. Data are presented for both idle conditions (with and without

simulated compressor bleed) since data were obtained at both conditions and correlation

of tile data to a single set of conditions was not practical. Generally, the test condition

with compressor bleed would be expected to produce higher emissions since the combustor

inlet pressure and temperature are lower, both of which are known to produce higher emissions, and because the fuel flow rate is lower, resulting in less effective fuel atomization.

As shown in tile table, nearly all Hybrid configurations closely approached or exceeded the

program idle pollutant goals. Best results at idle with bleed were achieved with Configuration 1t4, where carbon monoxide and total unburned hydrocarbon emission indices were

60 percent and 87.5 percent below the goal values, respectively. Best results at the unbled

idle condition wele achieved with Configuration I17 where the carbon monoxide and total

unburned hydrocarbon emission indices were 80 percent and 28 percent below the goal

values, respectively.

Tile tabulated data indicate significant differences in idle emissions for the different configurations. The differences appear to be related to the pilot zone equivalence ratio rather

than to changes in the basic emissions characteristics. This effect can be seen in Figure 22

which is a plot of carbon monoxide emissions versus overall fuel-air ratio. This shift is particularly noticeable for Configuration 1t5 for wldch the pilot zone equivalence ratio was increased 50 perccot at the design filel-air ratio over that of tile other configurations.

41
FABLI, Xlll

IIYBRII) COMBUSTOR EMISSIONS AT tl)LF, CONI)ITIONS

Emissions Index (g/k_ fuel)

Oxides Totai Corn bustion

of Cart,on Unburned Efficiency

Configuration Nitrogen Monoxide Hydrocarbons (Percent)

Goal 20.0 4.0 99.0 ,..

WITH BLEED

?. H1 3.1 12.0 8.0 98.8

H2 3.7 9.0 0.7 99.7

H3 3.4 10.0 0.4 99.7

H4 3.2 8.0 0.5 99.8

H5 3.7 17.0 0.5 99.5

H6 3.8 9.0 4.4 99.3

?

H7 3.3 44.0 4.1 98.5

WITHOUT BLEED

-_ H3 2.6 24.4 6.5 98.7

, H7 2.8 3.9 2.9 99.6

I

Notes:

--'_ With Bleed: Emissions data for oxides of nitrogen corrected to combustor inlet total

• temperature of 428K, inlet pressure of 2.93 atm, reference velocity of

18.3 m/s, humidity of 6.29 g/kg, and fuel-air ratio of 0.0126.

"- Emis.sions of carbon monoxide atd total unburned hydrocarbons cot-

: retted to combustor inlet pressure of 2.93 atm and fuel-air ratio of

.. 0.0126.

Efficiency based on carbon balance data.

Without Bleed: Emissions data for oxides of nitrogen corrected to combustor inlet total

temperature of 464K, inlet pressure of 3.95 arm, reference velocity of

20.6 m]s, humidity of 6.29 g/kg, and fuel-air ratio of 0.010.

Emissions of carbon monoxide and total unburned hydrocarbons corrected to combustor inlet pressure of 3.95 atln and fuel-air ratio of 0.010.

Efficiency based on carbon balance data.

42

L d
200.000.

CORRECTED TO:

160 • 000. • 2.03 ATM COMPRESSOREXIT PRESSURE

CONFIGURATION H1 m',

60 • 000, "1" CONFIGURATION H2

X CONFIGURATION 143 ....

In CONFIGURATION H4

4) CONFIGURATION H5

40°000- A CONFIGURAT_ONH6

_p CONFIGURATION H7

ou. 120,000

I00.000

X

0

2

0

Z

2

0

¢m 80,000.

g

DESIGN t

FUEL.AIR

60.000. _/,_ATIO

I

,o.ooo.

20.000.

0.0 , , j I

0.0 0_0060 0.0120 J 0.01' 60 '

OVERALL FUEL AIR RATIO

Figure22 Combustor CarbonMonoxide EttlissJot!Levels asa l.hJctlon of FueI-A r Ratio at Idle

Co;idtttot_sICithBleed

43
Take-off Rezults

"lqzetest results obtained at take-off conditions are presented ill Table XIV. None of the

configurations met the goal for oxides of nitrogen emission levels at sea-level take-off conditions, ltowever, levels of oxides of mtrogen significantly lower tim I those produced by the

current prodnctao 1 JT9D-7 eombustor were achieved. Tile best configurations demonstrated oxide of nitrogen emission levels that were approxinlately 60 percent lower than those

produced by the current JT9D-7 eombustor. These levels are still approxhnately 50 percent

above the emission index-level goal of 10....

The take-off results generally do not sbow significant changes in the emission levels as a result of the changes in configuration. Configuration H4 is prominant, however, with its

significantly higher emissions of oxides of nitrogen. This difference was caused by an increase in main zone equivalence ratio produced by shifting the main zone outer s_irler

flow to tile outer liner dilution holes.

Smoke numbers recorded at take-off conditions were very low, ranging from 1 to 4 for all

Hybrid col_figurations. }towever, it should be noted that the absolute SAE smoke numbers

reported were measured at rig pressures and were not corrected to engine pressure levels.

TABLE XIV

HYBRID CObHIUSTOR [:.MISSIONS AT SEA-LEVEL TAKE-OFF CONDITIONS

Emissions Index I_/k2 fuel)

Pilot Oxides Total Combustion

Fuel-Air of Carbon Unbunled liffieiency Smoke

Coflrlaulation Ratio Nitrogen Monoxide Ilydrecalbon_, (Percenl) Number

Goal 10.0 99.0 I $

Ill 0,u0455 13.89 IL79 9.40 98.5 --

0.00560 14.92 10.65 6.14 99.0 -

0.00681 16.28 _.12 2.01 99.7 ....

112 0,00435 18.59 19.36 6.09 98.8

0.00553 17 90 3.95 0.20 99.9 I

0,00791 14.61 3,38 0.17 99.9

114 0.00541 14.98 R.64 0.83 99.7

0.00731 24,52 1.37 0 100.O

0.00968 24,59 0.75 0 99,9

115 0.000329 16.61 ]3,04 Io.37 97.5

0.0057. _. 17.09 4.20 O.32 99.9

0.011970 25,88 _1.40 0.61 99.8

116 0.00223 22.02 16.31 8.C_3 9S.(_

O.O030b 20.05 20.28 13.47 97.9 4

0.004,13 19,06 I _.5.1 3.f_? 99,2 2

0.00759 16.57 5.t_8 O.t_9 99.8

tit 0.007¢,5 16.32 4,62 0.36 99.8

0.009,11 18.58 1L52 0.(_9 IJt).8

_u|¢$_ I:llli_OOfls dato for oxides tl f niltog_Lii COll_.ted Io t Ol!10u_tol inlel total

l¢lnllcraturt ' of 767K.._¢1 iltcssure of 21,09 atnl, r_fep:l_ce WIo,:ily of

25.g m/s, hunddily of (s.29 g]kg, and ¢olnbtzstor ex0 total tenlperaturv

of 1486S (overall fuel-air ra0o of 0.0215 ).

I!lldSOorl$of eMI)Oll fllOllOXid¢ Ind total urlbUl_led h_dro_sTholls ¢orttwled to conlhuslo? inl¢l pte_ute of _ 1.09 _llll.

I_ltirlen':y I as_d tin t:_l$I rLIlalatt_e dat,I

44
ApproachResults

The results obtained at _imulatcd app_.oaeh conditions are tabulated in Table XV. _lese

restllts showed that the lowest emissions at approach conditions were obtained with fuel

supplied only to tile pilot zone. With fuel supplied to both the pilot mid main zones, high

emissions of carbon monoxide and unburned hydrocm'bons resulted, although the emissions

of oxides of nitrogen were somewhat lower than when only the pilot zone was used. Supplying fuel to the main zone only produced lower emissions of carbon monoxide and unburned hydrocarbons than obtained when both zones were used, but the results were not

as good as those obtained with fuel supplied only to the pilot zone. Selected configurations

were tested with fuel supplied to only a portion of the mahl injectors, and in all cases, effieciency improved as the number of main.nozzles fueled was reduced.

TABLE XV

IIYBRID COMBUSTOR EMISSIONS AT APPROACII CONDITIONS

Emissions Index II_[klZI_el_.= -

Pilot Main Burner Oxides Total Combustion

Fuel-Air Fuel-Air of Carbon Unburned Efficiency

Configuration Katio Ratio Nitrogen Monoxide hydrocarbons (Percent) .

112 0.01039 0.00280 4.18 34.56 229.20 72.7

0.00683 0.00671 2.09 88.92 239.03 70.3

113 0.01319 0 8.72 1.50 0.95 99.9

1t3 - 11 ntain injectors 0 0.01352 7.75 10.57 1.59 99.6

1t3 - 7 nlain injectors 0 0.01349 7.13 9.19 0.32 99.7

n4 - 4 nlain injectors 0 0.01341 6.51 41.27 1.35 98.9

114 0.01381 0 9.34 1.71 0.41 99.9

115 0.01325 0 12.0i 1.13 1.45 99.8

115- II main injectors 0 0.01354 3.61 98.°-4 76.56 88.8

tl5 - 7 main injectors 0 0.01338 5.53 65.93 63.02 91.2

115- 4 mail1 injectors 0 0.01280 6.55 59.31 38.86 94.1

116 0.00690 0.00697 3.75 7(,.64 66.42 90.5

0.00480 0.00878 1.52 64.56 551.45 34,3-5-

117 0.01338 0 9.90 23.88 0.72 99.4

Notes: [:.nlissiollSdata for oxides of 1101ogt'ncorrected t_ combtlstor inlet total

telllpet'iIUlre of 586 K, inlet pressure of 8.51 aim, reference velocity of

23.3 lU/S,humidity of 6.2_I. s/ks, and ¢onlbuslor exit total telllperaturc

of I(J62K.

l_lllis_ioIt_ or =gilrb¢llt Intlnoxide _llld ttltal tlllbtn'n_d hydro_.'ar_on_ corrected

to colnbu_tor inlet prt._ule of 8.51 _llnl.

Efficiency ba_d ,JiI ¢;_Ibon balance data.

This behavior is related to the narrow stability limits of the pilot and main zones of the

llybrid combustor and to the relatively severe operating conditions at approach. At these

conditions, the overall fuel-air ratio is too low to permit simultaneous efficient operation

of both stages.

45

m .
'lllis result is unfortunate from ,'moperational standpoint since it would require staging of

the main zone during cllgine ;:cceleration for a missed al_proach, thereby increasing the enginezacceleration time from the approach condition.

Climb Results

Tile test results obtained at simulated climb conditions arc sllown in Table XVI. These data_

were obtained at fuel splits that provided acceptable efficiency levels at take-off conditions.

At climbout, combustion efticiencies were generally over 99.5 percent with oxides of nitrogen levels reduced 66 percent relative to the current production JT9D-7 levels.

TABLE XVI

HYBRID COMBUSTOR EMISSIONS AT CLIMB CONDITIONS

Emissions Index (_/k_ fuel)

Pilot Oxides . Total Combustion

Fuel-Air of Carbon Unburned Efficiency

Configuration Ratio Nitrogen Monoxide Hydrocarbons (Percent)

H2 0.00721 10.35 10.46 1.29 99.6

H4 0.00515 21.75 29.56 29.34 95.9

0.00607 18.81 33.56 35.05 95.2

0.00747 16.42 39.77 40.87 94.3

I-t5 0.00518 9.89 12.26 1,47 99.5

116 0.00757 12.22 I 1.24 1.36 99.6

117 0.00756 14.17 11.24 1.55 99.6

Notes: Emissions data for oxides of nitrogen corrected to conlbustor inlet total temperature

of 735K, inlet pressure of 18.51 atm, reference velocity of 25.5 m/s, humidity of

6.29 g/kg, and combustor exit total temperature of 1396 K.

Emissions of carbon nlonoxide and total uuburned hydrocarbons corrected to combustor inlet pressure of 18.51 atm.

Efficiency based on carbon balance data.

CruiseResults

Emissiort data were recorded at simulated altitude cruise conditions lot selected confignrations because of re:eat concern with the effects of pollutants released by aircr,_ft engines in

the stratosphere. The emission data arc presented in Table XVII.

46
I

I

TABLE XVII

IIYBRID COMBUSTOR EMISSIONS AT CRUISE CONDITIONS

Emissions hldex (_/k_ fuel)

Pilot Oxides Total Combustion

Fuel-Air of Carbon Unburned Efficiency

Configuration Ratio Nitrogen Monoxide Hydrocarbons (Percent)

_llb.,

JT9D-7* - 22.2 1.0 0.2 -

116 0.00768 7.89 23.93 4.26 98.9

117 0.00762 9.54 28.52 4.11 98.8

Notes: *F.stimated values

Emissions data for oxides of nitrogen corrected to combustor inlet total temperaturc of 704 K, bdet pressure of 9.31 arm, reference velocity of 23.8 m/s, humidity

of 6.29 g/kg, and combustor exit total temperature of 1413 K.

Emissions of carbon monoxide and total unburned hydrocarbons corrected to cornbuster inlet pressure of 9.31 arm.

Efficiency based on carbon balance data.

Best results were achieved with Configuration 116where combustion efficiency was 98.9

percent and tile emissions of oxides of nitrogen were reduced 64 percent relative to the current production JT9D-7 estimated level.

EPAP Results

The emission results for idle, approach, clilnb, and take-off conditions were used to calculate F.PAI' values for each configuration tbr whicl_u fficient data were available, and the results are presented in Tables XVIII and XIX. The emission levels used in the calculations

were selected from test data curves to provide the best overall emission levels when compromises between one class of emissions and another were required. The data used tbl the

approach condition correspond to operation of the pilot only for 'all configurations except

H2 and tt6, where both zones were fneled.

As shown in Tables XVlll and XIX, the best overall results were obtained with Configuration 1t7. Relative to the JT9D-7 production engine EPAP valnes, this configuration provided a 42-percent reduction in the emissions of oxides of nitrogen, a 56-percent reduction in

the emissions of carbon monoxide, and an 82-percent reduction in tile emissions of unburned hydroc:lrbons. 'l'be iolprovements in carbon monoxide and total unburned hydrocarbon emissions were obtained primarily at idle conditions while the improvement in the

emissions of oxides of nitrogen were obtained primarily at take-off and clinlb conditions.

Confignrations 112 ;rod 116, which were operated staged at approach, meet the oxides of nitrogen goal at tile expense o1"large increases in unburtlcd hydrocarbons and et_rbon monoxide.

47
'FABLI _,XVIII

IIYBRID COM IIUS'FOR EPA I)ARA MF,TI';R EMISSIONS RESULTS

E'I'A Parameter

(Ibm. pollutant/1000 lbf, thrust -

hr/_anding-takc-off cycle)

Oxides of Carbon Total Unburned SAE Smoke Idle

Configuration Nitrogen Monoxide Hydrocarbons Number Condition ..,

JT9D-7 (1) 5.9 8.5 3.9 4 Without Bleed

Goal (2) 3.0 4.3 0.8 15

H2 2.5 8.9 16.6 i With Bleed

tt4 (3) 4.5 4.7 3.2 - With Bleed

H5 (3) 3.3 4.6 0.4 - With Bleed

H6 2.9 8.2 5.5 4 With Bleed

H7 (3) 3.4 10.8 1.O - With Bleed

H7 (3) 3.4 3.7 0.7 - Without Bleed

Note: All data corrected to JT9D-7 engine conditions.

(1) Based on initial production test results for combustor EC 289386.

(2) 1979 EPA rule levels for class T2 engines.

(3) Pilot only fueled at the approach conditions.

IAlll I: XIX

I.ANIIING. I AK FOf:F ('Yt'I I P(IIN I I MISSION INDIt'FS FOR IIYIIRID COMIIIISI'OR

_4.1 e_H

Idle Ilfl¢¢dl A_ ..... (luuh --- Take ",q[ ('tutk

total Oxides Total (Ixtdcs 1.ill ()xtde_ T.lal • )xldes Told Oxid¢_

('album ihthumed ur Valhon IIitb_m_d q,I t "llb,m Ilnburned ,d (arbor. Unburned t,f ('llbon UnburlWd of

('onflgulllion Monoxkl.__r [lydle_IrlHm+ N2tm$¢n M_+Ia,x_¢ ll_dlq_i*t_,ll_ NtlluJtcn M,m,,_td¢ I[_,'_=ib_ms _ 51om'xid_ llJ_dlc_Id'ams _ _ _ _ .

Et'('p (;uml 20.20* 4LI. 4N* IOO

IT'B) 7

PltVJueUon 4(+ 7* 2I._* _ .)* 3 _ O 7 1._ tl '_ O 4 2g.lL O4 0.4 1t_1'+ lOt O2t 22.2t

III 110 ,_0 ) I 2 12 ?of la._x

112 _LII O 7 .t _ _iA,_ 2 lq.O 2 I IO _ I _ 10.4 L4 O 2 14 b

llJ 100,244 _ 0.4.65 + L4.26' 15+* tI.o_¢ IL72_

H$ I_fl O_t L7 II II 14_ _ 12()I_ 121 I _ +lq ,I 2 O] 11.7

H_ _0 44 _ 71,1_ Ld_.$ In I] _ 1.4 1_ _7 07 lie6 _J_ 4.J 7.r_

II1 44 O. _.')' 41._q t $.1.2._ = _I_ *_ 07 +" ,+,_+* I17 16 14_ 41, el4 II, I 2¢"+. 4 I 9_

• _lth,mt bleed

I_pl_ol truly [+ieled

.+ I _tlult|led _31ue

" 48 OIfl(;_,, -

uvlk/._
I | I 11

.............. i

I "

2. PERFORMANCE RESULTS

In addition to tile combustor emission lneasurements, performance measurements were

made during tile Phase 11progl"am to determine pressure loss, exit temperature pattern, idle

lean blowout limits, altitude stability and relight characteristics, and durability. The results

of these measurements for the Hybrid combustor are summarized below.

PressureLoss

mr,,,

The measured values of tile combustor, diffuser, and total combustion system pressure losses

are summarized in Table XX. As shown, tbe Hybrid combustor generally operated close to

the goal of six percent.

TABLE XX

SUMMARY OF PRESSURE LOSS TEST RESULTS

FOR HYBRID COMBUSTOR

Overall Diffuser

Pressure Combustor and Shroud

Configuration Loss (Percent) Loss (Percent) Loss (Percent)

Goal 6.0 ....

JT9D-7 5.4 ....

It I 6.5 2.6 3.9

H2 6.4 2.5 3.9

H3 6.7 2.3 4.4

1-14 6.7 2.3 4.4

1t5 5.6 2.1 3.5

116 5.8 2.4 3.4

1t7 6.4 2.8 3.6

Exit Temperature Data

Exit temperature traverse data were taken :_tidle conditions a,d at simulated take-off conditions. 'lhc resulting data were plotted as circunlfcrential profiles which were then reduced

to exit temperature I_attern factors and average radial temperature profiles. It should be

noted that tile maxinlunl allowable temperature for the traverse rakc was 1800 K, which precluded the acquisition of data at the full sea-level take-off fuel-air ratio. As a result, most

data taken at high power conditions were obtained at a fuel-air ratio of 0.016. For these tests,

pattern factors were calculated on thc basis of the inaxinlunl temperature observed with the

operational therulocouples. The average eonlbustor exit temperature was COUlpnfed f;'onl

the measured fuel-air ratio.

49
The circumferential temperature profiles for the llybrid combustor at idle aDd sea-level takeoff conditions are shown in Figures 23 and 24. These profiles show regions of increased temperature in line with the diffuser struts, particularly al idle conditions. These data suggest

that the sheltered regions dowustream of tile struts arc functioning as flameholders for the

pilot zone.

,CONFIGURATION H1 IDLE : , '

COMPRESSOR EXIT TEMPERATURE = 4281< : , !

COMPRESSCR EXIT PRESSURE = 2.03ATM ; .... _'-'" _'"':

1500 FUEL AIR RATIO = 0.0126 _ • ." .i w.,.

Q

¢. • : - ;- I

i il

¢' .; . . e - i--i ...........

14oo . _ ,. : i "

l

..... ,,....... , ! , . '. . _ +i :

I

1300 ' ' " : : : -

: , , [ . . '. ! i ! ,''i" i .... ! ....

. I _i:: !_ ,. ': ......: _,':,'_-!........ _,_=:: . .... . ,_-,, ....

+ i ,-; : • . Q = : I i I : : : : ; i '

1200 ..... r_ : -- • i --- ;-- • I ! , . J : "

, ' - • , i ! 4- ; - I ! ,. ;

_' : : + ' ! . . ! ! .

,.u I -' ] _ ," ii : -

: ! ! : .

I.- " ; , , " : . "

tw I . :

1000 ....

a. : - '. : + " " :. ' ' " " '. i "

1'- ' ! i : ; " : ; • RAO,IALAVERAGE

900 ' ....

RAKE READ 10% SPAN

i I , " RAKE HEAD 31% SPAN

800

x = ; : / x X RAKE HEAD 51% SPAN

: i • x J RAKE HEAD 71% SPAN

: i

• • = RAKE HEAD 90% SPAN

700 ; •i 1<

i I = ® =

600 _ ; ', ;

...... ="" F'i _=

, i . i• . STRU'r

.... i ' ' i",*i I'/

, I i •

R0( ' ; ' : " l! @ @ ] I I _ I ? I'. In I I I I I"

0.2 0.4 0.6 0.8 1.0 | ,2

CIRCUMFERENTIAL POSITION _" RADIAN$

I'igure 23 Circum_ren_ia_ E:dt Temperatt_re Patt_rt_ _ _bHd _m_t_st_r C_n_Nrati_n _ at1d_e

Condititms

5O
l,_$tre 24 Circttmferentia/ l:_it "l'emperat_treI'(tttcms Jbr Ilybrid C_smhustc_rCo_digztratim_112 at

Sea-Level Takc,.Off Conditions ICLthIq_eI.Air Ratio o.f O.016

51
! i

As shown in Table XXI, tile pattern factor values for tile llybrid eombustor ranged between

0.40 and 0.47 at sea-level take-off conditions. Although these values are significantly higher

than the goal of 0.25, there was no attempt made to improve the pattern factor during the

llybrid test program. Values arc comparable to those for the current production JT9D-7 enghre. Tile pattern factors for the Hybrid combustor at idle conditions are very high because

of tile radial staging used at idle conditions, but, since the average temperatures are low at

idle conditions, tile high pattern factors may not pose a turbine durability problem.

TABLE XXI

SUMMARY OF HYBRID COMBUSTOR EXIT

TEMPERATURE PATTERN FACTOR RESULTS

Pattern Factor

Idle Take-Off -

Configuration Conditions Conditions

Goal - - 0.25

JT9D-? - - 0.42

HI 1.33 0.45

H2 1.49 0.44

H3 1.02 - -

H4 1.10 0.47

H5 1.84 - -

H6 1.94 0.40

H7 1.43 - -

A typical radial temperature profile for the Hybrid combustor at idle conditions is shown in

Figure 25. The high temperatures near the inner wall of the combustor are a consequence

of the pilot zone location.

At sea-level take-off co,lditions, the Hybrid combustor produced a radial temperature profile

that approached the baseline profile for the JT9D-7 combustor, as shown in Figure 26. The

temperatures are shghtly higher than the baseline profile near the outer liner and slightly lower

near the inner liner.

Idle Lean Blowout Limits

Table XXII summarizes tile idle lean blowout results for the Hybrid combustor. The lean

blowout limits were all significantly higher than those for the JT9D-7 combustor and are

unacceptable. Poor blowout characteristics are characteristic of premix type combustors.

52

_ - 7
1300

120( C(_'_ COMPflESSOH EXIT TEMPERATURE = 428K

COMPf_ESSOI1EXIT PRESSURE 2.93 ATM

EL AIH RATIO 00120

_IIOG

W

3

F- 9_

8 7_

5oo I I i I I I __ 1 I I r

10 20 30 40 50 60 7O 80 SO ;EO

INNER LINER PERCENT TOTAL SPAN OF LINER OI'fER LINER

Figure 23 Radial Exit Temperature fol Hyb_d Combustor Configuration 1t2 at Idle Conditions

1700

_ 1600 :ONFIGURATION 0 H2 __

.¢ 150.0-- // /../ JTgD-7" ......

4#" / TARGET PROFILE

X

/

W 1400

0L

_:_ COMPRESSOR EXIT TEMPERATURE = 769K

1300 COMPRESSOR EXIT PRESSURE 6.8 ATM

FUEL AIR RATIO = 0.0160 1530°K NORMALIZED TO

¢J

120t] I I l l I l I I I" I I I I I I I i I I I

10 20 30 40 50 60 70 80 90 100

INNER OUTER LINE|_

LINER PERCENT TOTAL SPAN OF LINER

l'igure 26 Radial l';xit Temperature Pattem [or llybrid Contbustor Cot Jigurat_o H2 at ,_ a-Lew'l

Take.Off Conditions With Overall Fuel.A irRatio of O.0160 atul Pilot Zone lqteI.A ir Ratio

o[0.0052

53
._ I I I ............

i i

l

i

- TAIILE XXII

SUMMARY OF IlYBRID COMBUSTOR IDLE

• t

: LEAN BLOWOUT FEST RI.SULTS

E- Idle

, Lean Blowout

_- Configuration Fuel-Air Ratio

Goal _ -,,.

JT9D-7 < 0.004

.._' H1 0.0077

_: _ H6 0.0063

H7 0.0067

. , Notes: Data taken with sinlnlated compressor bleed. Nominal inlet conditions were

. inlet pressure of 2.93 aim and inlet temperature of 428K.

Altitude Stability and Relight Results

Altitude stability and relight characteristics were determined for Hybrid combustor Configurations H5 and H6, and the results are shown in Figure 27. The data have been plotted on

the JT9D-7 engine relight envelope for comparison purposes. The outer envelope indicates

the engine requirement for mlighting with a hot engine, while the rig baseline limit indicates

the requirement for lighting from a cold-soak condition, w, ich was the condition used for

the Phase II combustor tests. It should be noted, however, that the minimum pre=ure blowout limit must exceed the hot engine relight requirement by a considerable margin if the

engine is to accelerate successfully after ignition.

As indicated in Figure 27, only minimum pressure blowout data were obtained for the Hybrid combustor, since neither configuration tested could be lit at altitude conditions. Configuration H5 reached its minimum pressure blowout limit below the minimum required al-

- titudc for relight, and, therefore, has no potential for meeting the relight requirement. Configuration H6, however, demonstrated a blowout limit above the relight altitude requirement and, therefore, might be developed to meet the relight requirement through relocation

of the igniter plugs or modification of the fuel spray pattern. 1"he difference in altitude

characteristics is attributed to the 50-pereent leaner pilot zone equivalence ratio in Configuration 116.

Combustor Durability

Durability problems in the Ilybrid combustor were confined primarily to the pilot flameholder region. / series of modifications were made to correct the situation, and some ira-

• provement was obtained. In t'onfiguration 112, the design ofthe flanleholder attachment

to the inner liner was modified to permit thermal expansion of the fiameholder. In Configuration 113, cooling holes were added in the trailing edge of the flameholder, and in Configuration 116, cooling of the outer portion of the trailing edge of the flameholder was increased. "I]lcse changes were effective in improving the durahility amt penuitted 80 hours

P! 54
of testing without further modifications, llowever, significant deterioration still occurred,

primarily in the area near the attachment of the flameholder to the inner liner. The durability or the Hybrid eombustor Ill,ors and main burning zone appeared satisfactory throughout tile test program.

B.B81 K_JIE O,BB%KBIB 1,034 Kg/S 1,249 KglB

242 K 250 K 261 K 269 K

105OO

gO00

UJ

p-

_U "/500 =E

;¢ JT9D-7 RIG

-- BASELINE U.I

RELIGHT

6000 REQUIREMENT t'-

t-

,,,1

<¢

4500 RELIGHT ENVELOPE

LEGEND

3ooo ql, H-B

H-6

I I

0.2 B.3 B.4 0.5 B.7 0.8 0.9 1.B

FLIGHT IVlACH NUMBER M N

Figure27 Minimum PressureBlowout Test Resultsfor Hybrid Combustor Shown on JTBD-7Altftude

Rellght b.hvelope

C. VORBIX COMBUSTOR

1. EMISSION RESULTS

Idle Results

The emission resuhs obtained at idle conditions are summ.'lrized in Table XXIII.

Data were obtained both will) and without simulated bleed. Gel)erally, the trends were the

same for both sets of conditions for emissions of unburned hydroc:]rbons and oxides of nitrogen, but the emission levels for carbon monoxide were significantly lower without simulated bleed.

55
"- TABLE XXIII- --

VORBIX COMBUSTOR EMISSIONS AT IDLE CONDITIONS

Emissions Index (.,_/kgfuel)

Oxides Total Combustion

of Carbon Unburned Efficiency

Configuration Nitrogen Monoxide tlydrncarbons (Pereont)

Goal 20 4 99.0 ""'

WlTII BLEED

SI 1 3.8 39.0 1.2 98.9

S12 3.4 37.7 1.7 98.9

S13 3.4 37.7 1,7 98.9

S14 3.4 37.7 1,7 98.9

S15 3.4 37.7 1.7 98.9

S16 3.2 61.6 1.0 98.4

S17 2.7 55.3 1.3 98.5

S 18 2.9 69.8 1.1 98.2

S19 2.8 64.8 0.2 98.4

$20 3.1 46.3 6.2 98.2

$21 2.9 43,0 2.3 98.7

S22 3.1 54,4 2.8 98.4

, - $23 3.6 35.2 1.0 99.1

$25 2.9 40.:_ 0.5 99.0

WITIIOUT BLEED

$20 4.0 27.0 3.5 99.0

$21 3.4 28.0 i,6 99.2

- $22 3.7 39.5 2,8 98.8

- $23 4.1 35,0 2.0 99.0

$24. 3.8 32.0 2.7 99.0

$25 3,2 18.5 0.5 99,5

: Notes: Emissions data for oxides of nitrogen with bleed corrected to combustor inlet total

temperature of 428K, inlet pressure of 2.93 atm, reference velocity of 32.0 m/s,

humidity of 6.29 g/kg, and fuel-air ratio of 0.0126.

.... E'missions data for oxides of nitrogen without bleed corrected to combustor inlet

temperature of 464K, inlet pressure of 3.95 atm, reference velocity of 35.9 m/s,

humidity of 6.29 g/kg, and fuel-air ratio of 0.0100.

Emissions of carbon monoxide and total unburned hydrocarbons corrected to corn-

- " buster inlet pressure of 2.93 atm and fuel-air ratio of 0.0126 with bleed attd to corn-

"- buster inlet pressure of 3.95 and fuel-air ratio of 0.0100 without bleed.

Efficiency based on carbon balance data.

56
lI  ̧

I I

.- (Jenerally, idle pnllution results were dJsapl)oJatJnt' ill tha_ carhon nlouoxide conccntralJons

: exceeded the goals for nearly all conligurations investigated. Hnwever, the carbon illonoxide goal was achieved with ('nnfiguration S25. The total unburned hydrocarbon ,,,eels wcrc

achieved with most configurations.

Attempts to reduce cal'bon monoxide through primary zone modil]cations including airfow

redistributions, fnel spray modification, and volume modifications were ineffective. The

lack or response of the Vorbix eombustor to cotwentional, idle emissions reduction techniques suggests that the :hroat region and the massive introduction or dilution airflow ,..

through the main swirlers immediately downstream of the throat were prematurely quenching the combustio1: gases ant! thus inhibiting carbon monoxide oxidation. Test results from

the later Vorbix configurations, in which reduction of main swirler airflow was accompanied

by reduced idle carbon monoxide levels, tend to support this conclusion.

Take-off RasultL._

Tile test data for sea-level take-off conditions is summarized in Table XXIV.

The Vorbix combustors exhibited emission levels for oxides of nitrogen that were signifi-

- cmltly below tile current JT9D-7 production combustor level. This improvement was ob-

:_ taincd concurrently with high combustor efficiency. Several configurations approached or

exceeded the goal for emissions of oxides of nitrogen.

The modifications to reduce emissions of oxides of nitrogen involved changes in swirler and

dilution airflow distribution. Significan' reductions in emissions were achieved, with Configuration $23 producing an emissions index for oxides of nitrogen of only 9.0 with an

efficiency of 99.6 percent, surpassing the sea-level take-off emissions index goal for oxides

of nitrogen of 10. Configuration $25 achieved an emissions index for oxides of nitrogen of

10.8 with ,an efficiency of 99.8 percent. Although the lowest emissions for this configuration were slightly higher than those for Configuration $23, the overall emissions and efficiency characteristics of Configuration $25 were superior, as shown in Figure 28.

Reducing the number of main fuel injectol_ by 50 percent and the use of premixing tubes

were investigated, but neither produced significant reductions in the emissions of oxides

of nitrogen. The premixing tubes offer potential benefits of reduced tendency lbr nozzle

eoking aml relatively constant fuel spray penetration at all operating conditions, but exhibited localized durability problems. A full evaluation of the potential of the pr,:m,:xing

tubes was not pursued ill the Phase I1 program.

: As with the ltybrid combnstor, smoke levels measured at rigpressures were very low, ranging from SAE smoke numbci._ of I to 5. Configuration S12 was an exception, bowever,

producing an SAE smoke mtmber of 9. It should be noted that the absolute ShE smoke

nulnbcr$ were measured at rig pressures and were not corrected to engine plessure.

57
TABLE XXlV

VORBIX ('OMBUSTOR EMISSIONS AT SEA-LEVEL TAKF+-OFF C'ONI)ITIONS

|;,IllissioIISIndex _/.k_ fuel)

Pilot Oxides Total Combustion

Fuel-Air of Carbon Unburned Efficiency Smokc

Configuration Ratio Nitrogen Monoxide Hydrocarbons (l'crcent) Number

,p_-+

Goal 10.0 15

S11 0.00359 14.60 10.70 0.80 99.7 - -

0.00471 15.28 3.83 0.10 99.9 - -

0.00563 16.09 2.36 0.03 99.9 - -

0.00645 16.86 1.72 0.08 100.0 - -

Sl2 0.00357 15.62 18.35 1.22 99.4 9

0,00465 15.46 16.58 0.44 99.6 - -

0.00580 16.85 12,72 0.34 99,7 - -

0.00730 18.46 10.37 0.27 99.7 - -

S13 0.00884 21.68 6.26 0.16 99.8 - -

Sl4 0.00420 19.72 15.71 0.90 99,5 - -

0.00579 21.79 4,91 0.35 99,8 - -

0.00900 28.79 4.80 0.12 99.9 - -

Sl5 0.00412 13.51 23.24 1.26 99.3 - -

0.00579 16.07-- 19.75 0.71 99.5 - -

S16 0,00521 17.57 8.35 0.16 99.8 - -

0,00754 22.87 7.22 0.08 99.8 - -

0.00970 22.64 2.75 0 99.9 -- -

S17 0.00215 12,87 26.73 3.52 98,9 - -

0,00422 18.02 11.23 0.36 99.7 - -

0.00542 19.76 8.10 0.22 99,8 - -

0,00748 22.48 6.65 0.16 99.8 - --

0.00983 24.71 3.85 0,07 99,9 - -

SI8 0,00322 20,32 11.27 0.26 99+7 ....

0.00417 23,18 8.86 0.13 99.8 .....

0,00534 25.76 7.36 0,10 99.8 .....

0,00733 24.97 4,09 0.03 90.9 -- +

0.00970 24.09 2.63 0.02 99.9

SI9 0.00776 - - 3.45 0,07 99,9

58
TABLI_ XXIV (('ont'd)

F,nfissions Index (g/kg fuel)

Pilot Oxides Total Cond_ustion

Fuel-Air of Carbon Unburned Efficiency Smoke

Configuration Ratio Nitrogen Monoxkle ltydrocarbons (Percent) Number

$20 0,00224 13.06 20,63 4.02 99.0 4

0.00329 13.79 13.97 0,62 99.6 2 _..

0.00445 14.74 9.99 O.13 99.7 4

0.00767 18,09 3.40 0,03 99,9 3

821 0.00310 11.64 31.87 8.45 98.3 - -

0.00414 14.04 13,93 0,58 99.6 - -

0.00692 19,93 4.60 0.26 99.9 - -

0.00747 "•.1,_8 _ '_ 3.98 0.13 99.9 1

$22 0.00207 14.96 9.29 0.74 99,7 - -

0,00306 15,10 6.40 0.35 99.8 1

0.00422 16.26 4.17 0.26 99.9 - -

0.00521 16.84 2.36 0 99.9 - -

0.00600 17.67 1.89 0 100.0 1

0.00730 19.18 1.21 0.03 100.0 1

823 0.00208 9.01 0 72 1.05 99.6 2

0.00311 14.14 4,16 0.35 99.9 3

0,00427 14.80 3.39 0,03 99.9 2

0.00523 14.74 2.20 0,03 99.9 2

0.00739 18.22 1,24 0.13 100.0 1

824 0.00228 13.04 7.26 0.98 99.7 3

0.00333 12.94 4.52 0.29 99,9 3

0.00458 13.82 2,59 0.10 99.9 5

0,00555 13.82 1,22 0.10 100,0 5

$25 0.00555 .... 1.39 0,03 100.0 2

0.00439 - - 2.10 0.16 99,9 2

0.00334 11,15 2.49 0.16 99.9 2

0.00225 10.79 4.52 0.39 99.8 3

0.00587 _2.80 1.40 0.03 100.0

0.00132 11.49 7,81 0.94 99,7

Notes: Fnlissions data for oxides of nitrogen COrl-eclcdto ctmlbustor inlet total temperature

of 767 K, inlet pressure of 21.09 arm, reference velocity ot"45,0 m/s, humidity of

6,29 g/kg, and eombustor exit total tempcratul'c of 1486 K.

Emissions of carbon nlono._idc and total unburned hydrocarl,ons corrected to ,:ombustor inlet pressure of 21.09 aim.

Ffficiency based on carbon balance data.

59
I

2O

,.J

16 S 23

10

8 I 1 I I

_,9.e 99.7 _9.8 99.o loo.o

EFFICIENCY -_PERCENT

Figure28 RelationshipBetween F.missionsof Oxidesof Nitrogen trodE.fficencyfor Vorbix Combustor

Configurations$23 and$25

ApproachResults

The test results obtained at approach conditions are shown in Table XXV.

Unlike the llybrid combustor, the Vorbix combustor is capable of operating with high combustion efficiency with both the pilot and main zones fully fueled at approach.

ltowever, the overall emissions did not compare favorably with those of the current production JT9D-7 combustor at approach conditions. Although some configurations provided

emissions of oxides of nitrogen that were only one half those of the production combustor,

these configt,rations also produced emissions of carbon monoxide that were 12 times higher

than those of the production combustor. Vorbix configurations that produced emissions of

oxides of nitrogen comparable to those of the 9rodnction combustor produced emissions of

carbon monoxide thai were al;proximatcly 3 times those of the production combustor.

Climb Results

Emissions data obtained at climb t ,_ditions arc shown in Table XXVI. Tile data were obtained using the optimum pilot b, met fuel-air ratio determined at take-off conditions.

During climb, combustion efficiencics were generally over 99.5 percent with oxkles of nitrogen levels reduced 70 percent when compared to current production JT91)-7 levels.

60
TABLI_ XXV

VORBIX COMliUSTOR EMISSIONS AT API'ROACli CONDITIONS

Emissions Index (_/kg fuel)

Pilot Oxides Total Combustion

_lel-Air of Carbon Unburned Efficiency

Configuration Ratio Nitrogen Monoxide lfydrocarbons (Percent)

I _I_

S 1I 0,00744 5.07 20,40 1.65 99,3

0,00956 6.72 11.21 0,44 99,7 _,.

SI2 0.01300 9.97 6,01 0.30 99.8

S17 .0,01358 4.60 12.48 0.32 99,7

S18 0.00263 3,61 54,19 33.40 94.8

0,00547 7.04 20.76 1,53 99,3

0.00808 7.08 15.50 0.32 99,6

0.01070 6,05 15.34 1.14 99.5

$20 0.00688 5.54 15.63 2.16 99.4

0.00703 5.36 13.09 0.63 99.6

0.01040 8.57 9,48 0.16 99.8

0.01048 8,92 7.92 0,32 99.8

$21 0.00582 4.24 45,91 23.71 96.1

0.00754 5.16 19,74 2.82 99,2

$22 0,00458 3.55 47.68 14.79 97.1

0.00658 4,68 23,13 1.70 99,3

0.00935 6.5i 22.18 1.21 99,3

$23 0,00496 3.88 41.79 6.37 98.3

0.00705 4.62 22,03 1.18 99,3

0.00986 7.46 14.14 0,48 99.6

$25 0.00977 6.54 21,85 0.72 99,4

0.00682 4,34 30.94 1.92 99.0

0.00481 3.10 51.42 13.45 97,2

Notes: Emissions data fOl'oxides of nitrogen corrected to combustor inlet total temperature of 586 K, inlet pressure of 8.51 atm, reference velocity of 40,7 m/s, humidity

of 6.29 g/kg, and combustor exit total temperature of 106.. K.

Emissions of carbon n|onoxide and total unburned hydrocarbons corrected to combustor ildct pressure of 8.51 arm,

Efficiency based on c_lrbou balance data.

61
'I'AI] LE XXVI

VOI*,BIX COMI]USTOR EMISSIONS AT ('I,IMI_ CONI)IT1ONS

Emissions Index (g/kg fuel)

Pilot Oxides Total Corn bt_stion

Fuel-Air of Carbon Unburned Efficiency

Configuration Ratio Nitrogen Monoxide Hydrocarbons (Percent)

S 11 0.00418 12.03 9.12 0.54 99.7

S 13 0.00604 12.40 8.48 0.33 99.8

S17 - Pilot Only 0.01976 10.34 2.98 0.13 99.9

S18 0.00304 15.41 9.10 0.33 99.7

$20 0.00764 15.14 2.70 ....

$22 0.00209 10.50 12.93 2.37 99.4

$23 0.00389 11.37 4.85 0.14 99.9

$25 0.00228 8.53 8.80 0.91 99,5

Notes: Emissions data for oxides of nitrogen corrected to combustor inlet total temperalure of 735 K, inlet pressure of 18.51 atm, reference velocity of 44.5 m/s, humidity

of 6.29 g/kg, and combustor exit total temperature of 1396 K.

l'_'missions of carbon monoxide and total unburned hydrocarbons corrected to combustor inlet pressure of 18.51.

Efficiency based on carbon balance data.

CruiseResults

F-llliss_onsdalai obtained at sJmtdated crtlisc collditions are showr_ in Table XXVII. All configurations provided substantial improvements in the emissions of oxides of nitrogen relative

to the current production JT9D-7 engine levels, with the best configurat'o I provi_l ig an

improvement of approximately 80 percent. Emir," _ns of tot;d unburned hydrocarbons

and carbon monoxide were considerably higher tl.an those of the current production

JT9D-7 engine.

EPAP Results

t_PAP values were calctd:lted for each configuration for which sufficient data were available,

and the results arc shown in Tables XXVIII and XXIX. As for the llybrid eombustor, the :

emission levels used in l!le calculation were selected from test data curves to provide tile

best overall emission levels when compronliscs between the vario,:s classes of cmissiolls were

: required. Most ¢onfiguralions produced lower emissions of oxides of nitrogen and unburned hydrocarbons than the goal levels, hut none IIlet the goals for carbon monoxide

emissions.

-; 62
It appears th_,t ilnprovealcnts in carbon nlonoxide levels could be achieved by permitting

increases in oxides of nitrogen without exceeding that goal level. One way tiffs can be accomplished is by increasing tile pilot fuel-air ratio at each slagod simulated engine cycle point.

For example, configuration $25 produced EI'AP values of 2.2, 6.5 and 0.3 respectively for

.... oxides of nitrogen, carbon monoxide and total unburned hydrocarbons. }:or those data,

. pilot fuel-air ratios wore 0.00687 at approach, 0.00228 at climb and 0.00225 at takeoff.

Substituting data obtained with pilot fuel air ratios of 0.00977 at approach and 0.00587 at f.

_C_ takeoff produces EPAP values of 2.4, 5.7 and 0.2. It appears that increasing pilot fuel-air

ratios at the climb conditions would further reduce carbon monoxide levels. However, climbout data for $25 were obtained only at the 0.00228 condition.

Imb..

TABLE XXVII

VOP,BIX COMBUSTOR EMISSIONS AT CRUISE CONDITIONS

.. Emissions Index (g/kg fuel)

Pilot Oxides Total Combustion

Fuel-Air of Carbon Unburned Efficiency

Configuration Ratio Nitrogen Monoxide |lydrocarbons (Percent)

JT9D-7* 22.2 1.0 0.2

SI I 0.00402 6.52 28.39 41.38 94.5

S18 0._10498 13.47 15.72 0.25 99.6

0.00590 13.30 12.65 0.12 99.7

0.00692 14.06 I 1.02 0.07 99.7

$20 0.00405 7.3o 22.74 0.88 99.4

0.00606 8.57 11.65 O. 15 99.7

0.00817 10.46 5.68 0 99.9

$22 0.00270 6.99 28.40 7.10 98.5

$23 0.00395 7.02 11.96 0.65 99.6

$25 0.00245 4.48 22.29 3.97 99.0

* Estimated values

', Notes: Emissions data for oxides of nitrogen corrected to eombustor inlet total temperature of 704 K, in.ct pressure of 9.31 arm, rcfercnce velocity of 41.8 m/s, hulnidity

of 6.29 g/kg, :rod Colnbustor exit total tclnperatare of 1413 K.

Fmissioas of carbon monoxide and total unbunletl hydrocarbons corrected to cornbuster inlet pressure of 9.31 :ltlu.

Efficiency based on carbon b:dance data.

63

t
TABLE XXVIII

VORBIX COMBUSTOR EPA PARAMETER EMISSIONS RESULTS

EPA Paralneter

(Ibm pollutant/1000 Ibf

thrust-hr/landing-take-off cycle)

"Ox-icles..................... Totai ..... SAE

of Carbon Unburned Smoke Idle

Configuration Nitrogen Monoxide Hydrocarbons Number Condition .,,

JT9D-7 (1) 5.9 8.5 3.9 4 Without bleed

Goal (2) 3.0 4.3 0.8 _ 15

Sl 1 3.0 9. l 0.3 - With bleed

S18 3.4 14.8 0.3 - With bleed

$20 3.3 9.6 I. 1 2 With bleed

$20 3.5 6.2 0.7 Without bleed

$22 2.6 12.9 0.9 1 With bleed

$22 2.7 10.3 0.9 Without bleed

$23 2.5 8.7 0.3 2 With bleed

$23 2.6 8.6 0.5 Without bleed

$25 2.2 10.4 0.3 3 With bleed

$25 2,2 6,5 0,3 Without bleed

(I) Based on initial production test results for combnstor EC 289386

(2) 1979 EPA rule levels for class T2 engines

Note: All data corrected to JT9D-7 engine conditions

64
6

5

()RIGINM_ PAGE IS

or Poor QUAI21_
-: Overall, tile best results were obtained with Configuration $25, which provided EPA parameter values for all emissions species that were s_lbstanttally lower than tbo_ of the cuFrcnt

- production JT91)-7 engine, l{elative to the JT9D-7, Configuration $25 provided a 63-percent reduction in oxides of nitrogen, _ 24-percent reduction in carbon monoxide, and a 92-

" percent reduction in tetai unburned hydrocarbons.

Correlation of the efficiency data for idle, approach, climb, and sea-level take.off revea}s an

important characteristic of the Vorbix combustor, as shown in Figure 29. At idle conditions, the combustion efficiency drops off rapidly at pilot fuel-air ratios below 0.009, but

the efficiency becomes progressively higher for a given pilot fuel-air ratio as the power level --"

is increased. This characteristic indicates that fuel scheduling between the pilot and main

: zones can be effective in minimizi_g emissions at any intemlediate power level

2 99.2

(1:

_ 99.0

98.8

I . "ua 98.6

98.4

W

Z 98.2

98.0

a3

97.9 O

¢.Z,.-

97,6 O IDLE (WITHOUT BLEED)

APPROACH (OVERALL FUEL-AIR

97.,1 RATIO = 0.0#31

0 CL'.MB(OVERALL FLIEL-AIR RATIO • O.O194)

97.2 t_ O TAKE-OFF (OVERALL FUEL-AIR

RAT(O ,, 0,0215)

97,0 , I I I I t J J

0.002 0.004 0,0(](3 0,009 0.010 0.012 0.014

PILOTFUEL-AIRRATIO

Figure29 I':fflciencyCharacteristicsIbr VorbLx"Combustor Cott]igttration $23 asa l"uttctiotzof Pilot

_me b'aeI-AtrRatio

66

.... • I I I "/ ....... "' i-- --- ....
2, PERFORMANCE RESULTS

As for the ltybrid combustor, perfornlance measurements were made during tile l'hase II

program for tile Vorbix eombustor to determine pressure loss, exit temperature pattern,

idle lean blowout limits, altitude stability and rolight characteristics, and durability.

PressureLoss

Tile measured values of combustor, diffuser, and total combustion system pressure loss are

snmmarized in Table XXX. As shown, the initial configurations of tile Vorbix combustor ""

operated with pressure losses in excess of seven percent. This high loss level resulted primarily from the high shroud losses around the combustor front end. A significant reduction

in the losses was achieved in later configurations through incorporation of a hood to streamline the flow around the Vorbix front end. The hood was retained for Configuration $25,

but an increase in the losses occurred as a result of a change in the front-end airflow schedule, since the hood was specifically sized for the airflow requirements of Configuration $23.

TABLE XXX

SUMMARY OF PRESSURE LOSS TEST RESULTS

FOR VORBIX COMBUSTOR

Overall Diffuser

Pressure Combustor and Shroud

Configuration Loss (Percent) Loss (Percent) Loss (Percent)

Goal 6.0 ....

JT9D-7 5.4 ....

SI 1 7.2 2.7 4.5

S12 7.2 2.7 4.5

S13 7.2 2.7 4.5

S14 7.2 2.7 4.5

S15 7.2 2.7 4.5

SI6 7.3 2.7 4.6

S17 7.3 2.7 4.6

SI8 6.1 2.4 3.7

SI9 6.2 2.3 3.9

$20 5.6 2.3 3.3

$21 6.0 2.8 3.2

$22 5.6 2.5 3.1

$23 4.6 2.3 2.3

$24 4.6 2.3 2.3

S25 6.9 2.5 4.4

67

- q
Exit TemperatureData

When thermocoulflC durability I)roblcms limited tile amount of temperature data available,

pattern factors were calculated oil die basis of the hottest temperature reading obtained

from the ftulctioning tbermoeouples and an average temperature calculated ou the basis of

if tile meastwed fuel-air ratio. Most of the data for high power operating conditions was obtained with a fuel-air ratio of 0.016 to limit the maximum temperatures aild promote tiler-

---" ? mocouple life.

. The circumferential temperature profiles for the Vorbix combustor at idle and sea-level ..,.

. take-off conditions are shown in Figures 30 and 31. These data show far less scatter than

i those for the ltybrid combustor, and tile temperature patterns do not appear to be influ- enced adversely by the strut wakes.

I

i ? :, -.. + _ -, ....... . -

• ++'+ CONHGURAT ON S20 AT DLE I : + ; ; I [ ."

: ' COMPRESSOR EX I"TEMPERATURE = 428K .... I_. r_J ; . ' i " + : i ,

_ r" U 1_ i COMPRESSOREXITPRE$SURE':2.93ATM I: -I !+ ,

" i! ' !! FUEL,AIR RATLO =. 0.0126 ,- L._- _ |_ ]............. _ _ , ..

-- .... ! - _ ............ ' ,---i----'.---'+r-,-_ :-:-,--,....... ,

l+_ i , • ." , i " + ! " + ,.

m ....: I t-+" " -" ':" I r- .-r-: I_! T

30(]. i - i I _ , -

• : ; i i " - ; ! . , i • . ! !

: . : : • i I ,

l _ ; , i • j - _ I

5: • . 'I . . .: .: " i ' ,

' ; " ; ! i i i i

ul ,,oo.._ c ._ ' . .I , ' i I,L :

p_ i - ; ! i : i : I I

. J I , _l.- ' " "P " = ¥ _ '

i "_ . ! ! I " _ ' i . ' i. . l.i L I i , ! " .

leo0 _I_ +J If , _ I + ", L- _.L.x' _:_ : I_i I_lq" ill I _ _ " ,

i _l_ " 1+'_L'¢_ I . 1_[ II _# ! _%.' IIill !. ]I_X ::+'"l. t . I_-." _1 FiAOIA L_' AVt_ , ,lAG . : I . " .

-i,-.+':!A+t--', . __I';L: ......-' :

• • I i

| ,,I,

!%i V I !. -- ..... K I .

....... ,., ..... .....+........

ll_- ! ! I • , i ' ;i II I t+ I ] i _ '

i i i i "] , : i I ;! : i I . , :

I I i. I : , J

__ _ ; L _ J ! i _ I _ j i ! .

: i ' _ i I _ I_ RA_<EIIEAf.} IiE_,spAt'I

-+.... '+ I - .' .-.!+.4-:-4++ : --.-_---;:+-4 : ,,,',.<.:.,,,o:,,-+,,,,+

! , • / i .. i. I _KE EAD _I_ 5PAN +0_,:. , . I i

• . i i ..... , ...... , .... I I .._+. I /'_KL IIEAD 71, SPANi i ! . r', ! I , + • : i , : • : + RAK[ IIEAO 90:t SPAN

i ! STIIOT ' '- , ! " ,_ + ! - ] , i [ . I I

__:I_.J_.__.-J___.+_..t _l: .:LI_I ;_t.._i t + IJU ....

0,7 o,4 0 6 O.q 1.0 I 7

CIRCUMFERENTIAL PO,_ITION - RADIANS

I"i_+re30 _+Trcuniferent/;ial;_itTemperalun,PatternsSot PorhLrfTmthustor('oNSTgiir.lh)$20 ,l at Idle

Comfitiotis

68
t0% SPAN'. !

RAKE tt_AO 31% SPAN_-_

RAKE HEAD 51% SPAN - 1

RAKE ItEAO 7t% SPAN -

RAK_ HEAD 90% SPAN_

.:_. !--i .!-]--!... J

• ' STRUT!.. i _ ..... . _- ]

: i / ; i : i i

I " i ] I : ; I

CIRCUMFERENTIAL POSITION -- RADIANS

Figure 31 Circumferential l_:vitTempelzttitre Patterns lot" Vorbix Cmnbustor Configztration S20 at

Sea-Level Take-OJ]'C'cmditiolls With FueI-A it*Ratio of O.016

Pattern factors are shown in Table XXXI and range from 0.42 to 0.93 at take-off conditions. Although the best configuration provided a pattern tactor equal to that of the current JT9D-7 production eombustor, it remained 68 percent above the goal, indicaling the

desirability for fllrther development.

The radial teulperaturc protile obtained at idle conditions is shown in l"igure 32. Unlike the

llybfid combustor which produced high temperatures near the inner wall because of the r:ldial displaceinent of tile pilot and lnain burner zones, the Vorbix combustor produced a

relatively unifoml profile at idle conditions. This uiliforin profile is one benefit of axial

rather than radial separation of the two colnbnstor stages,

lligh-tempcrature radi=d temperature profiles for selected Vorbix combustor configurations

are shown in Figure 33. "ll=ese plots show tinct the profile is sensitive to the configurational

changes, Several configurations provided profiles simihir to the JTgl)-7 tt_rget profile.

b9
'|'AOLJ_X_XI

SUMMARY O1+VORI_IX('OMIffJSTOR EXIT

TI{MIq{I_A'I'URIP_ATI'I:I_N FA('I'OR RI_SUI,TS

PntIcm I:aclo_"

Idle 'l'ak¢4)ff

C+mO_llralio_ _rlditions Conditions

Goal -- - 0.25

- JTgl)-7 - -- [I.42

SI i 0.2? 0.57

SI2 0,46 0.93

- _ SI3 - - 0.58

___ sl4 ........

SIS .......

SI6 0,30 0,80

: SI7 0.30 0.80

SI8 0.27 0.78

SI9 0.57 = .

$20 0.40 0,65

$21 0,39 - -

$22 0.46 0.68

_23 0.36 0.43

$24 - -- 0.42

$23 0,44 0,48

1300

1200 COMPRESSOREXIT TEMPERATURE= 428K

COMPRESSOREXIT PRESSURE=2,03 ATM

FUEL AIR RATIO = 0,0126

: _ 1100

M1

ul

o.

uJ

_ 0_

• I_

0 7OO

SIX)

+.,

L

sco I, I I I I I I I I I

10 20 30 40 50 00 70 80 90 ISO

" INNER LINER PERCENT TOTAL SPAN OF LINER OUTER LINER

l,Tgtlre32 Radiall:Mt TemperaturePatternfor YorbLrCombustor C'onfigurction,$16at IdleConditions

70

• . gl - i _ - |
lflO0

0 S II COMlqlf_!,l)llIXII TIMP[IlhltlII[ Iii._l!_

51;I (;llMl_f_£n ll R iXIT I_II[;, F _lJill  ̧ fiJIA1M f_l_

.... _ Y _i_O FIj(I AIIIIIATIO II,01Gll V

[ - NOIIMALI_EO TO (I_KI

- I_,o I I i i I I I I I I

O IO 20 30 40 50 60 70 80 90 100

PERCENT TOTAL SPAN OF LINER

- P_gz_re33 RadialExit TemperaturePatternsfor Selected Vorbix Combustor Configurationsat

- Sea-LevelTake-OffConditions

Idle Lean Blowout Limits

The Vorbix eombustoridle lean blowout data arepresentedin Table XXXII. Asshown,

all Vorbix combustor configurations exceptConfigurations SI 1 and $21 provided lean

blowout limits lower than thosefor the JT9D-7 combustor. ConfiguraUon $21 hada high

pilot swider airflow which produced a leaner pilot zone aald consequently increased tile

overall combustor fuel-air ratio at lean blowout.

I"A BI.I_ XXXll

SLIMMARY ()1: VORBIX ('OMBUSTOR II)I.E

LEAN IR_OWOLrr 1 liS I ̧ R_!SI.ILTS

Idl_

l._an SlclwouI

('oiln;_u falien Fuel.Air Ratio

Goal

J] ftl).7 < 0.004

SII O.(IOr,3

I ? O.(IO1.1

_1_( 0.0023

SI') O.Lu3I'_

_1 O.OOI,4

_2_ 0.(IO30

S_4 0t_24

$2$ II O1)2_i

N_ltllllldl ird¢l rl_ndit ii_ll_ _¢r¢ Jlll_l ple_,_lt_

71
l f

I

i

Altitude Stability and Rolight Results

The altitude stability ,'rodrclight data for the Vorbix eoml_ustDrreve,'dedthat tile characteristicsaresensitiveto tile pilot swirler configuration. "lhe initi;d resultsobtained for Configur:itions SIO ;uld $20 are shown in Fig_.tre34. 'lbe results approach tile engine requiremcn t

and comparc favorably with results fro,n rig tests for the JT9D-7 comhustor. The effect of

the pilot swirlcr inodification incorporated in Configurations $24 and S25 is shown in Figurc 35. These data indicate significantly poorer altitudc stability and rclight characteristics.

In an attempt to improve pcrforlnauce, Configuration $24 was retcsted with a high-pressure ,..

drop fucl nozzle. Results are shown in Figure 36. As shown, this change was effective in

improving the altitude stability and rclight performanec, but it did not achieve the capabilities of Configurations S19 and $20.

0.681Kg/S 0,858 KglS 1.034 Kg/S 1.249 Ko/S

10500 242 K 250 K 261K 269 K

750(

I=

u.I

I'-

ILl

IIRED JTgD ENGINE

Z RELIGHT ENVELOPE

ILl

_3

=) LEGEND

I--

I

I-- 619 $19 --I

4-JOULE 2O-JOULE

IGNITION IGNITION $20

PRESSUnE• II

BLOWOUT

UGHT 0 &

NO LIGHT _ •

(

0,2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

FLIGHT MACH NUMBER M N

lqgute 34 Altitude Stability and Relight Cbaracteristfcs of Vorbb: Combustor Configurations S19 and

$20 Sbown tm JT9D-7 Altitude Relight I:)lvelope

72
e, f,i Nil

lull rtOW furl flow Itl_LrL¢l¢l_

O_JtlO_:gl';flT O0_L Kg/$1_ 011_¢,Kgl_l C

Id#dlMtJM

'_ ////

Ira..,

.-.=./ / f 4/_ 7" ,:_x'i,,_,,,.,%/" '._,",%_%,=

fiZZ',6 /

o I I I I I

o_ G_ O4 o5 o6 &t O8 09 10

FLIGHT MACH NUMBER MN

Figure35 Altitude Stability and Relight Characteristicsof Vorbix CombustorConfigurations$24 and

$2._ 14_thModified Riot Swirler

LEGEND

• _41NIktUhIPIIESSURE RLOWOU_

0 LIGIIT

• N(_ I I(_IIT

ALL I)ATA ODTAINEOWIIII

[ UEL I:LOW 0 ̧0170 I_glSI'C

AN{) _O..JOtlLE IGNITKIN

I100( r- o.68 t Kqt$ o.Sr_B1_91$ I.O34 Kgt$ 1.249 K0/$

10_. L 24? K 250 K 71 K/269 K

Z _

_ _ r jRT0( EL } I N(IIN[

IGHt ENVELOPE

FLIGHT MACH NUMBER M N

l,Tgure36 Altittlde Stability am/Religllt Characteristicsof Vorbix ComhustorConfiguration$24 Wl'til

h_creasedPressureI)mp I_lot FuelNozzles

73

(,ICL(;INALPAG/_I,_

I,' I'(101_QUA,_
It should bc noted that retention of the good altitude and relight capabilities demonstrated

in ('onfiguratiol_s S[ 9 and $20 could require compromise of the good emissions characteristics demonstrated in the later configurations. Achievenrent of both cllaraeteristics concurrcntly is expected to require considerable development.

CombustorDurability

The Vorbix combustor experienced only minor durability problems consisting of localized

hot spots on the eombustor liners. Configuration $21, bowcver, experienced local burning

in the vicinity of the main burner premixing tubes.

D, COMBUSTOR STATUS ASSESSMENT

Summaries of pollution results for the bust performing configuration of each combustor

concept, Hybrid Configuration H7 and Vorbix Configuration $25, are contained in Figures

37, 38, and 39. JT9D production engine data are also included for comparison. A summary

of performance results.and status for both concepts are contained in Tables XXXIH and

XXXIV.

74
sol- ...... k._6,_o- -I

I

-:' |19D.7ProI)UCIION[t_GINE!

" L"vor_lcxu_uste_.sj:_L,

oo 3°i [PAP= 8.5

_ _.5

/ [PAP: 3.7

o - 2o---T--4o 60 8o"-_f6o

PERCENTTHRUST

l_Tgure38 Emission Characteristics for Carbon Monoxide Versus Percent Engine 77trust

:_taJ

_:=_ 20

,.._,e:_ EPAP= 3.S

_=:c.:_12 o JTgD-7PROOUCTtO[HGINE If ,

_-,_ _, H,BRIOCOMBUSTORH-7'

_,j ¢,_

t-- __ " _.EPAP : 0.?

kPPROACH CLIMB

PERCETHRUST ,NT

Figure 39 1"3nissi,JnCharacteristics for Total Unbtm_ed llydrocarbon_ Versus Percent Engine 7hntst

75
TABLE, XXXIII

ItYBRII) COMBUSTOR PERFORMANCE RI'SULTS

Should Meet Should Meet

Currently Mcets Requirement With Requirement With

Requirement Nonnal Development Extensive Development

Pressure Loss X

Exit q'emperatnre

Pattern Factor X

Radial Exit Temperature Profile X

Idle Stability (Lean

Blowout) X

Altitude Relight

Characteristics X

Durability X

Carbon Fonnation X

TABLE XXXIV

VORBIX COMBIISTOR PERFORMANCE RESULTS

Should Meet Should Meet

Currently Meets Requirement With Requirement With

Requirement Normal Development Extensive Developlnent

I

Pressure Loss X

- Exit "l'emperature

Pattern Factor X

Radial Exit Temperature Profile X

Idle Stability (Lean

Blowout) X

Altitude Relight

Characteristics X

Durability X

Carbon Formation X

7_
I

t

On the basis of these results, the Vorbix comhustor was deemed tile more l_rnmising design

and the one most readily adaptable to engine installation. Tiros, the Vorbix concept was

selected for additional Phase 11testing at,d for the engine testing. The best all-around corn-

- bustor configurations, Vorbix combustor Configuration $25 and ltybrid combustor Configuration }17, exhibited significant EPA parmneter emissions reductions compared to the

JT91) production combustor while maintaining smoke below the visible threshold. Both

of these combustor.configurations have been developed to the exteJ_t that either of them

could have been chosen for the Phase III engine demonstration tests.

,lit...

Deternfining factors in the Phase III selection were file generally good performance aspects of the Vorbix eombustor, including the ability te operate full staged at approach.

With the exception of file program pattern factor goal, this design currently meets, or

should meet with nonnal development, all engine performance requirements. Vorbix pattern factors, while not at the goal level of 0.25, are equiv',dent to current l-roduction JT9D-?

levels.

E. PHASE III ENGINE DESIGN

A representative cross-section of the Phase III engine design is shown in Figure 40. This

confignration is expected to produce emissions comparable to those produced by Configuration $25 in Phase 11. The Phase 11Iconfiguration closely resembled Configuration $25

except for a number of minor changes required to incorporate th¢ eombustor into the engine and facilitate installation. A common cast fuel injector support has been designed for

both the pilot and the main burner zone to reduce fabrication cost. In addition, to reduce

diffuser case distrotion, the pilot fuel injector mount pads will be clustered in groups of

three. The pads will be mounted on ten bosses welded to the diffuser case. Clustering reduces the 0Jnount of case welding required and also eliminates interference with the diffuser

case bleed bosses.

Other minor changes inciude an increase in pilot zone swirler travel to allow radial growth

and a revised cooling airflow distribution. In addition, sturctural considerations dictated revision of the mount pin arrangement and stiffening of the throat cooling air scoops to prevent flutter.

: 77
i
CHAPTER IV

FUEL CONTROL DESIGN STUDY

A. INTRODUCTION

A fuel control design study was conducted as part of Phase II to identify control system requirements added by the staged cmnbustor concepts developed in the Experimental Clean

Combustor Program. A nmubcr of conceptual designs which satisfy tile functional requirements were specified, and the inost promising concepts were selected on the basis of available

technology and estimated life cycle cost.

The study was conducted in two parts. The first addressed modification of the current .]T9D -"

fuel control with a miuimum number of changes to meet the requirements of the staged Vorbix combustor. This concept was based on 1975 technology and included specific design

work to prepare a design for a breadboard control system for Phase 111engine testing. The

breadboard engine fuel control design for Phase 1II engine testing is described in this chapter. Items relating to identification of candidate schemes, concept evaluation and concept

selection are contained in Appendix C.

In the second part of the study, new design concepts using projected 1985 technology were

explored. These study results are also contained in Appendix-C.

B. FUEL CONTROL DESIGN REQUIREMENTS

The current JT9D-7 fuel control unit is a hydromechanical device which sets engine thrust by

scheduling and distributing total engine fuel flow as a function of power lever position, highpressure rotor speed, and limiting parameters such as combustor pressure.

The combustor pressure and fucl flow envelopes corresponding to the altitude-Mach number

operational envelope of the JT9D-7 engine are shown in Figure 41. The current production

single-stage combustor system employs duplex fuel nozzles, with the cut-in point of the seeondary fuel flow occurring at a fixed value of overall fuel flow. Effectively, the staging point

is independent of overall combnstor fuel-air ratio, and the primary/secondary split can be

quite different at various engine power levels and operating altitudes. In addition to the fuel

control function, the engine control unit schedules variable compres'_,Jr stator vane angle and

compressor bleeds.

- Two-stage combustors, of the type developed in the Experimental Clean Combustor Program,

are characterized by two separate combustion zones and two physically separate sets of fuel

_ injectors and Inanifolding. Since each colnbustor zone must be operated within generally narrow limits for optimum emission formation ;rod combustion efficiency, fuel distribution to

? each zone must be based on engine fuel-air ratio rather than total fuel flow. In addition, a

number of mechanical constraints such as inaxinlunl fuel pump pressure, minimum control-

- lable flow rate, fuel nozzle turn-down ratio, manifold head effect, etc., act to further limit the

fuel control designer's freedom in varying pilot to main fuel distribution. Preliminary

-" specification of the pilot/main fuel split for the two-stage Vorbix combustor operating at

sea-!evel is shown in Figure 42. Minimum and maximum limits are imposed on the pilot zone

•:" 79
I "i

,

fucl-_drratio to prevent lean blow out and excessive thernml stresses in tile pilot zone. These

limits were developed from the Phasc 11eombustor rig testing, and define tile practical operatillge_wclopc which can bc used for pilot/n_ain zone fuel schedule optimizatioll in the czlgine

during the Phase 111effort.

8 0 X I0 !'

%

i " o I ' I I I i I J I I

! .

_. 300x to"4

! _ 2_lx 304

_ i0o;_ I0 "1

-_lE ENVELOPE

f i =.G,.E=PeeD~._,..

t !_ 1;Igure 41 JTOD L)tgine Combltstor Pressure and l,'uel Flow l:_nvelopes

lOO

BLEEDSCLOSED

ON

80 ONLY WHILE STAGING

,..I

U.. THERMAL DISTRESS

P= LINE

0

,,.I

p= NOMINAL

2 ,OPERATING LINE

_ 40

LEAN STABILITY PILOT PUEL AIR RATfo

GUIDELINE

SEA-LEVEL

0.008 0.Ol0 0.012 0. 14 0.0_6 O.OlB 0,020 0.022 0,024

I OVERALL FUEL-AIR RATIO

[ Figure 42 I')_(,I,_'chedldin$l Requirements for Two-Stage VorbL*: Combustor Based on Phase I!

I:_'l;erit/teot¢l Clean Co//thttstor Results

80

=
i

An additional requirenlent iluposed by the staged Vorbix combustor is that passage tlu'ongll

the singing point (transition from 100 percent pilot to pilot plus main zone operation) must

be accolnpJ shed in a rapid and continuons inanner. Tiffs [s required for reasons of flight

safety, and is specified by lhe I:AA Airworthiness Stand.'u'ds (Reference 4) ill tcrlns of a fivesecond maximum allowable elapsed time for engine acceleration from flight idle to 95 percent thrusL The current production JT9D-7 fuel system is fully staged at ground idle, thereby eliminating "fill time" delays associated with the volume of tile secondary fuel manifold,

distribution tubes, and fuel nozzle SUpl_orls. llowever, the Experimcntal Clean Combustor

Vorbix combustor must st:lee between the idle and approach operating conditions. Onconlpensated manifold fill time delays will seriously impact engine transient response. For this

reason, signific.'mt effort was devoted to minimizing this problem in tile definition of candidate fuel control design concepts.

C. FUEL CONTROL BREADBOARD DESIGN FOR PHASE III

ENGINE TESTING

The breadboard fuel control and distribution system for the Phase III engine demonstration

testing is shown schematically in Figure 43. This arrangement utilizes existing JT9D fuel

control componems where feasible, and retains the engine control stator vane and bleed

schedule functions intact. The main fuel pump, main fuel control, and the fuel-oil cooler

are JT9D production components used without alteration. The flow distribution valve (percent split valve) will be similar to a Hamilton Standard Division component with appropriate

modifications to meet the flow requirements of the Vorbix combustor. The percent split

valve is controlled by a reprogrammable PDP 11/40 digital computer. Overall combustor

fuel-air ratio, required as a principal control parameter, is sensed by its proportionality to

total fuel flow divided by combustor inlet pressure. Mapifolds, staging valve operation, ete,

will be functionally identical to hardware intended for eventual flight use.

........................ *'_'_................ 7;"._.... ,_:'.;_t;..;;.

::,_, i _ :'", j---t.ktb'tgltre 43 I'reliminar.l' Desig_t o[ Breadboanl I"oel £'ontml fi*r Phase III Engine Testicle

81
- The staging mechanisnl will be implemented by a set of solenoid wdvcs labclcd S, 1., and R

' in Figure 43. These wdves are dcsigl,ed to provide three modes of operation. In the slmt-

- down and start ,node,the transfer valves will circulate all main zone fuel back to the pump

_, inlet. 111tile circulation mode, the transfer valves will circulate low-pressure fuel from tile ' I

interstage pump location through the mare zone namfold ant back to the pump inlet. In

the staged mode, tile transfer valves will direct the metered flow from tile percent split valve

to tile main zone fuel injectors.

Eleven solenoid valves are provided for the pilot zone fuel system. One of these initiates

fuel flow to ten fuel injectors arranged symmetrically around the engine. The remaining ten

• * l solenoids pernut.contro of the fuel flow to tile remaining twenty fuel injectors in pairs.

_" The sixty main zone fuel injectors are controlled by sixteen solenoid valves arranged to per-

_s_ mit circumferential zoning or sequencin.g_n staggered groups of four injectors.

, Flexibility in scheduling tile start fuel flow is provided by a system that overrides the start

_ flow schedule in the production JT9D fuel control. A start-flow augmentation solenoid

valve will .)rovide additional flow to the flow distribution valve on demand. The excess fuel

not required for starting will then be diverle.d to the pump inlet.

An imporlant design objective was to provide flexibility beyond that required for a flight

engine control in order to permit optimization of the sequencing logic and scheduling during

_-- .... test and to permit investigation of alternate fuel system arrangements. A nitrogen purge sys-

: tem was provided to prevent coke formation in the main zone fuel system under hot shut-

" . down conditions. An emergency fuel shutoffvalve was included to ensure fail-safe engine

shutdown in the event of a component or logic failure.

LL

Z

: 82
f I

CHAPTER V

CONCLUDING REMARKS

Tile Phase II program successfully documented.tile potential of the tlybrid and Vorbix eombustors and provided a firm basis for selection of the configuration for the Phase llLengine

test. Both concepts provided significant improvements in elnissions relative to tile current

production JT91)-7 combustor, although neitiler fully met tile program goals. Tile Vorbix ...

combustor was selected for continnation to the Phase III engine demonstration program

primarily on the basis of its greater potential for achieving the performance goals without

excessive compronfise of the emission levels. The Hybrid combustor was deficient in both

idle stability and altitude relight capability, and required unstaged (pilot only) operation

at the approach power_.setting to achieve acceptable combustion efficiency.

When projecting tile emissions reduction technology documented in Phase II, it must be

borne in mind that the quoted emission levels have been extrapolated from data obtained

from a sector rig tested at a maximum inlet pressure of six atmospheres. Furthermore, a

number of performance items will require additional development before production engine

applications can be considered. These are summarized in Chapter III. In particular, Vorbix

Configuration $25 is deficient in altitude relight capability. Earlier configurations showed

acceptable relight characterisitcs, but tile corresponding emission levels were less favorable.

These performance deficiencies are considered to be correctable with additional developnrent, but the eventual impact on emissions has not been fully defined.

The objectives of the Phase III effort are to verify the pollutant reductions achieved in test

rigs in an engine environment and to demonstrate the suitability of the ECCP combustor

concept for engine applications. Of primary interest will be assessment of transient acceleration/deceleration operation, which cannot be adequately tested in a component rig. At

the present time, achievement of satisfactory transient operation appears to be a fuel control requirement due to the axially-staged arrangement of the Vorbix eombustor. However,

engine testing is needed for proper evaluation. A potential problem area that cannot be assessed from the Phase 11rig testing is fuel decomposition and coking in the main zone fuel

system under conditions of hot unstaging (rapid engine deceleration). The Phase 111engine

fuel control will be equipped with a nitrogen purge system to minimize tiffs problem.

Although the Phase 11Iprogram will provide a realistic demonstration of technology, it is

recognized that tiffs program will not provide a combustor configuration that is fully developed for commercial use. In particular, assessment of the long-term hardware durability

and determination of the engine operating characteristics at altitude are beyond the scope

of the current effort. However, the results obtained from the engine testing will permit an

improved estimation of the pollution reduction technology achievable in practice, as well as

the magnitude of additional development still required.

The engine test program will be conducted using an experimental JT9D engine installed in a

sea-level static test cell. The first portion of the test progrqm will consist of a minimum of

30 hours of steady-state operation, with the objective of measuring emission levels and characterizing combnstor pcrformance. The test matrix is presented in Table XXXV. This test

83

.... j
matrix is structured around tile four power settings defined by tile F,I'A for tile calculation

of the EPA Parameter. Addition:d points h:lve hecn added to fully characterize tile part-power

emissions ch;n'actcristics. Vi,'iation of the pilot/main fuel split is a primary test wn'iable at the

higher ent,'il_e power settings. The engine will be fitted with extensive first turbine wine thermocouple instrumentation during the initial steady-state testing to determine combustor exit

temperature distribution. Following determination of tile temperature profile under actual

engine operating conditions, tile therlnocouplc instruulentation will be removed to permit

testing to tile maximum engine thrust allowed by turbine durability limits. Ill..,

TABLE XXXV

STI'ADY STATli I,ERFORMAN('E/EMISSION TEST POINTS

. / "

_,lalnPowetPoinls Idl© I Approach I Climb.otlt ' Take-off

I!ngtn¢ Conditions (3(YJ-powerl ] (85% power) ( I(X_ p_wet) Se¢ondaly Power Sub- Rich Sub- Rich Sub- Sub.

Engine Conditions Idle IdLe Appreach Approach I Climb-out Take.off

Pilot/blain

Fuel Sl_gillg Poinls (2XI (2X) (2X)

12-Poinl Fixed

Pollution Samplin_ X X X X

24-Point Fixed I

Pollution Sampling X X X X (3X) X X (lX) X (lX)

: Traverse San_pling I X X X X

Twenty hours of transient testing will be conducted following documentation of pollutant

emission levels and demonstration of satisfactory combustor performance under steady-state

conditions. This testing will consist of a series of progressively more rapid engine accelerations from idle to maxinlunl allowable thrust and deceleration back to idle. The objective

_ of this testing is to assess the transient operatiomd characteristics of the eombustor and fuel

control concepts. Particular attention will be paid to passage through tile pilot/main staging

point. The Phase 111fuel control is designed to tlermit considerable adjustment in fucl

management at tile staging point. No emissions data will be acquired during the transient

portion of the engine test program.

Three exhaust gas sampling techniques will be investigated, as indicated in Table XXXV.

The principal method will involve use of an exhaust plane rake having eight, equally spaced (45

degrees apart) radial armz, each having three sampling ports located at centers of equal areas.

This rake will provkle ,'ltot:d of 24 manifolded lationary sampling ports located within the

core engine exhaust stream. Variations to be investigated include reduction to four radial

arms, spaced 90 degrees apart, for a total of 12 salnl_ling ports lind traverse of tile eight and

probe ovcr a 45 degree sector, thus covering tile engine exhaust plane at 5-degree interwds.

A secondary objective of this portion of the program will be to assess sampling tcclmiques

for representative gas sllnlple acquisition.

84
I= ' I

APPENDIX A

: COMBUSTOR CONFIGURATIONS

: _

-_: 85
I "i 4 li 6 _ BIB

o -

1 2 I 14 1o I6 17 18 1O

COOLING ROLE PAI_rERN

INNER LINER OUI"ER LINER

DIA. AREA [ OIA. _,RE._

LOUVER mX 10.3 fill 2X 10"z LOUVER mX 10.3 ROLES 'mZx 10 .4

i _:ItOLES I

1 1,2;" B5 = 1.12 11 1;,32 lib 1.51

2 1.32 85 f,17 12 1.32 110 1,El

3 1.32 85 1.17 13 1.32 110 1.51

4 1.32 BB 1,17 14 1.32 110 1,51

5 1.32 92 1,26 15 1.93 60 I.'/5

6 1.32 Bt 1,11 16 '2.16 60 2.19

7 1.32 92 1,26 17 2.31 60 2.52

B 1,32 122 1.67 18 2.36 G0 2,63

9 1.59 BB 1,G8 19 1.59 110 2.17

10 1,59 85 1,68

AREA In2 X 10.4

PILOT BURNER FLAMEF*OLOEtl 94 (e O.B0 X 10"2m DIAMETER 47.26

P_LOT BUf_NER FLAMEFIC'LDER WEEP 25Q (.o0. ) 294 X 10"2m DIAMETER 12.67

MAIN BURNER OUTER SWIRLER 11 LE, HAND SWIRLERS 43.01 /_CO (EFFECTIVE AREA) MAiN BURN[:R If_NER SW[RLER

11 f11(RIT 1lAND SWH]LEHS 34.42 ACD tEFPECTIVE AREA}

BULKHEAD COOLING 170 ['_0.2¢a4X 10"2m DIAMETER 8.61

DILUTION • LOUVER 2 (0.105 R/-O. APART O.05P DAD. 14 _o 1.913 X 10"2m DIAMETER 40.23

OFF STRUT _. :

FINWALL '_)_(INNER WALL ) 1.01% WAD (BURNER AIRF [ OW)

ClNWALL _ (OU rEl_ WALL) 1.23% WAB (11URNER AIRF LOW}

SIDEWALL COOLING 9.00% WA4 {10TAL AIR F t OW STAIION 4)

TURBINE COOLING (INNER WALL) 7.9%WA4 (1DTAL AIRFLOW STATION 4)

TURBINE COOLING IOUTE_ WAL! ) 8.4% WA4 ITOTAI. AIRFLOW STATION 4}

PILOT BURNER NOZZLE PiN ()LN 2,%'00 11.10 LOCATIONS

MAIfk BUBNER NOZZLE P/N IOW,_.P:I1 LOCATIONS

MODI[:ICATIONS

PR^SE1IHYBmUBASEINE l RE_.PiNL._O4B_

I,'/_11rc'/IJ. Hl'flrid C.t_ h ,E'_I_G,"_IZJ_I_"z/liD I//

86
: I

j

12 13 14 IS 16 17 18 19

COOLING HOLE PATTERN

INNEI_ LINER OUTEI{ LINER

LOUVE_ II DIA. _I_EA I DIA AREA n_X IR "3 ::ROLES t_2X _R"z LOUVERIrnX 10.3 ::HOLES _n2X IR "4

1.32 SB 1.17 I I 1.32 110 I._I

2 1.32 B5 1.17 12 1.32 110 1.51

3 1.32 SS 1.17 13 | 32 11R 1.51

4 1.32 SS 1.11 _4 1.32 110 1.5}

5 1.32 92 1.26 I_ 1.93 BO 1.75

6 1.32 81 1.11 16 2 16 60 2.19

] 1.32 92 1.26 17 2.31 BR 2.52

8 1 32 122 l.Sl _8 2 36 BR 2.63

9 1 5_ B5 ?.68 19 1.59 11R 2.17

10 1 59 RS 1 6B

AREA m2 X IR "4

PILOT BURNER FI.AMEHOLDER

PILOT BURNER FI.AMEHOLDER WEEP 94 @ 0.80 X I0"2m DIAMETER 47,26

MAIN BURNER OUTER EWIHLER ?.43 ACD (EFFECTIVE AREAJ

MAIN BURNER INNER SWIRLER 11 LEFT HAND 5WlRLERS 43,01 ACD (EFFECTIVE AREA}

BULKHEAD COOLING 11 LEFT HAND SWIRI.£R$ 36,46 ACD (EFFECTIVE AREA)

170@0.254 X 10"_m OIAMETER 8.61 DILUTION • LOUVER 2 (O,IR5 RAD, APART 0,052 RAD.

BEE STRUT _. ) _4 @ ?,913 X ?0"2m DIAMETER 40.23

- FINWALL_ (INNER WALl.) 1.01% WAR (BURNER AIRFLOW)

FINWAI.L (_ (OUTER WALl.) _.23% WAB (_URNER AIRF LOW)

SIDEWALl. COOLING 5.00% WA4 (TOTAL AIRFLOW -- STATION 4)

TURBINE COOLING {INNER WALl.) 7.5% WA4 (TOTAL AIRFLOW STATION 41

TURBINE COOLING (OUTER WALL) 8.4%WA4 (TOTAL AIRFLOW STATION 41

PILOT BURNER NOZZLE PIN DEN 27700-11, 10 LO_TION5

MAIN BURNER NOZZLE PIN LOW _P, 11 LOCATIONS

MOOt EICAT_ONS REF ER ENCE HI

REPLACE RIGHT HAND MAIN PURNER INNER 5WIRLERSWITH LEFT RAND _-'_IRLERS

OPEN UP SPACING ON PILOT BURNER PEAMEHOLDER HOLES

REVISE PILOt BURNER FLAME 4oLDER ATTAC tMEN _,ANO weep AREA

FigureA-2 Ilybrid Crmd_ttstfJC' r _mfig,ttroli.ll 112
t i

J

1 2 3 4 b 6 7 010

..--=____

I

[

i

c -

I1 12 13 14 15 16 17 18 19

COOl ING HOLE PATTERN

INNEH LINEll OUT[it LINFR

IlIA AREA I OIA. IAREA

. LOUVEIt m X 10.3 :: ROLES m 2 X 10 4 LOUVER m X IO "3 :: ItOLES It]2 X IO "_

1 132 B5 1.17 11 1.32 110 I.EI

2 1.32 85 I .I 7 12 1.32 110 131

3 1.32 B8 I.I 1 13 1.32 It0 1.51

4 132 R5 1J 7 14 1.32 110 1,51

5 132 92 1.20 15 1.93 60 1.75

6 ;.32 Bt I.II 16 2.16 00 2.19

2 132 92 1125 I / 2.31 613 2.5;?

8 1 32 122 1.67 1B 2.36 60 2.63

9 1.59 85 1.6B 19 1.89 110 2.17

1(] I 50 85 1.(;B

" m

• PILOT BURNER F LAMEROLDER 94 (.'_0.80 X 10"2 DIAMETER 47.26 AREA m2 X 10"4

PILOT BURNER FLAMEHOLOER WEEP 7.43 ACO (EFFECTIVE AREA}

PILOT BURNER FLAMEHOLDER COOLING ION INNER WALL] 3Bf: _ 0.21.4 ): 10"2m DIAMETER 1+92

-+ MAIN BURNER INNER SWIRLER 11 RIGHT H,",ND SW!R[ ERE 2E.34 ACO (EFFECTIVE AREA)

BULKHEAPCOOLING 170@0.254X 10"2mDIAMETER 8.01

DILUTION. LOUVER 2(0 °APART 3 ° OFF STRUT(L) 14r_1.913X t0"2mOIAMETER 40.23

DILUTION LOLJVER 2 (0.105 RA[7. APART 0.0528 RAD. 11 SLOTS 1.E24 X 0.627 X lO'"_!tt 57.63

OFF 8TRUTq )

DILUTION LOUVER 14 (IN LINE WITH MAIN BURNER 12 (_ 1.885 X 10"2m DIAMETER 33.52

NOZZLE)

DILUTION LOUVER 15 (BE1WEEN MAIN BURNER

NOZZLEI

FINWALL I_ (INNER WALL) 1.01% WAS (BURNER AIRFLOW)

FINWALL_)(OUTER WALL) 1.23% WAg (BURNER AIRFLOW)

SIDEWALL COOLING E.O0% WA4 (TOTAL AIEIFLO'H STATION 4l

"_URBINE COOLING (INNER WALL) 1.5% WA4 (TOTAl_ AIRFLOW STATION 4)

TURBINE COOLING (OUTERWAII) 8.4%WA4(TOTALAIRFLOW -STATION4!

PILOT BURNER NOZZLE -- PIN 21700.11, 10 LOCATIONS

MAIN BURNER NOZZLE P/N LOW .%P* 11 LOCATIONS

:-- mO_),F,cATIONSII.I_EEDENCeIm.

BLOCK MAIN BURNER OUTER SWIRLER

REDUCE MAIN BURNER INNER SWIRl ER AREA

INSTALL CDOUNG IIOLESON INNER LINEN TRAILING EDGE OF PILOT BURNER FLAMEHOI_DER

AI)O DtLtJTION ROLES TO OUTER I Intn ROWS 14 AND 15

F_mreA-3 I/),_EridCNmlm,_m_rmJ'igur_timl113

C , I'_
I

) P S 4 5 6 7 B }0

11 12 13 14 15 16 17 18 19

COOLING _IOI.E PA_IERN

I INNEI_ LINEII OUTEI_ LINER

! i DtA. _,REA I DIA_ 1AREA '_

_LOUVEH] mx 10.3 - IIOLE5 it2X 104 LOUVEH JmX 1E3 ::HOLES m?X 10.4

" ' 1.32 85 1.17 11 I 1.32 110 1.51

I

;_ ] 1.32 85 1.IX I? I 1.3; lid

I.S}

-¢ i, 1.32 8,5 1.17 13 I 1.32 110 1.51

I 4 I 1.32 S5 1.17 14 I 1.32 lt0 1.51

I

_ i 1.32 92 1.26 15 i 1.93 GO 1.7S

' 6 I 1.32' B1 1.11 1G I 2.16 [ 60 2.19

I

1.32 92 t.;tG 17 I 2.31 60 2.52

I _ I 1.32 122 1.61 18 I 2.36 I G0 2.63

I B j lEO- 85 }.ES )_ ! _._ } r}o 2.17

ll) q 1.59 85 1 65 d

PILOTBURNE_ FLAMEHOLGER AREAm 2x 10.4

PI LOT BURNER F LAMEHOLDER WEEP

PILOT BURNER FLAMEHOLOER COOLING )ON INNER 04 @0.DO X 1O'Em DIAMETER 47,26

MAIN BURNER INNER EWIRLER WALL)

?,43 ACD }EFFECTIVE AREA}

BULKHEAD COOLING 3B@ 0.254 X tO'Era DIAMES"_R 1.92

D_LUTION - LOUVER 2 (E_ APART. 3° OFF STRUT t_l 11 I_IGtIT 1lAND SWIRl_ERE

170 @0,264 X 10"2m DIAMETER 25,34 ACU }EFFECTIVE AREAl

DILUTION LOUVER 2 (0.105 RAD. APART - 0.06;_ RAD. 14 @ 1.St3 X 1O'2m DIAMETER 8,61 OFFSTRUTCL _ }

DILUTION - LOUVER 14 (IN LINE WITH MAIN BURNER 11 SLOTS 1,E24 X 3.627 X 10"2m 40,23 S?,63

NOZZLE)

DILUTION LOUVER 15 (BETWEEN MAIN BURNER 12 _,1t,_b X 10"2m DIAMETER 33.5Z

NOZZLE}

EINWALL(_)(INNER WALL) 1,0)% WAD }BURNER AIRFLOW}

EtNWALLE_(OUTER WALl ) 1._3% WAD (BURNER AIRFLOW}

SIDEWALt. COOLING B-00%WA4 }]OTAL AIRFLOW. STATION 41

TURBINE COOLING (INNER WALL) 7,5% WA4 (TOTAL AtRF LOW STATION 4)

TURBINE GOD9 ING (OUTER WALL) B.4% WA4 }TOTAL AIRFLOW STAT}ON 4_

'PI LOT BURNER NOZZLE P}N OLN 348130 IO LOCATIONS (SOLID SPRAY CONE)

MAIN BURNER NOZZLE P,tN LOW _P | | LOCATIONS

MODIFICATIONS REFERENCE H3

INSTAl L SOLID SPRAY GONE NOZZLES IN PILOT

Iril_,uteA-4 II;,hPM Gmthu,r;/¢_Gr m[i_uratiun ?If

89
I 2 3 4 S G 7 910

Iw

11 12 13 14 15 10 17 1B 19

COOLING ROLE PATTERR

INN£R L_NE_t OUTEfl LJNER

I DIA. ' _.REA DtA. AREA

LDUVEr_  ̧ r_Xl0 3 HOLESI,2xl0 _ LOUVER mXlO'3]::,HOLES m2X10 "4

1 1.32 _ I 1.17 11 z.32 I 0 1.51 I

2 132 H5 I 1.17 12 1 32 I 110 1.51

' 3 1.32 _ t 1.17 13 ' 1,32 ( ItO 1.51

-- - I [

.. 4 1.32 _' I 1.17 14 1.32 I rio 1.51

5 L32 o) , 1.76 15 ' I.g3 _ SO 1._'5

I I

G 1.32 _1 t 1.11 10 2.1G I GO I 2A9

; 1.32 q9 1.26 17 I I GO | 2.52

1.32 122 I.GI 1B 2.3G I 60 I 2.03

' I

9 I.so 8s I 1.G8 Io I _-Go 11o I 2.17 I

I0 1.59 BE, I 1.68

i AREA m 2 X 10-4 I i

PILOT BURNER F LAMEHOLDER 75 @0.80 X 10"-2m DIAMETER 37.709

PILOT BURNER FLAMEHOLOER WEEP 7.43AC[ 3 (EFFECTIVE AREA)

MAIN BURNER OUTER RWIRLER 11 LEFT HAND SWlRLERB 62.31 ACO (EFFECTIVE AREA)

MAfN BURNER fNNER SWIRLER 11 RIGHT HAN[3 :_IR LERS 34.42 AC[3 (EFFECTIVE AREA)

.. BULKHEAD COOLING 17.0 @0.254 X 10-2m DIAMETER B,61

[31LUTION LOUVER 2 (6o APART. 3° OFF STRUT q_) 14 @ 1.913 X tO-2rn DIAMETER 40.23

F LAMEHOLDFR COOLING (ON INNER WALL) 38 @0.254 X 10-2m DIAMETER 1,92

FI NWALLI_(INNER WALL) 1.01% WAB (BURNER AIRFLOW)

EINWALL(_)OUTER WALL) 1.23% WAS (BURNER AIRFLOW)

SIDEWALL COOLING 5.00% WA4 (TOTAL AIRFLOW STATION 4)

TURBINE COOLING (INNER WALt.) 7.B% WA4 (TOTAL AIRFLOW STATION 41

TURB(NE COOLING (OUTER WALL) 8.4% WA4 (TOTAL AIRFLOW STATION 41

PILOT BURNER NOZZLE P/N [3LN 277OO-II 10 LOCATIONS

MAIN BURNER NOZZLE - PiN LOW ,._P11 LOCATIONS = .

MODIFICATIONS REFERENCE H4

REDUCE PILOT BURNER PREMIX PASSAGE AIRFLOW

INCREASE MAIN BURNER SWIRLER AIRFLOW (REMOVE BLOCKAGE RING)

INSTALL HOLLOW SPRAY CONE PILOT BURNER NOZZLES

FIEMOVE OUTER IINER [3LLUTION HOLES

Figure A-5 Hybrid Cot bustor C'o if(qt_rat[ml If5

9O
t f I ,

1 2 3 4 5 9 7 910

f I_ FF

11 12 13 14 15 16 17 18 19

COOLING F_OLEPATTERN

INNER LINER OUTER LINER

DIA. _REA i DIA, AREA

LOUVErt mX 10 .3 :_HOLES 2X 10  ̧_ I

LOUVER mX 10 .3 HOLES m2x 10 .4

I

I f.32 85 1.17 f 32 110 L91

2 t.32 85 1.17 12 1.32 110 1.51

3 1.32 85 I. 17 13 1.32 t 10 1.51

4 1.32 85 1.17 14 1.32 110 1.51

E 1.32 92 1.26 15 1.93 60 1.75

m

E 1.32 BI I 11 16 2.16 rio 2.19

F. 7 1.32 92 1.26 17 2,31 90 2.52

8 t.32 122 t.67 18 2.36 90 2.63

! 9 85 168 19 110

1.59 1.59 2.77

i _ 10 t.59 BE 1.BB AREA m2 x 10"4

PILOT BURNER FLAMEHOLDER 94 @O,BO k fO-2 DIAMETER 47,26 ACD (EFFECTIVE AREA_

PILOT BURNER FLAMEHOLDER WEEP 7.43 ACD (EFFECTIVE AREA)

MAiNBURNEROUTERSWIRLER 11 LEFT RANDSWIRLERS 62 •31A CD(EFFECTWEAREA)

MAIN' BURNER INNER SWIR LER ( l RIGHT HAND EWIR LERS 34.42 AGO (EFFECTIVE AREA)

:. RULKHEADCOOLING 172@0.411 X 10"2m DIAMETER 22,87

FLAMEHOLDER COOLING (ON OUTER WALL) 39 @0,254 X t O_'2mDIAMETER 1.97

FLAMEROLOER CODLIN(] (ON INNER WALLI 3_ C,_0.254 X 10"-2 fit DIAMETER 1,92

"- FINWALL (_) (INNER WALL( t,01%WAB (BURNER AIRF Lr_)

FINWALL(_)(OUTER WALL) 1,25% WAB (BURNER AIRFLOW)

SIDEWALL COOLING 5,GO%WA4 (TOTAL AIRFLOW - _TAT(ON 4)

TURBINE COOLING (INNER WALL) 7,5% WA4 (TOTAL AIRFLOW -- STATION 4)

TURBINE COOLING (OUTER WALLI 8,4% WA4 (TOTAL AIRFLOW - STATION 4)

PILOF BURNER NOZZLE - PIN DEN 27700.11 tO LOCAT(ON$

MAIN BURNER NOZZLE PIN LOW AP 11 LOCATIONS

MODIF ICATIONS REF EH ENCE H5

El IMINATE INNER (.fNER DILUTION COOLING

- ADD 25% CF DILUTION AIR TO BULKHEAD COOLING

ADD 65% OF DILUTION AIR TO OUTER LINER FLAMEHOLDER COOL.ING

PRIMARY FUEl INJECTORS EXTENOEO ONE INCI( OOWNSFREAM

INCREASE PILOT BURNER PREMIX PASSAGE AIRFLOW

Figure A.6 Ilyhr_d CcHnhRSIOP CON/_gurlltk)N 116
I 3 4 r_ G 010

I1 12 13 14 15 16 17 18 19

COOLING HOLE PATTERN

INNER LINER OUTER LINER I

!

OIA. _,REA [ DIA. ' AREA

LOUVE_ mXlO "3 =HOLES n2XI0 _4 LOUVER[mX1B "3 _=HOLES mZxto'4 I

1 1.32 85 1.17 I1 I 1.32 110 1.51

I

2 1.32 85 1.17 12 i 1.32 110 1.51 I

1

1

3 1.32 85 1.17 13 1.32 110 1.51 [

i

4 1.32 85 1.17 14 I 1.32 110 1.51

!

L32 92 1.26 15 ) 1.B3 60 1.75 _

I

1.3_ 8I 1.11 16 I 2.10 60 2.19 I

I

7 I 1.32 g2 1.26 17 | 2,3_ ' 60 2,52

I I

8 1_32 122 I_E7 1S i 2.30 00 2_03

B , 1.59 85 1.E8 f9 I 1.59 I 110 2.17 J

10 !1.59 85 1.68 i j Ii

AREA m2 X 10"4

PILOT BURNER FLAMEHOLDER 94 @0.80 X t0"2m DIAMETER 47.26

PI LOT BURNER F LAMEHOLDER WEEP 7.43 ACD (EF FECTtVE AREA)

MAtN BURNER OUTER SWIRLER 11 LEFT HAND SWIRLERS 62.31 ACD (EFFECTIVE AREA)

MAIN BURNER INNER SWfRLER | 1 RIGHT RAND SWIRLERE 34.43 ACO (EFFECTIVE AREAl

BULKHEAD COOLING 172 @0.411 X 10"2m DIAMETER 22.87

FLAMEHOLOER COOLING (ON OUTER WALL) 39 @ 0.254 X 10"2m DIAMETER I .g7

FLAMEHOLDER COOLING (ON INNER WALL) 38 @0.254 X 10"2m DIAMETER "1.92

FINWALL(_(INNER WALL) 1.01%WAR (BURNER AIRFLOW)

FINWALL (_ (OUTER WALL) 1.23%WAB (BURNER AIRFLOW)

SIDEWALL COOLING 5.00%WA4 (TOTAL AIRFLOW - STATtON 4)

TURBINE C_! LING (INNER WALL) 7.E% WA4 (TOTAL AIRFLOW - BTATION 4/

TUR01NE COOLING (OUTER WALL) B.4%WA4 (TOTAL AIRFLOW - STATION 4)

PILOT BURNER NOZZLE PIN DLN 34800 10 LOCATIONS

MAIN BURNER NOZZLE PIN LOW _,P 1t LOCATIONS

MODIFICATIONS REFERENCE H6

INSTALL SOLID CONE PILOT BURNER FUEL NOZZLES

Figure A- 7 I/ybrM Comhustor Configicratio_ 117

92
I I I I ,

2 3 4 S 6 7 5 S 10 I1

COOL)NG HOLE PATTERN

INNERUHER OUTEUNER R

OtA. ARE;, Om. AREA

LOUVER reX10 "3 #HOLES m2X|0 _4 LOUVER reX10 "3 #HOLES mZxlo "4

i

1 I 1,79 65 2.11 12 1,79 110 2.73

2 1.79 85 2,11 13 1.79 110 2.73

3 1.79 110 2,73 14 1.79 130 3.23

4 2.08 110 3,75 15 2.08 130 4.43

5 1.93 65 2.48 16 1.61 99 2.02

6 1.32 85 1,17 17 1.61 99 2.02

? 1.32 85 1.} 7 18 1.32 116 1.62

8 1,32 85 1,17 19 1.32 96 1,30

9 1.32 85 1.17 20 1,32 106 1A5

10 1.32 85 1,17 21 1.79 85 2.11

11 1.32 85 1.17 22 1.32 1!O 1.50

AREA m2X1O "4

PI LOT BURNER SWIR LER 7 LEFT HAND SWIR LERS 23,49 ACD (EFFECTIVE AREAJ

PILOT BURNER SWIRLER COOLING 224 @0,140 X 10"2m OIAM 3.43

MAIN BURNER SWIRLER 28 PIN L-164377 ALL RIGHT HAND 119.23 ACO (EFFECTIVE AREA)

BULKHEAD COOLING 20 @0.282 X 1O'Zm DIAMETER 558 @0. f02 X 19"2m OIAM 5.77

FINWALL_)(INNER WALL} 1.06% WA6 (BURNER AIRFLOW)

FINWALL(_)(OUTER WALL) 1.14% WAB (BURNER AIRFLOWi

FINWALL®(Cy LINDERS) 2.26% WAS (BURNER AIRFLOW_

SIDEWALL COOLING 5.00% WA4 (TOTAL AIRFLOW -- STATION 4)

TURBINE COOLING (INNER WALU 7.6% WA4 (TOTAL AIRFLOW L STATION 4)

TURBINE COOLING (OUTER WALLI 8.4% WA4 (TOTAL AIRFLOW-- STATION 41

PILOT NOZZLE PIN DLN 27700-13 7 LOCATIONS

MAiN BURNER NOZZLE PIN DLN 27700.11 13 LOCATIONS

MODIFICATIONS

PHASE U VORBIX BASELINE - REF, PIN L,104838

FigureA-8 Vorbix Combustor Gm.F_guratioSnI 1

93
) ."t 3 4 5 6 7 B 16 '_1

INNER LINEN OUTER LINER

DIA. AREA DIA. I AREA

LOUVER I rnXI0 "3 _HOLES mZXlO "4 LOUVER rnX18 "3 nHOLEE I m2X16. 4

1 1.79 65 2.11 12 1.79 116 2.73

2 _.79 85 2. I 1 13 !.TB 110 2.73

3 1.79 118 2.73 14 1.79 138 3.23

4 2.08 I1Q 3.75 15 2.08 136 I 4.43

5 1.63 85 2.4g 16 1.61 g9 2.02

8 1.32 85 1,17 17 I 61 89 2.62

7 1.32 85 I. 17 10 1.32 1'_8 'L.62

8 1.3Z 85 1.17 19 1.32 85 1.36

9 1.:_2 e5 _.I_' _o 1.3_, 186 1.45

"--_o I 1.32 e5 1.17 _1 1.76 e5 2.11

I 11' J 1.32 85 1.17 22 1.32 116 1.50

AREA m2 X 10.4

PILOT BURNER _IRLER ? LEFT HAND _IRLERS 23,49 ACD (EFFECTIVE AREA)

PILOT BURNER SWIRLER cOOLING 224 @0.140 X 10'2m 3.43

MAIN BURNER S'WIR LER 28 ALTERNATING SWIRL DIRECTION - DIAMETER 119.28 ACD (EFFECTIVE AREA)

P/N L-104377

BULKHEAD COOLING 20 @0.ZBZ X 16"2m DIAMETER 556 @0.102 X 10"2m 5.79

OIAMETER 2.83

FINWALL_)(INNER WALL) 1.06% WA8 (BURNER AIRFLOW)

FINWALL(_(OUTER WALL) 1.14%WAB (BURNER AIRFLOW)

FINWALL(_I)(CYLINDERS) 2.26% WAB (BURNER AIRFLOW)

SIDEWALL COOLING 5.00 WA4 (TOTAL AIRFLOW - STAT(ON 4)

TURBINE COOLING (INNER WALL) 7.8%WA4 (TOTAL AIRFLOW STATION 4)

TURBINE COOLING (OUTER WALL) 8,4% WA4 (TOTAL AIRF LOW - STATION 4)

PILOT BURNER NOZZLE PIN OLN 27768-13 7 LOCATIONS

MAIF! BURNER NOZZLE P/N DLN 27700-11 13 LOCATIONS

MODIF tCATIONS REFERENCE 11

INSTALL ALTERNATE ROTATING MAIN BURNER SWIRLERS

INSTALL MAIN BURNER NOZZLE COOLING

FigureA-9 VorbixCzTtnbustarOm[iguratioNS12

_:i 94
I 2 3 4 5 B 7 8 O 10 11

11"_ t4 16 17 1 1

COOLING HOLE PATTERN

INNER LINER OUTER LINER

OIA. AREA GIA. AREA

LOUVER mXtO "3 = HOLES m2XI0 `4 LOUVER mXlO "3 _ HOLES m2XlO 4

1 1.79 85 2.11 12 1,79 110 2.73

2 1.79 85 2,11 13 1.79 110 2.73

3 1.79 110 2,73 14 1.79 130 3.23

4 2.08 110 3,75 15 2.08 130 4.43

E 1.93 E5 2.48 16 1.61 g9 2,02

6 f .32 85 t. t 7 17 f .61 g5 2.02

7 1.32 85 1.17 18 1.32 118 1.62

8 1.32 85 1.17 19 1.32 95 1,30

9 1.32 85 1,17 20 1.32 106 1.45

10 1.32 85 1.17 21 1.79 85 2,1 I

11 1.32 85 1.17 22 1.32 110 1,50

PILOT BURNER SWIRLER 7LEFTHANDSWIRLERS ; AREAm 2 X 10 `4

PILOT BURNER 5WIRLER COOLING

MAIN BURNER BWIRLER 28 ALTERNATING BWIR L DIRECTION - 224 @0'140 X tO'2m DIAMETER 23,49 ACD (EFFECTIVE AREA)

BULKHEAD COOLING 20 _6.28_ X 10"2m DIAMETER PIN L-104377 558 @0,102 X 10"2m DIAMETER 8,43

MAIN BURNER NOZZLE COOLING AIR 52 @0.254 X 10"2m DIAMETER 119.23 &CD (EFFECTIVE AREA)

FINWALL(_)(INNER WALLI 1,86% WAS (BURNER AIRFLOW) " 5.77

FINWALL®(OUTER WALL) 1.14% WAB (BURNER AIRF LOW j 2,63

FINWALL"_(CYLINOERB) 2.26% WAB (BURNER AIRFLOW)

5(DEWALL COOLING 5.00% WA4 (TOTAL AIRFLOW STATION 4)

TURBINE COOLING ( iNNER WALL) 7,5% WA4 (TOTAt AIRF LOW - STATION 4)

TURBINE COOLING (OUTER WALL:( 8.4% WA4 (TOTAL AIRFLOW - STATION 4(

PILOT BURNER NOZZLE PIN DLN 27700-13 7 LOCATIONS

MAIN BURNER NOZZLE P/N DLN 27700-11 7 LOCATIONS

MDOIFICATIONS REFERENCE S12

:==,=:=:'o.----'Q. @ x@

PILOT INJECTOR@@_ l ] @ x y _)NJECTOR

MAIN tN3ECTOR MAIN INJECTOR

l:i,_treA.IO VorbLTCombustorCoJtfigura/ion,._13

95
I 7 3 4 E B 7 R 9 10 11

"n " " - ]

COOLING HOLE PATTERN

INNER LINER OUTER LINER

DIA. AREA ErA, AREA

LOUVER mXtO '3 =HOLES mZXI0 "4 LOUVER reX10 "3 _ HOLES m2XlO "4

I 1.79 SE 2,11 12 1.79 110 2.73

" 2 1.79 85 2.1t 13 1.79 f?O 2.73

3 1.79 110 2.73 14 1.79 130 3.23

4 2.08 110 3.75 15 2.05 130 4.43

5 1,93 85 2.48 16 1,61 SB 2.02

6 1.32 85 1.17 17 1.61 99 2,02

7 1,32 85 1,17 18 1.32 118 1,52

5 1.3_ 85 1. _7 1S 1.32 SB 1.30

B 1.32 $5 I. t 7 2_1 1.32 1E6 1,4B

10 1.32 BE 1.17 21 1.79 85 2,11

__ _" II 1.32 85 1.17 22 1.32 110 1.50 104

AREA m2 X

PILOT BIIRNER SWIRLEH 7 LEFT P_AND SWIRLERS 23,49AD0 |EEFECTIVE ARBA}

PILOT BURNER SWIRLER COOLING 224 @0.140 X 10"2m DIA 3,43

MAIN BURNER SWIRLBR 28 ALTERNATING SWIR L DIRECTION - 110.23 ACE (EFFECTIVE AREAJ

BULKHEAD COOLING 20 @0.2E2 X 10"2m DIA PIN L.104377 558 @0.102 X 10"2m DIA 5,79

MAIN BURNER NOZZLE COOLING AIR 52 @0.254 X 10"2m OIA 2,63

FINWALL_)(INNER WALL) 1.06% WAS (BURNER AIRFLOW{

FINWALLI_(OUTER WALLI 1,14%WAB (BURNER AIRFLOW)

FINWALL_)ICYL}NOERS{ 2.26_ WAS ((JURNER A(RFLOWI

_ SIDEWALL CODLING 5.00% WAd (TOTAL AIRF LOW • STATION 4)

TURBINE COOLING ( INNER WALLJ 7.5% WA4 {TOTAL AIRFLOW - STATION 4)

TURBINE COOLING (OUTER WALL) 5.4% WA4 (TOTAL AIRF LOW STATION 4)

PILOT BURNER NOZZLE P/N DLN 27700.13 7 LOCATIONS

MAIN BURNER NOZZLE P/N DLN 277_0-13 7 LOCATI0NS

MODIFICATIONS REFERENCE 513

7E

WritON0,99465

OAm2R UE'

/NI

'NJEOTO

( _

RtS

_

S Q _'_ QPILOT

.. PILOTINJECTOR._,,_,_ P _ Q _4_"_ NJEcTOR

MAIN INJECTOR MAtN INJECTOR

.... Figure A-I l Vorbix QTmbustor Cm_figuratlol_$14

96
t I 2 3 4 5 6 7 B 0 tO I1

COOLING HOLE PATTERN

INNER LINER OUTER LINER

DIA. AREA DIA. I AREA

LOUVER mXtO "3 :;HOLES m2Xt0 "4 LOUVER mXlO'3 It/HOLES m2X10 4

1 1.79 85 2.11 12 1,TB 110 2.73

2 1.79 85 2.11 13 1.79 110 2.73

3 1.79 I10 2.73 14 1.79 130 3.23

4 Z 08 110 3.75 15 2.OS 130 4.43

5 1.93 85 2,48 16 1.61 99 2.02

6 1.32 85 1.17 17 1.61 99 2.02

7 1.32 85 1.17 18 1.32 118 1.62

8 1.32 85 1. I 7 19 1.32 SS 1,30

9 1.32 85 1.17 20 1.32 106 1,46

10 1.32 85 1.17 21 1.75 85 2,11

I1 1.32 85 1.17 22 1.37 lt0 1.50 { AREA m2X 10 '4

PILOT RURNER SWIRLER 7 LEFT HAND SWIHLERS 224 @0.140 X 10"2m DIAMETER

PI LOT BURNER SWIRLER COOLING 224 @0,140 x 10.2 m 558 @0.102 X 10"2m DIAMETER

MAiN BURNER SWIR LER 28 ALTERNATING SWIPL DIRECTION - 52 @0,254 X 10-2m DIAMETER

BULKHEAD COOLING 20 @0.282 X 10"2 m PIN L-104377 558 @0.102 x 10.2 m 23,40 ACD (EFFECTIVE AREA)

MAIN BURNER NOZZLE COOLING AIR 52 @ DIAMETER 0.254 x 1O'2m 3A3

EINWALL_)(INNER WAL L) 1.06% WAB (BURNER AIR FLOW} It 9.23 ACD (EFF ECTIVE AREA)

FINWALL(_(OUTER WALL) 1A4%WA6 (6URNEH AIHFLOW} 5,77

FINWALLP_(cY LINGERS) Z 26% WAS (BURNER AIRF LOW) 2,63

SIDEWALL COOLING 5.00% WA4 (TO rAL AIRF LOW STATION 4)

TURBINE COOLING(INNERWALL) 7.5%WA4 (TOTAL AIRFLOW .STATION4)

TURBINE COOLING {OUTER WALL) S.4%WA4 (TOTAL AIRFLOW STATION 4)

PILOT BURNER NOZZLE PIN DLN 77700-13 7 LOCATIONS

MAIN 6U11NER NOZZLE PIN DLN 27700-13 13 LOCATIONS

MODIFICATIONS REFERENCE S14

RUN 13 SECONDARY FUEL INJECTORS

F_ure A-12 Vorblx Combustor Co7_figumtiolJ SIS

OI_,I(I.INPA(AL;{,I;S

01,"POOR qUA[,I1.'y 97
I _ 3 4 Ii 6 7 B 10 11

- C ING HOLE PAT1ERN ---

INNER LINER OUTER LINER

DIA. AREA DIA. AREA

-- LOUVER mXt0 "3 ::HOLES m2Xl0 "4 LOUVER reX10 "3 R ROLES m2XlO 4

1 I39 85 2.11 12 1.79 110 2,73

7 1.79 85 2.11 13 1.79 110 2,73

3 1.79 110 2.73 14 1.7g 130 3,23

4 2.08 110 3.75 15 2.08 130 4.43

B 1.93 BB 2.48 10 1.61 99 2,02

6 1.32 85 1.17 17 1.61 99 2,02

7 1.32 85 1.17 18 1,32 118 1,62

8 1.32 8S 1.17 19 1.32 95 1.30

9 1.32 65 1.17 20 1,32 106 1.45

10 1.32 85 1.17 21 1.70 85 2.11

11 1.32 85 1.17 22 1.32 110 1.50 AREA m2 X 104

PILOT BURNER SWlRLER 7 LEFT HAF:O SWIRLEHB 3.29 ACO (EFFECTIVE AREA)

PI LOT BURNER SWlRLSR COOLING 224 @0,140 X 10"2m DIAM 119,23 ACD (EFFECTIVE AREA)

MAIN BURNER SWIRLER 2B ALTEHNATINO SWIRL DIRECTION - 5.77

BULKHEAD COOLING 20 @0.282 X 10"2m DIAM PIN L-104377 558 @0,102 X 10"2m DIAM 2.63

MAIN BUHNEH NOZZLE COOLING AIR 52 @O,JOB DIAMETER 0AB8

FINWALL_)(INNEH WALL) 1.06% WAB (BURNEFt AIRFLOW)

E INWALL_)(OUTER WALL) 1.14%WAB )BURNER All]FLOW)

FINWALL_)ICYLINDEHS) 2.26% WAS (BURNER AIRFLOWi

SI[)EWALL COOLING 5.00%WA4 (TOTAL AIRFLOW. STATION 4)

TUHBINECOOLINGIINNERWALL) 7.5%WA4 (TOIALAIRFLOW. SIATION4)

TURBINECOOLtNG(OUTERWALLJ S.4%WA4 )TOTAL AIRFLOW -STATION4)

PRIMARY NO=ZLE P/N OLN 277(X]43, 7 LOCATIONS

MAIN BURNER NOZZLE P/N GLN 27700-11+ 13 LOCATIONS

MAIN BURNER [31LUTION (INNER WALL) 7 @ 1,604 X tO'2rn DIAMETER

MAIN BURNER DILUTION (OUTER WALL_ 7 @ 1.664 X tO'2m DIAMETER

MODIFICATIONS REFERENCE SlS

IN&TALL ALL RIGHT HAND MAIN BURNER SWIRLERS

REDUCE PILOT BURNER SWIRLER AIRFLOW WITH BLOCKAGE RING

ADD INNEFI AND OUTER LINER DILUTION AIR (ROWS 2 AND 18}

FigHr¢A-13 Vorbix Omzbustor Con.figumticm S/6

98
3 4 {_ 6 _ S 10 11

[ t2 13 14 15 16 I1 18 19 2_) 2_22_ _

COOLING HOLE PATTERN

INNER LINER OUTER LINER

OIA. AREA OIA. AREA

LOUVER nlxlo "3 :: ItOLES m2XtU "4 LOUVER reX10 "3 _ HOLES m2XlO "4

1 1179 85 2.11 12 1.19 110 2.73

2 1.79 65 2.11 13 1,79 110 2.73

3 1,79 110 2.73 14 1.79 130 3.23

4 2.08 110 3.75 15 2.08 130 4.43

5 1.93 85 2.46 16 1.61 99 2.02

6 1.32 SS 1.17 17 1.61 , SB 2.02

1,32 85 1.11 16 1.32 I tB 1,62

8 1.32 EB I.I 7 1B 1.32 B5 1,3B

9 1.32 85 1. I 7 20 1.32 106 1.45

10 1.32 BB 1.1 _ 21 1.19 SB 2.11

I1 1.32 85 1.17 22 1.32 1!0 1.50 AREA m2 X 10 .4

PI LOT BURNER SWIRLER 7 LEFT HAND _,WIRLERS 3.20 ACO (EFFECTIVE AREA)

PILOT BURNER SWIHLEH COOLING 224 @0.140 X 10"2m SIAM 3.43

MAIN BUBNER SWIR LEH 28 ALTERNATING SWIRL DIRECTION - 119.23 ACD (EFFECTIVE AREA(

BULKHEAD COOLING 20 @ 0.282 X 10"2m DIAM P/N L.104377 558 @ 0,102 X 10"2m GIAM 5.79

MAIN 6URNEFt NOZZI E COOLING AIR 52 @0.254 X tO'Zm DIAM 2.63

FINWALL(_(INNEH WALLI 1 06% WAB (BURNER AIHF LOW)

FINWALL_(OUTER WALL) 1.14%WA6 (BURNER AIRFLOW)

FINWALL_iCYLINDEHS} 2.26% WAO I(_URNER AIRFLOWJ

SIDEWALL COOLING S.B0% WA4 (TOTAL AIRF LOW STATION 41

TURBINE COOLINGiINNERWALL) 7.5%WA4(TOTALAIRrLOW= STATION 4)

TURBINE COOLING ( OUTER WALL) 8.4%WA4tTOTALAINFLOW STATION4)

PILOT BURNER NOZZLE P/N EX 10970 7 LOCATIONS

MAIN BURNER NOZZLE P/N DLN 277EO-11 13 LOCATIONS

MAIN BURNER DILUTION (INNER WALL) 7 @ t,664 X |O'2m DIAMETER 15,22

MAIN BURNER DILUTION (OUTER WALL) 7 @ 1,664 X 10"2m DIAMETER '15.22

, MO/JIFICATIONS REFERENCE S16

INSTALL LOW AP AERATING PILOT FUEL INJECTORS

:" FigureA-I4 Vorbix CombNstorConfigurationS] 7

- 99

i I
kL30 4 B 0 7 B 9 10 12 13

.I.._] L_NE_RILUTION t_''"' --""_.\ _I:F '_II _

IF' , _,/// lI_R# WITH G,ESINJ,ECI" NE,.E ORS .._'+ INN.,ADII.ODTINDT ON DLIIOL.INEE6 _R_", [

[1"/ j_/'_, z _p._ _._4_VJ.,_ _=,_ IN LINE WITH "_.,

Ir-J ][ IfIf d

COOLING HOLE PATTERN

INNER LINER OUTER LINER

DIA. AREA DIA. AREA

LOUVER mxlo +3 @HOLES m2XlO "4 LOUVER m XIO "3 @HOLES m2XlO +4

1 1.63 84 1.74 14 2,34 84-- 3.61

2 I,B3 84 I.]4 16 1,B3 64 1.74

3 1.63 64 1.74 16 1.63 84 1.74

4 1.63 84 1.74 17 1.63 '_4 1.74

5 1.96 84 2.52 16 1.80 64 2.15

6 2.27 84 3.45 '= 16 2,08 126 4.43

7 t .93 85 2.49 _0 1,63 99 2.05

8 1.32 85 1.17 21 1,63 99 2.65

9 ! .32 85 1.17 22 1.32 118 1.62

10 1.32 65 1.17 23 1,32 95 1.30

11 1.32 65 1.17 24 1,32 1_ 1.45

m

12 1.32 65 1.17 25 1.79 65 2.11

13 1.32 8B 1.17 26 1.32 110 1.51

AREA m2 X 10"4

PILOT BURNER SWIRLER 7 LEFT HAND b_VIRLERS 10.64 ACO (EFFECTIVE AREA)

BULKHEAD COOLING 147 @0.22B X 1O'2rn DIAMETER 5,6B

MAIN BURNER NOZZLE COOLING 52 @ 0.254 X IB'Zm DIAMETER 2,63

PILOT BURNER DILUTION (INNER NALL) ? @ 1.631 X IB'2m DIAMETER 14.62

PILOT BURNER DILUTION IOUTE_'I WALL) 7 @ 1.631 X IB'Zm DIAMETER 14.62

MAIN BURNER 6WIRLERS 28 RIGHT HAND SWIRLERS 119.23 ACD (EFFECTIVE AREA)

SIDEWALL COOLING 5% WA 4 (TOTAL AIRFLOW - STATfON 41

TURBINE COOLING (INNER WALL( 7.6% WA4 (TOTAL AIRFLOW - STATION 4}

TURBINE COOLING (OUTER WALLI 8.4% WA4 (TOTAL AIRFLOW - STATION 4}

PILOT BURNER NOZZLE EX 10970 7 LOCATIONS 3.61 ACt,J (EFFECTIVE AREA;

MAIN BURNER NOZZLE DLN2770B- 11 13LOCATIONS

MAIN BURNER DILUTION (iNNER WAr, L) 7 @1.664 X 1O'2m DIAMETER 15.22

MAIN BURNER DILUTION (OUTEH WALL) 7 @ 1,664 X 1O'2m DIAMETER 15.22

MODIFICATIDNS REFERENCE S 17

INCREASE PILOT VOLUME

ADD PILOT DILUTION AIR

MODIFY PILOT SWIR LER

MODIFY PILO T COOLING

FigNre A-15 Vorhix ComhNstor Configuration S18

100
• _ : LINER DILUTION c.._p_' ",_

COOLING HOLE PATTERN

:2 (NNER LINER OUTER LINER

DIA. AREA E'.IA AREA

LOUVER taXI0 "3 @HOLES m2XtO "4 LOUVER m Xt0 "3 @HOLES m2XlO "4

1 1.63 64 1.74 14 2.34 84 3.61

2 1.63 64 1.74 15 I 63 84 1.74

3 1.63 84 1.74 16 1.63 84 1.74

4 1.67 84 1.74 17 1.63 '84 1.74

S 1.96 84 2.52 18 I.BO 84 2.16

6 2.27 84 3.45 19 2.08 130 4.43

7 1.S3 BE 2.49 _0 1.63 SS 2.05

8 1.32 85 1.17 21 1.63 99 2.05

9 1.32 65 ; .17 22 1.32 118 1.62 +4

10 1.32 SS I .I _ 23 1.32 95 1.30

11 1.32 SE 1.17 24 1.32 106 1.45

_+ 12 1,32 66 1.17 i I _6 1,79 8S 2.11

13 1.32 85 1.17 I [ 26 1.32 110 1.51

AREA m2 X 10.4

PILOT BURNER 6WIRLER 7 LEFT HAND b_IRLERS 10,64 ADD (EFFECTIVE AREA)

BULKHEAD COOLING 147 @O.2Z6 X 102m DIAMETER 5.90

MAIN BURNER NUZZLE C3OLING 52 @0.254 X t0"2m DIAMETER 2.03

PILOT BURNER t 01LUTION t.NNER WALL) 7 @ 1.631 X 10"2m DIAMETER 14.62

PI LOT BURNER _ILUTION (OUTER WALL) 7 @ 1.631 X 10"2m DIAME t'ER 14.62

MAIN BURNER SWIR LERS 28 RIGHT HAND _tRLERS 110.23 ACD (EFFECTIVE AREA(

SIDEWALL COOLING 5% WA 4 (TOTAL AIRFLOW - STATION 4|

TURBINE COOLING [INNER WALL) 7.5% WA4 (TOTAL AIRF LOW - STATION 41

TURBINE COOLING (OUTEr1WALLI 8,4% WA4 ITOTAL AIRF LOW - STATION 4)

PILOT BURNER NOZZLE DLN 27700-11, 7 LOCATIONS

MAIN BURNER NOZZLE DLN 27700-11.13 LOCATIONS

MAIN BURNER DILUTION (INNER WALL) 7 _ 1.664 X 10.2 DIAMETER 15.22

MAIN BURNER DILUTION (OUTER WALL) 7 @ 1.664 X 10.2 DIAMETER 16.22

MODIFICATIONS REFER ENCE S1S

USED PRESSURE ATOMIZING PILOT BURNER NOZZLE

b'igureA-I 6 VorbLx Combustor Configuration SI 9

, _1;\4',_ , _,_

_£,_(.'_!_J_'...¢ ;,_,__"'

L . -........... - 1
, - "1":1HOS,N INNEANDROUTELINER R I

- F._ WITH INJEC ORS c,j_jF._tjb_p.,,.` DILUTION HOLES _ k

t/// ,..! ,NL,NEW--,T%" "

INNER LINER OUTER LINER

DIA. AREA D_A. AREA

LOUVER mX10 "3 _HOLES m2XlO "4 LOUVER nl XIO "3 :_HOLE5 mZX10 A

I 1.63 84 1.74 14 2.34 84 3.61

2 1.63 B4 184 IS 1.63 B4 1.74

3 1.63 84 1.74 16 1.63 84 1.74

4 1,63 B4 1,74 1"/ 1,63 '84 1,74

5 1.96 84 2.52 18 1.80 84 2.16

8 2.27 84 3-45 1S 2.08 130 4.43

7 1.93 85 2.49 20 1.63 99 2.05

B 1.32 85 1.17 21 1.63 99 2.05

9 1.32 85 1.1? 22 1.32 118 1.62

10 1.32 B5 1.17 23 1.32 95 1.30

1112 1.32 1.32 8585 1.17 1.! 7 2246 11.79 .32 106 B6 2.1.45I I

13 1.32 85 1.17 26 1.32 110 1.51

AREA m2 X 10-4

PILOT 6UHNER SWlRLER ? LEFT HAND b"WIRLERS 24.39 ACD (EFFECTIVE AREA)

8ULKHEAO COOLING 147 @0.226 X lO'2rn DfAMETER 6.90

MAIN dUHNER NOZZLE COOLING 52 @0.254 X 1O'2rn DIAMETER 2.63

PILOT BURNER DILUTION (INNER WALL) ? _ 1.C,31 X lO'2m DIAMETER 14.62

PILOT [tURNER DILUTION (OUTER WALL) 2 @ 1.631 X 10"2m DIAMETER 14.62

MAIN BURNER SWIRLER5 28 RIGHT RAND SWlRLER5 101.16 ACD (EFFECTIVE AREA)

SIDEWALL COOLING 6% WA 4 (TOTAL AIRF LOW STATION 4)

TURBINE Cnrll ING (INNER WALLI ?.5% WA4 [TOTAL AIRF LOW - STATION 4)

TURBINE COOLING (DUTER WALL) 8.4% WA4 (TOTAL AIRFLOW STATION 4|

PILOT BURNER NOZZLE DLN 27700-11.7 LOCATIONS

MAIN SURNER NOZZLE DLN 27700-11.13 LOCATIONS

MAIN 61;RNER DILUTION (INNER WALL) ? _ 1.664 X 10"2m DIAMETER 15.22

MAIN 8UHNE_ DILUTION [OUTER WALL) 7 @ 1.664 X lO'2m OIAMETEN 15.22

MODIFICATIONS REFERENCE S 19

INCREASE PILOT BURNER 8WIRLER AIR FLOW*

DECREASE MAIN BURNER 5WIRLER AIR FLOW

Fi_qtreA-17 VorhLvCombustorOmJ_gumtioN_20

102
" iINNERANO

_ HOLES IN LINE INNER AND OUTER LINE'R

,- W,TR,N,ECTORB __

_ J_//_j_._ "_ "- _ _ IN LINE WITH

19 21 22 23 2425_;_T_

COOUNGROLE_ATTERN

,NNERLINER OUTERL,NER

DIA. AREA DIA. AREA

LOUVER mXlO "3 @HOLES mZXl0 "4 LOUVER mXlO "3 #HOLES n12X10"4

1 i.63 84 1.74 14 2.34 84 3.61

2 1.63 84 1.74 15 1.63 84 1.74

3 1.63 $4 1.74 16 1 63 84 1.74

4 1.63 84 1.74 17 1.63 84 134

5 I.gB B4 2.52 18 1,80 84 2.1E

6 2.27 84 3.45 19 2.08 130 4.43

? 1.93 85 2.49 20 1.63 99 2.05

8 1.32 R5 1.17 21 1.63 99 2,U5

6 1.32 85 1.17 22 1.32 118 1.62

10 1.32 85 1.17 23 1.32 95 1.30

11 1.32 85 1.17 24 1.32 106 1.45

12 1.32 85 1.17 25 1.79 85 2.11

13 1.32 85 1.17 26 1.32 110 1.51

AREA mZx10 "4

PILOT BURNER SWIRLER 7 DUAL CO-ROTATING L H SWIRLERE 29,81 ACU (EFFECTIVE AREA)

BULKHEAD COOLING 147 @0.226 X 10"2m DIAMETEH 5,90

MAIN BURNER NOZZLE COOLING 52 @0.254 X 10"2m DIAMETER 2.63

PILOT BURNER DILUTION (INNER WALLI 7 @1,63t X 10"2m DIAMETER 14,62

PILOT BURNER DILUTION (OUTER WALL) 7 @1.63 X 10"Zm DIAMETER 14.62

MAIN BURNER SWIRLERS 28 RIGHT HAND SWIRLERS 94.21 ACO (EFFECTIVE AREA(

SIDEWA LL COOLING 5% WA4 (TOTA L AIRF LOW - STATION 4| 11.52 ACD (EFFECTIVE AREA}

TURBINE COOLING (INNER WALL) 7,5% WA4 (TOTAL AIRFLOW - STATION 4 16.91 ACO (EFFECTIVE AREA)

TURBINE COOI ING (OUTER WALL) 8.4% WA4 (TOTAL AIRFLOW - STATION 4 15,04 ACO (EFFECTIV6 AREA)

PILOT BURNER NOZZLE DLN 27700-1 I, 7 LOCATIONS

MAIN BURNER NOZZLE DLN 27700-1 I, 13 LOCATIONS

MAIN BURNER DILUTION (INNER WALLI 7 @1.664 X 1O'2m DIAMETER 2,359

MAIN BURNER DILUTION (OUTER WALL) 7 @1,664 X 1O'2m DIAMETER 2,359

MOOIF ICATIONS REF MODI FICATIONS REF. S 20

INCREASE PILOT BURNER SWIRLER AIR FLOW THROUGH THE USE OF LARGER BWIRLER

- INCREASE BLOCKAGE RINGS (TO 0.348 X 1B'2m INCH W{DTH} ON OUTER LINER MAiN BURNER SWIRLERS

ADDED CARBURETOR TUBES TO MAIN BURNER FUEL SYSTEM

ADDED TOROIDAL DEFLECTOR TO PILOT BURNER SWIRLER

FigureA-18 Vorbix Combustor CbN[iguration$21

103

iii
, I I 1 ,

1 2 3 d E 6 7 S 9 10 11 t213:

14-t B,, ,B ,S 2B21Z,'3

COOLING HOLE PATTERN

INNER LINER OUTER LINER

DIA. *_.REA DIA. AREA

LOUVER mXtO "3 _HOLES m2X10 "4 LOUVER mXtO "3 @HOLES m2Xl0 "4

1 i .53 B4 1.74 14 2.34 84 3.61

1.B3 84 1,74 1S 1.S3 84 1.74

3 1.63 84 1.74 16 1.63 84 1.74

4 1.63 84 1 74 17 1,63 '84 1.74

5 1.96 84 2.52 18 1.80 84 2.15

6 2.27 84 3.45 19 2.08 130 4.43

7 1.03 85 2.49 20 1.63 99 2.05

8 1.32 85 1.17 21 1.63 BB 2.05

9 1.32 85 1.17 22 1,32 118 1,62

10 1 32 85 1.17 23 1.32 95 1.30

1I 1.32 85 1.17 24 1.32 106 1.45

12 1.32 85 1.17 25 1.79 BE 2.1 I

3 1.32 B5 1. I 7 26 1.32 110 1,51

AREA m2 X 10-4

PILOT BURNER SWIRLER (INCLUDING SLOTS IN 7 LEFTHAND SWIRLERS 27,46 ACD (EFFECTIVE AREA( CENTER TUBE OF SWlRLER}

BULKHEAD COOLING 140 @0.234 X 10"2m DIAMETER 6.09

MAIN BURNER NOZZLE COOLING 52 @0,254 X 10"2m DIAMETER 2,63

PILOT BURNER DILUTION (INNER WALL ROW 1) 7 @ 1.63 X 10"2m DIAMETER 14.62

PILOT BURNER DILUTION (OUTER WALL ROW 14) 7 @ 1.63 X 10"2m DIAMETER 14.62

MAIN BURNER BWIRLERS 28 RIGHT HAND EWIRLERS

_ SIDEWALL COOLING 5% WA4 (TOTAL AIRFLOW -- STATION 4|

TURBINE COOLING (INNER WALL( 7.5% WA4 (TOTAL AIRFLOW -- STATION 4)

TURBINE COOLING (OUTER WALL( 8.4,%WA4 (_OTAL AIRFLOW - STATION 4)

PILOT BURNER NOZZLE DLN 27700-13, 7 LOCATIONS

MAIN BURNER NOZZLE DLN 27700-11, 13 LOCATIONS

MAIN BURNER DILUTION OUTER WALl 15.22

:- MAIN BURNER DILUTION INNER WALL 15.22

MODIFICATIONS REFERENCE S21

INSTALL HOOD

INSTALL NEW SWIRLER TORROIDAL DEFLECTOR WITH 3.3 X 10"2m DIAMETER ROLE.

.... ADD OUTER LINER SCOOP

REVISE BULKHEAD WITH COOLING AIR EN FERING TitROUGR RING CONCENTRIC WITH SWIRLER.

-* ADD TEMPERATURE-SENSITIVE PAINT ON LINER (INSIDE AND OUT)

REMOVE PREMIXING TUBE FROM MAIN SURNER.

-. l,'igNrcA-I9 I/_rhixCombt_sto_mr JigNratioSt_22 104

L r,
1 2 3 4 B S ? B 9 10 11 13

"" L.L

COOUNG HOLE PATTERN ,NNERL,NER OUTERL,NER

CIA, AREA DIA. AREA

LOUVER reX10 "3 _HULES m2Xl0 "4 LOUVER mXlB "3 #HOLES m2Xl0 _

I L63 84 1.74 14 2.34 84 3,61

2 1.63 84 1.74 IB 1.B3 84 1.74

3 1,63 84 1.74 16 1.63 84 1.74

4 1,63 84 I.?4 17 1.63 _4 1.74

S 1,56 84 2.52 18 1.8B 84 2,16

6 2.27 84 3.45 19 2.08 13B 4,43

7 1,93 BE 2.49 20 1.63 Sg 2,B5

8 1,32 85 1.17 21 1.63 99 2,B5

S 1.32 85 1.17 22 1.32 11B 1.62

10 1.32 85 1,17 23 1,32 95 1.30

11 1.32 E5 1.17 24 1.32 1C_ 1.4B

12 1.32 65 1.17 25 1.79 BE 2.11

13 1,32 85 1.17 26 1.32 11B 1.El

AREA m2 X 1B"4

PILOT BURNER SWIRLER (INCLUDING SLOTB ? LEFTHAND SWtRLERS 27.46 ACO (EFFECTIVE AREA)

IN CENTER TUBE OF BWlRLER)

BULKHEAO COOLING 140@ D,Z34 X IO'2m DJAMETER B.09

MAIN BURNER NOZZLE COOLING 52 @0,254 X 10"2m DIAMETER 2,63

PILOT BURNER DILUTION (INNER WALL ROW t) 7 @ 1.63 X 10"2m DIAMETER 14,62

PILOT BURNER DILUTION (OUTER WALL ROW 14) 7 @1,63 X 10"2m DIAMETER 14.$2

MAIN BURNER SWIRLERS 28 RIGHTHAND BWIRLERS 81.48 ACD (EFFECTIVE AREA)

SIDEWALL COOLING 5% WA4 (TOTAL AIRFLOW - STATION 4)

TURBINE COOLING [INNER WALL) 7.5% WA4 )TOTAL AIRFLOW _ STATION 4)

TURBINE COOLING (OUTER WALL) 8,4% WA4 (TOTAL AIRFLOW - STATION 4)

PILOT BURNER NOZZLE DLN 27700-13, 7 LOCATIONS

MAIN BURNER NOZZLE DLN 27700-11, 13 LOCATIONS

MAIN BURNER DILUTION OUTER WALL LOUVER 22 15,22

MAIN BURNER DILUTION OUTER WALL LOUVER 21 25.74

MODIFICATIONS REFERENCE S22

AD00,457 X 10"2mWIDE BLOCKAGE RING TO INNER LINER MAIN BURNER EWlRLER.

REMOVE INNER LINER MAIN BURNER DILUTION HOLES,

ADO A ROWOF DILUTION HOLES (I,?BSX 10"2m DIAMETER) BEHIND EACH MAIN BURNER FUEL NOZZLE (13) IN LOUVER NUMBER 22.

ADD A DILUTION HOLE (I.537 X 1O'Zm DIAMETER) BETWEEN EACH EXISTING HOLE IN LOUVER 23 AND CAPPED IT.

EXTENDED INNER LOUVER SCOOP (6,868 X 1O2m DIAMETER) AND OUTER LOUVER SCOOP (2,794 x 10"2m DIAMETER) UPSTREAM,

TRIP RAMP ADDED TO OUTER LOUVER.

I:igl_reA.20 V_rbL_ G_mbRslr_rConfiguration $2.?

ORI(AINALPAOt'_I.B 105

OP pOOl{,qUAL['['¥
I 2 3 4 6 8 7 8 9 IO I( 12 13

_

ll"llI l

14 ,6,, ,s,B =. 4dj2

COOLING HOLE PATTERN

INNER LINER OUTER LINER

DIA. AR EA DIA. AREA

LOUVER mXlO "3 #HOLES m2XI0 "4 LOUVER reX10 "3 _HOLES m2XI0 "4

1 _.63 84 1.74 14 2.34 84 3.61

2 1.63 B4 1.74 15 1.63 84 1.74

3 1.63 84 1.74 16 1.63 84 1.74

4 1.63 84 1.74 17 1.63 _4 1.74

5 1.96 84 2.52 18 1.80 84 2.15

6 2.27 84 3.45 19 2.08 130 4.43

7 1.93 S5 2.49 20 1.63 99 2.05

8 1.32 85 1.17 21 1.63 99 2.05

9 1.3Z E5 1. I 7 22 1.32 118 1.62

10 1.32 85 1.17 23 1.32 95 1.30

11 1.32 85 1.17 24 1.32 106 1.45

12 1.32 85 1.17 25 1.79 66 3.11

13 1.32 85 1.17 26 1.32 110 1.51

AREA m2 X 10.4

PILOT BURNER SWIRLER (INCLUDING SLOTS 7 LEFTHAND SWIRLER6 27.458 ACO (EFFECTIVE AREA)

iN OUTER TUBE OF SWIRLER)

BULKHEAD COOLING 140 @0.234 X 10"Zm DIAMETER 5.897

MAIN BURNER NOZZLE COOLING 62 _ 0.254 X 10"2m DIAMETER 2.$32

PILOT BURNER DILUTION (INNER WALL ROW 1) 7 @1.631 X IS'2m DIAMETER 14.619

PILOT BURNER DILUTION (OUTER WALL ROW 14) 7 @1.631 X 10"2m DIAMETER 14.619

MAIN BURNER SWlRLERS 28 RIGHTHAND SWIRLER$

SIDEWALL COOL1NG 6% WA4 (TOTAL AIRFLOW - STATION 4)

TURBINE COOLING IINNER WALL) 7,5% WA4 (TOTAL AIRFLOW - STA'FI ON 4)

TURBINE COOLING (OUTER WALL) 8,4% WA4 (TOTAL AIRFLOW - STATION 4)

PILOT BURNER NOZZLE DLN 27700-13, 7 LOCATIONS

MAIN BURNER NOZZLE DLN 2'700-11, 13 LOCATIONS

MAIN BURNER DILUTION OUTER WALL LOUVER 22 11.877

MAIN BURNER DILUTION OUTER WALL LOUVER 21 15._1

MODIFICATION REFERENCE 623

CAP EACH DILUTION HOLE IN LOUVER 22 BEHIND EACH PILOT BURNER NOZZLE AND OPEN ONE 1.588 X |O'2m DIAMETER

BEHIND EACH MAiN BURNER NOZZLE.

FigureA-21 Forbix Gmlbustor CoNfiguration$24

106
i r t x i

1 2 3 4 5 6 7 S 9 10 11 1213

I k .... J

INNER LINER OUTER LINER

DIA. AREA OIA. AREA

LOUVER mXtO "3 _HOLES m2Xt0 "4 LOUVER mX1O "3 #HOLES m2X10 JI

1 1.63 84 1.74 14 2,34 84 3.61

2 1.63 84 1.74 IS 1.63 84 1.74

3 1.63 84 1.74 16 1.63 $4 1.74

4 1.63 84 1.74 17 1.63 '84 1.74

5 1.96 84 2.52 18 1.80 84 2.1E

6 2.27 84 3.45 19 2,06 130 4,43

7 1.93 E5 2.49 20 1.63 gS 2.05

8 1.32 BE 1.17 21 1.63 99 2,05

9 1,32 85 1.f 7 Z2 1.32 118 1.62

10 1.32 85 1.17 23 1.32 05 1.30

11 1.32 85 I. 17 24 1.32 106 1.45

12 1,32 85 1.17 25 1.79 85 2.1 I

13 1.33 85 1.17 26 1,32 110 1,51

AREA m2 K 10.4

PILOT BURNER 5WIRLER 7 LEFT HAND SWIRLERS 21,968 ACD (EFFECTIVE AREA)

'_ BULKHEAD COOLING 140 @0.234 X 10"2m DIAMETER 6.09

MAIN BURNER NOZZLE COOLING 52 @0,254 X 10"2m DIAMETER 2.63

PILOT BURNER DILUTION (INNER WALL ROW 1} 7 @ 1.63 X tO'2m DIAMETER 14.62

PILOT BURNER DILUTION (OUTER WALL ROW 14} ? @ 2,06 X 10-2m DIAMETER 23.50

MAIN BURNER EWIRLERS 28 RIGHTHAND _NIRLERS

SIDEWALL COOLING 5% WA4 (TOTAL AIRFLOW - STATION 4)

TURBINE COOLING (INNER WALL) 7.5% WA4 (TOTAL AIRFLOW - STATION 4)

TURBINE COOLING (OUTER WALL) 8.4% WA4 (TOTAL AIRFLOW - STATION 4)

"_ PILOT EURNER NOZZLE DLN 27700-11 ? LOCATIONS

MAIN EURNER NOZZLE OLN 27;'00.11, 13 LOCATIONS

MAIN BURNER DILUTION OUTER WALL LOUVER 22 7,36

MAIN BURNER DILUTION OUTER WALL LOUVER 21 15.96

MAIN BURNER DILUTION OUTER WALL LOUVER 20 3.45

MODIFICATIONS REFERENCE $24

REDUCE PILOT BURNER SWtRLER AIRFLOW BY _0% (BLOCKAGE RING WIDTH 0.127 X 10"2m)

INCREASED OUTER LINER PILOT BURNER Dt LUTION TO MAINTAIN PILOT BURNER EQUIVALENCE RATIO

-- REDUCE INNER LINER MAIN BURNER SWIRLSR EFFECTIVE AREA BY ONE THIRD.

ADD ONE 1.fi88 X 10"2m DIAMETER HOLE tN LOUVER 20 BEHIND EACH STRUT (2},

FigNreA-22 VGrbixCombRstorConfig_mtioN$25

107
APPENDIXB

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APPENDIX C

FUEL CONTROL DESIGN STUDY

133
APPENDIXC

FUEL CONTROLDESIGNSTUDY

A....TECHNICALDISCUSSION

1. FUEL CONTROLDESIGNMODIFIEDFROMPRODUCTIONCONTROL

a. Identification of Candidate Schemes

Utilization of the existing engine control (including ft'.e!control), insofar as possible, will

minimize development cost and impact on the engine installation. This approach confines

changes to tile addition of a new flow distribution control system to distribute fuel to the

pilot and main zone manifolds. The flow distribution control system will replace the existing

pressurization and dump (P&D) valve, which now performs the fuel staging function based

on total fuel flow. The system isshown schematically in Figure 43. By leaving the engine

control intact, the engine fuel flow vs. thrust characteristic, and the assorted trims and biases,

remain unchanged. All other engine control functions such as bleed operation and high-pressure compressor vaue angle scheduling continue to be accomplished by the production control.

EXISTINGENGINECONTROL NEWHARDWARE

MAiNCONTROLSYSTEM FLOWDISTRIBUTION PILOT

SCHEDULESENGINE METERED CONTROLSYSTEM MANIFOLD FUELFLOW

FUELFLOW FUELFLOW • PROPORTIONS TOPILOTZONE

IN RESPONSETO "_- PILOTAND

i" POWERREQUEST MAINFUEL

.. ANDLIMITING FLOW

PARAMETER¢,I @ FILLSANDTURNS 4

CONTROLSVARIABLE ONMAIN ZONE MAIN FUELFLOW

VANESANDBLEEDc, MANIFOLD5 MANIFOLD TOMAINZONE

I,)'gure C-1 Com'ept 3hxlified From Productiott Cotltrol System

134
Candidate schelues were proposed to provide fuel distribution and staging between tile pilot

and main zones based on combustor fllel.air ratio, and to fill the main zone manifold prior to

staging. Fuel pump tnodifications were also evalt.ated since some schemes for meeting the

main zone manifold filling requirements and fuel distribution reqnired additional fuel pumping capability. The candidate schemes were defined by a matrix containing various approaches for fuel distribution, main-zone manifold filling and fuel pumping. This matrix is shown

in Table C-I, and the schetnatic diagrams for each of the schemes are presented in Section B

of this Appendix. All conceptual designs are technically feasible and all are capable of providing the nominal pilot-main distribution schedule within an accuracy band of +-2percent. ,.

TABLE <2-1

MATRIX FOR CANDIDATE CONTROL SCHEMES

FOR MODU:IED CONTROL

Scheme Number

1"I91) lb le ldle '!_f.llg ! Ih 2a 2h 2c 2d 2¢ ' 2f 3a 3b 3¢ 3d 3_

Flow Distribution

I:lowDistrihution Valve X X X X I I X X X X X X X X

Two Metering Valves X X : X X X X

Manifold FlU Technique

Fill Detector X X X X X X X

Cireulatioli

Prom InterstaBe of Main Pump X X X

From I)ischaq_e of blain Punlp X X X

Fronl Pilot Burner Fuel Line X X X

Itrom ]li_dl-Pressure Auxiliary Pump X

From Low-Pressure Auxiliary I'ump X

Pulnps t

One I_.nllp X X X X X X X X

One ntis Singlc-Slroke Mallifold Pump + j X

Two Pumps X I X X X X

Variable I)isplaccnxent Pump i I X X X X X

Two nlethods of flow distributiou were considered. In one, the total flow is passed through a

flow distribution valve which splits the flow in tile proper proportions to the pilot and main

zones. In the second, the fuel flows to tile pilot and main zones are separately metered

throngh two i_idependcnt valves.

Two basic approachcs were considered for filling thc main zone m-'lnifold prior to staging.

One of thcsc involves the use of a single-stroke fucl pump. When fuel is initially directed to

the inail_ zone, the single-stroke fuel punlp wouhl bc actuated by the inain zone fuel pressure,

providing a burst of fncl equal to the inanifold fill rcquirenlent,:. The second approach for

filling the inain inanifold involved fuel circulation through the systeln. Several different configurations were considered: systetns that operated with a filll-manifold detector to avoid the

need ror continuous circulation; the use of continttous circulation systctns operating between

eithcr the nlain putnp discharge and interstage Iocatiou or between tile nlaln punlp interst_,ge

location and inlcl; the use of continuous circulation systelns using auxiliary pmnps in either

thc inlet or discharge circuhttion lincs; aud the use of continuou circuhltion hled from the

pilot zone fuel supply line.

135
t I

The various pump options resulted prinlarily from the options selected for flow distribution

or main zone manifold filling. However, a variable displacement pump was also considered

indcpendcntly to determine if the simplification achieved in the control logic would offset

the increased cost of the variable displacement pump.

In addition to Schemes I b through 3f shown in Section B of this Appendix, a family of

schemes was studied to evaluate the possibility of adapting the existing JT9D-7 pressurizing

and dump valve to the distribution control system required for the Vorbix combustor. Preliminary studies revealed, however, that this system could not meet tile pilot-main distribu- ,.-

tion accuracy requirements and detailed evaluations ofabese concepts were not pursued.

b. Concept Evaluation

The schemes were evaluated initially on the basis of life cycle cost and those schemes with

excessive cost were eliminated from further consideration. The remaining schemes were then

evaluated on the basis of complexity of modifications required to tile production JT9D-7 engine control, the availability of the required t*chnology, and potential operational problems.

Each scheme was evaluated in both an all-hydromechanical configuration and in a hybrid hydromechanical and electronic configuration. The choice between these configurations was

made on the basis of life cycle cost.

Life cycle costs were estimated using the equation:

LCC = [NHP H + NEPEI + |MHC H + MECE] + F [wH + WEI

where;

LCC = Estimated life cycle cost

N Number of units required including spares

P = Cost of each unit

M = Nmnber of maintenance actions required during the life of the aircraft

C = Average cost of each maintenance action

F -- Fuel cost to carry one additional pound of weight during the life of the

aircraft

W = Control system weight

and tile subscripts:

|l = Relates to hydromechanical elements

E = Relates to electronic elements

136
Factors for the life cycle cost equation were determined by estimating production cost,

maintenance requirements, and weight for each of the candidate schemes. Where possible,

the actual costs, maintenance requirements, and weights of added items were used directly

in tile equation. When specific data were not available, correlations were used, as shown in

b]gures 0-2 and C-3,

++ _ lOG

g_

"++

o I I I I I

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_" IE,O--

_ I01)--

_.. _=1_

- m_8

o , I I I I I

2_ 40 60 81) IIXl

PERCENT OF NUMSE I_ OF INPUTS AND OUTPUTS

7

I'lgure02 CDrlelatJonofs Reliabiflty, Weight,aridCostfiJrI(vdromechanicalCont_d Elements

137
I I

I

_.2 elh,

_Q

ZlZ

ICe

I

'_ 40 IIO 80 100

F_RC_NT OF NUMBER OF INPUTS AND OUTPUTi

Plgure G3 Correlations of Reliability, Weight, atul Cost/*or Electronic Control Elements

The relative production costs, maintenance costs, and weights are summarized in Figures C.4,

G5 and (2-6, respectively, qllc resulting life cycle cost comparison is shown in Figure (2-7. Before proceeding to concept selection, the sensitivity of the life cycle cost estimates to errors

in the assumptions was determined by introducing 20-percent reductions into each of tlle factors in the life cycle cost equation for each of the schemes. The lowest life cycle cost estimate achieved by reducing one factor in the equation by 20 percent was then used with the

initial estimate to obtain life cycle cost estimate bands for each scheme. In no case did this

cause a change in the overall life cycle cost trends.

138
/

j . [

SCHEME NUMBER

l_lgureC-4 Relative Estimated Production Costs of Candidate Fuel Control Schemes

SCHEME NUMBER

l'ig_re C'-5 Relattve I',stin ated 31a't tet a t e Req r ren el ts of Candtdate I, _el('o tro/ St' en es

139
SCHEME NUMBER

"r

Figure C.6 Relative Estimated Weights of Cm_didate Fuel Control Schemes

Ld, !

_HI!ME NUMBER

, _ l'Tgure C-7 Relative I:3timated /,ire C.vcle Costs of CandMate I,iwl Contnff Schemes

:_ 140
c. Concept Selection

Concept selection was based on estimated life cycle cost, imp,'lct oil tile existing engine control system, and critical tec to ogy items, as sunmlarized in Figure C-8. The family of

schclnes designated 3a throngh 3f were eliminated iron1 further consideration on tile basis of

the life cycle cost estimales. These schemes all included the wlriable displacement pump, and

the high resulting life cycle cost indicated that this al_Pl'OaCh did not provide sufficient reduction in tile control logic requirements to offset tile high procurcm,_nt cost of tile pu up,

MOIREll_&N FILl OElt{;lOfl CIII(2LtI-A1IOU OF BiS_ OF I_'I

[ ;¢CE$SIV£t ¥ I*MI_IMUhtOtt_RGE 5 1ECIINOI_i¢ i$ MEIEIII:O FLOPt t, CCE,*I-ABI[

$CH[ME COSTLYOtJ_SIDE *O I'q15 _fCOi_/EAK tO _kltm£D TO CAt._E _Vai, EMS

NOXtOEn el' THE iltt I OF MA| L_IAQ. SEi,astl IVI i_¢ 0#,NO $U$11r ¥' $1tART_'ROItt LM[_

SySl Ft,r"

Zb X HYBR[OOFt

_e .t.'_ H'fORO

• 14YB_IDOP

tO }tVORO

t_ X

It X

*t X

ta X

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x

x

x

x

x

x

x

x

if,

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It

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L*I:IEflMI_$ TM_OO_NTI _'VAND OIS'[_tBUIIO_ OF FtJE• EXtSTING

IN A MANIfOLD BEIOI1_ 1}t_ MA_ffOI.O I$ TLtII_D ON.

Plgure C-8 Summary of Fuel Control Selection Process Results

The concepts requiring two metering valves (Schemes I f through li) and the concepts requiring two pumps (Schemes 2a through 2e) were eliminated on the basis of tile number of

changes required to the current JT9D-7 engine control. Schemes lb and 2f were then eliminated because of their rcquirelnent for a fnll manifold detector, and it appears that considerable development may be required before a satisfactory detector can be obtained. Scheme

le uses a metered circulation system in which fuel lbr main zone manifold filling and circulation is bled from the pilot zone fi_el flow. Although this concept appears technically feasible,

it could cause starting problems, and faihlre of the valve controlling the circulation flow

could starve the pilot zone of fuel. Since Schemes lc and Id provided all the benefits of

Scheme l e without being snbjcct to the potential problems of the metered circtdation system, Scheme le was eliminated :lntl Scllenles lc and Id were identified as the most promising systems.

141
Tile final ste ) i 1 tl c evaluation was to determine tile merits of tile hybrkl hydromechanicalelectronic systems relative to tile :d! Ilydromcehanie:ll systems for the two remaining candidate schemes. Figure C-7 indicates that the hybrid designs can be produced and operated for

lower life cycle cost. The sigllificancc' of tills result depends on tile scnsihvity of the Ffc

cycle cost estimate to clumgcs in tl_c assulnl_tions upon wllicl: tile estimate was based. Examination of each factor in the life cycle cost equation shows that the relative rmlking is

most seusitivc to procurement costs. Parity is rcachcd If either the hydromechanical t nit becomes 30 percent cheaper or tile hybrid unit becomes approximately 30 percent more expensive. All other factors require unreasonably large changes (or cannot physically be

changed enougb) to reverse the order of ranking.

In conclusion, both Schemes lc and ld appear acceptable in their hybrid configurations. The

final decision between these two would require a detailed design analysis beyond the scope

of the current program. Both systems are essentially the same except for the pump connections. Scheme lc, shown in Figure C-9, was selected as the model for the breadboard design

of the Phase Ill engine fuel control.

_tAl_ _,tAIN

_t/tt Etltt _U[ t OI$1IttBUTION

FU_P _IflOt VALVh _P, DILLIA Pfl[$$UH£

1

PILOI

V^vt F B_I_ER

FUEL _

_ F_I_I. BVFASI_

l_*lt _ip_ _I1KSSNII[0

J

l,Ig_re C:9 Schematic Diagram of Fuel Control Scheme lc

142
I l

!, " I

2. CONCEPTUAL DESIGNOF.FUEL CONTROL USING 1985 TECHNOLOGY

a. Fact Control Design Requirements

The fuel control design study for the Vorbix combustor was conducted in a manner parallel

to that discussed in Section IV-B of the basic document. The design requirements were reviewed, relative to the anticipated 1985 technology. The scope of the control requirements

was increased to respond to all engine control requirements, including high-pressure comprcs- ,..

sor vane angle scheduling and bleed valve control. No attempt was made to retain existing

control components in order to capitalize on the economies of integration of the-various

control functions. The required control parameters are listed in Table C-II.

TABLE C-II

CONTROl. SYSTEM REQUIREMENTS FOR CONTROL

USING 1985 TECHNOLOGY

Current Engine Requirements

Control fuel flow on basis of high-pressure rotor speed, power lever angle, speed set

bias, and acceleration and deceleration limits.

Control the variable vane stagger angles.

Control fuel flow for thrust reversing, flight idle, and cold starting with enricbment.

Control one station 3 bleed valve, four station 3.5 bleed valves, and 2 fifteenth-stage

bleed valves.

Provide compressor surge waruing signal.

Fail in an operationally safe manner.

New Reqnirements

Distr_nte fuel between pilot and main zones.

Provide electronic systems failure waruing signal.

Optional Capabilities at Small Additivaal Cost

Control fuel flow with engl_ e pressure ratio limiting.

Provide engine degradation warning signal.

143

............. m
b, ld¢_ltifieation of Candidate _llemes

Candidate schemes were identified through a matrix technique similar to that used for the

modified production control. The size of the matrix was reduced, to eliminate those

schelneS that were found to be noncompetitive in the study of the modified control. Both

tile shlglc-slroke pump used for main burner manifold filling and the variable displacement

puntp were eliminated. In addition, the metered circulation system was eliminated, and

only one auxiliary circulation pump scheme was included since no important differences

were identified between the two schemes considered in tile first study. All schemes considered usedigital.electronics for all functions, since current technology growth trends inai- ,..

cate that this approach will be substautially superior by 1985.

The resulting matrix included four schemes and is shown in Table-C-IIl. Schematic diagrams

for the schemes are presented in Section B of this Appendix.

TABLE C-Ill

MATRIX FOR CANDIDATE CONTROL SCHEMES FOR

CONTROL SYSTEMS USING 1985 TECHNOI..OGY

Scheme

A B C D

Flow Distribution

Flow Distribution Valve X X X

Two Metering Valves X

Manifold Fill Technique

Fill Detector X X

Circulation X X

Pumps

One Pump X X

Two Pumps X X

e. Concept Evaluation and Selection

Since the selection of sdlemcs had included all initial screening process to eliminate concepts determined to be unpromising from tile previous study, the evaluation for the all-new

schemes consisted only of a life cycle cost estimate for each scheme. The process was idenr •

tic'.d to that used previously, with individual estimates being made for production cost,

maintenance requircmeuts, and weight.

Actual costs, maintenance history, attd weights were used where sufficient data were available, and correlation curves were used where direct information was not available. The correlation factors for tile hydronlechanical components were tile same as those used previously, us it was assumed that the 1975 hydromcchanical technology has reached a level of

ntaturitywherc no significant improvements are anticipated. For tile electronic components, however, the factors were adjusted to reflect technology growth, as shown in Figure

C-IO.

144
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PERCENTOF NUMB[ R OF INPUTSAND OUTPUTS

14gure (.310 Correlations of ReliabiliO'. If'eight, and G_st,tbr Electronic Control l;lements Show#tg

,,Inti('ipated I']ffect of Techno/o&,l. Growth

The resulting components ill the life cycle cost Ctluation for Schemes A, B, C, and D are presented in Figure C-1 I and tire conlparcd to the current JT9D production control system. The

relative life cycle costs calculated from these factors :ire presented in Figure C-12. The comparison indicates that integrated all-electronic control systems similar to these schemes will

show a potential reduction in life cycle cost of approximately 40 percent rel_itive to the current production system, when modified to control il two-sttlge Vorbix coinbustor.

"these results indicate th_it Scheme B provides the lowest lilt" cycle cost. This concept incori. porates one fuel I_onip, _1flow distril:,ution valve, lind tl_c use of circulation flow 1okeep the

Illilin fuel tll;lliii'oltl full. It is inlporlilnt to note Ihal tills sehelne incorporates tile Siilne gellCrill design features ilS the add-o, illodificiilion Sclieln,'s Ic _lnd Id selected ill the prececding study.

145 2
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1;tgure C-II Relative Estimated Production Costs, Maintenance Requirements, and Estimated Weights of

Candidate New _bntrol Schemes

146

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B. STUDY SCHEMES

Schematic diagrams for thc schemes presented in Table C-I are containcd in Figures C-13

through C-26.

147

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t

APPENDIX D

NOMENCLATURE

173

, -lli
NOMENCLATURE

:' I'1 F.mission Index _ g pollutant/kg fnel

_" F/A kg fuel/kg air ir

It Specific Itumidity _ gl120/kg dry air

LBO Lean Blowout

P Pressure ~ atnl -=',

:- PF Pattern Factor

AI' Pressure Loss ~ atm

,.. T Temperature ~ K

V Velocity _ m/s

- Vref Combustor Reference Velocity _ m/s

Wa Airflow ~ kg/s

Wf Fuel Flow _ kg/s

M Flow Parameter ~ kgv'_K/m 2 atm s

6 Density ~ kg/m 3

r/c Combustion Efficiency

¢, I:"R Equivalence Ratio

:-- SUBSCRIPTS

- b Burner

s Static Conditions

t Total Condition

4 Compressor Fxit Station (JT9D-7)

5 Turbine Inlet Station (JT9D-7)

174
pb.,

_. APPENDIXE

REFERENCES

175
REFERENCES

I. _Environnlental Protection Agency, "Control of Air Pollution From Aircraft and Aircraft

Engines; I:mission Standards and Test Procedures for Aircraft," Federal Register 38

(136) Partll: 19076, July 13, 1973.

2. Odgers, .I., and Carrier, C., "Modeling of (;as Turbine Combustors; Consideration of

Combustor I'fficiency and Stability," ASME 72 WA>-I, September 1972. ,,_.

3. Sarli, V. J., Eiler, D. C., and Marshall, R. L., "Effects of Operating Variables on Gaseous

Emissions," APCA Conference, New Orleans, Louisiana, October..26-28, 1975.

4. Code of Federal Aviation Regulations, Part 33, "Airworthiness Standards: Aircraft

Engines," Subpart E, Paragraph 33.73, Power or Thrust Response, January 1, 1974.

176